4
good clean design and a wingi-
Lightplane Design A biplane of a given geometric aspect ratio is not as efficient as a monoplane of the same ratio, and the reduction in span is offset in part by the decreased efficien-
cy occuring in the form of increased induced drag. This is appreciable only at low speeds, but will result in an increase in landing speed. Just how small the difference in span between a monoplane and a biplane of equivalent aspect ratio is, can best be shown by an illustrative example. A biplane, of gap-chord ratio 1.00, must have a geometric aspect ratio of 10.0 in order to have a monoplane equivalent of 10.0. Comparing the spans of a monoplane with a wing area of 150 square feet, and a biplane of the same area with the upper and lower wings of the same areas AR equals b2/S (1) Where AR equals Geometric Aspect Ratio And for the biplane: b equals /10xl50/2 equals /750 equals 27 feet, 5 inches Thus, it may be seen that the saving in span, without loss of efficiency, is but slight and in most cases would be less than indicated above, as the area of the upper and lower wings of a biplane are seldom the same. A greater saving in span can be affected, of course, by using a geometric aspect ratio of 6.00 for the biplane, and suffering the loss in efficiency. In that case the span would be: b equals /6x75"
equals
/450
equals 21 feet, 4 inches. Another possibility is to use a
tapered wing on the monoplane, thus reducing the area at the tip. However, the simplest course to follow is to use a wing section with a high maximum lift coefficient, and then decrease the wing area as much as possible without making the landing speed greater than 35 to 40 mph.
Such a landing speed is not too high and is, in fact, preferable to one too low from a standpoint of controllability, as there is danger of losing lateral control at very low speeds, due to decreased rudder and aileron effectiveness. Stick
forces, landing speed and wing area small enough to make the ship stable in all kinds of air, are not the only considerations in the design of the plane, however. In order to get the economy mentioned above, it is necessary that the engine be of fairly low power,
so that the fuel consumption will not be too high, but at the same time the plane must be sufficiently fast to avoid throwing the gain away. This means, of course, a
section with a high ratio of maximum lift to minimum drag. This latter is needed in order to have a high speed range, and thus a high enough top speed without too high a landing speed. It is better to gain in speed range in this manner than through the use of flaps, or some similar device, as the flaps mean an additional control for our "dub" pilot to worry about, and possibly to misuse. The need for stability has been
mentioned above, and it has been seen that it is necessary to work out a compromise between the stability and the maneuverability of the plane. One other phase of stability to consider is spinning, and certainly the ship must have
a slow rate of spin, recover rapidly and have no tendency to spin flat. Not least important in designing a plane for the private flyer is the looks of the ship. And in respect to the appearance of the plane there is a great deal to say. The well-known fact that fancy paint jobs sell automobiles is also true of paint jobs and trim lines with airplanes. The prospective customer expects and demands a classy looking layout, and if the prospective customer is a woman, oh. . . . . .well! Many ships showing excellent performance h a v e failed to achieve a commercial success because they were not good looking. And, unfortunately, there have been very few good looking light planes. And, it is just as easy to design a ship that will create desire in the minds of the public as it is to lay out a poor appearing ship. If the average person is to be interested, the plane must be one he will be proud of showing! to his friends, and this will be more and more true as the market grows. *
*
***
The general features needed in the small plane for the private flyer were outlined in the preced-
ing article of this series. In order to develop the various points as
clearly as possible, it is best to have reference to a concrete base from time to time, so the design
of a typical plan will be carried through.
The plane selected will be a tandem two-place open cockpit lo-,v wing monoplane, with a full cantilever wing having a two to
one taper in plane form. The power-plane will consist of an air cooled engine developing about sixty-five horsepower. This fig-
ure does not represent the output developed by any particular engine now being built, but is. in
the opinion of the author, the minimum necessary for good performance in a two place airplane. It is perhaps, a trifle low—seventyfive horsepower might be a better figure — but since it was decided that economy was one of the most important features of this plane, it
seems best to set the figure as low as possible. There are unquestionably many readers who will take exception to the type of plane described above, so it seems advisable to spend a short time discussing the features which led to its choice. There are three definite advantages which come from the use of a low wing — a low center of gravity, the greatest possible "ground effect," and a maximum ol vision. A low center of gravity does two things — it gives better "rower on" and "power off" stability and reduces the danger of the plane nosing over in landing.
Considering the former, when the throttle on a plane is opened it produces two effects, (1) an increase in propeller thrust and (2) an increase in the down load
on the tail. The latter creates a stalling moment, and when the center of gravity is below the thrust-line, the former produces
a diving moment from the tail, and leaves the ship in the same attitude. As to the decreased tendency to nose over, anyone who has seen a low wing plane landed with the tail in the air, and instead of nosing over as most types of planes, rocked a bit, and remained right side up, will appreciate what this means to an inexperienced pilot. The "ground effect", by which is meant the reduction in induced drag when the airplane is close to the ground, is obviously a maximum for a low wing monoplane, since in this type the wing is nearer the ground than in any other, and results in a decreased landing speed. This, then, is one way of keeping the landing speed down without using an excessive amount of wing area. As to vision, a low wing plane gives a maximum in a climbing turn, where it is of vital
importance and at the same time, if the cockpits are properly located in respect to the wing, still
gives ample vision down. The choice of a tapered cantilever wing is less readily defensible. There is no doubt that, unless there is mass production, the tapered wing is much more expensive to build than is a rectangular one. It is also true that a cantilever wing must be considerably thicker than an externally braced one, so that the profile wing drag will be increas-
ed. This is offset, in part, by the elimination of wing bracing, but even so it is a factor of primary
importance. However, there are two important advantages which can be written against the disadvantages just discussed. The first was mentioned in the last article - the distribution of area in a tapered wing is such that the resultant force on the wings acts nearer to the plane of symmetry of the airplane than a rectangular wing, thereby giving smaller rolling moments from gusts of air and makies the plane less subject to weather conditions.
The second advantage calls for a momentary consideration of wing theory. Both from theory, and from pressure distribution tests, it has been found that the ideal spanwise lift distribution is elliptical in form, this pattern giving the minimum induced drag. Of course, the only time this distribution actually exists is for an elliptical wing. However, theory, once again borne out by wind tunnel tests, showfe that a wing with a plan form taper between two or three to one, approaches this pattern the most nearly. The induced drag of a rectangular wing of aspect ratio 6.0 is five percent greater than that of an elliptical wing, while that for a wing with a two to one taper and the same aspect ratio is only one percent higher than for the elliptical plan form. Thus, to summarize, the tapered cantilever wing makes the plane more expensive to build, and increase the profile drag, but makes the plsne more stable in rough air,
and decrease the induced dra^. Since the profile dra" increases w i t h the square of the velocity,
while the induced drag varies inversely with the square of the
velocity, the former will bo appreciable only at high speeds and the latter only at low ones. Consequently the effect will be to reduce both the maximum and minimum speeds of the plane. For the purpose of this plane the advantages of the tapered wing seem to outweigh the disadvantages, thereby justifying1 its use. It was decided in the previous articles that the geometric proportions of the plane should be determined from the calculations for the stick forces, so it is necessary to proceed with that phase of the design before anything can
be done toward a layout. There is only one thing which must first be decided upon, and that is the aspect ratio of the wing. In order to keep the span from get-
5
to use an aspect ratio greater than six, especially since the use of a tapered wing will make the span greater than it would otherAise be. At the same time using any
lower value would tend to offset the increase in efficiency due to the tapered wing. Consequently an aspect ratio of six will be used. It is now necessary to decide what stick force is desirable
at some particular
attitude of
flight, and since the stick forces are a maximum at landing that is the best criterion. The only
way in which to determine the proper force is to determine the value of the force on some plane which, according to the majority of pilots "feels right". A comparison of such values for several
planes leads to the conclusion that a stick force between four and five pounds at landing is
about the correct value.
News Note
Member Clyde A. Wilson, who lives at 1126 Valley Street in Minot, North Dakota acquired a Kari-Keen in 1952, and informs us that he just recently flew the ship for the first time after extensively reworking the ship. The Kari-Keen was a light plane with a full cantilever wing, and was built in 1929 at Sioux City, Iowa. The ship has a span of 30 feet, and a wing area of 160 sq. feet. Clyde repaired the entire ship, and installed a Continental A-75-8 engine. The ship is fabric covered, and seats two, side-by-side. It bears the identification number N-244K(X).
Letters To The Editor
Experimental Aircraft Association 3801 S. 56,th Street Milwaukee, Wisconsin
Attention: Mr. Paul H Poberezny Dear Paul: I would like to initiate a movement in the Experimental Aircraft Association to get a pledge from all members to install and
wear shoulder harness in all homebuilt aircraft as well as to provide
extra
structural
crash
protection for passengers and crew. Aside from being advertised as an established policy, such a pledge could be incorporated on the membership application form. Our members could be spared many injuries and the movement will be known as a safer hobby if we can eliminate most of the danger from any accidents.
We are pleased to welcome as a
new member, Mr. John Gelb of Buffalo, New York. Mr. Gelb is a
graduate aeronautical engineer, and
has served as aircraft designer with such firms as Heinkel in Germany, Boulton-Paul in England, and Kaman Aircraft Co. and Bell Aircraft
Co. in this country. He presently
flies a Cessna 170. Despite his affiliations with the industry, he states he is convinced that in order to make progress in light aircraft, and organization outside the present aircraft industry is needed, and that
the E.A.A. is that organization.
I am extremely conscious of such things as I had to make a forced landing with my little job last month into some impossible terrain. After I came down into a pile of boulders which stopped
me in just a few feet, my recording accelerometer read - 4.5 G's, plus 11.5 G's and I estimate my horizontal deceleration at 20 G's.
My harness limited the after affects to a two-day stiff neck. If anyone asks me, I say. "I Love shoulder
harnesses!"
My little
AIRMATE also took the punishment well and will be flying again in about two weeks. Sincerely yours, AIRMATE COMPANY, R. E. Schreder
Note: New membership blanks will now contain pledges as suggested by member Schreder.
12
Ccb/c] —[(St/Sx) (1/c) (dCL/da)t7tat] —[(Ss/Sw) (1/c) (dCL/da)t?tat
Aircraft Design In order to derive a method of finding the stick forces it is best to start with an expression for that force and then work back to some equation involving the general characteristics of the plane. This leads, obviously, to the equation for static longitudinal stability, as the expression for the pitching moment of the tail is the only one which will involve the load on the tail and the dimensions of the airplane. The value of the force which the
pilot must exert on the grip of the stick can be expressed by the equation F=kH/L . . . . . . . . . . . . . . . . . . (1) H=Elevator hinge moment (inch-lbs.) L=Length of stick (in.) k=Gear ratio in control system.
Expressing the value of the hinge moment in terms of the absolute moment coefficient, H=CH ce se q . . . . . . . . . . . . . . (2)
Where: ce=Elevator chord (in.). se= Elevator area (sq. ft.). q=Dynamic pressure.
s
[Th/cswq] . . . . . . . . . . . . . . . . (6) Where: CMO = Wing moment about chord quarter point, from wind tunnel
curves. CL=Lift coefficient from curves. a = Dist. from wing leading edge to center of gravity (in.+back.) c = Mean wing chord (in.)
St—Total tail area (sq. ft.). Sw=Total wing area (sq. ft.).
l = Dist. from .G. to C.P. of tail (assumed at tail post., in.) dCL/da)t=Slope of plot of tail lift coefficient
against angle of
attack. at=Angle of attack of tail = aw—E+ato+TB = as+TB . . . . . . . . . . . . . . . . . . (7) All of the symbols in this expression but T are discussed above. T=(dat/dB), or the slope of the plot of at against B. t=Tail efficiency factor. Cc = Wing chord force coefficient
= CD Cosa w — CL Sin aw. b = Dist. from wing chord to e.g. in chord.)
OH= Hinge moment coefficient. (See equation 4.). q=eV2 . —— . . . . . . . . . . . . . . . . . . . . (3) 2
Where: e = Air density (slugs per cu. ft.) V=:Velocity (ft. per sec.). CH=Clas+OB . . . . . . . . . . . . (4)
Ss=Tail area in slip-stream (sq. ft.) T +Propeller thrust (Ibs.)
— Total drag of airplane. D = Propeller diameter (ft.) h=Dist. from thrust line to e.g. in inches ( + when e.g. is below T.L.). In equation 6 it is to be
noted that the first bracket represents the pitching moment from the wing, the second that from
dC
Cl=———H and O =———H
(1075T/D2V2)]
inches (-f- when e.g. is below
(See equation 3.)
Where: de
:
.
das dB or the slopes of the plots of CH against ds and B respectively. The angle of attack of the stabilizer (ds) is expressed: ds=dw+E+dto . . . . . . . . . . . .(5) Where: dw=Angle of attack of the wing
chord (deg.). E=Angle of downwash (deg.). dto=Angle of stabilizer setting with respect to the wing chord (deg.). B=Angle of attack of elevator relative to the stabilizer. (In equation 4) The angle of the stabilizer setting (dto) can be found by solv-
ing the equation for the pitching moment of the airplane at cruising speed, since the stabilizer will be set to trim the plane at that speed, making the pitching moment and the hinge moment both equal to zero. Writing the equation in the form of non-dimensional coefficients, CMA=[CMO + CL (a/c—.25)
the tail neglecting the effect of the slip stream, and the last the thrust moment. Consequently, while the angle of attack of the tail (at) is expressed by equation 7 for both the second and third brackets, the value of downwash (E) must be
modified for the slip-stream in the latter case. For the case without the slip-stream, a very close approximation of the value of E can be obtained from the expression: E = 26 OL/R . . . . . . . . . . . . . . . . ( 8 )
Where: Rawing aspect ratio.
The additional affect of the slipstream can be accounted for by the expression: Et=26 CL/R+aw[(Vs—V)/Vs]
[1—(dE/da)] Where: Vs=Slip-stream velocity (ft. per sec.). V=Velocity of plane (ft. per sec.). dE/da = Slope of plot of E against aw.
The value of the elevator angle
(B) can be found from equatioii 4, since in the case now being considered CH=O. That leaves only T to be evaluated to have eliminated all of the unknown from equation 7 but the one for which
equation 6, is to be solved. The value of (dat/dB) or T depends on the shape of the tail surfaces, and on the ratio of mean elevator chord to the total mean
chord of the tail surface.
There
is no way in which it can be computed, but its value must be taken
from test data. Values of it for a given shape tail will be discussed later. The pends of the tween
tail efficiency factor deon the shape and location tail, and usually lies be0.65 and 0.80. It must be
noted that equation 6 involves
of 750 to 800 Ibs., both figures dependent upon the engine selected. With the help of any engineer who might consider helping him, Paul wants to work the design up into good construction drawings, and clear up all the other details preparatory to beginning the actual caonstruction. Anyone interested? Did you know that right now in France, there are well over three hundred home-builts flying, and a goodly number under construction? France is generally credited as being the leader in the home-building field. With as much aviation activity and interest shown in home building in this country, we see no reason why this country shouldn't capture that distinction before very long. How about it, men?
both 1/c and a/c, or the tail length
and the location of the center of gravity in terms of the chord. Since these are the two proportions of the airplane which are to be determined from a basis of the stick forces, it is obvious that the only possible procedure is to tssume values for them, solve through for stick forces, and if the lorce is not within the desired limits, to vary them slightly, and make another solution, until the desired result is obtained. While this sounds like a long drawn out and tedious method, after having carried through the solution a few times, one learns just about what proportions will be required for any particular type of airplane. Methods for determining the necessary amount of tail area, and any other factors in equation 6 which have not already been discussed, as well as the remainder of the solution for the stick forces having found the value of as, along with a solution for the forces of the stick on the plane being designed herein will be considered in the following article of this series.
EAA News Section Continued from page 4 It will be 19' 7" long with an equal wingspan of 21', and a continuous chord of 33" for both the upper and lower wing panels An engine of between 85 to 135 hp. is planned, and he also figures
full length flaps for the Ijwer wing panels. Expected performance would give it a maximum speed of 170-180 mph., cruising speed of 150 - 160 mph. and a landing speed of 65 or 70 mph. The rate of climb will be about 1400-1600 fpm. The wing area will be approximately 100 sq. ft. It will have a gross weight in the neighborhood ot 1350 Ibs., and an empty weight
EAA Headquarters has received a great many letters from readers and members indicating they would like to purchase materials, parts
and engines at a discount from dealers and suppliers as a member of EAA. As there are so few of us here at Headquarters to make contacts with suppliers, etc., possibly you members could make such contacts and forward the information to us. If you know of companies who supply materials that are useful in homebuilding, it is requested you contact them for possible advertising in the Experimenter as it will surely pay off.
Books For The Homebuilders Continued from page 6
and application to design of aircraft structures. 2)
(a) PROCEDURE HANDBOOK
FOR AIRCRAFT (b) PRACTICAL AERO ENGINEERING ($4.00 each---Aviation Press, 580 Market Street, San Francisco, California). They are 1938 and 1940 books, but are extremely good. Although dealing with aviation on an older level, they have good discussions on parasites drag, performance, etc. 3) (A) ANC-18 DESIGN OF WOOD AIRCRAFT TURES ($1.00)
STRUC-
(b) ANC-19 WOOD INSPECTION & FABRICATION ($1.25) Both are obtainable from the Superintendent of Documents, Washington, D.C., and are a must for design involving Wood construction. 4) NACA REPORTS 824 (Late sections--Laminar Airfoils) 485-518-522 (Wheels, gear, pants drag of 460 Airfoils-4 digit series) 5) AIRCRAFT MECHANICS POCKET MANUAL (Vale.... $6.75) This publication is good for
