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Title: Aeroplane Construction And Operation
Author: Rathbun, John B.
Language: English
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[Illustration: Curtiss Type JN4D Biplane Used by The United States Air
Mail Service]
*AEROPLANE*
*CONSTRUCTION AND*
*OPERATION*
Including Notes On Aeroplane Design
And Aerodynamic Calculation,
Materials, Etc.
A Comprehensive Illustrated Manual of Instruction for
Aeroplane Constructors, Aviators, Aero-Mechanics,
Flight Officers and Students. Adapted
Either for Schools or Home Study.
BY
JOHN B. RATHBUN
AERONAUTICAL ENGINEER
Consulting Aeronautical Engineer, Chicago Aero Works; Chief
Engineer,
Automotive Engineering Company. Formerly Instructor in
Aviation and Machine Design, Chicago Technical College.
*CHICAGO*
*STANTON and VAN VLIET CO.*
*PUBLISHERS*
1919
Copyrighted, 1918
By STANTON AND VAN VLIET CO.
AEROPLANE CONSTRUCTION AND OPERATION
INTRODUCTION
Many aeronautical books of a purely descriptive nature have been written
for the average man, but as a rule they contain little of interest for
the more serious student of the subject. Other books of a highly
technical and mathematical class have also been published, but their
contents are all but unintelligible to anyone but a trained engineer. It
is the purpose of the author to compromise between these two extremes,
and give only that part of the theory and description that will be of
practical use for the builder and flyer. The scope of the subjects
covered in this volume has been suggested by the questions asked by
students and clients, and is the result of many years’ correspondence
with beginner aviators and amateur aeroplane builders.
I have endeavored to explain the principles of the aeroplane in simple,
concise language, starting with the most elementary ideas of flight and
finishing with the complete calculations for the surfaces, power,
weight, etc. When mathematical operations are necessary they are simple
in form, and are accompanied by practical problems worked out
numerically, so that a man with even the most elementary mathematical
knowledge will have no difficulty in applying the principle to his own
work. In cases where the calculations would necessarily be complicated,
I have substituted tables of dimensions for the mathematical operations,
these dimensions being taken from a number of representative machines.
While flying cannot be taught by books, and is only the result of actual
experience, the chapter devoted to the use of controls under different
flight conditions will be of great benefit to the prospective aviator.
The portion of the book devoted to operation will be of use in flying
schools and training camps since both training methods and control
manipulation are covered in detail. In addition I have presented
considerable data on the requirements of the modern aeronautical motor.
So many new firms are now entering the aeroplane industry that there is
an ever increasing demand for trained mechanics, designers and flyers,
and many technical men now working along other lines are taking a keen
interest in aeronautical engineering. If the contents of this book will
serve to inspire the technical reader to deeper interest and practical
research in the fascinating subject of aeronautics, the author will be
more than satisfied with the result of his labor. The aeroplane is
rapidly assuming a great commercial importance, and there is no doubt
but what it will develop into an industry rivaling that of the
automobile.
To keep fully abreast of the times in aeronautic development, one should
be a constant reader of the excellent aeronautical magazines. Too much
praise cannot be given to the aeronautical press in its effort to
maintain an interest in this subject, and as with all pioneering
movements, these magazines have met with many discouragements and
financial setbacks in the earlier days of flying. To the American
magazines, "Aerial Age" and "Flying" (New York), the author owes a debt
of gratitude for the use of several of the cuts appearing in this book.
The English magazines, "Flight," "Aeronautics" and the "Aeroplane," have
been similarly drawn on. "Aviation and Aeronautical Engineering" (New
York) has suggested the arrangement of several of the tables included
herein. All of these papers are of the greatest interest and importance
to the engineer, aviator and aero-mechanic.
JOHN B. RATHBUN.
AERONAUTICAL MAGAZINES
The following list of American and English aeronautic publications will
be of interest to those who wish to keep in touch with the latest
developments in aeronautics:
- AVIATION AND AERONAUTICAL ENGINEERING (two issues per month). A
technical magazine published by The Gardner-Moffat Co., Inc., 120 W.
32d St., New York.
- AERIAL AGE (weekly). Popular and technical. The Aerial Age Co.,
Foster Bldg., Madison Ave. and 40th St., New York.
- AIR SERVICE MAGAZINE (weekly). Military and popular subjects.
Gardner-Moffat Co., Inc., 120 W. 32d St., New York.
- FLYING (monthly). Popular and military subjects. Published by Flying
Association, Inc., 280 Madison Ave., New York.
- AIR TRAVEL (weekly). Popular subjects. Published by Air Travel, New
York.
ENGLISH MAGAZINES.
- FLIGHT AND THE AIRCRAFT ENGINEER (weekly). Technical and popular.
Published by Flight and Aircraft Engineer, 36 Great Queen St.,
Kingsway, W.C.2, London, England.
- AERONAUTICS (weekly). Technical and industrial. Published by
Aeronautics, 6-8 Bouverie St., London, E.C.4, or may be had from
1790 Broadway, New York.
- THE AEROPLANE (weekly). Technical and popular. Published by "The
Aeroplane," 166 Piccadilly, London, W.1.
AEROPLANE CONSTRUCTION AND OPERATION ..............................
INTRODUCTION ....................................................
AERONAUTICAL MAGAZINES ..........................................
ENGLISH MAGAZINES. ..............................................
CHAPTER I. PRINCIPLES OF THE AEROPLANE. .........................
CHAPTER II. TYPES OF MILITARY AEROPLANES. .......................
CHAPTER III. ELEMENTARY AERODYNAMICS ............................
CHAPTER IV. EXPERIMENTAL LABORATORIES. ..........................
CHAPTER V. AERODYNAMICS OF LIFTING SURFACES (AEROFOILS). ........
CHAPTER VI. PRACTICAL WING SECTIONS. ............................
CORRECTION FACTORS FOR WING FORM AND SIZE. ....................
CHAPTER VII. BIPLANES AND TRIPLANES. ............................
CHAPTER VIII. EFFECTS OF PLAN FORM. (TANDEM AEROPLANES.) ........
CHAPTER IX. WING CONSTRUCTION. ..................................
CHAPTER X. WING CONSTRUCTION DETAILS. ...........................
CHAPTER XI FUSELAGE (BODY) CONSTRUCTION. ........................
CHAPTER XII DETAILS OF FUSELAGE CONSTRUCTION ....................
FUSELAGE WEIGHTS. .............................................
SIZE OF LONGERONS .............................................
CHAPTER XIII. CHASSIS CONSTRUCTION. .............................
CHAPTER XIV. ESTIMATION OF WEIGHT. ..............................
WEIGHTS OF AERONAUTICAL MOTORS. ...............................
CHAPTER XV. BALANCE AND STABILITY. ..............................
CHAPTER XVI. HEAD RESISTANCE CALCULATIONS. ......................
CHAPTER XVII. POWER CALCULATIONS. ...............................
CHAPTER XVIII. PROPELLERS. ......................................
CHAPTER XIX. OPERATION AND TRAINING. ............................
CHAPTER XX. AERONAUTICAL MOTORS. ................................
CHAPTER XXI. GLOSSARY OF AERONAUTICAL WORDS. ....................
A .............................................................
B .............................................................
C .............................................................
D .............................................................
E .............................................................
F .............................................................
G .............................................................
H .............................................................
I .............................................................
J .............................................................
K .............................................................
L .............................................................
M .............................................................
N .............................................................
O .............................................................
P .............................................................
P .............................................................
S .............................................................
T .............................................................
U .............................................................
V .............................................................
W .............................................................
CHAPTER I. PRINCIPLES OF THE AEROPLANE.
*Mechanical Flight*. Although the elementary principles of mechanical
flight are not of recent origin, the practical development of the flying
machine is confined almost entirely to the present century. Gravity
propelled gliders and small models have been flown with success from a
comparatively early date, but the first actual sustained flight with a
power driven machine was performed by the Wright Brothers in 1903. There
was no single element on this first successful machine that had not been
proposed many years before by Langley, Chanute, Montgomery, Henson,
Mouillard, and others, but this first flight must be attributed
principally to the fact that the Wrights started carefully and
painstakingly to learn how to operate (By practicing with gliders)
before starting on the first power machine. If Langley had studied the
operation of his machine as carefully as he did its theory and design,
he would have been flying long before the Wrights as his original
machine was afterwards successfully flown by Curtiss.
When once actual flight was achieved, and the success of the Wright
Brothers became generally known, the development proceeded with leaps
and bounds. All the resources of science and engineering skill were at
once applied to the new device until our present scientific knowledge of
the aeroplane compares very favorably with the older engineering
sciences. In the few years that have elapsed since the first flight, the
aeroplane holds all records for speed, endurance, and radius of action.
A great deal of the success so rapidly acquired can be credited to the
automobile and motorcycle industries, since it was the development of
the light internal combustion motors used on these machines that paved
the way for the still lighter aeronautic motor. Again, the automobile
industry was responsible for the light and powerful materials of
construction, such as alloy steel, aluminum alloys, and also for the
highly important constructional details, such as ball bearings,
pneumatic tires, carburetors, magnetos, steel tubing, etc. The special
methods developed in automobile work have helped to make the aeroplane
an immediate commercial proposition.
[Illustration: Curtiss Type JN4-B Primary Trainer]
*Types of Flying Machines*. In general, flight apparatus may be divided
into two classes, (1) The Lighter Than Air Type, such as the balloon and
dirigible, and (2) The Heavier Than Air Machine, represented by the
aeroplane, helicopter and ornithopter. The lighter than air machine is
supported in flight by "buoyancy" in much the same manner that a piece
of wood floats in water. When a balloon or dirigible, because of its
large volume, displaces a volume of air equal to its own weight, the
device will float. When the weight of air displaced exceeds the weight
of the balloon or dirigible, it will continue to rise until it reaches
an altitude where the diminished air density again results in an
equality between the weight of the device and the air displaced. At this
point it rests, or is in equilibrium. The flotation of such a device is
entirely due to static forces and hence (1) is often called an
"aerostat."
The sustenation of a Heavier Than Air Machine is due to an entirely
different application of forces. Forces in motion (Dynamic Forces) are
essential to the support of a heavier than air machine, and it is the
resultant of these forces that performs the actual lifting operation,
this resultant corresponding to the buoyant force of the aerostat.
"Dynamic" flight is obtained by an apparatus in which an arrangement of
surfaces are moved in such a way as to cause an upward component of the
forces generated by the impact of the air on the surfaces. The surfaces
drive the air down and when the force necessary for the continuous
downward deflection of air becomes equal to the weight of the machine it
is sustained in flight. Dynamic flight therefore depends on the
continuous downward deflection of masses of air, and when this motion
ceases, sustentation also ceases.
An aeroplane is provided with a deflecting surface that is fixed rigidly
in regard to the body of the machine, and the motion necessary for its
support is provided by driving the machine forward, the forward motion
being produced by the horizontal pull of air screws or propellers. It is
at once evident that the forward horizontal motion of the aeroplane must
be maintained for its support, for the surfaces are fixed and there is
no other possible way of producing a relative motion between the wings
and the air.
To overcome the objection of forward motion, several other machines have
been proposed in which the surfaces are moved in relation to the body,
as well as the air, thus making it possible for the device to stand
stationary while the revolving or reciprocating surfaces still continue
in motion in regard to the air. One type of the moving surface machine,
the "Helicopter," is provided with revolving surfaces arranged in the
form of vertical air screws or propellers, the blades of the propellers
being inclined so that they drive down a continuous stream of air and
produce the continuous upward reaction that supports the machine. While
such machines have succeeded in raising themselves off the ground they
are not yet practical flying devices. The "ornithopter" or "orthopter"
is a flapping wing machine that maintains flight after the manner of the
bird (Ornis). Like the helicopter, the ornithopter has not yet proved
successful.
[Illustration: Fig. 1. Comparison Between the Kite and Aeroplane; Fig.
2, Showing the Lift and Drag Forces Produced by the Air Stream. The
Propeller (P), Acts in a Manner Similar to the Kite String (S) in
Producing Relative Motion Between the Air and the Lifting Surfaces.]
*Principles of the Aeroplane*. In its elementary principles, the
aeroplane can be compared with the kite, as both are supported by the
impact of a horizontal stream of air. In Diagram 1, the kite surface is
indicated by X-X with the relative air stream W-W-W-W moving from left
to the right as indicated by the arrow heads. On striking the surface,
the air stream is deflected vertically, and in a downward direction, as
shown by the streams lines R-R-R-R. The reaction of the air deflection
produces the lift shown vertically and upwards by the arrow L. The kite
surface is held against the impact of the air stream by the string S so
that there is relative motion between the air and the kite, and so that
the surface will not be carried along with the air current toward the
right. If the kite were allowed to drift with the wind there could be no
relative motion between the surface and the air stream, hence the kite
would fall as soon as it attained the velocity of the wind. The
horizontal force exerted by the wind tending to carry the kite toward
the right is indicated by the arrow D and is known as the "drag" or
"drift" force. There are thus three forces, the lift (L), the drag (D),
and the resultant of the two forces indicated by the string (S). The
forces of lift and drag are nearly at right angles to one another. The
kite tail T is simply a stabilizing device whose purpose is to maintain
a constant angle between the surface and the wind and it performs an
almost negligible amount of lift.
A few more words in regard to the "relative velocity" between the
surface and wind. In the figure, the kite is assumed as being
stationary, while the wind moves from left to right. With a thirty mile
per hour wind, the relative air velocity in regard to the surface would
be 30 M. P. H. If the air particles are now considered stationary, and
if the kite is towed toward the left (opposite to figure) at 30 miles
per hour, the relative velocity between the surface and air would still
be 30 M. P. H. In other words, the kite may be stationary, or may move
in regard to the earth, but its lift is unaffected as long as the
relative motion between the surface and air remains constant. The motion
between an aeroplane and the earth depends upon the difference of the
aeroplane and wind velocities. For example, a aeroplane with a relative
speed of 60 miles per hour, flying against a headwind of 30 miles per
hour, moves 60-30 = 30 miles per hour in regard to the earth. The same
aeroplane flying with the above wind would have a velocity of 60+30=90
miles per hour past a fixed point on the earth’s surface, yet in both
cases, the relative velocity of the surface in regard to the air would
be the same.
Fig. 2 is a diagram of an aeroplane that shows the connection between
the kite and aeroplane principles. In this figure, the wing surface of
the aeroplane, X-X corresponds to the kite surface X-X. The relative air
W-WW-W striking the wing from the left is deflected down along the
arrows R-R-R-R and results in an equivalent lift force L, and a drag
force D as in the case of the kite. The resultant force required to
maintain the relative velocity between the air and wings is indicated by
D¹, opposite and equal to the drag force D. The resultant required for
overcoming the drag is provided by the screw propeller P instead of the
string S shown in Fig. 1. The propeller thrust (D¹) is parallel to the
air stream instead of being inclined as in the case of the string, but
the total effect is the same since both are "Resultants of the lift and
drag." To sustain the aeroplane, the lift (L) must be equal and opposite
to the weight shown by M. The fact that M and L are opposite and equal
makes it only necessary for the propeller to overcome the horizontal
drag, and hence the thrust can be made parallel to the air flow—or
nearly so. The aeroplane is provided with a small tail surface (T) that
corresponds to the kite tail (T). It maintains the lifting surfaces X-X
at a given angle with the air stream. The tail may, or may not aid in
supporting the machine, but in modern machines it is common to employ a
tail surface that is non-lifting under ordinary conditions of normal
flight. The body (B) contains the pilot, motive power, fuel, and such
useful load as it may be necessary to carry.
[Illustration: Fig. 3. Caudron Monoplane. Side Elevation.]
Fig. 3 shows a Caudron monoplane in side elevation. This view
illustrates the application of the principles shown by Fig. 2, except
for the vertical rudder at the rear. The latter is used for steering in
a horizontal direction. Fig. 4 shows the construction even more clearly
since it is a perspective view. The machine is a Morane "Parasol"
monoplane with the wing placed over the body. This location of the main
lifting surface is for the purpose of improving the view of the pilot
and in no way affects the principles just described. The wires shown
above the wing are bracing stays. The tail is hinged near the rear so
that the angle of the rear portion can be changed (Elevator flaps), and
permits the angle of the main wings to be altered in regard to the air
stream, thus causing the machine to ascend or descend. The tail also
damps out oscillations or vibrations due to irregularities in the air
current. The wheels and running gear (Chassis) allow the machine to be
run over the ground until the relative air speed is sufficient to
support the machine in flight.
[Illustration: Fig. 4. Morane Umbrella Type Monoplane. Courtesy of
"Flight."]
*Multiplanes*. In order to support a heavy load, and at the same time
have a small compact machine, it is necessary to have more than one
"layer" of wing surface. It is evident that the wing length or "span"
can be made much less than that of the monoplane surface shown, if the
total area could be divided into two or more parts. A machine having its
main lifting surface divided into two or more separate sections is known
as a "multiplane," this term becoming "Biplane," "Triplane," or
"Quadraplane," depending on whether there are two, three or four
independent lifting surfaces. There is almost a limitless variety of
arrangements possible, but the most common arrangement by far is that of
the biplane, in which there are two superposed surfaces as shown by Fig.
5. In this type, the two lifting surfaces are placed over one another
with a considerable "gap" or space between. The body is placed between
the wings and the tail surfaces and chassis remain the same as in the
monoplane. This is known as a "Tractor" biplane since the propeller is
in front and pulls the machine along while Fig.6 shows a "Pusher" type
biplane in which the propeller is mounted behind the wings and therefore
pushes the machine.
[Illustration: Fig. 4-A. Deperdussin Monoplane with Monocoque Body.
Gordon-Bennett Racer.]
*Biplanes*. Besides the advantages of size, the biplane has a number of
other good features. The deep spacing of the upper and lower surfaces
permits of a powerful and light system of trussing being placed in the
gap, and therefore the biplane can be made stronger (weight for weight)
than the monoplane in which no such trussing can be economically
applied. The vertical "struts" of the bracing can be clearly seen in the
figure. The efficiency of this interplane trussing greatly increases the
possible size and capacity of the aeroplane. With monoplanes it is
seldom possible to exceed a wing span of 36 feet without running into
almost unsurmountable structural difficulties. The weight of the large
monoplane also increases is leaps and bounds when this critical span is
once exceeded. To maintain an equal degree of strength the monoplane
requires very careful attention in regard to the design and
construction, and is correspondingly more expensive and difficult to
build than the biplane, although the latter has by far the greater
number of parts. By suitable arrangements in the location of the biplane
surfaces a very fair degree of stability can be obtained, an advantage
which is impossible with the monoplane.
[Illustration: Fig. 5. S. P. A. D. Tractor Biplane Speed Scout.]
A distinct disadvantage of the two superposed surfaces of the biplane is
due to the fact that there is "interference" between the upper and lower
wings, and that the lift for equal areas is less than in the case of the
monoplane. With a given form of wing, 100 square feet of monoplane
surface will lift considerably more than the same area applied in
biplane form. The amount of the "drag" for the support of a given load
is increased, and with it the amount of power required. The greater the
separation or "gap" between the wings, the greater will be the lift, but
when the gap is unduly increased to obtain a great lift the length of
the interplane bracing is increased to such an extent that the
resistance of the bracing will more than overcome the advantages due to
the large gap. There is a fixed limit to the gap beyond which it is not
practical to go. The bracing has a very material effect on the air
resistance, no matter how small the gap.
[Illustration: Fig. 6. Pusher Type Biplane in Which the Propeller Is
Placed Behind the Wings.]
*Triplanes*. Of late the triplane has been rapidly increasing in use,
and in certain respects has many advantages over either the monoplane or
biplane. This type has three superposed surfaces which still further
diminishes the size for a given area. The interference between the
surfaces is even greater than with the biplane, and hence the lift is
less for a given area and aspect ratio. This latter defect is partly, or
wholly overcome by the possibility of using long narrow wings, and
because of the reduced span there is a corresponding reduction in the
bracing resistance. It should be noted at this point that the efficiency
of a lifting surface is greatly increased when the ratio of the length
to the width is increased, that is, a long, narrow wing will be more
efficient than a short, wide shape. The relation of the length to the
width is called "aspect ratio," and will be described in more detail in
a following chapter.
[Illustration: Fig. 6-A. Farman Type Pusher Biplane.... Note the
Propeller At the Rear of Body, and the Position of the Pilot and
Passenger.]
[Illustration: Fig. 6-B. The Mann Two-Propeller Pusher Biplane. The
Propellers Are Mounted on Either Side of the Body, and Are Driven by a
Single Motor Through a Chain Transmission. This Drive Is Similar to the
Early Wright Machines.]
Fig. 7 is a sketch of a Sopwith Triplane Scout and shows clearly the
three superposed wings. The small amount of interplane bracing, and the
great aspect ratio, makes this type very suitable for high speed. The
body, tail and chassis arrangements are practically the same as those of
a biplane. The Curtiss Triplane Scout is the pioneer of this type of
machine, although experimental work on the triplane had been performed
in England by A. V. Roe many years ago. The Roe triplane was lightly
powered and for its time was successful in a way, but the Curtiss is the
first to enter into active competition against the biplane scout. Owing
to the small span required for a given area, and the possibilities of
very light and simple bracing, the triplane is an ideal type for heavy
duty machines of the "bombing" species. Enormous triplanes have been
made that are capable of a useful load running up into the tons, the
large Curtiss and Caproni’s being notable examples. As the triplane is
much higher than the biplane of equal area, the interplane bracing is
deeper and more effective without causing proportionately higher
resistance.
[Illustration: Fig. 7. Sopwith Triplane Speed Scout.]
*Quadraplane*. The use of four superposed surfaces has not been
extended, there probably being only one or two of these machines that
can be said to be successful. The small "quad" built by Matthew B.
Sellers is probably the best known. The power required to maintain this
machine in flight was surprisingly small, the machine getting off the
ground with a 4 horsepower motor, although an 8 horsepower was
afterwards installed to maintain continuous flight. The empty weight was
110 pounds with the 8 horsepower motor. The span of the wings was 18’ 0"
and the width or "chord" 3’ 0", giving a total area of about 200 square
feet.
[Illustration: Fig. 7-A. Curtiss Triplane Speed Scout. Courtesy "Aerial
Age."]
*Tandem Aeroplanes*. A tandem aeroplane may be described as being one in
which the surfaces are arranged "fore and aft." The Langley "Aerodrome"
was of the tandem monoplane type and consisted of two sets of monoplane
wings arranged in tandem. This pioneer machine is shown in Fig. 8, and
is the first power driven model to achieve a continuous flight of any
length. Instead of two monoplane surfaces, two biplane units or triplane
units can be arranged fore and aft in the same manner.
While there have been a number of tandem machines built, they have not
come into extensive use. Successful flight was obtained with a full size
Langley Aerodrome, and this machine flew with a fair degree of
stability. The failure of other tandem machines to make good was due, in
the writer’s opinion, to poor construction and design rather than to a
failure of the tandem principle. The Montgomery glider, famed for its
stability, was a tandem type but the machine was never successfully
built as a powered machine.
The wings must be separated by a sufficient distance so that the rear
set will not be greatly influenced by the downward trend of the air
caused by the leading wings. As the rear surfaces always work on
disturbed air they should be changed in angle, increased in area, or be
equipped with a different wing curvature if they are to carry an equal
proportion of the load. Usually, however, the areas of the front and
rear wings are equal, and the difference in lift is made by changes in
the wing form or angle at which they are set. In some cases the wings
are approximately the same, the difference in lift being compensated for
by moving the load further forward, thus throwing more of the weight on
the front wings.
[Illustration: Fig. 8. Langley’s "Aerodrome," An Early Type of Tandem
Monoplane.]
*The Aeroplane in Flight*. Up to the present we have only considered
horizontal flight at a continuous speed. In actual flight the altitude
is frequently varied and the speed is changed to meet different
conditions. Again, the load is not an absolutely constant quantity owing
to variations in the weight of passengers, and variations in the weight
of fuel, the weight of the latter diminishing directly with the length
of the time of flight. To meet these variations, the lift of the wings
must be altered to suit the loading and speed—generally by altering the
angle of the wings made with the line of flight.
Fig.9 shows an aeroplane in horizontal flight and lightly loaded, the
machine traveling along the horizontal flight path F-F. With the light
load, the angle made by the wings with the flight path is shown by (i),
the tail and body remaining horizontal, or parallel to the flight path.
With an increased load it is necessary to increase the angle of the
wings with the flight line, since within certain limits the lift
increases with an increase in the angle of incidence (i). Fig. 10 shows
the adjustment for a heavier load (W₂), the angle of incidence being
increased to (i’), and the body is turned down through a corresponding
angle. The increased angle is obtained by turning the elevator flaps (T)
up, thus causing a downward force (t) on the tail. The force (t) acts
through the body as a lever arm, and turns the machine into its new
position. It will be noted that when the angle of incidence is great
that the rear of the body drags down and causes a heavy resistance. This
position of a dragging tail is known to the French as flying "Cabré."
With high angles cabré flight is dangerous, for should the propeller
thrust cease for an instant the machine would be likely to "tail dive"
before the pilot could regain control. This sort of flight is also
wasteful of power. Cabré flight is unnecessary in a variable incidence
machine, the wing being turned to the required angle independently of
the body, so that the body follows the flight line in a horizontal
position, no matter what the angle of incidence may be. In this type of
machine the wings are pivoted to the body, and are operated by some form
of manual control.
[Illustration: Figs. 9-10-11-12. Showing the Use of Elevators in
Changing Angle of Incidence. - Machine Shown in Four Principal Attitudes
of Flight. As the Body and Wings Are in a Single Unit, the Body Must Be
Turned for Each Different Wing Angle.]
In Fig. 11, the large angle (i’) is still maintained, but the load is
reduced to the value given in Fig. 9. With an equal load, an increased
angle of incidence causes the machine to climb, as along the new flight
line f-f. With the load (W) equal to that in Fig. 9, the angle of
incidence will still be (i) but this will be along a new flight line if
the large angle (i’) is maintained with the horizontal as shown by Fig.
11. With the wings making an angle of (i’) with the horizontal, and
angle of incidence (i) with the flight line, it is evident from Fig. 11
that the new flight line f-f must make an angle (c) with the original
horizontal flight line F-F. This shows how an increased angle with a
constant load causes climbing, providing, of course, that the speed and
power are maintained. With a given wing and load there is a definite
angle of incidence if the speed is kept constant. Should a load be
dropped, such as a bomb, with the wing angle kept constant, the new path
of travel will be changed from F-F to f-f.
Fig. 12 shows the condition when the rear end of the body is elevated by
depressing the elevator flap T. This occasions an upward tail force that
turns the wings down through the total angle (i’). With the former
loading and speed, the angle of incidence is still (i) degrees with the
new flight path f-f, the new flight path being at an angle (c) with the
horizontal F-F. The body is turned through angle (i’), but the angle (i)
with the flight path f-f is still constant with equal loads and speeds.
To cause an aeroplane to climb, or to carry a heavier load, the elevator
"flap" is pointed up. To descend, or care for a lighter load, the
elevator is turned down. In normal horizontal flight the machine should
be balanced so that the tail is horizontal and thus creates no drag.
When the elevator *must* be used to keep the tail up in horizontal
flight, the machine is said to be "tail heavy."
*Longitudinal Stability*. In Figs. 9-10-11-12 the machine was assumed to
be flying in still air, the attitudes of the machine being simply due to
changes in the loading or to a change in altitude. The actual case is
more complicated than this, for the reason that the machine is never
operating in still air but encounters sudden gusts, whorls, and other
erratic variations in the density and velocity of the air. Each
variation in the surrounding air causes a change in the lift of the
wings, or in the effect of the tail surfaces, and hence tends to upset
the machine. If such wind gusts would always strike the wings, body, and
tail simultaneously, there would be no trouble, but, unfortunately, the
air gust strikes one portion of the machine and an appreciable length of
time elapses before it travels far enough to strike another. Though this
may seem to be a small fraction of time, it is in reality of sufficient
duration to have a material effect on the poise of the aeroplane.
Vertical gusts due to the wind passing over buildings, hills, cliffs,
etc., not only tend to upset the machine, but also tend to change the
altitude since the machine rises with an up gust and sinks with a down
trend in the Stream.
Assume a machine as in Fig. 9 to be traveling steadily along a
horizontal path in still air. A sudden horizontal gust now strikes the
machine from the front, thus causing a sudden lift in the main wings. As
this gust strikes the wings before the tail, the tail will stand at the
old altitude while the wings are lifted, thus giving the position shown
by Fig. 10. After passing over the wings it lifts the tail, this effect
probably not being sufficient to restore the wing and the tail to their
old relative attitude since the gust generally loses velocity after
passing the wings. A head gust of this type often strikes the front
wings diagonally so that it never reaches the tail at all. To remedy
this upsetting action of the gust, the pilot must move his rear elevator
so that the elevator is in the position shown by Fig. 12, that is, the
flap must be turned down so as to raise the tail.
A gust striking from behind may, or may not affect the elevator flaps,
this depending on their position at the time that the gust strikes. If
the flaps are turned up, the rear end will be raised by the gust and the
machine will head dive: if turned down, the gust will depress the tail,
raise the head and tend to "stall" the machine. If the tail is of the
lifting type, the rear entering gust will reduce the relative velocity,
and the lift, and cause the tail to drop. On passing over the tail and
striking the wings, the rear gust will reduce the velocity and cause a
loss in lift. This will either cause the machine to head dive or drop
vertically through a certain distance until it again assumes its normal
velocity.
All of these variations cause a continually fore and aft upsetting
movement that must be continually corrected by raising and lowering the
elevator flaps, and in very gusty weather this is a very tedious and
wearing job. It requires the continual attention of the pilot unless the
action is performed automatically by some mechanical device, such as the
Sperry Gyroscopic, or else by some arrangement of the surfaces that give
"inherent" stability. Control by means of the elevator flaps (which
raise and lower the body in a fore and aft direction, as shown) is known
as "longitudinal control," and when the machine is so built that
correction for the longitudinal attitude is obtained "inherently," the
machine is said to be "longitudinally stable." Modern machines can be
made very nearly longitudinally stable, and are comparatively unaffected
by any than the heaviest gusts.
*Lateral Stability*. The gusts also affect the side to side, or
"lateral" balance by causing a difference in lift on either end of the
wings. Should the gust strike one tip before the other, or should it
strike one tip harder than the other, the tendency will be to turn the
machine over sidewise. This is a more difficult problem to solve than
the longitudinal moment, although perfect inherent stability has been
attained in one or two machines without the use of additional automatic
control mechanism. Inherent lateral stability has always been attended
by a considerable loss in the efficiency of the aeroplane and speed due
to the peculiar arrangements in the main lifting surfaces. At present we
must make a decision between efficiency and stability, for one feature
must be attained at a sacrifice in the other. Contrary to the general
opinion, perfect stability is not desirable, for almost invariably it
affects the control of a machine and makes it difficult to maneuver.
Should the stability appliances be arranged so that they can be cut out
of action at will, as in the case of the Sperry Gyroscopic Stabilizer,
they will fulfill the needs of the aviator much more fully than those of
the fixed inherent type. The first thoroughly stable machine, both
longitudinally and laterally, was that designed by Lieutenant Dunne, and
this obtained its distinctive feature by a very peculiar arrangement of
the wing surfaces. It was excessively stable, and as with all very
stable machines, was difficult to steer in a straight line in windy
weather, and was correspondingly difficult to land.
[Illustration: Fig. 12.A. Diagram of the Tractor Biplane]
The first machine of the ordinary biplane type that proved inherently
stable was the R. E.-1 designed in England by Edward Busk. This machine
was flown from Farnborough to Salisbury Plain, and during this flight
the only control touched was the vertical rudder used in steering. Since
then, all English machines have been made at least partially stable, the
degree depending upon the service for which it was intended. It has been
found that in fighting, a very controllable machine is necessary, hence
stability must be sacrificed, or the control surfaces must be made
sufficiently powerful to overcome the stable tendency of the machine.
War machines are made to be just comfortably stable over the range of
ordinary flight speeds, and with controls powerful enough so that the
inherent stability can be overcome when maneuvering in battle. The
present war machine always contains an element of danger for the
unskilled pilot.
*Dihedral Angle*. This was the first lateral stability arrangement to be
applied to an aeroplane, but is only effective in still air. In rough
weather its general tendency is to destroy stability by allowing
dangerous oscillations to take place. Fig. 13 is a front view of a
monoplane in which the wings (w) and (w’) are set at an angle (d), this
angle being known as the "dihedral angle." The dotted line (m-m) shows
the line of a pair of perfectly horizontal wings and aids in
illustrating the dihedral. Assuming the center of lift at CL on the
wings, it will be seen that an increase in the dihedral raises the
center of lift above the center of gravity line C. G. by the amount (h).
With the center of gravity below the center of lift it is evident that
the weight of the machine would tend to keep it on a level keel,
although the same effect could, of course, be attained in another way.
The principal righting effect of the dihedral is shown by Fig. 14 in
which the wings (w) and (w’) are set as before. The machine is tipped or
"listed" toward the left (seen from aviator’s seat) so that wing (w’) is
down. By drawing vertical lines down until they intersect the horizontal
line X-X (the line of equilibrium), it will be seen that wing (w’)
presents more horizontal lift surface than (w) since the projected or
effective wing length (C) is greater than (b). Since (w’) presents the
greater surface, the lift (L) tends to restore the machine to its
original level position. If the wings were both set on the same straight
line, the projected lengths (b) and (c) would be the same and there
would be no restoring effect.
The dihedral would be very effective in still air, but in turbulent air,
and with the body swinging back and forth, the dihedral would act in the
nature of a circular guiding path, and thus tend to allow the swinging
to persist or increase rather than to damp it down, as would be the case
with level straight wings. Again, with the wing bent up at a
considerable angle, a side gust as at (S) would tend to throw the
machine still further over, and thus increase rather than diminish the
difficulty. In practical machines, the dihedral is usually made very
small (d = 176 degrees), the angle of each wing with the horizontal
being about 2 degrees, or even less. I think the advantage of such a
small angle is rather more imaginary than actual, and at present the
greater number of war machines have no dihedral at all. In the older
monoplanes the angle was very pronounced.
Fig. 15 shows the dihedral applied both to the upper wing (U) and the
lower wing (L), the usual method of applying dihedral to large biplanes.
Fig. 16 shows the method of applying the dihedral to small, fast
machines, such as speed scouts, the dihedral in this case being used
only on the lower wing. The dihedral on the bottom wing is usually for
the purpose of clearing the wing tips when turning on the ground rather
than for stability. A lower wing with a dihedral is less likely to
strike the ground or to become fouled when it encounters a side gust in
landing or "getting off." The use of straight upper wings makes the
construction much simpler, especially on the small machines where it is
possible to make the wing in one continuous length.
*Ailerons and Wing Warping*. Since the dihedral is not effective in
producing lateral stability, some other method must be used that is
powerful enough to overcome both the upsetting movements and the lateral
oscillations caused by the pendulum effect of a low center of gravity.
In the ordinary type of aeroplane this righting effect is performed by
movable surfaces that increase the lift on the lower wing tip, and
decrease the lift on the high side. In Some cases the lateral control
surfaces are separate from the wing proper (Ailerons), and in some the
tip of the wing is twisted or "warped" so as to produce the same effect.
These control surfaces may be operated manually by the pilot or by some
type of mechanism, such as the gyroscope, although the former is the
method most used. The lateral control, or side to side balancing of an
aeroplane, can be compared to the side to side balancing of a bicycle in
which the unbalance is continually, being corrected by the movement of
the handle bars.
Fig. 17 shows the control surfaces or "Ailerons" (A-A’), mounted near
the tips, and at the rear edge of the wing W. As shown, they are cut
into and hinged to the main wings so that they are free to move up and
down through a total angle of about 60 degrees. In a biplane they may be
fitted to the upper wing alone or to both top and bottom wings,
according to conditions. For simplicity we will consider only the
monoplane in the present instance.
In Fig. 18, a front view of the monoplane, the machine is shown "heeled
over" so that the wing tip (w) is low. To correct this displacement, the
aileron (A) on the low side, is pulled down and the aileron (A’) on the
opposite end is pulled up. This, of course, increases the lift on the
low end (w) and decreases the lift on the high side (w’). The righting
forces exerted are shown by L-L’. The increased angles made by A-A’ with
the wind stream affects the forces acting on the wings, although in
opposite directions, causing a left hand rotation of the whole machine.
In Fig. 19, conditions are normal with the machine on an even keel and
with both ailerons brought back to a point where they are level with the
surface of the wing, or in "neutral." Fig. 20 shows the machine canted
in the opposite direction with (w’) low and (w) high. This is corrected
by bringing down aileron (A’) and raising (A), the forces L-L’ showing
the rotation direction. By alternately raising and lowering the ailerons
we can correspondingly raise or lower the wing tips. It should be noted
here that in some machines the ailerons are only single acting, that is,
the aileron on the low side can be pulled down to increase the lift, but
the opposite aileron remains in the plane of the wings, and does not
tend to "push down" the high side. Since all of the aileron resistance
in a horizontal direction is now confined to the low side, it turns the
machine from its path, the high wing swinging about the lower tip with
the latter as a pivot. In the double acting control as shown in Figs. 17
to 23, the resistance is nearly equal at both tips and hence there is no
tendency to disturb the flight direction. With single acting ailerons,
the directional disturbance is corrected with the rudder so that when
the aileron is pulled down it is necessary to set the rudder to oppose
the turn. On early Wright machines the rudder and lateral controls were
interconnected so that the rudder automatically responded to the action
of the ailerons.
Fig. 21 is a detailed front elevation of the machine and shows the
control wheel (C) and cable connections between the wheel and the
ailerons A-A’. When the wheel C is turned in the direction of the arrow
K, the aileron A’ is pulled down by the flexible cable (i), and a
corresponding amount of cable (h) is paid off the wheel to the rising
aileron A. Aileron A is pulled up by the connecting cable (e) which is
attached to A’ at one end and to A at the other. Pulleys (f) and (g)
guide the interconnection cable. On turning the wheel in the opposite
direction, aileron A is pulled down and A’ is elevated. In flight,
especially in rough weather, there is almost continuous operation of the
control wheel. Figs. 22 and 23 are sections taken through the wing W and
the ailerons, showing the method of hinging and travel. Fig. 22 shows
the aileron depressed for raising the wing in the direction of L, while
Fig. 23 shows the aileron lifted to lower the wing. In normal flight,
with the machine level, the aileron forms a part of the wing outline (in
neutral position).
In the original Wright aeroplane, and in the majority of monoplanes, no
ailerons are used, the rear of main wing tip being bent down bodily to
increase the lift. This is known as "wing warping," and is practically a
single acting process since the depressing force on the high tip is
seldom as effective as the lift on the low. Warping is not generally
used on modern biplanes since it is impossible to maintain a strong
rigid structure with flexible tips. The control warping and twisting of
such wings soon loosens them up and destroys what remaining strength
they may have had.
[Illustration: Figs. 17-23. Showing Use of Ailerons in Maintaining
Lateral Balance.]
*Banking and Turning*. In making a sharp turn the outer wing tip must be
elevated to prevent slipping sidewise through the effects of the
centrifugal force (side slip). This is known as "banking." The faster
and sharper the turn, the steeper must be the "bank," or the angle of
the wings, until in some cases of "stunt" flying the wings stand almost
straight up and down. Should the bank be too steep there will be an
equal tendency to slip down, and inwardly, since the end resistance
against side slip is very slight. Some aeroplanes assume the correct
angle of bank automatically without attention from the pilot since the
extra lift due to the rapid motion of the outer tip causes it to rise.
On other machines the natural banking effort of the machine is not
sufficient, and this must be increased by pulling down the aileron on
the outer wing tip. Machines that have a tendency to "over-bank" must
have the ailerons applied in the reverse direction so as to depress the
outer tip. In cases of under, or overbanking machines it formerly
required experience and judgment on the part of the pilot to obtain the
correct banking angle. There are now "banking indicators" on the market
that show whether the machine is correctly banked or is side-slipping.
[Illustration: Fig. 24. A Deperdussin Monoplane Banking Around a Sharp
Turn at High Speed. Note the Elevation of the Outer Wing Tip and the
Angle Made with the Horizontal by the Wings. Speed, 105 Miles Per Hour.]
[Illustration: Standard "H-3" Training Biplane.]
CHAPTER II. TYPES OF MILITARY AEROPLANES.
*Divisions of Service*. In the army and navy, aeroplanes are used both
for offensive and defensive operations. They must carry out their own
work and intentions and prevent hostile craft from carrying out theirs.
In offensive operations the machines fly continuously over the enemy’s
country and attack every hostile craft sighted, thus creating a danger
zone within the enemy’s lines where no opposing machine can work without
being threatened with an overwhelming attack. The offensive also
includes bombing operations and the destruction of supply depots and
transportation centers. Defensive aerial operation consists in driving
out the enemy craft from our own lines, and in protecting working
machines when on photographing or observation trips. With a powerful
offensive there is of course little need for defense. The former method
is a costly one, and is productive of severe material and personal
losses.
At the present time there are eight principle functions performed by
military aeroplanes:
1. Offensive operations against enemy machines.
2. Reconnaissance, observation, special missions.
3. Bombing supply centers, railways, etc.
4. Photography.
5. Spotting for the artillery.
6. Signalling for infantry operations.
7. Submarine hunting.
8. Patrol and barrage.
[Illustration: Fig. 1. Curtiss "Baby". Biplane Speed Scout. Equipped
with 100 Horsepower Water Cooled Motor.]
Probably the most important service of all is performed by machines
under heading (1). If a successful offensive can be maintained over the
enemy’s lines he is unable to intelligently direct his artillery fire,
and can obtain no information regarding reinforcements, or troop
concentrations for an impending attack. With fighting aeroplanes
clearing the way for our own observation machines and artillery
spotters, the enemy is not only blinded, but is blocked in any attempt
to attack or concentrate his forces. The fact that the French aerial
offensive at Verdun was so efficiently and well maintained accounts for
the failure of the heavy German artillery. Driven far back over their
own lines, the German aviators were seldom able to observe the placing
of the shells, and as a result their gunners were practically trusting
to luck in reaching their target. An immediate and accurate bombardment
always followed one of the very infrequent German air raids over the
French lines. Whenever the French, partially abandoned their aerial
offensive in favor of a defensive campaign, they soon lost their mastery
of the air. As long as enemy machines can be kept back of their own
lines, new trench systems can be constructed, transportation lines can
be extended and ammunition dumps arranged, undertakings that would be
highly precarious with enemy observation machines continually passing
overhead.
[Illustration: Fig. 1. Italian "Pomilio" Two Seater Biplane. Courtesy
"_Flying_."]
To maintain an effective offensive places a tremendous strain on both
the men and the machines, for though the aeroplanes do not penetrate far
beyond the lines they usually meet with superior numbers, and in
addition are continually in range of the anti-aircraft guns. In an
attack over hostile country a slight mishap may cause the loss of a
’plane, for usually the distance from its base is so great as to prevent
a gliding return. Over its own lines an engine failure is usually only a
temporary inconvenience. Fighting aeroplanes, for the offensive, are
small high powered machines generally of the single seater type, and are
capable of high horizontal and climbing speeds. The armament consists of
a machine gun of the Lewis type, and occasionally a few light bombs may
be included in the equipment. As they do not carry out operations far to
the rear of the enemy’s lines they are provided with fuel for only two
or three hours, and this reduced fuel load is necessary for the high
speeds that must necessarily be attained. The area is limited to permit
of quick maneuvers in attack and escape, and at the same time to reduce
the head resistance and weight. The horizontal speed may run up to 150
miles per hour, with a climbing velocity that may exceed 1,000 feet per
minute. Such machines are variously known as "Speed Scouts," "Chasers,"
or "Pursuit Type" (French "DeChasse"). At the beginning of the war the
chasers were largely of the monoplane type, but at present the biplane
is in almost exclusive use.
[Illustration: Fig. 2. Machine Gun Mounting on Morane Monoplane. Gun
Fires Directly Through the Propeller Disc. The Deflecting Plate Attached
to the Root of the Propeller Blade Protects the Propeller When in Line
of Fire. Ammunition in This Gun Is Furnished in Straight Strips or
"Clips."]
The aeroplane employed for surveys of the enemy country and battle front
(2) are of an entirely different type and are much larger and slower.
These "Reconnaissance" machines are generally of the two-seater type,
the personnel consisting of an observer and the pilot, although in some
cases a third man is carried as an assistant to the observer, or to
handle a machine gun against an attack. Since their speed is
comparatively low, they are generally provided with an escort of
chasers, especially when employed on distant missions, this escort
repelling attacks while the observations are being made.
[Illustration: Fig. 2-a. Machine Gun Mounting on S. P. A. D. Biplane.
Gun Is Rigidly Attached to Fuselage Top in Front of Pilot.]
For accurate observation and mapping, the speed of an observation
machine must be necessarily low, and as they are additionally burdened
with a wireless set, an observer, a large fuel reserve, and other
impedimenta, they have a comparatively great area and are therefore
lacking in the maneuvering qualities of the chaser. The span will
average about 40 feet, and the weight carried per horsepower is greatly
in excess of that of the chaser. From a number of examples, the
reconnaissance type will average from 16 to 18 pounds per horsepower,
while the loading of the scout is from 8 to 12. This means that the
former has comparatively little reserve power for rapid climbing. The
present reconnaissance type is always armed, and must not be confused
with the early machine by that name, which, in fact, was merely an
enlarged training machine and had neither offensive nor defensive
powers. The Observer acts as gunner, and is located at a point where he
has the greatest possible range of vision, and where the angle of fire
is as little obstructed as possible.
The radius of action, or the distance traveled per tank of fuel, is
greater with the reconnaissance than with the chaser, present machines
having a capacity of from 10 to 12 hours on a single filling at normal
flight speed.
In bombing operations (3), the loading is very heavy and consequently a
"Bomber" must be a weight lifter to the exclusion of all other
qualities. Not only is the bomb load requirement severe, but the fuel
load is also of great importance, since bombing is usually carried out
at considerable distances from the base. Such machines may carry from
three to six men. All this calls for a tremendous area and a large power
plant. The Handley-Page "Giant," and the Caproni Triplane are examples
of Allied machines of this type while the German "Gotha," used in the
London air raids, is an equivalent enemy machine. As an example of the
weight carrying capacity of a typical bomber, the Handley-Page has
carried a test crew of 21 men, or a personnel load of 3,570 pounds. The
total weight, fully loaded, has been given as 11,500 pounds with a power
plant of 540 horsepower. The maximum speed is 90 miles per hour with a
climbing velocity of about 330 feet per minute. Duration is about 5%
hours at normal speed and full load.
[Illustration: Fig. 3. Handley-Page "Giant" Bombing Type Biplane.
Courtesy "_The Aeroplane_."]
[Illustration: Fig. 4. Curtiss Twin Motor Biplane-Type JN.]
Bombing is of great importance, not only because of the damage caused to
munition factories, transportation lines, store houses, etc., but also
because of the moral effect on both the enemy troops and the civil
population. A well-timed bombing raid will do more to disorganize an
army than almost any other form of attack, and this is attended with a
much less loss of life, and with less cost and equipment. Points in
enemy territory that could be reached in no other way are readily
attacked by bombing planes with all the disastrous effects of heavy
artillery fire. The aeroplane is better adapted for this service than
dirigibles of the Zeppelin type, for they require fewer men for their
operation, and in addition cost less to operate and build.
[Illustration: Fig. 4a. Curtiss "Wireless" Speed Scouts (S-2). By an
Ingenious Arrangement of the Interplane Struts There Is No Exposed Wire
or Cable.]
Bombing operations against well protected objectives are best made at
night since there is less chance of loss through anti-aircraft gun fire,
and also because of the difficulty that the defense machines have in
locating the raiders. Even when well equipped with searchlights and
listening stations, it is not the easiest thing in the world to pick out
and hold the location of an attacking squadron, for the searchlights
immediately betray themselves and can then be put out of action by fire
from the invaders. With the searchlights out of commission, it is almost
impossible for the defending chasers to locate and engage the raiders,
even before the bombs have been dropped. After the bomb dropping has
been accomplished (and with comparative accuracy because of the flares
dropped by the bombing party), the raiders are lightened of a
considerable portion of their load, and are correspondingly increased in
their ability to climb and to evade the enemy chasers.
[Illustration: Fig. 5. Sopwith Speed Scout or "Chaser."]
[Illustration: Fig. 6. Nieuport Biplane Scout with Machine Gun Pivoted
Above the Upper Wing. This Gun Fires Above the Propeller.]
Night flying in squadrons always introduces the danger of collision, and
to minimize this danger, by decreasing the number of machines, the size
and carrying capacity of the bombers has been continually increased.
Again, bombing requires the steady platform that only a large machine
can give, and for accuracy the span and area must be greater than that
of the reconnaissance type. In night flying a large machine is safer to
handle owing to its lower landing speed and ability to come to rest
quickly after landing, and this is of the greatest importance when
landing outside of the aerodrome. For daylight work at comparatively
short distances the smaller bomb carrier used at the beginning of the
war is probably preferable as it has better maneuvering qualities, and
as the bombs are divided among a greater number of machines they are not
so likely to be defeated before accomplishing their object. Because of
their great size, these bombing aeroplanes are nearly always of the
"twin motor" type with two, or even three, independent power plants. The
use of a twin power plant is an added insurance against forced landings
in hostile country, or over unsuitable ground, and even with one dead
engine the machine can be flown home at a fair speed.
[Illustration: Fig. 8. Fokker Synchronized Machine Gun. The Gun Is
Driven by the Motor in Such a Way That the Bullets Pass Between the
Propeller Blades. "_L’Aerophile_."]
"Spotting" for the guidance of the artillery is a duty usually performed
by the reconnaissance type, or small bombing type, and is usually done
under the escort of chasers. Their duty is to direct the battery as to
the placing of shots. The ideal machine for such a purpose would be the
direct lift type similar to a helicopter which could hover over one
particular spot until its object had been accomplished in making
measurements, and plotting enemy positions. Since no such machine is at
present available, the duty must be performed by a low speed aeroplane,
that is large enough to provide a fairly steady platform and at the same
time has sufficient speed for a quick getaway. A dirigible has the
necessary hovering qualities but lacks the speed necessary for avoiding
attack from even the slowest of aeroplanes, and in addition is a
magnificent target for anti-craft guns if kept at an altitude low enough
for accurate observation. A large speed range is a desirable
characteristic in such service.
Photography is of the greatest importance in reconnaissance, since the
camera distinctly records objects on the terrain, so small and obscure
that they may entirely escape the eye of the observer. Again, the
photograph is a permanent record that may be studied at leisure in
headquarters, or may be used in comparisons with photographs taken at an
earlier date in the same territory. Thus changes in the disposition of
enemy batteries, trenches, and troops can be quickly identified. With
modern aeroplane photographic equipment, a vast territory may be
investigated and mapped out by a single machine in a few hours.
Camouflage has but few terrors for the camera, and the photographs often
clearly reveal that which has been passed over time and time again by
the observers. When sent out on a specific mission, the aeroplane
returns the films in an amazingly short length of time, and within a few
minutes they are developed and are ready for the inspection of the
officers in charge. The analysis of these photographs has rapidly
developed into a science well worthy of a Sherlock Holmes. Changes in
the position of shadows, suspiciously sudden growths of underbrush,
changes in the direction of paths, and fresh mounds of earth all have a
definite meaning to the photographic expert.
[Illustration: Fig. 9. Types of Aeroplane Bombs. The Tail Surfaces Guide
the Bomb So That It Strikes on the Firing Pin and at the Same Time
"Safeties" the Bomb So That it Will Not Explode Until it Has Fallen for
Some Distance. In Falling, the Tail Blades Rotate and Release the Firing
Mechanism After the Bomb Has Fallen Clear of the Aeroplane. Courtesy of
"_Flying_."]
[Illustration: Fig. 9-a. Curtiss "JN" Twin Motor Biplane. Observer Is
Seated in Front.]
In the navy the aeroplane has proved of much value in scouting and
particularly in defense against the submarine. Because of its great
speed it has a daily radius of action many times that of a torpedo boat,
and because of its altitude the effective range of vision is still
further increased. At a fair height the observer can easily detect a
submarine even when submerged to a considerable depth, a feat impossible
when near the sea level. For disclosing the conditions existing in an
enemy harbor the aeroplane is fully the equal of the dirigible since it
can approach and retreat rapidly, and without much danger at
comparatively low altitudes. While the dirigible can float indefinitely
at one point, it must be done at an altitude that is safely out of range
of the enemy guns, and this is usually at a point where observation is a
difficult proposition. It does not take long to get the range of such a
target as a hovering dirigible, yet at a much lower altitude it is
difficult to handle naval anti-aircraft guns effectively against a
speeding aeroplane. The smaller scouting seaplanes can report the
position of a submarine to a torpedo boat or "sub-chaser," while the
larger machines are perfectly capable of dealing with the submarine at
first hand. On the large bombing type, a three-pounder gun and a number
of large bombs can be carried, either of which would be sufficient for
the purpose.
[Illustration: Fig. 10. Explosion of a German Aeroplane Bomb Near
Mesopotamia. Courtesy of "_Flying_."]
In land defense chasers and fighters are used for patrol, and to
maintain a barrage against the entrance of enemy machines into our
lines. The patrol machines work along the front line trenches, while the
machines maintaining the barrage are generally arranged in two parallel
lines back of the trenches, the first being about five miles, and the
second about ten miles from the front. All three lines are generally
placed between the enemy and the principal stations and railroad centers
to insure protection from enemy bombers and reconnaissance machines.
Should the first line patrol fail to keep raiders from crossing the
first line trenches, they will have to pass through at least two more
zones of organized fighting squadrons before reaching a vulnerable spot
in our lines. The machines used for patrol and barrage are of the high
speed and fast climbing chaser type. The response to an attack involves
rapid climbing, and a high degree of maneuvering.
[Illustration: Fig. 11. Caproni Triplane with Three Independent Power
Plants. The Motor in the Central Body Drives a Pusher Propeller, While
the Other Power Plants Are Mounted in the Two Outer Bodies and Drive
Tractor Screws. This is an Example of the Larger Bombing Aeroplanes. The
Gun Is Mounted in the Front of the Center Body. Courtesy "_The
Aeroplane_."]
Except for the bombers and battle planes, the machine gun or
"Mitraleuse" has been the only form of arm in common use on aeroplanes.
These use ammunition approximating service rifle caliber and are
furnished in bands, strips or drums according to the type of gun. With
larger guns, the weight of the ammunition has been found excessive with
all but the largest bombing machines, and the recoil of a large caliber
gun has also been difficult to overcome. In a modern American aeroplane
gun of large caliber the recoil has been reduced to almost a negligible
degree, even up to the four-pounder size, by a system of balanced
projectile reactions. This gun has met successful tests, but whether it
has met with general adoption would be difficult to say at the present
time. In Europe, large caliber aeroplane guns have been used on large
"battle planes" or "gun planes" for shelling dirigibles, or in
destroying searchlight stations in bombing raids. The battle planes are
nearly always of the "Twin" type with the gun mounted in the front end
of the fuselage.
*Summary of Types*. To sum up the types required in military operations,
we have: (1) High speed "Chaser" or "Scout" (Single seater), (2) High
speed "Chaser" (Two-seat type), (3) Reconnaissance type, (4) Bombing
type, (5) Gun or Battle Planes. This does not include the training
machines of the two place and "Penguin" types, but as these are simply
unarmed modifications of the two place reconnaissance and single seat
machine, respectively, we will not go into further details at this point
regarding their construction.
*The Chaser or Pursuit Type*. The most important factors in the design
of a chaser are speed and maneuvering ability. The speed must be at a
maximum in both the horizontal and vertical directions, for climbing
ability is fully of as much importance as horizontal speed. Second in
importance is the maximum altitude or "Height of ceiling" to which the
machine can ascend. This maximum "Ceiling" generally goes hand in hand
with the climbing speed, since a fast climber generally has a
correspondingly high maximum altitude. The combination of weight and
head resistance must be such that the climb interferes as little as
possible with the forward velocity.
[Illustration: De Havilland. V. Single Seat "Chaser" or "Speed Scout"
with a Single Rigidly Mounted Machine Gun on Top of Hood (To Left). It
will Be Noted That the Top Wing is Staggered Back Instead of Forward as
in the Usual Type, Thus Allowing the Pilot to Look Directly Up and in
Front of the Top Wing. Dimensions Are in Millimeters.]
Great climbing ability means a large power reserve, hence the weight
carried per horsepower is reduced to from 8 to 12 pounds on the fastest
machines, against the 16 to 18 pounds carried on the larger and slower
reconnaissance types. Strength must be sacrificed to meet these
conditions, so that instead of having a safety factor of from 8 to 12 as
in the larger machines, it is cut to about 5.5, or in other words, the
strength is relatively only half that of the usual type of aeroplane.
This great reduction in strength calls for careful handling, especially
in landing, and also painstaking care in the design and choice of
materials.
High speeds and maneuvering ability both call for small wing areas and
short spans, the areas being so adjusted that the resistance is at a
minimum at the highest speeds. The short spans have a minimum of exposed
interplane bracing and thus indirectly reduce both the head resistance
and the weight. Unfortunately, the most favorable areas at high speeds
are too small for safe landing speeds. With a fixed area, the minimum
landing speed is only a little less than one-half of the maximum flying
speed, hence with a maximum of 150 miles per hour the minimum will
probably be little less than 70 miles per hour. The most efficient wing
sections, and the greatest refinement in the body design, bracing, and
chassis are necessary at speeds of over 100 miles per hour. All other
conditions being equal, the resistance varies as the square of the
velocity, hence at 150 miles per hour, the resistance is 2.25 times that
at 100 miles per hour.
The following table gives approximately the apportionment of the head
resistance producing items in a typical speed scout or chaser.
─────────────────────────────────────────────────────────────────
Body (Fuselage) 68 per cent
─────────────────────────────────────────────────────────────────
Chassis, wheels, struts, etc. 15 per cent
─────────────────────────────────────────────────────────────────
Tail, rudder, fin, elevator 5 per cent
─────────────────────────────────────────────────────────────────
Wing structure, struts, wire, fittings. 12 per cent
─────────────────────────────────────────────────────────────────
The aerodynamic drag due to the lift of the wings is not included in the
above, the useless or parasitic resistance alone being considered. It
will be noted that the body causes by far the greater part of the
resistance, and as a result, the body of the speed scout requires the
most careful attention in regard to streamline form. Fortunately this is
possible with the short stumpy body of the chaser, since a true
streamline form approximates the average body dimensions of the scout.
In the larger machines, the body resistance is not as great in
proportion to the other items since there are more struts and stay
wires, the chassis is larger, and the tail surfaces are of greater area.
The chassis is the next largest item and is one of the most difficult
items to reduce. It has been suggested by several people that the
chassis could be stored away in the body while in flight, but this adds
additional mechanism and weight, and any automatic mechanism for folding
up the chassis members would likely prove unreliable.
*Chaser Armament*. A single seat chaser is provided with one or two
machine guns mounted on top of the fuselage, and directly in front of
the pilot, the length of the barrel being parallel with the fore and aft
center line. They may either be fixed rigidly to the fuselage top, or so
that they can be pointed up, and over the top of the upper wing. With
the machine guns fixed rigidly to the body, as in the early chaser
monoplanes used by Garros and Vedrines, it was necessary at all times to
fire directly through the disc area swept out by the propeller.
Two plans were tried for preventing the propeller from being broken by
the bullets. The first consisted of a device operated by the motor that
stopped the gun whenever the propeller blade came within the path of the
bullets. This early mechanism proved unreliable, since the frequent
stopping, with the propellers running 1200 revolutions per minute, soon
put the apparatus out of order. Soon after the failure of this method,
designers mounted curved protective steel plates on the inner portions
of the propeller blades at points where they were likely to be struck
with bullets. According to calculations in probability and chance, only
one bullet out of every eighteen will strike the protective plate on the
propeller blade, and hence only one out of eighteen bullets will be
wasted. This, however, was a makeshift, and on modern machines the gun
is driven, or "Synchronized" with the motor so that the bullets pass
between the blades.
[Illustration: Curtiss Biplane in Flight. Taken from Another Machine.
Courtesy "_Aerial Age_."]
Many modern single seat chasers have the gun pivoted to the top of the
fuselage so that the pilot can fire above the top plane and to either
side of the body. This does away with the difficulty of keeping the
machine headed directly at the enemy when in action, a method that is
imperative with the fixed type of gun. Two seater chasers are generally
arranged so that the gunner is seated back of the pilot, and the gun is
so pivoted and supported that it can be swung through a wide radius both
toward the front and on either side. This freedom of gun action at least
partly compensates for the slower maneuvering qualities of the two
seater type, since the gun may be swung with the target through quite a
range of field, and without changing the flight direction of the
machine. A gun of this type is provided with stops which prevent the
gunner from shooting into the outlying parts of his own machine. The gun
mounting in many cases of two seater construction consists of a light
circular track that runs around the edge of the cockpit opening. The gun
standard runs on this track, and the gun is pivoted at the upper end of
the standard so that the muzzle can be raised or lowered. The gun turns
in a horizontal plane by sliding on the track, and can be followed
around by the gunner who is seated in the center on a pivoted seat. With
this mounting it is possible to guard against a rear attack, to shoot
straight up, or nearly straight down over the sides of the fuselage.
In a few machines of the two seater type, two machine guns are provided,
one pivoted gun in the rear, and one gun rigidly fastened to the
fuselage in front of the pilot. It is very seldom that both guns can be
brought into action at once unless engaged with a number of enemy
machines, although the front gun is handy in pursuit, and at a time when
the rear gun is ineffective because of the pilot in front of him. Even
with the double equipment, the superior maneuvering qualities of the
single seater makes matters more even than would commonly be supposed.
An added advantage of the single seater is that it is smaller and
therefore more difficult to hit.
English speed scouts have largely adopted the American Lewis gun. The
cartridges in this gun are arranged radially in a circular drum, and are
fed to the gun as the drum revolves. The drum is mounted on the barrel
near the breech and is operated automatically by the successive
explosions. This feeds the cartridges and rejects the empty shells
without the attention of the pilot. It fires about 600 shots per minute.
When one drum is exhausted, another drum of new cartridges can be
quickly and easily inserted. The French use the belt system to a large
extent. In this system the cartridges are attached side by side on a
cotton web belt as in the older types of army machine guns. As in the
Lewis gun, the cartridges are fed automatically by the recoil of the
explosions, and the belt moves through the breech with a step by step
movement until the ammunition is exhausted. This is not nearly as
compact an arrangement as the Lewis gun, and is more difficult to pivot
on account of the dangling belts.
On the right hand side of the Nieuport body there is a drum on which the
belt with the loaded cartridges is wound. The empty end of the belt is
wound on a drum at the left, this drum being provided with a spring to
keep the belt taut. The empty cartridges are discharged through a tube
that passes through the side of the body. On the 1916 Fokker the gun is
of the Maxim type, and is immovably fastened above the engine cowl and
slightly to the right. To fire the gun, the pilot presses down a small
lever fastened to the control column, and from this lever the connecting
Bowden wire closes the motor clutch and starts the gun. A cam is fixed
to the motor shaft in relation to the propeller blades. When firing, the
elevator control is locked fore and aft, while the lateral control
movement is operated by the pilot’s knees. Steering is by the action of
his feet on the rudder bar. Thus the pilot can balance laterally, and
steer with his hands free for the manipulation of the gun, but he cannot
change his elevation.
[Illustration: Aeromarine Training Seaplane]
*Power Plant of the Chaser*. In the smaller speed scouts, the motor is
of the rotary air cooled type, the output ranging from 80 to 110
horsepower, but as the power demands increased the water-cooled motor
came into use, and at the present time has found favor with a large
number of builders. When the power exceeds about 120 horsepower it is
difficult to thoroughly cool the rotary engine, and although the Gnome,
Clerget and Le Rhone are extremely desirable on a chaser because of
their light weight, they cannot be used profitably on the larger scouts.
Up to the present time, the Nieuport and Sopwith use the Clerget and Le
Rhone rotary motors, but the S. P. A. D. and several others have adopted
the water-cooled type. Nearly all of the German chasers, such as the
Roland and Albatros, are water-cooled. Such motors must weigh well below
3 pounds per horsepower if there is to be sufficient power reserve for
fast climbing. The Curtiss scouts are also water-cooled, although the
rating is only 100 horsepower. The French and German machines are very
heavily powered, motors of 175 horsepower being very common, even on
single seaters. The fuel capacity is very limited, probably not
exceeding 2.5 to 3 hours in any case.
*General Dimensions of Scouts*. The following table will give a better
idea of the principal characteristics of these machines. It gives the
overall dimensions, power, speed, climb, etc. It will be noted that the
Nieuport biplane scouts have a smaller lower chord (*). The speeds given
are the sea level speeds since a great change in altitude affects the
performance to a marked degree.
*Reconnaissance Type Arrangement*. These machines are almost invariably
of the two seater type, and are equipped either with one machine gun for
the observer, or with a rigidly fixed gun for the pilot and a pivoted
gun at the rear for the observer. In the majority of cases the observer
is seated in the rear cockpit (Tractor types), and at a point where he
has a greater visual radius and field of fire. With the pusher type, the
observer is, of course, seated in the extreme front of the body, where
he has an extremely wide angle of vision. The pilot in the rear seat of
the pusher is effectually screened from any gun action, either from the
front, side or rear, as the propeller cuts off the field at the back and
the observer and interplane bracing blocks the way at the front and
sides. The observer’s cock-pit is equipped with the signalling
apparatus, photographic equipment, map boards, etc., as well as the
ammunition for the gun. The pilot’s compartment contains the navigating
instruments and controls.
*Armament*. At the beginning of the war nearly all of the French two
seaters were of the pusher type, this arrangement, of course, resulted
in almost a completely dead angle of fire in the rear, and a front
horizontal angle that was practically restricted to 160 degrees. Owing
to the forward position of the gun the vertical angle was quite good,
230 degrees or even better. In the tractor two seater, with a single
movable gun mounted "En barbette" at the rear, the horizontal angle is
about 180 degrees, but the vertical angle is less than with the pusher
type. When the rear gun is supplemented with a front rigidly mounted
gun, there is some protection at the front, but the rigid gun is far
from being as effective as the pivoted rear gun. The front gun of course
fires through the propeller. This armament is used by the German
machines "Aviatic," "Rumpler," "Albatros," and "L. V. G." The forward
rigid gun is usually of the infantry type, while the movable rear gun is
lighter. The latter is fed by drums, or rolled bands on spools, so that
reloading can be performed in the wind stream.
With the two seater type used in reconnaissance, artillery spotting, or
photography, the power is generally in the neighborhood of 220-260
horsepower, and the speed varies from 85 to 100 miles per hour. The area
is approximately 400 to 480 square feet. A single engine is generally
used.
*General Dimensions and Speeds*. Reconnaissance machines of various
types and makes are listed in the following table. A pusher is indicated
by (P) and a tractor by (T). The German aeroplanes (G), and the Allied
aeroplanes (A), are both listed for comparison: It will be noted that
several types of machines have been made by the same firms, and that in
some cases the same machines have different power plants. The Albatros
C-III has been furnished with both the 170 and 220 Mercedes motor. The
Ago biplane has a tapering wing, and the chord width (*) given is taken
at the body. While very recent machines cannot be described, because of
certain restrictions, the horsepower of the latest two seaters will
average about 240 horsepower. If the dates and power items are noted, it
will be seen that the machines used in 1917 have much larger motors
than those built in 1916. The weight per square foot of surface will
average about 6.5 pounds. The loading per horsepower rarely exceeds 17.0
pounds.
*Bombing Type Aeroplanes*. These large aeroplanes are fitted with either
two or three independent power plants. The German bombers are
represented by the Gotha, A. E. G., Friedrichshafen, and Rumpler G,
while the Allied bombers are the Caproni, Handley-Page, Farman, Voisin,
etc. The speed is about that of the reconnaissance type, and will seat
three or more men. The motors average 500 560 horsepower per power
plant, and the wing area is usually well over 1,000 square feet. The
small two seaters are generally equipped with two pivoted machine guns,
while the three seaters have a third machine gun arranged so that it can
be lowered and fired through a trap door in the bottom of the body.
Defense may thus be had from the rear, or below. In some of the pusher
types, a rapid fire gun of comparatively heavy caliber is mounted at the
front of the body in place of the usual machine gun. This is usually the
case with the sea planes used for submarine chasing.
In addition to bombing operations, these large machines are also used
for the protection of "spotting" aeroplanes, or for the direct
protection of the lines against land attacks. These heavily armed
bombers are very difficult to attack, even for the smaller and more
agile "Chasers," as they can fire from below as well as from the front,
top, or sides. In the bombers which have only a single gun in the rear,
the gunner is working at a disadvantage if his adversary forces him to
continually raise and lower his gun from the top of the body to the
lower trap door. This is very tiring to the rear gunner, and if the
chaser’s tactics are carried out for a sufficient length of time, it can
wear out the gunner by continually rising and dropping at the tail of
the bombing plane. In regard to the front gun, the twin motor type
offers many of the advantages of the pusher, and as a whole, the twin
arrangement will nearly double the field of fire of either the tractor
or pusher.
The bombing planes must have a very large radius of action, particularly
those that are used in night bombing operations. The Gothas in bombing
London fly several hundred miles from their base, and recently a
Handley-Page bombing plane flew from London to Constantinople, Turkey,
making only a few stops on the way. Starting out from Hendon, England,
the Handley-Page machine flew to Paris, down the Rhone valley to Lyons
and Marseilles, and then to Pisa, and Rome (Italy), where they landed.
From Rome the machine passed over Naples, over Oranto and then over the
Albanian Alps to the base at Salonica. Making preparations at this base
they flew the final stage of the trip to Constantinople, a distance of
250 miles over hostile country. The bombing of the Turkish capital was
done at night after a flight of 2 1/2 hours from Salonica. When over the
sea of Marmora, the ship "Goeben" was bombed, and in addition a hit was
scored on the two submarines lying at her side. Four bombs struck the
"Goeben" directly, from an altitude of 800 feet. Two more bombs were
dropped on the German ship, "General," which was the headquarters of the
German staff. Finally, after 30 minutes over the city of Constantinople,
the Turkish War Office was the recipient of two more bombs. In the words
of the Turkish communiqué this "Was not entirely destroyed." On its
return to Salonica it was found that the machine had been struck by 26
shrapnel bullets. This disabled one of the power plants so that the
greater part of the return journey was made on a single motor.
From London to Salonica five men were carried. In addition was their
luggage, bedding, two tool boxes, spare parts equivalent in weight to
one engine, and two 11’-6" spare propellers. Complete, the machine
weighed over 6 tons, with a useful load of about 6,000 pounds. In
crossing the Albanian Alps the machine frequently was at an altitude of
10,000 feet. The power plant consisted of two 275 horsepower Rolls-Royce
motors, and even at this high altitude, and with the heavy loading, no
trouble was experienced. During the bombing, only three men were
carried, the remainder of the useful weight being made up of bombs and
other ammunition. While this record will probably be beaten before this
book goes to press, it will at least give an idea as to the requirements
and capabilities of the bombing type aeroplane.
*Military Training Machines*. The military training machines used in the
United States are generally of the two seater tractor type, similar in
external appearance to the reconnaissance type machines already
described. They are low powered, 90 to 125 horsepower, and will have an
average span of 40’-0". The controls are in duplicate so that the
student’s controls move in unison with the instructor’s.
CHAPTER III. ELEMENTARY AERODYNAMICS
*Definition*. Aerodynamics treats of the forces produced by air in
motion, and is the basic subject in the study of the aeroplane. It is
the purpose of this chapter to describe in detail the action of the wing
in flight, and the aerodynamic behavior of the other bodies that enter
into the construction of the aeroplane. At present, aerodynamic data is
almost entirely based on experimental investigations. The motions and
reactions produced by disturbed air are so complex and involved that no
complete mathematical theory has yet been advanced that permits of
direct calculation.
*Properties of Air*. Air being a material substance, possesses the
properties of volume, weight, viscosity and compressibility. It is a
mechanical mixture of the two elementary gases, oxygen and nitrogen, in
the proportion of 23 per cent of oxygen to 77 per cent of nitrogen. It
is the oxygen element that produces combustion, while the nitrogen is
inert and does not readily enter into combination with other elements,
its evident function being to act as a dilutant for the energetic
oxygen. In combustion, the oxygen enters into a chemical combination
with the fuel while the nitrogen passes off with the products of
combustion unchanged.
Air is considered as a fluid since it is capable of flowing like water,
but unlike water, it is highly compressible. Owing to the difference
between air and water in regard to compressibility, they do not follow
exactly the same laws, but at ordinary flight speeds and in the open
air, the variations in the pressure are so slight as to cause little
difference in the density. Hence for *flight alone*, air may be
considered as incompressible. It should be noted that a compressible
fluid is changed in density by variations in the pressure, that is, by
applying pressure the weight of a cubic foot of a compressible fluid is
greater than the same fluid under a lighter pressure. This is an
important consideration since the density of the air greatly affects the
forces that set it in motion, and for this reason the density (weight
per cubic foot) is always specified in a test.
Every existing fluid resists the motion of a body, the opposition to the
motion being commonly known as "resistance." This is due to the cohesion
between the fluid particles and the resistance is the actual force
required to break them apart and make room for the moving body. Fluids
exhibiting resistance are said to have "viscosity." In early aerodynamic
researches, and in the study of hydrodynamics, the mathematical theory
is based on a "perfect fluid," that is, on a theoretical fluid
possessing no viscosity, and while this conception is an aid in studying
the reactions, the actual laboratory results are far from the computed
values. Such theory would assume that a body could move in a fluid
without encountering resistance, which in practice is, of course,
impossible.
In regard to viscosity, it may be noted that air is highly
viscuous—relatively much higher than water. Density for density, the
viscosity of air is about 14 times that of water, and consequently the
effects of viscosity in air are of the utmost importance in the
calculation of resistance of moving parts.
Atmospheric air at sea level is about 1/800 of the density of water. Its
density varies with the altitude and with various atmospheric
conditions, and for this reason the density is usually specified "at sea
level" as this altitude gives a constant base of measurement for all
parts of the world. As the density is also affected by changes in
temperature, a standard temperature is also specified. Experimental
results, whatever the pressure and temperature at which they were made,
are reduced to the corresponding values at standard temperature and at
the normal sea level pressure, in order that these results may be
readily comparable with other data. The normal (average) pressure at sea
level is 14.7 pounds per square inch, or 2,119 pounds per square foot at
a temperature of 60° Fahrenheit. At this temperature 1 pound of air
occupies a volume of 13.141 cubic feet, while at 0° F. the volume
shrinks to 11.58 cubic feet, the corresponding densities being 0.07610
and 0.08633 pounds per cubic foot, respectively. This refers to dry air
only as the presence of water vapor makes a change in the density. With
a reduction in temperature the pressure increases with the density
increase so that the effect of heat is twofold in its effect.
With a constant temperature, the pressure and density both decrease as
the altitude increases, a density at sea level of 0.07610 pounds per
cubic foot is reduced to 0.0357 pounds per cubic foot at an altitude of
20,000 feet. During this increase in altitude, the pressure drops from
14.7 pounds per square inch to 6.87 pounds per square inch. This
variation, of course, greatly affects the performance of aeroplanes
flying at different altitudes, and still more affects the performance of
the motor, since the latter cannot take in as much fuel per stroke at
high altitudes as at low, and as a result the power is diminished as we
gain in altitude. The following table gives the power variations at
different heights above sea level.
This air table also gives the properties of air through the usual range
of flight altitudes. The pressures corresponding to the altitudes are
given both in pounds per square inch and in inches of mercury so that
barometer and pressure readings can be compared. In the fourth column is
the percentage of the horsepower available at different altitudes, the
horsepower at sea level being taken as unity. For example, if an engine
develops 100 horsepower at sea level, it will develop 100 × 0.66=66
horsepower at an altitude of 10,000 feet above sea level. The barometric
pressure in pounds per square inch can be obtained by multiplying the
pressure in inches of mercury by the factor 0.4905, this being the
weight of a mercury column 1 inch high.
[Illustration: NOTE.-Densities marked * are interpolated from a graph,
but are close enough for all ordinary purposes.]
In aerodynamic laboratory reports, the standard density of air is
0.07608 pounds per cubic foot at sea level, the temperature being 15
degrees Centigrade (59 degrees Fahrenheit). This standard density will
be assumed throughout the book, and hence for any other altitude or
density the corresponding corrections must be made. Owing to the fact
that the temperature decreases as we gain altitude, further corrections
must be made in the tabular values, but as the changes are rather
difficult to make and are relatively small we will not take the matter
up at this point.
[Illustration: Fig. 1. Air Flow About a Flat Normal Plate. Pressure Zone
at Front and .#. Turbulent Zone at Rear (H). Arrows Show Direction O
OW.]
*Air Pressure on Normal Flat Plates*. When a flat plate or "plane" is
held at right angles or "normal" to an air stream, it obstructs the flow
and a force is produced that tends to move it with the stream. The
stream divides,as shown in Fig. 1 and passes all around the edges of the
plate (P-R), the stream reuniting at a point (M) far in the rear.
Assuming the air flow from left to right, as in the figure, it will be
noted that the rear of the plate at (H) is under a slight vacuum, and
that it is filled with a complicated whirling mass of air. The general
trend of the eddy paths are indicated by the arrows. At the front where
the air current first strikes the plate there is a considerable pressure
due to the impact of the air particles. In the figure, pressure above
the atmospheric is indicated by ***, while the vacuous space at the rear
is indicated by fine dots. As the pressure in front, and the vacuum in
the rear, both tend to move the surface to the right in the direction of
the air stream, the total force tending to move the plate will be the
difference of pressure on the front and rear faces multiplied by the
area of the plate. Thus if F is the force due to the impact pressure at
the front, and G is the force due to the vacuum at the rear, then the
total resistance (D) or "Drag" is the sum of the two forces.
Contrary to the common opinion, the vacuous part of the drag is by far
the greater, say in the neighborhood of from 60 to 75 per cent of the
total. When a body experiences pressure due to the breaking up of an air
stream, as in the present case, the pressure is said to be due to
"turbulence," and the body is said to produce "turbulent flow." This is
to distinguish the forces due to impact and suction, from the forces due
to the frictional drag produced by the air stream rubbing over the
surface.
Forces due to turbulent flow do not vary directly as the velocity of the
air past the plate, but at a much higher rate. If the velocity is
doubled, the plate not only meets with twice the volume of air, but it
also meets it twice as fast. The total effect is four times as great as
in the first place. The forces due to turbulent flow therefore vary as
the square of the velocity, and the pressure increases very rapidly with
a small increase in the velocity. The force exerted on a plate also
increases directly with the area, and to a lesser extent the drag is
also affected by the shape and proportions. Expressed as a formula, the
total resistance (D) becomes: D = KAV², where K = co-efficient of
resistance determined by experiment, A = area of plate in square feet,
and V= velocity in miles per hour. The value of K takes the shape and
proportion of the plate into consideration, and also the air density.
*Example*. If the area of a flat plate is 6 square feet, the
co-efficient K = 0.003, and the velocity is 60 miles per hour, what is
the drag of the plate in pounds? *Solution*. D = KAV² = 0.003 × 6 × (60
X 60) = 64.80 pounds drag. For a square flat plate, the co-efficient K
can be taken as 0.003.
*Aspect Ratio*. The aspect ratio of a plate is the ratio of the length
to the width. Thus, with an aspect ratio of 2.0, we understand that the
plate is twice as long as it is wide. The ratio of the length to the
width has a very considerable influence of the resistance or drag, this
increasing as the ratio is made greater. If the resistance of a square
plate is taken as 1.00, the resistance of a plate with an aspect ratio
of 20 will be about 1.34 times as great. The following table will give
the effects of aspect ratio on the resistance of a flat plane.
EFFECTS OF ASPECT RATIO ON FLAT PLATES.
─────────────────────────────────────────────────────────────────
Aspect Ratio. Resistance K as a Multiple of a Square
Length/Width Plate.
─────────────────────────────────────────────────────────────────
1.00 (square) 1.00
─────────────────────────────────────────────────────────────────
1.50 1.04
─────────────────────────────────────────────────────────────────
2.00 1.05
─────────────────────────────────────────────────────────────────
3.00 1.07
─────────────────────────────────────────────────────────────────
4.00 1.08
─────────────────────────────────────────────────────────────────
5.00 1.09
─────────────────────────────────────────────────────────────────
6.00 1.10
─────────────────────────────────────────────────────────────────
7.00 1.12
─────────────────────────────────────────────────────────────────
9.00 1.14
─────────────────────────────────────────────────────────────────
10.00 1.15
─────────────────────────────────────────────────────────────────
15.00 1.26
─────────────────────────────────────────────────────────────────
20.00 1.34
─────────────────────────────────────────────────────────────────
30.00 1.40
─────────────────────────────────────────────────────────────────
To convert the values of a square plate into a flat plate of given
aspect ratio, multiply the resistance of the square plate by the factor
under the "K" heading. For example: The resistance of a certain square
plate is 20 pounds, find the resistance of a plate of the same area, but
with an aspect ratio of 15. *Solution*. The factor for a ratio of 15
will be found to be 1.26, hence the resistance of the required plate
will be 20 × 1.26=25.2 pounds.
[Illustration: Fig. 2. Air Flow About a Streamline Body Showing an
Almost Complete Absence of Turbulence Except at the Extreme Rear Edge.
Resistance Is Principally Due to Skin Friction.]
*Streamline Forms*. When a body is of such form that it does not cause
turbulence when moved through the air, the drag is entirely due to skin
friction. Such a body is known as a "streamline form" and approximations
are used for the exposed structural parts of aeroplanes in order to
reduce the resistance. Streamline bodies are fishlike or torpedo-shaped,
as shown by Fig. 2, and it will be noted that the air stream hangs
closely to the outline through nearly its entire length. The drag is
therefore entirely due to the friction of the air on the sides of the
body since there is no turbulence or "discontinuity." In practical
bodies it is impossible to prevent the small turbulence (I), but in
well-designed forms its effect is almost negligible.
In poor attempts at streamline form, the flow discontinues its adherence
to the body at a point near the tail. The poorer the streamline, and the
higher the resistance, the sooner the stream starts to break away from
the body and cause a turbulent region. The resistance now becomes partly
turbulent and partly frictional, with the resistance increasing rapidly
as the percentage of the turbulent region is increased.
The fact that the resistance is due to two factors, makes the resistance
of an approximate streamline body very difficult to calculate, as the
frictional drag and the turbulent drag do not increase at the same rate
for different speeds. The drag due to turbulence varies as V squared
while the frictional resistance only varies at the rate of V to the
1.86th power, hence the drag due to turbulence increases much faster
with the velocity than the frictional component. If we could foretell
the percentage of friction, it would be fairly easy to calculate the
total effect, but this percentage is exactly what we do not know. The
only sure method is to take the results of a full size test.
Fig. 2 gives the approximate section through a streamline strut such as
used in the interplane bracing of a biplane. The length is (L) and the
width is (d), the latter being measured at the widest point. The
relation of the length to the width is known as the "fineness ratio" and
in interplane struts this may vary from 2.5 to 4.5, that is, the length
of the section ranges from 2.5 to 4.5 times the width. The ideal
streamline form has a ratio of from 5. to 5.75. Such large ratios are
difficult to obtain with economy on practical struts as the small width
would result in a weak strut unless the weight were unduly increased.
Interplane struts reach a maximum fineness ratio at about 3.5 to 4.5.
Fig. 3 shows the result of a small fineness ratio, the short, stubby
body causing the stream to break away near the front and form a large
turbulent region in the rear.
An approximate formula showing the relation of fineness ratio and
resistance (curvature equal) was developed by A. E. Berriman, and
published in "Flight" Nov. 12, 1915. Let D = resistance of a flat plate
at a given speed, and R = resistance of a strut at the same speed and of
the same area, then the relation between the resistance of the flat
plate, and the strut will be expressed by the formula R/D=4L/300d, where
L = length of section and d = width as in Fig. 2. This can be transposed
for convenience, by assuming the drag of a flat plate as D = 0.003AV²,
where A = area in square feet, and V = velocity in miles per hour. The
ratio of the strut resistance to the flat plate resistance, given by
Berriman’s formula, can now be multiplied by the flat plate resistance,
or strut resistance = R = 0.003AV² X 4L/300d. = 0.012LAV²/300d. It
should be understood that the area mentioned above is the greatest area
presented to the wind in square feet, and hence is equal to the length
of the strut (not section) multiplied by the width (d).
[Illustration: Fig. 3. Imperfect Streamline Body with a Considerable
Turbulence Due to the Short, Stubby Form. Fig. 4 Shows the Flow About a
Circular Rod or Cylinder.]
Assuming the length (L) of the section as 7.5 inches, and the width (d)
as one inch, the fineness ratio will be 7.5. Using the Berriman formula
in its original form, the relative resistance of the strut and flat
plate of same area will be found as R/D=4L/3000 = 0.1, that is, the
resistance of a streamline form strut of above fineness ratio will be
about 0.1 of a flat plate of the same area. It should be understood that
this is only an approximate formula since even struts of the same
fineness vary among themselves according to the outline. Results
published by the National Physical Laboratory show streamline sections
giving 0.07 of the resistance of a flat plate of the same area, with
fineness ratio = 6.5. In Fig. 4 the effects of flow about a circular rod
is shown, a case where the fineness ratio is 1. The stream follows the
body through less than one-half of its circumference, and the turbulent
region is very large; almost as great as with the flat plate. A circular
rod is far from being even an approach to a perfect form.
In all the cases shown, Figs. 1-2-3-4, it will be noticed that the air
is affected for a considerable distance in front of the plane, as it
rises to pass over the obstruction before it actually reaches it. The
front compression may be perceptible for 6 diameters of the object. From
the examination of several good low-resistance streamline forms it seems
that the best results are obtained with the blunt nose forward and the
thin end aft. The best position for the point of greatest thickness lies
from 0.25 to 0.33 per cent of the length from the front end. From the
thickest part it tapers out gradually to nothing at the rear end. That
portion to the rear of the maximum width is the most important from the
standpoint of resistance, for any irregularity in this region causes the
stream to break away into a turbulent space. From experiments it has
been found that as much as one-half of the entering nose can be cut away
without materially increasing the resistance. The cut-off nose may be
left flat, and still the loss is only in the neighborhood of 5 per cent.
*Resistance Calculations (Turbulency)*. In any plate or body where the
resistance is principally due to turbulent action, as in the flat plate,
sphere, cone, etc., the resistance can be computed from the formula R =
KAV², where R is the resistance in pounds and K, A, and V are as before.
The resistance co-efficient (K) depends upon the shape of the object
under standard air conditions, and differs greatly with flat plates,
cones, sphere, etc. The area (A) is the area presented to the wind, or
is the greatest area that faces the wind, and is taken at right angles
to its direction. The following table gives the value of K for the more
common forms of objects. See Figs. 4 to 12, inclusive:
There are almost an infinite number of different forms, but for the
present the above examples will fill our purpose. As an example in
showing how greatly the form of an object influences its resistance, we
will work out the resistance of a flat plate and a spherical ended cone,
both having the same presented diameter. The cone is placed so that the
spherical end will face the air stream. The area A of both objects will
be: 0.7854 × 2 × 2 = 3.1416 square feet. With an assumed wind velocity
of 100 miles per hour, the resistance of the circular flat disc will be:
R= KAV²=000282 × 3.1416 × (100×100) = 87.96 lbs. For the cone, R =
KAV²=0.000222 × 3.1416 × (100 × 100) = 6.97 lbs. From this calculation
it will be seen that it is advisable to surround the object with a
spherical cone shaped body rather than to present the flat surface to
the wind. In the above table the value of K is given for two positions
of the spherical based cone, the first is with the apex toward the wind,
and the second condition gives the value with the base to the wind. With
the blunt end forward, the resistance is about one-half that when the
pointed apex enters the air stream. This is due to the taper closing up
the stream without causing turbulence.
[Illustration: Figs. 4a-5-6-7-8-9-10-11-12. The Values of the Resistance
Co-efficient K for Different Forms and Positions of Solid Objects.
Arrows Indicate the Direction of the Relative Wind. (Eiffel.)]
With the apex forward there is nothing to fill up the vacuous space when
the air passes over the large diameter of the base as the curve of the
spherical end is too short to accomplish much in this direction.
*Skin Friction*. The air in rubbing over a surface experiences a
frictional resistance similar to water. At the present time the accepted
experiments are those of Dr. Zahm but these are still in some question
as to accuracy. It was found in these experiments that there was
practically no difference caused by the material of the surfaces, as
long as they were equally smooth. Linen or cotton gave the same results
as smooth wood or zinc as long as there was no nap or lint upon the
surface. With a fuzzy surface the friction increased rapidly. This is
undoubtedly due to a minute turbulence caused by the uneven surface, and
hence the increase was not purely frictional, but also due to
turbulence. In the tests, the air current was led parallel to the
surface in such a way that only the friction could move the surface. The
surface was freely suspended, and as the wind moved it edgewise, the
movement was measured by a sharp pointer. End shields prevented impact
of the air on the end of the test piece so that there was no error from
this source. The complete formula given by Dr. Zahm is rather
complicated for ordinary use, especially for those not used to
mathematical computations. If Rf = resistance due to friction on one
side of surface, L= length in direction of wind in feet, b = width of
surface in feet, and V= velocity in feet per second, then
Rf = 0.00000778L⁰.⁹³V¹.⁸⁶b.
It will be noted that the resistance increases at a lower rate than the
velocity squared, and at a less rate than the area. That is to say, that
doubling the area will not double the resistance, but will be less than
twice the amount. Giving the formula in terms of area and miles per hour
units, we have: Rf = 0.0000167A⁰.⁹³V¹.⁸⁶. Where A = area in square feet
and V = miles per hour. The area is for one side of the surface only. A
rough approximation to Zahm’s equation has been proposed by a writer in
"Flight," the intention being to avoid the complicated formula and yet
come close enough to the original for practical purposes. The latter
formula reads: Rf = 0.000009V² where Rf and V are as above. Up to 40
miles per hour the results are very close to Zahm’s formula, and are
fairly close from 60 to 90 miles per hour. This approximation is only
justified when the length in the direction of the wind is nearly equal
to the length. If the length is much greater, there is a serious error
introduced.
This formula is applied to surfaces parallel to the wind such as the
sides of the body, rudder, stabilizer, and elevator surfaces (when in
neutral). A second important feature of the friction formula is that it
illustrates the law of "similitude" or the results of a change in scale
and velocity, hence it outlines what we must expect when we compute a
full size aeroplane from the results of a model test.
*The Inclined Plane*. When a flat plate is inclined with the wind, the
resistance or drag will be broken up into two components, one at right
angles to the air stream, and one parallel to it. If the plate is
properly inclined, the right angled component can be utilized in
obtaining lift as with an aeroplane wing. This is shown in Fig. 13 where
L is the vertical lift force at right angles to the air stream and D is
the horizontal drag acting in the direction of the wind. As in the case
of the plate placed normal to the wind, there is pressure at the front
of the plate and a partial vacuum behind. The resultant force will be
determined by the difference in pressure between the front and the back
of the plate. The forces will vary as V² since the reaction is caused by
turbulent flow. Both the lift and drag will vary with the angle made
with the stream, and there will be a different value for the
co-efficient K for each change in the angle. The angle made with the air
stream is known as the "Angle of incidence" or the "Angle of attack."
The change of drag and lift does not vary at a regular rate with the
angle.
[Illustration: Fig. 13. Flow About, Inclined Plane and Forces Produced
by Stream. Fig. 14. Normal Plane with C.P. at center of Plate. Fig. 15..
C.P. Moves Toward Entering Edge When Plate Is Inclined to Wind.]
A line OR is the resultant of the lift and drag forces L and D, this
resultant being the force necessary to balance the two forces L-D. It is
on the point of application O that the plate balances, and this point is
sometimes known as the "Center of pressure." The center of pressure is
therefore the point at which the resultant intersects the surface of the
flat plate. The resultant OR is approximately at right angles to the
surface at small incident angles, and the point O is nearer the front or
"leading edge" (A) of the plane. The smaller the angle of incidence the
nearer will the point O approach the leading edge A. By drawing OL to
scale, representing the lift, and OD to scale representing the drag, we
can find the resultant OR by drawing LR parallel to the drag OD and DR
parallel to the lift line OL. All lines drawn through the intersection
of LR and DR will give the resultant OR to scale. All of the lines must
be started from the center of pressure at O.
The least resultant will, of course, occur when the plane is parallel to
the air stream. The maximum resultant will occur when the angle of
incidence is about 40 degrees, and on a further increase in the angle,
the value of the resultant will gradually decrease. When the plane is
parallel with the stream, the resultant is parallel to the plate, but
rapidly approaches a position at right angles at about an incidence of 6
to 10 degrees. Beyond 10 degrees incidence the angle of the resultant
increases past the normal.
The center of pressure (O), or the point where the resultant force
intersects the plane, moves forward as the angle of incidence is
decreased from 90°. When at right angles to the air current, the center
of pressure is exactly in the center of the plane as shown by Fig. 14.
In this case the drag (D) is the resultant, and acting in the center,
exactly balances the air forces. In Fig. 15 the angle of incidence is
reduced, consequently the center of pressure moves nearer the leading
edge (A). As the angle continues to decrease, the C. P. moves still
further forward until it lies directly on the front edge when the plate
becomes parallel with the air stream. The center of pressure movement is
due to the fact that more and more work is done by the front part of the
surface as the angle is decreased. Consequently the point of support, or
C. P., must move forward to come under the load. It should be understood
that the plane will balance about the C. P. if a knife edge bearing were
applied as at R in Fig. 15.
*Calculation of Inclined Planes*. We will now consider the inclined
plane as a lifting surface for an aeroplane, and make the elementary
calculations for such purpose. The lift will first be calculated for the
support of the given load, at the given velocity, and then the drag. For
several reasons, that will afterwards be explained, the flat plate or
plane is not used for the main lifting surfaces, but the experience
gained in computing the plate will be of great assistance when we start
calculating actual wings.
*Lift and Drag Co-efficients*. The lift component (L) of the inclined
flat plate depends on the velocity, area, aspect ratio and angle of
incidence. Instead of using the co-efficient (K) formerly used for the
total drag, we will use the lift co-efficient Ky. The formula for lift
now becomes: L = KyAV² where A = area in square feet, and V = velocity
in miles per hour. The lift co-efficient Ky, depends upon the angle of
incidence. The horizontal drag D will be calculated from the drag
co-efficient Kx, which is used in the same way as the co-efficient K in
the case of the normal plate. The subscript (x) is used to distinguish
it from the lift co-efficient. Both Ky and Kx must be corrected for
aspect ratio. The drag can be calculated from the formula: D = KxAV²
where the letters A and V are the same as above.
For the calculation of the drag, we will use a new expression—the
"Lift-Drag Ratio"—or as more commonly given, "L/D." This shows the
relation between the lift and drag, so that by knowing the lift and the
ratio for any particular case, we can compute the drag without the
necessity of going through the tedious calculation D = KxAV². The
lift-drag ratio for a flat plate varies with the angle of incidence, and
the aspect ratio, and hence a separate value must be used for every
inclination and change in aspect. To obtain the drag, divide the lift by
the lift-drag ratio. Hence if the lift is 1200 pounds, and the ratio
equals 6.00, the drag will be: 1200/6=200 pounds, or in other words, the
lift is 6 times the drag force. Changing the angle of incidence through
angles ranging from 1 degree to 7 degrees, the lift-drag ratio of a flat
plate will vary from 1.5 to 7.5. When the plane is parallel to the wind
stream and gives no lift, the drag is computed from Zahm’s skin friction
formula.
The following tables give the values of Ky, Kx, L/D, and center of
pressure movement for flat plates of various aspect ratios. The center
of pressure (C. P.) for each angle is given as a decimal fraction of its
distance from the leading edge, in terms of the width or "Chord."
[Illustration: Fig. 16. Plan View of Plate with Long Edge to Wind. Fig.
17. Plate with Narrow Edge to Wind, Showing Loss in Lift. 17a Shows
Effect of Raked Tips.]
Fig. 16 shows the top view or plan of a lifting surface, with the
direction of the wind stream indicated by the arrows w-w-w = w. The
longer side or "span" is indicated by S, while the width or chord is C.
Main lifting surfaces, or wings, have the long side at right angles to
the wind as shown. When in this position, the surface is said to be in
"Pterygoid Aspect," and when the narrow edge is presented to the wind,
the wing is in "Apteroid Aspect." The word "Pterygoid" means "Bird
like," and was chosen for the condition in Fig. 16, as this is the
method in which a bird’s wing meets the air. Contrary to the case with
true curved aeroplane wings, flat planes usually give better lift in
apteroid than in pterygoid aspect at high angles. The aspect ratio will
be the span (S) divided by the chord (C), or Aspect ratio = S/C.
It will be seen from the above that the lift coefficient Ky increases
with the aspect ratio, and that it generally declines after an angle of
30 degrees. The center of pressure moves steadily back with an increase
in angle.
*Example for Lifts*. A certain flat plane has an area of 200 square
feet, and moves at 50 miles per hour. The angle of incidence is 10
degrees, and the aspect ratio is 6. Find the total lift and the drag in
pounds. Also the location of the center of pressure in regard to the
leading edge, if the chord is 5.8 feet.
*Solution*. Under the table headed, "Aspect Ratio = 6" we find that Ky
at 10° = 0.00173, and that the lift drag ratio is 5.2. The center of
pressure is 0.333 of the chord from the front edge. The total lift then
becomes: L = KyAV² = 0.00173 x 200 x (50 x 50) = 865 pounds. Since the
lift drag ratio is 5.2, the drag = D = 865/5.2 = 166.3 pounds. The
center of pressure will be located 5.8 x 0.333 = 1.4 feet from the
leading edge.
Under the same conditions, but with an aspect ratio of 3, the lift will
become: L = KyAV² = 0.0014 x 200 x(50 x 50) = 700 pounds. In this case
the lift drag ratio is 5.1, so that the drag will be 137.8 pounds. Even
with the same area, the aspect ratio makes a difference of 865–700 = 165
pounds. If we were compelled to carry the original 865 pounds with
aspect 3 wing, we would also be compelled to increase the area, angle,
or speed. If the speed were to be kept constant, we would be limited to
a change in area or angle. In the latter case it would be preferable to
increase the area, since a sufficient increase in the angle would
greatly increase the drag. It will be noted that the lift-drag ratio
decreases rapidly with an increase in the angle.
[Illustration: Burgess Seaplane Scout.]
*Calculation of Area*: Let us assume that we are confined to the use of
an aspect ratio of 6, a speed of 50 miles per hour, weight = 2500
pounds, and wish to obtain the area that will give the most efficient
surface (Least lift-drag ratio.) The equation can be now transposed so
that the area = A = KyV². On examination of the table it will be seen
that the greatest lift-drag ratio is 6.4 at 5 degrees, and that the Ky
at this angle is 0.00103. Substituting these values in the equation for
area, we have A = L/KyV² = 2500/000103 x (50 x 50) = 971 square feet.
[Illustration: Wind Tunnel at Washington Navy Yard in Which the Air
Circulates Continuously Through a Closed Circuit]
CHAPTER IV. EXPERIMENTAL LABORATORIES.
*Test Methods in General*. As already explained, the behavior of a body
in an air stream cannot be predicted with any certainty by direct
mathematical calculation, and for this reason, each and every
aerodynamic body must be tested under conditions that are as nearly
similar to the actual working conditions as possible. Prior to Professor
Langley’s first experiments in 1887, mechanical flight with a heavier
than air machine was derided as an impossibility, even by such
scientists as Navier, Von Helmotz, Gay-Lussac, and others, who proved by
the most intricate calculations that a body larger than a bird could not
be supported by its own energy. Such calculations were, of course, based
on a wrong understanding of air flow, and as no experimental work had
been done up to that time, the flow was assumed according to the
individual taste and belief of the demonstrator. The presence of a
vacuum on the back of a plate was not understood, and as this
contributes full two-thirds of the lift, it is an easy matter to see why
all of the early predictions fell short of the actual lifting forces. To
quote one classic absurdity, the scientist Navier proved mathematically
that if mechanical flight were possible, then 17 swallows would be
capable of developing one horsepower.
In spite of these discouraging computations, Langley proceeded with a
very carefully conducted series of experiments, first investigating the
laws of surface sustenation on various forms of plates, and when the
data collected was sufficient for his needs, he started to construct a
number of model flyers with various wing arrangements and aerofoil
forms. It was Langley’s experiments upon aerofoils that cleared the way
for the Wright Brothers, who started a further and more complete
investigation in 1896. Experiments were made on the effect of curvature,
aspect ratio and angle of incidence, and the results obtained in their
"wind tunnel" were afterwards applied to their successful full size
machine. During 1901 to 1902 the Wrights investigated the properties of
at least 100 different aerofoil forms. Both Langley’s and Wrights’
experiments were with models, although they were made in a different
manner. It was in this way that experimental evidence gained precedence
over theory.
Langley’s specimens were mounted at the end of a revolving arm, so that
with the arm revolving, a relative air stream of known velocity could be
had. The aerofoil was mounted in such a way that the lift and drag could
be measured. In the early experiments of the Wrights, the models were
placed in an enclosed channel through which a stream of air was
maintained by a fan. The model was attached to a balance system so that
the lift and resistance could be measured. This is what is now known as
a "wind tunnel," and at present is almost exclusively used in model
tests. Several investigators immersed their model aerofoils in running
water so that the direction of flow could be visibly observed. While
this latter method is of great service in determining disturbances,
stream line flow, and general characteristics, it is qualitative rather
than quantitative, and cannot be used in obtaining accurate numerical
results. A more accurate method of mapping out the direction of flow,
eddies, etc., is to introduce smoke into the air stream.
*Full Size Experiments*. The old "rule of the thumb" method of building
a full size machine without model test data or other experimental
evidence to begin with has seen its day. It is not only exceedingly
expensive, but is highly dangerous, and many a flyer has met his death
in the endeavor to work out untried principles on a full size machine.
The first cost of the machine, the continual breakage and operating
expense, to say nothing of the damage suits and loss of time, make a
preliminary full size tryout an absurdity at the present time. Again,
the results of full size experiments are not always reliable, as so much
depends upon the pilot and weather conditions. The instruments used on a
large machine are far from being as accurate as those used in model
tests. These are also likely to be thrown out of adjustment unknowingly
by falls or collisions. The great number of variables that enter into
such a test make it almost an impossibility to obtain accurate data on
the result of minor alterations, and, in fact, it is almost impossible
to get the same results twice without further alterations than changing
the pilot. Full scale tests are necessary after sufficient data has been
obtained and applied in a scientific manner to the design of the
machine, but successful performance cannot be expected from a powered
machine built by guess work.
When performed in connection with a wind tunnel, or based on dependable
data from other sources, full size wing tests are very instructive and
useful if care is taken to have the tests conducted under uniform and
known conditions. Many full size experiments of this nature have been
carried out by Saint-Cyr University in France, and by the Royal Aircraft
Factory in Great Britain. Both of these institutions have a wind tunnel
and an almost unlimited fund of performance data, and last but not
least, have the services of skilled observers.
At Saint-Cyr, the full size wings, or the entire machine, are carried on
an electric car or "chariot." The speed of the car, the lift and drag,
can be determined at any moment during the run through suitable
recording devices. Actual flying tests have also been made, the
measurement of the propeller thrust giving the drag, while the lift is
known as being equal to the weight of the machine. The R.A.F. have
carried out a very extensive series of flight tests, the experiments on
the old "B. E.-2" probably being the best known.
The greater part of the experiments performed with the car at Saint-Cyr
differed considerably from the results obtained by model tests, and
apparently these differences followed no specific law. According to
theory, and the results obtained by different laboratories, the
performance of a full size wing should be better than with a model, but
the Saint-Cyr tests showed that such was not always the case. The center
of pressure movement differed in almost every case, and as a direct
result, the pressure distribution of the large wings was materially
different than with the model. The lift-drag ratio results varied,
sometimes being better for the model than for the large wing. These
differences can probably be explained as being due to variation in air
currents, side winds, etc.
*Model Tests*. Since lift and resistance are due to relative motion
between a body and the air stream, a model can either be towed through
the air, or it can be held stationary while the air is forced past it.
There has been some controversy on the relation between the results
obtained by the two methods, but for the present we will accept the
common belief that the results obtained by either method are the same.
In testing ship models, they are always towed through the tank, but in
the case of aero-dynamic bodies this is complicated and not desirable.
In towing models through the air a very high velocity is needed and this
necessitates either a very long track or a short time length for making
the observations. Again, it is almost impossible to avoid errors because
of vibration, inequality of movement due to uneven track, or air eddies
caused by differences in temperature and by the movement of the towing
device. In fact, the same difficulties apply to towed model tests as to
the full size "electric chariot."
The whirling arm method of testing as used by Langley, Maxim-Vickers,
and others, is a form of "towed testing," but is also open to serious
objections. Unless the arm is very long, every part of the model surface
will not move at the same velocity, the outer portions moving the
faster. As the forces produced by an air stream vary as the square of
the velocity, this may introduce a serious error. The fact that the body
passes repeatedly over the same path introduces error, as the body after
the first revolution is always working in disturbed air. The centrifugal
force, and the currents set up by the arm itself all reduce the accuracy
of the method.
When a model is placed in a uniform current of air in a properly
designed channel or tunnel, the greater part of the errors due to towed
tests are eliminated. The measuring instruments can be placed on a firm
foundation, the air stream can be maintained at a nearly uniform speed
and with little error due to eddies, and the test may be continued under
uniform conditions for an indefinite period. While there are minor
errors due to wall friction and slight variations in the velocity at
different points in the cross section of the tunnel, they are very small
when compared to the errors of towing. For this reason the wind tunnel
is the accepted means of testing.
*Eiffel’s Wind Tunnel*. The Eiffel Laboratory at Auteuil, France, is
probably one of the best known. The results in Chapters III and V were
obtained in this laboratory and thousands of similar experiments have
been carried out at this place. Two tunnels, a large and small, are
placed side by side in the main laboratory room, the tunnels being
supported midway between the floor and ceiling. The air is drawn from
this room into an airtight experimental chamber through a bell-mouthed
circular opening. A grill or honeycomb baffle is placed in the opening
to straighten out the flow, and from this point the air passes across
the chamber and exits through a circular duct to the suction side of a
large fan. From the fan the air is discharged into the room. The same
air thus circulates through the tunnel continuously. The test chamber is
considerably wider than the openings so that the walls do not influence
the flow around the model. A cylinder of air passes through the chamber
at a remarkably uniform velocity, and without any appreciable eddies.
Diameter of the stream approximates 6.6 feet in the large tunnel and 3
feet in the smaller. In the large tunnel the maximum velocity is 105
feet per second, and 131 feet per second is attained in the smaller. A
50-horsepower electric motor is used with a multiblade fan of the
"Sirocco" type.
The observer and weighing mechanism are supported above the air stream
on a sliding floor, and a standard extends from the model in the wind
stream to the balances on the weighing floor. These balances determine
the lift and drag of the models, the center of pressure, etc.
*The N. P. L. Tunnel*. The National Physical Laboratory at Teddington,
England, has a remarkably complete and accurate aerodynamic equipment.
This consists of a large tunnel of 7 square feet area, a small tunnel of
4 square feet, and a whirling table house. The large tunnel is 80 feet
in length with an air flow of 60 feet per second, the air being
circulated by a four-bladed propeller driven by an electric motor of 30
horsepower. The velocity is uniform within one-half per cent, and the
most accurate of results have been obtained. The smaller tunnel is about
56 feet long and the wind velocity is about 40 miles per hour maximum.
The propeller revolves at 600 revolutions and is driven by a
10-horsepower electric motor. There is no chamber and the models are
suspended in the passage half way between the "Diffuser" in the entering
end, and the baffles in the exit. The Massachusetts Institute of
Technology, and the Curtis Aeroplane Company both have similar tunnels.
*United States Navy Tunnel*. In this tunnel the air is confined in a
closed circuit, the return tunnel being much larger than the section in
which the tests are performed. The cross-sectional area is 8 square feet
at the point of test, and the stream is uniform within 2 per cent. The
balance and controls are mounted on the roof of the tunnel, with an arm
extending down through the air stream to the model, as in the Eiffel
tunnel. The balance is similar to Eiffel’s and is sensitive to less than
2/1000 pound. A velocity of 75 miles per hour may be attained by the
500-horsepower motor, but on account of the heating of the air stream
through skin friction, the tests are generally made at 40 miles per
hour. Models up to 36-inch span can be tested, while the majority of
models tested at M. I. T. are about 18 inches.
CHAPTER V. AERODYNAMICS OF LIFTING SURFACES (AEROFOILS).
*General Wing Requirements*. The performance of flat plates when used as
lifting surfaces is very poor compared with curved sections or wing
forms. It will be remembered that the greatest lift-drag ratio for the
flat plate was 6.4, and the best Ky was 0.00294. Modern wing sections
have a lift-drag ratio of over 20.0, and some sections have a lift
coefficient of Ky–0.00364, or about 60 per cent higher than the lift
obtained with a flat plate. In fact, this advantage made flight
possible. To Langley, above all other men, we owe a debt of gratitude
for his investigations into the value of curved wing surfaces.
*Air Flow About an Aerofoil*. To distinguish the curved wing from the
flat plane, we will use the term "Aerofoil." Such wings are variously
referred to as "Cambered surfaces," "Arched surfaces," etc., but the
term "Aerofoil" is more applicable to curved sections. The variety of
forms and curvatures is almost without limit, some aerofoils being
curved top and bottom, while others are curved only on the upper
surface. The curve on the bottom face may either be concave or convex,
an aerofoil of the latter type being generally known as "Double
cambered." The curves may be circular arcs, as in the Wright and
Nieuport wings, or an approximation to a parabolic curve as with many of
the modern wings.
Fig. 1-b shows the general trend of flow about an aerofoil at two
different angles of incidence, the flow in the upper view being
characteristic for angles up to about 6 degrees, while the lower view
represents the flow at angles approximating 16°. At greater angles the
air stream breaks away entirely from the top surface and produces a
turbulence that greatly resembles the disturbance produced by a flat
plate. It will be noted in the top figure (At small angles) that the
flow is very similar to the flow about a streamline body, and that the
air adheres very closely to the top surface. The flow at small angles is
very steady and a minimum of turbulence is produced at the trailing
edge.
[Illustration: Figs. 1a, 1b. Aerofoil Types and Flow at Different
Angles.]
When increased beyond 6°, turbulence begins, as shown in the lower
figure, and a considerable change takes place in the lift-drag ratio.
This is known as the "Lower Critical Angle." The turbulence, however, is
confined to the after part of the wing, and little or no disturbance
takes place in the locality of the lower surface. We observe that an
increase in angle and lift produces an increased turbulent flow about
the upper surface, and hence the upper surface is largely responsible
for the lift. Below 10° the trend of the upper portion of the stream is
still approximately parallel to the upper surface.
[Illustration: Fig. 2. Showing How Lift Is Obtained When an Aerofoil Is
Inclined at a Negative Angle, the Line of Flight Being Along X-X.]
From 16° to 18°, the stream suddenly breaks entirely away from the wing
surface, and produces an exceedingly turbulent flow and mass of eddies.
The lift falls off suddenly with the start of the discontinuous flow.
The angle at which this drop in lift takes place is known variously as
the "Second Critical Angle," the "Burble Point," or the "Stalling
Angle." Any further increase in angle over the stalling angle causes a
drop in lift as the discontinuity is increased. With the flat plane, the
burble point occurs in the neighborhood of 30° and movement beyond this
angle also decreases the lift. In flight, the burble point should not be
approached, for a slight increase in the angle when near this point is
likely to cause the machine to drop or "Stall." The fact that the
maximum lift occurs at the critical angle makes the drop in lift at a
slightly greater angle, doubly dangerous.
A peculiar feature of the aerofoil lies in the fact that lift is still
obtained with a zero angle of incidence, and even with a negative angle.
With the aerofoil shown in Fig. 2 there will be a considerable lift when
the flat bottom is parallel with the direction of travel, and some lift
will still be obtained with the front edge dipped down (Negative Angle).
The curved upper surface causes the air stream to rise toward the front
edge, as at E, hence the wing can be dipped down considerably in regard
to the line of motion X-X, without going below the actual air stream.
*Action in Producing Lift*. At comparatively high angles of incidence,
where there is turbulent flow, the lift and drag are due principally to
the difference in pressure between the upper and lower surfaces as in
the case of the flat plate.
There is a positive pressure below as in the front of a flat inclined
plane, and a vacuous region above the upper surface. The drag with the
plane below the burble point, and above the "Lower Critical Angle," is
due both to skin friction and turbulence—principally to the latter.
Below the first critical angle (6°), the skin friction effect increases,
owing to the closeness with which the air stream hangs to the upper
surface.
Since there is but little turbulence at the small angles below 6°, the
theory of the lift at this point is difficult to explain. The best
explanation of lift at small angles is given by Kutta’s Vortex
Hypothesis. This theory is based on the fact that a wing with a
practically streamline flow produces a series of whirling vortices
(Whirlpools) in the wake of the wings, and that the forward movement of
the plane produces the energy that is stored in the vortices. The
relation between these vortices is such, that when their motion is
destroyed, they give up their energy and produce a lifting reaction by
their downward momentum. The upward reaction on the wing is thus equal
and opposite to the downward momentum of the air vortices.
*Drag Components*. At large angles of incidence where turbulence exists,
the lift and also the drag are nearly proportional to the velocity
squared (V²). Where little turbulence exists, and where the air stream
hugs the surface closely, the drag is due largely to skin friction, and
consequently this part of the drag varies according to Zahm’s law of
friction (V²). For this reason it is difficult to estimate the
difference in drag produced by differences in velocity, since the two
drag components vary at different rates, and there is no fixed
proportion between them. Since the frictional drag does not increase in
proportion to the area, but as A⁰.⁹⁸, difficulty is also experienced in
estimating the drag of a full size wing from data furnished by model
tests.
*Incidence and Lift*. Up to the burble point the lift increases with an
increase in the angle; but not at a uniform rate for any one aerofoil,
nor at the same rate for different aerofoils. The drag also increases
with the angle, but more rapidly than the lift after an incidence of
about 4° is passed, hence the lift-drag ratio is less at angles greater
than 4°. Decreasing the angle below 4° also decreases the lift-drag, but
not so rapidly as with the larger angles. At the angle of "No Lift" the
drag is principally due to skin friction.
Fig. 3 shows a typical lift and incidence chart that gives the relation
between the angle of incidence Ɵ and the lift coefficient. This curve
varies greatly for different forms of aerofoils both in shape and
numerical value, and it is only given to show the general form of such a
graph. The curve lying to the left, and above the curve for the "Flat
plate," is the curve for the particular aerofoil shown above the chart.
The "Lift-Coefficients" at the left hand vertical edge correspond to the
coefficient Ky, although these must be multiplied by a factor to convert
them into values of Ky. As shown, they are in terms of the Absolute
units used by the National Physical Laboratory and to convert them into
the Ky unit they must be multiplied by 0.0051V² where V is in miles per
hour, or 0.00236v² where v = feet per second. The incidence angle is in
degrees.
[Illustration: Fig. 3. Chart Showing Relation Between Incidence And
Lift.]
It will be noted that the lift of the aerofoil is greater than that of
the plate at every angle as with nearly every practical aerofoil. The
aerofoil has a lift coefficient of 0.0025 at the negative angle of -3°,
while the lift of the flat plate of course becomes zero at 0°. As the
incidence of the aerofoil increases the lift coefficient also increases,
until it reaches a maximum at the burble point (Stalling angle) of about
11.5°. An increase of angle from this point causes the lift coefficient
to drop rapidly until it reaches a minimum lift coefficient of 0.46 at
17°. The flat plate as shown, reaches a maximum at the same angle, but
the lift of the plate does not drop off as rapidly. The maximum
coefficient of the aerofoil is 0.58 and of the plate 0.41. The rapid
drop in pressure, due to the air stream breaking away at the burble
point, is clearly shown by the sharp peak in the aerofoil curve. The
sharpness of the drop varies among different aerofoils, the peaks in
some forms being very flat and uniform for quite a distance in a
horizontal direction, while others are even sharper than that shown.
Everything else being equal, an aerofoil with a flat peak is the more
desirable as the lift does not drop off so rapidly in cases where the
aviator exceeds the critical angle, and hence the tendency to stall the
machine is not as great. This form of chart is probably the simplest
form to read. It contains only one quantity, the lift-coefficient, and
it shows the small variations more clearly than other types of graphs in
which the values of Kx, lift-drag, and the resultant force are all given
on a single sheet.
*Center of Pressure Movement*. As in the case of the flat plate the
center of pressure on an aerofoil surface varies with the angle of
incidence, but unlike the plate the center of pressure (C. P.) moves
backward with a decrease in angle. The rapidity of travel depends upon
the form of aerofoil, in some types the movement is very great with a
small change in the angle, while in others the movement is almost
negligible through a wide range. In general, aerofoils are inherently
unstable, since the C. P. moves toward the trailing edge with decreased
angles, and tends to aggravate a deficiency in the angle. If the angle
is too small, the backward movement tends to make it still smaller, and
with an increasing angle the forward movement of the center of pressure
tends to make the angle still greater.
Fig. 4 is a diagram showing the center of pressure movement for a
typical aerofoil with the aerofoil at the top of the chart. The left
side of the chart represents the leading edge of the aerofoil and the
right side is the trailing edge, while the movement in percentages of
the chord length is shown by the figures along the lower line. Thus
figure ".3" indicates that the center of pressure is located 0.3 of the
chord from the leading edge. In practice it is usual to measure the
distance of the C. P. from the leading edge in this way.
[Illustration: Fig. 4. Chart Giving Relation Between Incidence and C.P.
Movement.]
For an example in the use of the chart, let us find the location of the
C. P. at angles of 0°, 3° and 7°. Starting with the column of degrees at
the left hand edge of the chart, find 0°, and follow along the dotted
line to the right until the curve is reached. From this point follow
down to the lower row of figures. It will be found that at 0° the C. P.
lies about half way between 0.5 and 0.6, or more exactly at 0.55 of the
chord from the leading edge. Similarly at 3° the C. P. is at 0.37 of the
chord, and at 7° is at 0.3 of the chord. From 11° to 19°, the C. P. for
this particular aerofoil moves very little, remaining almost constant at
0.25 of the chord. Reducing the angle from 3° causes the C. P. to
retreat very rapidly to the rear, so that at –1° the C. P. is at 0.8 of
the chord, or very near the trailing edge of the wing.
*Other Forms of Charts*. The arrangement of wing performance charts
differs among the various investigators. Some charts show the lift,
drag, lift-drag ratio, angle of incidence, center of pressure movement,
and resultant pressure on a single curve. This is very convenient for
the experienced engineer, but is somewhat complicated for the beginner.
Whatever the form of chart, there should be an outline drawing of the
aerofoil described in the chart.
Fig. 5 shows a chart of the "Polar" variety in which four of the factors
are shown by a single curve. This type was originated by Eiffel and is
generally excellent, except that the changes at small angles are not
shown very clearly or sharply. The curve illustrates the properties of
the "Kauffman" wing, or better known as the "Eiffel No. 37." A more
complete description of this aerofoil will be found under the chapter
"Practical Wing Sections." A single curve is marked at different points
with the angle of incidence (0° to 12°). The column at the left gives
the lift-coefficient Ky, while the row at the bottom of the sheet gives
the drag-coefficients Kx. At the top of the chart are the lift-drag
ratios, each figure being at the end of a diagonal line. In this way the
lift, drag, liftdrag and angle of incidence are had from a single curve.
Take the characteristics at an angle of 10 degrees for example. Find the
angle of 10° on the curve, and follow horizontally to the left for Ky.
The lift-coefficient will be found to be 0.0026 in terms of miles per
hour and pounds per square foot. Following down from 10°, it will be
found that the drag-coefficient Kx = 0.00036. Note the diagonal lines,
and that the 10° point lies nearest to the diagonal headed 7 at the top
of the chart. (It is more nearly a lift-drag ratio of 7.33 than 7.) In
the same way it will be found that an angle of 8 degrees lies almost
exactly on the lift-drag diagonal marked 9. The best lift-drag is
reached at about 2 degrees at which point it is shown as 17.0. The best
lift-coefficient Ky is 0.00276 at 12 degrees.
[Illustration: Fig. 5. Polar Type Chart Originated By Eiffel.]
A third class of chart is shown by Fig. 6. This single chart shows three
of the factors by means of three curves; one for the lift-coefficient,
one for the drag-coefficient, and one for the C. P. movement. Follow the
solid curves only, for the dotted lines are for comparison with the
results obtained by another laboratory in checking the characteristics
of the wing. The curves refer to the R.A.F.-6 section described in the
chapter on "Practical Wing Sections." The lift-coefficients Ky will be
found at the right of the chart with the drag-coefficients Kx at the
left and in the lower column of figures. The upper column at the left is
for the C. P. movement and gives the C. P. location in terms of the
chord length. The angles of incidence will be found at the bottom.
Values are in terms of pounds per square foot, and miles per hour.
[Illustration: Fig. 6. Chart of R.A.F.-6 Wing Section with Three
Independent Curves.]
In using this chart, start with the angle of incidence at the bottom,
and follow up vertically to the lift or drag curves. If the value of Ky
is desired, proceed from the required incidence and up to the "Lift"
curve, then horizontally to the right. To obtain the drag, follow up
from the angle of incidence to the "drift" curve, and then horizontally
to the left. For the position of the C. P., trace up from angle until
the "Center of Pressure" curve is reached, and then across horizontally
to the left. If the angle of 8 degrees is assumed, the lift-coefficient
will be found as Ky = 0.0022, the drag Kx =0.00016, and the center of
pressure will be located at 0.32 of the chord from the leading edge.
This test was made with the air density at 0.07608 pounds per cubic
foot, and at a speed of 29.85 miles per hour. The peak at the burble
point is fairly flat, and gives a good range of angle before the lift
drops to a serious extent. The aerofoil R.A.F.-6 is a practical wing
form used in many machines, and this fact should make the chart of
special interest.
*Surface Calculations*. The calculation of lift and drag for an aerofoil
are the same as those for a flat plate, that is, the total lift is
expressed by the formula: L = KyAV² where A is the area in square feet,
and V is the velocity in miles per hour. From this primary equation, the
values of the area and velocity may be found by transposition.
A = L/KyV² and V = L/KyA.
The drag can be found from the old equation, D = KxAV², or by dividing
the lift by the lift-drag ratio as in the case of the flat plate.
*Example*: A wing of the R.A.F.-6 form has an area of 200 square feet,
and the speed is 60 miles per hour. What is the lift at 6° incidence?
*Solution*. From Chart No. 6 the lift coefficient Ky is 000185 at 6°,
hence the total lift is: L = KyAV² = 0.00185 x 200 x (60 x 60) = 1332
pounds. With an angle of 8 degrees, and with the same speed and area,
the lift becomes,
L = 0.0022 x 200 x (60 x 60) = 1584 pounds. The drag coefficient Kx at
an angle of 6° is 0.00012, and at 8° is 0.00016. The drag at 6° becomes
D = KxAV² = 0.00012 x 200 x (60 x 60) = 86.4 pounds. The lift-drag ratio
at this angle is L/D = 1332/864 = 15.4. The drag at 8° is D = KxAV² =
0.00016 x 200 x (60 x 60) = 115.2 pounds. The lift-drag at 8° is L/D =
1584/115.2 = 13.8.
*Forces Acting on Aerofoil*. Fig. 7 is a section through an aerofoil of
a usual type, with a concave under-surface In an aerofoil of this
character all measurements are made from the chordal line X-X which is a
straight line drawn across (and touching) the entering and trailing
edges of the aerofoil. The angle made by X-X with the horizontal is the
angle of incidence (i). The width of the section, measured from tip to
tip of the entering and trailing edges, is called the "Chord." In this
figure the entering edge is at the left. The direction of lift is "Up"
or as in the case of any aerofoil, acts away from the convex side.
In the position shown, with horizontal motion toward the left, the lift
force is indicated by L, and the horizontal drag force by D, the
direction of their action being indicated by the arrow heads. The force
that is the resultant of the lift and drag, lies between them, and is
shown by R. The point at which the line of the resultant force
intersects the chordal line X-X is called the "Center of Pressure" (C.
P.) The resultant is not always at right angles to the chordal line as
shown, but may lie to either side of this right angle line according to
the angle of incidence (i). A force equal to and in the same direction
as R, will hold the forces L and D in equilibrium if applied at the
center of pressure (C. P.) Owing to the difference in the relative
values of L and D at various angles of incidence, the angle made by R
with the chordal line must vary. The lift and drag are always at right
angles to one another. The resultant can be found by drawing both the
lines L and D through the C. P., and at right angles to one another, and
then closing up the parallelogram by drawing lines parallel to L and D
from the extreme ends of the latter. The resultant force in direction
and extent will be the diagonal R drawn across the corners of the
parallelogram.
[Illustration: Fig. 7. Forces Acting on an Aerofoil, Lift, Drag, and
Resultant. Relative Wind Is from Left to Right.]
The forces acting on the upper and lower surfaces are different, both in
direction and magnitude, owing to the fact that the upper and lower
surfaces do not contribute equally to the support of the aerofoil. The
upper surface contributes from 60 to 80 per cent of the total lift. A
change in the outline of the upper curved surface vitally affects both
the lift and lift-drag, but a change in the lower surface affects the
performance to an almost negligible amount.
In the case of thin circular arched plates the curvature has a much more
pronounced effect. When the curvature of a thin plate is increased, both
the upper and lower surfaces are increased in curvature, and this
undoubtedly is the cause of the great increase in the lift of the sheet
metal aerofoils tested by Eiffel.
The drag component of the front upper surface is "Negative," that is,
acts with the horizontal force instead of against it. The lower surface
drag component is of course opposed to the horizontal propelling force
by enough to wholly overcome the assisting negative drag force of the
front upper surface. The resultants vary from point to point along the
section of the aerofoil both in extent and direction. A resultant true
for the entering edge would be entirely different at a point near the
trailing edge.
*Distribution of Pressure*. To fully understand the relative pressures
and forces acting on different parts of the aerofoil we must refer to
the experimental results obtained by the Eiffel and the N. P. L.
laboratories. In these tests small holes were drilled over the aerofoil
surface at given intervals, each hole in turn being connected to a
manometer or pressure gauge, and the pressure at that point recorded.
While the reading was being taken, the wind was passed over the surface
so that the pressures corresponded to actual working conditions. It was
found that the pressure not only varied in moving from the entering to
trailing edge, but that it also varied from the center to the tips in
moving along the length of the plane. The rate of variation differed
among different aerofoils, and with the same aerofoil at different
angles of incidence.
On the upper surface, the suction or vacuum was generally very high in
the immediate vicinity of the entering edge. From this point it
decreased until sometimes the pressure was actually reversed near the
trailing edge and at the latter point there was actually a downward
pressure acting against the lift. The positive pressure on the under
surface reached a maximum more nearly at the center, and in many cases
there was a vacuum near the entering edge or at the trailing edge. With
nearly all aerofoils, an increase in the curvature resulted in a decided
increase in the vacuum on the upper surface, particularly with thin
aerofoils curved to a circular arc.
[Illustration: Fig. 8. Pressure Distribution for Thin Circular Section.
Fig. 9. Shows the Effect of Increasing the Camber. (Eiffel)]
By taking the sum of the pressures at the various parts of the surface,
it was found that the total corresponded to the lift of the entire
aerofoil, thus proving the correctness of the investigation. The sum of
the drag forces measured at the different openings gave a lower total
than the total drag measured by the balance, and this at once suggests
that the difference was due to the skin friction effect that of course
gave no pressure indication. The truth of this deduction is still
further proved by the fact that the drag values were more nearly equal
at large angles where the turbulence formed a greater percentage of the
total drag.
Figs. 8, 9, 10, 11, 12 are pressure distribution curves taken along the
section of several aerofoil surfaces. These are due to Eiffel. In Fig. 8
is the pressure curve for a thin circular aerofoil section, the depth of
the curve measured from the chordal line being 1/13.5 of the chord. The
vacuum distribution of the upper surface is indicated by the upper
dotted curve, while the pressure on the bottom surface is given by the
solid curve under the aerofoil. The pressures are given by the vertical
column of figures at the right and are in terms of inches of water, that
is, the pressure required for the support of a water column of the
specified height. Figures lying above 0 and marked (-), refer to a
vacuum or negative pressures, while the figures below zero are positive
pressures above the atmospheric. The entering edge is at the right, and
the angle of incidence in all cases is 6°.
It will be seen that the vacuum jumps up very suddenly to a maximum at
the leading edge, and again drops as suddenly to about one-half the
maximum. From this point it again gradually increases near the center,
and then declines toward the trailing edge. It will also be seen that
the pressure on the lower surface, given by the solid curve, is far less
than the pressure due to the upper surface. Since the lower pressure
curve crosses up, and over the zero line at a point near the trailing
edge, it is evident that the supper surface near the trailing edge is
under a positive pressure, or a pressure that acts down and against the
lift. The pressures in any case are very minute, the maximum suction
being 0.3546 inch of water, while the maximum pressure on the under
surface is only 0.085 inch.
Fig. 9 shows the effect of increasing the curvature or camber, the
aerofoil in this case having a depth equal to 1/7 the chord, or nearly
double the camber of the first. The sharp peak at the entering edge of
the pressure curve is slightly reduced, but the remaining suction
pressures over the rest of the surface are much increased, indicating a
marked increase in the total pressure. The pressure at the center is now
nearly equal to the front peak, and the pressure is generally more
evenly distributed. There is a vacuum over the entire upper surface and
a positive pressure over the lower. The general increase in pressure due
to the increased camber is the result of the greater downward deviation
of the air stream, and the corresponding greater change in the momentum
of the air. The speed at which the tests were made was 10 meters per
second, or 22.4 miles per hour. The curves are only true at the center
of the aerofoil length and for an aspect ratio of 6.
The average pressure over the entire surface in Fig. 8 is 1.202 pounds
per square foot, and that of Fig. 9 is 1.440 pounds, a difference of
0.238 pound per square foot due to the doubling camber (16.5 per cent).
Another aerofoil with a camber of only 1/27 gave an average pressure of
0.853 pound per square foot under the same conditions. A flat plane gave
0.546. Tabulation of these values will show the results more clearly.
──────────────────────────────────────────────────────────────────────
Camber of Av. Pres. Per Inc. in Pres. in Efficiency
Surface Sq. Ft. Lbs./Sq. Ft. ─────────────────
Top. Bottom.
──────────────────────────────────────────────────────────────────────
Flat Plane. 0.546 0.000 0.89 0.11
──────────────────────────────────────────────────────────────────────
1/27 0.853 0.307 0.72 0.28
──────────────────────────────────────────────────────────────────────
1/13.5 1.202 0.349 0.71 0.29
──────────────────────────────────────────────────────────────────────
1/7 1.440 0.238 0.59 0.41
──────────────────────────────────────────────────────────────────────
In this table, the "Efficiencies" are the relative lift efficiencies of
the top and bottom surfaces. For example, in the case of the 1/7 camber
the top surface lifts 59 per cent, and the bottom 41 per cent of the
total lift.
Fig. 10 is a thin aerofoil of parabolic form, while Fig. 11 is an
approximation to the comparatively thick wing of a bird. In both these
sections it will be noted that the front peak is not much greater than
the secondary peak, and that the latter is nearer the leading edge than
with the circular aerofoils. Also that the drop between the peaks is
small or entirely lacking. The lower surface of the trailing edge is
subjected to a greater down pressure in the case of the thin parabola,
and there is also a considerable down pressure on the upper leading
edge. The pressure in Fig. 10 is 1.00 pound per square foot, and that of
No. 11 is 1.205, while the efficiency of the top surfaces is
respectively 72 and 74 per cent.
[Illustration: Fig. 10. Thin Parabolic Aerofoil with Pressure
Distribution. Fig. 11. Pressure Distribution of Thick Bird’s Wing Type.
(Eiffel)]
Fig. 12 shows the effect of changing the angle of the bird wing from
zero to 8 degrees. The lift per square foot in each case is shown at the
upper left hand corner of the diagram while the percentages of the upper
and lower surface lifts are included above and below the wing. For these
curves I am indebted to E. R. Armstrong, formerly of "Aero and Hydro."
As the angle is increased, the suction of the upper surface is much
increased (0.541 to 1.370 pounds per square foot), and the pressure at
the leading edge increases from depression to a very long thin peak. The
maximum under pressure is not much increased by the angle, but its
distribution and average pressure are much altered. At 0° and 2° the
usual pressure is reduced to a vacuum over the front of the section as
shown by the lower curve crossing over the upper side of the wing, and
at this point the under surface sucks down and acts against the lift.
[Illustration: Fig. 12. Effect of Incidence Changes on the Pressure
Distribution of a Thick Bird’s Wing. (After Eiffel)]
*Distribution of Drag Forces*. The drag as well as the lift changes in
both direction and magnitude for different points on the wing. In the
front and upper portions the drag is "Negative," that is, instead of
producing head resistance to motion it really acts with the propelling
force. Hence on the upper and front portions the lift is obtained with
no expenditure of power, and in fact thrust is given up and added to
that of the propeller. The remaining drag elements at the rear, and on
the lower surface, of course more than overcome this desirable tendency
and give a positive drag for the total wing. The distribution is shown
by Fig 13 which gives the lift, drag and resultant forces at a number of
different points on two circular arc aerofoils having cambers of 1/13.5
and 1/7 respectively. In this figure, the horizontal drag forces are
marked D and d, and the direction of the drag is shown by the arrows.
The lift is shown by L and the resultant by R as in the Fig. 7.
As shown, the arrows pointing to the right are the "Negative" drag (d)
forces that assist in moving the plane forward, while the drag indicated
by arrows (D) pointing to the left are the drag forces that oppose or
resist the horizontal motion. With the smaller camber (1/13.5) the drag
forces are very much smaller than those with the heavier camber of 1/7,
and the negative drifts (d) are correspondingly smaller. All of the drag
due to the lower surface, point to the left (D), and hence produce head
resistance to flight. The drag to the rear of the center of the upper
surface are the same. In front of the upper center we have right hand,
or negative drifts (d), that aid the motion. These forward forces
obtained by experiment prove the correctness of Lilienthal’s "Forward
Tangential" theory advanced many years ago.
[Illustration: Fig. 13. Direction of Drag Over Different Portions of
Circular Arc - Aerofoils.]
[Illustration: Fig. 14a. Distribution for Wright Wing. (b) M. Farman
Wing. (c) Breguet Wing. (d) Bleriot Wing. (e) Bleriot 11-Bis.]
[Illustration: Fig. 15. Pressure Distribution at Various Points Along
the Length of a Nieuport Monoplane Wing.]
*Distribution on Practical Wings*. With the exception of the bird wing,
the distributions have been given for thin plates that are of little
value on an aeroplane. They do not permit of strong structural members
for carrying the load. The actual wing must have considerable thickness,
as shown by the aerofoils in Figs. 1, 2, 3, etc., and are of
approximately stream line form. Fig. 14 shows the distribution for
actual aeroplane wings: (a) Wright, (b) M. Farman, (c) Breguet, (d)
Bleriot 11.(d), (e) Bleriot 11-bis. The Wright wing is very blunt and
has an exceedingly high lift at the leading edge. The M. Farman, which
is slightly less blunt, has a similar but lower front peak. The Breguet
is of a more modern type with the maximum thickness about 25 per cent
from the leading edge. The latter shows a remarkably even distribution
of pressure, and is therefore a better type as will be seen from the
relative lifts of 0.916 and 0.986 pounds per square foot. The lift-drag
ratio of the Breguet is also better, owing to the greater predominance
of the negative drag components. Decreasing the thickness and the
undercamber of Bleriot 11, resulted in an unusual increase of 10 per
cent of the under pressure, and a decrease in the Suction, shown by
Bleriot 11-bis. The Bleriot has the sharpest entering edge and the least
upper pressure. In the above practical wing sections the aspect ratio is
variable, being the same in the test model as in the full-size machine.
The Bleriot being a monoplane has a lower aspect ratio (5), than the
biplanes (a), (b) and (c). The Breguet with an aspect of 8 has a lift of
0.986 pounds per square foot as against the 0.781 of the Bleriot, and
undoubtedly part of this difference is due to the aspect ratio. The
pressure falls off around the tips as shown by the successive sections
taken through a Nieuport monoplane wing in Fig. 15. Section (f) was
taken near the body and shows the greater lift. Section (g) is midway
between the tips and body, and (h) and (i) are progressively nearer the
tips. As we proceed toward the tips from the body the pressure falls off
as shown in the sections, this reducing from 1.07 to 0.55 pounds per
square foot. This wing also thins down toward the tips or "washes out,"
as it is called.
[Illustration: Fig. 16. Showing Pressure Distribution on the Plan View
of a Typical Wing, Leading Edge Along A-A, Trailing Edge D-C. Center of
Pressure. Marked "C.P." The Proportion Pressures Are Indicated by the
Shading on the Surface, the Pressure Being Negative at the Tips and Near
the Rear Edge.]
CHAPTER VI. PRACTICAL WING SECTIONS.
*Development of Modern Wings*. The first practical results obtained by
Wright Brothers, Montgomery, Chanute, Henson, Curtiss, Langley, and
others, were obtained by the use of cambered wings. The low value of the
lift-drag ratio, due to the flat planes used by the earlier
experimenters, was principally the cause of their failure to fly. The
Wrights chose wings of very heavy camber so that a maximum lift could be
obtained with a minimum speed. These early wings had the very fair
lift-drag ratio of 12 to 1. Modern wing sections have been developed
that give a lift-drag ratio of well over 20 to 1, although this is
attended by a considerable loss in the lift.
As before explained, the total lift of a wing surface depends on the
form of the wing, its area, and the speed upon which it moves in
relation to the air. Traveling at a low speed requires either a wing
with a high lift co-efficient or an increased area. With a constant
value for the lift-drag ratio, an increase in the lift value of the wing
section is preferable to an increase in area, since the larger area
necessitates heavier structural members, more exposed bracing, and
hence, more head resistance. Unfortunately, it is not always possible to
use the sections giving the heaviest lift, for the reason that such
sections usually have a poor lift-drag ratio. In the practical machine,
a compromise must be effected between the drag of the wings and the drag
or head resistance of the structural parts so that the combined or total
head resistance will be at a minimum. In making such a compromise, it
must be remembered that the head resistance of the structural parts
predominates at high speeds, while the drag of the wings is the most
important at low speeds.
In the early days of flying, the fact that an aeroplane left the ground
was a sufficient proof of its excellence, but nowadays the question of
efficiency under different conditions of flight (performance) is an
essential. Each new aeroplane is carefully tested for speed, rate of
climb, and loading. Speed range, or the relation between the lowest and
highest possible flight speeds, is also of increasing importance, the
most careful calculations being made to obtain this desirable quality.
*Performance*. To improve the performance of an aeroplane, the designer
must increase the ratio of the horsepower to the weight, or in other
words, must either use greater horsepower or decrease the weight carried
by a given power. This result may be obtained by improvements in the
motor, or by improvements in the machine itself. Improvements in the
aeroplane may be attained in several ways: (1) by cutting down the
structural weight; (2) by increasing the efficiency of the lifting
surfaces; (3) by decreasing the head resistance of the body and exposed
structural parts, and (4) by adjustment of the area or camber of the
wings so that the angle of incidence can be maintained at the point of
greatest plane efficiency. At present we are principally concerned with
item (2), although (4) follows as a directly related item.
Improvement in the wing characteristics is principally a subject for the
wind tunnel experimentalist, since with our present knowledge, it is
impossible to compute the performance of a wing by direct mathematical
methods. Having obtained the characteristics of a number of wing
sections from the aerodynamic laboratory, the designer is in a position
to proceed with the calculation of the areas, power, etc. At present
this is rather a matter of elimination, or "survival of the fittest," as
each wing is taken separately and computed through a certain range of
performance.
*Wing Loading*. The basic unit for wing lift is the load carried per
unit of area. In English units this is expressed as being the weight in
pounds carried by a square foot of the lifting surface. Practically,
this value is obtained by dividing the total loaded weight of the
machine by the wing area. Thus, if the weight of a machine is 2,500
pounds (loaded), and the area is 500 square feet, the "unit loading"
will be: w = 2,500/500 = 5 pounds per square foot. In the metric system
the unit loading is given in terms of kilogrammes per square meter.
Conversely, with the total weight and loading known, the area can be
computed by dividing the weight by the unit loading. The unit loading
adopted for a given machine depends upon the type of machine, its speed,
and the wing section adopted, this quantity varying from 3.5 to 10
pounds per square foot in usual practice. As will be seen, the loading
is higher for small fast machines than for the slower and larger types.
A very good series of wings has been developed, ranging from the low
resistance type carrying 5 pounds per square foot at 45 miles per hour,
to the high lift wing, which gives a lift of 7.5 pounds per square foot
at the same speed. The medium lift wing will be assumed to carry 6
pounds per square foot at 45 miles per hour. The wing carrying 7.5
pounds per square foot gives a great saving in area over the low lift
type at 5 pounds per square foot, and therefore a great saving in
weight. The weight saved is not due to the saving in area alone, but is
also due to the reduction in stress and the corresponding reduction in
the size and weight of the structural members. Further, the smaller area
requires a smaller tail surface and a shorter body. A rough
approximation gives a saving of 1.5 pounds per square foot in favor of
the 7.5 pound wing loading. This materially increases the horsepower
weight ratio in favor of the high lift wing, and with the reduction in
area and weight comes an improvement in the vision range of the pilot
and an increased ease in handling (except in dives). The high lift types
in a dive have a low limiting speed.
As an offset to these advantages, the drag of the high lift type of wing
is so great at small angles that as soon as the weight per horsepower is
increased beyond 18 pounds we find that the speed range of the low
resistance type increases far beyond that of the high lift wing.
According to Wing Commander Seddon, of the English Navy, a scout plane
of the future equipped with low resistance wings will have a speed range
of from 50 to 150 miles per hour. The same machine equipped with high
lift wings would have a range of only 50 to 100 miles per hour. An
excess of power is of value with low resistance wings, but is
increasingly wasteful as the lift co-efficient is increased. Landing
speeds have a great influence on the type of wing and the area, since
the low speeds necessary for the average machines require a high lift
wing, great area, or both. With the present wing sections, low flight
speeds are obtained with a sacrifice in the high speed values. In the
same way, high speed machines must land at dangerously high speeds. At
present, the best range that we can hope for with fixed areas is about
two to one; that is, the high speed is not much more than twice the
lowest speed. A machine with a low speed of 45 miles per hour cannot be
depended upon to safely develop a maximum speed of much over 90 miles
per hour, for at higher speeds the angle of incidence will be so
diminished as to come dangerously near to the position of no lift. In
any case, the travel of the center of pressure will be so great at
extreme wing angles as to cause considerable manipulation of the
elevator surface, resulting in a further increase in the resistance.
*Resistance and Power*. The horizontal drag (resistance) of a wing,
determines the power required for its support since this is the force
that must be overcome by the thrust of the propeller. The drag is a
component of the weight supported and therefore depends upon the loading
and upon the efficiency of the wing. The drag of the average modern
wing, structural resistance neglected, is about 1/16 of the weight
supported, although there are several sections that give a drag as low
as 1/23 of the weight. The denominators of these fractions, such as "16"
and "23," are the lift-drag ratios of the wing sections.
Drag in any wing section is a variable quantity, the drag varying with
the angle of incidence. In general, the drag is at a minimum at an angle
of about 4 degrees, the value increasing rapidly on a further increase
or decrease in the angle. Usually a high lift section has a greater drag
than the low lift type at small angles, and a smaller drag at large
angles, although this latter is not invariably the case.
*Power Requirements*. Power is the rate of doing work, or the rate at
which resistance is overcome. With a constant resistance the power will
be increased by an increase in the speed. With a constant speed, the
power will be increased by an increase in the resistance. Numerically,
the power is the product of the force and the velocity in feet per
second, feet per minute, miles per hour, or meters per second. The most
common English power unit is the "horsepower," which is obtained by
multiplying the resisting force in pounds by the velocity in feet per
minute, this product being divided by 33,000. If D is the horizontal
drag in pounds, and v = velocity of the wing in feet per minute, the
horsepower H will be expressed by:
H = Dv / 33,000
Since the speed of an aeroplane is seldom given in feet per minute, the
formula for horsepower can be given in terms of miles per hour by:
H = DV / 375
Where V = velocity in miles per hour, D and H remaining as before. The
total power for the entire machine would involve the sum of the wing and
structural drags, with D equal to the total resistance of the machine.
*Example*. The total weight of an aeroplane is found to be 3,000 pounds.
The lift-drag ratio of the wings is 15.00, and the speed is 90 miles per
hour. Find the power required for the wings alone.
*Solution*. The total drag of the wings will be: D = 3,000/15 = 200
pounds. The horsepower required: H = DV/375 = 200 × 90/375=48
horsepower. It should be remembered that this is the power absorbed by
the wings, the actual motor power being considerably greater owing to
losses in the propeller. With a propeller efficiency of 70 per cent, the
actual motor power will become: Hm = 48/0.70=68.57 for the wings alone.
To include the efficiency into our formula, we have:
H = DV/375E
where E = propeller efficiency expressed as a decimal. The greater the
propeller efficiency, the less will be the actual motor power, hence the
great necessity for an efficient propeller, especially in the case of
pusher type aeroplanes where the wings do not gain by the increased slip
stream.
The propeller thrust must be equal and opposite to the drag at the
various speeds, and hence the thrust varies with the plane loading, wing
section, and angle of incidence. Portions of the wing surfaces that lie
in the propeller slip stream have a greater lift than those lying
outside of this zone because of the greater velocity of the slip stream.
For accurate results, the area in the slip stream should be determined
and calculated for the increased velocity.
Oftentimes it is desirable to obtain the "Unit drag"; that is, the drag
per square foot of lifting surface. This can be obtained by dividing the
lift per square foot by the lift-drag ratio, care being taken to note
the angle at which the unit drag is required.
*Advantages of Cambered Sections Summarized*. Modern wing sections are
always of the cambered, double-surface type for the following reasons:
1. They give a better lift-drag ratio than the flat surface, and
therefore are more economical in the use of power.
2. In the majority of cases they give a better lift per square foot of
surface than the flat plate and require less area.
3. The cambered wings can be made thicker and will accommodate heavier
spars and structural members without excessive head resistance.
*Properties of Modern Wings*. The curvature of a wing surface can best
be seen by cutting out a section along a line perpendicular to the
length of the wing, and then viewing the cut portion from the end. It is
from this method of illustration that the different wing curves, or
types of wings, are known as "wing sections." In all modern wings the
top surface is well curved, and in the majority of cases the bottom
surface is also given a curvature, although this is very small in many
instances.
Fig. 1. shows a typical wing section with the names of the different
parts and the methods of dimensioning the curves. All measurements to
the top and bottom surfaces are taken from the straight "chordal line"
or "datum line" marked X-X. This line is drawn across the concave
undersurface in such a way as to touch the surface only at two points,
one at the front and one at the rear of the wing section. The
inclination of the wing with the direction of flight is always given as
the angle made by the line X-X with the wind. Thus, if a certain wing is
said to have an angle of incidence equal to 4 degrees, we know that the
chordal line X-X makes an angle of 4 degrees with the direction of
travel. This angle is generally designated by the letter (i), and is
also known as the "angle of attack." The distance from the extreme front
to the extreme rear edge (width of wing) is called the "chord width" or
more commonly "the chord."
In measuring the curve, the datum line X-X is divided into a number of
equal parts, usually 10, and the lines 1-2-3-4-5-6-7-8-9-10-11 are drawn
perpendicular to X-X. Each of the vertical numbered lines is called a
"station," the line No. 3 being called "Station 3," and so on. The
vertical distance measured from X-X to either of the curves along one of
the station lines is known as the "ordinate" of the curve at that point.
Thus, if we know the ordinates at each station, it is a simple matter to
draw the straight line X-X, divide it into 10 parts, and then lay off
the heights of the ordinates at the various stations. The distances from
datum to the upper curve are known as the "Upper ordinates," while the
same measurements to the under surface are known as the "Lower
ordinates." This method allows us to quickly draw any wing section from
a table that gives the upper and lower ordinates at the different
stations.
A common method of expressing the value of the depth of a wing section
in terms of the chord width is to give the "Camber," which is
numerically the result obtained by dividing the depth of the wing curve
at any point by the width of the chord. Usually the camber given for a
wing is taken to be the maximum camber; that is, the camber taken at the
point of greatest depth. Thus, if we hear that a certain wing has a
camber of 0.089, we take it for granted that this is the camber at the
deepest portion of the wing. The correct method would be to give 0.089
as the "maximum camber" in order to avoid confusion. To obtain the
maximum camber, divide the maximum ordinate by the chord.
[Illustration: Fig. 1. Section Through a Typical Aerofoil or Wing, the
Parts and Measurements Being Marked on the Section. The Horizontal Width
or "Chord" Is Divided Into 10 Equal Parts or "Stations," and the Height
of the Top and Bottom Curves Are Measured from the Chordal Line X-X at
Each Station. The Vertical Distance from the Chordal Line Is the
"Ordinate" at the Point of Measurement.]
*Example*. The maximum ordinate of a certain wing is 5 inches, and the
chord is 40 inches. What is the maximum camber? The maximum camber is
5/40 = 0.125. In other words, the maximum depth of this wing is 12.5 per
cent of the chord, and unless otherwise specified, is taken as being the
camber of the top surface.
The maximum camber of a modern wing is generally in the neighborhood of
0.08, although there are several Successful sections that are well below
this figure. Unfortunately, the camber is not a direct index to the
value of a wing, either in regard to lifting ability or efficiency. By
knowing the camber of a wing we cannot directly calculate the lift or
drag, for there are several examples of wings having widely different
cambers that give practically the same lift and drift. At the present
time, we can only determine the characteristics of a wing by experiment,
either on a full size wing or on a scale model.
In the best wing sections, the greatest thickness and camber occurs at a
point about 0.3 of the chord from the front edge, this edge being much
more blunt and abrupt than the portions near the trailing edge. An
efficient wing tapers very gradually from the point of maximum camber
towards the rear. This is usually a source of difficulty from a
structural standpoint since it is difficult to get an efficient depth of
wing beam at a point near the trailing edge. A number of experiments
performed by the National Physical Laboratory show that the position of
maximum ordinate or camber should be located at 33.2 per cent from the
leading edge. This location gives the greatest lift per square foot, and
also the least resistance for the weight lifted. Placing the maximum
ordinate further forward is worse than placing it to the rear.
Thickening the entering edge causes a proportionate loss in efficiency.
Thickening the rear edge also decreases the efficiency but does not
affect the weight lifting value to any great extent. The camber of the
under surface seems to have but little effect on the efficiency, but the
lift increases slightly with an increase in the camber of the lower
surface. Increasing the camber of the lower surface decreases the
thickness of the wing and hence decreases the strength of the supporting
members, particularly at points near the trailing edge. The increase of
lift due to increasing the under camber is so slight as to be hardly
worth the sacrifice in strength. Variations in the camber of the upper
surface are of much greater importance. It is on this surface that the
greater part of the lift takes places, hence a change in the depth of
this curve, or in its outline, will cause wider variations in the
characteristics of the wing than would be the case with the under
surface. Increasing the upper camber by about 60 per cent may double the
lift of the upper surface, but the relation of the lift to the drag is
increased. From this, it will be seen that direct calculations from the
outline would be most difficult, and in fact a practical impossibility
at the present time.
[Illustration: Fig. 2. (Upper). Shows a Slight "Reflex" or Upward Turn
of the Trailing Edge. Fig. 3. (Lower) Shows an Excessive Reflex Which
Greatly Reduces the C.P. Movement.]
By putting a reverse curve in the trailing edge of a wing, as shown by
Fig. 2, the stability of the wing may be increased to a surprising
degree, but the lift and efficiency are correspondingly reduced with
each increase in the amount of reverse curvature. In this way, stability
is attained at the expense of efficiency and lifting power. With the
rear edge raised about 0.037 of the chord, the N. P. L. found that the
center of pressure could be held stationary, but the loss of lift was
about 25 per cent and the loss of efficiency amounted practically 12 per
cent. With very slight reverse curvatures it has been possible to
maintain the lift and efficiency, and at the same time to keep the
center of pressure movement down to a reasonable extent. The New U.S.A.
sections and the Eiffel No. 32 section are examples of excellent
sections in which a slight reverse or "reflex" curvature is used. The
Eiffel 32 wing is efficient, and at the same the center of pressure
movement between incident angles of 0° and 10° is practically
negligible. This wing is thin in the neighborhood of the trailing edge,
and it is very difficult to obtain a strong rear spar.
*Wing Selection*. No single wing section is adapted to all purposes.
Some wings give a great lift but are inefficient at small angles and
with light loading. There are others that give a low lift but are very
efficient at the small angles used on high speed machines. As before
explained, there are very stable sections that give but poor results
when considered from the standpoint of lift and efficiency. The
selection of any one wing section depends upon the type of machine upon
which it is to be used, whether it is to be a small speed machine or a
heavy flying boat or bombing plane.
There are a multitude of wing sections, each possessing certain
admirable features and also certain faults. To list all of the wings
that have been tried or proposed would require a book many times the
size of this, and for this reason I have kept the list of wings confined
to those that have been most commonly employed on prominent machines, or
that have shown evidence of highly desirable and special qualities. This
selection has been made with a view of including wings of widely varying
characteristics so that the data can be applied to a wide range of
aeroplane types. Wings suitable for both speed and weight carrying
machines have been included.
The wings described are the U.S.A. Sections No. 1, 2, 3, 4, 5 and 6; the
R.A.F. Sections Nos. 3 and 6, and the well known Eiffel Wings No. 32, 36
and 37. The data given for these wings is obtained from wind tunnel
tests made at the Massachusetts Institute of Technology, the National
Physical Laboratory (England), and the Eiffel Laboratory in Paris. For
each of these sections the lift co-efficient (Ky), the lift-drift ratio
(L/D), and the drag co-efficient (Kx) are given in terms of miles per
hour and pounds per square foot. Since these are the results for model
wings, there are certain corrections to be made when the full size wing
is considered, these corrections being made necessary by the fact that
the drag does not vary at the same rate as the lift. This "Size" or
"Scale" correction is a function of the product of the wing span in feet
by the velocity of the wind in feet per second. A large value of the
product results in a better wing performance, or in other words, the
large wing will always give better lift-drag ratios than would be
indicated by the model tests. The lift co-efficient Ky is practically
unaffected by variations in the product. If the model tests are taken
without correction, the designer will always be on the safe side in
calculating the power. The method of making the scale corrections will
be taken up later.
Of all the sections described, the R.A.F.-6 is probably the best known.
The data on this wing is most complete, and in reality it is a sort of
standard by which the performance of other wings is compared. Data has
been published which describes the performance of the R.A.F.-6 used in
monoplane, biplane and triplane form; and with almost every conceivable
degree of stagger, sweep back, and decalage. In addition to the
laboratory data, the wing has also been used with great success on full
size machines, principally of the "Primary trainer" class where an "All
around" class of wing is particularly desirable. It is excellent from a
structural standpoint since the section is comparatively deep in the
vicinity of the trailing edge. The U.S.A. sections are of comparatively
recent development and are decided improvements on the R.A.F. and Eiffel
sections. The only objection is the limited amount of data that is
available on these wings—limited at least when the R.A.F. data is
considered—as we have only the figures for the monoplane arrangement.
*WING SELECTION.*
*(1) Lift-Drag Ratio*. The lift-drag ratio (L/D) of a wing is the
measure of wing efficiency. Numerically, this is equal to the lift
divided by the horizontal drag, both quantities being expressed in
pounds. The greater the weight supported by a given horizontal drag, the
less will be the power required for the propulsion of the aeroplane,
hence a high value of L/D indicates a desirable wing section—at least
from a power standpoint. In the expression L/D, L = lift in pounds, and
D = horizontal drag in pounds. Unfortunately, this is not the only
important factor, since a wing having a great lift-drag is usually
deficient in lift or is sometimes structurally weak.
The lift-drag ratio varies with the angle of incidence (i), reaching a
maximum at an angle of about 4° in the majority of wings. The angle of
incidence at which the lift-drag is a maximum is generally taken as the
angle of incidence for normal horizontal flight. At angles either
greater or less, the L/D falls off, generally at a very rapid rate, and
the power increases correspondingly. Very efficient wings may have a
ratio higher than L/D=20 at an angle of about 4°, while at 16° incidence
the value may be reduced to L/D = 4, or even less. The lift is generally
greatest at about 16°. The amount of variation in the lift, and
lift-drag, corresponding to changes in the incidence differs among the
different types of wings and must be determined by actual test.
After finding a wing with a good value of L/D, the value of the lift
co-efficient Ky should be determined at the angle of the maximum L/D.
With two wings having the same lift-drag ratio, the wing having the
greatest lift (Ky) at this point is the most desirable wing as the
greater lift will require less area and will therefore result in less
head resistance and less weight. Any increase in the area not only
increases the weight of the wing surface proper, but also increases the
wiring and weights of the structural members. With heavy machines, such
as seaplanes or bomb droppers, a high value of Ky is necessary if the
area is to be kept within practical limits. A small fast scouting plane
requires the best possible lift-drag ratio at small angles, but requires
only a small lift co-efficient. At speeds of over 100 miles per hour a
small increase in the resistance will cause a great increase in the
power.
*(2) Maximum Lift (Ky)*. With a given wing area and weight, the maximum
value of the lift co-efficient (Ky) determines the slow speed, or
landing speed, of the aeroplane. The greater the value of Ky, the slower
can be the landing speed. For safety, the landing speed should be as low
as possible.
In the majority of wings, the maximum lift occurs at about 16° of
incidence, and in several sections this maximum is fairly well sustained
over a considerable range of angle. The angle of maximum lift is
variously known as the "Stalling angle" or the "Burble point," since a
change of angle in either direction reduces the lift and tends to stall
the aeroplane. For safety, the angle range for maximum lift should be as
great as possible, for if the lift falls off very rapidly with an
increase in the angle of incidence, the pilot may easily increase the
angle too far and drop the machine. In the R.A.F.-3 wing, the lift is
little altered through an angle range of from 14° to 16.5°, the maximum
occurring at 15.7°, while with the R.A.F.-4, the lift drops very
suddenly on increasing the angle above 15°. The range of the stalling
angle in any of the wings can be increased by suitable biplane or
triplane arrangements. If large values of lift are accompanied by a
fairly good L/D value at large angles, the wing section will be suitable
for heavy machines.
*(3) Center of Pressure Movement*. The center of pressure movement with
varying angles of incidence is of the greatest importance, since it not
only determines the longitudinal stability but also has an important
effect upon the loading of the wing spars and ribs. With the majority of
wings a decrease in the angle of incidence causes the center of pressure
to move back toward the trailing edge and hence tends to cause nose
diving. When decreased beyond 0° the movement is very sharp and quick,
the C. P. moving nearly half the chord width in the change from 0° to
-1.5°. The smaller the angle, the more rapid will be the movement.
Between 6° and 16°, the center of pressure lies near a point 0.3 of the
chord from the entering edge in the majority of wing sections. Reducing
the angle from 6° to 2° moves the C. P. back to approximately 0.4 of the
chord from the entering edge.
There are wing sections, however, in which the C. P. movement is
exceedingly small, the Eiffel 32 being a notable example of this type.
This wing is exceedingly stable, as the C. P. remains at a trifle more
than 0.30 of the chord through nearly the total range of flight angles.
An aeroplane equipped with the Eiffel 32 wing could be provided with
exceedingly small tail surfaces without a tendency to dive. Should the
elevator become inoperative through accident, the machine could probably
be landed without danger. This wing has certain objectionable features,
however, that offset the advantages.
It will be noted that with the unstable wings the center of pressure
movement always tends to aggravate the wing attitude. If the machine is
diving, the decrease in angle causes the C. P. to move back and still
further increase the diving tendency. If the angle is suddenly
increased, the C. P. moves forward and increases the tendency toward
stalling.
If the center of pressure could be held stationary at one point, the
wing spars could be arranged so that each spar would take its proper
proportion of the load. As it is, either spar may be called upon to
carry anywhere from three-fourths of the load to entire load, since at
extreme angles the C. P. is likely to lie directly on either of the
spars. Since the rear spar is always shallow and inefficient, this is
most undesirable. This condition alone to a certain extent
counterbalances the structural disadvantage of the thin Eiffel 32
section. Although the spars in this wing must of necessity be shallow,
they can be arranged so that each spar will take its proper share of the
load and with the assurance that the loading will remain constant
throughout the range of flight angles. The comparatively deep front spar
could be moved back until it carried the greater part of the load, thus
relieving the rear spar.
With a good lift-drag ratio, and a comparatively high value of Ky, the
center of pressure movement should be an important consideration in the
selection of a wing. It should be remembered in this regard that the
stability effects of the C. P. movement can be offset to a considerable
extent by suitable biplane arrangements.
*(4) Structural Considerations*. For large, heavy machines, the
structural factor often ranks in importance with the lift-drag ratio and
the lift co-efficient. It is also of extreme importance in speed scouts
where the number of interplane struts are to be at a minimum and where
the bending moment on the wing spars is likely to be great in
consequence. A deep, thick wing section permits of deep strong wing
spars. The strength of a spar increases with the square of its depth,
but only in direct proportion to its width. Thus, doubling the depth of
the spar increases the strength four times, while doubling the width
only doubles the strength. The increase in weight would be the same in
both cases.
While very deep wings are not usually efficient, when considered from
the wing section tests alone, the total efficiency of the wing
construction when mounted on the machine is greater than would be
supposed. This is due to the lightness of the spars and to the reduction
in head resistance made possible by a greater spacing of the interplane
struts. Thus, the deep wing alone may have a low L/D in a model test,
but its structural advantages give a high total efficiency for the
machine assembled.
*Summary*. It will be seen from the foregoing matter that the selection
of a wing consists in making a series of compromises and that no single
wing section can be expected to fulfill all conditions. With the purpose
of the proposed aeroplane thoroughly in mind, the various sections are
taken up one by one, until a wing is found that most usefully
compromises with all of the conditions. Reducing this investigation to
its simplest elements we must follow the routine as described above: (1)
Lift-drift ratio and value of Ky at this ratio. (2) Maximum value of Ky
and L/D at this lift. (3) Center of pressure movement. (4) Depth of wing
and structural characteristics.
*Calculations for Lift and Area*. Although the principles of surface
calculations were described in the chapter on elementary aerodynamics,
it will probably simplify matters to review these calculations at this
point. The lift of a wing varies with the product of the area, and the
velocity squared, this result being multiplied by the co-efficient of
lift (Ky). The co-efficient varies with the wing section, and with the
angle of incidence. Stated as a formula: L = KyAV² where A = area in
square feet, and V = velocity of the wing in miles per hour. Assuming an
area of 200 square feet, a velocity of 80 miles per hour, and with K =
0.0025, the total life (L) becomes: L= KyAV² =0.0025 x 200 x (80 × 80) =
3,200 pounds. Assuming a lift-drag ratio of 16, the "drag" of the wing,
or its resistance to horizontal motion, will be expressed by D =
L/r=3,200/16=200 pounds, where r = lift-drag ratio. It is this
resistance of 200 pounds that the motor must overcome in driving the
wings through the air. The total resistance offered by the aeroplane
will be equal to the sum of the wing resistance and the head resistance
of the body, struts, wiring and other structural parts. In the present
instance we will consider only the resistance of the wings.
When the lift co-efficient, speed, and total lift are known, the area
can be found from A = L/KyV², the lift, of course, being taken as the
total weight of the machine. The area of the supporting surface for a
speed of 60 miles per hour, total weight of 2,400 pounds, and a lift
co-efficient of 0.002 is calculated as follows:
A = L/KyV² = 2,400/0.002 × (60 × 60) = 333 sq. ft.
A third variation in the formula is that used in finding the value of
the lift co-efficient for a particular wing loading. From the weight,
speed and area, we can find the co-efficient Ky, and with this value we
can find a wing that will correspond to the required co-efficient. This
method is particularly convenient when searching for the section with
the greatest lift-drag ratio. Ky = L/AV², or when the loading per square
foot is known, the co-efficient becomes Ky = L’/V². For example, let us
find the co-efficient for a wing loading of 5 pounds per square foot at
a velocity of 80 miles per hour. Inserting the numerical values into the
equation we have, Ky = L’/V² =5/(80 × 80) = 0.00078. Any wing, at any
angle that has a lift co-efficient equal to 0.00078 will support the
load at the given speed, although many of the wings would not give a
satisfactory lift-drag ratio with this co-efficient.
It should be noted in the above calculations that no correction has been
made for "Scale," aspect ratio or biplane interference. In other words,
we have assumed the figures as applying to model monoplanes. In the
following tables the lift, lift-drag and drag must be corrected, since
this data was obtained from model tests on monoplane sections. The
effects of biplane interference will be described in the chapter on
"Biplane and Triplane Arrangement," but it may be stated that
superposing the planes reduces both the lift co-efficient and the
liftdrag ratio, the amount of reduction depending upon the relative gap
between the surfaces. Thus with a gap equal to the chord, the lift of
the biplane surface will only be about 80 per cent of the lift of a
monoplane surface of the same area and section.
*Wing Test Data*. The data given in this chapter is the result of wind
tunnel tests made under standard conditions, the greater part of the
results being published by the Massachusetts Institute of Technology.
The tests were all made on the same size of model and at the same wind
speed so that an accurate comparison can be made between the different
sections. All values are for monoplane wings with an aspect ratio of 6,
the laboratory models being 18x3 inches. The exception to the above test
conditions will be found in the tables of the Eiffel 37 and 36 sections,
these figures being taken from the results of Eiffel’s laboratory. The
Eiffel models were 35.4x5.9 inches and were tested at wind velocities of
22.4, 44.8, and 67.2 miles per hour. The tests made at M. I. T. were all
made at a wind speed of 30 miles per hour. The lift co-efficient Ky is
practically independent of the wing size and wind velocity, but the drag
co-efficient Kx varies with both the size and wind velocity, and the
variation is not the same for the different wings. The results of the M.
I. T. tests were published in "Aviation and Aeronautical Engineering" by
Alexander Klemin and G. M. Denkinger.
*The R.A.F. Wing Sections*. These wings are probably the best known of
all wings, although they are inferior to the new U.S.A. sections. They
are of English origin, being developed by the Royal Aircraft Factory
(R.A.F.), with the tests performed by the National Physical Laboratory
at Teddington, England. The R.A.F.-6 is the nearest approach to the all
around wing, this section having a fairly high L/D ratio and a good
value of Ky for nearly all angles. It is by no means a speed wing nor is
it suitable for heavy machines, but it comprises well between these
limits and has been extensively used on medium size machines, such as
the Curtiss JN4–B, the London and Provincial, and others. The R.A.F-3
has a very high value for Ky, and a very good lift-drag ratio for the
high-lift values. It is suitable for seaplanes, bomb droppers and other
heavy machines of a like nature that fly at low or moderate speeds. The
outlines of these wings are shown by Figs. 7 and 8, and the camber
ordinates are marked as percentages of the chord. In laying out a wing
rib from these diagrams, the ordinate at any point is obtained by
multiplying the chord length in inches by the ordinate factor at that
point. Referring to the R.A.F.-3 diagram, Fig. 8, it will be seen that
the ordinate for the upper surface at the third station from the
entering edge is 0.064. If the chord of the wing is 60 inches, the
height of the upper curve measured above the datum line X-X at the third
station will be, 0.064 × 60 = 3.84 inches. At the same station, the
height of the lower curve will be, 0.016 × 60= 0.96 inch.
The chord is divided into 10 equal parts, and at the entering edge one
of the ten parts is subdivided so as to obtain a more accurate curve at
this point. In some wing sections it is absolutely necessary to
subdivide the first chord division as the curve changes very rapidly in
a short distance. The upper curve, especially at the entering edge, is
by far the most active part of the section and for this reason
particular care should be exercised in getting the correct outline at
this point.
[Illustration: Figs. 7-8. R.A.F. Wing Sections. Ordinates as Decimals of
the Chord.]
*Aerodynamic Properties of the R.A.F. Sections*. Table 1 gives the
values of Ky, Kx, L/D, and the center of pressure movement (C. P.) for
the R.A.F.-3 section through a range of angles varying from -2° to 20°.
The first column at the left gives the angles of incidence (i), the
corresponding values for the lift (Ky) and the drag (Kx) being given in
the second and third columns, respectively. The fourth column gives the
lift-drag ratio (L/D). The fifth and last column gives the location of
the center of pressure for each different angle of incidence, the figure
indicating the distance of the C. P. from the entering edge expressed as
a decimal part of the chord. As an example in the use of the table, let
it be required to find the lift and drag of the R.A.F.-3 section when
inclined at an angle of 6° and propelled at a speed of 90 miles per
hour. The assumed area will be 300 square feet. At 6° it will be found
that the lift co-efficient Ky is 0.002369. From our formulae, the lift
will be: L = KyAV² or numerically, L=0.002369 × 300 × (90 × 90) = 5,7567
lbs. At the same speed, but with the angle of incidence reduced to 2°,
the lift will be reduced to L = 0.001554 × 300 x (90 × 90) = 3,776.2
pounds, where 0.001554 is the lift co-efficient at 2°. It will be noted
that the maximum lift co-efficient occurs at 14° and continues at this
value to a little past 15°. The lift at the stalling angle is fairly
constant from 12° to 16°.
The value of the drag can be found in either of two ways: (1) by
dividing the total lift (L) by the lift-drag ratio, or (2) by figuring
its value by the formula D = KxAV². The first method is shorter and
preferable. By consulting the table, it will be seen that the L/D ratio
at 6° is 14.9. The total wing drag will then be equal to 5,756.7/149 =
386.4 lbs. Figured by the second method, the value of Ky at 6° is
0.000159, and the drag is therefore: D = KxAV² = 0.000159 × 300 × (90 ×
90) = 386.4. This checks exactly with the first method. The lift-drag
ratio is best at 4°, the figure being 15.6, while the lift at this point
is 0.001963. With the same area and speed, the total lift of the surface
at the angle of best lift-drift ratio will be 0001963 × 300 × (90 × 90)
= 4,770 lbs.
At 4° the center of pressure is 0.385 of the chord from the entering
edge. If the chord is 60 inches wide, the center of pressure will be
located at 0.385 × 60 = 23.1 inches from the entering edge. At 15°, the
center of pressure will be 0.29 × 60 = 17.4 inches from the entering
edge, or during the change from 4° to 15° the center of pressure will
have moved forward by 5.7 inches. At -2°, the pressure has moved over
three-quarters of the way toward the trailing edge -0.785 of the chord,
to be exact Through the ordinary flight angles of from 2° to 12°, the
travel of the center of pressure is not excessive.
The maximum lift co-efficient (Ky) is very high in the R.A.F.-3 section,
reaching a maximum of 0.003481 at an incidence of 14°. This is second to
only one other wing, the section U.S.A.-4. This makes it suitable for
heavy seaplanes.
Table 2 gives the aerodynamic properties of the R.A.F-6 wing, the table
being arranged in a manner similar to that of the R.A.F.-3. In glancing
down the column of lift co-efficients (Ky), and comparing the values
with those of the R.A.F.-3 section, it will be noted that the lift of
R.A.F.-6 is much lower at every angle of incidence, but that the
lift-drag ratio of the latter section is not always correspondingly
higher. At every angle below 2°, at 6°, and at angles above 14°, the L/D
ratio of the R.A.F.-3 is superior in spite of its greater lift. The
maximum L/D ratio of the R.A.F.-6 at 4° is 16.58, which is considerably
higher than the best L/D ratio of the R.A.F.-3. The best lift
co-efficient of the R.A.F.-6, 0003045, is very much lower than the
maximum Ky of the R.A.F.-3.
The fact that the L/D ratio of the R.A.F.-3 wing is much greater at high
lift co-efficients, and large angles of incidence, makes it very
valuable as at this point the greater L/D does not tend to stall the
plane at slow speed. A large L/D at great angles, together with a wide
stalling angle tends for safety in slow speed flying.
Both wing sections are structurally excellent, being very deep in the
region of the rear edge, the R.A.F.-6 being particularly deep at this
point. A good deep spar can be placed at almost any desirable point in
the R.A.F-6, and the trailing edge is deep enough to insure against rib
weakness even with a comparatively great overhang.
Scale corrections for the full size R.A.F. wings are very difficult to
make. According to the N. P. L. reports, the corrected value for the
maximum L/D of the R.A.F.-3 wing is 18.1, the model test indicating a
maximum value of 15.6. I believe that L/D = 17.5 would be a safe full
size value for this section. The same reports give the full size L/D for
the R.A.F.-6 as 18.5, which would be probably safe at 18.0 under the new
conditions.
*Properties of the Eiffel Sections (32-36-37)*. Three of the Eiffel
sections are shown by Figs. 10, 11 and 12, these Sections being selected
out of an enormous number tested in the Eiffel laboratories. They differ
widely, both aerodynamically and structurally, from the R.A.F.
aerocurves just illustrated.
[Illustration: Fig. 10-11-12 Ordinates for Three Eiffel Wing Sections]
Eiffel 32 is a very stable wing, as has already been pointed out, but
the value of the maximum L/D ratio is in doubt as this quantity is very
susceptible to changes in the wind velocity—much more than in the
average wing. Since Eiffel’s tests were carried out at much higher
velocity than at the M. I. T., his lift-drift values at the higher
speeds were naturally much better than those obtained by the American
Laboratory. When tested at 67.2 miles per hour the lift-drift ratio for
the Eiffel 32 was 184 while at 22.4 miles per hour, the ratio dropped to
13.4. This test alone will give an idea as to the variation possible
with changes in scale and wind velocity. The following table gives the
results of tests carried out at the Massachusetts laboratory, reported
by Alexander Klemin and G. M. Denkinger in "Aviation and Aeronautical
Engineering." Wind speed, 30 miles per hour.
The C. P. Travel in the Eiffel wing is very small, as will be seen from
Table 3. At -2° the C. P. is 0.33 of the chord from the leading edge and
only moves back to 0.378 at an angle of 20°, the intermediate changes
being very gradual, reaching a minimum of 0.304 at 6° incidence. The
maximum Ky of Eiffel 32 is 0.002908, while for the R.A.F.-6 wing, Ky =
0.003045 maximum, both co-efficients being a maximum at 16° incidence,
but the lift-drag at maximum Ky is much better for the R.A.F.-6.
Structurally, the Eiffel 32 is at a disadvantage when compared with the
R.A.F. sections since it is very narrow at points near the trailing
edge. This would necessitate moving the rear spar well up toward the
center with the front spar located very near the leading edge. This is
the type of wing used in a large number of German machines. It will also
be noted that there is a very pronounced reverse curve or "Reflex" in
the rear portion, the trailing edge actually curving up from the chord
line.
Eiffel 36 is a much thicker wing than either of the other Eiffel curves
shown, and is deficient in most aerodynamical respects. It has a low
value for Ky and a poor lift-drag ratio. It has, however, been used on
several American training machines, probably for the reason that it
permits of sturdy construction.
[Illustration: Fig. 13 Characteristic Curves for Eiffel Wings Sections]
Eiffel 37 is essentially a high-speed wing having a high L/D ratio and a
small lift co-efficient. The maximum lift-drag ratio of 20.4 is attained
at a negative angle -08°. The value of Ky at this point is 0.00086, an
extremely low figure. The maximum Ky is 0.00288 at 14.0°, the L/D ratio
being 4.0 at this angle. Structurally it is the worst wing that we have
yet discussed, being almost "paper thin" for a considerable distance
near the trailing edge. The under surface is deeply cambered, with the
maximum under camber about one-third from the trailing edge. It is
impossible to use this wing without a very long overhang in the rear of
the section, and like the Eiffel 32, the front spar must be very far
forward. For those desiring flexible trailing edges, this is an ideal
section. This wing is best adapted for speed scouts and racing machines
because of its great L/D, but as its lift is small and the center of
pressure movement rapid at the point of maximum lift-drag, it would be
necessary to fly at a small range of angles and land at an extremely
high speed. Any slight change in the angle of incidence causes the
lift-drag ratio to drop at a rapid rate, and hence the wing could only
be manipulated at its most efficient angle by an experienced pilot.
Again, the angle of maximum L/D is only a few degrees from the angle of
no lift.
*U.S.A. Wing Sections*. These wing sections were developed by the
Aviation Section of the Signal Corps, United States Army, and are
decided improvements on any wing sections yet published. The six U.S.A.
wings cover a wide range of application, varying as they do, from the
high speed sections to the heavy lift wings used on large machines. The
data was first published by Captains Edgar S. Gorrell and H. S. Martin,
U.S.A., by permission of Professor C. H. Peabody, Massachusetts
Institute of Technology. An abstract of the paper by Alexander Klemin
and T. H. Huff was afterwards printed in "Aviation and Aeronautical
Engineering." While several of the curves are modifications of the
R.A.F. sections already described, they are aerodynamically and
structurally superior to the originals, and especial attention is called
to the marked structural advantages.
U.S.A.-1 and U.S.A.-6 are essentially high speed sections with a very
high lift-drag ratio, these wings being suitable for speed scouts or
pursuit machines. The difference between the wings is very slight,
U.S.A.-1 with K-000318 giving a better landing speed, while U.S.A.-6 is
slightly more efficient at low angles and high speeds.
[Illustration: Fig. 14. U.S.A. Wing Sections Nos. 1-2-3-4-5-6, Showing
the Ordinates at the Various Štations Expressed as Decimals of the
Chord. U.S.A.-4 is a Heavy Lift Section, While U.S.A.-1 and U.S.A.-6.are
High Speed Wings. For Any Particular Duty, the Above Wings Are Very Deep
and Permit of Large Structural Members. The Center of Pressure Movement
Is Comparatively Slight.]
With 0° incidence, the ratio of U.S.A-1=11.0 while the lift-drag of
U.S.A.-6 at 0° incidence is 13.0. The maximum lift of U.S.A.-1 is
superior to that of Eiffel 32, and the maximum lift-drag ratio at equal
speeds is far superior, being 17.8 against 14.50 of the Eiffel 32.
Compared with the Eiffel 32 it will be seen that the U.S.A. sections are
far better from a structural point of view, especially in the case of
U.S.A.-1. The depth in the region of the rear spar is exceptionally
great, about the same as that of the R.A.F.-6. While neither of the
U.S.A. wings are as stable as the Eiffel 32, the motion of the C. P. is
not sudden nor extensive at ordinary flight angles.
Probably one of the most remarkable of the United States Army wings is
the U.S.A.-4 which has a higher maximum lift co-efficient (Ky) than even
the R.A.F.-3. The maximum Ky of the U.S.A.-4 is 0.00364 compared with
the R.A.F.-3 in which Ky (Maximum)=0.003481. Above 4° incidence, the
lift-drag ratio of the U.S.A.-4 is generally better than that of the
R.A.F.-3, the maximum L/D at 4° being considerably better. This is a
most excellent wing for a heavy seaplane or bomber. The U.S.A.-2 has an
upper surface similar to that of the R.A.F.-3, but the wing has been
thickened for structural reasons, thus causing a modification in the
lower surface. This results in no particular aerodynamic loss and it is
much better at points near the rear edge for the reception of a deep and
efficient rear spar.
U.S.A.-3 is a modification of U.S.A.-2, and like U.S.A.-2 would fall
under the head of "All around wings," a type similar, but superior to
R.A.F.-6. These wings are a compromise between the high speed and heavy
lift types—suitable for training schools or exhibition flyers. Both have
a fairly good L/D ratio and a corresponding value for Ky.
U.S.A.-5 has a very good maximum lift-drag ratio (16.21) and a good
lift-drag ratio at the maximum Ky. Its maximum Ky is superior to all
sections with the exception of U.S.A.-2 and 4. Structurally it is very
good, being deep both fore and aft.
In review of the U.S.A. sections, it may be said that they are all
remarkable in having a very heavy camber on both the upper and lower
surfaces, and at the same time are efficient and structurally excellent.
This rather contradicts the usual belief that a heavy camber will
produce a low lift-drag ratio, a belief that is also proven false by the
excellent performance of the Eiffel 37 section. The maximum Ky is also
well sustained at and above 0.003. There is no sharp drop of lift at the
"Stalling angle" and the working range of incidence is large.
*Curtiss Wing and Double Cambered Sections*. An old type of Curtiss wing
is shown by Fig. 15. It is very thick and an efficient wing for general
use. It will be noticed that there is a slight reflex curve at the
trailing edge of the under surface and that there is ample spar room at
almost any point along the section. The nose is very round and thick for
a wing possessing the L/D characteristics exhibited in the tests. The
conditions of the test were the same as for the preceding wing sections.
Fig. 16 shows a remarkable Curtiss section designed for use as a
stabilizing surface. It is double cambered, the top surface being
identical with the lower, and is therefore non-lifting with the chord
horizontal. The force exerted by the surface is equal with equal
positive or negative angles of incidence, a valuable feature in a
control surface. In spite of its great thickness, it is of excellent
stream line form and therefore has a very good lift-drag ratio. At 0°
angle of incidence the resistance is at a minimum, and is much less than
that of a thin, square edged, flat plate. This double cambered plane
reduces the stay bracing and head resistance necessary with the flat
type of stabilizer surface.
[Illustration: Fig. 15. Old Type of Curtiss Wing. 16. Curtiss Double
Camber for Control Surfaces.]
The Curtiss sections mentioned above were described in "Aviation and
Aeronautical Engineering" by Dr. Jerome C. Hunsaker, but the figures in
the above table were obtained by the author on a sliding test wire
arrangement that has been under development for some time. At the time
of writing several of the U.S.A. sections are under investigation on the
same device.
CORRECTION FACTORS FOR WING FORM AND SIZE.
*Aspect Ratio*. As previously explained, the aspect ratio is the
relation of the span to the chord, and this ratio has a considerable
effect upon the performance of a wing. In the practical full size
machine the aspect ratio may range from 5 in monoplanes, and small
biplanes, to 10 or 12 in the larger biplanes. The aspect in the case of
triplanes is even greater, some examples of the latter having aspects of
16 to 20. In general, the aspect ratio increases with the gross weight
of the machine. Control surfaces, such as the rudder and elevator,
usually have a much lower aspect ratio than the main lifting surfaces,
particularly when flat non-lifting control surfaces are used. The aspect
of elevator surfaces will range from unity to 3, while the vertical
rudders generally have an aspect of 1.
With a given wing area, the span increases directly with an increase in
the aspect ratio. The additional weight of the structural members due to
an increased span tend to offset the aerodynamic advantages gained by a
large aspect ratio, and the increased resistance due to the number and
size of the exposed bracing still further reduces the advantage.
*Effects of Aspect Ratio*. Variations in the aspect ratio do not give
the same results in all wing sections, and the lift co-efficient and L/D
ratio change in a very irregular manner with the angle of incidence. The
following tables give the results obtained by the N. P. L. on a Bleriot
wing section, the aspect ratio being plotted against the angle of
incidence. The figures are comparative, an aspect factor of unity
(1,000) being taken for an aspect ratio of 6 at each angle of incidence.
To obtain an approximation for any other wing section at any other
aspect ratio, multiply the model test (Aspect=6) by the factor that
corresponds to the given angle and aspect ratio. At the extreme right of
the table is a column of rough averages, taken without regard to the
angles.
The column of average values is not the average of the tabular values
but is the average of the results obtained by a number of investigators
on different wing sections. Through the small angles of 0° and 2° the
low aspect ratios give a maximum Ky greater than with the larger
aspects. The larger aspects increase the lift through a larger range of
angles but have a lower maximum value for Ky at small angles. Beyond 2°
the larger aspect ratios give a greater Ky.
*Aspect for Flat Plates*. For flat plates the results are different than
with cambered sections. The lift-drag ratios are not much improved with
an increase in aspect, but the highest maximum lift is obtained with a
small aspect ratio. For this reason, a small aspect ratio should be used
when a high lift is to be obtained at low speeds with a flat plate as in
the case of control surfaces. An aspect ratio of unity is satisfactory
for flat vertical rudders since a maximum effect is desirable when
taxi-ing over the ground at low speeds. The flat plate effects are not
important except for control surfaces, and even in this case the plates
are being superseded by double cambered sections.
*Reason for Aspect Improvement*. The air flows laterally toward the wing
tips causing a very decided drop in lift at the outer ends of the wings.
The lift-drag ratio is also reduced at this point. The center of
pressure moves back near the trailing edge as we approach the tips, the
maximum zone of suction on the upper surface being also near the
trailing edge. The lift-drag ratio at the center of the plane is between
4 or 5 times that at a point near the tips. All of the desirable
characteristics of the wing are exhibited at a point near the center.
When the aspect ratio is increased, the inefficient tips form a smaller
percentage of the total wing areas, and hence the losses at the tips are
of less importance than would be the case with a small aspect. The end
losses are not reduced by end shields or plates, and in attempts to
prevent lateral flow by curtains, the losses are actually often
increased. Proper design of the form of the wing tip, such as raking the
tips, or washing out the camber and incidence, can be relied upon to
increase the lift factor. This change in the tips causes the main wind
stream to enter the wings in a direction opposite to the lateral leakage
flow and therefore reduces the loss. Properly raked tips may increase
the lift by 20 per cent.
*Effects of Scale (Size and Velocity)*. In the chapter "Elementary
Aerodynamics" it was pointed out that the lift of a surface was obtained
by the motion of the air, or the "turbulence" caused by the entering of
the plane. It was also explained that the effect of the lift due to
turbulence varied as the square of the velocity and directly as the area
of the wings. This would indicate that the lift of a small wing (Model)
would be in a fixed proportion to a large wing of the same type. This
holds true in practice since nearly all laboratories have found by
experiment that the lift of a large wing could be computed directly from
the results obtained with the model without the use of correction
factors. That is to say, that the lift of a large wing with 40 times the
area of the model, would give 40 times the lift of the model at the same
air speed. In the same way, the lift would be proportional to the
squares of the velocities. If the span of the model is taken at "1"
feet, and the velocity as V feet per second, the product IV would
represent both the model and the full size machine. The lift is due to
aerodynamic forces strictly, and hence there should be no reason why the
"V²" law should be interfered with in a change from the model to the
full size machine.
In the case of drag the conditions are different, since the drag is
produced by two factors that vary at different rates. Part of the drag
is caused by turbulence or aerodynamic forces and part by skin friction,
the former varying as V² while the skin friction varies as V¹.⁸⁸. The
aerodynamic drag varies directly with the area or span while the skin
friction part of the drag varies as 1⁰.⁹³, where 1 is the span. From
considerations of the span and the speed, it will be seen that the
frictional resistance increases much slower than the aerodynamic
resistance, and consequently the large machine at high speed would give
less drag and a higher value of L/D than the small model. In other
words, the results of a model test must be corrected for drag and the
lift-drag ratio when applied to a full size machine. Such a correction
factor is sometimes known as the "Scale factor."
Eiffel gives the correction factor as 1.08, that is the liftdrag ratio
of the full size machine will be approximately 1.08 times as great as
the model.
A series of full size tests were made by the University of St. Cyr in
1912-1913 with the object of comparing full size aeroplane wings with
small scale models of the same wing section. The full size wings were
mounted on an electric trolley car and the tests were made in the open
air. Many differences were noted when the small reproductions of the
wings were tested in the wind tunnel, and no satisfactory conclusions
can be arrived at from these tests. According to the theory, and the
tests made by the N. P. L., the lift-drag ratio should increase with the
size but the St. Cyr tests showed that this was not always the case. In
at least three of the tests, the model showed better results than the
full size machine. There seemed to be no fixed relation between the
results obtained by the model and the large wing. The center of pressure
movement was always different in every comparison made.
One cause of such pronounced difference would probably be explained by
the difference in the materials used on the model and full size wing,
the model wing being absolutely smooth rigid wood while the full size
wing was of the usual fabric construction. The fabric would be likely to
change in form under different conditions of angle and speed, causing a
great departure from the true values. Again, the model being of small
size, would be a difficult object to machine to the exact outline. A
difference of 1/1000 inch from the true dimension would make a great
difference in the results obtained with a small surface.
*Plan Form*. Wings are made nearly rectangular in form, with the ends
more or less rounded, and very little is now known about the effect of
wings varying from this form. Raking the ends of the wing tips at a
slight angle increases both the lift-drag and lift by about 20 per cent,
the angle of the raked end being about 15 degrees. Raking is a widely
adopted practice in the United States, especially on large machines.
*Summary of Corrections*. We can now work out the total correction to be
made on the wind tunnel tests for a full size machine of any aspect
ratio. The lift co-efficient should be used as given by the model test
data, but the corrections can be applied to the lift-drag ratio and the
drag. The scale factor is taken at 1.08, the form factor due to rake is
1.2, and the aspect correction is taken from the foregoing table. The
total correction factor will be the product of all of the individual
factors.
*Example*. A certain wing section has a lift-drag ratio of 15.00, as
determined by a wind tunnel test on a model, the aspect of the test
plane being 6. The full size wing is to have an aspect ratio of 8, and
the wing tips are to be raked. What is the corrected lift-drag ratio of
the full size machine at 14°?
*Solution*. The total correction factor will be = 1.08 × 1.10 × 12 =
1.439. The lift-drag ratio of the full size modified wing becomes 15.00
x 1.439 = 21.585.
As a comparison, we will assume the same wing section with rectangular
tips and an aspect ratio of 3. The total correction factor for the new
arrangement is now 1.08 X 0.72 = 0.7776 where 0.72 is the relative
lift-drag due to an aspect of 3. The total lift-drag is now 15.00 X
0.7776 = 11.664.
Having a large aspect ratio and raked tips makes a very considerable
difference as will be seen from the above results, the rake and aspect
of 8 making the difference between 21.585 and 11.664 in the lift-drag.
Area for area, the drag of the first plane will be approximately
one-half of the drag due to an aspect ratio of three.
*Lift in Slip Stream*. The portions of a monoplane or tractor biplane
lying in the propeller slip stream are subjected to a much higher wind
velocity than the outlying parts of the wing. Since the lift is
proportional to the velocity squared, it will be seen that the lift in
the slip stream is far higher than on the surrounding area. Assuming for
example, that a certain propeller has a slip of 30 per cent at a
translational speed of 84 miles per hour, the relative velocity of the
slip stream will be 84/0.70 = 120 miles per hour. Assuming a lift factor
(Ky)=0.0022, the lift in the slip stream will be L = 0.0022 × 120 × 120
= 31.68 pounds per square foot. In the translational wind stream of 84
miles per hour, the lift becomes L = 0.0022 X 84 X 84 = 15.52 pounds per
square foot. In other words, the lift of the portion in the slip stream
is nearly double that of the rest of the wing with a propeller
efficiency of 70 per cent.
CHAPTER VII. BIPLANES AND TRIPLANES.
*Biplane Characteristics*. From an aerodynamic standpoint, the monoplane
wing is more efficient than the superposed wings of the biplane type,
since the proximity of the two surfaces in the latter causes a decided
loss in the total lift. Other practical advantages, however, offset the
losses due to the superposed surfaces, and hence the total efficiency of
the complete biplane may be even greater than that of the monoplane. For
the same area the structural parts of the biplane are lighter, and this
advantage increases rapidly with the size of the machine so that when a
span of 36 feet is exceeded, any other arrangement than that of the
biplane or triplane becomes almost a practical impossibility. A biplane
is easier and cheaper to make than a monoplane, since the wing bracing
of the former can be arranged to better advantage, the load-bearing
members can be simpler, and the safety factor made higher for an equal
weight. By suitable adjustments between the wings of a biplane, it is
possible to obtain a very high degree of inherent longitudinal stability
without incurring much loss in efficiency, an arrangement that is of
course impossible with a single monoplane surface. By "staggering," the
view of the pilot is increased, and the generally smaller size of the
machine permits of better maneuvering qualities for a given load.
*Interference*. Due to "interference," or to the choking of the air
stream between the upper and lower surfaces, the lift of both wings is
reduced, with the drag remaining about the same as with a single
surface. This, of course, reduces the total lift-drag ratio at all
except certain angles. The relative lift-drag ratios of the monoplane
and biplane depend to some extent upon the form of the wing.
Interference causes a loss on the opposing faces of the wings, the lift
being reduced on the top surface of the lower wing, and on the bottom
surface of the top wing. Since the upper surface of the lower wing is
under suction, and therefore produces the greater proportion of lift, it
is natural that the lower wing lift should be reduced to a greater
extent than in the upper wing, since it is only the lower surface of the
latter that is affected. At normal flight angles the upper wing carries
about 55 per cent of the total load. At zero degrees incidence, the
upper wing carries as high as 62 per cent of the total load, while at 12
degrees this may be reduced to 54 per cent.
*Gap-Chord Ratio*. Calling the distance between the upper and lower
wings the "gap," it may be said that the ratio of the gap to the wing
chord greatly influences the lift. This ratio is called the "gap-chord
ratio," and may vary from 0.8 to 1.0 in small machines or 1.0 to 1.2 in
slow, heavy aeroplanes. With the drag remaining practically constant,
the lift-drag is of course affected by a change in the gap-chord ratio,
this quantity being diminished at small gap ratios. Compared with a
monoplane, the lift of a biplane is about 0.77 when the gap is 0.8 of
the chord, and about 0.89 of the monoplane value when the gap-chord
ratio is increased to 1.6. In this range the lift-drag approximates 0.82
and 0.89, respectively. The center of pressure movement is not greatly
changed with any gap-chord ratio, and to all practical purposes remains
the same as with the monoplane. It should be understood that these
remarks apply only to the "Orthogonal" biplane arrangement in which the
wings are vertically over one another.
While biplane efficiency is increased by having a large gap-chord ratio
(wing efficiency alone), the total efficiency of the aeroplane is not
always increased by a large gap, principally because of the great head
resistance due to the longer struts and interplane bracing. At high
speeds the longer bracing members often more than offset the gain due to
wing efficiency, and as a result the gap of high speed scouts will
generally be found in the neighborhood of 0.8 the chord. With slow,
heavy machines, where lift is of great importance, and where slow speed
does not affect the structural resistance to so great an extent, the
gap-chord ratio will range from 1.0 to 1.2.
In making the above comparisons between monoplanes and biplanes, equal
aspect ratios have been assumed for both types, but in actual practice
the aspect ratio of biplanes is always greater than with monoplanes, and
as a result the biplane loss is usually less than indicated above. When
correction has been made for the aspect ratio, the disparity in the
monoplane and biplane values of Ky and L/D is not as great as commonly
supposed. "Biplane reduction factors," or the factors used in reducing
monoplane values to those of the biplane, depend to a great extent upon
the wing section as well as upon the gap, and for exact values of the
factors we should have the tests report of the wings in biplane form.
Lacking this information, we can adopt the values obtained by the N. P.
L. for an old type of wing in order to get approximate results. To
obtain the biplane values, multiply the monoplane values obtained by the
wind tunnel test by the factors found under the required gap-chord
ratio. These factors apply to an aspect ratio of 6.
BIPLANE REDUCTION FACTORS (N. P. L.) (At Normal Flight Angles)
───────────────────────────────────────────────────────────────────
Gap-Chord Ratio. 0.8 1.0 1.2 1.6
───────────────────────────────────────────────────────────────────
Ky Reduction Factor 0.77 0.82 0.86 0.89
───────────────────────────────────────────────────────────────────
L/D Reduction Factor 0.82 0.84 0.85 0.89
───────────────────────────────────────────────────────────────────
Dr. Hunsaker conducted experiments at the Massachusetts Institute of
Technology on biplane and triplane combinations, and the results were
reported in "Aviation and Aeronautical Engineering," Nov. 1, 1916. The
R.A.F.-6 section was used with a gap-chord ratio of 1.2. The biplane
portions of the experiments are as follows, the actual Ky and L/D values
and reduction factors being arranged according to the angle of
incidence:
It will be noted that there is steady improvement in the lift factor
with an increase in the angle from 2° up (except at 8°), and that the
same holds true with the L/D factor. That is, the biplane values become
nearly monoplane values at high angles, and in the case of the L/D ratio
the biplane actually is 24 per cent greater than the monoplane value at
an angle of 16°. The lift coefficient Ky above, is not far from the
corresponding Ky, for gap-chord ratio = 1.2 in the first table. The
maximum biplane value of L/D occurs at the same point as in the
monoplane wing, that is, at 4°. The fact that the lift-drag is so high
at 16° is very favorable, since the biplane would be less likely to
stall when flying slowly, and with a big demand on the engine. The range
of angles at the stalling angle is much greater than with the monoplane
wing, and the lift does not fall off so rapidly after the maximum is
reached.
*Biplane Arrangements*. In the foregoing data we have assumed that the
upper wing was placed directly above the lower, and with the leading
edges on the same vertical line as shown by Fig. 3. This is known as an
"Orthogonal" biplane, and the gap is indicated by G and the chord by C.
In Fig. 4 the forward edge of the top wing is advanced beyond the lower,
or is "Staggered," the amount of the stagger being indicated by S. This
allows of better view, and slightly increases both the lift and L/D
values. With a comparatively large stagger the range of the stalling
angle is increased, and the lift does not fall off as rapidly after the
maximum is reached as with the orthogonal type. In Fig. 5 the top wing
is given a backward stagger, but the exact effects of this arrangement
are not generally known. There are few machines using the reversed
stagger, the only example, to the writer’s knowledge, being the De
Havilland speed scout. By staggering, the resistance of the interplane
bracing struts (3) is somewhat reduced, because of their inclination
with the wind, although they are longer for the same gap than in Fig. 3.
Fig. 6 shows the chord of the lower wing (C’) shorter than the upper
chord, a type used in the Nieuport speed scout. In effect, this is a
form of stagger, and it undoubtedly widens the view of the pilot, and to
some extent increases the efficiency and the range of the stalling
angle. Neither the stagger in (4) nor the small lower chord alone
improves the stability to any extent. To obtain any marked advantage
with the short lower chord, the chord C’ must be very much shorter than
the upper chord, say from 0.80C to 0.50C. The loss of area is so great
that this would not be permissible on any except the fastest machines,
where lift is not a primary consideration. The pilot’s view, however, is
very much improved with the short lower chord, and in battle this is an
important consideration.
Fig. 7 shows the chord of the upper wing inclined at an angle with the
lower chord by the amount (d). This is known as "Decalage" and is
productive of a great degree of longitudinal stability when taken in
combination with stagger. The stability attained by decalage and stagger
is without a great loss in the L/D ratio, while the lift and stalling
angle range are both increased. This latter stable combination is shown
by Fig. 8, in which the wings are given both stagger and decalage.
[Illustration: Slow Speed, Two-Seat Biplane, with a Large Gap-Chord
Ratio. The Large Gap Is Permissible in a Slow Machine, as the Strut
Resistance Is Less Than the Gain in Lift-Drag Ratio Obtained by the
Greater Gap. It Will Be Noted That These Wings Have a Considerable
Amount of Stagger. The Position of the Bottom Wing Allows the Observer
to See Almost Directly Below.]
[Illustration: A High Speed, Two-Seat Fighting Biplane, with a Small
Gap-Chord Ratio. In This Case, the Strut Resistance Would Be Greater
Than the Aerodynamic Gain of the Wings with a Greater Gap Chord Ratio.
The Gunner Is Located in the Rear Seat, and Behind the Trailing Edge of
the Lower Wings. He Has a Clear Field to the Rear an Over the Top Wing.]
*Forward Stagger*. Eiffel performed experiments with Dorand wings, and
found that when the top surface was staggered forward by 1/2.5 of the
chord (0.4C), and with a gap-chord ratio of 0.9, an increase in lift of
from 6 to 10 per cent was obtained. The L/D was the same as with no
stagger. With thin circular plates, 1/13.5 camber, and a gap-chord ratio
= 0.66, the lift-drag was better (than with no stagger) only when the
value of Ky was greater than 0.066 (metric). Then the L/D improved
progressively with the amount of stagger. Ky was improved by 5 per cent
when the stagger was equal to half the chord, and by 10 per cent when
the stagger was equal to the chord. The N. P. L. with a Bleriot wing,
aspect ratio=4, found that Ky was increased by 5 to 6 per cent with a
stagger of 0.4C, and the L/D was increased by about 4 percent. The
gap-chord ratio was 1.00.
[Illustration: A Single Seat Biplane Speed Scout with an Air Cooled
Motor.]
In a series of tests made by A. Tcherschersky, the backward stagger as
in Fig. 5 gave about 15 per cent greater lift than the orthogonal
biplane, or about 4 per cent less lift than a monoplane surface of the
same area. The stagger in this experiment was about 0.33C. In default of
more accurate information, it would seem that backward stagger would
give better results than forward stagger, since the air swept down by
the upper surface would pass further to the rear of the lower plane and
hence would not so greatly affect the vacuum on the upper surface of the
lower wing. This would, however, destroy the view of the pilot to a
greater extent than any of the other arrangements.
Stagger always introduces structural difficulties, makes the wings
difficult to assemble, and the wires are of varying lengths. A simple
orthogonal cell is more compact and better from a manufacturing
standpoint, as it simplifies the fittings, and to a slight extent
decreases the weight. When combined with sweep back, the complication is
particularly in evidence. It is pleasing to note the prevalence of
orthogonal cells on modern battle-planes.
*Influence of Camber*. The amount of air swept down by the upper wing is
largely determined by the curvature of the under surface of the upper
wing. By decreasing, or flattening out the curvature of this surface,
the velocity is increased in a horizontal direction and reduced in a
vertical direction, so that the lower wing is less affected. The upper
surface of the upper wing is not influenced by interference. It should
be noted at this point that air in striking a convex surface is
increased in horizontal speed while the reverse is true of the lower
concave surface. If the under surface of the upper wing were made
convex, the down trend of the air would be still further reduced, and
the loss on the lower wing reduced in proportion.
Increasing the camber on the upper surface of the lower wing increases
its horizontal velocity and hence affects the upper wing to a less
extent, but as the upper wing loss is comparatively slight, the camber
increase below is not of great consequence. This has only been tried in
one machine to the writer’s knowledge, one of the Standard seaplanes, in
which the upper wing was an R.A.F.-6 and the lower wing was a deeply
cambered U.S.A.-2 section. The lower surface of the R.A.F.-6 is
comparatively flat.
*Effects of Decalage*. When the upper wing incidence is increased in
regard to that of the lower wing, or is given decalage, the stability is
increased with a slight increase in the power or drag. This angle shown
by (d) in Figs. 7 and 8, must be accompanied by stagger to obtain
stability, the angle (d) ranging from 1° to 4°. With a decalage of 2.5°,
and a stagger of half the chord, a high degree of stability is attained
with a loss in the lift-drag of from 4 to 6 percent. The lift and the
range of the stalling angle are both increased, the former by about 3
percent, while the latter is nearly double. By increasing the decalage
to 4°, the lift-drag is still 4 percent less than with the orthogonal
cell, but the range of the stalling angle is nearly tripled. The 4°
decalage is very stable and is suitable for training machines or for
amateurs. In either case, the stagger-decalage system is usually better
than sweep back, reflex curves or negative wing tips.
Without regard to the stability, and only with the idea of a greater L/D
in mind, it has been usual in several European machines to adopt a
"negative" decalage; that is, to increase the angle of the lower wing in
regard to the upper chord. With the top chord horizontal, a negative
decalage of 4° would make the incidence of the lower wing equal to 4°.
This has not been generally found advantageous in model tests, but in
full size machines there is a considerable increase in the L/D ratio.
The greater incidence of the lower wing also improves the lift of this
surface and thus requires less surface for obtaining the same total
lift, especially when top wing is staggered forward. Incidence of top
wing of Nieuport = 1°-30’. Lower wing is set at 3°.
*Varying Incidence*. With several types of European speed scouts, and in
the case of the old Handley-Page monoplane, the angle of incidence is
reduced from the center of the wing to the tip. Thus in one speed scout,
the incidence at the body is 4°, and 2° at the tips. A decrease in angle
toward the tips has much the same effect as an increase in aspect ratio;
that is, it decreases the lateral flow and end leakage. It also has an
effect in aiding the lateral stability because there is less lift at the
tips, and hence they are less affected by side gusts. "Washed out"
incidence is an aid to longitudinal stability, as the center of pressure
at the tips is moved further back than at the center of the wing, and
therefore the C. P. is distributed over a longer distance fore and aft
than it would be with a uniform angle of incidence.
In driving the propeller, the motor tends to turn the body in a
direction opposite to that of the propeller rotation, and if no other
provision is made this must be overcome by means of the ailerons. The
"Motor torque" on small span machines is particularly difficult to
overcome in this way, owing to the short lever arm length of the
ailerons. To practically overcome the torque, without excessively
loading the ailerons, it is usually the practice to set the lower left
wing tip at a greater angle than the lower right wing. The greater angle
at the left gives a lift that opposes the turning moment of the motor.
This compensation can never be complete, for the motor torque varies
with the motor output, hence an average angle is selected so that the
incidence will cover the usual horizontal flight speeds.
*Triplane Arrangement*. When a biplane exceeds a certain weight the area
required for a given landing speed makes it desirable to increase the
number of lifting surfaces to more than two, if the span and stress are
to be kept down within reasonable limits. Thus the biplane has its
limits as well as the monoplane, and in the biplane this limit is
generally reached when the span approaches 80 feet. In addition to the
increased weight due to spans of over 80 feet, there are other troubles
in regard to the space required for housing, and awkwardness in
maneuvering. On the smaller and faster aeroplanes, the triplane
arrangement permits of space condensation, and also allows of larger
aspect ratios than with the biplane. The greater depth of the triplane
structure makes the interplane bracing even more effective than in the
case of the biplane. For equal spans there is less bracing exposed to
the wind, and the weight of the wing spars and ribs can be considerably
reduced. The shorter ribs of the triplane alone contribute in no small
degree to the saving in weight.
Considering the wings alone, without reference to the head resistance of
the bracing, etc., there is a greater loss of lift and L/D when three
tiers of wings are superposed than with a biplane. In experiments by Dr.
Hunsaker upon R.A.F.-6 and Curtiss wing sections, it was found that at
about 4°, that the triplane required about 6 percent more power than the
corresponding biplane. At this angle, the L/D for the triplane was 12.8,
against the ratio of 13.8 for the biplane. The gap-chord ratio in each
case was maintained at 1.2. Both the R.A.F.-6 and the Curtiss wings gave
results of the same general character, and there was not a great deal of
difference in the numerical values. At very high angles, 12° to 16°, the
lift of the biplane and triplane only differed by about 2 percent, but
at very small angles such as are used at normal flight speeds, the
reduction of lift in the triplane was very marked.
The drag was not greatly different below 12°, but at 16° the
drag-coefficient is less than that of either the biplane or monoplane,
and for machines flying at low speeds, or heavily loaded, this decrease
is of great advantage since it relieves the motor at a time when power
is particularly required. At this point it should be noted that at high
angles, the L/D generally is better for multiplanes in an almost direct
proportion to the number of surfaces. In this experiment, the lift-drag
ratios for a monoplane, biplane, and triplane were respectively 4.5,
5.6, and 6.5. The drop in lift after the point of maximum lift, or the
stalling angle, is not as rapid as in the case of the biplane or
monoplane, and hence there is less danger of stalling the triplane. With
the same area, and loading, the landing speed of the biplane and
triplane will be about the same.
The following tables give the lift, and lift-drag ratios as determined
in these experiments, the factors being in terms of the monoplane values
of an R.A.F.-6 wing. Thus to obtain triplane values, multiply the given
monoplane values by that number opposite the required angle of
incidence. Aspect ratio = 6.
[Illustration: Curtiss Triplane Speed Scout. Note the Great Aspect Ratio
of the Wings, and the Relatively Great Gap-Chord. Ratio. Only One Set of
Struts Are Used in a Single Row, Hence the Head Resistance Is at a
Minimum. The Span Is 25’-0" and the Chord 2’-0", Giving an Aspect Ratio
of 12.5.]
Thus, if the monoplane lift value for the R.A.F.-6 wing at 4° is Ky =
0.001.45, then the triplane value will be 0.00145 + 0.757 = 0.001097 as
given in the table. The monoplane lift-coefficient of any other wing
section can be handled in the same way with fair accuracy. To obtain the
corrected lift-drag ratio for any wing section, multiply the lift-drag
of the monoplane wing by the factor in the above table corresponding to
the incidence of the monoplane test wing.
[Illustration: The Italian Caproni Triplane of the Heavy Lift or Bombing
Type. Motors Are Installed in Each of the Three Bodies, Tractor
Propellers Being Used in the Two Long Outer Bodies, While a Pusher Screw
Is Used at the Rear of the Central Passenger Body. The Enormous Size of
This Triplane Can Be Seen by Comparing it with the Caproni Monoplane
Shown at the Right. Courtesy "_Flying_."]
The upper wing gives the greatest percentage of lift, and the middle
wing the least, since the latter suffers from interference on both
sides. It has been found that the sum of the top and bottom wings of a
triplane group gives the same lift as the two wings of a biplane under
equal conditions. It was also found that the lift-coefficients and
lift-drag of the upper plane alone was very nearly equal to the lift of
the combined effects of all three wings, and at all angles. Calling the
lift of the middle wing 1.00 (4°), the lift of the upper wing will be
1.91 and the lower wing 1.64. Calling the L/D of the middle wing 1.00
(4°), the relative life-drag will be L/D = 2.59 for the upper wing and
1.69 for the lower. With the middle wing still assumed at unity, the
lift of the top plane is at 1.49 at 16°, and the lower wing 1.20. The
liftdrag at 16 degrees will be respectively 1.00, 1.22, and 1.117 for
middle top and bottom. At 0°, the upper wing will carry 2.68, the middle
1.00, and the bottom 1.82. At 0°, the lift-drag of the top is 3.63, the
middle 1.00, and bottom 2.30. These relative figures are only useful in
comparing the loading when computing the strength of the structural
parts. See "Aviation and Aeronautical Engineering" Nov. 1, 1916.
*Overhanging Wing Tips*. In many American machines, and in some European
machines, such as the Farman, the upper wing is given a much greater
span than the lower. Of late, the tendency has been to make the wings of
equal, span and fully 90 per cent of the modern machines will be found
to be arranged in this way. While the overhanging tips may slightly
increase the efficiency of the biplane by reducing interference at the
ends, it makes the span unduly long and difficult to brace at the end.
The added end bracing due to the overhang probably offsets any
aerodynamic advantage to be obtained, although I have no accurate data
on this point. Compactness is certainly not a feature. It is said that
ailerons are more effective when mounted on the upper overhang, and this
may be so, but I note that the area is about the same in any case. With
overhanging tips, the ailerons are generally placed on upper wings, only
while with equal or nearly equal spans, they are placed top and bottom.
The overhanging section and the ailerons form a single detachable unit
as a general rule. With nearly equal spans, the upper and lower ailerons
are generally interconnected with a small strut in such a way that they
act together.
Small speed scouts, rarely if ever, have any overhang since the object
of these machines is to make them as small and compact as possible.
CHAPTER VIII. EFFECTS OF PLAN FORM. (TANDEM AEROPLANES.)
*General Notes*. Up to the present we have considered only two wing
outlines, the rectangular and the wing with raked tips. In addition, we
have considered only the effect taking place on monoplane surfaces. For
the purpose of obtaining longitudinal stability, or for distributing the
lift upon two or more following surfaces, the plan view in some
aeroplanes has been somewhat modified as in the Dunne, Langley and Ago.
Undoubtedly the simple rectangular wing, or the wing with raked tips,
have proved the most efficient from an aerodynamic standpoint, but as
the layman is usually interested in distorted wing shapes, or
odd-looking outlines, I will describe the effect of such stabilizing
forms upon the lift and drag.
Figs. 1-9 show the usual range of wing forms, at least those that have
been used on well known machines, and all of them have flown with
varying results. In any case, the variations in the values of the lift
and lift-drag are not excessive, the extreme cases varying possibly not
more than 20 per cent on either side of the values for a plain
rectangular wing. While almost perfect longitudinal stability can be
obtained in other ways, by less loss than by changing the plan form,
certain manufacturers still adhere to one or more deviations from the
more usual rectangular form.
Fig. 1 shows a plan view of a machine with a rectangular wing, and No. 2
shows the machine provided with raked tips. Fig. 3 is a wing with an
inclined entering edge, as used on the English Mann biplane. Fig. 4 is
the German Ago with a diamond form surface, the evident purpose of which
is to simplify the wing spar bracing as shown by the dotted lines. By
bringing the wing spars ’together at the tips, the spars themselves form
a triangle to resist the drag stresses. Fig. 5 is the common form of
"sweep back" or "retreated wing" as used in the Standard H-3 training
biplane, and several other modern biplanes. While this arrangement
undoubtedly assists longitudinal stability, it causes certain losses
that will be described later. The inherently stable Dunne is shown by
Fig. 6 in which the sweep back is increased to almost 90 degrees. In
fact the retreat is so great that no tail or stabilizing surfaces are
used at all, the elevator functions being performed by the ailerons. It
should be noted that the swept back tips really act as stabilizers since
they trail back far behind the center of gravity and center of pressure.
The ailerons (a) are not needed for lateral balance, hence ascent and
descent—and also turning in a horizontal plane—are performed by the
ailerons in setting them in different relative positions. Being far to
the rear they are very effective elevators, although their action and
the extreme retreat of the wings, causes a considerable drag. This is
somewhat offset by the absence of tail resistance. An Austrian or German
"Taube" is shown by Fig. 7 with the negative trailing wing tips (a),
that greatly assist longitudinal stability, but which are decidedly
inefficient. This is evidenced by the fact that neither the Germans nor
Austrians build this machine at present—at least for active service. The
tips (a) are bent up at the rear and thus form a "negative" angle of
incidence with the main lifting surface. This was the original invention
of Igo Etrich, an Austrian, and as with everything else, the idea was
promptly grabbed by the Germans at the beginning of the war and claimed
as their own idea. Etrich was one of the pioneers in aviation, a science
that did not prosper in Germany until the impracticability of the
Zeppelin for universal service was an accepted fact.
[Illustration: Figs. 1-9. Plan Views of Different Wing Arrangements and
Wing Outlines.]
Fig. 8 is the wing outline of the Bleriot monoplane, a representative
wing used on the earlier monoplanes. This is really reversed rake, and
hence does not stand for efficiency in lift. A tandem aeroplane is shown
by Fig. 9 in which the leading surface is (m) and the trailing wing is
(n). The tail (t) may, or may not be included. This is a type that has
been neglected in its practical development although it has been
repeatedly proposed. The Langley machine, the Montgomery glider, and the
Richardson are of this type.
In all the figures the wing ribs are indicated by the thin full lines
passing across the width of the wing, with the tail at (t). The arrow
represents the line of flight. It will be noted that with any but
rectangular wing outlines the rib lengths are different, throughout the
wing. This makes this wing a bad manufacturing proposition, and a
difficult and expensive wing to repair. To provide against emergencies
the aviator must keep a complete extra wing in reserve for repair parts,
while the manufacturer is put to the expense of a great number of rib
molds, and must also keep a large number of ribs in stock. There is a
constant difficulty due to the fact that mistakes are often made in
ordering the ribs for repair, and altogether, anything but a rectangular
wing is a decided nuisance.
*Sweep Back or Retreat*. With swept back wings, as shown in Fig. 5, the
center of pressure movement is peculiar. The C. P. moves forward when
reducing the incidence at small angles, and thus tends to reduce head
diving, but at very large angles the C. P. again moves forward, tending
to increase the angles further and thus stall the machine. This reversal
of C. P. movement takes place at about 10° to 12°, and the movement is
sharper and further with each increase in the sweep back. At ordinary
angles of flight, say at from 0° to 6° the forward C. P. movement is
satisfactory, but at low speeds and high angles stability is only
partially secured, and hence for the total performance sweep back is not
to be desired. The wing section used in the above investigation was an
R.A.F.-6 with an aspect ratio of 6.
In regard to the lift coefficient Ky, it was noted that this factor was
decreased with every degree in the angle of sweep back. At the incidence
angle of 4° for the maximum value of L/D, the value of Ky decreased from
0.00143 with a sweep back of 0° (Straight wings), to 0.00120 with a
sweep back of 30°. At the same incidence, but with a sweep back of 10°,
the lift became: Ky = 0.00130. With a retreat of 20°, Ky- 0.00129. The
lift-drag also suffered with an increase in retreat, this being 17.00
with straight wings, at 16.5 at 10°, 16.2 at 20°, and 12.8 at 30°. Up
to, and including a retreat of 20°, the loss in lift-drag is not so bad,
but in the change from 20° to 30° there is a very great loss.
[Illustration: Fig. 10. (Upper) Various Angles Made by Wings During
Experiments. (Below) The Center of Pressure Movement with Varying
Angles.]
The "angle of retreat" herein specified is such that each wing is moved
back through an angle of one-half the total given retreat (r) angle in
Fig. 5. That is, with a retreat of 30°, each wing section makes an angle
of 15° with the entering edge of a pair of straight wings. It is more
usual to specify the included angle between the leading edges as
indicated by (S) in Fig. 5. In the upper portion of Fig. 10 is shown the
various angles of the wings during the experiments, while below is the
C. P. movement according to the different angles of retreat. The above
is based on experiments made by H. E. Rossell and C. L. Brand, assistant
Naval Constructors, U. S. Navy, and published in "Aerial Age."
The center of pressure referred to is that at the forward point of the
middle longitudinal section of each wing, and with a given incidence the
C. P. is thrown to the rear by about 0.2 of the chord by a sweep back of
10 degrees, and 0.4 of the chord for a sweep back of 20°. The center of
gravity of the machine will thus have to be moved to the rear if sweep
back is employed.
In making a turn a machine with sweep back has a natural tendency to
bank up in the correct attitude, and even a retreat of 10° will add
nearly 100 per cent to the banking tendency when compared to a pair of
straight wings. It will be seen that with swept back wings, the leading
edge is radial, consequently meets the air stream at more nearly a right
angle. This gives more lift to the outer wing than would be the case
with straight entering edges, while the inner end losses increase
correspondingly in lift and hence tends further to depress the inner
tip. The greatest value of sweep back is found in its resistance to side
slip. If the machine should be "over banked," and tend to slip down
toward the inner side of the turn, the backward angle of the inner wings
will cause the entering edge to meet the side stream at more nearly
right angles, and thus tend to reduce the bank and the inner slide slip.
In this respect the retreat is an aid to lateral stability. On the other
hand, the sweep back tends to keep the machine rolling in rough weather
for side gusts then meet the inclined wing edges at an effective angle.
This is particularly noticeable in landing.
*Raked Tips*. This subject was discussed under "Standard Wing Section,"
but it may be repeated here, that Eiffel in an experiment with the
Coanda Wing (Eiffel 38), found that the L/D of the raked wing was 20 per
cent higher than with the same wing in rectangular form. This value
would not be safe to assume with all sections.
[Illustration: Fig. 11. Influence of "Wash Down" on the Rear Wings of a
Tandem Pair.]
[Illustration: Fig. 12. Tandem Arrangements Used in Eiffel Experiments.
(1) Chords in Straight Line, (2) Rear Wing at 2.5°, (3) Rear Wing at
5°.]
*Tandem Arrangement*. In tandem wings as shown by Fig. 9, the downward
wash of the front wing (m) will affect the rear wing (n) by causing a
change in the relative direction of the flow. If the front wings are at
an incidence of (i) degrees (Fig. 11), the deviation or washdown of the
air stream to the rear will be expressed by: d = (0.5i + 1). If i =
incidence of rear wing measured from the chord of the front wing, then
the angle of incidence made by the rear chord to the horizontal will be:
I = (i-i’) - (0.5i + 1), where I is the incidence of the rear plane with
the horizontal.
Experiments by Eiffel on tandem planes with circular cambered aerofoils
gave exceedingly good results for certain combinations (Fig. 12). These
arrangements were used, (1) Chords in a straight line (2), Rear aerofoil
tilted down at a negative angle of 2.5°, (3), Rear Plane tilted down at
a negative angle of 5°. In all cases the camber was 1/13.5 of the chord,
and the front and rear wings were spaced two chord widths apart. While
the drag did not change much for any of the arrangements, the lifts
varied widely, and arrangement (2) is by far the more efficient in
lifting capacity. No. 2 is 50 per cent greater than (1), and has twice
the lifting ability of (3). For the same angle of incidence, the front
wing does the same amount of lifting in all cases, the difference being
entirely due to the changes in the rear surface. In (2) the lift of the
rear aerofoil is actually 13 per cent greater than the front plane. The
following tables give the results:
Demo Table TABLE OF LIFTS IN GRAMS FOR THREE ARRANGEMENTS.
──────────────────────────────────────────────────────────────────
ANGLE OF ARRANGEMENT ARRANGEMENT ARRANGEMENT
INCIDENCE (1) (2) (3)
──────────────────────────────────────────────────────────────────
3 665 1094 334
──────────────────────────────────────────────────────────────────
6 987 1568 703
──────────────────────────────────────────────────────────────────
9 1315 2068 965
──────────────────────────────────────────────────────────────────
12 1540 2326 1347
──────────────────────────────────────────────────────────────────
Average at all 1127 1764 837
angles
──────────────────────────────────────────────────────────────────
Percentage 0.64 1.00 0.47
──────────────────────────────────────────────────────────────────
[Illustration: Fig. 13. Drzewiecki Tandem Arrangement for Longitudinal
Stability.]
The lifts in the above table are for the two planes working together,
and the angle of incidence is the angle of the front aerofoil, or rather
the angle of the combination. The wings were 15 x 90 centimeters, aspect
ratio = 6. Fig. 12 shows the construction clearly. This is only true for
circular arched surfaces of the camber given.
[Illustration: Fig. 14. Drzwiecki Tandem Wing Arrangement for
Stability.]
M. Drzewiecki working with Eiffel’s results on the above combinations,
produced an inherently stable tandem monoplane, in which the front and
rear wings were of different cambers and were set at different
incidences. The front wing is Eiffel No. 8 set normally at 8° incidence,
and the rear wing is Eiffel No. 13-bis (Bleriot 11-bis), set normally at
5°. The center of gravity is approximately half-way between the two
wings, and the front is smaller than the trailing surface. Because of
the difference in area, the lift of the front wing varies less rapidly
than the rear when the angle of the machine changes because of disturbed
air. Should the machine "head up," the rear wing increases faster in
lift than the front, and hence restores the machine to a horizontal
position. Should the front surface drop, the incidence is reduced, but
as incidence of the rear wing is less than the front (8°), the rear wing
is reduced to nearly a zero angle of incidence—(With little lift). The
front wing is still inclined at a considerable incidence: (3°) when the
rear is at zero. This drops the rear, and raises the front wing so that
the normal attitude is restored. Lateral stability is obtained by moving
the two halves of the front wing in relation to one another, the
relative movement being similar to that of ailerons.
[Illustration: Typical Wing Assembling Shop]
CHAPTER IX. WING CONSTRUCTION.
*General Wing Frame Layout*. In many ways, the frame of the wing is one
of the most important structural parts of the aeroplane. It not only
maintains the proper aerodynamic form of the aerofoil, but also
transmits the air pressure and lift to the body of the machine, and
therefore carries the entire weight of the aeroplane when in flight. In
spite of the heavy loading on this frame it has been brought to a
remarkable degree of strength and lightness. Not only is "Brute"
strength necessary, but it must also be rigid enough to properly retain
the outlines of the aerofoil with the heaviest loadings, hence the
efficiency of the aeroplane greatly depends upon the stiffness as well
as strength. The contour of the entering edge must be particularly
accurate and well supported since it is at this point that the greater
part of the lift is obtained, and where a slight deviation in form will
materially affect the lift and drag.
The fabric surface, on which the air pressure is exerted, must transmit
the pressure and lift to the main structural members through the parts
that give form to the surface. The fabric surfacing, being flexible and
pliant, must be supported at frequent intervals by the forming members
which in effect are similar to the joists of a floor system. The forming
members are then supported in turn by longitudinal beams, or girders,
that transmit the pressure to the point where the load is applied. The
girders not only carry the lifting force, but must also take care of the
drag which acts at right angles to the lift. To pass girders that are
sufficiently strong, and yet within the limits of weight, through the
narrow space between the top and bottom surfaces of the wing is not
always the simplest of problems.
Figs. 1 and 2 show typical wing frames in diagrammatic form, the upper
views are the plans, while at the bottom are sections taken through the
wing. The outlines of the sections are curved to the outlines of the
aerofoil adopted for the wings, and after this outline is drawn out to
scale, we must maneuver our structural members so that they will lie
entirely between the surfaces.
In Fig. 1, the forming ribs are indicated by R, these being the members
curved to the aerofoil form. They are spaced along the length of the
wing at intervals of about one foot and the fabric is applied to the top
and bottom edges of the rib. The ribs are fastened to the front spar F,
and the rear spar S. The spars are equivalent to beams, and are for the
purpose of transmitting the lift of the ribs to the body. A thin strip E
(nosing) running along the entering edge of the wing, serves to hold the
fabric taut at this point and also forms it to the shape of the aerofoil
entering edge. The thin trailing edge strip (T) performs the same
purpose, and the wing outline is completed by the "End bow" (A) which
retains the fabric at the wing tips. Between the front spar F and the
rear spar S is the trussed "Drag bracing," which binds the two spars
into a truss in a horizontal direction, and against the drag of the
surfaces. This consists of the "Drag" wires or cables (d) and the short
wood struts (e), although in many cases the ribs are strengthened at the
point of attachment of the drag wires and serve as struts. The aileron G
is located at the outer tip and is hinged to the rear spar or to an
extension of the rear spar. Between the spars are thin strips known as
"battens" which stiffen the ribs sideways, these are shown by (F).
[Illustration: Fig. 1. (Left) Wing Assembly with Spar to the Rear of the
Entering Edge. Fig. 2. (Right) Assembly with the Front Spar at the
Entering Edge.]
Metal connection clips C, at the end of the wing spars, are for
attaching the wings to the body, or for connection of the two halves of
the upper wing of a biplane. Looking at the lower sectional view we see
the interplane struts of a biplane attached to the front and rear spars
as at (m) and (n). Referring to the plan view, the location of the
struts is indicated by * * * at the points where the drag-bracing is
attached to the spars.
Fig. 2 is a form of wing in which the spar F’ also forms the entering
edge, thus eliminating one part of the wing. One objection to this
construction is that the front spar must necessarily be shallower than
the spar shown in Fig. 1. The rear spar is in the usual location at S’,
the two spars being connected through the usual end bow A’. The trailing
edge T’ may be either a thin strip, or it may be a thin cable as
indicated. This wing is similar to the wing used on the early Wright
machines, and is still used by Farman, Voisin and other European
manufacturers of biplanes. Usually the trailing ends of the ribs
overhang the rear spar for quite a distance, in this type of wing,
giving a flexible trailing edge. The front and rear interplane struts
(m) and (n) are shown, the former connecting with the front spar at a
point near the entering edge.
[Illustration: Fig. 3. Sub-Rib Construction, the Sub-Ribs (r) Are Placed
at the Entering Edge.]
Fig. 3 shows the usual construction except that short "Sub-ribs" marked
(r) are placed between the main ribs R at the entering edge. These short
ribs increase the support and accuracy of the curve at the entering
edge, or else allow wider spacing of the main ribs R. The fabric must be
well supported at this point, not only to maintain the best efficiency
of the aerofoil, but to relieve the stress on the fabric, as it is here
(Top surface) that the greatest suction pressure comes. Should there be
a rip or tear near the entering edge, in the lower surface, the upper
fabric will be subjected to both the pressure underneath and the vacuum
above. This adds fully 25 per cent to the load on the upper facing.
The main spars may be of wood or steel tubing, although the former
material is generally used. They are of a variety of forms, the "I" beam
section, solid rectangular, hollow box, or a combination of plate and I
sections, the total object being to obtain the greatest strength with
the least possible weight. When made up of several pieces of wood they
are known as "Built up" spars.
[Illustration: Fig. 4. Effects of C.P. Movement on Spar Loading]
The load on the spars varies with the total weight carried, and also
with the movement of the center of pressure due to changes in the angle
of incidence. When the center of pressure moves to any extent, the loads
on the two spars may vary between wide limits, and in extreme cases,
either spar may carry the full load. This is shown clearly by Fig. 4, a
section taken through the wing. The front spar F and the rear spar S are
spaced by the distance L, the respective spar loads being indicated by Y
and Z. As before explained, the center of pressure moves forward at
large angles (CP), while at small angles it moves back say to position
(CP’). Should it move back as far as CP-2, the load will come directly
under the rear spar and this member will therefore carry the entire
load. When at the forward position CP, the greater part of the load will
come on the front spar, and only a small portion will now come on S. In
the same way, when at a small angle of incidence, the center of pressure
will be at CP’, a distance (K) from the rear spar. The greater part of
the load will now be on S. The action is the same as if the entire
weight W or lift, were concentrated at the center of pressure.
When intermediate between the two spars, the center of pressure causes a
bending moment in the rib R, and is at a maximum when the CP is midway
between the two spars. It will be seen that the C. P. movement has an
important effect outside of the question of stability, and this travel
must be taken into careful consideration when the strength of the spars
is calculated. To find the load on the rear spar, for example, with the
center of pressure at CP, multiply the lift W by the distance P, and
divide by the spar spacing L. This will give the load Z. With the C. P.
in the same position, the load on the front spar will be the difference
between the total lift W and the load on the rear spar, or Y = W-Z. With
the load at CP’, the load on the front spar will be: Y = WxK/L, and the
load on S will be Z = W - Y.
For example, we will assume that the lift W = 1000 pounds, and the
distance P = 12 inches. The spar spacing L = 30 inches and the center
pressure is at CP. The load Z on the rear spar, will be: Z = WxP/L =
1000 x 12/30 = 400 pounds. The load on the front spar can be found from
the formula, Y = W – Z = 1000–400 = 600 pounds.
[Illustration: Fig. 5. Perspective View of Wing Construction (Rear Spar
Omitted), Showing Hollowed Entering Edge and Built-Up Spar. Rib Is of
the "I" Beam Type. Courtesy "_Flight_."]
Fig. 5. shows a typical form of wing construction (rear spar omitted).
The front spar is of the "Built up type," and the trailing edge is a
flattened steel tube. The rear spar is simply a solid rectangular beam.
A central ash "I" beam is used as the front spar, with vertical spruce
plates on either side. The spruce entering edge, or "nosing," is formed
to the shape of the entering edge and is hollowed out for lightness. The
rib is also of the built up type, the upper and lower flanges are of
spruce and the middle portion (Web) is cotton-wood. At the point where
the spar passes through the rib, the rib flanges pass over, and are
tacked to the spar. The spruce nosing fits closely over the front web of
the rib. The rib flanges are cut away so that the outside of the nosing
will come flush with the flange line of the rib.
[Illustration: Fig. 6. Caudron Monoplane Wing with Steel Tube Spars and
Flexible Trailing Edge. A Slot in the Rear of the Rib Web Permits the
Deflection of the Trailing Edge. Drag Wire Bracing Is Used Between the
Front and Rear Spars. Courtesy "_Flight_."]
A wing of decidedly different construction is the Caudron monoplane wing
shown by Fig. 6, the front and rear spars of this wing being steel tubes
with an entering edge of thin wood. The drag bracing wires may be seen
connected at alternate ribs by small steel plates and the latter also
serve to attach the ribs to the spars. Instead of being cut out
entirely, the webs of the ribs are hollowed out between the spars.
Probably the most unique feature is the construction of the long
flexible trailing ends of the rib at the right. The trailing rib edge is
divided into an upper and lower section by a long slot, the upper
sections being rigid, while the lower edge is thin and flexible. The
flexible edges allow the lower ends of the ribs to give locally, and
reduce the camber when struck by a heavy gust. This aids in the lateral
stability, since the lift is thus considerably reduced at the point of
impact, and it also relieves the wing of unnecessary stresses. The rigid
upper section of the rib acts as a limit stop to the lower half, and
prevents the flexure from exceeding a certain amount. Owing to the
flexibility of the trailing edges a steel wire or cable must be used for
the trailing edge.
[Illustration: Fig. 7. Standard H-3 Wing Construction. Spar, Rib and
Drag Strut Connections at Left. Body Connection Fitting or Hinge at
Right. Note Drag Wire Fittings. Courtesy "_Aerial Age_."]
Details of the framing of the Standard H-3 are shown by Fig. 7. The
figure at the left gives a clear idea of the connections between the
drag struts and spar, while the view at the right shows the body
connection at the end of the spar. I am indebted to "Aerial Age" for
these sketches. The main spar is in a solid piece, channeled out to "I"
beam form, except at the point where the spruce drag strut is attached.
At the end of this strut is attached a sheet steel fitting that affords
a means of connecting the drag wires, and for fastening the strut to the
main wing spar. At the point of attachment, wooden plates are fastened
to either side of the spar. These prevent the fitting bolts and the
fitting from sliding along the spar when subjected to an uneven pull in
the wires. A veneer top and bottom plate still further strengthen this
joint and hold the sub-rib in place. The main ribs are strengthened, at
the point where the spar passes through the rib web, by small vertical
blocks. In the right hand figure the steel clevis is shown bolted to the
spar. A lug for the wing drag wire is brought out from the fitting. The
clevis on the wing engages with a similar clevis on the body of the
machine, and the two are fastened together with a bolt or pin.
[Illustration: Fig. 8. Typical Biplane Wing. Gap for Aileron Shown at
Right End of Wing. Left End Rests Against Fuselage, the Observation Port
Being Cut Out at the Upper Left Hand Corner. Drag Wire Bracing Clearly
Shown. Courtesy "_Aerial Age_."]
Fig. 8 is a photograph of a biplane wing with the framing members
completed and ready for the application of the fabric. At the right is
the opening left for the aileron, and at the left is the observation
port, the latter coming next to the body. As this is a lower wing, the
sockets for the attachment of the interplane struts can be seen on the
upper and near edges of the main spar. Between the spars are very thin
wood strips running with the length of the wing. These are the "Battens"
used for stiffening the ribs between the points of support at the spars.
As the distance between the spars is comparatively great, in respect to
the thickness of the rib flanges, some sidewise support of this kind is
necessary. The drag-bracing cables cross three rib spaces, or are
fastened to every fourth rib. Between the front spar (at the bottom),
and the entering edge, are the small strips that serve as sub-ribs.
Double cross bracing is used at the inner end of the wing (left), while
additional knee braces are placed at the aileron opening, and at the
outer tips. This is necessary to withstand the stresses due to
assembling and handling, rather than for the flight stresses.
Fig. 9 is a Standard Wing ready for covering. Before the fabric is
applied, a narrow cloth strip is wrapped over the trailing edge, as
shown, and is stitched to the frame. This forms a means of stitching the
main covering at the rear edge, where the ends of the upper and lower
surfaces meet.
*Wing Fabric or Covering*. At the present time unbleached Irish linen is
used almost exclusively for covering the wing structure, although in the
early days of flying rubberized fabrics were used to a great extent.
[Illustration: Fig. 9. Standard H-3 Wing Ready for Covering. Opening for
Aileron Flap Shown at Upper Left Hand Edge (Trailing Edge).]
After the linen is stretched on the wing frame, it is given several
coats of a special preparation commonly known as "Dope" to proof the
fabric against moisture. In addition to waterproofing, the dope adds
considerably to the strength of the fabric and shrinks it tightly on the
ribs—much more evenly than could be done by hand. When completely
"Doped," the linen should be proof against the effects of salt water,
moisture, or extreme dryness, and the fabric must be "Drum tight" at
every point on the surface of the wing.
The linen should have a tensile strength of at least 75 pounds per inch
of width in any direction, and weigh from 3.75 to 4.4 ounces per square
yard. It must be wet spun, free from filling matter and uncalendered. As
a usual thing, the width should not be less than 36 inches, although the
width can be altered to meet conditions of rib spacing, etc. The U.S.A.
seaplane specifications (1916) require a minimum strength along the warp
of 75 pounds per inch width, and 85 pounds per inch of width along the
weft. The following table gives the properties of well known wing
fabrics:
*Wing Dope*. Wing dopes are in nearly every case based on
cellulose—either cellulose acetate or nitrate being the most common
base. This has proven far superior to the resin, copal, gum or oil bases
contained in ordinary varnish, since the cellulose of the dope seems to
amalgamate with the cellulose of the flax fiber and bond the whole into
an integral structure. The fact that the dope must be elastic bars the
use of shellac or other hard resin solutions. The solvents used for the
cellulose dopes vary with the makers, some using amyl-acetate,
tetrachlorethane, etc., while others use special secret compounds that
are best adapted for their bases. Many of the solvents give off
poisonous gases in drying, and this must be guarded against by good
ventilation. The vapor of tetrachlorethane is particularly dangerous,
and has resulted in many deaths.
[Illustration: Fig. 10. Complete Framing Plan of Typical Monoplane
Structure. (A) Pilot. (B) Passenger. (M) Motor. (S) Stabilizer. (E)
Elevator. (R) Vertical Rudder.]
Doped surfaces have from 10 to 25 per cent greater tensile strength and
resistance to tearing than the undoped linen, and increases the weight
of the fabric by about 0.7 ounce per square yard for each coat applied.
Under ordinary weather conditions, dope will require from 20 to 40
minutes per coat for drying, and at least one-half hour should be
allowed between each coat. Weather conditions have a great effect on the
action of dope, and with cellulose compounds the best results are
obtained in a clean dry room, well warmed, and without drafts. On rainy
days the linen is damp and the dope does not set well, and this trouble
is not greatly helped by artificial heat. Drafts cause white spots and
streaks, especially if cold air is allowed to enter directly upon the
warm wing surface. To prevent drafts the ventilating ducts should be
near the floor, and as the vapor is heavier than the air, and flows
downwards, this means of ventilation is entirely practicable.
*Applying the Dope*. The number of coats depend upon the character of
the job, but at least five coats should be applied, and preferably
seven. On the best grade of work, the dope is generally covered with
three or more final coats of spar varnish, although this is not
absolutely necessary. For ordinary work, dope alone on Irish linen has
proved very satisfactory for land machines, five coats being the usual
amount applied on exhibition aeroplanes and planes for amateur use.
Seven coats of dope with three coats of spar varnish are specified for
military machines that are to be used on salt water. Seaplanes are
subjected to conditions that are particularly hard on fabric and must be
protected accordingly.
In applying the dope, at least one-half hour should be allowed for
drying between each coat, and more if possible. The first two coats
should be painted on lightly, the purpose being simply to fill the pores
of the fabric and to prevent the succeeding coats from sinking through.
If the first two coats are too heavy, the dope filters through the mesh
of the linen and drops on the lower surface, causing spots and a waste
of a very expensive material. Dope is expensive even with the greatest
care exercised in its application, and the writer has seen cases where
the first two coats were so heavily applied that fully 50 per cent of
the fluid ran through and caked in among the structural parts of the
machine. This ran the doping expense up to a terrific figure. The cloth
should be dry, and the work performed, if possible, on a dry day. To
save dope, never take out of the supply drum more than can be used for
one coat, for the dope soon becomes tacky on exposure to the air, and a
satisfactory job is hard to obtain if it gets in this condition.
[Illustration: Fig. 10a. Testing the Wing Structure of a Curtiss Biplane
by Means of Sand Bag Loading. The Wings Are Turned Upside Down and a
Sand Load Is Laid Uniformly Over the Wings. So That They Produce a Load
Equal to, or Greater Than, the Flight Load.]
In placing the fabric on the wings, particular care must be taken in
stretching so as not to have it too tight when cellulose dope is used.
The dope shrinks the linen to a very considerable extent, and if it is
too tight to begin with, the stress due to shrinkage will place an
excessive stress on both the fabric and the structure. When of the
proper tautness, the fabric should sound like a drum when snapped with
the finger. Any less tension than this will permit the fabric to sag
badly when under the air pressure and reduce the efficiency of the
wings. In fastening the cloth it should be just stretched taut and no
more. In damp weather the cloth can be stretched a little tighter than
in dry weather.
*Transparent Coverings*. In some types of battle-planes and scouts, a
part of the wing section directly above the pilot is covered with a
transparent fireproof cellulose sheet, much resembling celluloid. This
permits the pilot to see above him through the overhanging wing, and is
of great value in action. In some cases, a strip is placed on the lower
wings along the sides of the body so that the ground is also easily
visible. These cellulose sheets will not crack nor splinter, and are
nearly as flexible as rubber. Celluloid of the ordinary variety must not
be used, for this is easily ignited and is likely to start a disastrous
fire.
*Placing the Fabric on the Wings*. In some aeroplanes the seams of the
fabric are run parallel to the ribs, and are tacked or sewed directly to
them, while in other cases the seams are run diagonally across the plane
or on the "bias." Diagonal seams are most satisfactory, and if care is
taken there is no more waste of linen than with the straight seam. The
seams should be of the double-lapped or "English welt" type, and this of
course necessitates sewing before the fabric is placed on the wings. The
seams used on overcoats are satisfactory for this purpose, and give a
covering that will not stretch nor bag. Some use linen thread and others
use silk, but the linen is preferable, since dope often causes silk to
rot. The seams should be covered with linen to protect them from the
weather and to prevent the entrance of water to the interior woodwork.
Ordinarily, the wing is turned upside down for covering, with the
concave side uppermost. The seams are sewed together so that the
completed fabric is wider than the length of the wing and is a little
longer than is necessary to wrap entirely around the width of the wing.
The fabric is then temporarily fastened along the trailing edge, is
passed under the wing to the front edge, and over the concave upper side
back to the trailing edge. At this point the excess material will hang
down over the rear edge. With the wing in its upside down position, the
convex side will be at the bottom, and if a weight is hung on the
overhanging material at the rear edge, the cloth will be pulled tight
against the lower convex side and straight across and above the concave
side. The fabric at the top is then stretched along the cordal line of
the ribs. By laying a narrow board on top of the fabric, and near the
entering edge, the fabric can be brought down uniformly along the
concave edge of the ribs, and by tacking or sewing as the board is moved
back the concave face can be covered without further trouble. After the
concave face is disposed of, the wing can be turned over and the fabric
is then fastened to the convex side of the ribs.
[Illustration: Fig. 11. Method of Stretching Fabric on Wings. Fabric
Passes Under and Then Over Concave Side and Is Pressed Down into Hollow
by a Board as Shown.]
One method of fastening the linen is to lay tape over the ribs, and then
drive tacks through the tape and fabric into the rib. The tape keeps the
tacks from tearing through the linen. The tape should be heavy linen of
from 3/4 to 1 1/4 inches wide, and laid in cellulose before tacking, so
that the tape will be cemented to the fabric and the solution will be
driven into the tack holes. After the tape is in place, it should be
covered with not less than three coats of cellulose dope before the main
surface is treated. This gives an additional three coats over the tape
where it is most needed for protection against moisture. In any case,
the seam or tacking strip should be pressed down so that it projects as
little as possible above the general surface of the wing.
Tacking is not desirable, for tacks and nails always tend to split the
thin members of the rib, and very often corrode the cloth and weaken the
fabric. This has resulted in the whole fabric being ripped off while in
flight. Iron or steel tacks should never be used, as they destroy the
strength of the fabric very rapidly by the formation of rust, and
particularly with sea-planes used on salt water. Sewing the fabric to
the ribs with linen thread is the most satisfactory method and is in
general use among high-grade builders.
The fabric can be stitched to bands of thread or tape, the latter being
wrapped about the rib. Stitches can also be taken from one surface
straight through to the other. The thread or tape bands on the ribs are
merely wrappings taken around the rib flange, and through the lightening
holes, these bands usually being about 4 inches apart. They at once tie
the flange to the web and form a soft surface into which to take the
stitches. Fig. 4 in Chapter X shows the thread bands (d) wrapped around
the rib flange (G), and through the lightening hole, the fabric lying
above and below the rib flanges as shown. A section view through the rib
and fabric is shown at the right.
*Varnishing*. When varnish is to be used over the dope, only the best
grade of spar varnish should be used, since any other kind is soon
destroyed by moisture. From two to three varnish coats will be
sufficient, and each coat should be thoroughly dried and sandpapered
before the next is applied, care being taken not to injure the fabric
with the sandpaper. Sandpapering between the dope layers is not
necessary, since each successive coat partially dissolves the preceding
coat, and thus welds the layers together. Varnish, however, does not act
in this way, and the coats must be roughened. Shellac rots the linen and
should not be used.
The Government specifies one coat of flexible white enamel in which a
small quantity of lead chromate is mixed. This is applied over the last
coat of varnish. The lead chromate filters out the actinic rays of the
sun, thus reducing the injurious effect of the sunlight on the covering.
If coloring matter is added to the dope, it should be in liquid form, as
powders destroy the strength and texture of the dope deposit.
[Illustration: Fig. 12. Wing Structure of Handley-Page Giant Biplane.
Courtesy "_Aerial Age_."]
*Patching Fabric*. The majority of dopes can be used as cement for
patching, but as dope will not stick to varnish, all of the varnish
around the patch must be thoroughly removed with some good varnish
remover. The varnish must be thoroughly cleaned off or there will be no
results. Before applying the dope, the patch must be well stitched all
around the edges, then cemented with the dope. The patch must now be
covered with at least five coats of thin dope as in finishing the
surface. Particular attention must be paid to filling the dope in the
stitching.
CHAPTER X. WING CONSTRUCTION DETAILS.
*Types of Ribs*. The rib first used by the Wright Brothers consisted of
two spruce strips separated by a series of small pine blocks.
Practically the same construction was used by Etrich in Austria. With
the coming of the monoplane, and its deep heavy spars, the old Wright
rib was no longer suitable for the blocks were not efficient in thick
wing sections. The changes in the wing form then led to the almost
universal adoption of the "I" type rib in which an upper and lower
flange strip are separated by a thin vertical web of wood. At present
the "I" rib is used on nearly every well known machine. It is very
strong and light, and is capable of taking up the end thrust of the drag
wires, as well as taking care of the bending stresses due to the
vertical loading or lift.
Fig. 1 shows the original Wright rib with the "Battens" or flanges (g)
and the spacing blocks B. The front spar is at the leading edge (F), and
the rear spar at S. An "I" beam, or "Monoplane" type is shown by Fig. 2,
and as will be seen is more suitable for deep spars such as (F’) and S’.
The upper and lower flanges (g) are separated by the thin perforated web
(w), the sectional view at the right showing the connection between the
flanges and the web. Lightening holes (h) reduce the weight of the web,
with enough material left along the center of the web to resist the
horizontal forces. The web is glued into a slot cut in the flanges, and
the flanges are then either tacked to the web with fine nails, or bound
to it by turns of thread around the flange.
On the average machine, the web is about 3/16 thick, while the flanges
are from 3/4 to 1 inch wide, and from 3/16 to 1/4 inch thick. On the
very large machines, the dimensions of course are materially increased.
At the strut locations in biplanes, and the point of cable bracing
attachment in monoplanes, the ribs are increased in strength unless the
end thrust of the stay wires is taken up by a separate strut. At the
point of stay connection in the old Nieuport monoplane the rib was
provided with a double web thus making a hollow box form of section
great enough to account for the diagonal pull of the stays.
[Illustration: Fig. 1. Wright Type Rib with Battens and Block
Separators. Fig. 2. Monoplane or "I" Type of Rib with Solid Web.]
Fig. 3 shows the Nieuport monoplane ribs, which are good examples of box
ribs. The sections at the left are taken through the center of the ribs.
The wing chord tapers from the body to the wing tip, while the thickness
of the wing section is greatest at the middle, and tapers down both
toward the tips and toward the body. The upper section is located at the
body, the second is located midway between the body and the tip, and the
other two are near the tips, the bottom being the last rib at the outer
end. The ribs shown are of box form as they are at points of connection,
but the intermediate ribs are of the "I" type shown by Fig. 2.
[Illustration: Fig. 3. Nieuport Monoplane Ribs. This Wing Is Thickest in
the Center and Washes Qut Toward Either End, Thus Making All of the Ribs
Different in Curvature and Thickness. At the Point of Stay Wire
Attachment Double Webbed Box Ribs Are Used.]
*Rib Material*. In American aeroplanes, the flanges of the ribs are
generally made of spruce. The webs are of poplar, whitewood, cottonwood
or similar light material. There is not a great deal of stress on a rib,
and the strongest material is not necessary, but as there are a great
many ribs in a wing assembly lightness is a primary consideration. A few
ounces difference on each rib makes a great deal of difference in the
total weight, especially when there are 80 or more ribs in a complete
machine. Exception to the above materials will be found in the Curtiss
"Super-American" Flying Cruiser which has ribs with pine webs and birch
flanges. European aeroplane practice makes use of hardwood in the ribs.
*Web Stiffeners*. The webs being thin and deep, and cut for lightening
as well, need bracing at the points where concentrated loads are placed,
such as at the front and rear spars, and at points between the
lightening holes. By gluing thin strips to the webs (in a vertical
direction), and so that the tops and bottoms of the strips come tight
against the upper and lower flanges, a great deal of the strain on the
web can be avoided. The stiffening blocks are shown by (x) in Fig. 4,
and are placed on both sides of the front and rear spars F and S, and
also between the lightening holes H.
[Illustration: Fig. 4. Details of "I" Type Rib, Showing Lightening Holes
and Stiffeners.]
*Flange Fastenings*. In the section at the right of Fig. 4 it will be
seen that the web is inserted into a groove cut in the flanges and is
then glued into place. It would be unsafe to trust entirely to the glue,
owing to the effects of aging, moisture and heat, and consequently some
additional means of fastening must be had. It has been customary to nail
through the flange into the web, but as the web is only about 3/16 inch
thick it is likely to split.
An approved method is shown in Fig. 4 in which Irish linen thread wraps
(d) are passed through the lightening holes (H), and over the flanges
(G). The thread is coated with glue before wrapping and after, and when
dry it is thoroughly varnished for protection against moisture. The
bands are spaced from 3 to 4 inches apart. If nails are used they should
be brass nails—never steel or iron.
At the points (F) and (S) where the spars pass through the web, the web
is entirely cut out so that the flanges ordinarily lie directly on the
spars. In this case it is necessary to bevel the spar so that it at
least approximately fits the curve of the flange. Sometimes when a full
size spar is impossible, as in cases where the spar tapers toward the
tips, wood packing pieces may be placed between the flange and the spar;
tapered to make up for the curve. The flanges in any case must be
securely fastened to the spar by brass wood screws as at (e), and the
edges of the web should fit tightly against the sides of the spars.
*Wing Battens*. The wing battens run along the length of the wings, from
end to end, and between the spars, and serve to brace the ribs sideways
as shown in some of the general views of the wing assembly. To
accommodate the battens, the openings (f) are cut directly under the
flange. Usually the battens are thin spruce strips from 3/16 to 1/4 inch
thick and 1/2 inch wide, and should be run through the web at a point
near the stiffeners. The thickness from the top of the flange to the
under side of the lightening hole is from 1/2 to 3/4 inch as indicated
by (C).
*Strength of Wood Ribs*. The strength of a rib for any individual case
can be found by the method used in computing beams, the rib usually
being assumed to have a uniformly distributed load, although this is not
actually the case, as before explained. The greater part of the load in
normal flight is near the front spar, but this shifts back and forth
with the angle of incidence so that there is no real stationary point of
application, and the rib must be figured for the maximum condition. The
total load carried by one of the intermediate ribs is due to the area
between the ribs, or to the unit loading multiplied by the rib spacing
and chord. The portion of the rib between the spars can be calculated as
a uniformly loaded beam, supported at both ends. The entering edge in
front of the spar, and the trailing edge to the rear may be taken as
uniformly loaded beams supported at one end. The proportion of the loads
coming on the ends and center position can be taken from the pressure
distribution diagrams as shown under "Aerofoils."
A number of tests were made on ribs by Mr. Heinrich of the Heinrich
Aeroplane Company, the ribs being built up on short pieces of spar so
that actual conditions were approached. Instead of using a distributed
load, such as usually comes on the rib, a concentrated load was placed
at the center. If the rib were uniform in section the equivalent
uniformly distributed load could be taken as one-half the concentrated
load, but because of the lightening holes this would not be very exact.
It would be on the safe side, however, as such a test would be more
severe than with a uniform load. The ribs were of the same type as shown
in Fig. 5, and were placed 32.5 inches apart. The front spar was 2 7/16
inches deep, the rear spar 2 1/16 inches deep, and the overall depth of
the rib at the center was 3 1/8 inches. The rib flanges were of white
wood 3/4 inch wide, and 3/16 inch thick. In rib No. 1 the web was solid
whitewood, 3/16 inch thick, and in ribs Nos. 2 and 3, the webs were
mahogany three-ply veneer (5/32 inch thick).
*Test of Rib No. 1*. There are 5 lightening holes between spars, with 2
inches of material left between the holes, and 1/2 inch between first
hole and front spar opening. With 95 pounds concentrated load at the
center, the first rupture appeared as a split between the first hole and
spar opening. At 119 pounds, the flanges had pulled away from one side
of the spar, and 1/8 inch away from the web. Full failure at 127.5
pounds. Web was split between each lightening hole with a complete cross
break at center of web, the latter being caused by a brad hole in the
web.
*Test of Rib No. 2*. Laminated web, with no brads driven opposite
lightening holes. At 140 pounds rib deflected 5/16 inch, and when
relieved sprang back only 3/16 inch. With 175 pounds, the deflection
again was 5/16 inch, but the rib continued to bend slowly, the flanges
pulling away from the web and spars. The wood was not broken anywhere,
the failure being in the brads and glue.
*Test of Rib No. 3*. Same materials as No. 2, but web was fitted inside
of the "I" beam spar and the rib flanges were screwed to the spar. At
175 pounds there was no sign of rupture anywhere, and the deflection was
5/16 inch. At 185 pounds the rib broke very suddenly and cleanly, and in
such a way as to indicate that this was the true strength of the rib.
The normal loading on the rib in flight was 17.5 pounds, uniformly
distributed, so that with a concentrated load of 370 pounds equivalent,
the safety factor was 21.1.
The conclusions to be arrived at from this test are as follows:
(1) When a solid soft wood web is used, there should be at least 2 1/2
to 3 inches between lightening holes.
(2) A laminated or three-ply web is the best.
(3) No brads should be driven opposite lightening holes.
(4) The web should fit closely to the spar sides and the flange of the
rib should be tightly screwed to the top and bottom of spar.
The above gives an idea as to the strength of the usual form of wood
rib, and can be used comparatively for other cases if the reader is not
familiar with strength calculations.
[Illustration: Fig. 7. Rib Bending Press for Curving the Rib Flanges.]
*Making the Rib*. Wooden webs are cut out on the band saw, and the webs
are so simple that there is not much more to be said on the subject. The
flanges, however, must be steamed and bent to the nearly correct form
before assembly. After planing to size and cutting the groove for the
reception of the web, the ribs are placed in the steamer and thoroughly
steamed for at least an hour. A rib flange press shown by Fig. 7
consists of two heavy blocks with the inner faces cut approximately to
the rib outline. The steamed ribs are then placed between the blocks,
the bolts are screwed down tight, and is left for 24 hours so that the
strips have ample time to cool and dry. For the amateur or small
builder, the steamer can be made of a galvanized "down-spout" connected
with an opening cut in the top of an ordinary wash boiler. One end of
the spout is permanently sealed, while the other is provided with a
removable cover so that the strips can be inserted. A hole cut near the
center of the spout is connected to the opening in the boiler cover by a
short length of spouting or pipe. The spout should be made large enough
in diameter to contain all of the ribs that can be pressed at one time,
and should be long enough to accommodate longer pieces such as the
fuselage longerons, etc.
When removed from the press, the rib flanges can be glued to the webs
taking care that the glue is hot, and that it thoroughly covers the
groove surface. The rib must now be held accurately in place in a second
form, so that the true rib outline will be retained until the glue drys.
A great deal depends upon the accuracy of the second form, and the
accuracy with which the web outline is cut. The larger manufacturers use
metal rib forms or "jigs," but the small builder must be content with a
wooden form consisting of a board fitted with suitable retaining cleats,
or lugs. The outline of the aerofoil is drawn on the board, the tips of
the cleats are brought to the line, and are screwed to the board so that
they can be turned back and forth for the admission and release of the
ribs. The strip bending press in Fig. 7 is only intended to bend the
flanges approximately to form, and hence two layers may be put in the
press at one time without much error.
*Wing Spars*. In American aeroplanes these members are usually of the
solid "I" form for medium size exhibition and training machines, but for
small fast aeroplanes, where every ounce must be saved, they are
generally of the built up type, that is, made up of two or more members.
In Europe, built up construction is more common than in this country,
and is far preferable for any machine that justifies the additional time
and expense. The wing spars are the heaviest and most important members
in the wing and no trouble should be spared to have them as light as the
strength and expense will permit. They are subjected to a rather severe
and complex series of stresses; bending due to the load carried between
supports, compression due to the pull of the stay wires, bending due to
the twist of non-central wire fittings, stresses due to drag and those
caused by sudden deviations in the flight path and by the torque of the
motors. These should be accurately worked out by means of stress
diagrams if the best weight efficiency is to be obtained.
[Illustration: Fig. 9. Types of Wing Spars. (A) Is the "I" Beam Type.
(B) Box Spar. (C) Is Composite Wood and Steel, Wrapped with Tape.]
A number of different wing spar sections are shown by Figs. 9, 10, 11.
Spar (A) in Fig. 9 is the solid one piece "I" type (generally spruce),
channeled out along the sides to remove the inefficient material at the
center. The load in this case is assumed to be in a vertical direction.
In resisting bending stresses, it should be noted that the central
portion of the material is not nearly as effective as that at the top
and bottom, and that the same weight of material located top and bottom
will produce many times the results obtained with material located along
the center line. At points of connection, or where bolts pass through
the spar, the channeling is discontinued to compensate for the material
cut away by the bolt and fittings.
Spar (B) is of the hollow type, made in two halves and glued together
with hardwood dowel strips. The doweling strips may be at the top and
bottom as shown, or on the horizontal center line as shown by Spar (J).
The material of the box portion is generally of spruce. This is a very
efficient section as the material lies near the outer edge in every
direction, and offers a high resistance to bending, both horizontally
and vertically. Unfortunately a great deal depends upon the glued
joints, and these require careful protection against moisture. There is
absolutely no means of nailing or keying against a slipping tendency or
horizontal shear. The best arrangement to insure against slipping of the
two halves is to tape around the outside as shown by spar (E). This is
strong linen tape and is glued carefully to the spar, and the whole
construction is proofed against moisture by several coats of spar
varnish and shellac. In addition to the strengthening effect of the
tape, it also prevents the wood from splintering in accidents.
[Illustration: Fig. 10. Four Types of Wing Spars, the Spar D Being a
Simple Steel Tube as Used in the Caudron and Breguet Machines.]
Spar (C) consists of a central ash "I" section, with steel strips in the
grooves. Two spruce side strips are placed at either side as stiffeners
against lateral flexure, and the entire construction is taped and glued.
This is very effective against downward stresses, and for its strength
is very compact. Since spruce is much stiffer than either the thin steel
strip, or the ash, it is placed on the outside. Spar (D) in Fig. 10 has
been described before.
Fig. (F) consists of two spruce channels placed back to back, with a
vertical steel strip between them. Again the spruce is used as a side
stiffener, and in this case probably also takes a considerable portion
of the compression load. Spar (G) is a special form of box spar used
when the spar is at the entering edge of the wing, the curved nose being
curved to the shape of the aerofoil nose. In Fig. 11 (H), a center ash
"I" is stiffened by two spruce side plates, the ash member taking the
bending moment, and the spruce the compression. Spar (I) has a compound
central "I," the upper and lower flanges being of ash and the center web
of three ply veneer. The two outer plates are of spruce. This should be
a very efficient section, but one that would be difficult and costly to
build. Fig. (J) is the same as (B), except that the parting lies in a
horizontal plane. Spar (K) has ash top and bottom members, and spruce or
veneer side plate. The resistance of this shape to side thrust or twist
would be very slight. The sides are both screwed and glued to the top
and bottom members.
[Illustration: Fig. 11. Built-Up Wooden Wing Spars, Commonly Used with
European Aeroplanes.]
The front end of a Hansa-Brandenburg wing is shown by Fig. 12, the box
spar and its installation being drawn to scale and with dimensions in
millimeters. The top and bottom are sloped in agreement with the rib
flange curve, and the rib web is strengthened by stiffeners at either
side of the spar. The hardwood dowel strips are at top and bottom as in
Fig. B, and when placed in this position the glued joint is not
subjected to the horizontal shear forces. The walls are thicker at top
and bottom than at the sides, in order to resist the greater vertical
forces. For the same reason is deeper than it is wide. As will be
remembered, the drag is very much less than the lift, and again, the
drag stress is greatly reduced by the internal drag wire bracing.
*Leading Edge Construction*. In the early Bleriot monoplanes the leading
edge was of sheet aluminum, bent into "U" form over the nose of the rib.
In modern biplanes, this edge is generally of "U" form hollow spruce,
about 3/16 inch thick. Another favorite material is flattened steel
tubing, about 1/2" x 1/4", and of very light gauge, the long side being
horizontal. The tube has the advantage of being much thinner and much
stiffer than the other forms, and the thin edge makes it very suitable
for certain types of aerofoils. The wing tip bows are generally of
hollowed ash and are fastened to the spar ends, leading edges, and
trailing edges with maple dowels, the joint being of a long scarfed
form. When the scarfed joint is doweled together, it is wrapped with one
or two layers of glued linen tape. In some types of machines the top
surface of leading edge is covered with thin two ply wood from the
extreme front edge to the front spar. This maintains the aerofoil curve
exactly at the most critical point of lifting, and also stiffens the
wing against the drag forces.
[Illustration: Fig. 12b. An Old Type of Curtiss Biplane Strut
Socket, at Left. At the Right Is a More Modern Type in Which the
Bolts Do Not Pierce the Spar.]
*Trailing Edges*. These are either of thin beveled ash or Steel tube. On
army machines, the rear part of the trailing edge fabric is pierced with
holes about 3/16 inch diameter, the holes being provided with rust proof
eyelets. This relieves an excess of pressure due to rips or tears; one
opening being located between each rib and next to the body.
[Illustration: Fig. 12c. A Standard H-3 Interplane Strut Socket
Is Shown at the Left, the Bolts in This Case Passing on Either
Side of the Spar. Note the Stay Wire Attachment Clips and Pinned
Strut Connection. A German Strut Socket at Right. Courtesy
"_Aerial Age_."]
*Protection of Wing Wood Work*. In protecting the wood framework of the
wings from the effects of moisture, at least three coats of good spar
varnish should be carefully applied, with an extra coat over the glued
surfaces and taping. Shellac is not suitable for this purpose. It cracks
with the deflection of the wings and finally admits water. The steel
parts of the wing should be given two coats of fine lead paint, and then
two coats of spar varnish over the paint. Wires are treated with some
flexible compound, as the vibration of the thin wires, or cables, soon
cracks off any ordinary varnish. The use of shellac cannot be too
strongly condemned; it is not only an indifferent protection, but it
causes the fabric to rot when in contact with the doped surface.
*Monoplane Wing Spars*. A few representative monoplane wing spars are
shown by Fig. 13, the R. E. P., Bleriot XI, and the Nieuport. Except at
points where the stay wires are connected, the Bleriot spar is channeled
out into "I" beam form as indicated in the figure. It will be noted that
the top and bottom faces of the spar are slanted to agree with the
curvature of the ribs. A. steel connection plate is bolted to the sides
of the spar by through bolts, and with a lug left top and bottom for the
top and bottom guy wires. The R. E. P. is also an "I" type, the section.
"A" being taken through the channeled portions, while "B" is taken
through one of the connection points where the beam is a solid
rectangle. The channeling should always stop at connection points;
first, so that the plates have a good bearing surface, and second, to
allow for the material removed by the bolt holes.
[Illustration: Fig. 13. Typical Monoplane Wing Spar
Construction.]
Probably the most interesting of all the spars is the Nieuport, which is
a combined truss and girder type. This spar tapers down from the center
to both ends, being thickest at the points where the guys are attached.
The top flange (J), and the bottom flange (L), are ash, while the side
plates and diagonals (H) are spruce. The diagonals resist the shear, and
are held in place by the tie bolts (I). At the left, the spruce cover
plate is removed, while at the right it is in place with the interior
construction shown in dotted lines. The dimensions are in millimeters.
[Illustration: Fig. 14. Plate Connection for Monoplane Stay Wire
Connection to Spar. A Compression Member or Drag Strut Is Shown
in the Center of the Spar Which Takes Up the Thrust Due to the
Angularity of the Stays and Also the Drag Stress.]
*Location of Spars*. There are a number of items that affect the
location of the spars in regard to the leading edge. The most important
factors in the choice of location are: (1) Shape and depth of wing
section, (2) Center of pressure movement, (3) Drag bracing requirements,
(4) Width of ailerons, (5) Method of attaching the interplane struts.
CHAPTER XI FUSELAGE (BODY) CONSTRUCTION.
*Purpose of Fuselage*. The fuselage of a monoplane or tractor biplane is
the backbone of the machine. It forms a means of connecting the tail
surfaces to the main wing surfaces, carries the motor, fuel and pilot,
and transmits the weight of these items to the wings and chassis. With
the exception of the wing structure, the fuselage is the most important
single item in the construction of the aeroplane. Fig. 1 shows a typical
arrangement of a two-place biplane fuselage equipped with a water-cooled
motor. The motor E, propeller Y, and radiator R are placed in the front
of the fuselage, and considerably in advance of the wings W and D.
Immediately behind the motor is the passenger’s seat A and the fuel tank
F. The pilot’s seat B is placed behind the trailing edge of the wings
and is behind the passenger’s seat. The cockpit openings G and H are cut
in the fuselage top for the passenger and pilot respectively. The rear
extension of the fuselage carries the control surfaces, L being the
vertical fin, M the rudder, O the fixed tail or stabilizer, and P the
elevator.
*Resistance*. To reduce the resistance in flight, the fuselage is of as
perfect streamline form as possible, the fuselage being deepest at a
point about one-third from the front. From this point it tapers out
gradually to the rear. With the motors now in use it is only possible to
approximate the ideal streamline form owing to the front area of the
radiator and to the size of the motor. Again, the projection of the
pilot’s head above the fuselage adds considerably to the resistance. The
wind shields I disturb the flow of air. The connections to the tail
surfaces and to the chassis members K also add to the total resistance.
An arched "turtle deck" J is generally provided, of such a shape that
the pilot’s head is effectively "streamlined," the taper of this deck
allowing the disturbed air to close in gradually at the rear. The flat
area presented by the radiator R is probably the greatest single source
of resistance, and for this reason the radiator is sometimes placed at
the side of the fuselage, or in some other position that will allow of a
better front end outline. An example of this construction is shown by
Fig. 2 in which the radiator R is placed behind and above the motor E.
The front fuselage end Z can now be made of a more suitable streamline
form.
[Illustration: Figs. 1-2-3. Fuselage and Motor Arrangement of
Tractor Biplanes.]
Fig. 3 is a view of the front end of a typical fuselage in which an
air-cooled rotary type of motor is installed. Since the diameter of this
type of motor is seldom much less than three feet, it is necessary to
have a very great diameter in the extreme front. The motor housing or
"cowl" marked E has a diameter "d" which should smoothly blend into the
outline of the fuselage at "b." In the older types of construction there
was often a very considerable break in the outline at this point,
especially in cases where the circular cowl was abruptly connected to a
square fuselage. A break of this sort greatly increases the head
resistance. A "spinner" or propeller cap marked Z in Fig. 3 is an aid in
reducing the resistance offered by the motor cowl and also reduces the
resistance of the inner, and ineffective, portion of the propeller
blades. The cap in any case is smaller than the cowl opening in order
that cooling air be admitted to the enclosed cylinders.
*Distribution of Loads*. Returning to Fig. 1, we note that the weight of
the fuselage, pilot, passenger, fuel, control surfaces and motor are
carried to the upper wing W by the "cabane" strut members C-C and stays,
the lower wing being connected directly into the sides of the fuselage.
Continuations of the cabane members on the interior of the fuselage
inter-connect the upper and lower wings (shown in dotted lines). The
interplane stays in connection with the cabane members bind the wings
and fuselage into a unit. A vertical line "CP" passes through the center
of pressure of both wings, and approximately through the center of
gravity of the machine. In other Words, the machine is nearly balanced
on the center of pressure line. The turning moments of weights behind
the CP must approximately balance the opposite turning moments of the
masses located in front of the CP. The exact relation between the center
of pressure and the center of gravity will be taken up later.
Variable loads such as the passenger, gasoline and oil, are placed as
nearly as possible on the center of pressure line, so that variation in
the weight will not affect the balance. In the figure, the passenger’s
seat A, and the fuel tank F are on the CP line, or nearly so. Thus, a
reduction in the weight of the fuel will not affect the "trim" of the
machine, nor will a wide variation in the weight of the passenger
produce any such effect. As shown, the passenger’s seat is placed
directly on the top of the fuel tank, an arrangement widely used by
European constructors. In the majority of American machines the fuel
tank is placed at the top of the fuselage instead of in the position
illustrated. As the pilot is considered as a constant weight, his
location does not affect the balance.
When at rest on the ground, the weight of the rear end of the fuselage
is supported by the tail skid N. The length of this skid must be such
that the tail surfaces are kept well clear of the ground. The center of
the chassis wheel Q is placed in front of the center of gravity so that
the weight of the machine will cause the tail skid to bear on the ground
when the machine is at rest. If the wheel were behind the center of
gravity, the machine would "stand on its nose" when making a landing.
The wheels must be located so that the tendency to "nose over" is as
small as possible, and yet must not be set so far forward that they will
cause an excessive load on the tail skid. With too much load on the
skid, the tail will not come up, except after fast and prolonged
running, and heavy stresses will be set up in the framework due to the
tail bumping over the ground at high speed. The skids should not be
dragged further than absolutely necessary, especially on rough ground.
With proper weight and wheel adjustment, the tail should come up in a
short run. The wheel adjustment will be taken up under the head of
"Chassis."
*Position in Flight*. In normal horizontal flight, the center line of
thrust CT is horizontal or nearly so. This line of thrust passes through
the center of the motor crankshaft and propeller. In an upward climb,
the CT is inclined at the angle of climb, and since the CT indicates the
line of flight, the streamline curves of the body should be laid out so
that the axis of least body resistance will coincide with the line of
thrust. When flying horizontally at the normal speed, the body must
present the minimum of resistance and the wings must be at the most
efficient angle of incidence. In climbing, or flying at a very low
speed, the tail must necessarily be depressed to gain a large angle of
wing incidence, and hence the body resistance will be comparatively high
owing to the angle of the body with the flight line. It is best to have
the least resistance of the fuselage coincide with the normal horizontal
flight speed. This condition at once establishes the angle of the wings
in regard to the fuselage center line.
*Center Line of Resistance*. The center line of thrust should pass
through the center of total head resistance as nearly as possible. The
total resistance referred to is composed of the wing drag, body and
chassis resistances. In an ordinary military type of aeroplane this line
is located approximately at one-third of the gap from the bottom wing.
Owing to variations in the drag of the Wings at different angles, this
point varies under different flying conditions, and again, it is
affected by the form and size of the fuselage and chassis. The exact
location of the center of resistance involves the computation of all of
the resistance producing items.
In addition to passing through the center of resistance, the center line
of thrust should pass slightly below the center of gravity of the
machine. In this position the pull of the motor tends to hold the head
up, but in-case of motor failure the machine immediately tends to
head-dive and thus to increase its speed. The tendency to dive with a
dead motor automatically overcomes the tendency to "stall" or to lose
headway. With the centerline of thrust determined, and with given motor
dimensions, the fuselage position can be located at once in regard to
the wings. This is good enough for a preliminary layout, but must be
modified in the final design. As before explained, the centerline of
thrust is located at a point between the two wings, approximately
one-third of the gap from the lower wing.
In machines having a span of 35 feet and over, it is a trifle less than
one-third the gap, while in small speed scouts it is generally a trifle
over one-third. This rule checks very closely with the data obtained
from 22 standard machines. Thus, in a machine having a 6-foot gap, the
propeller centerline will be located about 2 feet above the lower wing.
The top of the fuselage (measured under the stabilizer surface) is from
5 to 8 inches above the center line of thrust. At the motor end, the
height of the fuselage above the CT is controlled by a number of
factors, either by the type of motor, or by the arrangements made for
access to the motor parts. In a number of European machines, the motor
sits well above the top of the fuselage, this always being the case when
a six-cylinder, vertical water-cooled motor is used. With an air-cooled
type, the top is governed by the cowl diameter.
*Motor Compartment*. The space occupied by the motor and its accessories
is known as the "motor compartment," and in monoplane and tractor
biplane fuselage it is located in the extreme front of the body. The
interior arrangement varies with different types of motors and makes of
machines. With rotary cylinder motors, the "compartment" is often
nothing but a metal cowl, while with large water-cooled motors it
occupies a considerable portion of the body. Water-cooled motors are
generally covered with automobile type hoods, these usually being
provided with "gilled" openings for ventilation. Owing to the heat
generated in the motor, some sort of ventilation is imperative at this
point. Whatever the type, the compartment is always cut off from the
rest of the fuselage by a fireproof metal partition to guard against
fire reaching the passenger or fuel tanks. The fuel and oil should
always be separated from the motor by substantial partitions since a
single carbureter "pop" may cause serious trouble.
[Illustration: Fig. 4. Mounting and Cowls for Rotary Cylinder
Motors. Courtesy "_Flight_."]
Accessibility is a most important feature in the design of the motor
end, and hence the hood should be of the hinged automobile type so that
it can be easily raised for inspection or repairs. In the Curtis JN4-B
Military Training Tractor, the cylinder heads and valve mechanism
project slightly above the top of the hood so that these parts are amply
cooled and are entirely accessible. Access to the carbureter can be had
through a small hand-hole in the side of the hood. The radiator in this
Curtiss model is located in the extreme front end of the
fuselage—automobile fashion. The propeller shaft passes through a
central opening in the radiator. In Fig. 6 the vertical motor E is set
down low in the frame, the upper part of the fuselage F ending at H. The
engine bearer B, which carries the motor, forms the top part of the
fuselage at this point. The engine is thus in the clear and access can
easily be had to every part of the motor. The radiator is in front of
the motor at R. When in flight the motor is covered by a sheet metal
hood similar to the folding hood used on automobiles. This type is used
in the Martin, Sturtevant, and several European machines.
[Illustration: Fig. 5. Motor Compartment of a Curtiss Tractor
Biplane Using a Front Type Radiator. Note the Two Exhaust Pipes
Which Carry the Gases Over the Top of the Wings.]
Fig. 7 is a very common front end arrangement used with side radiators.
The top fuselage member F is brought down in a very low curve, leaving
the greater part of the motor projecting above the fuselage. At the
extreme front, the upper fuselage member F joins the engine bearer B,
the connection being made with a pressed steel plate. The radiator R is
shown at the side of the fuselage. The cylinders are not usually covered
when in flight. In the front view it will be noted that the radiators
are arranged on either side of the fuselage. A side view of the H. and
M. Farman Fighter is shown by Fig. 10. This is a very efficient French
machine which has seen much active service in the war. The front end is
much like that shown in Fig. 7 except that a spinner cap is fitted to
the propeller boss. A "V" type motor allows of the radiator being
mounted between the two rows of cylinders, and in a position where it
will cause the least possible head resistance.
[Illustration: Figs. 6-7. Various Motor Arrangements, and
Radiator Locations.]
[Illustration: Fig. 10. A Type of H. and M. Farman Tractor
Fighting Biplane. This Machine Uses a "V". Type Motor, with the
Radiator in the Valve Alley. The Gunner and the Machine Gun Are
Mounted in the Rear Cock-pit. It Will Be Noted That the Body Is
Raised Well Above the Lower Wing So that the Gun Field Is
Increased. The Pilot Is Well Ahead of the Entering Edge of the
Lower Wing. Courtesy of "_Aerial Age_."]
The fuselage is of excellent streamline form and shows careful study in
regard to the arrangement of the power plant. Unlike the majority of
machines, the fuselage is raised well above the bottom wing, this being
done evidently to increase the range of the gun in the rear cockpit. The
increased height allows the gun to shoot over the top plane at a fairly
small angle, and the height above the ground permits the use of a very
large and efficient propeller without having an excessively high
chassis. At the rear the fuselage tapers down to a very thin knife edge
and therefore produces little disturbance.
[Illustration: Fig. 11. Radiator Mounted at Leading Edge of
Upper Wing. This Type Is Used with the Sturtevant and Lawson
Aeroplanes and Is Very Effective Because of the Improved
Circulation.]
Fig. 11 shows a Sturtevant Training Biplane in which the radiator is
mounted at the front edge of the upper plane. This arrangement was
originally introduced by the Sturtevant Company in their steel biplanes
and has proved a very efficient type for cooling, although the radiator
must affect the lift of the top plane to a very considerable extent.
*Pilot and Passenger Compartments*. These compartments contain the
seats, controls, and instruments, and in the military types contain the
gun mounts and ammunition. In some battle-planes, the passenger or
"observer" occupies the rear seat, as this position gives a wider range
of fire against rear or side attacks. This arrangement is true of the H.
and M. Farman machine just illustrated and described. In the large
German "Gotha" the gunner occupies the rear position and fires through,
or above, a tunnel built through the rear end of the fuselage. In some
forms of training machines, the pilot and passenger sit side by side
instead of in tandem, as this arrangement allows better communication
between the pilot and student, and permits the former to keep better
watch over the movements of the student. A notable example of this type
is the Burgess Primary Trainer. A side-by-side machine must have a very
wide fuselage and therefore presents more head resistance than one with
the seats arranged in tandem, but with proper attention to the
streamline form this can be reduced so that the loss is not as serious
as would be imagined from a view of the layout.
The seats may be of several types, (a) the aluminum "bucket" type
similar to, but lighter than, the bucket seats used in racing
automobiles; (b) the woven wicker seat used in many types of German
machines, or (c) the modified chair form with wooden side rails and
tightly stretched leather back and bottom. Whatever the type, they
should be made as comfortable as possible, since the operation of a
heavy machine is trying enough without adding additional discomfort in
the form of flimsy hard seats. In the older machines the seats were
nothing more than perches on which the pilot balanced himself
precariously and in intense discomfort. A few pounds added in the form
of a comfortable seat is material well spent since it is a great factor
in the efficient operation of the machine. Wicker seats are light,
yielding and comfortable, and can be made as strong or stronger than the
other types. It seems strange that they have not been more widely
adopted in this country.
All seats should be slightly tilted back so that the pilot can lean back
in a comfortable position, with a certain portion of his weight against
the back of the seat. Sitting in a rigid vertical position is very
tiring, and is especially so when flying in rough weather, or on long
reconnaissance trips. The backs of the seats should be high and head
rests should be provided so that the pilot’s head can be comfortably
supported against the pressure of the wind. If these head rests are
"streamlined" by a long, tapering, projecting cone running back along
the top of the fuselage, the resistance can be considerably reduced.
This arrangement was first introduced in the Gordon-Bennett Deperdussin
and has been followed in many modern machines. In the Deperdussin, the
pilot’s head was exposed directly to the full blast of the propeller
slipstream and a head support was certainly needed. Small, transparent
wind shields are now fastened to the front edge of the cockpit openings
which to a certain extent shield the pilot from the terrific wind
pressure. These are quite low and present little resistance at high
speeds.
[Illustration: Fig. 12. Hall-Scott Motor and Side Type Radiator
Mounting on a Typical Tractor.]
A heavy leather covered pad, or roll, should be run entirely around the
edge of the cockpit opening as a protection to the pilot in case of an
accident or hard landing. The roll should be at least 3 1/2" in diameter
and should be filled with horse hair. All sharp edges in the cockpit
should be similarly guarded so that in the event of the pilot being
thrown out of his seat, he will not be cut or bruised. Each seat should
be provided with an improved safety strap that will securely hold the
pilot in his seat, and yet must be arranged so that it can be quickly
and easily released in an emergency. In flight the occupants must be
securely strapped in place to prevent being thrown out during rapid
maneuvers or in rough weather. Buckles should be substantial and well
sewed and riveted to the fabric so that there will be no danger of their
being torn out. The straps must be arranged so that they will not
interfere with the free movement of the pilot, and so that they will not
become entangled with the controls. It is best to copy the sets approved
by the government as these are the result of long continued experiment
and use.
[Illustration: Fig. 13. Deperdussin Monoplane with Monocoque
Body. Note the Streamline Form of the Body and the Spinner Cap
at the Root of the Propeller.]
Care must be taken to have the seats located at the correct height from
the floor so that the legs will not become cramped. In the majority of
machines, the vertical rudder is operated by the feet. Unless the seat
is at the proper height, the pilot will be in a strained position as he
cannot shift around nor take his feet from the rudder bar. Either the
rudder bar or the seat should be adjustable for different lengths of
legs. Usually the adjustment is made in the rudder bar since it is not
usually advisable to shift the seats owing to the necessity of having
the pilot’s weight in a fixed position. In some old types of monoplanes,
the pilot sat on a small pad placed on the floor of the fuselage.
Needless to say, this was a horribly uncomfortable position to be in,
but as the flights of that time were of short duration it did not matter
much. If the feet could be removed occasionally from the rudder bar the
matter of seat position would not be of so much importance, but to sit
flat on the floor, with the legs straight out, for a couple of hours is
a terrible strain and has undoubtedly caused many accidents through
cramps.
As both the passenger and the fuel are varying weights, the fuel tank
seat idea is good. This allows both of these items to be placed at the
center of gravity of the machine where weight variation will have no
effect on the balance of the plane. In this position, however, the fuel
must be pumped up to a higher auxiliary tank since the main tank would
be too far below the carbureter for gravity feed.
The flooring of the cockpits can be either of veneer, or can be built up
of small spruce slats about 1/2" x 1/2", the slats being spaced about
1/2" apart. The latter floor is specified by the government for seaplane
use, and is very light and desirable. The floor is placed only at points
where it will be stepped on. Observation holes are cut in the floor on a
line with the edge of the seats so that the occupants can view the
ground without looking over the edge of the fuselage. The observation
port holes are about 9 inches in diameter. Glass should never be used in
the cockpits except for the instrument covers, unless it is of the
non-splintering "triplex" laminated type of glass. The use of
inflammable celluloid should also be avoided as being even more
dangerous than the glass. The triplex glass is built up of two or more
layers of glass, which are cemented together with a celluloid film
applied under heavy pressure. This form of construction is very strong,
and while it can be broken, it will not fly apart in the form of
splinters.
All instruments should be placed directly in front of the pilot so that
he can take observations without turning his head. Usually all of the
instruments, with the exception of the compass, are mounted on a single
"instrument board" placed in front of the pilot and directly under the
forward edge of the cockpit opening. The compass can be placed on the
floor as in American machines, or inserted in the upper wing as in some
European machines. The motor control apparatus is placed where it can be
reached conveniently from the seat. Oil gages, gasoline gages, and
revolution counters are generally placed on the instrument board where
they can be easily observed. If a wireless set is carried, the switches
are placed on, or near, the instrument board. Owing to the uses to which
the different machines are put it is impossible to give a list of
instruments that would be suitable for every machine. The simplest
machine should have the following instruments:
*Altimeter*
for measuring altitude.
*Clock*
of special aeroplane type.
*Incidence indicator.*
*Air speed indicators*
for measuring the speed of the machine relative to the air.
*Gasoline, oil and pressure gages*
for determining amount of fuel.
Instruments for Navy and Army machines are of course more complete. In
the specifications for Army Hydroaeroplanes (twin motor type 1916) the
following instruments are specified:
*Aneroid Barometer*.
Graduated in feet, and reading from sea-level to 12,000 feet.
*Compasses*.
One in each cockpit. To be of the Sperry Gyroscopic type with an
elastic suspension and properly damped. Shall be attached to, and
synchronized with, the ground drift indicator.
*Ground Drift Indicator*.
Located in observer’s cockpit. For noting drift due to side
winds. See illustration on page 244.
*Clock*.
Special aeroplane type, built to resist vibration.
Located in pilot’s cockpit.
[Illustration: Fig. 14. Aeroplane Compass of the McCreagh-Osborn
Type. (Sperry)]
*Gasoline Supply Gage*.
To indicate the amount of fuel in gravity service tank at all
attitudes of flight, and shall be visible from pilot’s seat. A
gage in the main tank will also be desirable that will register
the approaching exhaustion of fuel. This indicator should
register when 75% of the fuel in the main tank is exhausted, and
then record the remainder continuously.
*Air Speed Indicator*.
One in pilot’s cockpit.
*Angle of Incidence Indicator*.
Sperry type. To be located in pilot’s cockpit.
*Inclinometer*.
For measuring angle of inclination of longitudinal axis of
machine. In pilot’s cockpit, and placed on instrument board in
the vicinity of tachometers.
*Bank Indicator*.
For indicating the proper amount of bank on turns. In pilot’s
cockpit.
*Map Board*.
One revolving map board placed in pilot’s cockpit.
*Map Desk*.
One folding map desk in observer’s cockpit.
[Illustration: Fig. 15. Sperry Ground Drift Indicator.]
*Tachometers*.
For measuring speed of motors in revolutions per minute. Pilot’s
cockpit.
*Self-Starter Switch*.
For operating self-starter. On instrument board in pilot’s
cockpit.
Among the other accessories specified in the cockpit for the above
machines are a Pyrene fire extinguisher; a 2-gallon water breaker; a
speaking tube for communication between the pilot and observer (1" to 1
1/8"); a flashlight signal for speaking tube; and a tool kit. The weight
of the tool kit shall not exceed 11 lbs.
[Illustration: Fig. 16. Cock-pit of a "London and Provincial"
Biplane. Control Lever in Foreground and Instrument Board Under
Cowl. Courtesy of "_Flight_."]
*General Proportions of the Fuselage*. The total length of the fuselage
depends upon the type of power plant, upon the span or chord of the
wings, and upon the arrangement of the tail surfaces. The rear end of
the fuselage should be far enough away from the wings to insure that the
rear surfaces are not unduly affected by the "down-wash" or the
"wake-stream" of the wings. A very short fuselage gives a short lever
arm to the control surfaces, hence these surfaces must be very large
with a short body. With the stabilizer surface close to the wings, the
"damping" effect is slight, that is, the surface does not effectively
kill or "damp down" oscillations. Large tail surfaces are heavy,
difficult to brace, and cause a very considerable amount of head
resistance. The extra weight of a long fuselage is generally offset by
the increased weight caused by the large tail area of the short body
type.
When machines are crated and shipped at frequent intervals, a very long
fuselage is objectionable unless it is built in two sections. It also
requires much storage space and a very long hangar. Machines for private
use must often be sacrificed from the efficiency standpoint in order to
keep the dimensions within reasonable limits. An aeroplane requiring an
enormous hangar has certainly no attraction for the average man. Every
effort must be made to condense the overall dimensions or to arrange the
extremities so that they can be easily dismounted. Exhibition flyers
require specially portable machines since they ship them nearly every
day, and the expense of handling a long awkward fuselage may alone
determine the choice of a plane. It is usually best to divide the body
at a point just to the rear of the pilot’s seat, although many flyers
look upon a two-part body with disfavor unless they can be shown that
the joint connections are as strong as the rest of the fuselage.
[Illustration: Fig. 19. Fuselage Dimension Chart for Two Place
Aeroplane Fuselage. Upper Diagram Is the Water Cooled Type and
the Lower Figure Applies to a Machine with a Rotating Air Cooled
Motor. See Table of Dimensions on Page 248.]
As a guide in the choice of fuselage proportions, a set of diagrams and
a table are attached which gives the general overall dimensions of
several prominent makes of machines. The letters in the diagram refer to
the letters heading the columns in the tables so that the general
dimensions of any part can be readily determined. I do not claim that
these dimensions should be followed religiously in every case, but they
show what has been done in the past and will at least suggest the limits
within which a new machine can be built.
[Illustration: Fig. 20. Fuselage Dimension Diagrams Giving the
Principal Dimensions of Speed Scout Machines. Upper Figure Shows
Typical Scout with Water Cooled Motor (Curtiss), While Lower
Diagram Shows an Arrangement Common with Rotary Air Cooled
Motors (Nieuport)]
Fig. 19 gives the outline and dimension letters for two-place machines
of what is known as the "Reconnaissance Type." Both water-cooled and
air-cooled motor equipments are shown, the top machine being of the
water-cooled type while the lower figure shows a typical two-place
machine with a rotary air-cooled motor. Underneath this side elevation
is a front view of the fuselage, and a section taken through the point
of greatest depth. As shown, the fuselage is of square cross-section,
but the dimension B applies equally to the diameter of a circular
cross-section. Dimension C gives the height of the curved upper deck, or
"turtle deck" of the fuselage. Dimension D shows the extreme length
extending in front of the leading edge of the lower wing, and T shows
the length of the rear portion back of the leading edge of the lower
wing, the leading edge being taken as a base of measurement. The
location of the deepest section, measured from the extreme front of the
fuselage, is given by E, the depth at this point being indicated by B.
The extreme width is shown by I. In the machine shown, side radiators
are used, the blunt front end dimensioned by G and K being the dimension
of the front engine plate. When front radiators are used, the dimensions
G and K also apply to the size of the radiator. The amount of advance,
or the distance of the chassis wheel center from the leading edge is
given by S, and the distance of the wheel center below the leading edge
is given by R. V is the length of the engine projecting above the
fuselage top. The passenger or observer is indicated by 1 and the pilot
by 2. The top plane is 3 and the bottom plane 4. The engine is located
by 5, and the top fuselage-rail, or "longeron," by 7. Turtle deck is 6.
[Illustration: * Round monocoque body, dimensions (B) and (F)
measured from outer diameter or top of circle]
Fig. 20 gives the diagrams of speed scout machines, both of the
air-cooled and water-cooled types. These are the small, fast, single
seat machines so much used in the European war for repelling air attacks
and for guarding the larger and slower bombing and observation machines.
The upper drawing shows a Curtiss Speed Scout equipped with a "V" type
water-cooled motor and a circular front radiator. While the front of the
body is circular, it gradually fades out into a square cross-section at
the rear. The lower machine is a Nieuport speed scout equipped with a
rotating cylinder air-cooled motor. In this scout, the diameter of the
motor cowl is given by dimension K. Body is of square cross-section. It
will also be noted that the chord of the lower plane is less than that
of the upper plane and that the deep body almost entirely fills the gap
between the two wings.
[Illustration: Fig. 21. Curtiss J N 4-B Fuselage Boxed for
Shipment.]
With some of the later speed scouts, the body entirely fills the gap
between the wings and the top plane is fastened directly to the top
members of the fuselage. This makes windows necessary in the sides of
the fuselage. When vertical water-cooled motors are used on speed
scouts, the front view is entirely cut off, for these are very large
motors and project above the fuselage for a considerable distance. This
obstruction is avoided in the Curtiss speed scout shown, by the use of a
"V" type motor. It will be noted that these two scouts, especially the
Nieuport, are of excellent stream line form, a very important item with
such high speed machines. The propeller of later Nieuports is provided
with a conical spinner cap which evidently reduces the head resistance
to a considerable extent. The different portions of the machine are
indicated by the same figures as in the case of the reconnaissance
machine.
CHAPTER XII DETAILS OF FUSELAGE CONSTRUCTION
Classification of types. While there are a number of methods adopted in
building up the fuselage structure, the common type is the "wire truss"
in which wood compression members are used in connection with steel wire
or cable tension members. Four wooden "longerons" or "longitudinals" run
the entire length of the body and are bent to its general outline. The
longitudinals are spaced at the correct distance by wood compression
members, which in turn are held in place by wire cross bracing. This
method of trussing forms a very strong and light structure, although
rather complicated, and difficult to build. The cross-section is
rectangular, although in many cases the body is converted into a
circular or elliptical section by the use of light wood formers fastened
to the main frame.
Another well known type is the "Monocoque" body, first used on the
Gordon-Bennett Deperdussin monoplane. This fuselage is a circular shell
built up of three-ply tulip wood, thus forming a single piece body of
great strength. The three-ply shell is really a veneer, the layers
proceeding spirally around the body, each layer being securely glued to
its neighbor. Between each layer is a scrim layer of treated silk, and
another fabric layer is generally glued to the outside of the shell. The
shell is very thin, the total thickness of the three layers of wood and
fabric in modern machines being rather less than 1.5 millimeters (about
1/16 inch). In the original "Deps" this was somewhat greater, 0.15 inch.
Monocoque construction as a rule is heavy and expensive, but offers the
great advantage of strength, perfect alignment at all times, and of
offering resistance to rifle and shell fire. If the longitudinals of a
truss type are struck with a bullet, or shell fragment, the entire
fuselage is likely to fail, but a monocoque body may be well perforated
before failure is likely to take place.
The American L. W. F. Tractor Biplane has a monocoque body in which
spruce laminations are used instead of hardwood. One ply runs
longitudinally while the other two layers are spiralled to the right and
left respectively. Between each layer is a scrim layer of treated silk,
the whole construction being covered with a final layer of fabric,
several coats of waterproof compound, and four final coats of spar
varnish. When used for seaplanes the wood plies are stitched together
with strong wires to prevent separation due to dampness. Since spruce is
used in place of hardwood, the construction is lighter than in European
models, and the L. W. F. Company claim that it is lighter than the usual
truss construction. An additional advantage of the monocoque
construction is that the pilot is protected against splinters or
penetration by the limbs of trees when making a forced landing in the
brush.
Another form of monocoque construction was adopted by the French
builder, Bleriot, at the beginning of the war. The fuselage of this
machine was covered with papier-mache, the ash longitudinals being
buried in this mixture. The papier-mache is built up with glue and silk
threads. This construction is very light and strong, but is expensive
and difficult to protect against moisture. The front of the fuselage is
protected with a 3 millimeter steel armor plate to protect the pilot
against bullets and shrapnel. The papier-mache portion of the body is
not easily splintered by bullets.
A third form of monocoque, experimented upon by the author, is the steel
shell type in which the three-ply wood veneer is supplanted by a thin
steel shell. This outer shell is strengthened by suitable stiffener
angles. With a shell thickness of 0.013 inch, the strength is equal to
the strength of a wood shell and is slightly less in weight. It has the
advantage of being easily and cheaply formed into shape and is
absolutely proof against the influences of heat and moisture. It cannot
splinter, will not catch fire and offers a maximum resistance against
penetration. There is yet much experimental work to be done before the
construction is perfected.
About midway between the truss fuselage and the monocoque is the veneer
construction used on many of the modern German aeroplanes. In general,
this may be described as being a veneer shell fastened to the
conventional wood longitudinals. Stay wires are not in general use, the
veneer taking the shear due to the bending movement. Six longerons are
used instead of four, the two additional members being located midway on
the vertical sides. Transverse wood frames take the place of the
transverse stay wires used in the truss type. Examples of this type are
met with in the "Albatros de Chasse" and in the "Gotha" bomb dropper.
The single seater, "Roland," has a fuselage of circular section, with a
true monocoque veneer construction, but German-like, reinforces the
construction with a number of very small longitudinals. In this machine
there are 6 layers, or plies, of wood reinforced by fabrics. The entire
thickness of the wood and fabric is only 1.5 millimeters (1/16 inch).
Steel tube fuselage dates back to the beginning of the aeroplane
industry. In this type the wood longitudinals of the wood truss type are
replaced with thin gage steel tubes, the cross struts being also of this
material. The diagonal bracing may be either of steel wire, as in the
wood frames, or may be made up of inclined steel tube members that
perform both the duty of the stay wires and struts. For the greatest
weight-efficiency, a steel tube body should be triangular in section
rather than square. A triangular section saves one longitudinal and a
multitude of wire struts and connections since no transverse bracing is
necessary. Connections on a steel tube fuselage are difficult to make
and are heavy. They require much brazing and welding with the result
that the strength is uncertain and the joint is heavy.
A very modern type of steel construction is that developed by the
Sturtevant Company. The members of the Sturtevant fuselage are in the
form of steel angles and channels, similar in many respects to the
sections used in steel buildings and bridges. The joints are riveted and
pinned as in steel structural work. The longitudinals are angles and the
struts are channels. Crystallization of the steel members is prevented
by the use of special pin-connected joints provided with shock absorbing
washers. Owing to the simplicity of the riveted joints, there is
practically no weight due to connections, and since the weight of
connections is a large item in the total weight of a fuselage, the
Sturtevant is a very light structure. According to G. C. Loening,
engineer of the company, the fittings of a large wood fuselage weigh at
least 60 pounds. This is almost entirely saved with the riveted
connections.
[Illustration: Truss Type Fuselage of Curtiss R-4 Biplane,
Showing Motor and Front Radiator Mounted in Place. It Will Be
Noted That the Upper and Lower Longerons Are Channeled Out for
Lightness and Hence These Members Are of the "I" Beam or Channel
Form. Propeller Flange Is Shown Projecting Through the Radiator
Opening.]
A novel type of wood fuselage has been described by Poulsen in "Flight."
Eight small longitudinals are used which are held in place by three-ply
wooden formers or diaphragms. Wire bracing is used in a longitudinal
direction, but not transversely in the plane of the diaphragms. The
cross-section is octagonal, and the completed structure is covered with
fabric. For the amateur this offers many advantages since the wiring is
reduced to a minimum and all of the members are small and easily bent to
shape. It is fully as light as any type of body, for the connections are
only thin strips of steel bolted to the diaphragms with small machine
screws. No formers are needed for the deck, and the machine can be given
a close approximation to the ideal stream-line form with little trouble.
*Truss Type Fuselage*. We will now take up the construction of the truss
type of fuselage in more detail, and investigate the merits of the
different types of connections used in fastening the frame together.
Like every part of the aeroplane, the fuselage must either be right or
wrong, there is no middle course. Fig. 23 shows a side elevation of a
typical truss type fuselage built up with wood longitudinals and struts,
the tension members being high tensile strength steel wire and cable. L
and L’ are the upper and lower longitudinals, S-S-S are the vertical
struts, and T-T-T are the horizontal cross struts which run across the
frame. The engine bed is the timber marked B at the front of the body.
The upper wing is attached to the body through the "cabane" struts C,
and the chassis connections are shown at D. The stern post E closes the
rear end of the body in a knife edge and acts as a support for the
rudder and the rear end of the stabilizer. F is the seat rail which
carries the seats and supports the control yokes.
All cross bracing is of high tensile strength steel wire, or of high
strength aviation cable, these strands taking the tensile stresses while
the wood struts are in compression. In the forward portion, double
stranded cables are generally used, with solid wire applied to the after
portions. The longitudinals are of ash from the motor to the rear of the
pilot’s seat, while the rear longitudinals are generally of spruce. In
some machines, however, the entire length of the longitudinals is ash.
The latter arrangement makes a heavier, but stronger body. The struts
are usually of spruce as this material is stiffer than ash and much
lighter.
[Illustration: Diagram of Typical Truss Type Fuselage, Showing
Principal Members in Place.]
[Illustration: Fuselage of Hansa-Brandenburg Fighting Biplane.
See Page 268. Figs. 27-28-29-30. Fuselage Details. (Truss Type)]
Both the struts and longitudinals are frequently channelled out for
lightness, as shown by Fig. 27, the wooden member being left rectangular
in section only at the points where the connections are made with the
struts and cables. The channelling-out process, if correctly followed,
gives very strong stiff members with a minimum of cross-sectional area
and weight. Many captured German machines, on the contrary, have solid
longitudinals of rectangular section, wrapped with linen fabric. This
fabric strengthens the construction and at the same time reduces the
chances of splintering the wooden members in a hard landing. The fabric
is glued to the wood and the entire wrapping is then given several coats
of a moisture repelling varnish. In the older types of fuselage, the
longitudinals were often of the "laminated" class, that is, were built
up of several layers of wood glued together in a single rectangular
mass. This reduced the tendency toward splitting, but was very
unreliable because of the uncertainty of the glued joints when exposed
to the effects of heat and moisture. Laminated longitudinals are now
seldom used, particularly in the region of the motor where water and oil
are certain to wreck havoc with the glued up members.
As the stresses rapidly diminish toward the tail, it is the general
practice to taper down the section of the longitudinal toward the rear
and to reduce the section of the struts. The longitudinals are generally
kept constant in section from the motor to the rear of the pilot’s seat,
the taper starting at the latter point and continuing to the rear end.
For example, if the longitudinal section at the motor is 1 1/4" x 1
1/2", the section at the rear will be 1" x 1", the width of the struts
corresponding to this taper. While tapering is very desirable from the
weight standpoint, it makes the fitting problem very difficult since
each fitting must be of a different dimension unless the connections can
be designed so that they are adjustable to changes in the section of the
longitudinals. In one machine, the width and depth of the longitudinals
are kept constant, the variation in weight and section being
accomplished by increasing the depth of the channelling as the rear is
approached. With this design, the same fittings can be applied from one
end to the other.
[Illustration: Figs. 27-28-29-30. Fuselage Framing Members and
Details.]
Since the loading of the struts is comparatively light, they can be much
reduced in section by channelling or by chamfering, as shown by Fig. 28.
If the width and thickness is maintained, much of the interior material
can be removed without danger of reducing the strength. Sketch (A) in
Fig. 28 shows a very common method of strut reduction, the strut being
of rectangular section throughout its length, but tapered in such a way
that it is thickest at the center (d) and thinnest at the two ends (e).
To obtain the correct relation between the center end thickness requires
very careful calculation. As shown, the strut is attached to the upper
and lower longitudinals by sheet steel fittings or "sockets." Sketch (B)
shows a simple method, the rectangular strut being chamfered off at each
of the four corners, and left full size at either end where the fittings
connect it with the longitudinals. This form is not correct from a
technical standpoint, but is generally good enough for lightly loaded
struts, and has the advantage of being cheaply and easily constructed.
In sketch (C) a channelled strut is shown, the center portion being
channelled out in a manner similar to the channelling of the
longitudinals. This lightening process is most commonly adopted with the
large heavily loaded struts in the front portion of the fuselage, and at
the points where the motor bed is suspended or where the wings and
chassis are attached to the body. The black dots at the ends of the
struts indicate the bolt holes for the fittings, it being permissible to
drill holes in the ends of the struts but not in the longitudinal
members. If the strut is large enough to resist the bending stresses at
the center it will generally allow of holes being drilled near the ends
without danger of strength reduction. Again, the struts are always in
compression and hence the bolts may be depended upon to partly take the
place of the removed material in carrying the compressive stresses.
Holes should never be drilled in the longitudinals since these members
may be either in tension or compression, depending upon the angle at
which the elevator flaps are set. The hole not only destroys the
strength at the point at which it is drilled, but this reduction also
extends to a considerable distance on either side of the hole, owing to
the fibrous nature of the wood. In steel members the effect of the hole
is purely local and does not usually extend much beyond the edge of the
hole. Considering the wood beam as consisting of a series of parallel
fibers, it will be seen that severing any one of the fibers will
decrease the strength of the wood through a distance equal to the length
of the cut fiber, or at least through a distance equal to the natural
shear value of the resins that bind the fibers together.
Fuselage fittings are almost numberless in the variety of design. They
must be very light and strong, must be applied without drilling the
longerons, and should be simple and cheap to construct. They are usually
made of sheet steel of from 0.20 to 0.30 point carbon, and may be either
bent or pressed into shape. At the points where the struts are joined to
the longitudinals, the fittings connect struts and wires in three
planes, the vertical struts and fore and aft wires; the transverse wires
and horizontal struts, and the top and bottom wires that lie in a
horizontal plane. There are at least 6 connections at every strut, four
of the connections being made to the stay wires or cables. A simple
connection is therefore very hard to design.
Fig. 29 shows a typical fuselage "panel" and the interconnected members
in their usual relation. LU and LL are the top and bottom longitudinals
at the right, while LU’ and LL’ are the longitudinals at the right hand
side. The vertical struts SV and SV’ separate the top and bottom
longitudinals, while the horizontal struts SH and SH’ separate the right
and left hand sides of the fuselage body. The wires w-w-w-w brace the
body fore and aft in a vertical plane. The wires t-t lie in a horizontal
plane, produce compression in the horizontal struts SH-SH’, and stiffen
the frame against side thrust. The transverse rectangle SV-SH-SV’-SH’ is
held in shape by the transverse stay wires W-W, this rectangle, and the
stays resisting torsional stress (twisting), act against the struts
composing the sides of the rectangle. In some European machines, the
wires WW are eliminated, and are replaced by thin veneer panels, or
short wood knee braces as shown by Fig. 30. The section shows the
longitudinals L-L-L-L and the struts SV-SV’-SH-SH’ braced by the veneer
sheet or diaphragm D. This diaphragm is well perforated by lightening
holes and effectually resists any torsional stress that may be due to
motor torque, etc. Since the transverse wires W-W in Fig. 29 are rather
inaccessible and difficult to adjust, the veneer diaphragm in Fig. 27
has a great advantage. In this regard it may be stated that wire bracing
is not a desirable construction, and the substitution of solid veneer is
a step in advance.
Wire bracing has always seemed like a makeshift to the author. The
compression and tension members being of materials of widely different
characteristics are not suitable in positions where a strict alignment
must be maintained under different conditions of temperature and
moisture. The difference in expansion between wire and the wood
compression members produces alternate tightness and slackness at the
joints, and as this is not a uniform variation at the different joints,
the frame is always weaving in and out of line. Under the influence of
moisture the wood either swells or contracts, while the wire and cable
maintain their original lengths and adjustments. The result is that a
frame of this kind must be given constant attention if correct alignment
is desired.
The adjustment of a wire braced wood fuselage should be performed only
by a skilled mechanic, as it is easily possible to strain the members
beyond the elastic limit by careless or ignorant handling of the wire
straining turnbuckles. In the endeavor to bring an old warped fuselage
back into line it is certain that the initial tension in the wires can
be made greater than the maximum working stress for which the wires were
originally intended. Shrinkage of the wood also loosens the bond between
the wooden members and the steel fittings unless this is continually
being taken up. Some form of unit construction, such as the monocoque
body, is far more desirable than the common form of wire trussed wood
body.
[Illustration: Fuselage Details of De Havilland V. Single Seat
Chaser. A Rotary Le Rhone Motor Is Used in a Circular Cowl. The
Diagonal Bracing in the Front Section is Reinforced by Laminated
Wood Plates Instead of by Wires. Dimensions in Millimeters.]
*Fuselage Fittings*. In the early days of aviation the fuselage fittings
on many machines were made of aluminum alloy. This metal, while light,
was uncertain in regard to strength, hence the use of the alloy was
gradually abandoned. At present the greater part of the fittings are
stamped steel, formed out of the sheet, and are of a uniform strength
for similar designs and classes of material.
The steel best adapted for the fittings has a carbon content of from
0.20 to 0.30, with an ultimate strength of 60,000 pounds per square
inch, and a 15 per cent elongation. The steel as received from the mill
should be annealed before stamping or forming to avoid fracture. After
the forming it can be given a strengthening heat treatment. A lower
steel lying between 0.10 and 0.15 carbon is softer and can be formed
without annealing before the forming process. This material is very
weak, however, the tensile strength being about 40,000 pounds per square
inch. Fittings made of the 0.15 carbon steel will therefore be heavier
than with the 0.30 carbon steel for the same strength. The thickness of
the metal will vary from 1/32" to 1/16", depending upon the load coming
on the fitting.
A typical fuselage strut fitting is shown by Fig. 31-A in which L-L-L
are the longerons, d is the fitting strap passing over the longerons, S
and So are the vertical and horizontal struts respectively. The stay
wires are fastened to ears (b) bent out of the fitting, the wires being
attached through the adjustable turnbuckles (t). The struts are provided
with the sheet steel ferrules marked (F). There are no bolts passing
through the longitudinals L-L’, the fitting being clamped to the wooden
member. This is very simple and light fitting. Fig. 31-B is a similar
type, so simple that further discussion is unnecessary.
Fig. 32 shows a fuselage strut fitting as used on the Standard Type H-3
Biplane. We are indebted to "Aerial Age" for this illustration. This
consists of a sheet metal strap of "U" form which is bent over the
longitudinal and is bolted to the vertical strut. At either side of the
strut are through bolts to which bent straps attach the turnbuckles.
These straps are looped around the bolts and form a clevis for the male
ends of the turnbuckles.
[Illustration: Fig. 31. Typical Fuselage Strut Fittings.]
An old form of fuselage connection used on the Nieuport monoplane is
shown by Fig. 33, an example of a type in which the bolts are passed
through the longeron member. This fitting is very light but
objectionable because of the piercing of the longeron.
An Austrian aeroplane, the Hansa-Brandenberg, has a wood fuselage in
which no stay wires are used. This fuselage is shown by Fig. 23a. Both
the vertical and inclined members are wood struts. The outer covering of
wood veneer makes the use of stay wires unnecessary since the sheath
takes up all horizontal stresses, and hence forms a sort of plate girder
construction. The German Albatros also employs a wireless veneer
fuselage, the construction being shown in detail by Figs. 36 and 36a.
Three longerons are located on either side of the body, the third member
being placed at about the center of the vertical side. As will be seen,
the veneer makes the use of wire bracing and metal connections
unnecessary. The veneer also insures perfect alignment.
[Illustration: Fig. 32. Fuselage Strut Fittings of the Standard
H-3 Training Biplane.]
*Wing Connections*. The lower wings are attached to the lower
longitudinals by a special sheet steel fitting which also generally
connects to a vertical strut at this point, and to an extra heavy
horizontal strut. A sheet metal clevis, or socket, on the wing spar is
pinned to the fuselage half of the fitting so that the wing can be
easily detached when the machine is to be dissembled. At this point a
connection is also provided for the end of the inner interplane stay
wires. The horizontal strut at the point of wing attachment is really a
continuation of the wing spar and takes up the thrust due to the
inclination of the interplane stays. In the majority of cases the
horizontal thrust strut is a steel tube, with the hinged connection
brazed to its outer ends. This is one of the most important and heavily
loaded connections on the machine and should be designed accordingly.
[Illustration: Fig. 33. Fuselage Fittings of the Nieuport
Monoplane.]
Fig. 37 shows a typical wing to fuselage connection of the hinge type.
The wing spar (G) is covered with a sheet steel ferrule (A) at its inner
end. Two eye bars (B) are bolted to the wing spar, and over the ferrule,
the eyes of the bar projecting beyond the end of the spar. This forms
the wing half of the connecting hinge. The eyes are fastened to the
fuselage hinge member (H) by means of the pin (E). This pin has a
tapered end for easy entry into the joint, and is pierced with holes at
the outer end for cotter pins or a similar retaining device. The
fuselage hinge member (H) is brazed to the end of the steel tube strut
(T). This tube runs across the fuselage from wing spar end to wing spar
end.
Strut tube (T) lies on, and is fastened to, the fuselage longeron (L),
and also lies between the two halves of the vertical strut (S). The
vertical strut is cut out at its lower end for the receipt of the steel
tube (T). A steel plate is brazed to the tube, is wrapped about the
longeron (L) and is bolted to the vertical strut (S). The interplane
stay (F) is attached to the pin (E) at the point of juncture of the wing
spar eye and the fuselage member of the hinge. A collar (I) is brazed to
the tube, and forms a means of attaching the fuselage stays (D). The
drift wires (C) of the wings are attached to an eye at the end of one of
the wing spar bolts. As shown, the fitting (H) is a steel forging, very
carefully machined and reduced in weight. The inside wing ribs are
indicated by (K), from which it will be seen that there is a gap between
the end of the wings and the outside face of the fuselage.
[Illustration: Fig. 36. Veneer Fuselage Construction of the
German "Albatros" Speed Scout. Body Outline Is Obtained by
Veneer Diaphragms and no Stay Wires Are Used.]
Fig. 36-a shows the construction of the wing joint of the German
Albatros machine. The fuselage is of monocoque construction which allows
of a simple attachment to the outer shell. This is a very sturdy and
simple connection. Fig. 38-Z is the wing attachment detail of the
English London and Provincial Biplane (1916), the fuselage in this case
being of the wire trussed wood type. We are indebted to "Flight" for
this illustration.
[Illustration: Fig. 36-a. Details of Albatros Veneer Fuselage
Construction.]
In some machines the interplane stay wires are attached to a lug formed
from the attachment plate, but we do not consider that this construction
is as good as the type in which the wire is attached directly to the
wing spar pin. While the former may be easier to assemble, the
attachment of the wire to the pin eliminates any eccentricity, or
bending moment, due to the pull of the interplane stay. The attachment
in the L. W. F. insures against any eccentricity in the stay attachment,
and at the same time makes the assembly and dismounting a very simple
matter.
[Illustration: Fig. 37. Wing Connection to Fuselage.]
*Chassis Member Attachment*. The attachment of the chassis struts
generally involves some difficulty as these members usually intersect
the line of the longerons at a very awkward angle. If the wing
attachment is near the same point, as it generally is, the detail is
made doubly difficult. The chassis must be pin connected as in the case
of the wing joint so that the chassis members can be easily and quickly
removed. A detail of a chassis to body connection is shown by Fig. 39.
In this figure (L) is the lower longeron, (S) is the vertical fuselage
strut, and (C) is one of the chassis members. The upper end of the
chassis member is enveloped in a sheet steel ferrule (D) which is bolted
in place, and which is provided with a clevis at its upper end for the
attachment pin (P).
A plate (E) is bolted to the fuselage strut (S) and is passed around the
lower longeron (L), a hinge joint (H) being provided for attachment to
the chassis ferrule through the pin (P). Ears or lugs are left at (G-G)
for the attachment of the fuselage stays (B-B). On the inner side of the
plate (E) are attachment lugs for the horizontal strut (H). It will be
noted that the plate (E) is well provided with lightening holes so that
the weight can be kept down to a minimum. The pin is tapered at the end,
and is provided with cotter pin holes. The fitting in general is small,
and does not produce any great degree of head resistance, the small part
exposed being of good streamline form.
[Illustration: Fig. 38x. Wing Connection of the Albatros Reconnaissance
Biplane. Fig. 38y. Wing Attachment of Albatros Fighter with Pin Joint.
Fig. 38z. Wing Connection of London and Provincial Biplane.]
Great care should be taken in brazing or welding these fittings, since
the heat changes the structure of the metal and greatly reduces its
strength. The brazing temperature varies from 1,500 to 1,700 degrees, a
point well above the tempering heat of steel. Attempts have been made to
heat treat the metal after the brazing operation, but with very little
success, owing to the fact that the heat treating temperature is
generally at or above the melting point of the brazing spelter, hence is
likely to cause holes and openings in the brazed joints. With acetylene
welded joints the parts can, and should be, heat treated after the
welding. While this is an apparent advantage of acetylene welding, all
parts cannot be successfully handled in this manner. The welding torch
can only join edges, while the brazing spelter can be applied over
almost any area of surface. Welding is very successful in joining thin
steel tubes while in many fittings made of sheet metal, brazing is the
only feasible operation.
[Illustration: Fig. 39. Chassis Connection.]
Both methods have a common fault, in that they are unreliable. Imperfect
welds and brazing are not always apparent from the outside, actual
breakage of the part being necessary to determine the true nature of the
joint.
FUSELAGE WEIGHTS.
*Distribution of Weight*. The weight of a fuselage depends upon the span
of the wings, upon the seating capacity, and upon the weight and type of
the power plant. The weight also varies considerably with the type of
construction, that is, whether of truss, veneer, or monocoque
construction. A heavily powered machine, or one carrying more than a
single person, requires heavier structural members and hence weighs more
than a small single seater. The amount of fuel carried also has a
considerable bearing on the fuselage weight.
Probably the best method of treating this subject is to give the
fuselage weights of several types of well known machines. The reader
will then have at least a comparative basis for determining the
approximate weight. (Truss type only.)
There are so many variables that the weight cannot be determined by any
set rule or formula. Alexander Klemin in his "Course in Aerodynamics and
Airplane Design" says that the approximate weight of a bare wood truss
type fuselage is about 150 pounds for a machine having a total weight of
2,500 pounds. For small biplane and monoplane scouts weighing
approximately 1,200 pounds total, the bare fuselage frame will weigh
about 70 pounds. These figures are for the bare frame alone and without
seats, controls, tail skids or other fittings. The weights given under
the column headed "Wt. Bare" include the engine beds, tail skids,
flooring, cowling and body covering, and hence exceed the "bone bare"
estimate of Klemin by a considerable amount.
The all-steel fuselage of the large Sturtevant battle-plane (Model A)
weighs 165 pounds inclusive of the steel engine bed. A wooden, wire
braced fuselage of the same size and strength weighs well over 200
pounds, the metal fittings and wires weighing about 60 pounds alone. Ash
is used in the wood example for the longerons. The struts and diagonal
members in the Sturtevant metal fuselage are riveted directly to the
longitudinals, without fittings or connection plates. The safety factor
for air loads is 8, and for the ground loads due to taxi-ing over the
ground, a safety factor of 4 is used.
After a minute comparison of the items comprising the fuselage of the
Curtiss JN4-B and the Standard H-3, Klemin finds that the fuselage
assembly of the Standard H-3 amounts to 13.6 per cent of the total
loaded weight, and that the fuselage of the Curtiss JN4-B is 15.5 per
cent of the total. Tanks, piping and controls are omitted in both cases.
For machines weighing about 2,500 pounds, Dr. J. C. Hunsaker finds the
body weight averaging 8.2 per cent of the total, this figure being the
average taken from a number of machines.
On careful examination it will be found that the fuselage assembly
(bare) amounts to a trifle less than the wing group for biplanes having
a total weight of from 1,900 to 2,500 pounds. The relation between the
wing weight and the fuselage weight seems to bear a closer relation than
between the fuselage and total weights. We will set these different
relations forth in the following table:
PERCENTAGE OF FUSELAGE WEIGHT
──────────────────────────────────────────────────────────────────────
Fuselage Weight as
Percentage of the Total
Load
Name of ─────────────────────────────────── Wing Weight As
Plane or Body Body Body Percentage of Total
Investigator Assembly Assembly Assembly Weight
Bare and and
Equipment Power
Plant
──────────────────────────────────────────────────────────────────────
Curtiss 15.50% 17.86% 43.96% 14.15%
JN4-B
──────────────────────────────────────────────────────────────────────
Standard H-3 13.60% 17.70% 45.40% 14.52%
──────────────────────────────────────────────────────────────────────
J.C. 8.20% 11.50% 34.30% 16.50%
Hunsaker
──────────────────────────────────────────────────────────────────────
Author’s 14.96% 17.62% 46.66% 14.60%
Experience
──────────────────────────────────────────────────────────────────────
Average of 13.06% 16.17% 42.58% 14.94%
Above
──────────────────────────────────────────────────────────────────────
In the above table, the column headed "Body Assembly and Equipment"
includes the body frame, controls, tanks and piping. In the fourth
column, the radiator, motor, propeller, water, and exhaust pipe have
been added. For the average value it will be seen that the bare fuselage
is about 1.88 per cent lower than the weight of the wings. It should be
noted that the wing weight given is the weight of the surfaces alone,
and does not include the weight of the interplane struts, wires and
fittings. The weight of the wing surfaces as above will average about
0.75 pounds per square foot.
Based on the above figures, we can obtain a rough rule for obtaining the
approximate weight of the fuselage, at least accurate enough for a
preliminary estimate. If A = the total area of the wings, then the total
weight of the wings will be expressed by w = 0.75A. The weight (f) of
the fuselage can be shown as f = 13.06/1494 x w = 0.65A.
*Example*. The area of the Standard H-4 is 542 square feet total. Find
the approximate weight of the fuselage. By the formula, f = 0.65A = 0.65
x 542 = 352 pounds. The actual bare weight is 302.0 pounds. For several
other machines, the actual weight is greater than the weight calculated
by the formula, so that the rule can be taken as a fair average,
especially for a new type that is not as refined in detail as the H-3.
SIZE OF LONGERONS
The size of the longerons, that is, the section, is influenced by many
factors. As these members must resist flying loads, the leverage of
elevator flaps, stresses due to control wires, landing stresses and the
weight of the motor and personnel it is always advisable to itemize the
loading and then prepare a diagram to obtain the stresses in the
different members. This latter method is a method for a trained
engineer, but an exhaustive description of the method of procedure will
be found in books on the subject of "Strength of Materials." For the
practical man, I give the following list of longeron dimensions so that
he will have at least a guide in the selection of his material.
The length of the fuselage and power of motor are given so that the
reader can obtain sizes by comparison, although this is a crude and
inaccurate method. As the longerons taper from front to back, the sizes
of the section are given at the motor end, and also at the tail. The
size of the front members depends principally upon the weight of the
motor and the passenger load, while the rear longerons carry the
elevator loads and the tail skid shock. If the rudder is high above the
fuselage it introduces a twisting movement that may be of considerable
importance. The loads on the stabilizer, elevators and the vertical
rudder are very severe when straightening out after a steep dive or in
looping, and the pull on the control wires exerted by the aviator at
this time greatly adds to the total stress. In the front of the
fuselage, the motor exerts a steady torque (twist) in addition to the
stress due to its weight, and to this must be added the gyroscopic force
caused by the propeller when the machine is suddenly changed in the
direction of flight. The combination of these forces acting at different
times makes the calculation very difficult.
In the case of the Curtiss R-4, the front longerons taper down from the
motor 1.63" x 1.25" to a point directly behind the pilot’s seat, the
section at the latter point being 1.25" x 1.25". From this point the
rear longerons taper down to 1" x 1" at the tail. At the motor, the
section is 1.63" x 1.25". The longitudinals of the Bleriot monoplane are
laminated and are built up of alternate layers of spruce and ash. This
is an old type of machine and this practice has since been discontinued.
It will be noted that as the power is increased, the size of the front
longerons is generally increased, although this is not always the case
in speed machines. The Chicago Aero Works’ "Star" fuselage could easily
carry a 90 horsepower motor, although this size is not regularly
installed.
*Pusher Type Fuselage (Nacelle)*. Compared with the tractor biplane and
the monoplane fuselage, the body of the pusher is very short and light.
The latter body simply acts as a support for the motor and personnel
since the tail loads are carried by the outriggers or tail booms. The
motor is located at the rear end of the body and may be either of the
air or water-cooled type. The accompanying figure shows a typical pusher
type body, or "Nacelle" as it is sometimes called.
The advantages of the pusher type for military service are obvious. The
observer or gunner can be placed immediately in the front where his
vision is unobstructed, and where the angle of fire is at a maximum.
[Illustration: Typical Pusher Body Showing Wings, and Outrigger to Tail
Surfaces.]
*Twin Motored Fuselage*. Twin motored aeroplanes generally have the
power plants mounted at a point about midway between the fuselage and
tips of the wings. In almost every case, the power plants are of unit
construction, that is to say, consist of the motor, radiator and
propeller complete on one support, only the fuel and oil tanks being
mounted in the fuselage. The fuselage of the twin may be similar in
length and general construction to that of the tractor biplane, or it
may be a short "nacelle" similar to that used in the pusher type. In any
case, the observer can be located in the extreme front of the body.
An interesting and unusual construction is the body of the Caproni
Biplane (1916). A center nacelle carries the passengers, a pusher screw
being located at the rear of the central body as in the case of the
pusher biplane. On either side of the center are the motors driving the
tractor screws, each motor being encased in a long tractor type fuselage
that also supports the tail surfaces. The latter fuselage serves to
streamline the motors and takes the place of the usual outrigger
construction. There are three bodies, two tractor screws, and one pusher
screw. Somewhat similar in design is the famous German "Billy
Two-Tails," this machine being equipped with two tractor type bodies. A
motor is located in the front of each body. Each fuselage is provided
with accommodations for passengers, and is long enough to support the
tail surfaces. The Caproni and the German machine are both very large
machine and heavily powered.
*U.S.A. Sea-Plane Specifications (1916)*. These government
specifications cover a twin motored sea-plane with a central nacelle.
The body is arranged so that the forward man (observer) can operate the
forward machine gun through a horizontal arc of at least 150°, and
through a vertical arc of at least 270°, with the gun at an angle of
about 75° with the center line of the body. The muzzle must be forward
of the propeller plane. The rear man (pilot) operates a machine gun
through a vertical arc of at least 150° to the rear, and through a
vertical arc of at least 180°, with the gun at an angle of about 105°
with the fuselage center line. The muzzle must be to the rear of the
plane of propeller rotation.
The number of stays and other important connections which extend across
the plane of propeller rotation shall be reduced to a minimum. It is
considered advisable to incorporate in the design of the body such a
structure (in the plane and 8 inches forward of propeller plane) as will
prevent a broken propeller blade from severing the main body. The system
used in the construction of the cage masts used on battleships is
suggested, with a number of spruce compression members in place of
stay-wires. The clearance of the propeller tips from the sides of the
central body shall be from 5 to 12 inches. No part of the gas tanks
shall lie in the plane of propeller rotation, nor within a space 6
inches ahead of this plane.
A space extending at least 9 inches back from the rear of the observer’s
seat, and entirely across the body, must be left open and unoccupied in
order that any desired instruments can be installed therein. In the
center line of the body, a circular hole 9 inches in diameter shall be
cut in the floor of the observer’s cock-pit, the rear of the hole being
5 inches forward of the forward edge of the observer’s seat. The
flooring of the pilot’s and observer’s cockpits shall consist of spruce
strips 1/2" x 1/2 " spaced at 1/2" intervals along the longerons. No
flooring is to be placed under the seats.
The safety factor of the body and tail structure shall not be less than
2.5, the air speed being taken at 100 miles per hour with the elevator
at an angle of 20° and the fixed stabilizer surface at 6°. All wire
tension members not readily accessible for inspection and adjustment are
to be single strand high tensile steel wire. All tension stays that are
easily accessible shall be of non-flexible stranded steel cable. For
turnbuckle safetying No. 20 semi-hard copper wire shall be used. All
cable shall be well stretched before making up the connections. A load
equal to 20 or 30 per cent of the breaking load shall be applied for a
period of from two to three hours. The hard wire must undergo a bending
test by bending at a right angle turn over a radius equal to the
diameter of the wire, back and forth four times each way. No more than
four sizes of turnbuckles shall be used on the entire aeroplane
structure. The strengths and size numbers of the turnbuckles will be as
follows: No. 1 = 8,000 lbs. No. 2 = 4,600 lbs. No. 3 = 2,100 lbs. No. 4
= 1,100 lbs. Controls and fittings in the vicinity of the compasses
shall, as much as possible, be of non-magnetic material. All steel plate
and forged fittings shall be protected against the action of salt water
by baking enamel, the best standard three coat process being used. All
covered wiring and turnbuckles shall be coated by at least two coats of
Flexible Compound.
All steel tubing shall be thoroughly cleaned, slushed with mineral oil
inside, and plugged at both ends by wood plugs impregnated with mineral
oil or paraffine. All steel nuts, bolts, pins and cotter pins shall be
protected by heavy nickel plating over copper. All wood members,
especially faying surfaces, end grain butts, scarfs and joints, shall be
protected against the access of moisture before final assembly by the
best grade of varnish, or by impregnation by paraffine. All wood shall
be straight grained, well seasoned, of uniform weight, and free of
knots, pitch pockets, checks or cracks. Spruce to be of the very highest
grade of selected straight, even grained, clear spruce. It shall be air
seasoned, preferably for two years. Kiln dried wood is not acceptable.
It is highly desirable to have all bolts, pins, plate fittings and
turnbuckle ends made of chrome vanadium steel (S. A. E. Specification
6.130), heat treated to obtain the best physical characteristics. All
parts and fittings that must be bent shall be heat treated after all
bending operations are completed, and by such a sequence of treatment as
will produce the desired grain and toughness, and relieve all stresses
due to the bending. This includes sheet and forged steel fittings,
turnbuckle ends and bolts and pins. All steel parts and fittings
submitted to stress or vibration shall be heat treated in such a manner
as to produce the highest possible refinement of grain and give the
greatest possible resistance to alternating and vibratory stresses.
Where plate fittings are in contact with wooden members, sharp edges
next to the wood shall be removed. In making up and connecting steel
fittings, welding shall be used wherever possible. If impracticable to
weld, and in such cases only, brazing will be used, proper heat
treatment to be employed to restore strength and toughness of metal
after such welding or brazing. Extreme care should be taken to avoid
nicking or kinking any wire, cable or fitting. Fittings, sheet or
forged, must be free from sharp corners and supplied with generous
fillets.
In general the S.A.E. Standards will be acceptable, and these standards
for screw threads shall be used wherever possible. U.S. Standard threads
will be accepted where threaded into cast iron, cast aluminum or copper
alloys. All nuts and pins must be provided with one or more positive and
durable safety devices. In general, where it must be expected that a
structural fitting will be disassembled a number of times during the
life of the aeroplane, castellated nuts with split pins, in accordance
with S.A.E. Standards, shall be used. Wherever this is not the case,
pins or bolts shall be riveted in a workmanlike manner.
Seats shall be securely braced against both horizontal and vertical
stresses. Arrangement and dimensions of cock-pits shall be as nearly as
practicable to that indicated by the drawings (not published in this
chapter). In addition, if practicable, the pilot should be provided with
quick release arm rests. Sections of best grade of khaki on each side of
seats, in which pockets are made, should be fastened to longerons and
vertical posts in such a way as to be securely in place and yet readily
detachable for inspection of structural wiring and fittings. Safety
belts shall be provided for both seats and securely fastened. The belts
shall safely support at any point a load of 2,000 pounds applied as in
practice. Rubber shock absorbers in the safety belt system are
considered to be an advantage. The quick release device shall be as
indicated in drawings and shall reliably and quickly function. Seat pads
shall be quickly detachable in order that they may be used as life
preservers. They will be filled with Kapok or other similar material and
covered with real leather to protect it against the action of salt
water.
Suitable covers shall be provided over the top of the rear end of the
fuselage. These must be easily removed and capable of being securely
fastened in place during flight. Space shall be allowed in the body
directly in the rear of the observer’s seat for the stowage of the sea
anchor. When in use, the sea anchor shall be attached by suitable and
convenient fastening hooks to the two points along the lower longerons,
and at the junction of the two vertical struts in the rear of the front
seat. The structure must be such that it will successfully withstand the
stresses imposed by the sea anchor. Controls shall be of the standard
Deperdussin type, installed in the rear cock-pit only. The tanks for the
main supply of gasoline shall be in the fuselage and located so that the
longitudinal balance will not be disturbed by the emptying of the tank
during flight.
The above data is not in the exact form of the original specifications
and is not complete, but gives only the specifications that affect the
design of the body. These were picked out part by part from the
original.
*Army Specification 1003 (Speed Scout)*. These specifications cover the
design of land machines, the extracts given here referring only to the
safety factor. Body forward of the cockpit shall be designed for safety
factor of 10 over static conditions, with the propeller axis horizontal.
Body in rear of cockpit shall be designed to fail under loads not less
than those imposed under the following conditions:
(a) Dynamic loading of 5 as the result of quick turns in pulling out of
a dive. (b) Superposed on the above dynamic loading shall be the load
which it is possible to impose upon the elevators, computed by the
following formula: L = 0.005AV², where A is the total area of the
stabilizing surface (elevators and fixed surface), and V is the
horizontal high speed of the machine. The units are all in the metric
system. (c) Superposed on this loading shall be the force in the control
cables producing compression in the longerons.
*Fuselage Covering*. Disregarding the monocoque and veneer constructed
types of fuselage, the most common method of covering consists of a
metal shell in the forward end, and a doped linen covering for that
portion of the body that lies to the rear of the rear seat. The metal
sheathing, which may be of sheet steel or sheet aluminum, generally runs
from the extreme front end to the rear of the pilot’s cockpit. Sheet
steel is more common than aluminum because of its stiffness. Military
machines are usually protected in the forward portions of the fuselage
by a thin armor plate of about 3 millimeters in thickness. This is a
protection against rifle bullets and shrapnel fragments, but is of
little avail against the heavier projectiles. Armor is nearly always
omitted on speed scouts because of its weight. Bombers of the
Handley-Page type are very heavily plated and this shell can resist
quite large calibers.
The fabric used on the rear portion of the fuselage is of linen similar
to the wing covering, and like the wing fabric is well doped with some
cellulose compound to resist moisture and to produce shrinkage and
tautness. On the sides and bottom the fabric is supported by very thin,
light stringers attached to the fuselage struts. On the top, the face is
generally curved by supporting a number of closely spaced stringers on
curved wooden formers. The formers are generally arranged so that they
can be easily removed for the inspection of the wire stay connections
and the control leads. On some machines the top of the fuselage consists
entirely of sheet metal supported on formers, while in others the metal
top only extends from the motor to the rear of the rear cockpit.
CHAPTER XIII. CHASSIS CONSTRUCTION.
*General Notes*. The chassis or landing gear carries the weight of the
aeroplane when resting on or running over the ground, and is subjected
to very heavy shocks, especially when landing. It is provided with
pneumatic tired wheels, an elastic shock absorbing device, and the
structural members that connect the axle with the fuselage. In some
forms of landing gear, the wheels are supplemented by long horizontal
skids which serve to support the machine after the elastic shock
absorbers are fully extended or when the wheels collapse. The skids also
protect the aeroplane in cases where the wheels run into a ditch and
also prevent the machine from nosing over in a bad landing. Since the
skids and their structural members cause a high resistance they are now
seldom used except on the larger and slower machines. In running over
the ground, or in making a hard landing, part of the shock is taken up
by the deflection of the tires and part by the deflection of the shock
absorber. The greater the movement of the tires and absorber, the less
will be the stress in the frame.
In the majority of cases, the shock absorbers consist of rubber bands or
cords, these being wound over the axle and under a stationary part of
the chassis members. Since rubber is capable of absorbing and
dissipating a greater amount of energy per pound of weight than steel,
it is the most commonly used material. Rubber causes much less rebound
or "kick" than steel springs. The principal objection to rubber is its
rotting under the influence of sunlight, or when in contact with
lubricating oil. The movement of the axle tube is generally constrained
by a slotted guide or by a short radius rod.
The design of a suitable chassis is quite a complicated problem, for the
stresses are severe, and yet the weight and resistance must be kept at a
minimum. In running over rough or soft ground for the "Get off," the
shocks and vibration must be absorbed without excessive stress in the
framework, and without disturbing the balance or poise of the machine.
There must be little tendency toward nosing over, and the machine must
be balanced about the tread so that side gusts have little tendency in
throwing the machine out of its path. It must be simple and easily
repaired, and the wheels must be large enough to roll easily over
moderately rough ground.
[Illustration: Fig 1. "V" Type Chassis as Applied to "Zens" Monoplane.
Courtesy "Flight."]
*Types of Chassis*. The simplest and most extensively used landing gear
is the "Vee" type shown by Fig. 1, and is equally applicable to
monoplanes, biplanes or triplanes. Primarily, the Vee chassis consists
of two wheels, an axle, a rubber shock absorber, and two sets of Vee
form struts. The chassis shown by Fig. 1-a is that of the
Hansa-Brandenburg and is typical of biplane chassis. The winding of the
rubber cord and the arrangement of the chassis struts are clearly shown.
The two struts are connected at the bottom by a metal fitting, and the
rubber is wound over the axle and under this fitting. No guiding device
is used for the axle, the machine being freely suspended by the chord.
The struts are made as nearly streamline form as possible.
[Illustration: Fig 1a. "V" Type Chassis Used on Hansa-Brandenberg
Biplane.]
Fig. 2 is a front view of a typical Vee chassis, and Fig. 3 is side view
of the same device, the same reference letters being used in each view.
The vertical struts C run from the fuselage at F to the connecting axle
guide plate G. The wheels W-W are connected with the steel tube axle A,
and the struts are braced against side thrust by the cross-tube D and
the stay wire braces B-B. In Fig. 3 the metal fitting G is provided with
the guiding slot S for the axle A. The elastic rubber cord absorber
passes over the axle and is fastened to the plate G by the studs I. Fig.
4 is a side view of the chassis of the Lawson trainer, which like many
other primary training machines, uses a front pilot wheel to guard
against nosing over. The rear two wheels (W) are elastically supported
between the Vee struts C and F, while the front wheel X is attached to
the fuselage by the vertical strut E, and to the rear wheel frame by the
tube G. It will be noted that the front wheel is smaller than the rear
main wheels, as this wheel carries but little load. The tail skid T is
hinged to the fuselage and is provided with elastic cord at the upper
end so that the shock is reduced when the tail strikes the ground. Fig.
5 shown directly above the Lawson trainer, is the complete assembly of
the Hansa-Brandenburg already described. The tail skid of the
Hansa-Brandenburg is indicated by T.
[Illustration: Figs. 2-3. Typical "V" Chassis With Axle Guide.]
The metal shod ash skid stick is hinged to the lower face of the
fuselage, and at the upper end is attached to a stationary fuselage
member through four turns of elastic cord. When the skid strikes an
obstacle the rubber gives and allows the tail to move in relation to the
ground. By this arrangement the greater part of the device is enclosed
within the fuselage and, hence, produces little head resistance.
[Illustration: Fig 4. (Below). Lawson Training Tractor Biplane. Fig. 5
(Above). Hansa-Brandenburg Fighting Biplane Showing Chassis and Tail
Skid (t).]
Fig. 7 is the skid chassis of the Farman biplane which shows clearly the
arrangement of the skids and the shock absorbing suspension. A metal
bridge is attached to the axle, and a series of short rubber bands are
used in connecting the axle bridge, and the bridge on the skid. A
triangular tubular radius rod is attached to the axle and hinged to the
skid. This restrains the travel of the axle in a fore and aft direction.
Another form of skid shock absorber is given by Fig. 8, in which the
rubber rings pass over a spool on the axle. The guiding links or radius
rods on the inside of the skids regulate the axle travel. In general,
the use of a radius rod is not desirable as it transmits a percentage of
the shock to the machine.
[Illustration: Fig. 7. (Left). Farman Skid Type Chassis. Fig. 8. Another
Type of Skid Chassis in Which the Axle Is Guided by a Radius Rod or
Lever.]
[Illustration: Fig. 9. Chassis Details of the Nieuport Monoplane. This
Has a Central Skid and Uses an Automobile Type Steel Spring Instead of
Rubber Cord. Fig. 10 Is a Detail of the Nieuport Spring. (At Right.)]
Fig. 9 is an older form of Nieuport monoplane chassis, a steel cross
spring being used in place of the usual rubber bands. This is simple,
but comparatively heavy, and is subject to frequent spring breakage. To
guard against spring failure, a long ash skid is placed under the axle.
The spring system is connected with the body by three sets of oval steel
struts. An old type of Curtiss chassis is given by Fig. 11. This has
been widely used by amateurs and exhibition flyers, but requires a
fairly smooth landing ground as there are no shock absorbers. The only
shock absorption is that due to the deflection of the tires. The extreme
forward position of the front wheel effectually prevents any tendency
toward nosing over when landing.
[Illustration: Fig. 11. An Old Type of Curtiss Exhibition Chassis With
Three Wheels.]
[Illustration: Fig. 12 (Left). Standard H-3 Shock Absorber. Fig. 13.
(Right). Rubber Cord on Axle.]
A Standard H-3 shock absorbing system is given by Fig. 12. This has a
bracket or hanger attached to the axle over which the elastic cord is
wrapped. The cord is wrapped in continuous turns between the axle hanger
and the bottom of the Vee support members. As shown, the upper
streamlined bar is the axle, while the lower is the cross bar brace
which serves to hold the lower ends of the U’s. I am indebted to "Aerial
Age" for this cut. In order to guide the axle in a straight line in its
up and down movement, two radius links are attached between the axle and
the front vertical strut. One decided advantage of the "Standard
construction" is that the cords are wound without crossing the strands,
thus reducing cutting and wear between the cord turns. Fig. 13 is a
variation of Fig. 12, the cord being wound directly around spools on the
axle and the lower stationary cross tube. The axle is guided by a slot
in the guide plate at the right, while end motion is controlled by a
radius link. Fig. 14 is the double wheel arrangement of a large "Twin"
bombing plane. Two wheels are placed directly under each of the motor
units so that a portion of the load is communicated to the chassis by
tubes. Diagonal tubes transmit the body load to the chassis.
[Illustration: Fig. 14. Chassis for Twin Motored Biplane of Bombing
Type.]
*Folding Chassis*. Owing to the great relative resistance of the chassis
it has been suggested by many designers to provide a folding frame which
will automatically fold up into the body after the machine has left the
ground. This would be a decided advantage but the gear is complicated
and probably not altogether reliable.
*Height of Chassis*. The height of the chassis is made as small as
possible with a sufficient clearance for the propeller tips. It is usual
to have the tips of the propeller blades clear the ground by from 10 to
12 inches when the aeroplane is standing with the body in a horizontal
position. Any smaller clearance is almost certain to result in broken
blades when landing at a sharp angle or when running through high grass.
If the chassis is excessively high the resistance will be high and the
machine is also likely to be top heavy.
[Illustration: Figs. 15-16. Methods of Calculating Wheel Position on Two
Wheel Chassis. This Is an Important Item in the Design of an Aeroplane.]
*Location of Wheels*. The exact location of the wheels, in a fore and
aft direction, is of the greatest importance. If they are too far ahead
of the center of gravity, too much weight will be placed on the tail
skid and excessive running will be required to get the tail off the
ground. If the wheels are too far back, the machine will be likely to
nose over when landing or running over the ground. In any case, the
wheels must be well ahead of the center of gravity so that the weight
will resist a forward overturning moment. In the majority of orthogonal
biplanes, in which the leading edges of the upper and lower wings are on
the same vertical line, the center of the wheel is from 3 to 6 inches
back of the leading edges. In staggered biplanes the wheel center is
from 6 inches to one foot in front of the lower leading edge. This
difference is caused by the fact that the center of gravity is nearer
the leading edge of a staggered wing than with the Orthogonal type, and
hence the wheels must be further forward.
Fig 15 (upper diagram) shows the conditions when the machine is running
over the ground with the body horizontal. The vertical line a-a passing
through the center of gravity C G is a distance N from the center of the
wheel. The weight acting down has a tendency to pull the tail down, this
moment being equal to the weight of the machine multiplied by the
distance N, or W x N. The elevator flap M exerts a lifting force Ky
which acts through the lever arm L, and opposes the moment due to the
weight. The force K must be equal to K = WN/L. The distance I is the
distance of the wheel center line from the entering edge of the wing.
The weight on the tail skid S when the machine is resting on the ground
will be equal to S = WN/M, and this may range anywhere from 40 to 200
pounds, according to the size of the aeroplane.
Fig. 16 illustrates a principle of wheel location advanced by Capt.
Byron Q. Jones, and published in "Aviation and Aeronautical Engineers,"
Nov. 16, 1916. The body is shown in a horizontal position with the
propeller axis X-X horizontal. The center of gravity is at G on X-X, the
weight acting down as at P with the line prolonged meeting the ground
line at B. A line E-E is a line drawn tangent to the wheels and the tail
skid at D, the angle of this line with the ground determining the
maximum angle of incidence. E-E is the ground line when the machine is
at rest. For the best conditions, Capt. Jones finds that the line
connecting the point of tangency C, and the center of gravity at G,
should make an angle of 13 degrees and 10 minutes with the vertical GB
dropped through the center of gravity. With the line GA drawn
perpendicular to the resting line E-E, the angle BGA should be 10
degrees as nearly as possible. This is for a two-wheel Vee chassis, but
with a third front wheel as with the training of type the angle CGB can
be made less. Capt. Jones has found that with the wheels in the above
location there will be no tendency to nose over even with very poor
landings, and this method has been applied to the training machines at
the San Diego Signal Corps aviation school. If the angle BGA is greater
than 10 degrees it is difficult to "taxi" the machine on the ground,
this tending to make the machine spin or turn into the wind. Capt. Jones
claims that a two-wheel chassis arranged according to these rules is
superior to the three-wheel type for training purposes since the
tendency toward spinning is less.
The location of the tail skid S should be such that the elevator and
rudder surfaces are well off the ground with the skid fully deflected,
and yet the skids must be low enough to permit of the maximum angle of
incidence or an angle of EXX = 10 degrees. To a certain extent, the
maximum angle of incidence determines the chassis height. If the angle
EXX is made greater than the greatest angle of incidence, the wings can
be used as air brakes in bringing the machine to a quick stop after
landing.
The track, or the distance between the centers of the wheels measured
along the axle, must be about 1/7 or 0.15 of the span of the lower wing.
This makes the track vary from 5 to 7 feet on the usual types, and as
high as 15 feet on the large bombing planes. The track must be great
enough to prevent overturning when making a landing on soft ground or
with a cross wind. If the track is excessive, there will be a heavy
spinning moment in cases where one wheel strikes a depression or soft
spot in the ground.
*Shock Absorbers*. The axle movement allowed by the elastic shock
absorbers and guiding appliances averages from 5 to 6 inches. The
greater the movement, the less will be the stresses induced by a given
drop, but in practice the movement is generally limited by
considerations of chassis height and propeller clearance. It can be
proved that a movement of 5 inches will produce a maximum stress equal
to 8.6 times the weight of the machine under conditions of a one-foot
drop, while with an absorber movement of 6 inches the stress is reduced
to 7.5 times the weight. This calculation takes the tire deflection into
consideration. With the absorber movement limited to one inch, the
stress may be as high as 35 times the weight of the machine.
F=W (2 + 2.77/x) where W = weight of machine in pounds, F = the stress
produced by the fall, and x = the absorber movement in inches.
*Landing Gear Wheels*. The wheels are generally of the tangent laced
wire spoke type, and are enclosed with discs to reduce the resistance.
They must have very wide hubs to resist the heavy end stresses caused by
landing sidewise. The length of the hub should be at least twice the
diameter of the tire and a greater width, say three times the tire
diameter, is preferable. The narrow hubs used on motorcycle wheels are
not safe against side blows, although they may be capable of
withstanding the vertical load. The wheels are rated according to the
outside diameter over the tire, and by the diameter of the tire casing.
A 26" x 4" wheel signifies that the outside diameter is 26 inches with a
casing diameter of 4 inches.
The 26 x 4 tires are used on the majority of training machines of the
two-wheel type, while a 20 x 4 wheel is used for the front wheel of the
three-wheel trainer. Two larger sizes, 30 x 4 and 34 x 4, have also been
used to some extent, particularly on the Ackerman spring wheels.
CHAPTER XIV. ESTIMATION OF WEIGHT.
*Effect of Weight*. Weight is an all important consideration and is most
difficult to estimate unless one has accurate data on existing machines
of the same type. The total weight in flying order depends upon the
useful load to be carried, and upon the weight of the power plant. The
weight of the latter varies both with the useful load and with the
speed, climb, and duration of flight. The type of aeroplane determines
the relative head resistance which again reflects back to the weight of
the power plant.
The only reason for the existence of an aeroplane is to carry a certain
useful load for a given distance, and this useful load is the basis of
our weight calculations. The basic useful load consists of the
passengers and cargo, although in some specifications the live load may
be construed as including the weight of the fuel, oil and instruments,
and in the case of military aeroplanes, the weight of the armament,
armor, ammunition, wireless and cameras. For comparison, the elements
constituting the live load should always be specified.
For a given horsepower, speed and climb, it is obvious that the dead or
structural weight should be at a minimum for a maximum live load
capacity. The dead load carried in present aeroplanes will be
undoubtedly reduced in the future by the adoption of lighter and
stronger materials, better methods of bracing, and by reductions in the
weight of the power plant. Just as the automobile industry developed
light and powerful materials of construction, so will the aeroplane
designer develop more suitable materials for the aircraft. While the
present power plant has been refined to a remarkable extent when
compared with the older types, it is still far from the lowest possible
limit. At present the complete power unit—the motor, radiator,
propeller, water, etc.—will weigh from 2 to 5 pounds per horsepower.
With a given aeroplane, the performance is determined by the total
weight and power. The duration and flight range can be increased by
increasing the fuel weight at the expense of the passenger or cargo
weight. The power available for climbing is the excess of the total
power of the motor over the power required for horizontal flight. Since
the power for horizontal flight depends principally upon the weight, it
is at once evident that the weight is a regulating factor in the
climbing speed. In fact the climbing speed may be almost directly
determined from the weight carried per horsepower at normal flight
speed. A fast climbing scout may weigh from 8 to 12 pounds per
horsepower, while the large low climbing machine will weigh from 16 to
20 pounds per horsepower, the respective climbing speeds being
approximately 1,200 and 350 feet per minute.
*Fuel Efficiency and Weight*. The efficiency of the motor, or its fuel
consumption for a given output, has a very marked effect upon the total
weight of the aeroplane. Under certain conditions a very light motor
with a high fuel consumption will often contribute more to the total
weight than a heavier but more economical motor. In short flights, up to
3 hours, the very light rotating cylinder motor with its high fuel
consumption probably gives the least total weight, but for longer
flights the more efficient and heavier water-cooled type is preferable.
For flights of over three hours the fuel weight is a considerable
percentage of the total weight. The proper motor for any machine must be
selected by computing the weight of the fuel and oil required for a
given duration and then adding this to the total weight of the engine
and its cooling system.
*Distribution of Weight*. Practically the only way to predict the weight
of a proposed machine is to compare it with a similar existing type.
After the ratio of the useful load to the total load has been
determined, the useful load of the proposed machine can be divided by
the ratio factor to obtain the total weight. It should be noted in this
regard that if the proposed machine is much larger than the nearest
existing example, a liberal allowance must be made to compensate for the
increase in the proportional weight of the structural members. There
have been many mathematical formulas advanced for predicting the weight,
but these are very inaccurate in the majority of cases.
As a rough estimate, based on a number of successful machines, the
weight of the actual aeroplane structure without power-plant, live load,
fuel, oil, or tanks, is very nearly 32 per cent (0.32) of the total
weight. The remaining 68 per cent is divided up among the power-plant,
fuel and live load. Thus, the aeroplane structure proper of a machine
weighing 2000 pounds total will be 2000 x 0.32 = 640 pounds. Taking the
weight of the power plant, tanks and piping at 28 per cent, the total
dead load of the bare machine without fuel or oil will be 60 per cent of
the total. With a training aeroplane built for a 6-hour flight, the fuel
and oil will approximate 16 per cent, so that the total percentage
possible for the crew and cargo will be 24 per cent. With a given live
load, the total load can now be calculated by dividing the live load by
its percentage. Using the above value, for example, the total weight in
order of flight with a live load of 720 pounds becomes: W = 720/0.24 =
3000 pounds.
In government specifications the total weight of the pilot and passenger
are taken at 330 pounds, or 165 pounds per man. Gasoline and oil are for
a 4-hour flight. A safer average figure will be 170 pounds per man, and
a fuel allowance of 6 hours. The floats of a seaplane or flying boat
bring the percentage of the dead load much higher than with the land
type of chassis.
The following table will give an idea as to the weight distribution
expressed both in pounds, and as a percentage of the total weight. It
covers a wide range of types, varying from the training types Curtiss
JN-4B and the Standard H-3, to the Handley-Page Giant bomber and the
Nieuport speed scout. The average values found by Hunsaker for a number
of machines weighing in the neighborhood of 2500 pounds is given in the
fourth column. Under each heading are the actual weights and the
percentages of the total weight for each item. Items marked (*) include
both gasoline and oil. Mark (C) is the power plant complete, and (@)
includes radiator.
*Weight Per Horsepower*. As already explained, the weight carried per
horsepower varies with the type of machine. When the total weight is
determined for any aeroplane, the power requirements can be calculated
by dividing the total weight by the weight per horsepower ratio. A fair
value for a training or exhibition machine is from 18 to 20 pounds per
horsepower, while for a very high speed machine, such as a chaser, the
weight will be taken at 10 pounds per horsepower. For two-seater
fighters 16 to 18 pounds is fair practice. For a comparison of the
horsepower-weight ratios used on different well-known machines see
tables in Chapter II. Thus, if our total weight is found to be 2400
pounds as determined from the above table, and if this is a training
machine, the horsepower will be: 2400/20 = 120 horsepower. Using the
same total weight, but powered for two-seater fighter conditions, the
power will be increased to 2400/16= 150 horsepower. As a scout the power
will be increased still further to 2400/10=240 horsepower.
As a problem in solving the weight and horsepower from the data, we will
assume that we are to design a two-seater fighter with a total useful
load of 1200 pounds. This load consists of the following items:
Personnel (2) = 330 pounds; gas and oil = 500 pounds; guns and
ammunition = 370 pounds. The nearest example that we have to this live
load is that of the Standard H-3, which carries 744 pounds and in which
the percentage of live load is 28.1 per cent. As our machine will be
somewhat larger, we will not be far from the truth if we take the
percentage as 0.27 instead of 0.281. The total weight, in flying order,
will now be 1200/0.27 = 4440 pounds. At 16 pounds per horsepower the
motor will be: 4440/16=277 horsepower.
An empirical formula for a high-speed scout was set forth in "Aviation
and Aeronautical Engineering" by D. W. Douglas. This is based on the
horsepower unit. A unit wing loading of 8.45 pounds per square foot, and
a low speed of 55 miles per hour was assumed. The wing section chosen
was the U.S.A.-1. In the formula, H = horsepower:
- Power plant weight = 3 H.
- Chassis weight = 0.7 H.
- Tail weight = 0.25 H.
- Fuel for 2.25 hours = 1.4 H.
- Military load = 250 pounds.
- Tanks and piping = 0.42 H.
- Fuselage weight = 1.84 H.
- Wing weight = 1 lb. sq. ft.
- Propeller = 2.8/H.
- (Total) = (7.61 H + 2.5/H + 250)/7.45 = Weight of aeroplane fully
loaded in the order of flight.
*Weight of Wings*. The weight of the wings depends upon the span, very
small machines having wings that weigh only 0.38 pounds per square foot,
while the wings of very large machines may run as high as 1.1 pounds per
square foot. For average size biplanes from 0.75 to 0.80 pounds per
square foot would probably be safe—that is, for areas ranging from 450
to 550 square feet. The weight of the upper wing of the Nieuport is
0.815 pounds per square foot, while the lower wing (short chord) is
0.646 pounds per square foot. The wings of the Standard H-3 trainer will
average 0.77 pounds per square foot, the lower wing and center section
being heavier than the upper wing. The wings of the Curtiss JN-4B will
average 0.75 pounds per square foot. These weights do not include the
interplane wires or struts, nor the fittings. The total weight of the
interplane struts of the JN-4B, the Aviatic, and machines of similar
size will average from 28 to 30 pounds. The ailerons will weigh about 12
pounds each.
*Weight of Motors*. There is a considerable difference in the weight of
air-cooled and water-cooled motors. The water, water piping, radiators
and jackets of the water-cooled motors adds considerably to the weight
of the complete power plant. The mountings are heavier for the
water-cooled motors, and because of the tandem arrangement of the
cylinders, the crankshaft and crankcase weigh more. In taking the bare
weight of the power plant all of the accessories must be included. In
the following table, the "bare engine" includes the carbureter, magneto,
and necessary integral accessories, but does not include the jacket
water, mounting, radiator, oil in base, water piping, nor controls.
Water-cooled motors are marked by (W) and air-cooled by (A). Rotary
air-cooled are (RA), and gallons (G).
WEIGHTS OF AERONAUTICAL MOTORS.
The bare radiator will weigh from 0.48 to 0.56 pounds per horsepower,
the average being safe at 0.52. The water contained in the radiator will
average 0.35 pounds per horsepower. The weights of the piping and the
water contained therein will be computed separately. The circular sheet
metal cowl used over the rotary cylinder air-cooled motor is equal to
twice the square root of the motor weight, according to Barnwell.
Propeller weight varies considerably with the diameter, pitch, etc., but
a safe rule will give the weight as 2.8 √H where H = horsepower. The
tanks will weigh from 0.75 to 1.2 pounds per gallon of contents, or
approximately 1/5 the weight of the contents when completely filled.
*Chassis and Wheel Weight*. The chassis of a two-wheel trainer will
weigh about 90 pounds complete, although there are chassis of training
machines that weigh as much as 140 pounds. The chassis of speed scouts
will be from 22 to 40 pounds complete. Tail skids can be taken at from 6
to 8 pounds.
Tangent wire wheels complete with tires are about as follows: 26 x 4 =
21 pounds; 26 x 5 = 28 pounds; 26 x 3 = 14 pounds. Ackerman spring spoke
wheels are estimated as follows: 20 x 4 = 17.5 pounds; 26 x 3 = 22
pounds; 26 x 4 = 32 pounds; 30 x 4 = 35 pounds; 34 x 4 = 45 pounds.
*Military Loads*. A 20-mile wireless outfit devised by Capt. Culver
weighed 40 pounds with storage batteries, while the 120-mile outfit
weighed 60 pounds with a 180-watt generator. The 140-mile U.S.A.
mule-back wireless of 1912 weighs 45 pounds. The "Blimp" specifications
allow 250 pounds.
The Lewis gun as mounted on the "11" Nieuport weighs 110 pounds,
including mount, gun and ammunition. Lewis gun bare is 26 pounds. The
Davis 6-pounder, Mark IV, weighs 103 pounds with mounting but without
ammunition, while the same make of 3-inch 12-pounder weighs 238 pounds
under the same conditions.
*Controls and Instruments*. The Deperdussin type controls used on the
Curtiss JN-4B weigh 16 pounds per control, while those installed in the
Standard H-3 weigh about 13 pounds. An average of 15 pounds per control
is safe. An instrument board for the aviators’ cock-pit, fully equipped,
weighs from 20 to 24 pounds. The front, or students’ instrument board
will average 10 pounds. Pyrene extinguisher and brackets = 7 pounds;
Speaking tube = 3 pounds; Oil pressure line and gage = 3 pounds; Side
pockets = 3 pounds; Tool kit = 10 pounds.
*Control Surfaces*. The rudder, stabilizer, fin, and elevator can be
made so that the weight will not exceed 0.60 to 0.65 pounds per square
foot.
*General Notes on Weight*. Before starting on the weight estimates of
the machine the reader should carefully examine the tables in Chapter II
which give the weights, and general characteristics of a number of
modern machines.
*Weights and Wing Area*. When the weight of the machine is once
determined, the next step will be to determine the wing area. For speed
scouts or very large heavy duty machines the choice of a wing section
must be very carefully considered. For the speed scout several wings
giving a minimum high speed resistance should be examined, such as the
Eiffel 37 or the U. S. A-1 or U.S.A.-6. For the low-speed aeroplane to
be designed for great lift, a number of sections such as the U.S.A.-4 or
the R.A.F.-3 should be tried for a number of speeds and angles. For
training machines a wing of the "All around" type such as the R.A.F.-6
should be adopted, the structural characteristics in the case of a
trainer having an important bearing on the subject. If W = weight of the
machine in pounds, V = low speed in miles per hour, A = total area in
square feet, and Ky=lift coefficient, then the area becomes A=W/KyV².
Compensation must be made for biplane interference for aspect ratio, and
stagger as previously explained. For an ordinary training machine with
the usual gap/chord ratio, and aspect ratio, the correction factor of
0.85 may be safely employed.
*Example*. We will take the case of an aeroplane carrying a personnel
load of 340 pounds, oil and gasoline 370 pounds, and baggage amounting
to 190 pounds, instruments 100 pounds. Total live load will be 1000
pounds. Taking the live load percentage as 0.30, the total load will be
1000/0.30 = 3333 pounds. If the low speed is 50 miles per hour, and the
maximum Ky of the chosen wing is 0.003 at this speed, the area will be A
= W/KyV² = 3333/0.003 x (50 x 50) = 444 square feet. Since this is a
biplane with a correction factor of 0.85, the corrected area will be:
444/0.85 = 523 square feet. The unit loading, or weight per square foot
will be: 3333/523 = 6.36 pounds. The corrected area includes the
ailerons and the part of the lower wing occupied by the body.
*Empirical Formula for Loading*. After investigating a large number of
practical biplanes, the author has developed an expression for
determining the approximate unit loading. When this is found, the
approximate area can be found by dividing the total weight by the unit
loading. This gives an idea as to the area used in practice.
It was found that the unit loading increased with the velocity at nearly
a uniform rate. This gave an average straight line formula that agreed
very closely with 128 examples. If V = Maximum velocity in miles per
hour, and w = weight per square foot, then the unit loading becomes:
w = 0.065V - 0.25 for the average case. For high speed scouts this gives
a result that is a trifle low, the formula for a fast machine being more
nearly w = 0.65V - 0.15, for speeds over 100 miles per hour.
A two-seat machine of average size weighs 2500 pounds, and has a maximum
speed of 90 miles per hour. Find the approximate unit loading and area.
The loading becomes: w = 0.065V - 0.25 = (0.065 x 90) - 0.25 = 5.6
pounds per square foot. The approximate area will be: 2500/5.60 = 446
square feet.
If the above machine had a speed of 110 miles per hour, the formula
would be changed for the high-speed type machine, and the loading would
become:
w = 0.065V - 0.15 = (0.065 x 110) - 0.15 = 7.00 pounds per square foot.
The required area will be: 2500/7.0 = 372 square feet. When the unit
load is also determined in this way it is a very simple matter to choose
the wing section from Ky = w/V².
*Area From Live Load and Speed*. By a combination of empirical formula
we can approximate the area directly. For the average size machine, w =
0.065V - 0.25. And the total weight W = U/0.32 where U is the useful or
live load. Since A = W/w, then A = U/(0.65V - 0.25) x 0.32 = U/0.021V -
0.08.
Thus if an aeroplane travels at 90 miles per hour and has carried a
useful load of 800 pounds (including gas and oil), the approximate area
is: A = U/0.021.V = 0.08 = 800/(0.021 x 90) - 0.08 = 442 square feet.
This assumes that the useful load is 0.32 of the total load and that the
speed is less than 100 miles per hour.
CHAPTER XV. BALANCE AND STABILITY.
*Elements of Stability*. When we balance a board on a fulcrum so that it
stands in a perfectly horizontal position, the board is said to be "In
equilibrium," or is supported at its "Center of gravity." There is only
one point at which a body will balance, and this point is at the center
of gravity or "C. G." In an aeroplane, the combined mass of the body,
motor, wings, fuel, chassis, tail and live load has a center of gravity
or a balancing point at which the lift must be applied if the machine is
to rest in equilibrium. When the center of lift (or center of pressure)
does not pass through the center of gravity of the aeroplane, some other
force must be applied to overcome the unbalanced condition. When the
machine is unbalanced in a fore and aft direction with the tail low, a
force must be applied by the elevator flaps that is opposite and equal
to the moment of the unbalanced forces. An aeroplane is stable when it
is balanced in such a way that it returns to a state of equilibrium
after meeting with a disturbance.
When disturbed, a stable body does not usually return instantly to its
position of equilibrium, but reaches it after a series of decreasing
oscillations. The heavier the body, and the more compact its form, the
longer will it oscillate about its fulcrum before coming to rest. By
arranging broad surfaces at the ends of the oscillating body, a portion
of the energy will be expended in creating air currents, and the motion
will be readily "damped out." If the damping effect is so great that the
body does not swing back after once reaching the position of
equilibrium, the body is said to be "dead beat," or "dynamically
stable." There is a great difference between the static forces that tend
to return the body to a position of equilibrium and the dynamic
retarding forces that tend to damp out the oscillations. Usually, a body
with excessive static stability is far from being stable in a true
sense, since such a body tends to oscillate longer, and more violently,
than one in which the static restoring forces are not so strongly
marked. A body may be statically but not dynamically stable, but a
dynamically stable body must of necessity be statically stable.
Static stability in calm air is determined by the location of the center
of gravity, the center of lift, the center of propeller thrust, the
center of area of the surfaces, and the center of the forward
resistance. The forces acting through these centers are: (1) The weight;
(2) The lifting force; (3) The propeller thrust; (4) The resistance. The
weight and lift are vertical forces equal and opposite in direction. The
thrust and resistance are horizontal forces, also equal and opposite in
direction. When all of these forces intersect at a common point, they
will completely neutralize one another and the body will be in
equilibrium.
Dynamic stability is attained by the use of large damping surfaces such
as the stabilizer surface, fins, and the elevator. These act to kill the
oscillations set up by the static righting couples or forces. Without
suitable damping surfaces the machine would soon be out of control in
gusty weather since successive wind gusts will act to increase the
oscillations of the righting forces until the machine will turn
completely over. On the other hand, an aeroplane can be too stable and
therefore difficult to steer or control in gusts because of its tendency
toward changing its attitude with every gust in order to restore its
equilibrium. A machine should only be partially stable, and the majority
of pilots are firmly set against any form of mechanical or inherent
control. No matter how simple the method, mechanical control always
introduces a certain amount of mechanism that may go wrong. The question
of stability has already been solved to a sufficient extent.
A disturbance that simply changes the direction of travel is not
considered an unstable force since it normally does not tend to endanger
the machine. Nearly any machine, equipped with any possible form of
control apparatus, tends to change its direction when being righted.
*Axes of Stability*. An aeroplane has six degrees of freedom or motion.
Three are of translation or straight line motion, and three are of
rotation about rectangular axes. It can travel forward in a straight
line, rise and fall in a vertical plane, or skid sidewise. When it rolls
from side to side about the fore and aft axis (X axis) it is laterally
unstable. When pitching up and down in a fore and aft direction, and
around an axis parallel with the length of the wings (Y axis), the
machine is said to be longitudinally unstable. When swinging or "Yawing"
from right to left about a vertical axis (Z axis) it is unstable in
"Yaw."
Rolling is resisted by the ailerons, pitching by the elevators and
stabilizer, and yawing by the vertical directional rudder. Lateral
oscillation are damped out by the wing surfaces and by vertical surfaces
or "Fins." Longitudinal oscillations are damped mostly by the stabilizer
and elevator surfaces. Directional or yawing vibrations are corrected by
the damping action of the vertical tail fin, vertical rudder and the
sides of the body, the latter also serving to damp out longitudinal
vibrations. On an absolutely calm day, the pilot can shut off the motor
and glide down without touching the controls if the machine is
longitudinally stable. The glide generally starts with a few pitching
oscillations, but these gradually are damped out by the tail as soon as
the machine picks up its natural gliding angle and speed, and from this
point it will continue without oscillating.
*The Spiral and Nose Dive*. There are two forms of instability that have
not yet been fully corrected, and both are highly dangerous. One of
these is known as the "spiral dive" or nose spin, and the other as the
straight nose dive. The aeroplane in a spiral nose dive rotates rapidly
about a vertical axis during the dive. Spiral instability resulting from
lateral instability, can be minimized by decreasing the area of the
vertical rudder and by the proper placing of fins so that there is not
so great an excess of vertical area to the rear of the C. G.
The covered-in body acts as a fin and will be productive of spiral
instability if the area is not properly distributed. In the majority of
cases the rear of the body is equivalent to a large fin placed to the
rear of the C. G. A fin above the G. G. tends to reduce all spiralling.
*Stability and Speed*. An aeroplane in straight horizontal flight must
be driven at such an angle, and such a speed, that the weight is just
sustained. To be inherently stable the machine must always tend to
increase its speed by diving should the power be cut off in any way. An
aeroplane that does not tend to increase its speed in this way, "Stalls"
or becomes out of control. Any machine that will automatically pick up
its gliding angle after the propeller thrust has ceased is at least
partially inherently stable, and if it does not possess this degree of
stability, other forms of stability are practically worthless. The
machine having the smallest, flattest gliding angle is naturally safest
in cases of power failure, and hence the gliding angle is somewhat
related to the subject of stability.
[Illustration: A Spanish Aeroplane Using a Peculiar Form of Upper Fin.
These Fins Also Perform the Duty of Vertical Rudders as Well as Acting
as Stabilizers.]
The longitudinal stability decreases with a decrease in the speed, the
fore and aft vibrations becoming more rapid due to the decreased effect
of the tail surfaces, and to the reduction of wing lift. Instability at
low speeds is common to all aeroplanes, whether inherently stable or
not, and at a certain critical speed the machine becomes absolutely
unstable in a dynamic sense. If a machine is to be stable at low speeds,
it must not fly at too great an angle of incidence at these speeds, and
it should have a very large tail surface acting at a considerable
distance from the wings. Hunsaker states that the lowest speed should
not require more than 80 per cent of the total lift possible.
*Inertia or Flywheel Effect*. The principal weights should be
concentrated as nearly as possible at the center of gravity. Weights
placed at extreme outer positions, as at the wing tips, or far ahead of
the wings, tend to maintain oscillations by virtue of their flywheel
effect. The measure of this inertia or flywheelage is known as the
"Moment of Inertia" and is the sum of the products of all the masses by
the squares of their distances from the center of gravity. A great
amount of inertia must be met by a large damping surface or control area
if the vibrations are to be damped out in a given time. In twin-motored
aeroplanes the motors should be kept as close to the body as the
propellers will permit.
*Wind Gusts and Speed*. A machine flying at high speed is less affected
by wind gusts or variations in density than a slow machine, since the
disturbing currents are a smaller percentage of the total speed. In
addition, a high speed results in smaller stresses due to the gusts.
*Gyroscopic Instability*. The motor gyroscopic forces do not affect the
stability of a machine to any great extent, and in twin motored
aeroplanes the gyroscopic action of the propellers is almost entirely
neutralized. At one time the gyroscopic torque was blamed for every form
of instability, but on investigation it was found that the practical
effect was negligible.
*Instability Due Power Plant*. The power plant affects stability in a
number of ways. The thrust of the propeller may cause a fore and aft
moment if the center line of thrust does not pass through the center of
resistance. This causes the machine to be held head up, or head down,
according to whether the line of thrust is below or above the C. G. If
the propeller thrust tends to hold the head up in normal flight, the
machine will tend to dive, and assume its normal gliding velocity with
the power off, hence this is a condition of stability. With the effect
of the thrust neutral, or with the thrust passing through the center of
resistance, the machine will not tend to maintain the speed, and hence
it is likely to stall unless immediately corrected by the pilot. With
the line of thrust above the C. G., the stall effect is still further
increased since with this arrangement there is a very decided tendency
for the machine to nose up and increase the angle of incidence when the
power is cut off.
[Illustration: Steel Elevator and Rudder Construction Used on a European
Machine. The Elevators Also Act as Stabilizers, the Entire Surface
Turning About the Tube Spar.]
The slip stream of the propeller has a very decided effect on the tail
surfaces, these being much more effective when the propeller slip stream
passes over them. With lifting tails, or tails that normally carry a
part of the load, the stoppage of the slip stream decreases the lift of
the tail and consequently tends to stall the machine. Non-lifting tails
should be arranged so that the slip stream strikes down on the upper
surface. This tends to force the tail down, and the head up in normal
flight, and when the power ceases the tail will be relieved and there
will be an automatic tendency toward diving and increase in speed. On a
twin aeroplane, a similar effect is obtained by making the upper tips of
both propellers turn inwardly. The air is thus thrown down on the tail.
With a single motor, the torque tends to turn the aeroplane in a
direction opposite to the rotation of the propeller. Lateral stability
is thus interfered with when the motor is cut off or reduced in speed.
With right-hand propeller rotation, for example, the machine will be
turned toward the left, forcing the left tip down. To maintain a
horizontal attitude, the left aileron must be held down by an amount
just sufficient to overcome the torque. In some machines one wing tip is
given a permanent increase in incidence so that the down seeking tip is
given permanent additional lift.
*Lateral Stability*. When an aeroplane is turned sharply in a horizontal
plane, or "Yaws," the outer and faster moving wing tip receives the
greater lift, and a lateral rolling moment is produced about the fore
and aft axis. In the opposite condition, a lateral rolling moment tends
to yaw or to throw the aeroplane off a straight course. Below a certain
critical speed, the lateral or rolling oscillations increase in
amplitude, with a strong tendency to side slip, skid or spiral. The tail
fin or rudder retards the tail velocity in a side slip, and thus turns
the slipping or skidding machine into a vertical spiral or spinning nose
dive. This spin increases the angle of bank and hence the side slip.
This in turn increases the turning or yawing velocity, and the spiral
starts. This tendency toward a spiral dive can be corrected by a
vertical fin placed forward, and above the center of gravity, or by
raising the wing tips. An upper fin of this type will give a force that
tends to break up the bank when side slip starts and thus will prevent
spinning.
[Illustration: Sperry Gyroscopic Control System for Automatic Stability.
The Gyroscopic Control at the Left Controls the Movements of the
Electric Servo-Motor at the Extreme Right. This Motor Operates the
Control Surfaces Through the Pulley Shown. A Small Electric Generator
Between the Servo-Motor and Gyroscope Provides the Current and Is Driven
by a Small Wind Propeller.]
At normal speeds the rolling is damped down by the wing surfaces, and
can be further controlled by the application of the ailerons. At the
lower critical speed when the machine is stalled, one wing tip has no
more lift than the other, and hence the damping effect of the wings and
the action of the ailerons becomes negligible.
*Dutch Roll*. In "Dutch Roll," the rolling is accompanied by an
alternate yawing from right to left. This is aggravated by a fin placed
high above the C. G., and hence corrections for spiral dive conflict
with corrections for Dutch roll. The rolling is accompanied by some side
slip, and the motion is stable providing that there is sufficient fin in
the rear and not an excessive amount above the C. G.
*Degree of Stability*. Excessive stability is dangerous unless the
control surfaces are powerful enough to overcome the stable tendency.
Since a stable machine always seeks to face the relative wind, it
becomes difficult to handle in gusty weather, as it is continually
changing its course to meet periodic disturbances. This is aggravated by
a high degree of static stability, and may be positively dangerous when
landing in windy weather.
*Control Surfaces*. A non-lifting tail must give no lift when at a zero
angle of incidence. It must be symmetrical in section so that equal
values of lift are given by equal positive and negative angles of
incidence. Square edged, flat surfaces are not desirable because of
their great resistance. A double cambered surface is suitable for such
controls as the stabilizer, elevator and rudder. It has a low
resistance, permits of strong internal spars, and is symmetrical about
the line of the chord. Some tails are provided with a cambered top and a
flat bottom surface so that the down wash of the wings is neutralized.
Under ordinary conditions this would be an unsymmetrical lifting
surface, but when properly adapted to the wings the lifting effect is
completely neutralized by the down wash.
The curvature of the section should be such that the movement of the
center of pressure is as small as possible. With a small movement of the
center of pressure, the surface can be accurately balanced and hinged on
the center of pressure line. It is desirable to have the maximum
thickness of section at, or near to the C. P., so that a deep spar can
be used for the support of the hinge system. Usually the movement of the
control surfaces is limited to an angle of 30 degrees on either side of
the center line, as the lift of all surfaces start to decrease after
this point is reached. The surface movement should be limited by the
maximum lift angle of the section in any case, since an accident will be
bound to occur if they are allowed movement beyond the angle of maximum
lift.
In locating the control surfaces, careful attention should be paid to
the surrounding air conditions so that they will not be unduly affected
by the wash-down of the wings or body. The effectiveness of the tail
surfaces is very much reduced by bringing them close to the wings, and
the lift is always reduced by the wash of a covered fuselage.
The wash-down effect of the wings on the tail is proportional to the
chord and not to the span, and for this reason an increase in span does
not always necessitate an increase in the length of the body. An
adequate damping effect requires a large surface at the end of a long
lever arm.
*Balancing the Aeroplane*. Figs. 1 to 6 show the principles involved in
the balancing of the aeroplane. In Fig. 1 a number of weights 1’-2’-3’
and 5M are supported on a beam, the load being balanced on the fulcrum
point M. The load 2’ being directly over the fulcrum, has no influence
on the balance, but load 1’ at the left tends to turn the mass in a
left-hand direction, while 3’ and 5M tend to give it a right-hand
rotation. This turning tendency depends upon the weights of the bodies
and their distance from the fulcrum. The turning tendency or "Moment" is
measured by the product of the weight and the distance from the fulcrum.
If weight 1’ should be 10 pounds, and its distance A’ from the fulcrum
should be 20 inches, then it would cause a left-hand moment of 10 x 20 =
200 inch pounds. If the system is to be in balance, then the left-hand
moment of 1’ should be equal to the sum of the moments of 3’ and 5M.
Thus: 1’ x A = (3’ x B) + (5M x C’).
[Illustration: Figs. 1-6. Methods of Balancing an Aeroplane About Center
of Lift.]
The application of this principle as applied to a monoplane is shown by
Fig. 4, in which X-X is the center of pressure or lift. The center of
lift corresponds to the fulcrum in Fig. 1, and the weights of the
aeroplane masses and their distance from the center of lift are shown by
the same letter as in Fig. 1. The engine 1’ is at the right of the C. P.
by the distance A, while the fuel tank 2 is placed on the C. P. in the
same way that the weight 2’ in Fig. 1 is placed directly over the
fulcrum. By placing the tank in this position, the balance is not
affected by the emptying of the fuel since it exerts no moment. The
chassis G acting through the distance E is in the same direction as the
engine load. The body 5 with its center of gravity at M acts through the
distance C, while the weight of the pilot 3 exerts a right-hand moment
with the lever arm length B. If the moments of all these weights are not
in equilibrium, an additional force must be exerted by the tail V.
Fig. 2 shows an additional weight 4’ that corresponds to the weight of
the passenger 4 in Fig. 5. This tends to increase the right turning
moment unless the fulcrum is moved toward the new load. In Fig. 2 the
fulcrum M remains at the same point as in Fig. 1, hence the system
requires a new force P’ acting up at the end of the beam. If the load
was in equilibrium before the addition of 4’, then the force P’ must be
such that P’ x T’ = 4" x D’. In the equivalent Fig. 5, the center of
gravity has moved from its former position at S to the new position at
R, the extent of the motion being indicated by U. To hold this in
equilibrium, an upward force P must be exerted by the elevator at Y, the
lever arm being equal to (T + U).
Fig. 6 shows the single-seater, but under a new condition, the center of
pressure having moved back from X-X to Z. To hold the aeroplane in
equilibrium, a downward force must be provided by the tail V which will
cause a right-hand moment equal to the product of the entire weight and
the distance U. For every shift in the center of pressure, there must be
a corresponding moment provided by the elevator surface. The condition
is shown by the simple loaded beam of Fig. 3. In this case the fulcrum
has been moved from M to N, a distance equal to the center of pressure
movement in Fig. 6. This requires a downward force P’ to maintain
equilibrium.
*Center of Pressure Calculation*. Fig. 7 is a diagram showing the method
of calculating the center of gravity. The reference line R is shown
below the elevators and is drawn parallel to the center of pressure line
W-W, the latter line being assumed to pass through the center of
gravity. The line R may be located at any convenient point, as at the
propeller flange or elsewhere, but for clearness in illustration it is
located to the rear of the aeroplane. The weight of each item is
multiplied by the distance of its center of gravity from the line R,
these products are added, and the sum is then divided by the total
weight of the machine. The result of this division gives the distance of
the center of gravity from the line R. Thus, if the center of gravity of
the body (11) is located at (10), then the product of the body weight
multiplied by the distance B will give the moment of the body about the
line R. The weight of the motor (2) multiplied by the distance F gives
the moment of the motor about R, and so on through the list of items.
[Illustration: Center Of Gravity Table]
The distance of the center of gravity (or center of pressure) from the
reference line R is given by H + K. This gives the numerical value
219350/1375 = 1596 inches. Thus if we measure 159.6 inches from R toward
the wings we will have located the center of gravity. The location of
the C. G. can be changed by shifting the weights of the motor,
passenger, or other easily moved items. In any case, the C. G. should
lie near the center of pressure.
*Tail Lever Arms*. The effective damping moment exerted by the fixed
stabilizer surface (12) will be the product of its area by the distance
(I), measured from the center of pressure of the wing to the center of
pressure of the stabilizer. The lever arm of the elevator is the
distance (H) measured from the centers of pressure as before.
[Illustration: Fig. 7. Method of Determining the Center of Gravity of an
Aeroplane.]
*Resultant Forces and Moments in Flight*. The aeroplane is in
equilibrium when all of the forces pass through a common center, as
shown by Fig. 8. In this figure the lift (L), the weight (W), the line
of propeller thrust (T), and the resistance (R) all pass through the
center of gravity shown by the black dot C. G. There are no moments and
hence no correction is needed from the elevator (T). In Fig. 9, the
thrust and resistance pass through the center of gravity as before, but
the center of lift (L) does not pass through the center of gravity, the
distance between the two being indicated by (n). This causes a moment,
the length of the lever arm (n) being effective in giving a right-hand
rotation to the body. If horizontal flight is to be had this must be
resisted by the upward elevator force (E).
In Fig. 10, the lift passes through the center of gravity, but the line
of resistance lies below it by the amount (m). The thrust (T) tends to
rotate the machine in a left-handed direction. The elevator must exert a
downward force (e) to resist the moment caused by (m). This is a bad
disposition of forces, as the machine would tend to stall or tail-dive
should the propeller thrust cease for even an instant. The stability of
Figs. 8 and 9 would not be affected by the propeller thrust, as it
passes through the C. G. in both cases. In Fig. 11, the center line of
thrust is below the line of resistance (R), so that the thrust tends to
hold the nose up. Should the motor fail in this case, the nose would
drop and the machine would start on its gliding angle and pick up speed.
In Fig. 12 none of the forces intersect at a common point, the lift and
weight forming a right-handed couple, while the thrust (T) and the
resistance (R) form a left-handed couple that opposes the couple set up
by the weight and lift forces. If the thrust-resistance couple can be
made equal to the lift-weight couple, the aeroplane will be in
equilibrium and will need no assistance from the elevator. As the
weights in the aeroplane are all located at different heights, it is
necessary to obtain the center of gravity of all the loads in a vertical
plane as well as horizontally. Thus in Fig. 13 the line C. G. is the
center of gravity of the engine weight (1), the wing weight (2), the
pilot’s weight (3), the chassis weight (4), the fuselage weight (5), and
the fuel tank weight (6). The line C. G. is the effective center of all
these loads, and is calculated by taking the products of the weights by
the distance from a reference line such as R-R. The center of resistance
is the effective center of all the resistance producing items such as
the wings, body, struts, chassis, etc.
[Illustration: Figs. 8-15. Forces Affecting the Longitudinal Stability
of an Aeroplane.]
A suggestion of the method employed in obtaining the center of
resistance is shown by Fig. 14, the center line of resistance R-R being
the resultant of the wing resistance (D), the body resistance (B), and
the chassis resistance (C). It will be noted that the wing resistance of
biplane wings (W-W’) does not lay midway between the wings but rather
closer to the upper wing, as shown by (E). This is due to the upper wing
performing the greater part of the lift. In locating the center of
resistance, the resistance forces are treated exactly like the weights
in the C. G. determination. Each force is multiplied by its distance
from a horizontal reference line, and the sum of the products is divided
by the total resistance. As shown, the center of resistance R-R passes
through the center of gravity C. G. The center of pressure line X-X also
contains the center of resistance.
A staggered biplane cell is shown by Fig. 15, the center of pressure of
the upper and lower wings being connected by the line X-X as before. The
center of resistance of the pair is shown at (D), where it is closer to
the upper wing than to the lower. A vertical line Y-Y dropped through
the center of resistance gives the location of the center of lift. As
shown, the center of lift is brought forward by the stagger until it is
a distance (g) in front of the leading edge of the lower wing. The
center of lift and the center of resistance both lie on a line
connecting the center of pressure of the upper and lower wings.
*Calculation of Control Surfaces*. It is almost impossible to give a
hard and fast rule for the calculation of the control surfaces. The area
of the ailerons and tail surfaces depends upon the degree of stability
of the main wings, upon the moment of inertia of the complete machine,
and upon the turning moments. If the wings are swept back or set with a
stagger-decalage arrangement, they will require less tail than an
orthogonal cell. All of these quantities have to be worked out
differently for every individual case.
*Aileron Calculations*. The ailerons may be used only on the upper wing
(2 ailerons), or they may be used on both the upper and lower wings.
When only two are used on the upper wing it is usually the practice to
have considerable overhang. When the wings are of equal length either
two or four ailerons may be used. Roughly, the ailerons are about
one-quarter of the wing span in length. With a long span, a given
aileron area will be more effective because of its greater lever arm.
If a = area of ailerons, and A = total wing area in square feet, with S
= wing span in feet, the aileron area becomes: a = 3.2A/S. It should be
borne in mind that this applies only to an aeroplane having two ailerons
on the upper wing, since a four-aileron type usually has about 50 per
cent more aileron area for the same wing area and wing span. For,
example, let the wing span be 40 feet and the area of the wings be 440
square feet, then the aileron area will be: a = 3.2A/S = 3.2 x 440/40 =
35.2 square feet. If four ailerons were employed, two on the upper and
two on the lower wing, the area would be increased to 1.5 x 35.2 = 52.8
square feet. As an example in the sizes of ailerons, the following table
will be of interest:
[Illustration: Aileron Sizes Table]
In cases where the upper and lower spans are not equal, take the average
span—that is, one-half the sum of the two spans.
*Stabilizer and Elevator Calculations*. These surfaces should properly
be calculated from the values of the upsetting couples and moments of
inertia, but a rough rule can be given that will approximate the area.
If a’ = combined area of stabilizer and elevator in square feet; L =
distance from C. P. of wings to the C. P. of tail surface; A = Area of
wings in square feet, and C = chord of wings in feet, then:
a’ = 0.51AC/L. Assuming our area as 430 square feet, the chord as 5.7
feet, and the lever arm as 20 feet, then:
a’ = 0.51AC/L = 0.51 x 430 x 5.7/20 = 62.5 square feet, the combined
area of the elevators and stabilizer. The relation between the elevator
and stabilizer areas is not a fixed quantity, but machines having a
stabilizer about 20 per cent greater than the elevator give good
results. In the example just given, the elevator area will be: 62.5/22 =
28.41 square feet, where 2.2 is the constant obtained from the ratio of
sizes. The area of the stabilizer is obtained from: 28.41 x 1.2 = 34.1
square feet.
*Negative Stabilizers*. A considerable amount of inherent longitudinal
stability is obtained by placing the stabilizing surface at a slight
negative angle with the wings. This angle generally varies from -2° to
-6°. At small angles of wing incidence the negative angle of the tail
will be at a maximum, and acting down will oppose further diving and
tend to head the machine up. At large wing angles, the tail will be
depressed so far that the tail angle will become positive instead of
negative, and thus the lift on the tail will oppose the wings and will
force the machine to a smaller angle of incidence. The negative angle
can thus be adjusted to give longitudinal stability within the ordinary
range of flight angles.
*Stabilizer Shapes and Aspect Ratio*. Stabilizers have been built in a
great number of different shapes, semicircular, triangular, elliptical,
and of rectangular wing form. Measured at the rear hinged joint, the
span or width of the stabilizer is about 1/3 the wing span for speed
scouts, and about 1/4 the wing span for the larger machines. Nearly all
modern machines have non-lifting tails, or tails so modified that they
are nearly non-lifting. Since flat plates give the greatest lift with a
small aspect ratio, and hence are most effective when running over the
ground at low speeds, the stabilizers and elevators are of comparatively
low aspect. In general, an aspect ratio of 3 is a good value for the
stabilizer. Vertical rudders generally have an aspect ratio of 1, and
hence are even more effective per unit area than the stabilizers. This
is particularly necessary in ground running.
[Illustration: Aileron Control Diagram of Curtiss JN4-B.]
[Illustration: Elevator Control Diagram of Curtiss JN4-B.]
*Vertical Rudders*. The calculation of the vertical rudders must take
the moment of inertia and yawing moments into effect, and this is rather
a complicated calculation for the beginner. As an approximation, the
area of the rudder can be taken from 9 to 12 square feet for machines of
about 40 feet span, and from 5 to 8 square feet for speed scouts.
[Illustration: Stick Control Used on the Caudron Biplane. Wing Warp Is
Used Instead of Ailerons. Back and Forth Movement Actuates Elevator.]
[Illustration: German Stick Control With Double Grips. A. Latch on the
Side of the Stick Acts on a Sector So That the Lever Can Be Held at Any
Point. It Is Released by the Pressure of the Knees.]
*Wing Stability*. Under wing sections, the subject of the center of
pressure movement has already been dealt with. The variation of the
center of pressure with the angle of incidence tends to destroy
longitudinal stability since the center of pressure does not at all
times pass through the center of gravity. On some wings, the camber is
such that the variation in the position of the center of pressure is
very little, and hence these are known as stable wings. A reflex curve
in the trailing edge of a wing reduces the center of pressure movement,
and swept back wings are also used as an aid in securing longitudinal
stability. Introducing stagger and decalage into a biplane pair can be
made to produce almost perfect static longitudinal stability. It should
be noted that stability obtained by wing and camber arrangements is
static only, and requires damping surfaces to obtain dynamic stability.
[Illustration: Form of Control Used on the Nieuport Monoplane.]
*Manual Controls*. In flight, the aviator has three control surfaces to
operate, the ailerons, elevator, and rudder. In the usual form of
machine the ailerons and elevator are operated by a single lever or
control column, while the rudder is connected with a foot bar. In the
smaller machines "Stick Control" is generally used, the ailerons and
elevator being moved through a simple lever or "Joy Stick" which is
pivoted at its lower end to the floor. The Deperdussin or "Dep" control
is standard with the larger machines and consists of an inverted "U"
form yoke on which is mounted the wheel for operating the ailerons.
*Stick Control*. With the stick pivoted at the bottom, a forward
movement of the lever causes the machine to descend while a backward
movement or pull toward the pilot causes the aeroplane to head up or
ascend. The stick is connected with the elevators with crossed wires, so
that the flaps move in an opposite direction to the "Stick." Moving the
stick from side to side operates the ailerons.
[Illustration: Standard Stick Control and Movements Used in the U.S.A.]
*Deperdussin Control*. A "U" shaped yoke, either of bent wood or steel
tube, is pivoted the bearers at the sides of the fuselage. Wires are
attached to the bottom of the yoke so that its back and forth movement
is communicated to the elevator flaps. On the top, and in the center of
the yoke, is pivoted a hand wheel of the automobile steering type. This
is provided with a pulley and is connected with the aileron flaps in
such a way that turning the wheel toward the high wing tip causes it to
descend. Pushing the yoke forward and away from the aviator causes the
machine to descend, while a reverse movement raises the nose. The "Dep"
control is reliable and powerful but is bulky and heavy, and requires a
wide body in order to allow room for the pilot.
[Illustration: Foot Rudder Bar Used in the Standard H-3. Courtesy
"Aerial Age."]
*Rudder Control*. Foot bar control for the rudder is standard with both
the stick and Dep controls. The foot bar is connected with the rudder in
such a way that the aeroplane turns opposite to the movement of the foot
bar in the manner of a boat. That is, pushing the right end of the bar
forward causes the machine to turn toward the right.
[Illustration: Automatic Control System (Sperry) Installed in Fuselage
of Curtiss Tractor Biplane.]
CHAPTER XVI. HEAD RESISTANCE CALCULATIONS.
*Effect of Resistance*. Resistance to the forward motion of an aeroplane
can be divided into two classes, (1) The resistance or drag due to the
lift of the wings, and (2) The useless or "Parasitic" resistance due to
the body, chassis and other structural parts of the machine. The total
resistance is the sum of the wing drag and the parasitic resistance.
Since every pound of resistance calls for a definite amount of power, it
is of the greatest importance to reduce this loss to the lowest possible
amount. The adoption of an efficient wing section means little if there
is a high resistance body and a tangle of useless struts and wires
exposed to the air stream. The resistance has a much greater effect on
the power than the weight.
*Weight and Resistance*. We have seen that the average modern wing
section will lift about 16 times the value of the horizontal drag, that
is, an addition of 16 pounds will be equal to 1 pound of head
resistance. If, by unnecessary resistance, we should increase the drag
by 10 pounds, we might as well gain the benefit of 10 x 16 = 160 pounds
of useful load. The higher the lift-drag efficiency of the wing, the
greater will be the proportional loss by parasitic resistance.
*Gliding Angle*. The gliding angle, or the inclination of the path of
descent when the machine is operating without power, is determined by
the weight and the total head resistance. With a constant weight the
angle is greatest when the resistance is highest. Aside from
considerations of power, the gliding angle is of the greatest importance
from the standpoint of safety. The less the resistance, and the flatter
the angle of descent, the greater the landing radius.
Numerically this angle can be expressed by: Glide = W/R, where W = the
weight of the aeroplane, and R = total resistance. Thus if the weight is
2500 pounds and the head resistance is 500 pounds, the rate of glide
will be: 2500/500 = 5. This means that the machine will travel forward 5
feet for every foot that it falls vertically. If the resistance could be
decreased to 100 pounds, the rate of glide would be extended to 2500/100
= 25, or the aeroplane would travel 25 feet horizontally for every foot
of descent. This will give an idea as to the value of low resistance.
*Resistance and Speed*. The parasitic resistance of a body in uniform
air varies as the square of the velocity at ordinary flight speeds.
Comparing speeds of 40 and 100 miles per hour, the ratio will be as 40°
is to 100° = 1600: 10,000 = 6.25, that is, the resistance at 100 miles
per hour will be 6.25 times as great as at 40 miles per hour.
The above remarks apply only to bodies making constant angle with the
air stream. Wings and lifting surfaces make varying angles at different
speeds and hence do not show the same rate of increase. In carrying a
constant load, the angle of the aeroplane wing is decreased as the speed
increases and up to a certain point the resistance actually decreases
with an increase in the speed. The wing resistance is greatest at
extremely low speeds and at very high speeds. As the total resistance is
made up of the sum of the wing and parasitic resistance at the different
speeds, it does not vary according to any fixed law. The only true
knowledge of the conditions existing through the range of flight speeds
is obtained by drawing a curve in which the sums of the drag and head
resistance are taken at intervals.
*Resistance and Power*. The power consumed in overcoming parasitic
resistance increases at a higher rate than the resistance, or as the
cube of the speed. Thus if the speed is increased from 40 to 100 miles
per hour, the power will be increased 15.63 times. This can be shown by
the following: Let V = velocity in miles per hour, H = Horsepower, K=
Resistance coefficient of a body, A = Total area of presentation, and R
= resistance in pounds. Then H = RV/375. Since R = KAV², then H = KAV² x
V/375 = KAV³/375.
*Resistance and Altitude*. The resistance decreases with a reduction in
the density of the air at constant speed. In practice, the resistance of
an aeroplane is not in direct proportion to a decrease in the density as
the speed must be increased at high altitudes in order to obtain the
lift. The following example given by Capt. Green will show the actual
relations.
Taking an altitude of 10,000 feet above sea level where the density is
0.74 of that at sea level, the resistance at equal speeds will be
practically in proportion to the densities. In order to gain
sustentation at the higher altitude, the speed must be increased, and
hence the true resistance will be far from that calculated by the
relative densities. Assume a sea level speed of 100 ft./sec., a weight
of 3000 pounds, a lift-drag ratio of L/D = 15, and a body resistance of
40 pounds at sea level.
Because of the change in density at 10,000 feet, the flying speed will
be increased from 100 feet per second to 350 feet per second in order to
obtain sustentation. With sea level density this increase in speed (3.5
times) would increase the body resistance 3.5 x 3.5 = 12.25 times,
making the total resistance 12.25 x 40 = 490 pounds. Since the density
at the higher altitude is only 0.74 of that at sea level, this will be
reduced by 0.26, or 0.26 x 490 = 364 pounds. Thus, the final practical
result is that the sea level resistance of the body (40 pounds) is
increased 9.1 times because of the speed increase necessary for
sustentation. Since the wing angle and hence the liftdrag ratio would
remain constant under both conditions, the wing drag would be constant
at both altitudes, or 3000/15 = 200 pounds. The total sea level
resistance at 100 feet per second is 200 + 40 = 240 pounds, while the
total resistance at 10,000 feet becomes 364 + 200 = 564 pounds.
The speed varies as the square root of the change in density percentage.
If V = velocity at sea level, v = velocity at a higher level, and d =
percentage of the sea level density at the higher altitude, then v =
V/√D. When the velocity at the high altitude is thus determined, the
resistance can be easily obtained by the method given in Capt. Green’s
article. The following table gives the percentage of densities referred
to sea level density.
───────────────────────────────────────────────────────────────────
Altitude Feet Density Percent Altitude Feet Density Percent
───────────────────────────────────────────────────────────────────
Sea-level 1.00 7,500 0.78
───────────────────────────────────────────────────────────────────
1,000 .97 10,000 .74
───────────────────────────────────────────────────────────────────
2,000 .95 12,500 .66
───────────────────────────────────────────────────────────────────
3,000 .91 15,000 .61
───────────────────────────────────────────────────────────────────
5,000 .85 20,000 .52
───────────────────────────────────────────────────────────────────
If the velocity at sea level is 100 miles per hour, the velocity at
20,000 feet will be 100/0.72 = 139 miles per hour, where 0.72 is the
square root of the density percentage, or the square root of 0.52 = .72
at 20,000 feet.
*Total Parasitic Resistance*. Aside from the drag of the wings, the
resistance of the structural parts, body, tail and chassis depends upon
the size and type of aeroplane. A speed scout has less resistance than a
larger machine because of the small amount of exposed bracing, although
the relative resistance of the body is much greater. The type of engine
also has a great influence on the parasitic resistance. The following
gives the approximate distribution of a modern fighting aeroplane:
─────────────────────────────────────────────────────────────────
Body 62 percent
─────────────────────────────────────────────────────────────────
Landing gear 16 "
─────────────────────────────────────────────────────────────────
Tail, fin, rudder 7 "
─────────────────────────────────────────────────────────────────
Struts, wires, etc. 15 "
─────────────────────────────────────────────────────────────────
The body resistance is by far the greatest item. A great part of the
body resistance can be attributed to the motor cooling system, since in
either case it is diverted from the true streamline form in order to
accommodate the radiator, or the rotary motor cowl. The body resistance
is also influenced by the necessity of accommodating a given cargo or
passenger-carrying capacity, and by the distance of the tail surfaces
from the wings. A body is not a streamline form when its length greatly
exceeds 6 diameters.
*Calculation of Total Resistance*. The nearest approach that we can make
to the actual head resistance by means of a formula is to adopt an
expression in the form of R = KV² where K is a factor depending upon the
size and type of machine. The true method would be to go over the planes
and sum up the individual resistance of all the exposed parts. The parts
lying in the propeller slip stream should be increased by the increased
velocity of the slip stream. The parasitic resistance of biplanes
weighing about 1800 pounds will average about, R = 0.036V² where V =
velocity in miles per hour. Biplanes averaging 2500 pounds give R =
0.048V². Machines of the training or 2-seater type weigh from 1800 to
2500 pounds, and have an average head resistance distribution as
follows:
─────────────────────────────────────────────────────────────────
Body, radiators, shields 35.5 percent.
─────────────────────────────────────────────────────────────────
Tail surface and bracing 14.9 "
─────────────────────────────────────────────────────────────────
Landing gear 17.2 "
─────────────────────────────────────────────────────────────────
Interplane struts, wires and fittings 23.6 "
─────────────────────────────────────────────────────────────────
Ailerons, aileron bracing, etc. 8.8 "
─────────────────────────────────────────────────────────────────
The averages in the above table differ greatly from the values given for
the high speed fighting machine, principally because of the large
control surfaces used in training machines, and the difference in the
size of the motors.
With the wing drag being equal to D = Kx AV², and the total parasitic
resistance equal to R = KV², the total resistance can be expressed by Rt
= KxAV² + KV², where K = coefficient of parasitic resistance for
different types and sizes of machines. The value of K for training
machines will average 0.036, for machines weighing about 2500 pounds K =
0.048. Scouts and small machines will be safe at K = 0.028. The wing
drag coefficient Kx varies with the angle of incidence and hence with
the speed. For example, we will assume that the wing drag (Kx) of a
scout biplane at 100 miles per hour is 0.00015, that the area is 200
square feet, and that the parasitic resistance coefficient is K = 0.028.
The total resistance becomes: R = (0.00015 x 200 x 100 x 100) + 0.028 x
100 x 100 = 300 + 280 = 580 pounds. The formula in this case would be R
= KxAV² + 0.028V².
*Strut Resistance*. The struts are of as nearly streamline form as
possible. In practice the resistance must be compromised with strength,
and for this reason the struts having the least resistance are not
always applicable to the practical aeroplane. From the best results
published by the N. P. L. the resistance was about 12.8 pounds per 100
feet strut at 60 miles per hour. The width of the strut is 1 inch. A
rectangular strut under the same conditions gave a resistance of 104.4
pounds per 100 feet. A safe value would be 25 pounds per 100 feet at 60
miles per hour. If a wider strut is used, the resistance must be
increased in proportion. With a greater speed, the resistance must be
increased in proportion to the squares of the velocity. When the struts
are inclined with the wind, the resistance is much decreased, and this
is one advantage of a heavy stagger in a biplane.
The "Fineness ratio" or the ratio of the width to the depth of the
section has a great effect on the resistance. With the depth equal to
twice the width measured across the stream, a certain strut section gave
a resistance of 24.8 pounds per 100 feet, while with a ratio of 3.5 the
resistance was reduced to 11.4 pounds per hundred feet. Beyond this
ratio the change is not as great, for with a ratio of 4.6 the resistance
only dropped to 11.2 pounds.
*Radiator Resistance*. For the exact calculation of the radiator
resistance it is first necessary to know the motor power and the fuel
consumption since the radiator area, and hence the resistance, depends
upon the size of the motor and the amount of heat transmitted to the
jacket water. An aeronautic motor may be considered to lose as much
through the water jackets as is developed in useful power, so that on
this basis we should allow about 1.6 square feet of radiation surface
per horsepower. This figure is arrived at by J. C. Hunsaker and assumes
that the wind speed is 50 miles per hour (73 feet per second). The most
severe cooling condition is met with in climbing at low speed, and it is
here assumed that 50 miles per hour will represent the lowest speed that
would be maintained for any length of time with the motor full out. For
a racing aeroplane that will not climb for any length of time, one-half
of the surface given above will be sufficient, and if the radiator is
placed in the propeller slip stream it can be made relatively still
smaller as the increased propeller slip at rapid rates of climb
partially offsets the additional heating.
In the above calculations, Hunsaker does not take any particular type of
radiator into consideration, merely assuming a smooth cooling surface.
The Rome–Turney Company states that they allow 1.08 square feet of
cooling surface per horsepower for honeycomb radiators, and 0.85 square
feet for the helical tube type. The surface referred to means the actual
surface measured all over the tubes and cells, and does not refer to the
front area nor the exterior dimensions of the radiator. While a radiator
may be made 25 percent smaller when placed in the slipstream, the
resistance is increased by about 25 per cent, with a very small saving
in weight, hence the total saving is small, if any. Side mounted
radiators have a lower cooling effect per square foot than those placed
in any other position, owing to the fact that the air must pass through
a greater length of tube than where the broad side faces the wind.
In the radiator section tested by Hunsaker, there were about 64 square
feet of cooling surface per square foot of front face area, but for
absolute assurance on this point one should determine the ratio for the
particular type of radiator that is to be used. The Auto Radiator
Manufacturing Corporation, makers of the "Flexo" copper core radiators,
have published some field tests made under practical conditions and for
different types and methods of mounting. The four classes of radiators
described are: (1) Front Type, in which the radiator is mounted in the
end of the fuselage; (2) Side Type, mounted on the sides of the body;
(3) Overhead Type, mounted above the fuselage and near the top plane;
(4) Over-Engine Type, placed above and connected directly to the motor,
as in the Standard H-3.
The following table gives the effectiveness of the different mountings
in terms of the frontal area required per horsepower and the cooling
surface, the area being in square inches (Front face area of radiator).
Area in wind of type (3) is half the calculated frontal area since one
core lies behind the other: Taking the value of the Rome–Turney
honeycomb radiator as 60 square feet of cooling surface per horsepower,
the frontal area per horsepower will be 0.0169 square feet, assuming
that the radiator is approximately 6 inches thick. This amounts to 2.43
square inches of frontal area per horsepower.
*Example*. Find the approximate frontal area of a Rome-Turney type
honeycomb radiator used with a motor giving 100 brake-horsepower. Find
Resistance at 50 miles per hour (73 feet per second).
─────────────────────────────────────────────────────────────────
Class of Mounting Square Inches Cooling Surface Per H.P
Per H.P. Square Inch
─────────────────────────────────────────────────────────────────
Front Type 4.00 117.00
─────────────────────────────────────────────────────────────────
Side Type 7.20 104.00
─────────────────────────────────────────────────────────────────
Overhead Type 2.70 112.00
─────────────────────────────────────────────────────────────────
Over-Engine Type 5.00 121.00
─────────────────────────────────────────────────────────────────
*Solution*. Area = A = 0.0169 HP = 0.0169 x 100 = 1.69 square feet. The
honeycomb portion of surface for a square radiator of the above area
will measure 16.2" x 16.2". Allowing a 1-inch water passage or frame all
around the core, the side of the completed square radiator will measure
16.2" + 1" = 18.2". The diameter of a circular radiator core of the same
1.69 x 144 area will be 17.4 inches, since D = 1.69x144/0.7854. Adding
the water passage, the overall diameter becomes 17.4 + 1 + 1 = 19.4
inches. The round honeycomb front radiator used on the 100 horsepower
Curtiss Baby Scout measures 20 inches. Hunaker found the resistance of a
honeycomb radiator to be R=0.000814 AV², there being 4 honeycomb cells
per square inch. A = area of radiator in square feet, and V = velocity
in feet per second. Adopting, for example, a speed of 73 feet per
second, and an area equivalent to a 19.4-inch diameter circular radiator
as above, the total resistance becomes:
R = 0.000814 AV² = 0.000814 x 3.1x (73 x 73) = 13.32 pounds, at 50 M. P.
H. where A = 3.1 square feet.
*Resistance of Chassis*. Disc wheels (Enclosed spokes) have a resistance
of about one-half that of open-wire wheels. The N. P. L. and Eiffel have
agreed that the resistance of a wheel approximating 26" x 4" has a
resistance of 1.7 pounds at 60 miles per hour (Disc type). For any other
speed, the wheel resistance will be R = 1.7 V²/3600, where V = speed in
miles per hour. We must also take into consideration the axle, chassis
struts, wiring, shock absorbers, etc. The itemization of the chassis
resistance, as given by the N.P.L. for the B.E.-2 biplane is as follows
(60 miles per hour):
─────────────────────────────────────────────────────────────────
Wheels 2(a)1.75 pounds 3.5 pounds
─────────────────────────────────────────────────────────────────
Axle 2.0 "
─────────────────────────────────────────────────────────────────
Chassis struts and connections 1.1 "
─────────────────────────────────────────────────────────────────
Total chassis resistance(3)60 MPH. 6.6 pounds
─────────────────────────────────────────────────────────────────
At any other speed, the resistance for the complete chassis can be given
by the formula R = 6.6V²/3600. This allowance will be ample, as the B.
E.-2 is an old type and is equipped with skids.
*Interplane Resistance*. The interplane struts and wires are difficult
to estimate by an approximate formula, the only exact way being to
figure up each item separately from a preliminary drawing. The
resistance varies with the form of the strut or wire section, the
length, and the thickness. The fact that some of the struts lie in the
propeller slipstream, and some outside of it, makes the calculation
doubly difficult. The only recourse that we have at present is to
analyze the conditions on the B. E.-2. With struts approximating true
streamline form, a great percentage of the total resistance is skin
friction, and as before explained, this item varies at a lesser rate
than the square of the speed.
According to a number of experiments on full size biplanes averaging
1900 pounds, it has been found that the interplane resistance (Struts,
wires and fittings) amounts to about 24 per cent of the total parasitic
head resistance of the entire machine, the drag of wings not being
included. The maximum observed gave 29 per cent and the minimum 15 per
cent. The resistance of the interplane bracing of speed scouts will be
considerably less in proportion, as there are fewer exposed struts and
cables on this type, the resistance probably averaging 15 per cent of
the total head resistance. Based on these figures the resistance of the
interplane bracing can be expressed by the following formula, in which I
= resistance of interplane bracing in pounds, and V = translational
speed in miles per hour:
I = 0.009 V² (For two-place biplanes weighing 1900 pounds).
I = 0.0054V² (For biplane speed scouts or racing type biplanes).
*Strut Resistance*. The above estimate includes wiring, strut fittings,
etc., complete, and also takes the effect of the slipstream into
consideration. A more accurate estimate can be made on the basis of
strut length. To obtain this unit value we have recourse to the B. E.-2
tests. The translational speed in 60 miles per hour (88 feet per second)
and the slipstream is taken at 25 feet per second. This gives a total
velocity in the slipstream of 113 feet per second. The struts are 1%
inches wide, and vary in length from 3’ 0" to 6’ 0". In the slipstream
the increased velocity increases the resistance of the items by 64 per
cent.
Total running length = 110’–0". Total resistance = 10.81 pounds. The
resistance per foot = 10.81/110 = 0.099 pounds.
*Resistance of Wire and Cable*. In this estimate we will take the
resistance given in the B. E.-2 tests, since values are given in the
slipstream as well as for the outer portions. In the translational
stream there is 240’ 0" of cable, 70’ 0" of No. 12 solid wire, and 52
turnbuckles, the total giving a resistance of 38.10 pounds. In the
slipstream there is 50’ 0" of cable and 30’ 0" of solid wire with a
resistance of 11.00 pounds. The total wire and cable resistance for the
wings is therefore 49.10 pounds. The resistance of the wire and cable
combined is 0.127 pounds per running foot.
*Summary of Interplane Resistance*. The total interplane resistance
includes the struts, wires, cables and turnbuckles, a portion of which
are in the slipstream. Since the total head resistance of the entire
machine (B.E.-2) is 140 pounds at 60 M. P. H., and the interplane
resistance = 10.81 + 49.10= 59.91 pounds, the relation of the interplane
resistance to the total resistance is 43 per cent. This is much higher
than the average (24 per cent), but the B.E.-2 is an old type of machine
and the number of struts and wires were much greater than with modern
aeroplanes.
*Control Surface Resistance*. The resistance of the control surfaces is
a variable quantity, since so much depends upon the arrangement and
form. Another variation occurring among machines of the same make and
type is due to the various angles of the surfaces during flight, or at
least during the time that they are used in correcting the attitude of
the machine. With the elevator flaps or ailerons depressed to their
fullest extent, the drag is many times that with the surfaces in
"neutral," and as a general thing the controls are depressed at the time
when the power demand is the greatest—that is, on landing, flying slow,
or in "getting off."
Ailerons "in neutral" can be considered as being an integral part of the
wings when they are hinged to the wing spar. In the older types of
Curtiss machines the ailerons were hinged midway between the planes and
the resistance was always in existence, whether the ailerons were in
neutral or not. Wing warping, in general, can be assumed as in the case
where the wings and ailerons are combined. With ailerons built into the
wings, the resistance of the ailerons, and their wires and fittings, can
be taken as being about 4 per cent of the total head resistance. With
the aileron located between the two wings, the resistance may run as
high as 20 per cent of the total.
Like the ailerons, the elevator surfaces and rudder are variable in
attitude and therefore give a varying resistance. In neutral attitude
the complete tail, consisting of the rudder, stabilizer, elevator, fin
and bracing, will average about 15 per cent of the total resistance, it
being understood that a non-lifting stabilizer is fitted. With lifting
tails the resistance will be increased in proportion to the load carried
by the stabilizer. In regard to the tail resistance it should be noted
that these surfaces are in the slipstream and are calculated
accordingly, although the velocity of the slipstream is somewhat reduced
at the point where it encounters the tail surfaces. The total tail
resistance of the B. E.-2 is given as 3.3 pounds.
*Resistance of Seaplane Floats*. The usual type of seaplane with double
floats may be considered as having about 12 per cent higher resistance
than a similar land machine. Some forms of floats have less resistance
than others, owing to their better streamline form, but the above figure
will be on the safe side for the average pontoon. Basing our formula on
a 12 per cent increase on the total head resistance, the formula for the
floats and bracing will become: Rt = 0.00436V² where R1 = resistance of
floats and fittings.
*Body Resistance*. This item is probably the most difficult of any to
compute, owing to the great variety of forms, the difference in the
engine mounting, and the disposition of the fittings and connections.
The resistance of the pilot’s and passenger’s heads, wind shields, and
propeller arrangement all tend to increase the difficulty of obtaining a
correct value. Aeroplanes with rotary air-cooled motors, or with large
front radiators have a higher resistance than those arranged with other
types of motors or radiator arrangements. Probably the item having the
greatest influence on the resistance of the fuselage is the ratio of the
length to the depth, or the "fineness ratio." In tractor monoplanes and
biplanes, of the single propeller type, the body is in the slipstream,
and compensation must be made for this factor.
If it were not for the motor and radiator, the tractor fuselage could be
made in true dirigible streamline form, and would therefore present less
resistance than the present forms of "practical" bodies. The necessity
of placing the tail surfaces at a fixed distance from the wings also
involves the use of a body that is longer in proportion than a true
streamline form, and this factor alone introduces an excessive head
resistance. The ideal ratio of depth to length would seem to range from
1 to 5.5 or 1 to 6. The fineness ratio of the average two-seat tractor
is considerably greater than this, ranging from 1 to 7.5 or 8.5. A
single-seat machine of the speed-scout type can be made much shorter and
has more nearly the ideal proportions.
The only possible way of disposing of this problem is to compare the
results of wind tunnel tests made on different types of bodies, and even
with this data at hand a liberal allowance should be made because of the
influence of the connections and other accessories. Eiffel, the N. P.
L., and the Massachusetts Institute of Technology have made a number of
experiments with scale models of existing aeroplane bodies. It is from
these tests that we must estimate our body resistance, hence a table of
the results is attached, the approximate outlines being shown by the
figures.
As in calculating the resistance of other parts, the resistance of the
body can be expressed by R = KxAV², where Kx = coefficient of the body
form, A = Cross-sectional area of body in square feet (Area of
presentation), and V = velocity in miles per hour. The area A is
obtained by multiplying the body depth by the width. The "area of
presentation" of a body 2’ 6" wide and 3’ 0" deep will be 2.5 x 3 = 7.5
square feet.
The experimental data does not give a very ready comparison between the
different types, as the bodies not only vary in shape and size, but are
also shown with different equipment. Some have tail planes and some have
not; two are shown with the heads of the pilot and passenger projecting
above the fuselage, while the remainder have either a simple cock-pit
opening or are entirely closed. The presence of the propeller in two
cases may have a great deal to do with raising the value of the
experimental results. The propeller was stationary during the tests, but
it was noted that the resistance was considerably less when the
propeller was allowed to run as a windmill, driving the motor. This
latter condition would correspond to the resistance in gliding with the
motor cut off. In all cases, except the Deperdussin, the bodies are
covered with fabric, and the sagging of the cloth in flight will
probably result in higher resistance than would be indicated by the
solid wood or metal model used in the tests. The pusher type bodies give
less resistance than the tractors, but the additional resistance of the
outriggers and tail bracing will probably bring the total far above the
tractor body.
In the accompanying body chart are shown 7 representative bodies: (a)
Deperdussin Monocoque Monoplane Body, a single-seater; (b) N. P. L.-5
Tractor Biplane Body, single-seater; (c) B. F.-36 Dirigible Form,
without propeller or cock-pit openings; (d) B. E.-3. Two-Place Tractor
Body, with passenger and pilot; (e) Curtiss JN Type Tractor Body, with
passengers, chassis and tail; (f) Farman Pusher type, with motor,
propeller and exposed passengers; (g) N. P. L. Pusher Body, bare. Body
(a) was tested with a 1/5 scale model at a wind tunnel speed of 28
meters per second, the resistance of the model being 0.377 kilograms
(0.83 pounds). Body (d) in model form was 1/16 scale and was tested at
20.5 miles per hour, at which speed the resistance was 0.0165 pounds.
Model (e) was 1/12 scale and was tested at 30 miles per hour. These
varying test speeds, it will be seen, do not allow of a very accurate
means of comparison. The resistance of model (e) was 0.1365 pounds at
the specified wind-tunnel air speed.
The speeds given in the above table are simply translational speeds, and
are not corrected for slipstream velocity. With a slipstream of 25 per
cent, increase the body resistance by 40 per cent. It would be safe to
add an additional 10 per cent to make up for projecting fittings, baggy
fabric, and scale variations.
Since a body of approximately streamline form has a considerable
percentage of skin friction, scale corrections for size and velocity are
even of more importance than with wing sections. No wind-tunnel
experiments can determine the resistance exactly because of the
uncertainty of the scale factor. The resistance as given in the table is
also affected by the proximity of the wing and tail surfaces, and by
projections emanating from the motor compartment. It will be noted that
the dirigible form B.F.-36 is markedly better than any of the others,
being almost of perfect streamline form. The nearest approximation to
the ideal form is N.P.L.-5, which has easy curves, low resistance, and
is fairly symmetrical about the center line. Because of their small
size, the pusher bodies or "nacelles" have a small total resistance, but
the value of Kx is high.
[Illustration: Chart Showing Forms of 7 Typical Aeroplane Fuselage]
*Problem*. Find the resistance of a Curtiss Tractor Type JN body with a
breadth of 2’ 6" and a depth of 3’ 3", the speed being 90 miles per
hour. The slipstream is assumed to be 25 per cent, with an additional 10
per cent for added fabric loss, etc.
[Illustration: Typical Stream Line Strut Construction.]
*Solution*. The cross-sectional area = 2’ 6" x 3’ 3" = A = 8.13 square
feet. The velocity of translation is 90 M. P. H., or V² = 8100. The
value of the resistance coefficient is taken from the table, Ko-0.00273.
The total resistance R = KxAV² = 0.00273 x 8.13 x 8100 = 178.2 pounds.
Since a slipstream of 25 per cent increases the resistance by 40 per
cent, the resistance in the slipstream is 1782 x 1.4 = 249.48 pounds.
The addition of the 10 per cent for extra friction makes the total
resistance = 249.48 x 1.1 = 274.43 pounds. The resistance of this body,
used with "twin" motors, would be 178.2 x 1.1 = 196.02, but as a tractor
with the body in the slipstream, the resistance would be equal to 274.43
pounds as calculated above.
CHAPTER XVII. POWER CALCULATIONS.
*Power Units*. Power is the rate of doing work. If a force of 10 pounds
is applied to a body moving at the rate of 300 feet per minute, the
power will be expressed by 10 x 300 = 3000 foot-pounds per minute. As
the figures obtained by the foot and pound units are usually
inconveniently large, the "Horsepower" unit has been adopted. A
horsepower is a unit that represents work done at the rate of 33,000
foot-pounds per minute, or 550 foot-pounds per second. Thus if a certain
aeroplane offers a resistance of 200 pounds, and flies at the rate of
6000 feet per minute, then the work done per minute will be equal to 200
x 6000 = 1,200,000 foot-pounds. Since there are 33,000 foot-pounds of
work per minute for each horsepower, the horsepower will be:
1200000/33000 = 36.3.
As aeroplane speeds are usually given in terms of miles per hour, it
will be convenient to convert the foot-minute unit into the mile per
hour unit. If H = horsepower, R = resistance of aeroplane, and V = miles
per hour, then H = RV/375, the theoretical horsepower, without loss. If
an aeroplane flies at 100 miles per hour and requires a propeller thrust
of 300 pounds, then the horsepower becomes:
H = RV/375 = 300 x 100/375 = 80 horsepower. This is the actual power
required to drive the machine, but is not the engine power, as the
engine must also supply the losses due to the propeller. The propeller
losses are generally expressed as a percentage of the total power
supplied. The percentage of useful power is known as the "Efficiency."
The efficiency of the average aeroplane propeller will vary from 0.70 to
0.80. If e = propeller efficiency expressed as a decimal, the motor
horsepower becomes: H = RV/375e. To obtain the motor horsepower, divide
the theoretical horsepower by the efficiency. Using the complete formula
for the solution of an example in which the flight speed is 100 M. P.
H., the resistance 225 pounds and the efficiency 0.75, we have:
H = RV/375e = 225 x 100/375 x 0.75 = 80 horsepower.
*Power Distribution*. Since power depends upon the total resistance to
be overcome, part of the power will be used for driving the lifting
surfaces and a part for overcoming the parasitic resistance. The power
required for driving the wings depends upon the angle of incidence,
since the drag varies with every angle. The wing power varies with every
flight speed, owing to the changes in angle made necessary to support
the constant load. The power for the wings will be least at the speed
and angle that corresponds to the greatest lift-drag ratio. Owing to the
low value of the L/D at very small and very large angles, the power
requirements will be excessive at extremely low and high speeds.
As the parasitic resistance increases as the square of the speed, the
power for overcoming this resistance will vary as the cube of the speed.
It is the parasitic resistance that really limits the higher speeds of
the aeroplane, since it increases very rapidly at velocities of over 60
miles per hour.
The total power at any speed is the sum of the wing power and power
required for the parasitic resistance. Owing to variations in the wing
drag and resistance at every point within the flight range, it is
exceedingly difficult to directly calculate the total power at any
particular speed. The wing drag and the resistance should be calculated
for every speed, and then laid out by a graph or curve. The minimum
propeller thrust, or the minimum total resistance, occurs approximately
at the speed where the body resistance and wing drag are equal. The
minimum horsepower occurs at a low speed, but not the lowest speed, and
this will differ with every machine.
[Illustration: Fig. 1. Power Chart of Bleriot Monoplane, With Outline of
Wing Section. The Results Were Taken From Full Size Tests Made by the
English Government.]
Fig. 1 is a set of performance curves drawn from the results of tests on
a full size Bleriot monoplane. At the bottom the horizontal row of
figures gives the horizontal speed in feet per second. The first column
to the left is the horsepower, and the second column is the resistance
or drag in pounds. The four curves represent respectively the body
resistance, wing or "plane" resistance, horsepower, and total
resistance. The horizontal line "AV" shows the available horsepower. It
will be noted that the body resistance increases steadily from 9 pounds
at 50 feet per second, to 180 pounds at 100 feet per second. The wing
resistance, on the other hand, decreases from 350 pounds at 56 feet per
second to a minimum of 130 pounds at 83 feet per second. It will be
noted that the angles of incidence are marked along the wing-drag curve
by small circles. The incidence is 6° at 75 feet per second, and 4° at a
little less than 85 feet per second.
The available horsepower "AV" is 42. This is shown as a straight line,
although in the majority of cases it is slightly curved owing to
variations in power at the higher speeds, and to variations in the
propeller efficiency. At 90 feet per second the actual horsepower curve
crosses the line of available horsepower "AV." Beyond this point
horizontal flight is no longer possible, as the power requirements would
exceed the available horsepower. It will be noted that the lowest total
resistance occurs near the point where the body and wing resistance
curves intersect, or in other words, where the body and wing resistance
are equal. The minimum horsepower takes place at 63 feet per second, or
at a point nearly 1/3 between the lowest flight speed and the highest
speed attained by the available horsepower in horizontal flight (90
ft/sec).
The actual range of flight speeds is limited to points between the
intersection of the "Horsepower required" curve, and the "Available
horsepower" curve. By increasing the propeller efficiency, or by
increasing the power of the motor, the available horsepower line is
raised and the flight range increased.
*Horsepower For Climbing*. Up to the present we have only considered
horizontal flight. The power available for climbing is the difference
between the power required to maintain horizontal flight at any speed,
and the actual horsepower that can be delivered by the propeller. Thus,
if the actual power delivered by a motor through the propeller is 85
horsepower, and the power required for horizontal flight at that speed
is 45, then we have: 85–45= 40 horsepower available for climbing. Since
the difference between the driving power and the power required for
horizontal flight is less at extremely low and high speeds, it is
evident that we will have a minimum climbing reserve at the high and low
speeds. Consulting the power curve for the Bleriot monoplane, we see
that the power required at 56 feet per second is 40 horsepower, and at
85 feet per second it is 38 horsepower. At the low speed we have a
climbing reserve of 44–40 = 4 H. P., and at the higher speed 44–38 = 6
H. P. The maximum available horsepower "AV" is 44 horsepower. The
minimum horizontal power required is found at 63 feet per second, the
climbing reserve at this point being 44–28 = 16 horsepower. At 55 feet
per second, and at 90, we would not be able to climb, as we would only
have sufficient power to maintain horizontal flight.
If W = total weight of aeroplane, c = climbing speed, and H = horsepower
reserve for climbing, then the climbing speed with a constant air
density will be expressed by: c = 33000H/W. Assuming that the weight of
the Bleriot monoplane is 800 pounds, and that we are to climb at the
speed of the greatest power reserve (16 horsepower), our rate of climb
is:
c = 33000H/w = 33000 x 16/800 = 660 feet per minute. It should be
understood that this is the velocity at the beginning of the climb.
After prolonged climbing the rate falls off because of diminishing power
and increasing speed. Much depends upon the engine performance at the
higher altitudes, so that the reserve power for climb usually diminishes
as the machine rises, and hence the rate of climb diminishes in
proportion.
The following table taken from actual flying tests will show how the
rate of climb decreases with the altitude. These machines were equipped
with 150 H. P. Hispano-Suisa motors. It will be noted that the S. P. A.
D. and the Bleriot hold their rate of climb constant up to 7800 feet
altitude, which is a feat that is undoubtedly performed by varying the
compression of the engine.
Besides increasing the power, the rate of climb can also be increased by
decreasing the weight of the aeroplane.
*Maximum Altitude*. The maximum altitude to which a machine can ascend
is known as its "Ceiling." This again depends on both the aeroplane and
the motor, but principally on the latter. It has been noted that
machines having the greatest rate of climb also have the greatest
ceiling. Thus the ceiling of a fast climbing scout is higher than that
of a larger and slower machine. Based on this principle, a writer in
"Flight" has developed the following equation for ceiling, which, of
course, assumes a uniform decrease in density. Let H = maximum altitude,
h = the altitude at any time t after the start of the climb, and a = the
altitude after a time equal to twice the time t, then:
H = h / (2 - a/h)
Approximate values of h and a may be had
from the following table, which are the results of a test on a certain
aeroplane:
*Time* (Minutes)
0.0 2.5 5.0 7.5 10.0 12.5 15.0 17.5 20.0 22.5
*Altitude* (Feet)
0.03300 6150 8730 10760 12610 14190 15530 16650 17600
If we assume that the height is 10760 feet after the first 10 minutes,
and that the altitude after twice this time (20 minutes) is 16650 feet,
then the maximum ceiling attained will be:
H = h / (2 - a/h) = 10760 / 2 - 16650/10760 = 23,770 feet. The use of
this formula requires that the climb be known for certain time intervals
before the ceiling.
*Gliding Angle*. The gliding angle of the wings alone is equal to the
lift-drag ratio at the given angle. The best or "Flattest gliding angle"
is, of course, the best lift-drag ratio of the wing—say on the average
about 1 in 16. The gliding angle of the complete machine is considerably
less than this, owing to the resistance of the body and structural
parts. This generally reduces the actual angle to less than 12, and in
most cases between 6 and 8. Expressed in terms of degrees, tan ø = R/W
where R = head resistance and W = weight in pounds.
Fig. 2 is a diagram giving the gliding force diagram. The plane descends
along the gliding path AC, making the angle of incidence (ø). When in
horizontal flight, the lift is along OL and the weight is OW. When
descending on the gliding path the lift maintains the same relation with
the wing, but the relative angle of the weight is altered. The weight
now acts along OG. The drag is represented by OD, with the propeller
thrust OP equal and opposite to it. With the weight constant, the lift
OL is decreased by the angle so that the total life = L = W cos ø. The
action of the weight W produces the propelling component OP that gives
forward velocity. The line AB is the horizontal ground line. If the
total lift-drag ratio is 8, then the gliding angle will be 1 in 8, or
measured in degrees, tan ø = R/W = 1/8 = 0.125. From a trigonometric
table it will be found that this tangent corresponds to an angle of 7° –
10’. It should be noted that R is the total resistance and not the
wing-drag.
[Illustration: Fig. 2. Gliding Angle Diagram Showing Component of
Gravity That Causes Forward Motion. The Gliding Angle Depends Upon the
Ratio of the Resistance to the Weight.]
*Complete Power Calculations*. Knowing the total weight and the desired
speed, we must determine the wing section and area before we start on
the actual power calculations. This can either be determined by
empirical rules in the case of a preliminary investigation, or by actual
calculation by means of the lift coefficients after the approximate
values are known. Sustaining a given weight, we can vary the angle,
area, wing section, or the speed, the choice of these items being
regulated principally by the power. Given a small area and a great angle
of incidence, we can support the load, but the power consumption will be
excessive because of the low value of the L/D ratio at high angles. If
small area is desired, a large value of Ky due to a high lift-wing
section is preferable to a low lift wing at high angles. In general, the
area should be so arranged that the wing is at the angle of the maximum
lift-drag ratio at the rated speed. A low angle means a smaller motor,
less fuel, and hence a lighter machine. This selection involves
considerable difficulty, and a number of wing sections and areas must be
tried by the trial and error method until the most economical
combination is discovered.
[Illustration: Graphical Gliding Diagrams of Several Aeroplanes Recorded
in British Army Contest of 1912.]
The first consideration being the total weight, we must first estimate
this from the required live load. This can be estimated from previous
examples of nearly the same type. Say that our required live load is 660
pounds, and that a live load factor of 0.30 is used. The total weight
now becomes 660/0.30=2200 pounds. To make a preliminary estimate of the
area we must find the load per square foot. An empirical formula for
biplane loading reads: w = 0.065V - 0.25 where V = maximum speed in M.
P. H., and w = load per square foot. If we assume a maximum speed of 90
M. P. H. for our machine, the unit loading is w = (0.065 x 90) - 0.25 =
5.6 pounds per square foot. The approximate area can now be found from
2200/5.6 = 393 square feet. (Call 390.) The minimum speed is about 48
per cent of the maximum, or 43 M. P. H. We can now choose one or more
wing sections that will come approximately to our requirements by the
use of the basic formula, Ky = w/V².
At high speed, Ky = 5.6/(90 x 90) = 0.000691. At low speed, Ky = 5.6/(40
x 40) = 0.003030. We must choose the most economical wing between these
limits of lift, and on reference to our wing section tables we find:
It would seem from the above that the chosen area is a little too large,
as the majority of the L/D ratios at high speed are poor, the best being
11.00 of the U.S.A.-1. The angles are small, being negative in most
cases at high speed. While the lift-drag of the R.A.F-3 is very good at
low speed, it is very poor at high, hence the area for this section
should be reduced to increase the loading. The R.A.F.-6 and the U.S.A.-1
show up the best, for they are both near the maximum lift at low speed
and have fair L/D ratios at high speed. It will be seen that for the
best results there should be a series of power curves drawn for the
various wings and areas. This method is too complicated and tedious to
take up here, and so we will use U.S.A.-1, which does not really show up
so bad at this stage. Both the R.A.F.-6 and the U.S.A.-1 have been used
extensively on machines of the size and type under consideration. While
we require Ky = 0.003030, and U.S.A-1 gives 0.003165, we will not
attempt to utilize this excess, as it will be remembered that we should
not assume the maximum lift for reasons of stability.
The wing-drag at high speed will be 2200/11.0 = 200 pounds, and at low
speed it will be: 2200/10.4 = 211 pounds. Since the maximum L/D is 17.8
at 3°, where Ky is 0.00133, the least drag will be: 2200/17.8 = 124
pounds. This least drag will occur at V = V5.6/000133 = 65 M. P. H.
The wing drag for each speed must now be divided by the correction
factor 0.85, which converts the monoplane values of drag into biplane
values. Since this is practically constant it does not affect the
relative values of Kx in comparing wings, but it should be used in final
results. For this type of machine we will take the total parasitic
resistance as r = 0.036V². At 90 M.P.H., r = 0.036 x 90 x 90 = 291.6
pounds. At 65 M. P. H., the resistance is: 0.036 x 65 x 65 = 152.1. At
the extreme low speed of 43 M. P.H. we have r - 0.036x43 x 43 = 66.56
pounds. The total resistance (R) is equal to the sum of the wing-drag
and the parasitic resistance. At 90 M. P. H. the total resistance
becomes 200 + 291.6 = 491.6 pounds. At 65 M.P.H. the total is 124 +
152.1 = 176.1, and at 43 M.P.H. it is 211 + 66.56 = 277.56 pounds. The
horsepower is computed from H = RV/375e, and at 90 M. P. H. this is : H
= 491.6 x 90/375 x 0.80 = 147.5 H. P. where 0.80 is the assumed
propeller efficiency. At 65 M. P. H. the horsepower drops to H = 176.1 x
65/375 x 0.8 = 38.1 H. P., assuming the same efficiency. In the same way
the H. P. at 43 M. P. H. r is 39.8.
A table and power chart should be worked out for a number of sections
and areas according to the following table. The calculations should be
computed at intervals of 5 M. P. H., at least the lower speeds. Wing
drag is not corrected for biplane interference:
*Weight and Power*. The weight lifted per horsepower varies in the
different types of aeroplanes, this difference lying principally in the
reserve allowed for climbing and horizontal speed. A speed scout may
carry as little as 8 pounds per horsepower, while a slow two-seater may
exceed 20 pounds per horsepower. A rough estimate of the horsepower
required may be had by dividing the total weight by the weight per
horsepower ratio for that particular type. Thus if the unit H. P.
loading is 16 pounds and the total weight is 3200, then the horsepower
will equal 3200/16 = 200 horsepower. Assuming that the live load w’ is
0.32 of the total weight W, then W = w’/0.32. If m = lbs. per H.P., then
H = W/m or H = w’/0.32m. Taking the case of a training machine where m =
20, and the live load is 640 pounds, the approximate horsepower will be:
H = w’/0.32m = 640/0.32 x 20 = 100 horsepower. A speed scout carrying
320 pounds useful load, with m = 10, will require H = 320/0.32 x 10 =
100 horsepower.
CHAPTER XVIII. PROPELLERS.
*Principles and Use of Propellers*. A propeller converts the energy of
the engine into the thrust required to overcome the resistance of the
aeroplane. To maintain flight the thrust, or force exerted by the
propeller, must always equal the total resistance of the aeroplane. A
total resistance of 400 pounds requires a propeller thrust of 400, and
as the resistance varies with the speed, the engine revolutions must be
altered correspondingly. The propeller is the most complicated and least
understood element of the aeroplane, and we can but touch only on the
most elementary features. The inclined blades of the propeller throw
back an airstream, the reaction of which produces the thrust. The blades
can also be considered as aerofoils moving in a circular path, the lift
of the aerofoils corresponding to the thrust of the propeller. The
reactions in any case are quite complicated and require the use of
higher mathematics for a full understanding.
*Pitch and Velocity*. When in action the propeller rotates, and at the
same time advances along a straight line parallel to its axis. As a
result, the tips of the propeller blades describe a curve known as
"Helix" or screw-thread curve. The action is very similar to that of a
screw being turned in a nut. For clearness in explanation we will call
the velocity in the aeroplane path the "Translational velocity," and the
speed of the tips in their circular path as the "Rotational velocity."
When a screw works in a rigid nut it advances a distance equal to the
"Pitch" in each revolution, the pitch of a single threaded screw being
equal to the distance between the threads. Since the propeller or "Air
screw" works in a fluid, there is some slip and the actual advance does
not correspond to the "Pitch" of the propeller blades. The effective
pitch is the distance traveled by the propeller in one revolution. The
actual pitch or the angle of the blades must be greater than the angle
of the effective helix by the amount of slip.
If N = Revolutions per minute, P = effective pitch in feet and V =
translational velocity in miles per hour, then V= NP/88. With an
effective pitch of 5 feet, and 1200 revolutions per minute, the
translational velocity of the aeroplane will be: V = 1200 x 5/88=68.2
miles per hour.
[Illustration: Excelsior Propeller, an Example of American Propeller
Construction. This Propeller Is Built Up of Laminations of Ash.]
The actual pitch of the blades would be from 15 to 25 per cent greater
than the effective pitch because of the slip. To have thrust we must
have slip. With the translational velocity equal to the blade-pitch
velocity, there is no airstream accelerated by the blades, and
consequently there is no thrust due to reaction. The air thrown to the
rear of a propeller moves at a greater speed than the translation when
thrust is developed, and this stream is known as the "slipstream." The
difference between the translational and slipstream velocity is the
slip.
The angle of the blade face determines the pitch. The greater the angle
of the blade with the plane of propeller rotation, the greater is the
pitch. This angle is measured from the chord of the working face of the
table, or from that side faced to the rear of the blade. In the majority
of cases the working face is flat. The front face is always heavily
cambered like a wing section, with the greatest thickness about
one-third the chord from the entering edge. As in the case of the wing,
the camber is of the greatest importance.
A uniform pitch propeller has a varying blade angle, smallest at the tip
and increasing toward the hub. With a uniform pitch propeller, every
part of the blade travels through the same forward distance in one
revolution, hence it is necessary to increase the angle toward the hub
as the innermost portions travel a smaller distance around the circle of
rotation. Theoretically, the angle at the exact center would be 90
degrees. The blade angles at the different points in the length of a
uniform pitch propeller are obtained as follows: Draw a right angle
triangle in which the altitude is made equal to the pitch, and the base
is equal to 3.1416 times the propeller diameter. The angle made by the
hypotenuse with the base is the blade angle at the tip. Divide the base
into any number of equal spaces and connect the division points with the
upper angle. The angles made by these lines with the base are the angles
of the different blade sections.
*Blade Form*. The blade may be either straight-sided or curved. In the
latter case the most deeply curved edge is generally the entering edge,
and the maximum width is about one-third from the tip. Much care is
exercised in arranging the outline so that the center of pressure will
not be located in an eccentric position and thus harmfully distort the
blade when loaded. If this is not attended to, the pitch will vary
according to the load. In one make of propeller the blade is purposely
made flexible so that the pitch will accommodate itself correctly to
different flight speeds and conditions. This, however, is carefully laid
out so that the flexure is proportional throughout the blade to the
changes in the load.
[Illustration: The Lang Propeller, Having Straight Edges, Slightly
Tapering Toward the Tips. The Tips Are Sheathed With Thin Copper for
Protection Against Spray. This Outline Is Often Known as the "Normale."
Type From the French Propeller First Using This Outline.]
[Illustration: A "Paragon" Propeller With a Curved Leading Edge. The
Maximum Width Is About One-Third the Blade Length from the Tip and §o.
Toward the Tip So That It Is Very Narrow at the Outer End. The Steel
Propeller Flange Is Shown in Place on the Hub.]
*Propeller Diameter*. The largest propellers are the most efficient. The
propeller should be as large as can be safely swung on the aeroplane.
Large, slow revolution propellers are far superior to the small high
speed type. It is more economical to accelerate a large mass of air
slowly with a large diameter than to speed up a small mass to a high
velocity. The diameter used on any aeroplane depends upon the power
plant, propeller clearance, height of chassis and many other
considerations. Approximately the diameter varies from about 1/3 the
span on small speed scouts, to 1/5 or 1/6 of the span on the larger
machines.
*Air Flow*. The greater part of the air is taken in through the tips,
and is then expelled to the rear. This condition prevails until the
blade angle is above 45 degrees, and from this point the flow is
outward. Owing to the great angles at the hub, there is little thrust
given by the inner third of the blade, the air in this region being
simply churned up in a directionless mass of eddies. At the tips the
angle is small and the velocity high, which results in about 80 per cent
of the useful work being performed by the outer third of the blade. In
some aeroplanes a spinner cap is placed around the hub to reduce the
churning loss and to streamline the hub. The blade section is very thick
at the hub for structural reasons.
The "Disc area" of a propeller is the area of the circle swept out by
the blades. It is the pressure over this area that gives the thrust, and
in some methods of calculation the thrust is based on the mean pressure
per square foot of disc area. The pressure is not uniformly distributed
over the disc, being many times greater at the outer circumference than
at the hub. The average pressure per square foot depends upon the blade
section and angle. Because of the great intensity of pressure at the
circumference, the effective stream is in the form of a hollow tube.
*Number of Blades*. For training, and ordinary work, two-bladed
propellers are preferable, but for large motors where the swing is
limited, three or four blades are often used. A multiple-bladed
propeller absorbs more horsepower with a given diameter than the
two-blade type. In general, a four-bladed propeller revolving slowly may
be considered more efficient than the two-blade revolving rapidly. Where
the swing and clearance are small, a small four-blade may give better
results than a larger and faster two-blade. A three-blade often shows
marked superiority over a two-blade even when of smaller diameter, and
the hub of the three-blade is much stronger than the four-blade,
although neither the three or four is as strong as the two-blade type.
*Effects of Altitude*. At high altitudes the density is less, and
consequently the thrust is less with a given number of revolutions per
minute. The thrust can be maintained either by increasing the speed, or
by increasing the pitch. For correct service at high altitudes the
propeller should undoubtedly be of the variable pitch type, in which the
pitch can be controlled manually, or by some automatic means such as
proportional blade flexure.
*Effects of Pitch*. Driven at a constant speed, both the thrust and
horsepower increase with the pitch up to a certain limiting angle.
For a given horsepower the static thrust depends both on the diameter
and the pitch. If the pitch is increased the diameter must be decreased
in proportion to maintain a constant speed. As the pitch is regulated by
the translational speed and revolutions, the static thrust of a high
speed machine is very small. As the translational speed increases, the
pitch relative to the wind is less, and consequently the thrust will
pick up until a certain limiting speed is reached.
*Thrust and Horsepower*. The calculation for thrust and power are very
complicated, but the primary conditions can be given by the following:
Let V = the pitch velocity in feet per minute, T = thrust in pounds, and
H = horsepower, then H = TV/33000E from which T = 33000HE/V, the
efficiency being designated by E. Since the pitch velocity is NP, where
N = revs. per minute and P = pitch in feet, then T = 33000HE/PN.
Assuming a 5-foot pitch, 1200 revs., the efficiency = 0.75, and the
horsepower 100, the thrust will be:
T = 33000 x 100 x 0.75/5 x 1200 = 412.5 pounds. The pitch in this case
is the blade pitch, and the great uncertainty lies in selecting a proper
value for E. This may vary from 0.70 to 0.85. The diameter is also an
unknown factor in this primitive equation.
*Materials and Construction*. The woods used for propeller construction
are spruce, ash, mahogany, birch, white oak, walnut, and maple. Up to 50
H. P. spruce is suitable, as it is light, and strong enough for this
power. In Europe walnut and mahogany are the most commonly used,
although they are very expensive. Birch is very strong and comparatively
light for its strength, and can be used successfully up to 125
horsepower. Ash is strong, light and fibrous, but has the objectionable
feature of warping and cannot withstand moisture. Maple is too heavy for
its strength. White oak, quarter-sawed, is the best of propeller woods
and is used with the very largest engines. It is strong for its weight
and is hard, but is very difficult to work and glue. For tropical
climates, Southern poplar is frequently used as it has the property of
resisting heat and humidity.
One-inch boards are rough dressed to 7/8 inch and then finished down to
13/16 or 3/4 inch. After a thorough tooth planing to roughen the surface
for the glue, they are thoroughly coated with hot hide glue, piled
together in blocks of from 5 to 10 laminations, and then thoroughly
squeezed for 18 hours in a press or by clamps until the glue has
thoroughly set. Only the best of hide glue is used, applied at a
temperature of 140°F. and at a room temperature of 100°. The glue must
never be hotter, nor the boards cooler than the temperatures stated. The
propeller after being roughed out is left to dry for ten days so that
all of the glue stresses are adjusted. If less time is taken, the
propeller will warp out of shape. The propeller is worked down within a
small fraction of the finished size and is again allowed to rest. After
a few days it is finished down to size by hand, is scraped, and tested
for pitch, tracking and hub dimensions.
The finish is glossy, and may be accomplished by several coats of spar
varnish or by repeated applications of hot boiled linseed oil well
rubbed in, finishing with three or four coats of wax polish. There
should be at least 5 applications of linseed oil, the third coat being
sandpapered with No. 0 paper. The wood should be scraped to dimension
and must not be touched with sandpaper until at least two coats of
varnish or oil have been applied.
The wood must be absolutely clear and straight grained, and without
discolorations. The boards must be piled so that the edge of the grain
is on the face of the blades, and the direction of the annular rings
must be alternated in the adjacent boards.
[Illustration: Plan and Side Elevation of the S.P.A.D. Speed Scout.
Courtesy "_Aerial Age_".]
CHAPTER XIX. OPERATION AND TRAINING.
*Self-Training*. In the early days of aviation, there were few schools,
and these were so expensive to attend that the majority of the
aeronautical enthusiasts taught themselves to fly on home-made machines.
While this was a heroic method, it had the advantage of giving the
student perfect confidence in himself, and if his funds were sufficient
to outlast the crashes, it resulted in a finished and thorough flyer. In
general, this process may be described as consisting of two hours of
practice followed by a week or more of repairing.
The present-day beginner has many advantages. He has the choice of many
excellent schools that charge a reasonable tuition, and where the risk
of injury is small. He has access to the valuable notes published in the
aeronautical magazines, and the privilege of consulting with experienced
aviators. The stability and reliability of the planes and the motors
have also been improved to a remarkable degree, and the student no
longer has to contend with a doubtful aeroplane construction nor with
the whims of a poorly-constructed motor.
*Training Methods*. In the majority of American schools, the instructor
accompanies the student in the first flights. The controls are "Dual,"
or interconnected, so that the instructors’ controls act in unison with
those of the student, thus giving the latter an accurate knowledge of
the movements necessary for each flight condition. After the first few
flights the instructor can relax his controls at times so that the
student can take charge. This continues until the student has shown the
ability to handle the machine alone under ordinary conditions and is
then ready for his first "Solo" or flight alone. The first solo is a
critical period in his training, for when once in flight he is beyond
all human aid.
At the navy training school at Pensacola, the student is first taken for
a ride with one of the instructors without giving him access to the
controls. This is simply to give the student an experience in the
sensation of flight. After this he is taken for a series of short
flights on a dual control machine, the instructor gradually allowing him
to take charge to a greater and greater extent as he develops the "Air
feel." During this time the intricacies of the maneuvers are also
gradually increased, so that after about ten hours of this sort of work
he is allowed to take his first solo. It has been found that the average
student will require from 10 to 20 hours of dual control instruction
before he is fit to fly alone. When his work has proven satisfactory he
is then allowed to fly in rough weather, execute spirals, and attempt
high altitude and long distance flying.
Some instructors believe in showing what can be performed in the air
from the very beginning. During the first dual flights, the pilot
indulges in dives, vertical banks, side slip, or even looping. After an
experience of this sort, the student is far more collected and easy
during the following instructions in simple straight flying. If this
preliminary stunt flying has a very material effect on the nerve of the
student it may be taken for granted that he is not adapted for the work
and can be weeded out without further loss of time. If he is of the
right type, this "rough stuff" has a beneficial influence on his work
during the succeeding lessons. During this time numerous landings are
made, for it must be understood that this is one of the most difficult
features of flying. With 15 minute lessons, at least 6 landings should
be made per lesson.
A second method of instruction, and to the author’s mind the most
desirable, is by means of the "Penguin" or "Roller." This is a
low-powered machine with very small wings—so small that it cannot raise
itself from the ground. By running the penguin over the ground, the
student learns how to manage his engine and to steer with his feet. In
this way he obtains a certain delicacy of touch without endangering
either himself or an expensive machine. After he has progressed
satisfactorily on this machine he graduates to a faster penguin or else
to a very slow aeroplane with which he can actually leave the ground.
Since the second penguin, or the slow aeroplane are much faster than the
first machine, the student finds that the sensitiveness of the rudder
and controls are greatly increased. They require more careful handling
than in the first instance, and the slightest mistake or delay will send
the machine skidding. The aeroplanes used at this stage are very
low-powered, and are capable of rising only a few feet from the ground,
but they give the student an opportunity of learning the aileron and
tail controls in comparative safety. The same result can be obtained
with a standard aeroplane by a permanent set in the throttle control,
and by adjusting the stabilizer surface. The beginner is allowed to work
only during calm weather, as the low speed and small lifting capacity is
likely to cause an accident if the machine is caught by a side or
following gust. He only learns how to get the machine off the ground, to
keep the tail up and hold it in a straight line for a few seconds.
The man taught by the penguin method is alone when he first leaves the
ground, and hence is generally more self-reliant than one who has been
"Spoon fed." His experience in handling the controls has made his
movements instinctive, so that when he first actually flies he is in a
better position to analyze the new problems before him. It is a better
and cheaper method for the school as the breakage is less expensive and
allows the unfit students be weeded out before they cause damage to
themselves or to the school property.
*Ground Instruction*. Before attempting flight, the student should be
thoroughly versed in the principles an constructional details of the
aeroplane and the aeronautic motor. He should know how to take down,
time and repair every type of motor with which he is likely to come into
contact. He should be able to tell at a glance whether the machine is
rigged or trued up properly, and have a general knowledge of the
underlying principles of aerodynamics. The study of these subjects is
the function of the ground school. At this school the student should
learn the assembling and adjusting of the aeroplane structure and its
balance.
*Types Suitable for Pilots*. There is a great diversity of opinion as to
the type of man best suited for flying. In this country the government
requirements regarding age and physical condition are very exacting,
while in Europe it has been found that physical condition is not an
index to a man’s ability as a pilot. Many of the best French pilots were
in such bad shape as to be rejected by the other branches of the army.
Our men are well under 30 years of age, while in European service there
are many excellent pilots well over 40. It is almost impossible to tell
from external appearances whether a man can become a good pilot.
In general he must be more intelligent and better educated than the
average infantryman. He should not be subject to an attack of "Nerves,"
nor become easily rattled, for such a man courts disaster in flying.
Many exhibition flyers of reputation have proved absolute failures in
military service. A knowledge of mechanics will be of great benefit and
has been the salvation of many a pilot in active service. Automobile or
motorcycle experience is particularly valuable. Recklessness, or a
dare-devil sort of a disposition, are farthest from being qualifications
for an aviator. Such a man should not be permitted to fly, for he is not
only a constant menace to himself but to everyone else concerned.
*Learning to Fly Alone*. It is with the greatest hesitancy that the
author enters into a "Ground course" of flight instruction. I can,
however, list the principal things to avoid and some of the things to
do, but this will never take the place of actual field instruction and
experience. The first and last thing to remember is to "Proceed slowly
and with caution." Never try a new stunt until you are absolutely sure
that you have thoroughly mastered the preliminary steps in straight
flying. Over-confidence at the beginning is almost as bad as no
confidence at all, and the greatest difficulty met with by instructors
during the first solo flights is to keep the student from imitating the
maneuvers of the more experienced flyers. Spend plenty of time rolling
or "Grass-cutting" before attempting to leave the ground. Be sure that
you can handle the rudder with accuracy, and at fairly high speeds
before attempting to lift. A few days spent in sitting in the machine
(motor dead), and acquainting yourself with the controls is excellent
practice and certainly is not a loss of time. With the machine in the
hangar, move the controls for imaginary turns, dips and other maneuvers
so that the resistance, reach and limit of control movement will come
more naturally when the machine is moving.
During the ground rolling period, the elevator or stabilizer should be
set so that it is impossible to leave the ground, and the motor should
be adjusted so that it cannot develop its full thrust. This will provide
against an accidental lift. Be easy and gentle in handling the controls,
for they work easily, and have powerful effect at high speeds. The
desperate fervor with which the beginner generally yanks at the
"joystick" is generally the very reason for his accidents. Do not start
off at full speed without first getting used to the effect of the
controls. Learn to find the location of the various devices so that you
can reach them without looking or without fumbling.
*The First Straight*. By adjusting the stabilizer and elevators so that
the latter has a greater degree of freedom, and by changing the motor so
that it can be run at a slightly higher speed, we are in a position to
attempt our first flight. Be careful that the adjustment will limit the
climb of the machine, and choose only the calmest of weather. It should
be remembered that the aeroplane will get off the ground at a lower
speed than that required for full flight at higher altitudes, this being
due to the cushioning effect between the wings and the earth. A machine
traveling at a speed capable of sustaining flight at a few feet above
the earth will cause it to stall when it is high enough to lose this
compression. The adjustment should be such that the machine cannot rise
above this "Cushion," and in this condition it is fairly safe for the
beginner.
In making the first runs under the new conditions of adjustment, the
student should learn to manipulate the elevators so that they will hold
the tail up in the correct position, that is, with the chord of the
wings nearly horizontal. Do not allow the tail skid to drag over the
ground further than necessary. At this point the student should be
strapped in the seat by a quick-detachable safety belt.
Now comes the test. Get under full headway with the tail well up, taking
care to run against the breeze. The speed increases rapidly, then the
motion and jar seem softer, and the motor ceases to roar so loudly.
There is now a very distinct change in the note of the motor. You are
off. At this point a very peculiar illusion takes place, for your
elevation of a few feet seems about a thousand times greater than it
really is. With this impression the student usually tries to correct
matter by a sudden forward push on the control lever causing fine dive
and a smash. It must be borne in mind that only the slightest movement
of the controls should be made, and if this does not prove sufficient
after a moment or so, advance them still further but very gently. Sudden
movements must be avoided. At first the "Hops" should not extend over a
hundred yards or so until the student is sure of his controls. Little by
little they can be increased in length and height. He should practice
for some time before attempting a flight of more than a mile. By this
time, the student will have learned that the landing is by far the most
difficult feature in flying, and he should practice this incessantly
before trying flights in windy weather.
The machine should be headed directly into the wind, both in getting off
and in landing, especially in the latter case, as a sudden following
gust will tend to stall a machine or upset it. With a head wind, the
lift is maintained at a low speed and hence is an aid in a safe landing.
When flying in still air there is little if any use for the ailerons,
but in gusts the student will need their aid in maintaining lateral
balance. After the rudder and elevator controls have been well learned
the effect of the ailerons can be tried. Gusty or squally weather must
be avoided at this point in the training, and no turns should yet be
attempted.
When the student attains heights greater than a few feet he should take
great care in obtaining a sufficient ground speed before trying to get
off, for if lifted before the full flying speed is attained it is likely
to stall. Fast climbing at sharp angles is dangerous unless a sufficient
ground speed has been attained. Sustentation is due to forward speed,
and this must not be forgotten. The quickest climb for getting over
trees and other obstructions is obtained by gaining full speed on the
ground before the climb begins, as the power of the engine is aided by
the momentum of the machine.
In landing in small fields it is necessary to bring the machine to rest
as soon as possible, and this stopping distance depends to a great
extent upon the attitude of the machine when it first touches the
ground. If it is landed so that the chassis wheels and tail skid strike
the ground simultaneously, the incidence is so great that the wings act
as air brakes. On landing, the angle in any case should be quickly
increased past the angle of maximum lift. The lift is much reduced and
the drag is increased by quickly pulling the control toward the aviator.
This also reduces the tendency toward nosing over.
A normal landing in a large field can be affected by first starting down
at the normal gliding angle, and when from twenty to thirty feet above
the ground the elevator control is pulled back so that the machine will
describe a curve tangent to the ground. In student’s practice the curve
should not be exactly tangent to the ground, but tangent to a level two
or three feet above the ground. The machine is now losing speed, and to
prevent settling the elevator should be pulled back a trifle. The speed
continues to decrease until it settles down through the small remaining
distance with the elevator full back. The points of support should
strike simultaneously. It is difficult for the beginner to make this
sort of a landing, as there always seems to be an uncontrollable desire
to jam the machine down on the ground. If a puff of wind happens to
strike the machine when a few feet off, the student becomes rattled by
the suddenly increased elevation and jams her down doubly hard.
*Wind Flying*. The nature of wind at low altitudes is determined to a
great extent by the contour of the ground. Eddies are caused by trees,
embankments, fences, small hills, etc., which tend to disturb the
equilibrium or change the course of the aeroplane. As the altitude
increases, the effects of these obstructions are less pronounced, until
at from 2000 to 3000 feet the effect is practically negligible. Winds
that may be "Bumpy" near the ground are fairly regular when 3000 feet is
attained. At the higher altitudes the velocity increases, and if the
machine is flying against the wind the progress will naturally be much
slower at the higher altitudes. When starting in a strong wind it is
advisable to attain an altitude of at least 300 to 400 feet before
turning. Turning in with the wind carries the possibility of a drop or
stall.
A short gust striking the machine, head on, tends to retard the velocity
in regard to the earth, but in reality increases the relative air speed
and thus causes the machine to climb momentarily. A prolonged head gust
may produce a stall unless corrected by the elevator or met with by
reserve power. A rear gust reduces the relative wind velocity and tends
to make the machine stall, although there are a few cases where the gust
velocity has been great enough to cause a precipitate drop. The higher
the speed, the less the danger from rear gusts.
The gusts are much more pronounced with low winds, say winds of about 5
to 15 miles per hour, and hence it is usually more tricky to fly in a
wind of this velocity than with a higher wind. It is not the speed of
the wind so much as it is its variation from the average velocity. One
should start to work on a "bump" at the moment it first starts to
appear.
When flying with the wind, the total speed in regard to the earth is the
sum of the wind speed and the aeroplane speed. When flying against it is
the difference between the aeroplane and air speeds. Thus, if the air
speed of the aeroplane is 60 miles per hour, the speed in regard to the
ground will be 75 miles per hour with a following wind of 15 miles per
hour, and 45 miles per hour when flying against a 15-mile wind. The
speed when flying across the wind would be represented by the diagonal
of a parallelogram, one side of which represents the aeroplane speed,
and the other side the wind speed. The angle of the diagonal is the
angle at which the machine must be pointed. When viewed from the ground,
an aeroplane in a cross wind appears to fly sideways.
*Turning*. After the beginner is able to maintain longitudinal and
lateral balance on straight away flights, he next attempts turns. At
first, the turns must be of great radius. As the radius is gradually
shortened, the effects of centrifugal force become greater, increasing
the tendency toward skidding or outward side slip. To prevent skidding,
the outer wing tip must be raised so that the lift will oppose the
centrifugal force. The shorter the turn, and the faster it is made, the
greater will be the banking angle. Should the bank be too steep, the
gravitational force will pull the machine down, and inwardly in a
direction parallel to the wings. This is known as an "Inner side slip."
The banking may be performed by the natural banking tendency of the
aeroplane or may be assisted by depressing the aileron on the outer wing
tip. Unless the speed is well up to normal, the machine will be likely
to stall and drop on a turn, as the head resistance is much greater
under these conditions. For safety one should take a short downward
glide before starting the turn, so that the speed will surely be
sufficient to carry it around the turn. A turn should never be attempted
when climbing unless one has a great reserve power. The combined effects
of the turning resistance, and absorption of energy due to the climb,
will be almost certain to stall the machine. There are banking
indicators on the market which will prove of great service. These
operate on the pendulum principle and indicate graphically whether the
aeroplane is being held at the correct angle of bank.
*Proper Flight Speed*. An aeroplane should always be provided with an
air speed meter, giving the speed of the machine in relation to the air.
When flying with the wind the pilot is likely to be confused by the
tremendous ground speed at which his machine is flying. While the
machine may be moved at a fast clip in regard to the earth, it may be
really near the stalling speed. This error is particularly dangerous
when one turns in with the direction of the wind, after flying against
it for some time. The sudden increase in the earth speed, when fully in
with the wind, always creates a sudden desire to throttle down at the
very time when the relative air speed has already been greatly reduced
by the turn. Stalling due to this cause has resulted in many accidents,
and the beginner should always attain an altitude of a least 500 feet
before he tries turning in with a strong wind. An accurate speed
indicator eliminates this danger to a great extent, but it should be
proved that the instrument itself is accurate before too much reliance
is placed on it.
Before the advent of the indicator, pilots were compelled to estimate
the speed by the sense of feel, some depending upon the feel of the wind
pressure on their faces, and others by the relative resistance offered
to the movement of the control surfaces. The sense of "Air feel"
developed by the late Lincoln Beachey was marvelous, for without
instruments he would repeatedly climb nearly into a stall when only 50
feet from the ground, and then recover with his chassis nearly dragging
in the weeds.
*Gliding (Fr. Vol Plan)*. "Gliding" is a descent along an inclined path
without power, and is possible with any aeroplane. By suitably inclining
the wings with the horizontal, gravity is made to produce a forward
propelling component that moves the machine forward at the expense of a
loss in altitude. The angle of the gliding path made with the horizontal
is known as the "Gliding angle," and indicates the efficiency of the
aeroplane, for with machines having very low head resistance the angle
is very "flat," and more nearly approaches the horizontal. The best or
flattest gliding angle is an inherent feature of the aeroplane design,
and this cannot be exceeded by any effort on the part of the pilot. It
is generally expressed in terms of the ratio of the descent to the
forward distance traveled, thus a gliding angle of 10 means that the
aeroplane travels 10 feet horizontally for every foot of descent. Any
angle steeper than the flattest angle can be produced by pushing forward
on the elevator controls, thus depressing the elevator tips.
A very flat gliding angle is a most important feature from the
standpoint of safety, as it determines the extent of the area within
which a landing can be made with a dead engine. If the gliding angle is
taken as 12, and the height is 2000 feet, then the radius of the
circular area in which a landing is possible is 2000 × 12 = 24000 feet,
and the diameter is twice this or 48,000 feet, so that we can land
anywhere within a distance of over 9 miles. If the best gliding angle of
the machine were only 10, this will be reduced to 2000 × 10 = 20000
feet, hence our chance of choosing a safe landing space would be cut
down in proportion. The best gliding angle corresponds to a certain
speed and wing angle, and must be determined by experiment, but in many
machines the adjustment of the weight is such that the machine
automatically picks up the best glide as soon as the motor is cut off,
and needs but little correction by the elevators. Such a machine is
dived slightly when the motor is cut out, and then after a few
oscillations settles down and travels steadily along the proper gliding
path. In trying to improve this performance, the speed indicator and
incidence indicator should be carefully watched so that neither the
stalling angle nor the stalling speed are approached. The best glide
angle corresponds to the best flight speed and will be increased if the
incidence is much below or above the incidence for the most economical
flight speed.
*Vertical Nose Dive*. When the aeroplane is diving vertically, nose
down, the center of pressure movement in some machines may oppose the
elevators, thus making it difficult to straighten out into the
horizontal. If pulling full back on the elevator control does not remedy
matters, the control should then be quickly reversed so that there is a
momentary tendency to throw the machine over on its back. This breaks up
the lock, and when accomplished, the controls should be again pulled
back to bring the machine into the horizontal with the elevators in the
original straightening out position. The momentum swings the machine out
and against the locking position, thus aiding the controls in overcoming
the moment of the C. P.
[Illustration: Typical Gliding Angle Diagram Showing Path Inclination of
Deperdussin Monoplane.]
*Tail Spin (Spinning Nose Dive)*. Spinning is due to side slipping or
stalling, and sooner or later every pilot gets into this position either
through accident or intention. If an accident, it may be due to the
design of the aeroplane through an improper distribution of the vertical
surfaces, or again it may be caused by very steep banking without an
equivalent rate of turning. Incorrect manipulation of the ailerons when
the machine is near stalling speed, or when gliding in a spiral of
gradually decreasing radius, also causes this result. At any rate, the
side slip and stall are the final cause of spin. In "Stunt flying,"
where a spin is desired, one of the quickest methods of getting a spin
is to pull the controls all the way back and push the rudder hard over
in the desired direction of spin with the motor shut off. Another way to
get a spin with lots of "Pep" in it, is pull the stick clear back with
the motor on, and climb until the machine is stalled, then rudder over
hard with the controls still held back. The aeroplane will now fall over
on its side in the direction of the rudder, and assisted by the motor
which has again cut in after the peak of the climb, will give all the
spin that any critic could ask for. After the stall occurs, the motor
should be throttled down for it is likely to strain the plane or even
break it.
There are several ways of coming out of the spin. Probably the best way,
and the one that causes the least loss of altitude, is to keep the
controls pulled back all the way, and rudder in the opposite direction
to the spin (Motor cut out). The rudder will stop the spin, and the
elevator will cause the plane to level out of the dive simultaneously,
but the controls should be put into neutral as soon as leveled out or
there will be another spin started in the opposite direction. A very
common method used by exhibition flyers is first put the controls into
neutral, and rudder opposite to the spin until it stops turning and it
is then put into a nose dive. The straight nose dive can then be easily
corrected by pulling back on the controls until it levels out. This
latter method develops an excessive speed and requires a high altitude.
When the aeroplane is overbanked at normal speed, and the turn is not
correspondingly rapid, the plane will slip down sideways into an "Inside
side slip." The strong upward wind against the side of the body will
turn the nose into a dive, the nose drops, and the tail will then start
to swing around in a circle larger than the circle described by the
nose—the dive continuing. When much below the normal flight speed, or
near the stalling point, the inner ailerons are not as effective when
making a sharp turn for their velocity is much reduced. When fully
depressed, the inner ailerons give very little lift toward righting the
machine but add to the drag and tend to spin the machine around with the
inner tips acting as a center of rotation. The outer ailerons are very
effective and because of the high speed of these tips, there is a strong
banking tendency that eventually will result in side slip and a spin if
the pilot is not experienced. Either the spin due to overbanking, or
that due to low speed may be straightened out according to instructions
already given.
When a turn is attempted at low speed near the ground, the student
generally fails to bank up sufficiently through fear of striking the
ground with the lower wing tip, and therefore gets into an outward side
slip. In the frantic effort to keep the low wing up and off the ground
he depresses the low aileron to the full, thus increasing the drag on
the low side and starting the spin. Very much to his surprise he finds
that this actually drops the low tip further instead of raising it as
the outer tip is now speeding around at a tremendous clip, and the outer
lift is increasing the bank against his will. Given time, and altitude,
the plane will bank up until it stands on end, and in any event a bad
side slip results, and the fun is on. If near the ground as assumed,
either the side slip or the resulting nose dive will soon terminate
matters. The moral to be derived from this experience is to keep up to
speed in making a turn, to maintain a safe altitude, and in case the
speed should fall off, to depress the *outer aileron*. The outer aileron
will resist the spin if depressed, as the drag acts against the spin,
and the bank thus obtained will act against the outer side slip, without
destroying the velocity of the machine as a whole. In turning at
stalling speed, the aileron effects are reversed, and as soon as
straightened out the engine should be opened up so that the speed will
be increased and the landing made as easy as possible.
If the fin and rudder surface is not sufficient for the machine, little
is gained by turning the rudder to an angle greater than 15°, and in
such cases it is much more effective when held parallel to the wind. If
correction has been started before the spin has developed great
rapidity, the rudder can first be turned to check the rotation and then
turned back parallel to the wind. It is always best to shut off the
engine when getting out of a tail spin, especially if the engine
rotation is in the direction of the spin, since the motor torque aids
the spin and acts against the controls. In case of a smash there is no
danger from fire with the engine cut out.
*Stunt Flying*. When the student has had 20 hours or more of solo
flying, and is capable of performing the ordinary maneuvers with
confidence and accuracy, he is in a position to undertake stunt flying
under the directions of a good instructor in a dual control machine.
This tremendously increases the confidence of the student if gone about
in the right way, and in his after flying experience enables him to get
out of tight places that would otherwise often prove impossible. There
is no doubt but what stunt flying has decreased the percentage of
accidents when properly taught, and that Pegoud’s original stunt of
looping the loop has been one of the greatest steps in the advancements
of aeronautics that we have had, if only for the fact that it taught the
flyer that there was no flying attitude so bad but what there was a
solution for it.
*Flying Upside Down*. With the machine on its back, then wings are very
inefficient, and it is impossible to maintain horizontal flight in this
position, and the machine is also very unstable. It should really be
called gliding instead of flying since the aeroplane constantly loses
altitude along an inclined gliding path. The distance that a machine can
be glided in this way depends upon the skill of the pilot, and it will
also be found that upside down flight with a large dihedral is more
difficult than with straight wings. The upside down flight begins with a
glide to gain speed, the path being about 20° with the horizontal, and
this speed gain is imperative since it requires both the power of the
motor and the momentum of the machine to overcome the sharp climb for
the turnover.
[Illustration: Upside Down Glide Diagram, Showing Successive Positions
of Aeroplane.]
After sufficient speed has been attained, the controls are pulled back
for a climb at about a 60° angle, as between (A) and (C), this maneuver
being best performed with the gliding path (C-D) against the wind. With
the control pulled back at (A), the rudder is thrown over sharply in the
desired direction of the turnover, and this will turn the machine over
as indicated by at (B), the machine finally getting on its back at the
peak of the climb (C). With the machine on its back, reverse rudder to
stop the overturning, and when the wings are horizontal, the rudder
should be put in neutral to hold it in this position. At (C) the motor
is shut off, and the glide continued to (D) where it is leveled out by a
backward pull on the controls. This should always be performed at an
altitude well over 2000 feet.
[Illustration: Looping Diagram Showing Successive Positions of Machine.]
*Looping*. This is probably the easiest of all stunts outside of the
spiral glide. It starts with about a 20° glide as at (A) to increase the
velocity (Motor on), and at the beginning of the loop at (B), the
control lever is pulled back slowly. The controls must be pulled back
faster and faster as the plane approaches the top of the loop, a steady
pull producing nearly the correct effect because of the decreasing
elevator resistance as the machine reaches the top of the loop at (C).
At the top, the lever should be clear back and must be held in this
position until at the bottom (D) where the machine leaves the loop along
the inclined path (D-E) At (D), the stick is pushed slowly forward to
neutral, gradually bring the machine into the horizontal. The loops must
always be made when flying into the wind, and the faster they are made
the better, for there is less strain on the frame and speed also
prevents the motor from cutting out at the top of the loop.
[Illustration: A Few Straight Loops and Backward Reverse Loops Performed
by Niles.]
[Illustration: Photograph of Night Looping by Charles Niles. The Machine
was Provided with Railroad Flares which Left the Trace or Path of the
Aeroplane on the the Dry Plate.]
*Immelmann Turn*. This maneuver was originated by the German flyer
Immelmann, and is much used in combat by both the Allied and German
armies, for it subjects the enemy to a maximum field of fire and enables
the machine to make a quick getaway with a single seat machine. With the
enemy machine at (X), and with our machine provided with two machine
guns, it will be seen that the enemy is under the fire of either the
rigid front gun or the pivoted cock-pit guns through nearly
three-quarters of the twisting loop. The pivoted gun which fires over
the top wing is the most effective as it can reach the enemy machine (X)
at (A), (B), (D) and (E), the only blind spots occurring during the
climb from (B) to (D) as indicated by the partly rolled over position at
(C).
[Illustration: Successive Positions in Immelmann Turn, Enemy Machine
Either Being at Y or X.]
It begins with the usual power glide (Motor on) at (A) in order to gain
speed, and at the beginning of the 60° at (B) the elevator controls are
pulled back and the rudder given a quick turn to the extreme position in
the direction of the desired turnover. The rudder action turns the
machine over on its back at the peak of the climb at (D) without the use
of ailerons. At the top (D), the rudder is thrown to the opposite
direction to stop the roll over and is then brought back to neutral to
hold the machine flat on its back. The elevator controls are held back
until the machine comes out of the reverse loop extending from (D) to
(E) and until it leaves on the horizontal at (F). As the object is to
get away quick, the finish along (E-F) should be made with the wind, and
preferably should be started across wind. Motor should be throttled down
from (D) to (E) to prevent coming out with excessive speed.
[Illustration: Positions in "Turn Over," the Machine Continuing to
Rotate in the Same Direction from B to D.]
*The Roll-Over*. The start of this stunt is exactly like the start of
the upside down glide or the Immelmann, while the finish is a sort of
reversed Immelmann, the machine being straightened out without going
around a loop. When at the top of the turnover climb, the rudder is not
reversed and straightened out as in the Immelmann, but a little rudder
is kept on so as to continue the turnover and bring the plane out on the
horizontal. The rudder action is assisted by a slight application of the
aileron while the elevator control is pushed forward after the machine
leaves the peak of the climb. I am indebted to Lieutenant Charles W.
Keene for suggestions on the Immelmann and roll over, and to the late
Lieutenant R. C. Saufly of the U. S. Navy for other items on training.
[Illustration: Flying Upside Down With a Bleriot Monoplane. The Plane is
Far Too Low to Recover Its Normal Position, and as a Result, the Glide
Ended Fatally.]
[Illustration: An Aeroplane Equipped with the Light Boat Hull Shown, in
this Figure is Known as a "Flying Boat". It Differs from the "Seaplane,"
as the Floats or Pontoons of the Latter Do Not Enclose the Passenger and
Pilot.]
CHAPTER XX. AERONAUTICAL MOTORS.
*General Notes*. It is assumed that the reader understands the
principles of the automobile motor and its accessories, for a minute
description of gas-engine principles does not fall within the scope of
this book. If more information is desired on this subject, the reader is
referred to the author’s "Practical Handbook of Gas, Oil and Steam
Engines." Only those features peculiar to aeronautic motors will be
discussed in this chapter.
*Aeronautic Requirements*. The principal requirements of an aeronautic
motor are light weight, low oil and fuel consumption, reliability and
compactness. The outline as viewed from the shaft end is also very
important, for the motor must be mounted in a narrow streamline body.
The compression pressures are much higher than those employed on auto
motors, and the speed is generally lower. With one or two exceptions the
four-stroke cycle has been universally adopted.
Aeronautic service is a severe test for the motor. From the start to the
finish of a flight, the aeroplane motor is on a steady grind, loaded at
least to 75 per cent of its rated power. The foundations are light and
yielding and the air density varies rapidly with changes in the
altitude. As the fuel and oil require an expenditure of power for their
support, the fuel consumption becomes of great importance, especially in
long flights. Because of the heavy normal load the lubricating system
must be as nearly perfect as it is possible to make it.
A motor car runs normally at from 10 to 25 per cent of its rated
horsepower, while the aero motor may develop as high as 75 per cent to
100 per cent for hours at a time. A car engine of 672 cubic inches
displacement is rated at 65 horsepower, while the same size aero engine
has a rating of 154. On the basis of normal output, this ratio is about
7 to 1, and taking the weight of the aero motor as one-half that of the
auto type, the true output ratio becomes 14 to 1. Up to the time of a
complete overhaul (50 hours), and at 100 miles per hour, the average
distance traveled by the aero motor is 5000 miles. The equivalent motor
car mileage is 25,000, and the duration is about 1000 hours. This
suggests the necessity for improved materials of construction. Even on
the present aeronautic motors the fiber stress in the crank-shaft ranges
from 120,000 to 140,000 pounds per square inch against the 80,000-pound
stress used in auto shafts. The crank case of an aeronautic motor must
be particularly rigid to withstand the stresses due to the light
mounting, and this demands a higher grade metal than that ordinarily
used with automobiles. Unlike the car engine, quality comes first and
price is a secondary consideration.
*Cooling Systems*. Both the air and water cooling system is used, the
former for light fast aeroplanes such as speed scouts, and the latter
for the larger and more heavily powered machines. Even in some types of
speed scouts the air-cooled motor has been displaced by the
water-cooled, owing to the fact that the air-cooler cannot be built
satisfactorily for outputs much greater than 110 horsepower. By
increasing the revolutions of the stationary water-cooled type an
increase in power may be had with the same cylinders, but in the case of
the rotary air-cooled type the speed is limited by the centrifugal
forces acting on the cylinders.
[Illustration: A 6-Cylinder Hall-Scott Motor Installed in a Martin
Biplane.]
While the weight of the radiator, water and piping increase the weight
of the water-cooled motor very considerably, the total weight is not
excessive. When the fuel is considered, the total weight is below that
of the rotary when long flights are attempted. The radiator and water
add complication and are a source of danger. The radiators increase the
head resistance and add very considerably to the maintenance cost.
[Illustration: A Motor Installation in a Pusher Type Biplane, Showing
the Motor at the Rear and the Double Radiator Sections Over the Body.]
Each type of cooling has its limitations, and it is hoped that an
improvement in cooling may be had in the near future. This system should
primarily reduce the size and resistance of the power plant, and if
possible the weight, although the latter is a secondary consideration.
At present the cooling system prevents even an approach to the true
streamline form of the body.
*Propeller Speed*. For the best results, the propeller speed should not
exceed 1200 revolutions per minute, and for structural reasons this is
generally limited to 1500 R. P. M. This at once puts a limiting value on
the output of a given size engine unless a gear down arrangement is
used. It should be understood, between certain limits, that the power
output increases roughly as the speed. With direct drive arrangements in
which the propeller is mounted directly on the end of the engine shaft,
the motor revs. are necessarily the propeller revs., and the only way of
increasing the speed is by increasing the length of the stroke or by
gearing down. An increase in stroke adds rapidly to the weight by
increasing the cylinder length, length of connecting rod, length of
crank throws, etc.
*Horsepower Rating*. At present there are many methods of calculating
the horsepower of gasoline engines. Formula applying to auto or boat
motors does not apply to flight conditions, for the aero motor is
essentially a high compression type and has a greater output per unit of
displacement. It is not practical to give the rated horsepower as the
maximum output possible under ideal conditions, for this would give no
idea as to the practical capabilities except by long tedious
calculation. The brake horsepower would give no overload capacity at a
fixed propeller speed, and the conditions are entirely different from
those regulating the rating of auto motors. The latter can be forced up
to the wrecking speed, or many times the normal automobile speed of 30
miles per hour.
As aero engines are generally well kept up, and well tuned at all times,
the rated horsepower may be taken from 15 to 20 per cent below that of
the maximum brake horsepower. In geared-down motors, the gear efficiency
is still to be considered. The question of the quality of the mixture,
and barometric pressure, also enter into the problem whether the power
is rated on the maximum obtained with a rich mixture, or is calculated
from the output at the maximum efficiency. A writer in "Aviation"
suggests that the rated horsepower be taken as 95 per cent of the power
developed at a point midway between the maximum output, and the output
at the greatest efficiency. Barometric pressure to be 30 inches and the
revolutions 1200.
Owing to the great diversity in the bore-stroke ratio, a power formula
must include the bore and stroke. This makes the S.A.E. formula for auto
motors impossible. A formula is proposed by a writer in "Aviation." The
writer has checked this up with the published performance of several
well-known aeronautical motors.
H = B²SNR/12,500 Where B = bore in inches, S = stroke in inches, N =
number of cylinders, R = Crankshaft revolutions per minute, and H =
rated horsepower. This applies only to the four-stroke cycle type.
*Power and Altitude*. The power drops off rapidly with an increase in
altitude unless corrections are made for compression and mixture. With
constant volume, the decreased density causes decreased compression. As
the weight of air taken in per stroke is reduced, this also reduces the
amount of fuel that can be burned per stroke. By holding the compression
constant through adjustment of the clearance or valve motion, a fairly
constant output can be had through a wide range of altitudes.
A compression of 115 pounds per square inch (commonly used) is difficult
to handle with a light construction, but this pressure must be obtained
if the output is to be kept within practical limits. Engines having a
compression ratio of as high as 6 are running satisfactorily at sea
level, this ratio giving a mean effective working pressure of 134 pounds
per square inch. With this ratio the engine cannot be used with full
open throttle at sea level for more than 10 or 15 minutes without
causing damage to the shaft, bearing and valves. At about 10,000 feet
the compression is normal.
At great altitudes carburetion has become a great problem, and as aerial
battles have already taken place at elevations of 20,000 feet, it is
quite possible that future motors will be equipped with some device that
will force a measured fuel charge into the cylinders. The air necessary
for the combustion will also have to be pumped in by some means.
*Weight Per Horsepower*. The weight per horsepower of the engine is a
very loose term since so much depends upon the equipment included in the
weight. As many as 20 items may be considered as being in the doubtful
list, and among these are the radiator water, piping, mounting,
propeller hub, oil in sump, wiring, self-starter, etc. The only true
unit weight is that obtained by taking the plant complete (ready to
run), with the cooling system, gasoline for an hour’s flight, and the
oil. The weight of the bare engine signifies nothing. The weights of the
various items used on well known motors are given in a table under the
chapter "Weight Calculations." While the bare weight of a certain engine
may be very low per brake horsepower, an excessive fuel consumption will
often run the effective weight up and over that of a type in which the
bare weight is far greater. The weight of the engine per horsepower,
including the magneto and carbureter, will run from 2.2 to 5.0 pounds,
according to the type.
[Illustration: Two Examples of Cowls Used Over Rotary Cylinder Motors
(Air Cooled).]
*Fuel Consumption*. The fuel consumption of water-cooled motors varies
from 0.48 to 0.65 pounds per horsepower hour, an average of 0.6 being
safe. The fuel consumption of a rotary air-cooled motor will range from
0.6 to 0.75. The oil consumption varies from 0.18 gallons per horsepower
in the air-cooled type to 0.035 with the water-cooled stationary motor.
*Radiators*. Owing to extremes in the temperature of the air at
different altitudes, the radiating surface should be divided into
sections so that a constant cooling effect can be obtained by varying
the effective surface of the radiator. The temperature can also be
controlled by an automatically regulated by-pass which short circuits a
part of the radiator water at low temperatures. Constant water
temperature has much to do with the efficiency and general operation of
the motor, and there will be only one temperature at which the best
results can be obtained.
[Illustration: Typical Radiators. A) Side or Top Type.]
[Illustration: Typical Radiators. (B) Front Type.]
Hunsaker finds that 0.83 square feet of actual cooling surface per
horsepower is correct at 60 M. P. H., while others give a value of about
100 square foot under similar conditions. The front or projected area
varies with the thickness of the radiator, the thicknesses varying from
2 to 5 inches. The Livingston radiator gives a cooling surface of 50
square inches per square inch of front surface. The total cooling effect
depends upon the speed, the location in regard to the slipstream, and
the position on the body. A radiator maker should always be consulted
when making the final calculations. See Chapter XVI.
*Fuel Tanks and Piping*. The fuel tanks may be of copper, aluminum or
tin-coated steel, and all joints should be welded or riveted. Never
depend upon solder, as such joints soon open through the vibration of
the engine. Gasoline should not come into contact with steel, nor the
zinc used on galvanized iron. Splash plates are provided to keep the
fluid from surging back and forth while in flight. All gas should be
supplied to the engine through a filter or strainer placed in the main
gas line. The valves in the fuel lines should be provided with
stopcocks, so arranged that they can be closed from the pilot’s seat.
In general, the carbureters should be fed by gravity from an overhead
service tank, this tank being supplied from the main reservoir by air
pressure or a gasoline pump. The air can be compressed by a pump on the
engine or by a paddle driven pump operated by the airstream, and as a
rule the latter is preferable, as it can be operated with the aeroplane
gliding and with the engine dead. Air pressure systems are likely to
fail through leaks, while with a good gasoline pump conditions are much
more positive. The gravity service tank should be located so that it
will feed correctly with the aeroplane tilted at least 30 degrees from
the horizontal.
[Illustration: Two Views of the "Monosoupape" Gnome Rotary Cylinder
Motor. This Motor Has 9 Cylinders Arranged Radially Around the
Crankshaft and Develops 100 Horsepower. The Cylinders Are Air Cooled.]
The gasoline piping should be at least 5/16 inch inside diameter, and
should be most securely connected and supported against vibration. To
guard against crystallization at the point of attachment, special
flexible rubber hose is generally used. This must be hose made specially
for this purpose, as ordinary rubber hose is soon dissolved or rotted by
gasoline and oil. Air pockets must be avoided at every point in the fuel
and oil system.
[Illustration: Hall-Scott "Big Six" Aeronautical Motor of the Vertical
Water-cooled Type. 125 Horsepower.]
*Rotating Cylinder Motors*. The first rotating cylinder motor in use was
the American Adams-Farwell, a type that was soon followed by the better
known French "Gnome." Other motors of this type are the Clerget,
LeRhone, Gyro and Obereusel. They are all of the air-cooled type—cooled
partly by the revolution of the cylinders about the crank-shaft, and
partly by the propeller slipstream. While the pistons slide through the
cylinder bore, the rotating cylinder motor is not truly a reciprocating
type, as the pistons do not move back and forth in regard to the crank
shaft. The cylinders revolve about the crank shaft as a center, while
the pistons and connecting rods revolve about the crank pin, the
difference in the pivot point causing relative, but not actual,
reciprocation.
[Illustration: Hall-Scott 4-Cylinder Vertical Water-cooled Motor. 80-90
Horsepower.]
The original Gnome motor drew in the charge through an inlet valve in
the piston head. The gas passed from the mixer, through the hollow
crank-shaft, and then into the crank-case. The exhaust valve was in the
cylinder head. This valve arrangement was not entirely satisfactory, and
the company developed the "Monosoupape" or "Single valve" type. The 100
H. P. Monosoupape Gnome has 9 cylinders, 4.3" x 5.9". The total weight
is 272 pounds and the unit weight is 2.72 pounds per horsepower. It
operates on the four-stroke cycle principle. The gas consumption is 12
gals. per hour, and it uses 2.4 gals. of castor oil. The cylinders and
cooling fins are machined from a solid steel forging, weighing 88
pounds. The finished cylinder weighs 5.5 pounds after machining. The
walls are very thin, probably about 1/16 inch, but they stand up well
under service conditions.
[Illustration: Sturtevant "V" Type 8-Cylinder Water Cooled Aeronautical
Motor. This Motor Is Provided With a Reduction Gear Shown at the Rear of
the Crankcase.]
Assuming the piston to be on the compression stroke, the ignition will
occur from 15° to 20° before the top dead center. Moving down on the
working stroke, and at 85° from top dead center, the exhaust valve
begins to open, and the exhaust continues until the piston returns to
the upper dead center. With the valve still open, pure air now begins to
enter through the exhaust valve and continues to flow until the valve
closes at 65° below the bottom center. Still descending, the piston
forms a partial vacuum in the cylinder, until at 2° before the lower
center the piston opens the ports and a very rich mixture is drawn in
from the crank case. This rich mixture is diluted to the proper density
by the air already in the cylinder, and forms a combustible gas. The
upward movement of the piston on the compression stroke closes the ports
and compression begins. The mixture enters the crank case through a
hollow shaft, with the fuel jets near the crank throws. A timed fuel
pump injects the fuel at the proper intervals.
[Illustration: Dusenberg 4-Cylinder Vertical Water Cooled Motor With
Reduction Gear. Four Valves Are Used Per Cylinder. Note Peculiar Valve
Motion.]
*Curtiss Motors*. The Curtiss motors are of the water-cooled "V" type,
with 6 to 8 cylinders per row. These are probably the best known motors
in America and are the result of years of development, as Curtiss was
the first to manufacture aero motors on a practical scale.
[Illustration: Curtiss Type OX-5 Eight Cylinder Aeroplane Motor]
*Hall-Scott Motors*. These motors are made by one of the pioneer
aeronautical motor builders, and have met with great favor. They are of
the vertical water-cooled type, and with the exception of minor details
and weight are very similar in external appearance to the automobile
motor. Four and 6-cylinder types are built.
*Sturtevant Motors*. These are of the "V" water-cooled type, and are
provided with or without a reducing gear. At least one model is provided
with lined aluminum cylinders.
*Dusenberg Motor*. This is a four-cylinder, water-cooled, vertical motor
with a very peculiar valve motion. The valves are operated by long
levers extending from the camshaft. Two inlet, and two exhaust valves,
are used per cylinder. The motor is generally furnished with a reducing
gear.
*Roberts Motor*. This is a solitary example of the two-stroke cycle
type, and has been used for many years. It is simple and compact, and is
noteworthy for the simplicity of its oiling system. The oil is mixed
with the gasoline, and is fed through the carbureter. This is one of the
many advantages of a two-stroke cycle motor.
*Table of Aeronautical Motors*. The following table will give an idea as
to the general dimensions of American aeronautical motors:
*The Liberty Motor*. The necessity of speed and quantity in the
production of aeronautical motors after the declaration of war caused
the Government to seriously consider the design of a highly standardized
motor. This idea was further developed in a conference with
representatives of the French and British missions on May 28, 1917, and
was then submitted in the form of sketches at a joint meeting of our
allies, the Aircraft Production Board, and the Joint Army and Navy
Technical Board. The speed with which the work was pushed is remarkable,
for on July 3rd, the first model of the eight cylinder type was
delivered to the Bureau of Standards. Work was then concentrated on the
12 cylinder model, and one of the experimental engines passed the 50
hour test August 25, 1917.
It is of the "V" type with the cylinder blocks at an angle of 45 degrees
instead of 60 degrees as in the majority of 12 cylinder "V" motors. This
makes the motor much narrower and more suitable for installation in the
fuselage, and in this respect is similar to the arrangement of the old
Packard aviation motor. It has the additional advantages of
strengthening the crank case. The bore and stroke is 5" x 7" as in the
Hall-Scott models A-5 and A-7. The cylinders combine the leading
features of the German Mercedes, the English Rolls-Royce,
Lorraine-Dietrich, and Isotta-Fraschini. Steel cylinder walls are used
with pressed steel water jackets, the latter being applied by means of a
method developed by the Packard Company. The valve cages are drop
forgings, welded to the cylinder heads.
The camshaft and valve gear are above the cylinder head as in the
Mercedes, but the lubrication of the parts was improved upon by the
Packard Company.
The crankshaft follows standard 12 cylinder practice except as to the
oiling system, the latter following German practice rather closely. The
first system used one pump to keep the crankcase empty delivering the
oil to an outside reservoir. A second pump took the oil from the
reservoir and delivered it to the main crankshaft bearings under
pressure. The overflow from the main bearings traveled out over the face
of the crank throw cheeks to a "Scupper," which collected the excess for
crank pin lubrication. In the present system, a similar general method
is followed except that the pressure oil is not only fed to the main
crankshaft bearings, but also through holes in the crank cheeks to the
crank pins instead of by the former scupper feed.
A special Zenith carburetor is used, that is particularly adapted to the
Liberty motor. A Delco ignition system of special form is installed to
meet the peculiar cylinder block angle of 45 degrees. This ignition is
of the electric generator type and magnetos are not used.
Several American records have been broken by the new motor, and it is
reported to have given very satisfactory service, but full details of
the performance are difficult to obtain owing to the strict censorship
maintained in regard to things aeronautic. The motor is particularly
well adapted to heavy bombing and reconnaissance type machines, or for
heavy duty. It is reported that the use of the motor has been
discontinued on speed scouts, although further developments along this
line may not have been reported.
The following gives the principal characteristics of the Liberty motor,
issued by the National Advisory Committee for Aeronautics.
────────────────────────────────────────────────────────────────────
Year Horse-power Weight Weight Per Gasoline H. P.
(Model) Pounds H. P. Hour
────────────────────────────────────────────────────────────────────
1917 400 801 2.00 0.50
────────────────────────────────────────────────────────────────────
1918 432 808 1.90 0.48
────────────────────────────────────────────────────────────────────
1918 450 825 1.80 0.46
────────────────────────────────────────────────────────────────────
The motors listed are all 12 cylinder models, and the output and unit
weights are based on a crank-shaft speed of 1800 R. P. M. The 5" x 7"
bore and stroke give an output of 37.5 horsepower per cylinder in the
latest model. In 1917, the Liberty motor was 65 per cent more powerful,
and 28 per cent lighter, than the average stock motor in service during
that year.
CHAPTER XXI. GLOSSARY OF AERONAUTICAL WORDS.
In the following list are the most common of the aeronautical words and
phrases. Many of these words are of French origin, and in such cases are
marked "Fr." In cases of English words, the French equivalents follow in
parentheses. When a French word or term is given it is in italics,
unless it is in common use in this country. Words marked (*) are the
revisions adopted by the National Advisory Board of Aeronautics at
Washington, D. C., and include the term "Airplane," which was intended
to supplant the more common "Aeroplane." These revisions have not met
with universal adoption, for the older words are too well established to
admit of change.
A
ABSOLUTE UNITS.
Units given in terms of mass. For expression in terms of pounds
(Gravitational units) they must be multiplied by some factor
involving the value of gravitation. Thus, to convert units of
mass into pounds, the mass must be multiplied by the value of
gravitation, 32.16 being the average figure taken for this
quantity. To convert the absolute lift factors given by the
N.P.L. into pounds per square foot per mile per hour, multiply
the absolute value by O.005IV².
ABSOLUTE ZERO.
The temperature at which heat ceases to exist. This is 461
degrees below the Fahrenheit zero, or 273 degrees below the
Centigrade zero.
ACCELERATE.
To increase in speed.
_ACIER_ (Fr.)
Steel.
AERODONETICS.
A word originated by Lanchester to denote the science of
stability.
AEROCURVE.
See AEROFOIL.
AEROFOIL.*
A thin wing-like structure designed to obtain lift by the
reaction of moving air upon its surfaces.
AERODYNAMICS.
A science investigating the forces produced by a stream of air
acting upon a surface.
AERODYNAMIC RESISTANCE.
The resistance caused by turbulence or eddies.
AERODROME.
A flying field. This word was also used by Langley to describe an
aeroplane.
AEROSTAT.
A lighter than air machine.
AEROSTATICS.
The science of lighter than air machines, or devices sustained by
flotation.
AEROPLANE. (Fr. L’Avion.)
A heavier than air craft sustained by fixed wing surfaces driven
through the air at the same velocity as the body of the machine.
Auxiliary surfaces are provided for stabilizing, steering, and
for producing changes in the altitude. The landing gear may be
suitable for either land or water, although in the latter case it
is generally known as a "Seaplane." The Committee equivalent is
"Airplane."
AILERON.*
A movable auxiliary surface used in maintaining lateral balance.
_AILE_ (Fr.)
Wing.
AIRBOAT.
An aeroplane provided with a light boat hull in which the pilot
and passenger are enclosed.
AIRCRAFT.*
Any form of craft designed for the navigation of the air. This
includes aeroplanes, balloons, dirigibles, helicopters,
ornithopters, etc.
AIRPLANE.*
See aeroplane.
AIRSHIP.
A lighter than air craft provided with means of propulsion.
AIR-SCREW.
See PROPELLER.
ALTIMETER.*
An instrument used for determining the height of aircraft above
the earth.
ALTITUDE.
Height of aircraft above sea level—generally given in feet.
AMPHIBIAN.
An aeroplane equipped with landing gear for both land and water.
ANEMOMETER.*
An instrument for measuring the velocity of the wind.
ANGLE OF INCIDENCE.
The angle made by a surface or body with an air stream. In the
case of curved wings, the angle is measured from the chord of the
curve.
ANGLE OF ATTACK.*
See Angle of Incidence.
ANGLE OF ENTRY.
The angle made with the chord of a wing section by a line drawn
tangent to the upper curved face, and at the front edge.
ANGLE OF TRAIL.
The angle made by a line drawn tangent to the upper surface at
the trailing edge.
ANEMOGRAPH.
An instrument used for graphically recording the velocity of air
currents.
APTEROID ASPECT.
A wing is in apteroid aspect when the narrow edge is toward the
wind.
_ARETIER ARRIERE_ (Fr.)
See Trailing Edge.
_ARETIER AVANT_ (Fr.)
See Leading Edge.
_ARBRE_ (Fr.)
Shaft.
ASPECT RATIO.*
The ratio of the wing span to the chord (length divided by
width).
_ATTERRISSAGE_ (Fr.)
Landing Gear.
_ATTACHES_ (Fr.)
Fastenings.
AUXILIARY SURFACE.
A surface used for stability or for the control of the aeroplane.
AVIAPHONE.
An electric system of communication between the passenger and
pilot.
_AVION_ (Fr.)
See Aeroplane.
AXIS OF PITCH.
The axis taken parallel to the length of the wings, and through
the center of gravity. This is sometimes called the "Y" axis.
AXIS OF ROLL.
An axis passing fore and aft parallel to the center line of the
propeller. This axis is sometimes called the "X" axis.
AXIS OF YAW.
A vertical axis, passing through the center of gravity, around
which the machine swings when being steered in a horizontal
direction under the action of the rudder. This is the "Z" axis.
B
BALSA WOOD.
A very light wood obtained from South America. It is lighter than
cork.
BALLOON.*
A form of aircraft of the lighter than air type comprising a gas
bag and car. It is not provided with a power plant, and depends
upon the buoyancy of the gas for its sustentation. A balloon
restrained from free flight by means of a cable is known as a
"Captive balloon." A kite balloon is an elongated form of captive
balloon, fitted with a tail to keep it headed into the wind, and
is inclined at an angle so that the wind aids in increasing the
lift of the gas.
BALLONET.*
A small air balloon within the main gas bag of a balloon or
dirigible used for controlling the ascent or descent, and for
keeping the fabric of the outer envelope taut when the pressure
of the gas is reduced. The ballonet is kept inflated with air at
the required pressure, the air being controlled by a valve or by
regulating the speed of the blower.
BANK.* To incline the wings laterally when making a turn so that a
portion of the lift force will be opposed to the centrifugal
force.
BAROGRAPH.* (Fr. _Barographe_.)
An instrument used for recording pressure variations in the
atmosphere. The paper charts on which the records are made are
used for determining the altitude of aircraft.
BAROMETER.
An instrument used for measuring variations in the atmospheric
pressure, but is not provided with a recording mechanism as in
the case of the barograph.
BEAUME.
A scale of density or a hydrometer unit used in measuring the
density of fluids. On the Beaume scale water is 10.00, while on
the "Specific gravity" scale water is 1.00. The Beaume scale is
generally used for gasoline and oils.
BENDING-MOMENT.
The moment or "Leverage" that tends to bend a beam.
_BEQUILLE_ (Fr.)
Tail Skid.
_BERCEAU de MOTEUR_ (Fr.)
Engine Bed.
BIAS LAID FABRIC.
Fabric laid on the wing structure with the seams at an angle with
the ribs.
_BIPLACE_ (Fr.)
Two Seater.
BIPLANE.* (Fr. _Biplan_.)
An aeroplane with two superposed lifting surfaces.
BODY. See FUSELAGE.
BODY RAILS.
See LONGERONS.
_BOIS_ (Fr.)
Wood.
_BOIS CREUS_ (Fr.)
Hollow wood construction.
BOMBER.
An aeroplane used for bombing operations.
BOOM. The fore and aft beams running from the wings to the tail in a
pusher type biplane.
_BORD de ATTACQUE_ (Fr.)
Entering or leading edge.
_BORD de SORTIE_ (Fr.)
Trailing edge.
BLADE, PROPELLER.
(Fr. _Pale-Helice_.)
_BOULON_ (Fr.)
Bolt.
_BOUSSOLE_ (Fr.)
Compass.
_BRAS de AILE_ (Fr.)
Wing Spar.
_BREVET_ (Fr.)
Flying permit or license.
BRAKES, AIR.
Small adjustable flaps used in increasing the head resistance
during a landing, thus decreasing the speed.
BURBLE POINT.
The angle at which the lift of a wing section reaches a maximum.
BUOYANCY.
The static force due to a difference in density. The difference
in density between the gas in a balloon envelope and the outside
air determines the sustaining or buoyant force of a balloon.
BUS. A slow fairly stable aeroplane used in training schools.
C
CABLE. (Fr. _Cable_.)
A wire rope built up of a number of small strands.
_CABRE’_. * (Fr.)
A flying attitude in which the angle of incidence is larger than
normal with the tail well down.
CAMBER.*
The convexity, or rise of a lifting surface, measured from the
chord of the curve. It is usually given as the ratio of the
maximum height of the curve to the length of the chord. Top
camber refers to the upper surface, and bottom camber to the
lower surface.
_CABANE_ (Fr.)
The center struts rising from the top of the body to the upper
wing, or the short struts used for the bracing of the overhanging
portions of a biplane wing. Usually cabane denotes the center
cell struts.
_CANARD_ (Fr.)
A machine in which the elevator and stabilizer are in front. The
canard type flies "Tail first."
_CAPOT_ (Fr.)
Cowl or motor hood.
CAPTIVE BALLOON.*
See Balloon.
CAPACITY.*
The lifting capacity is the maximum flying load of an aircraft.
The carrying capacity (live load) is the excess of the lifting
capacity over the dead weight of the aeroplane, the latter
including the structure, power plant and essential accessories.
_CARLINGUE_ (Fr.)
Cock-pit.
CATA-HEDRAL.
A negative dihedral, or wing arrangement, where the wing tips are
lower than the center portion.
CATAPULT.
A device for launching an aeroplane from the deck of a ship or
other limited space. The first Wright machines were launched with
a catapult.
CELL. (Fr. _Cellule_.)
The space included between adjacent struts of a biplane. The
space between the center struts is the "Center Cell."
CEILING.
The maximum altitude to which an aeroplane can ascend.
_CEINTURE de SURETE_ (Fr.)
Safety Belt.
CENTER OF PRESSURE (C. P.)*
The point of application of the resultant of all aerodynamic
forces on an aeroplane wing. If the wing is supported at the
center of pressure it will be in equilibrium.
CENTER OF GRAVITY.
The point at which an aeroplane will balance when freely
suspended.
CENTER OF BUOYANCY.*
The point at which the resultant of all the buoyant forces act.
_CHARNIERE_ (Fr.)
Hinge.
CHASSIS.
The landing wheels and their frame. This is also called the
"Landing gear" in English, or the "Train de Atterrissage" in
French. The chassis carries the load when resting on the ground
or when running over the surface.
CHORD.*
This has two meanings. The chord is the width of a wing or its
shortest dimension. The chord is also the straight line drawn
across the leading and trailing edges of a wing section.
CHASER. (Fr. _Avion de Chasse_.)
A small, fast machine used in scouting or fighting. This type is
also known as a "Speed scout."
_CHARA-A-BANC_ (Fr.)
A two seater aeroplane in which the pilot and passenger are
seated side by side.
_CLOCHE_ (Fr.)
A type of control column used on the old Type XI Bleriot.
COCK-PIT.
The part of the body occupied by the pilot or passenger. The
openings in the body cut for entrance and exit are the "Cock-pit
Openings."
_COMMANDES A PONT_ (Fr.)
Control bridge or Deperdussin yoke.
_COMPTE TOURS_ (Fr.)
Tachometer or speed indicator.
COMPONENTS.
The individual forces that make up a total resultant force.
CONTROLS.* (Fr. _Commandes_.)
The complete system used for steering, elevating, balancing, and
speed regulation. When controls are operated by hand they are
known as "Manual Controls."
CONTROL BRIDGE. (Fr. _Commandes A Pont_.)
The "U" shaped lever used with the Deperdussin control system.
Sometimes known as the "Yoke."
CONTROL STICK. (Fr. _Manche A Balai_.)
A simple control lever capable of being moved in four directions
for elevation, depression and lateral balance.
CONTROL SURFACES.
The adjustable surfaces used for directing and balancing
aircraft. On an aeroplane these are represented by the rudder,
elevator, and ailerons.
_CONTREPLAQUE_ (Fr.)
Three-ply wood.
_CORDE_ (Fr.)
Cord or wire.
_CORD A PIANO_ (Fr.)
Piano or solid hard wire.
CORD WINDING. (Fr. _Transfil_.)
A winding wrapped around wooden struts to prevent splintering or
complete fracture.
_COSSE_ (Fr.)
Thimble for cable connections.
_COUSS IN_ (Fr.)
Cushion.
COVERING, WING. (Fr. _Entoilage_.)
The fabric used in covering the wing structure.
CRITICAL ANGLE.*
The angle of attack or incidence at which the lift is a maximum.
COWL. (Fr. _Capot_.)
The metal cover surrounding a rotary cylinder motor.
_CROISILLONS_ (Fr.)
Bracing wires.
D
DAMPING.
The reduction of oscillation or vibration by the resistance of
the stabilizing surfaces.
DEAD LOAD.
The weight of the structure, power plant, and essential
accessories.
DEAD WATER.
The wake directly in the rear of a moving body or surface.
_De CHASSE_ (Fr.)
See CHASER.
DECALAGE.*
The difference in the angle of incidence between the upper and
lower wings of a biplane.
DEMOISELLE TYPE.
A small monoplane type developed by Santos Dumont.
DENSITY.
The specific weight, or the weight per cubic foot.
_DIEDRE_ (Fr.)
Dihedral angle.
_DERIVE_ (Fr.)
Fin.
DIHEDRAL ANGLE. (Fr. _Diedre_.)
When the tips of the wing are higher than at the center, the two
wing halves form an angle. The included angle between the two
halves, taken above the surface, is known as the "Dihedral
angle."
DIPPING FRONT EDGE.
A wing section in which the leading edge is well bent down below
the rest of the lower surface.
DIRIGIBLE.*
A lighter than air craft in which sustentation is provided by a
gas bag. It differs from a balloon in having a power plant, and
is thus capable of flying in any desired direction regardless of
the wind.
DISC AREA OF PROPELLER.*
The total area of the disc swept out by the propeller tips.
DISCONTINUITY.
Interruption in direction, or breaks in stream line flow. A body
causing eddies or turbulence causes "Discontinuous flow." The
surface separating the eddies and the continuous stream is called
a "Surface of Discontinuity."
DISPLACEMENT.
The volume or space occupied by a floating body.
DOUBLE CAMBER.
A wing section in which both the top and bottom surfaces are
given a convex camber or curvature.
DOPE. (Fr. _Enduit_.)
A solution used for protecting and stretching the wing fabric.
DRAG. The resistance offered to the forward motion of a surface or body
moving through the air. As defined by the Advisory Committee this
is the total resistance offered by the craft and includes both
the resistance of the wings and body. This conception is
confusing, hence the author has considered drag as being the
forward resistance of the wings alone. The resistance of the
structure is simply called the "Head resistance," and the sum of
the resistances is the "Total resistance." This nomenclature was
in existence before the Advisory Board proposed their definition.
DRIFT. As defined by the Advisory Board, the drift is the horizontal
resistance offered by the wings alone. This is confusing since
previous works defined "Drift" as the amount by which an aircraft
was driven out of its normal path by wind gusts. According to
usage, "Drift" is the sidewise deviation from the normal flight
path.
DRAG WIRES.
The bracing wires used for resisting the drag stresses set up in
the wing.
DRIFT INDICATOR.
An instrument for indicating the amount by which an aircraft is
blown out of its path by side winds.
DUAL CONTROL.
A double system of control that can be operated both by the pilot
and passenger.
DUTCH ROLL.
A combined side roll and fore and aft pitch. The machine rolls
from side to side in combination with an up and down motion of
the nose.
DYNAMIC PRESSURE.
The pressure due to the impact of an air stream.
E
ECCENTRIC LOAD.
A load acting to one side of the center line of a beam or strut.
_ECOLE_ (Fr.)
School.
_ECROU_ (Fr.)
Nut.
EDDY. An irregularly moving mass of air caused by the breaking up of a
continuous air stream, or by "Discontinuity."
EFFICIENCY.
The efficiency of a lifting surface is generally expressed by the
ratio of the lift to the drag, or the "Lift-drag ratio." The
efficiency of a propeller is the ratio of the work usefully
applied to the air stream in regard to the power supplied to the
propeller.
ELEVATOR.*
The hinged horizontal tail surface used for maintaining
longitudinal equilibrium and for ascent or descent.
_EMPENNAGE_ (Fr.)
The group of tail surfaces, including the elevator and
stabilizer.
_ENDUIT_ (Fr.)
Dope.
ENGINE ROTATION.
According to the Advisory Board, an engine is turning in
right-hand rotation when the output shaft stub is turning
anti-clockwise.
ENGINE BEARERS (BED). (Fr. _Berceau du Moteur_).
The timbers or fuselage members upon which the engine is
fastened.
ENGINE SPIDER or BRACKET. (Fr. _Arraignee Support de Moteur_.)
A perforated metal support for a rotary cylinder motor.
ENTERING EDGE. (Fr. _Bord D’Attaque or Aretier Avant_.)
The front edge, or air engaging edge, of an aerofoil or lifting
surface. It is also called the "Leading Edge."
_ENTOILAGE_ (Fr.)
Wing fabric or covering.
ENVELOPE.
The gas bag of a balloon or dirigible.
_ESSIEU_ (Fr.)
Axle.
_ENVERGURE_ (Fr.)
Wing span.
EXPANDING PITCH.
A system in which the pitch increases or "expands" towards the
tips of the propeller.
F
FABRIC, WING. (Fr. _Entoilage_.)
The cloth used for covering the wing and control surface
structures.
FAIRING. (Fr. _Fusele_.)
Wood coverings used to streamline steel struts or other
structural members.
FERRULES.
Sheet steel caps used for the ends of the interplane Struts.
FIN. (Fr. _Derive_.)
A fixed vertical stabilizing surface used for damping out
horizontal vibration and oscillations.
FINENESS RATIO.
The ratio of the maximum length to the width of a streamline
body.
FITTINGS. (Fr. _Ferrures, Godets_.)
The metal parts used for making connections between the
structural parts of an aeroplane.
FIXED TAIL.
(See STABILIZER.)
FLACCID BLADE PROPELLER.
A propeller having a cloth covered frame work on which the fabric
is free to adjust itself to the air pressure.
FLAPS, ELEVATOR. (Fr. _Volets de Profondeur_.)
See ELEVATOR.
FLEXIBLE SHAFT. (Fr. _Transmission flexible_.)
Used for tachometer drive.
FLOORING.
(Fr. _Plancher_.)
_FLASQUE D’HELICE_ (Fr.)
Propeller flange.
FOOT LEVER. (Fr. _Palomnier_.)
The foot lever generally used to operate the rudder.
FORMERS.
Supports used in giving a certain outline to the fuselage. The
formers are attached to the fuselage frame and in turn support
small stringers on which the fabric is fastened.
FRICTIONAL WAKE.
The following current of air in the rear of a moving body or
surface. Because of the friction, a portion of the air is drawn
in the direction of the motion.
FUSELAGE.
A structure, usually enclosed, which contains and streamlines the
power plant, passengers, fuel, etc. Sometimes called the "Body."
FUSELAGE BIPLANE.
See TRACTOR BIPLANE.
FUSIFORM.
Of streamline form.
G
GAP. The vertical distance between leading edges of the superposed
planes of the biplane or triplane.
GLIDING. (Fr. _Vol Plan_.)
With an aeroplane the weight of the machine can be made to
provide a forward component that will allow the machine to
descend slowly (without power) along an inclined line. This line
is known as the "Gliding Path."
GLIDING ANGLE.
The angle made by the gliding path with the horizontal is known
as the gliding angle. This may be expressed in degrees or in the
units of horizontal distance traveled per foot of fall.
GLIDER.
A small form of aeroplane without a power plant, which is capable
of gliding down from an elevation in the manner of an aeroplane.
With a proper direction and velocity of wind it can be made to
hold a constant altitude and can be made to hover over one spot
continuously.
_GOUVERNAIL_ (Fr.)
Rudder.
GUY WIRE.
A bracing wire.
H
HARD WIRE.
A solid tempered wire of high tensile strength used for aeroplane
bracing systems.
H.T. WIRE.
Another expression for hard or high tensile strength wire.
HEAD RESISTANCE.
The resistance of the structural parts of an aircraft. In an
aeroplane, the head resistance is the sum of the resistances of
the body, stays, struts, chassis, tail, rudder, elevators, etc.;
in fact, this includes everything with the exception of the wing
drag.
HELICOPTER.
A type of direct lift machine in which sustentation is performed
by vertical air screws or propellers.
HELIX. A geometrical curve formed by the combined advance and revolution
of a point.
_HELICE_ (Fr.)
Propeller or screw.
_HELICE TRACTIVE_ (Fr.)
Tractor propeller.
HYDROAEROPLANE.
See SEAPLANE.
HOLLOW WOOD CONSTRUCTION.
(Fr. _Bois Creus_.)
HOOD OF ENGINE.
(Fr. _Capot_.)
HYDROMETER.
An instrument for measuring the density of liquids.
I
ICTHYOID.
Fish or stream lined shape.
INCLINED PLANE.
A plane inclined to the wind stream so that the energy of the air
stream is broken up into the two components of lift and drag.
INCLINOMETER.
An instrument used for determining the angle of the flight path.
INCIDENCE.
See ANGLE OF INCIDENCE.
INCIDENCE, NEGATIVE.
The angle formed with the air stream when front edge of the
lifting surface dips below the apparent flight path.
INHERENT STABILITY.
Stability due to some fixed arrangement of the main or auxiliary
surfaces. A machine that requires mechanism or moving parts for
its stability is automatically but not inherently stable.
INTERFERENCE.
The crowding of the airstream in the gap of a biplane or triplane
causes the surfaces to "Interfere," and results in a loss of
lift.
J
JOY STICK.
See CONTROL STICK.
_JAMB de FORCE_ (Fr.)
Bracing strut.
_JANTE_ (Fr.)
Rim of wheel.
K
KITE BALLOON.
See BALLOON.
KILOMETER.
French metric unit of distance. One kilometer equals O.621
statute mile or O.5396 nautical mile.
KILOGRAM.
Metric unit of weight. One kilogram equals 2.205 Avoir. pounds.
KNOCKOUT HUB.
An aeroplane chassis wheel hub provided with removable bronze
bushings.
KEEL PLANE AREA.*
The total effective side area of an aeroplane which tends to
prevent skidding or side slipping.
L
LATERAL STABILITY.
Stability about the fore and aft axis.
_L’AVION_ (Fr.)
Aeroplane.
LAMINATED.
Built up in a series of layers.
LEEWAY.*
The angular deviation from a given course due to cross currents
of wind.
LEADING EDGE.
See ENTERING EDGE.
LIFT. The vertical component of the forces produced on an aerofoil by
an air current.
LIFT COEFFICIENT.
The lift per unit of area at a unit velocity (Ky). The American
lift coefficient is the lift in pounds per square foot at one
mile per hour.
LIFT CAPACITY.
See CAPACITY.
LIVE LOAD.
The live load generally includes the passengers, pilot, fuel,
oil, instruments, and portable baggage, although in some cases
the instruments are included in the dead load. The live load is
the difference between the total lift and the dead load.
LOADING (UNIT).
The unit loading is the load carried per square foot of wing
surface, or is equal to the total weight divided by the area.
LONGERONS.
The principal fore and aft structural members of the fuselage.
LONGITUDINALS.
See LONGERONS.
LONGITUDINAL STABILITY.
Stability in a fore and aft direction about the "Y" axis.
M
MASS. The quantity of matter. Is equal to the weight in pounds divided
by the gravitation, or generally to the weight divided by 32.16.
_MANDRIN de BOIS_ (Fr.)
Wood spar.
_MAROUFLAGE_ (Fr.)
Strut taping with fabric bands.
_MANCHE A BALAI_ (Fr.)
Control stick.
_MANETTE_ (Fr.)
Throttle.
_MAR, MONTANT_ (Fr.)
Interplane struts.
METACENTER.*
The point of intersection of a straight vertical line passing
through the center of gravity of the displaced fluid or gas, and
the line that formerly was a vertical through the center of
gravity before the body was tipped from its position of
equilibrium. There is a different metacenter for each position of
a floating body.
MONOPLANE. (Fr. _Monoplan_.)
A type of aeroplane with a single wing surface.
MONOCOQUE BODY (Fr.)
A body built up in tubular form out of three-ply wood, thus
virtually forming a single piece body.
_MONOPLACE_ (Fr.)
Single seater.
_MONOSOUPAPE_ (Fr.)
Single valve Gnome motor.
_MONTGOLFIER_ (Fr.)
Hot air balloon.
MULTIPLANE.
An aeroplane having the main lifting surface divided into a
number of parts.
N
NACELLE.
The body or fuselage of an aeroplane or dirigible. It generally
signifies a dirigible body. The short fuselage of a pusher type
is often called the nacelle.
NATURAL STABILITY.
See INHERENT STABILITY.
NEGATIVE AILERONS.
Ailerons making a negative angle with the wind when in normal
flight. The negative incidence of the ailerons is decreased on
the low side and increased on the high side so that the high side
is pushed down. This decreases the drag on the lower, inner wing
in making a turn, and therefore does not tend to stall the
machine.
_NERVURES_ (Fr.)
Wing ribs.
NORMAL PLANE.
A flat plane placed with its surface at right angles to the air
stream.
NORMAL PRESSURE.
The pressure at right angles to the surface of a plane.
NON-LIFTING TAIL.
A stabilizing surface arranged so that it carries no load in
normal flight.
NOSING.
The member used for the entering edge of the wing.
NOSE. The front end of the aeroplane.
NUT. (Fr. _Ecru_.)
O
_OBLITEUR RINGS_ (Fr.)
The special piston rings used on the Gnome motor.
ORTHOPTER.
Any type of wing flapping machine.
ORNITHOPTER.
A wing flapping machine that imitates bird flight.
ORTHOGONAL BIPLANE.
A biplane with the upper and lower leading edges in line.
OSMOSIS.
The transfer of hydrogen or other gas through a balloon envelope
by a molecular process. This must not be confused with leakage
due to holes.
OUTRIGGER.
See Boom.
O.W.L. TYPE.
A type of machine adapted for use over "Water and Land."
P
PANCAKE.
A straight vertical drop due to stalling.
PATH OF FLIGHT.
The path of the center of gravity of an aircraft in reference to
the air.
_PALONNIER_ (Fr.)
Foot bar or lever.
PANELS.
The wing sections included between adjacent struts.
_PATIN, PATINNAGE_ (Fr.)
Skids.
PARASOL TYPE.
A monoplane in which the wing is located above the body.
PENGUIN.
A training machine which cannot leave the ground.
PERSONNEL.
Pilot and passengers.
PETROL.
An English term for gasoline.
PHILLIP’S ENTRY.
See DIPPING FRONT EDGE.
PILOT. The operator of aircraft.
PILOT BALLOONS.
Small balloons sent up to determine the direction of the wind.
PITCH. The forward distance traveled through by one revolution of the
propeller.
PITCHING.
A fore and aft oscillation, first heading up and then diving.
PISCIFORM.
Fish form.
_PIQUE, VOL_ (Fr.)
Dive.
_PLAFOND_ (Fr.)
"Ceiling" or maximum altitude obtainable.
_PLAN CENTRAL_ (Fr.)
Center panel.
_PLANCHER_ (Fr.)
Flooring.
_PLAN FIXE de QUEUE_ (Fr.)
Stabilizer surface.
_PLAN de DERIVE_ (Fr.)
Stabilizing fin.
PITOT TUBE.
An instrument for measuring the velocity of an air current.
PONTOON.
Seaplane floats.
_POIGNEE_ (Fr.)
Handle.
_POMPE_ (Fr.)
Pump.
_POMPE A PRESSION_ (Fr.)
Pressure pump.
_POULIE_ (Fr.)
Pulley.
PROPELLER. (Fr. _Helice_.)
A device used in converting the energy of a motor into the energy
required for the propulsion of an aircraft. It consists of two or
more rotating blades which are inclined in regard to the relative
wind, and hence they act as rotary aeroplanes in creating a
tractive force.
PROPELLER ROTATION.
The direction of rotation is determined when standing in the slip
stream.
_PNEU_ (Fr.)
Pneumatic tire.
PTERYGOID ASPECT.
A wing flying with the long edge to the wind is said to be in
"Pterygoid Aspect."
PUSHER TYPE.
An aeroplane with the propeller in the rear of the wings.
PYLON. A marking post on an aeroplane course.
P
RACE OF A PROPELLER.
The air stream thrown by the propeller.
RADIAL MOTOR.
A motor with the cylinders arranged in radial lines around the
crankcase.
RAKED TIPS.
The tips are arranged at an angle with the wing so that the span
of the trailing edge is greater than that of the leading edge.
_RAYONS_ (Fr.)
Spokes.
REFLEX CURVE.
An aerofoil in which the trailing edge is given an upward turn.
_REMOUS_ (Fr.)
A downward current of air.
_RESERVOIR_ (Fr.)
Tank.
_RESERVOIR SOUS PRESSION_ (Fr.)
Pressure tank.
RESULTANT.
The total force resulting from the application of a number of
forces.
RETREAT.
Back swept wings with the tips to the rear of the wing center.
RIBS. The fabric forming member of the wing structure.
RUDDER, VERTICAL.
A control surface used for steering in a horizontal plane.
ROLL. Oscillation about the fore and aft axis.
S
SCREW. (Fr. _Helice_.)
See PROPELLER.
SEAPLANE.*
An aeroplane equipped with floats or pontoons for landing on
water.
SCOUT TYPE.
See CHASER.
SERVICE TANK.
The fuel tank feeding directly into the carburetors.
SET BACK WINGS.
A type of wing in which the leading edge is inclined backward as
in the Mann biplane. The trailing edge is straight.
SHOCK ABSORBERS.
An elastic device on the chassis or landing gear that absorbs
vibration by allowing a limited axle movement.
SIDE SLIP.*
Sliding down sideways, and toward the center of a turn. This is
due to an excessive angle of bank.
_SIEGE_ (Fr.)
Seat.
SIMILITUDE, LAWS OF.
The drag or resistance of a small aerodynamic body does not
increase in direct proportion with the area and speed. The laws
governing the relation between a model and a full size machine
are known as the laws of "Similitude."
SKIDS.* (Fr. _Patin, Pattinage_.)
Long wood or metal runners attached to the chassis to prevent the
"nosing over" of a machine when landing, or to prevent it from
dropping into holes or ditches on rough ground. It also acts when
the wheels collapse.
SKID CURTAINS.
Vertical side curtains or surfaces provided to reduce the
skidding action on turns or to prevent side slip.
SKIDDING.*
Sliding sideways away from the center of the turn. It is due to
insufficient banking on a turn.
SKIN FRICTION.
The resistance caused by the friction of the air along a surface.
SLIP.* Applied to propeller action, the slip is the difference between
the actual advance of an aircraft and the theoretical advance
calculated from the product of the mean pitch and the revolutions
per minute. When the propeller is held stationary, the slip is
said to be 100 percent.
SLIP STREAM.
The wind stream thrown by a propeller.
SOARING FLIGHT.*
The sustentation of a wing surface due to wind currents and
without the expenditure of other power. Soaring flight is
performed by gulls, buzzards and vultures, but no practical
machine has yet been built that will fly continuously without the
aid of power.
SPAN. (Fr. _Envergure_.)
The length or longest dimension of a wing, generally taken at
right angles to the wind stream.
SPAR. (Fr. _Bras D’Aile_.)
The main wing beams that transmit the lift to the body.
SPOKES.
(Fr. _Rayon_.)
SPREAD.*
See SPAN.
STABILITY.*
The property of an aircraft that causes it to return to a
condition of equilibrium after meeting with a disturbance in
flight.
STAGGER.*
The advance of the leading edge of the upper wing over that of
the lower wing.
STABILIZER.* (Fr. _Stabilisateur_.)
A horizontal tail surface (fixed) used for damping out
oscillations and for promoting longitudinal stability.
STALLING.*
The condition of an aeroplane that has lost the speed necessary
for steerage way or control.
STALLING ANGLE.
See CRITICAL ANGLE.
STEERING WHEEL.
(Fr. _Volant_.)
STATOSCOPE.*
An instrument for detecting a small rate of ascent or descent.
Used principally with balloons.
STAY WIRE. (Fr. _Tendeur_.)
A wire or cable used as a tie to hold members together, or to
give stiffness to a structure.
STEP.* A break in the form of a float or flying boat bottom.
STREAMLINE.
A form of body that sets up no turbulence or eddies in passing
through air or liquid.
STRUT.* (Fr. _Mar, Montant_.)
A compression member used in separating the upper and lower wings
of a biplane, or the longerons of the fuselage.
SWEEP BACK.
See RETREAT.
T
TACHOMETER. (Fr. _Compte Tours_.)
An instrument for directly indicating the revolutions per minute.
TAIL.* (Fr. _Queue_.)
The rear part of an aircraft to which usually are attached the
rudder, stabilizer, and elevator.
TAIL SKID. (Fr. _Bequille_.)
A flexibly attached rod which holds the tail surfaces off the
ground, and breaks the landing shocks on the tail structure.
TAIL BOOM.
See BOOM.
TAIL DIVE.
A very dangerous backward dive.
TAIL SPIN.
A condition in which the tail revolves about a vertical line
passing through the center of gravity.
TANDEM PLANES.
A form of aeroplane in which the wings are placed one after
another.
TAUBE. An old type of German or Austrian aeroplane with back swept wing
tips.
TAXI. To run along the ground.
THIMBLE. (Fr. _Cosse_.)
An oval grooved metal fitting used for the protection of a cable
loop at the point of attachment.
THREE-PLY. (Fr. _Contreplaque_.)
A wood sheet composed of three layers of wood glued together, the
line of grain crossing at each joint.
THRUST.
The propulsive force exerted by a propeller.
THRUST DEDUCTION.*
The reduction of thrust due to a reduction of pressure under the
stern of the aircraft.
_TIRANT_ (Fr.)
Bracing tubes.
TORQUE.
The turning force or moment exerted by the motor.
_TOILE_ (Fr.)
Linen.
TORQUE WARP.
The amount of warp, or permanent set in the ailerons necessary to
overcome the torque or twisting effect of the motor. In some
machines the torque is overcome by changing the angle of
incidence at the wing tips.
TRACTOR BIPLANE.
A type of aeroplane in which the propeller is placed in front of
the wings so that it pulls the machine along.
TRAILING EDGE.
The edge of a wing at which the air stream leaves the surface.
_TRAIN D’ATTERRISSAGE_ (Fr.)
Landing gear.
_TRANSFIL_ (Fr.)
Cord winding on the struts.
TRIPLANE. (Fr. _Triplan_.)
An aeroplane with three superposed wings.
TURBULENCE.
The eddies or discontinuity caused by a body or surface passing
through the air.
U
USEFUL WEIGHT. The difference between the total lift and the dead load.
This comprises the pilot and passenger, the weight of the fuel, baggage
and instruments.
UNIT LOADING. The weight per square foot of main wing surface.
V
_VOL PLANE_ (Fr.)
See GLIDE.
_VOL PIQUE_ (Fr.)
See DIVE.
_VOLANT_ (Fr.)
Steering wheel.
_VERNIS_ (Fr.)
Varnish.
_VRIL_ (Fr.)
Spinning nose dive.
_VOLETS de PROFONDEUR_ (Fr.)
Elevator flaps.
W
WARP CONTROL.*
Lateral control obtained by twisting the wing tips.
WASHOUT.
Decreased camber or incidence toward the wing tips.
WEATHER-COCK STABILITY.
Stability in the line of travel, or with the relative wind, so
that the machine always tends to head into the wind.
WHALE TYPE.
A speed type biplane in which the body entirely fills the gap
between the upper and lower wings.
WHIRLING TABLE.
A testing device in which a model wing or body is placed at the
end of a revolving arm.
WINGS.* (Fr. _Aile_.)
The main supporting surfaces of an aeroplane.
WORKING FACE.
The face of a propeller blade lying next to the slip stream.
WAKE GAIN.*
Due to skin friction and eddies, a moving aircraft drags a
certain amount of the surrounding air with it. This reduces the
effective resistance of the hull and increases the effective
pitch of a pusher propeller since the latter acts on a forward
moving mass of air. This is "Wake gain."
[Illustration: Launching with Catapult from Deck of Battleship]
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