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