aiaa- aerodynamic design of low speed aircraft

AI AA-86-2693 Aerodynamic Design of Low-Speed Aircraft With a NASA Fuselage/Wake- .Propeller Configuration F.R. Goldschmied, F.R. Goldschmied P.E., Monroeville, PA

AI ANAH S/AS E E Aircraft Systems, Design ti

Technology Meeting October 20-22 1986/Dayton, Ohio

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1633 Broadway, New York, NY 10019

AERODYNAMIC DESIGN OF LOW-SPEED AIRCRAFT W I T H A NASA FUSELAGE/WAKE-PROPELLER CONFIGURATION

F A B I O R , G O L D S C H M I E D * M O N R O E V I L L E PA 15146

AIAA 86-2693

Abstract

4 A brief parametric study has been carried out on the application of a NASA axismmetric fuselage/ wake-propel l e r configuration , using f u l 1 -Scale wind-tunnel data, t o low-speed general aviation a i r c r a f t w i t h conventional NACA wings . The design matrix comprises t w o wing aspect ra t ios , ( 8 and IO), tHo wing loadings (15 and 21 PSF), f i v e fuselage diameters (40", 42", 48", 55" and 65") and f ive corresponding gross weights (800, 1200, 1400, 2400 and 3200 l b ) . The experimental propulsive efficiency of the NASA fuselagdwake-propeller configuration goes from 103% f o r the bare fuselage t o a range between 96% and 85% f o r the a i rc raf t . The aerodynamic eff ic iency index ranges from 17.95 t o 11.47 while tha t f o r conventional a i r c r a f t ranges from 7.95 t o 5.00, a s shown by a survey Of 76 general aviation and sport a i r c r a f t . A 50% power reduction, f o r the same gross weight and speed, i s a very practical possibi l i ty .

Finally, a specif ic comparison has been carried out between the 55" diameter 4-seat design and one of the l a t e s t 4-seat pusher a i r c r a f t , f o r the same 2400-lb gross weight and 180 MPH speed, yielding 78 HP f o r the former and 180 HP f o r the l a t t e r ; t h i s demonstrates t h a t , while a wake-propeller iS a pusher, a pusher propeller i s not necessarily a wake-propeller.

Nomenclature

Propeller advance ra t io

Wing area, f t 2

Aerodynamic efficiency index

Wing aspect ra t io

Wing span, f t

Volume axial drag coeff. of fuselage/wake-propel l e r

Volume a x i a l drag coeff. of bare fuselage

Volume axial thrust coeff. of fusel age/wa ke-propel 1 e r

Area axial drag coeff. of Wings e tc .

Wing area l i f t coeff.

Lift/drag ra t io

Propeller power coeff.

HP Volume power coeff. of fuselage/ CHP = u ~ 0 . 6 6 wake-propeller 90 0 d Propeller diameter, f t

*Consultant. Associate Fellow, AIAA. "Copyright 1986 by Fabio R. Goldschmied. Published

by AIAA w i t h permission.

D

f = D L

F

FO

T = -F

TO

L

n

qo = kPU0 2

RL - - LUO

Rv = _ - vo.33u0 "

u, = u, V

YO. 66

~ 0 . 3 3

WO

B

TI TOUO HP

q FOUO ? --Ti-

"

Fuselage diameter, f t

Fuselage fineness r a t i o

Net axial drag force on fuselage/ wake-propel l e r , 1 b

Axial drag force on bare fuselage (no propeller) , l b

Net axial thrust force on fuselage/wake-propeller, l b

Thrus t produced by propeller, l b

Fuselage length, f t

Propeller speed, RPS

Free-stream dynamic pressure, PSF

Length Reynolds Number

Volume Reynolds Number

Free-stream velocity, FPS

Fuselage volume, f t3

Fuselage equivalent area, f t z

Fuselage equivalent length, f t

aircraft grossweight-wing l i f t , l b

Propeller blade angle

Propeller efficiency

Propulsive efficiency

Air kinematic viscosity. f tz /sec

Air mass density, I b sec2/ft4

I . Introduction

The concept of body wake regeneration f o r propulsion originated i n 1865 with Froudel; his t h i n k i n g was based on the overall momentum balance of the moving vehicle and he believed tha t the Rankin drag/thrust concept was an anachronism from the days when canal barges were towed by horses. When a force i n one medium must be overcome by power i n p u t i n another medium, o r more generally when there i s an impedance matching problem, the drag/thrust concept may have great merit. In steady motion through a single f l u i d , however, a s w i t h an a i r c r a f t , a LTA or a submarine, the drag/ thrust concept i s misleading i n i t s apparent sim- p l i c i t y and i t invariably resu l t s i n the adoption of reduced performance targets . For instance, when

a fuselage i s sa id t o have a c e r t a i n drag a t a given speed, i t i s imp l i ed t h a t the fuselage wake's momentum i s condemned t o useless d i ss ipa t i on , wi thout poss ib le recourse o f any k ind; i t i s imp l i ed t h a t the drag can on ly be balanced by an equal pro- p e l l e r t h r u s t , according t o the Rankin concept. S t i l l today, general a v i a t i o n a i r c r a f t . are viewed e s s e n t i a l l y as powered g l i de rs , w i t h the t h r u s t e r ( p r o p e l l e r o r j e t u n i t ) i n s t a l l e d i n a manner no t conducive t o e f f i c i e n t wake regenerat ion. Even fuselage-mounted pusher p rope l l e rs a re too l a r g e and are no t t a i l o r e d t o the spec i f i c wake, as w i l l be seen i n Sect ion V, where the wake-propeller a i r - c r a f t i s compared t o a conventional pusher-propel ler a i r c r a f t .

The r a t i o n a l approach i s t o f o l l o w Froude's concept' and t o expend power t o prevent o r t o m in i - mize the occurrence of the wake. Smith and Rob- e r t s , * Kuchemann and Weber,3 Edwards," 5 David- son,6 Goldschmied7 and many others have con- t r i b u t e d t o the development o f wake regenerat ion, w i t h and wi thout a c t i v e boundary-layer con t ro l . For instance, Goldschmied7 8 has shown tha t , f o r an axisymmetric fuselage, the wake drag can be reduced by a f a c t o r o f 10 w i t h an e f f i c i e n t s ing le- s l o t suct ion boundary-layer contro l aftbody design; t he o v e r a l l power was reduced by a fac to r o f 2, f o r equal fuselage volume and speed.I0 I n hydro- dynamics, t he a p p l i c a t i o n o f wake-propellers has become standard p r a c t i c e f o r a l l high-speed under- water veh ic les such as submarines, torpedoes, e t c . An exce l l en t reference on t h i s subject i s given by Huang, Wang, S a n t e l l i and Groves."

It has been t h i s au tho r ' s experience t h a t commu- n i c a t i o n s between hydrodmamic is ts and aerodynami- c i s t s have no t been good; even w i t h i n t h e A I A A i t s e l f , papers i n the Journal of Hydronautics f a i l e d t o get many readers from the a i r c r a f t commu- n i t y . As a t y p i c a l example, a very s i g n i f i c a n t 1976 paperi2 was ignored by a t l e a s t four A I A A authors.13 l6

The ob jec t i ve o f t h i s b r i e f p r e l i m i n a r y parame- t r i c study i s t o acquaint the general a v i a t i o n com- muni ty w i t h the aerodynamic p o t e n t i a l of fuselage/ wake-propel ler con f igu ra t i ons f o r s ingle-engine low-speed a i r c r a f t . Th i s p o t e n t i a l was demonstra- ted exper imenta l ly by a NASA wind-tunnel t e s t o f a f u l l - s c a l e fuselage (50.88'' diameter, 246" l eng th ) w i t h i t s custom-tai lored wake-propeller (24.00" diameter) a t 100 MPH.

The very extens ive t e s t r e s u l t s were reported by McLemore" i n 1962 bu t i t appears t h a t they have been ignored f o r a i r c r a f t app l i ca t i on s imply be- cause the p r o j e c t had been funded by the LTA program and the fuselage was designated as a "1/20- Scale A i r s h i p Model . ' I

A recent NASA-funded survey of p r o p e l l e r propul- s ion system i n t e g r a t i o n by M i l e y and van Lavante18 does n o t i nc lude any reference t o McLemore'sl7 ex- c e l l e n t work, a l though a m a j o r i t y of the survey's e f f o r t i s claimed t o have been d i rec ted t o the time pe r iod before 1964; no o the r i n v e s t i g a t i o n o f fuselage/wake-propeller i n t e g r a t i o n i s t o be found among the 121 references reviewed by M i ley and van Lavante. In t h e i r a u t h o r i t a t i v e textbook on a i r - p lane aerodynamics and p e r f o m n c e , Lan and Roskam'g s imply do n o t recognize even the possi- b i l i t y of wake-propulsion; on pase 269 of Ref. 19 i t i s stated: " In pusher con f igu ra t i ons .... the p r o p e l l e r w i l l be operat ing i n s i d e the nace l l e

( fuselage?) wake. I n t h i s s i t ua t i on , t he p r o p e l l e r e f f i c i ency can be g r e a t l y reduced."

The power assessment o f a i r c r a f t i s c a r r i e d out on the bas i s o f t he Aerodynamic E f f i c i e n c y Index (AEI) which i s def ined as fo l l ows :

AEI = Gross W t . x Speed Prope l l e r Power

4-J

_ _ _ - Lift x Propuls ive E f f i c i e n c y Drag

Since there i s no r e a d i l y a v a i l a b l e A E I data base f o r general a v i a t i o n a i r c r a f t , such a data base has been compiled f o r 76 a i r c r a f t and i t i s tabu- l a t e d i n Appendix I and I 1 f o r convenient r e f e r r a l . The A E I data p o i n t s have a l s o been p l o t t e d against a i r c r a f t speed i n F ig . 12.

11. NASA Wind-Tunnel Test

The NASA wind-tunnel t e s t program was c a r r i e d out i n the Langley 30 ' x 60' Wind-Tunnel and was reported by McLemore.17 The axisymmetric fuselage had 50.88" diameter and 246.46" length, w i t h a fineness r a t i o f = 4.84 and an enclosed volume V = 184 f t 3 . The p r o f i l e shape o f the body was o f t he well-known "Akron" a i r s h i p f a m i l y 2 0 21 b u t shortened from the Akron 's f = 5.9 t o f = 4.84. It can be noted t h a t a body w i t h f = 4.5 had a l s o been tested i n the wind-tunnel by Abbott,22 y i e l d - i n g no h ighe r drag. I t i s i n t e r e s t i n g t o note t h a t the Akron wind-tunnel t e s t mode12Q 2ihad a l eng th o f 236" and a diameter o f 40", w i t h a l eng th Rey- no lds number o f RL = 1.8 x IO6 and a volume drag c o e f f i c i e n t Cg = 0,022.

On the o the r hand, the drag c o e f f i c i e n t o f McLemore's'7 model was (w i thou t p r o p e l l e r ) CD = 0.021, i n c l u d i n g the three t a i l f i n s and the "gondola" a t t he same Reynolds number. The ne t drag w coe f f i c i en t o f t he bare body can be estimated from the t e s t r e s u l t s o f Abbott** (F ig . 8 o f Ref. 22): the drag c o e f f i c i e n t increment due t o the presence o f the f i n s i s ACD = 0.002 and the drag increment due t o the gondola i s ACD = 0.001. Thus an i nc re - ment of 0.003 must be added t o the ne t t h r u s t data, so as t o t r u l y p o r t r a y the bare body perfotmance.

It can be noted t h a t t he minimum pressure occurs a t 18% length, thus laminar f l ow can p l a y Only a minor r o l e i n the performance; Abbott** found no drag d i f f e rence f o r h i s f = 4.5 body when the t e s t model was pol ished a l l over (F ig . 5 of Ref. 22). The maximum diameter occurs a t 40% length; Table I below l i s t s the l ena th and diameter coordinates of the t e s t body for -convenient reference, s ince they a re no t given by McLemore.17

The wake-propel ler had 4 blades and 24" diameter ( l e s s than ha l f body diameter) ; the diameter corresponds t o the measured diameter of the wake (w i th - out p r o p e l l e r ) .

Although the t e s t model was labeled a 1/20-Scale A i r s h i p Model, i t happens a l s o t o be a f u l l - s c a l e fuselage f o r general a v i a t i o n a i r c r a f t , being l a r g e enough f o r side-by-side seat ing!

F igure 1 presents the schematic o f the t e s t model w i t h 3 t a i l f i n s and the "gondola"; i t i s reproduced from Fig. 1 o f Ref. 17. I t can be noted t h a t the drag o f an a i r c r a f t empennage would no t be s i g n i f i c a n t l y h igher than the t o t a l drag o f t h e t a i l f i n s and o f the gondola. W

2

Max. D i ameter 50.88

a

FIG.^ - SCHEMATIC LAYOUT OF NASA WIND-TUNNEL TEST ~KODEL, SHOWING

- .- HULL, GONDOLA AND T A I L , REPRODUCED FROVI FIG, 1 OF NASA TN D-1026.

0 .6 .E 1.0 .1 .2

_ _ ~ . ~

i ' I G . 2 - AXIAL FORCE C O € F F . C n V S . P R O P E L L E R

CP

.1c

.08

.06

.04

.02

c .6 .a 1.0 1.2

" F I G , 3 - PROPELLER POWER C O E F F . Cp Vs.

ADVANCE RATIO. REPRODUCED F R O M

FIG.^ D) OF NASA TN D-1026 PROPELLER ADVANCE RATIO. REPRODUCED FROM FIG.^ B) OF NASA TN D-1026.

3

http://COFF.Cn

Table I NASA Test Body Coordinates

Lenqth, in. Diameter, i n .

24.624 36.304 36.936 42.564 49.248 47.020 61.560 48.885 73.872 50.149 86.184 50.753 98.496 50.880

110.808 50.880 123.120 50.633 135.432 50.642 147.745 48.888 160.056 47.200 172.369 49.732 184.68 41.299 196.993 36.904 209.305 215.461 225.000 227.773 233.929 240.085 246.46

31.426 27.867 24.564 20.977 16.677 9.493

0

Figure 2 presents the net body/propeller d r a g coefficient Cg = F/qOVo.@j against propeller advance r a t io a = Uo/nd f o r 3 blade angles 15*, 20" and 25'; i t i s reproduced from F i g . 4d) Of Ref. 17. I t can be noted that negative d r a g Coefficient Co values represent net th rus t ; i t can be seen that more than enough net thrust was generated to tow two o t h e r identical bodies. Indeed the maximum thrust coefficient was Cg = -0.053 and the drag coefficient of the body ( w i t h o u t propeller) was

Figure 3 presents t h e propeller power Coeffi- cient Cp = HP/pn3d5 against the propeller advance r a t io a = Uo/nd fo r 3 blade angles 15", 20' and 25O; i t i s reproduced from F i g . 4b) of Ref. 17. I t must be noted t h a t th i s power Coefficient Cp does not include any body dimensions as such; i t has to be translated into the volume power coefficient CHP = HP/qoUoV0.66 so a s t o be used fo r a parametric fuselage analysis.

Figure 4 presents the propeller efficiency 'I

= ToUo/HP against the propeller advance ra t io a = Uo/nd fo r 3 blade angles 15", 20" and 25'; i t i s reproduced from Fig . 4c) of Ref. 17. To i s the actual measured thrust of the propeller, Uo i s the measured free-stream velocity and HP i s the measured shaft power t o the propeller. The seemingly impossible evidence i s t h a t propeller eff ic iency i s up t o 122% f o r advance ra t ios yielding zero net thrust (equilibrium f l igh t of body): t h i s i s con- clusive proof t h a t the wake-propeller does indeed operate within an area of lower veloci t ies and t h a t the eff ic iency does not have t o decrease, as pre- dicted by Lan and R~skam. '~

Cg = 0.021!

1.4

1.2

1.0

.8

.6

r

.4

.2

0 .6 .8 1.0 1.2 1.4

FIG.4 - P R O P E L L E R E F F I C I E N C Y VS P R O P E L L E R ADVANCE RATIO. R E P R O D U C E D FROM

F l G , q C ) O F NASA TN 0-1026 W'

However, the propulsive eff ic iency n = FoUo/HP i s down to a mere 103% a t equilibrium gecause the propeller-induced flow f i e ld aver the body causes additional d r a g , as compared t o the bare body, o r a so-called "thrust-deduction" of the propeller th rus t . The same propeller, mounted in a conven- t i o n a l free-stream ins ta l la t ion , achieved 73% prope l le r a n d propulsive eff ic iencies . T h u s despite the "thrust-deduction." the power g a i n of the wake- propeller was an impressive 41% over the freestream propel le r . This i l l u s t r a t e s the power that can be extracted from the fuselage wake's kinetic energy and which i s going t o waste today in the general aviation a i r c ra f t . I t can be added t h a t the "thrust-deduction" i s n o t a necessary ev i l : i t can be eliminated by designing the body shape t h r o u g h a complex i t e r a t ive process tha t includes the propel ler ' s e f fec t on the body pressure d i s t r i - bution and boundary-layer development. The f inal body shape i s such t h a t maximum pressure recovery i s achieved on the fuselage 's aftbody with the prope l le r , while the aftbody flow would be f u l l y separated without the propel le r ' s e f fec t . This maximum pressure recovery i s assumed t o be tha t given by the Goldschmied turbulent separation criterion.27

The NASA experimental t e s t d a t a are tabulated below in Table I1 f o r convenient reference; in addition, the volume power e f f ic ien t CHP has been computed from the propeller power coefficient Cp.

w

4

Table I1 Make-Pmpeller Test

Advance Axial Propeller Volume Propeller Ratio Drag Power Power Fff iciency

Experimental Parameters of URSA Fuselage/

a Coeff. Cg Coeff. Cp Coeff. CHP n ___ - ___ - PROPELLER BLADE ANGLE 6 = 15"

0.505 -0.053 0.050 0.0960 0.545 -0.041 0.048 0.0730 0.640 -0.017 0.0435 0.0412 0.685 -0.0085 0.0410 0.0317 0,700 -0,0060 0.040 0.0289 0.740 -0.0015 0.0370 0.0227

PROPELLER BLADE ANGLE B = 20'

0.63 -0.0370 0.077 0.0764 0.69 -0.0245 0.073 0.0552 0,765 -0.0135 0.069 0.0383 0.825 -0,0055 0.065 0.0287 0,860 -0,0035 0.063 0.0245 0.890 0.00 0.061 0.0214

PROPELLER BLADE ANGLE E = 25"

0.720 0.775 0.845 0.930 0.965 0.990

-0.0355 -0.0255 -0.0175 -0.0120 -0.0055 4 . 0 0 3 5

0.118 0.112 0.108 0.103 0.101 0.099

0.0785 0.0597 0.0444 0.0318 0.0279 0.0253

0.87 0.92 1.04 1.08 1.09 1.10

0.96 1.01 1.08 1.14 1.18 1.21

0.80 0.90 0.97 1.05 1.08 1.10

Table 111 below presents the parameters t o be used fo r the parametric aerodynamic design procedure. The net thrust o f the body/wake-propeller system T has been corrected f o r the t a i l and gondola drag, i . e . 0.003 has been added t o the given experimental

~4 values of the thrust coeff ic ient .

Advance Ratio a __

Table 111 FuselageNake-Propeller Parameters (Corrected f o r Tail & Gondola Drag)

Thrust Volume Power Coeff. CT Coeff. CHP

0.505 0.545 0.640 0.685 0.700 0.740

0.63 0.69 0.765 0.825 0.860 0.890

PROPELLER BLADE ANGLE 8 = 15"

0.056 0.044 0.020 0.0115 0.009 0.0045

PROPELLER BLADE ANGLE E = 20-

0.0960 0.0730 0.0412 0.0317 0.0289 0.0227

Figure 5 presents the plot o f CHP vs CT and Fig. 6 presents the plot of a vs. Cy.

I t design, 120 MPI As sug should

0.0:

0.08

0.07

a I: V

0.06 LL L L W 0 V

0.05 Id 3 0 a W 004 E ' 3 J 0 >

0.03

C.02

0.01

; t o be noted t h a t the propeller was fo r speeds of 100 MPH; a t speeds above

the propeller tip-speed becomes excessive. ;ked by McLemore, a larger number of blades ! used, such a s 6 o r 8.

NASA BODY/WAKE-PROPELLER TN 3-1026

NOTE: T H R U S T HAS B E E N - C O R R E C T E D FOR G O N D O L A R T A I L D R A G

0 0.01 0.02 0.03 0.04 0.05

THRUST CGEFF, C, F I G . 5 - VOLUME P O W E R C O E F F ~ C H P V S ~ T H R U S T

rnccc r C T C Y L l r ,

111. Aerodynamic Design Objectives

The aerodynamic design objective of t h i s brief preliminary study i s t o determine the Aerodynamic Efficiency Index AEI of the NASA fuselagelwake- propeller a i r c ra f t w i t h conventional NACA wings a t the maximum cruise speed.

These AEI values wiIl then be comnared t o t h e

0.040 0.0764

. . - corresponding data of U.S. Business , -Uti i i ty and Personal Aircraft (tabulated in Appendix I ) and of Spor t and Home-built Aircraft (tabulated i n

0.720 0.775 0.845 0.930 0.965 0.990 ier*i

PROPELLER BLADE ANGLE E = 25"

0.0385 0.0785 0.0285 0.0597 0.0205 0.0444 0.0150 0.0318 0.0085 0.0279 0.0065 0.0253

a

Appendix 111, a s computed from gross weight, maximum engine power and maximum a i r c r a f t speed.

From the data of Appendix I1 i t i s seen that one-seat a i r c ra f t range from 500 t o 800 Ib, from 115 t o 150 MPH and from 22 t o 60 HP. For the present study an appropriate fuselage diameter i s 40", so a s t o allow ample elbow room t o the p l lo t .

5

1.0

0.9

0

4 aL

W 0

> 4

7 0.8

0.7 n

5 0.6 - W a 0 LL a 0.5

0.4

4

NASA BODY/WAKE-PROPELLER \ TM 0-1026

6 Wind-Tunnel Model

From the data of Appendix I and 11, i t i s seen t h a t two-seat a i r c r a f t range from 950 t o 2400 l b , from 110 t o 276 MPH and from 50 t o 300 HP. For the present study an appropr ia te fuselage diameter i s 42" f o r tandem seat ing and 48" f o r s ide-by-side seat ing.

t h a t four-seat a i r c r a f t range from 2400 t o 3900 l b , f rom 138 t o 230 MPH and from 160 t o 235 HP. For the present study an appropr ia te fuselage diam- e t e r i s 55". From the data of Appendix I , i t i s seen t h a t s ix -seat a i r c r a f t range from 2750 t o 6775 l b , f rom 158 t o 302 MPH and from 200 t o 760 HP. For the present study an appropr ia te fuselage diam- e t e r i s 65", so as t o a l l o w room between the seats.

For t h i s p r e l i m i n a r y study, t he speed range has been l i m i t e d t o 140-180 MPH and the gross weight range has been l i m i t e d t o 800-3200 l b . The t a b l e below l i s t s the m a t r i x of con f igura t ions t o be analyzed.

From the data of Appendix I and 11, i t i s seen \ur/

NOTE:THRUST H A S BEEN CORRECTED

FOR GONDOLA 8 TAIL D R A G

I n o rder t o assess the roominess of t he fuselage . diameters, i t can be noted t h a t a cross-sect ion

of 24" x 36" i s o f ten quoted f o r spor t s ing le -seat a i r c r a f t a s s a t i s f a c t o r y t o accommodate the p i l o t , as aga ins t the selected 40" diameter. A cross- sec t ion of 42" x 40" i s quoted as sa t i s fac to ry f o r side-by-side two-seat a i r c r a f t , a s aga ins t the selected 48" diameter f o r the 2-seat side-by-side and 55" diameter f o r 4-seat a i r c r a f t . Wing aspect- r a t i o s of 8 and 10 w i l l be considered, as we l l as

I V . Parametric Ana lys is THRUST COEFF. c T The f i r s t wing parameter t o be considered i s

: i ~ . 6 - PROPELLER ADVANCE R A T I O V S ~ T H R U S T the wing aspec t - ra t i o ( A R ) ; F ig . 7 presents the general a v i a t i o n survey of wing aspec t - ra t i o aga ins t wing l i f t c o e f f i c i e n t from the data tabu- l a t e d i n Appendix 111. I t can be seen t h a t there

. -. C O E F F . cT

Number of Seats

Fuselage Diameter D i n .

Fuselage Length L i n .

Fuselage Volume v f t 3

Equiv. Area ~ 0 . 6 6 f t z

P rope l l e r Diameter d i n .

Gross W t . Wo l b

Speed UO MPH

doesn ' t seem t o be any trend: t he range i s from 6 t o 10, w i t h a preponderance o f p o i n t s between 7.25 and 8.50. For t h i s b r i e f p r e l i m i n a r y study, i n o rder t o keep a s o l i d wind-tunnel experimental basis, NACA AR = 10 and AR = 8 wings have been selected; the wind-tunnel t e s t data a re shown i n F ig . 8 (reproduced from Fig. 18 of Ref. 23) f o r the AR = 8 NACA 652-415 wing and i n F ig . 9 f o r the AR = 10 NACA 653-418 wing.

Table IV Sunmary o f A i r c r a f t Parameters

1

40

193

90.0

20.0

18.85

800

Tandem Side-by-Side

104.0 155.0

22.0 28.7

19.80 22.65

1400

140 - 180

2:5 1 ::3 ~ 5:ir 234.0 386.0 I 184

37.8 52.8 32.2

26.00 30.65 24.00

2400 3200 _ _ 100

j

- I W

6

9

MA"LE

2X

v

X

X

V

I , /

I

4 4

r x x x

/ /

/ I

x I'

X I ' * e /

I X I /

/ m I'

, 0 0 /

I m , m m I

0.2 0.3

W I N G L IFT COEFF, C, 3 M A X - S P E E D SEA-LEVEL

FIG,^ - GENERAL A V ~ A T I O N SURVEY : W I N G ASPECT RATIO AR VS.WING LIFT C O E F F . ~ a MAX.SPEED 8 SEA LEVEL

The l i f t / d r a g r a t i o s o f t h e two wings are p l o t - t e d i n F i g . 10; i t can be seen t h a t top l i f t / d r a g va lues a r e i n the l i f t c o e f f i c i e n t range from 0.25 t o 0.50.

The nex t quest ion i s t o examine the general a v i a t i o n p r a c t i c e i n terms o f wing l i f t c o e f f i c i e n t aga ins t a i r c r a f t speed: such a survey i s presented i n F ig . 11 from t h e data tabu la ted i n Appendix 111. Three p l o t s a re shown w i t h d o t t e d curves: t h e top p l o t i n d i c a t e s the h ighest l i f t c o e f f i - c i e n t s a c t u a l l y used i n general a v i a t i o n a i r c r a f t design and i t shows a very cons is ten t trend. I f CL = 0.30 i s chosen, i t can be seen t h a t i t i n t e r - sects t h e p l o t @ 140 MPH, y i e l d i n g a wing l o a d i n g of 15 PSF, wh i le if CL = 0.25 i s chosen, i t i n t e r - sects t h e p l o t @ 180 MPH, y i e l d i n g a wing l o a d i n g o f 21 PSF.

Table V presents t h e summary of wing areas and wing spans f o r t h e m a t r i x o f AR = 8 and AR = 10 and of 15 PSF and 21 PSF wing loading. The wing area ranges from 38.1 f t 2 (800 l b , 21 PSF) t o 211.2 - f t 2 (3200 l b , 15 PSF); t h e wing span ranges from

F r G . 8 - WING DRAG C O E F F . ~ ~ V S . W I N G LIFT C O E F F , C ~ , NACA 652-415 8 2415 WINGS OF AR=8. REPRODUCED FROM

FIG.^^ OF NACA REPORT 824 . 17.4 ft (800 l b , AR = 8, 21 PSF) t o 45.9 f t (3200 l b , AR = 10, 1 5 PSF).

Table V I presents t h e summary o f wing l i f t coef- f i c i e n t s CL f o r 140, 160 and 180 MPH and f o r 15 PSF and 21 PSF wing loading. The l i f t c o e f f i c i e n t ranges from 0.180 (180 MPH, 15 PSF) t o 0.415 (140 MPH, 21 PSF).

Table V I 1 presents t h e summary o f wing l i f t / d r a g r a t i o s (as obtained from Fig. 10) f o r 140, 160 and 180 MPH. AR = 8 and AR = 10 and f o r 15 PSF and 21 . . .. PSF wing loading. The r a t i o ranges f rom 32.0 (180 MPH, AR = 8, 15 PSF) t o 41.5 (140 MPH, AR = 10, 1 5 PSF).

From the gross weight W, and t h e l i f t / d r a g r a t i o , t h e wing d r a g is computed; a 10% wing/fuselage in te r fe rence drag i s added. The empennage drag i s assumed t o be 27.5% o f t h e wing drag. From t h e t o t a l drag F, t h e p r o p e l l e r t h r u s t c o e f f i c i e n t CT i s computed:

CT i_ -~ F Note: The Vo.66 va lues a r e given qoy0.66 i n Table I V

From Fig. 5, t h e corresponding va lues of t h e volume power c o e f f i c i e n t CHP are determined f o r 15O b lade angle (bes t p r o p e l l e r performance).

The summary o f CT and CHP va lues i s presented i n Table VI11 f o r convenient reference. The t h r u s t coeff. CT ranges from 0.0163 t o 0.0473, wh i le t h e volume power c o e f f . CHP ranges f rom 0.0370 t o 0.0780. It can be noted t h a t i n Table 111 t h e h ighes t experimental value of t h e power c o e f f i c i e n t i s CHP = 0.0960 and t h e corresponding t h r u s t c o e f f i c i e n t i s CT = 0.056.

7

Lif t coefficient CL

FIG.^ - WING DRAG C O E F F . C ~ VS.WING LIFT C O E F F . ~ ~ NACA 553-418 ~ N D 23Cl8 WINGS OF AR=10. R E P R O D U C E D FROM

- FIG.^^ O F NACA R p y , 8 2 4 .

NACA 65-015 AR=S \4lNG 8 i 0 0 . 2 0 . 4 0.6 0.8

W I N G LIFT COEFFICIENT CL . ~ ~~ .. ... . . ~ -.

F I G , ~ O - W I N G LIFT/DRAG RATIO VS.WING LIFT C O E F F . ~ ~ . NACA 65+18 A R = l 0 WING NACA 652-415 M = 8 VING

~ . ~ ~ W'

. ..- . . ... . . . I 300 - .

250 ~

~~. . . ;- i

.. . .. . . . . . .~ . . .~~ : 150 MAXIMUM AIRCRAFT SPEED. MPH !

..~. 120 ' j

Fic.11 - GENERAL AVIATION SURVEY : WING LIFT C O E F F . C ~ @ S E A - L E V E L Vs. M A X I M U M A I R C R A F T SPEED, MPH

8

W

Gross W t . W, l b

Wing Span, ~b i t 17.4 21.3

Wing Span, b ft j 22.9 : 28.1

Wing Span, b ft 119 .5 i 23.9

A R = 8 21 PSF j

AR = 10 15 PSF I

AR = i o 21 PSF i

800 I 1200 1400

92.4

66.6

27.1t

23.0

30.4

25.8

Table V Sunmry o f Wing Areas and Spans The next parameter t o be considered i s the pro- p u l s i v e e f f i c i e n c y i n the h igh t h r u s t area requi red

2400 1 3200 f o r , a i r c r a f t design. The p ropu ls i ve e f f i c i ency - i s computed as fo l lows: 158.4 . 211.2

I np = ( T ' FO)uO = cT + = CT + 0.0190

114.2 ' 152.3 HP CHP CHP

15'59 41.10 The bare body drag F, must be added t o the ne t t h r u s t T t o compute the t o t a l t h r u s t power; t he r a t i o o f 30.2 ' 34.9 t h i s power over the actual p r o p e l l e r power i s the p ropu ls i ve e f f i c i e n c y o f t he a i r c r a f t . Table IX

39'8 45.9 oresents the summary o f oroouls ive e f f i c i e n c v : i t 'can be seen t h a t i t - r a n g e s i rom 91.2% t o 96 . i% f o r t he 800 l b gross weight, from 87.6% t o 9 3 . a f o r t he 1250 l b , from 88.5% t o 92.8% f o r t he 1400 l b , from 85.0% t o 92.8% f o r the 2400 l b and from 85.6%

33'8 39'0

t o 93.8% f o r t he 3200 l b . These values can be compared t o 103% f o r t he propuls ion o f t he body alone and t o 65% f o r a t y p i c a l t r a c t o r a i r c r a f t .

The next step i s the computation o f the actual

Table VI S m r y o f Wing L i f t C o e f f i c i e n t s -_ _- 21

0.415 p r o p e l l e r power 0.317 0.25

Wing Loading PSF

140 MPH Speed 0.30 160 MPH 0.228 ' 180 MPH

CHP qpUoVo.66 550

0.180 .- __--_

HP =

Table VI1 Smmary of Wing L i f t /D rag Ra t ios

I Wing AR = 10 ! Wing AR = 8 Table X presents the summary of the p r o p e l l e r Speed vower r e s u l t s . The power ranges from 18.30 HP f o r

15 PSF ' 21 PSF .____ l 5 '"w the 800 l b a i r c r a f t @ 140 MPH t o 117 HP f o r t he 140 MPH 41.5 41.0 37.5 3200 l b a i r c r a f t @ 180 MPH.

180 : 35.5 j 40.0 32.0 I 37.0 160 39.0 41.5 36.5 ! 37.25

Table VI11 S m r y of Thrust and Power C o e f f i c i e n t s __- .- ___ -

I 800 1200 1400 I 2400 Gross W t . W, l b

AR = 10, 15 PSF Winq Uo = 140 MPH CT

CHP U, E 160 MPH CT

CHP U, = 180 MPH CT

CHP AR = 10, 2 1 PSF Winq

Uo = 140 MPH CT CHP

U, = 160 MPH

Uo = 180 MPH 2 P

AR = 8, 15 PSF Wins U, = 140 MPH

2 P U, = 160 MPH

U, = 180 MPH CT CHP

AR = 8, 21 PSF Winq U, = 140 MPH CT -

CHP U, = 160 MPH

Uo = 180 MPH CT 4 CHP

0.0262 0.0485 0.0212 0.0425 0.0184 0 3390

0.0265 0.0490 0.0200 0.0410 0.0163 0.0370

0.0290 0.0525 0.0227 0.0440 0.0205 0.0415

0.0298 0.0535 0.0223 0.0440 0.0177 0.0380

0.0357 0.0610 0.0290 0.0525 0.0251 0.0475

0.0361 0.0615 0.0271 0.0500 0.0223 0.0440

0.0396 0.0665 0.0310 0.0550 0.0278 0.0505

0.0406 0.0680 0.0304 0.0540 0.0241 0.0460

0.0319 0.0560 0.0260 0.0485 0.0224 0.0440

0.0323 0.0570 0.0244 0.0465 0.0185 0,0390

0.0353 0.0610 0.0277 0.0505 0.0249 0.0470

0.0363 0.0625 0.0272 0.0505 0.0215 0.0430

9

0.0415 0.0695 0.0338 0.0590 0.0292 0.0530

0.0420 0.0700 0.0317 0.0555 0.0260 0.0485

0.0460 0.0760 0.0361 0.0615 0.0324 0.0570

0.0473 0.0780 0.0354 0.0610 0.0280 0.0510

3200

0.0400 0.0675 0.0322 0.0570 0.0279 0.0510

0.0400 0.0675 0.0303 0.0540 0.0232 0.0450

0.0439 0.0730 0.0344 0.0600 0.0310 0.0550

0.0452 0.0750 0.0338 0.0590 0.0268 0.0495

Gross W t . Wo l b 800

Gross W t . Wo lb

1200 1400

A R = 10, 15 PSF Winq U, = 140 MPH

Uo = 140 MPH A R = 10, 15 PSF Winq

Uo = 160 MPH

U, = 160 Uo = 180

I I i !

0.932 0.896 i 0.909 0.946 ! 0.914 1 0.928

A R = 10, 21 PSF Winq U, = 140 MPH

Uo = 140 MPH Uo = 160 MPH Uo = 180 MPH

Uo = 140 MPH A R = 8, 15 PSF Wing

U i = 160 Uo = 180

0.922 0.933 1 0.961

0.890 ! i ; 0.881

0.928 0.951 0.954 , 0.938

0.914

AR = 8, 15 PSF Wing Uo = 140 MPH Uo = 160 Uo = 180

Uo = 140 MPH Uo = 160 Uo = 180

A R = 8, 21 PSF Winq

Uo = 160 MPH 0.948 0.909 0.952 1 0.927

0.912 0.876 0.939 0.915 0.966 0.937

i Uo = 180 MPH

Uo = 140 MPH Uo = 160 MPH Uo = 180 MPH

AR = 8, 21 PSF Wing

Table X Sumnary of Pmpeller Power, BHP

0.925 0.934

0.885 0.915 0.942

800

18.30 24.03 31.38

18.46 23.14 29.83

19.80 24.85 33.46

20.12 24.85 30.63

1200

25.23 32.66 42.05

25.49 31.10 38.95

27.63 34.19 44.79

28.13 33.55 40.80

In regard to propeller speed, the advance r a t io can be found from F i g . 6 and the speed can be computed; however, i t has been found tha t fo r speeds above 120 MPH the NASA 4-blades propeller requires excessively high tip-speeds. T h u s the propeller RPM has n o t been reported f o r the 140-180 MPH speed range. McLemore himself suggested the use of larg- e r number of propeller blades: 6 and 8 blades would be en t i re ly practical t o reduce the tip-speed while maintaining o r improving the efficiency. The new propfan technology can be p u t t o excellent use f o r the development of optimized wake-propellers i n the 200-350 MPH speed range.

Finally the aerodynamic eff ic iency index can be computed from the given gross weight and speed and the computed propeller power. The resu l t s a re given i n Table X I ; the range i s from 11.47 t o 17.95.

In order t o assess the meaning of the AEI num- bers, F i g . 12 presents the general aviation survey of the aerodynamic eff ic iency index against a i r - c r a f t velocity. I t can be seen tha t the range i s from a low of 5.00 to a high of 7.95.

2400 i 3200

0.870 0.874 0.895 0.898 W 0.909 0.919

0.871 0.874 0.913 0.913 0.928 0.938

0.855 0.862 0.896 0.890 0.902 0.909

0.895 0.925

0.850 0.892 0.921

1400 i 2400

30.29 49.50 39.35 63.06 50.89 I 80.74

I 30.83 37.74 45.12

33.00 40.94 54.39

33.81 40.60 49.75

49.86 59.12 73.89

54.13 65.66 86.87

55.53 65.13 77.72

3200

67.26 85.11

108.50

67.18 80.54 95.72

72.62 kd 89.48

117.00

74.64 88.00

105.35

On the other hand, the data of Table XI are presented in the plot of Fig. 13 f o r convenient visual evidence: i t seems t o be c lear tha t i n the 140-180 MPH, 800-3200 lb range the NASA fuselage/wake- propeller a i r c r a f t o f fe rs substantial aerodynamic efficiency index improvements over 100%.

I t can be argued tha t the AEI data of F i g . 12 represent actual operational "d i r ty" a i r c r a f t while the AEI data of F i g . 13 represent "clean" wind- tunnel model performance, with neglect of several miscellaneous drags.

A 15% speed reduction is a most generous allow- ance f o r such miscellaneous drags; a s an example, f o r the single-seat Moni a i r c ra f t . the fixed land- ing gear causes a speed reduction of only 5 MPH (4.1%) a s compared t o the retracted gear. T h u s the highest points of Fig. 13 would drop t o the 15.0 level and the lowest points t o the 10.0 level ; the aerodynamic efficiency index improvements a re s t i l l very substant ia l , a s compared t o the Fig. 12 range from 5.0 t o 7.95.

10

Table X I Summry of Aemdynamic Efficiency Index AEI

Gross W t . W, l b I 800 1200

AR = 10, 15 PSF Wing ! i ~ u, = 180 i

1 i

16.28 14.15 12.23

U, = 140 MPH U, = 160

4

AR = 10, 21 PSF Winq 16.09 14.72 12.86

U, = 140 MPH Uo = 160 U, = 180

U, = 140 15.09 U, = 160 13.72 u, = 180 11.47

AR = 8, 15 PSF Wing

AR = 8, 21 PSF Wing U, = 140 MPH 14.77 U, = 160 13.72

12.50 Uo = 180 14.06 i i

17.74 15.63 13.66

17.53 16.35 14.73

16.21 14.95 12.87

15.88 15.23

1400 2400

17.26 17.95 15.08 16.14 13.12 14.22

16.86 17.89 15.80 17.23 13.79 15.57

15.84 16.52 14.57 15.58 12.35 13.26

16.16 15.69 14.81

15.45 14.68 13.45

3200

17.72 16.00 14.14

17.71 16.93 15.00

16.47 15.23 13.10

16.00 15.50 14.53

F I G , 1 2 - GENERAL A V I A T I O N SURVEY : A E R O D Y N A M I C E F F I C I E N C Y I N D E X AEI V S . M A X I M U M SPEED

11

c - q _ c l 4 0 MPH AR=10.21 PSF -

-160 MPH AR=10.21 PSF

A--140 MPH AR=8.15 PSF u

!3 v- - - - - - - - -

4-- -b-- - - - - - _-_-- - - - /-- \I80 MPH AR=8,15 PSF 8' , I P

140 MFH 160 MPH 180 MPH

V

II ~~~~. .

c , P'

~.

1 2 3

AIRCRAFT GROSS WEIGHT LE x 1000

F l G , l 3 - NASA FUSELAGE/WAKE-PROPELLER : AIRCRAFT GROSS WEIGHT FOR 140 .

V . Fuselaqe Power Ra t io

The Dower reau i red f o r the oroou ls ion of t he fuselage i s the ' l a r g e s t c o n t r i b k o k t o the t o t a l conventional a i r c r a f t power. For a t y p i c a l s ing le - engine t r a c t o r genera l -av ia t ion a i r c r a f t , t he drag d i s t r i b u t i o n i s as fo l lows:

To ta l A i r c r a f t Drag Coeff . COA = 0.0275 100%

Empennage Drag Coeff. CDA = 0.0025 9%

Fuselage Drag Coeff.

(Note:

CDA = 0.0150 54%

CDA i s based on wing area)

The fuselage i s the l a r g e s t c o n t r i b u t o r t o a i r c r a f t drag; desp i te t h i s obvious f a c t , aerodynamic o p t i - m iza t i on o f t he fuselage has n o t received much a t t e n t i ~ n . ? ~ Most of t he research has been focused on the wing, which accounts f o r o n l y 36% o f t he t o t a l drag. I n conclusion, w i t h the assumption t h a t t he p r o p e l l e r t h r u s t can o n l y make the fuse- l age drag higher, then the t y p i c a l fuselage i s respons ib le f o r 54% of the power, a t the l e a s t .

It i s most i n t e r e s t i n g t o examine the fuselage power r a t i o (fuselage power/ total power) f o r t he NASA fuselage/wake-propel l e r con f igura t ion w i t h NACA AR = 10 and AR = 8 wings, w i t h wing load ings o f 15 and 21 PSF.

From F ig . 5 i t i s seen t h a t the fuselage power coef f . CHP (cor rec ted f o r gondola and t a i l drag)

AERODYNAMIC EFFICIENCY INDEX A E I Vs , . 18G MPH SPEED RANGE

i s 0.0190 f o r zero n e t t h r u s t , i . e . f o r e q u i l i b r i u m f l i g h t cond i t ions , w i t h 103% propu ls ive e f f i c i ency . The Dower r a t i o 0.019/CHP has been ComDuted and i s dresented i n Table XI1 f o r t he f i v e arass weights, t he two wing aspect r a t i o and the two-wing loadings. 4

It can be seen t h a t the range i s from a maximum o f 0.513 (corresoondina t o A E I = 12.86) t o a min i - ,~~ - mum o f 0.243 (corresponding t o A E I = i6.52). A lso i t can be noted t h a t t he maximum A E I = 17.95 corresponds t o a r a t i o of 0.273 and the minimum A E I = 11.47 corresponds t o a r a t i o of 0.457.

I n conclusion, t he NASA fuselage/wake-propeller con f igura t ion has made i t poss ib le t o reduce the fuselage power c o n t r i b u t i o n from 54% t o 25%.

V I . Assessment of 4-Seat Pusher A i r c r a f t

The 4-seat 2400 l b con f igu ra t i on has turned ou t t o have t h e h ighes t A E I values, as shown above i n Table X and i n Fig. 13, i n the speed range between 140 and 180 MPH. Also the 4-seat a i r c r a f t i s t he most popu lar general a v i a t i o n model to-date; economy, roominess and range are among the most des i rab le cha rac te r i s t i cs .

An idea l engine would be the Teledyne Continen- t a l GR-36 85 HP r o t a r y u n i t , w i t h RPM i n the area Of 7000, a d r y weiaht of 110 l b and a soec i f i c f u e l consumption o f 0:45 lb/HP hr, as keported by DeMeis.25 As shown i n Table I X , t he power requ i red @ 180 MPH w i t h 21.0 PSF winq l oad inq i s 77.72 HP f o r AR = 8 and 73.89 f o r AR = - l o wing:

An assessment should be c a r r i e d ou t aga ins t t he l a t e s t and most advanced 4-seat pusher a i r c r a f t a v a i l a b l e today; one such a i r c r a f t i s t he Prescot t Pusher, as descr ibed by Cox.26 Table XI11 below ~

12

p r e s e n t s a s ide-by-side l i s t i n g of a l l r e l evan t parameters f o r the two des igns , so a s t o allow easy comparison.

~ Gross W t . Wo l b ! 800 1200

Table X I 1 Summary o f Fuselage Power Ra t io

1400

Fuel Tank 45 gal . = 270 l b

P r o p e l l e r d i a . 26" d i a . RPM 7000

No. of Blades 8

A R = 10, 15 PSF Wing I i 0.390 0.311 I 0.339

0.391 0.358

0.362 0.447 0.400 , 0.431

Uo = 140 MPH U, = 160 U, = 0.487

I

A R = 10, 21 PSF Wing Uo = 140 MPH 0.387 0.308 0.333 U, = 160 0.463 0.380 0.408 Uo = 180 0.513 0.431 0.487

AR = 8, 15 PSF Winq Uo = 140 MPH Uo = 160 U, = 180

U, = 140 MPH U, = 160 U, = 180

A R = 8, 21 PSF Wins

0.362 0.431 0.457

0.355 0.431 0.500

0.285 0.311 0.345 0.376 0.376 0.404

0.279 0.304 0.376 0.441

0.352 0.413

2400

0.273 0.322 0.372

0.271 0.342 0.391

0.250 0.309 0.333

0.243 0.311 0.372

Table X I 1 1 Tabula t ion of 4-Seat Pusher Aircraft Parameters

3200

0.281 0.333

0.281 0.352 0.422

0.260 0.316 0.345

0.253 0.322 0.383

Parameter ! Fuselage/Wake-Propeller I P r e s c o t t Pusher

No. of S e a t s

Gross Weight, l b

4

2400

Fuselage Length, f t 22.08

Fusel age Cross-Sect i on

Wing Area, f t 2 114.2

%&d 55" d iameter

I

Wing Span, f t

Wing Aspect-Ratio

30.2

8.0

Wing Loading, PSF 21.0

Power @ Speed

Engine: Type Power RPM Dry Weight, l b

77.72 HP @ 180 MPH 65.13 HP @ 160 MPH 55.53 H P @ 140 MPH

GR-36 Continental 85 HP

4

2400

20.25

42" width x 40" h t .

111.0

29.33

7.75

21.6

180 HP @ 184 MPH

0-360 Lycoming 180 HP 2700 257

45 g a l . = 270 1 b

72" d i a . 2700

2

0.42 lb/HP hr

12.0 g a l . / h r

15.3

690

13

The economy i s enhanced by the 85 HP engine aga ins t the 180 HP; t he roominess i s enhanced by the 55" d ia . cross-sect ion aga ins t the 42" x 40"; t he range i s increased from 690 t o 1300 mi les, us ing the same 45-gal. tank.

The A E I value f o r the Prescot t Pusher @ 184 MPH i s 6.5 wh i l e i t i s 15.57 f o r AR = 10 21 PSF wing and 14.81 f o r AR = 8 21 PSF wing; i t i s c l e a r t h a t a very subs tan t ia l imDrovement p o t e n t i a l has been made - a v a i l a b l e by the NASA fuselage/wake- p r o p e l l e r con f igura t ion .

V I I . Conclusions

1.

2.

3 .

4.

5.

1.

2.

3.

The NASA fuselage/wake-propel l e r con f igura t ion , as app l ied t o a m a t r i x o f a i r c r a f t designs i n the 140-180 MPH speed range, has shown con- c l u s i v e l y t h a t a 50% power reduc t ion i s a prac- t i c a l p o s s i b i l i t y , f o r the same gross weight and speed; the bas i s f o r t h i s comparison i s provided by a survey of 76 general a v i a t i o n a i r c r a f t (Appendix I and 11).

The propu ls ive e f f i c i e n c y f o r t he a i r c r a f t design m z t r i x ranges from 85% t o 96%; t h i s can be s u b s t a n t i a l l y improved by e l i m i n a t i o n o f fuselage drag increments induced by the p r o p e l l e r through fuselage shape op t im iza t i on . Th is i s t he area where f u r t h e r t h e o r e t i c a l and wind-tunnel research i s h i g h l y recommended. It can be remembered t h a t conventional t r a c t o r a i r c r a f t have 65% p ropu ls i ve e f f i c i e n c i e s .

While the convent ional fuselage accounts f o r 54% o f t he t o t a l drag and power, the fuselage power of t he NASA conf igura t ion ranges from 51% down t o 24%; t h i s represents a subs tan t ia l design improvement over conventional p r a c t i c e .

The most popu lar general a v i a t i o n a i r c r a f t i s t he 4-seat model i n the 140-180 MPH speed range. A d e t a i l e d comparison has been c a r r i e d ou t aga ins t t he 4-seat Prescot t Pusher, f o r t he same 2400 l b gross weight and 180 MPH speed, i n d i c a t i n g 78 HP aga ins t 180 HP, 5.8 ga l l ons /h r aga ins t 12.0, 30.9 MPG aga ins t 15.3 and 1300 m i l e s range aga ins t 690.

Clvde McLemore was 20 vears ahead o f h i s t ime: hi; e x c e l l e n t work w i i l bear f r u i t i o n now f o r t he general a v i a t i o n indus t ry .

L i s t of References

Froude, W., "Discussion o f Paper by W.J.M. Rankin," Trans. I n s t . Naval A rch i tec ts , Vol. 6, 1865, pp. 35-37.

Smith. A.M.O. and Roberts. E.. "The J e t A i r - p l a n e ' U t i l i z i n g Boundary-Layer ' A i r f o r Propul- sion," J . Aero. Sciences, Vo1. 14, 1947.

Kuchemann, D. and Weber, J., "Aerodynamics of Propulsion," McGraw H i l l Book Company, New York, N.Y., 1953.

4. Edwards. J . 8.- 'Fundamental Aspects o f Propul- s ion f o r Laminar Flow Aircraf t , ' Boundary-Layer and Flow Cont ro l , G . V. Lachmann, ed. Pergamon Press, 1961, pp. 1077-1122.

5. Edwards, B., "Laminar Flow Control-Concepts, Experiences and Speculat ions," AGARD R-654, 1977, pp. 4.1-4.41.

6. Davidson, I. M.. "Some Notes on A i r c r a f t Pro- pL l s ion b j Aage Regenerztion, In te rna t iona l - Conqress. & Suosonic Aerona.t ics, hew Yom Academy o f Sciences, ?member 1968, pp. 641- 651.

7. Goldschmied, F. R., " In tegra ted Hu l l Design, Boundary-Layer Control and Propuls ion o f Sub- merged Bodies," A I A A Journal of Hydronautics, Vol. 1, No. 1, J u l y 1967, pp. 2-11.

8. Goldschmied, F. R., " In tegra ted H u l l Design, Boundary-Layer Control and Propuls ion f o r Sub- merged Bodies: Wind-Tunnel V e r i f i c a t i o n , " A I A A Paper 82-1204, 1982.

9. Goldschmied, F. R., "Jet-Propuls ion o f Subsonic Bodies w i t h J e t Total-Head Equal t o Free- Streams," A I A A Paper 83-1790, 1983.

10. Goldschmied, F. R., "On the Aerodynamic Optimi- z a t i o n of Mini-RPV and Small GA A i r c r a f t , " A I A A Paper 84-2163, 1984.

11. Huang, T. T., Wang, H. T . , S a n t e l l i , N. and Groves, N . C. , "Propeller/Stern/Boundary-Layer I n t e r a c t i o n i n Axisymmetric Bodies: Theory and Experiment," David W. Tay lo r Naval Ship R&D Center Report 76-0113, December 1976.

12. Farn, C.L.S., Goldschmied, F. R. and Whir- low. D . K.. "Pressure D i s t r i b u t i o n P red ic t i on f o r . Two-Dimensional Hydro fo i l s w i t h Massive Turbu len t Separation," A I A A Journal o f Hydro- nau t i cs , Vol. 10, July-Aug. 1976, pp. 95-101.

w 13. Goldschmied, F. R., "Comment on Separation Model f o r Two-Dimensional A i r f o i l s i n Transonic Flow," A I A A Journal , Vo1. 12, No. 7, p. 1138, 1985.

14. Blascovich, J . D., "Charac te r i s t i cs o f Sepa- ra ted Flow A i r f o i l Ana lys is Methods," A I A A Paper 84-0048, Jan. 1984.

15. Goldschmied. F. R . . "Comments on An Inverse Boundary-Layer Method f o r Compressible Laminar and Turbulent 8oundary-Layers," A I A A Journal o f A i r c r a f t , Vol. 14, No. 5, p. 509, May 1977.

16. Goldschmied, F. R . , "Comments on Experimental I n v e s t i ga ta i on o f Subsonic Turbulent Separated Boundary-Layers on an A i r f o i l ,'I A I A A Journal of A i r c r a f t , Vo1. 14, No. 9, pp. 927-928, Sept. 1977.

17. McLemore, H. Clyde, "Wind-Tunnel Tests o f a I/ZO-Scale A i r s h i p Model w i t h Stern Propel- le rs , " NASA TN 0-1026, Jan. 1962.

18. Mi ley , S. J . and von Lavante, E., "P rope l l e r Propuls ion System I n t e g r a t i o n - S ta te o f Tech- nology Survey," NASA Cont rac tor Report CR-3882, J u l y 1984.

14

19. Lan, C. E. and Roskam, J., "A i rp lane Aerody- namics and Performance," The U n i v e r s i t y of Kansas, Lawrence, Kansas, 1981.

20. Freeman, Hugh e., "Force Measurements on a I/4O-Scale Model of t he U.S. A i r s h i p Akron," NACA Report 432, 1932.

& 21. Freeman, Hugh B., "Measurements o f Flow i n

the Boundary Layer o f a 1/40-Scale Model o f t he U.S. A i r s h i p Akron," NACA Report 430, 1932.

22. Abbott, I r a H., "The Drag o f Two Streamline Bodies as Af fec ted by Protuberances and Append- ages," NACA Report 451, 1932.

23. Abbott, I r a H., von Doenhoff, A l b e r t E. and St ivers , Louis S., "Summary of A i r f o i l Data," NACA Report 824, 1945.

24. Parsons, J . S., Goldschmied, F. R. and Goodson, R. E., "Shapinq o f Ax i smmet r i c Bodies f o r Min- imum Drag i n Incompressible Flow," A I A A Journal o f Hydronautics, Vol. 8, No. 3, pp. 100-107, J u l y 1974.

25. DeMeis, Richard, "Rotary Grows Up," Aerospace America, June 1986, pp. 48-51.

26. Cox, B i l l , "Pusher f o r t he Zlst Century," Home- b u i l t A i r c r a f t , May 1986, pp. 30-63.

27. Goldschmied, F. R., "An Approach t o Turbu len t Incompressible Separation Under Adverse Pres- sure Gradients," A I A A Journal o f A i c r a f t , Vol. 2, March/Apri l 1965, pp. 108-115.

~

APPENDIX I AERODYNAMIC EFFICIENCY INDEX OF U.S. BUSINESS, UTILITY AND PERSONAL AIRCRAFT

Max. Power I Max. Speed

I HP I MPH

Designat ion I Seats 1 Wt I

BEECH AIRCRAFT CORP.

C-23 Sundowner C-24R Sie r ra F-33A Bonanza V-358 Bonanza A-36 Bonanza B-36TC Bonanza B-55 Baron E - 5 5 Baron 58 Baron 58P Baron 58TC Baron 860 Duke 76 Duchess C / R

k-' 77 Skipper

CESSNA AIRCRAFT CO.

152 152 Aerobat 172 Skyhawk 172 Cutlass 172 Cutlass RG 182 Skylane 182 Skylane RG Turbo 182 Skylane Turbo 182 Skylane RG 185 Skywagon 206 S t a t i o n a i r 6 T-206 Turbo S t a t i o n a i r 6 207 S t a t i o n a i r 8 T-207 Turbo S t a t i o n a i r 8 210 Centurion T-210 Turbo Centur ion P-210 Centurion T-303 Crusader 340-A 402-C Business L i n e r 402-C U t i l i l i n e r 414 Chancel lor 421 Golden Eagle

MAULE AIRCRAFT CORP.

M5-180C MS-210TC M5-235C M6-235

4 4- 6 4- 5 4- 5 4- 6

6 4- 6 4- 6 4- 6 4- 6 4- 6 4- 6

4 2

2 2 4 4 4 4 4 4 4 6 6 6 8 8 6 6 6 6 6

6- 10 6- 10 6-8 6-8

4 4 4 A

I

!

I

i

I i I I

I

!

! i

I I

!

2450 2750 3400 3400 3650 3850 5100 5300 5500 6200 6200 6775 3900 1675

1675 1675 2407 2558 2658 3110 3110 3100 3112 3362 3612 3616 3812 3816 3812 4016 4016 5176 6025 6885 6885 6785 7500

2400 2500 2500 2500

~

15

180 200 285 285 300 300 520 570 600 650 650 7 60 360 115

108 108 160 180 180 230 235 235 235 300 300 310 300 310 300 310 310 500 620 650 650 620 7 50

180 210 235 735

!

I I I i

~

! !

I i i ~

i

i

1 i !

!

145 158 198 198 212 230 216 230 239 256 277 275 191 121

122 122 138 140 161 163 180 182 189 169 169 192 165 185 197 226 208 226 264 242 242 258 277

156 196 172 150

AE I

5.26 5.77 6.28 6.28 6.86 7.85 5.63 5.62 5.82 6.49 7.02 6.52 5.50 4.68

5.03 5.03 5.52 5.29 6.32 5.86 6.34 6.38 7.00 5.03 5.41 5.97 5.57 6.05 6.65 7.75 7.16 6.22 6.82 6.81 6.81 7.51 7.36

5.53 6.20 4.86 4.24

Desi gnat ion

800 1650 2400 1808

506 500 560

2400 1800

MOONEY AIRCRAFT CORP.

M20J201 M20K231

P I P E R AIRCRAFT CORP.

60 100 300 160 22 22 22

180 160

PA-28-161 War r io r 2 PA-28-181 Archer 2 PA-28RT-201T Arrow 4 PA-28-236 Dakota

500 700

1275 1700 1000 1500 950

1250

PA-32-301 Saratoga PA-32R-301 Saratoaa SP

40 44

100 180

50 118 65

100

PA-3ZR-301T Saratzga SP PA-32-301T Turbo Saratoga PA-34-220T Seneca 3 PA-38-112 Tomahawk 2 PA-46-310P Ma l ibu PA-60-602P Aeros ta r PA-60-700P Aeros ta r

Seats

4 4

4 4 4 4 6

10 7 6 6 6 6 6 2 6 6 6

Gross W t l h

2740 2900

2440 2550 2990 3000 6500 7000 7200 3600 3600 3600 3600 4750 1670

Max. Power HP

200 225

160 180 200 235 650 700 700 300 300 300 300 440 112

Max. Speed 1 AEI MPH

201 7.32 231 7.94 \ray

5.92 5.61 7.63 5.60 6.72 6.78 7.35 5.52 5.84 6.51 6.06 6.37 5.03 7.35 7.81 7.24

NOTE: Weights, Powers and Speeds quoted from: A v i a t i o n Week & Space Technology, March 12, 1984, p. 144.

APPENOIX I1 AERODYNAMIC EFFICIENCY INDEX OF U.S. SPORT AND HOME-BUILT AIRCRAFT

Desi gnat i o n Ref.

T a y l o r Mini-Imp T a y l o r B u l l e t Swearingen SX-300 Fa1 co Windex 1100 Moni Moni Tr i -Gear P r e s c o t t Pusher Whisper S t a r - L i t e S i l h o u e t t e Lanca i r (Lancer) 200 Glasai r FT-180 Sparrow Hawk MK I 1 cozy Whitehawk-65 Whi tehawk- 100

Seats I Max. HP Power Gross W t

l b Max. Speed

MPH

182

A E I

5.33 6.60 5.88 6.38 7.66 ~

7.27 7.80 6.50 6.75 4.00 5.09 7.24 6.04 5.86 6.80 6.12 6.06

- NOTE: Weights, Powers and Speeds a re quoted f rom t h e f o l l o w i n g sources as l i s t e d below.

(1) Spor t Av ia t i on , August 1984, p. 24. (10) Spor t Av ia t i on , A p r i l 1986, p. 20. (2 ) Homebuil t A i r c r a f t , May 1986, p. 24. (11) Spor t A v i a t i o n , March 1985, p. 42. (3 ) Homebuil t A i r c r a f t , October 1985, p. 29. (12) Spor t A v i a t i o n , A p r i l 1985, p. 14. (4 ) Homebuil t A i r c r a f t , August 1985, p. 27. (13) Spor t Av ia t i on , August 1985, p. 15. (5 ) B u i l d e r ' s s p e c i f i c a t i o n s . (14) Spor t Av ia t i on , June 1985, p. 52. (6) B u i l d e r ' s s p e c i f i c a t i o n s . (15) Homebuil t A i r c r a f t , April 1986, p . 29. ( 7) B u i l d e r ' s spec i f i ca t i ons . (16) B u i l d e r ' s s p e c i f i c a t i o n s . (8) Homebuil t A i r c r a f t , May 1986, p. 62. (17) B u i l d e r ' s spec i f i ca t i ons . ( 9 ) Sport Av ia t i on , March 1986, p . 47.

NOTE: 76 a i r c r a f t see l i s t e d i n Appendix I and 11. Category A E I between 4 and 5 has 4 a i r c r a f t o r 5.2% ... Category A E I between 5 and 6 has 26 a i r c r a f t o r 34.2% ... Category A E I between 6 and 7 has 28 a i r c r a f t o r 36.8% ... Category A E I between 7 and 8 has 18 a i r c r a f t o r 23.6% ... The lowes t AEI va lue i s 4.50 and t h e h ighes t 7.94.

-

16

~

APPENDIX 111 WING PARAMETERS OF U.S. BUSINESS, UTILITY, PERSONAL SPORT & HOME-BUILT AIRCRAFT

Designat ion

‘ a 4 BEECH AIRCRAFT CORP.

C-23 Sundowner C-24R S i e r r a F-33A Bonanza V-356 Bonanza A-36 Bonanza B-36TC Bonanza 8-55 Baron E-55 Baron 58 Baron 58P Baron 58TC Baron 860 Duke

CESSNA AIRCRAFT CO.

151 152 Aerobat 172 Skvhawk

182 Skylane 182 Skylane RG Turbo 182 Skylane Turbo 182 Skylane RG 185 Skywagon 206 S t a t i o n a i r 6 T-206 Turbo S t a t i o n a i r 207 S t a t i o n a i r 8 T-207 Turbo S t a t i o n a i r 8

T-210 Turbo Centur ion P-210 Centur ion T-303 Crusader 340-A 402-C Business L i n e r 402-C U t i l i n e r 414 Chancel lor

4 210 Centur ion

421 Golden Eagle

MAULE AIRCRAFT CORP.

M5-180C M5-219TC M5-235C M6-235

MOONEY AIRCRAFT CORP.

M20J201 M20K231

Wing Wing Aspect Loading Ra t io I PSF

7.36 7.36 6.20 6.20 6.20 7.60 7.17 7.17 7.17 7.59 7.59 7.21 7.97 6.92

6.88 6.93 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.44 7.73 7.73 7.73 8.03 7.93

8.61 8.61

a. 61

9.48

6.04 6.04 6.04 6.37

7.44 7.44

16.78 18.83 18.78 18.78 20.16 20.47 25.60 26.60 27.60 32.96 32.96 31.82 21.54 12.88

10.46 10.46

14.70 15.27 17.87 17.84 17.81

19.32 20.75

21.90 21.93

22.94 22.94 27.35 32.74 30.49 30.49 30.04 33.18

13.83

17.88

20.78

21.78

15.20 15.80 15.80 14.40

15.60 16.50

Power Loading

1 b/HP

13.60 13.75 11.92 11.92 12.16 12.83 9.80 9.30 9.16 9.53 9.53 8.91

10.83 14.50

15.50 15.50 15.00 14.20 14.70 13.50 13.20 13.10 13.20 11.20 12.00 11.60 12.70 12.30 12.70 12.90 12.90 9.70 9.20

10.60 10.60 10.90 10.00

13.30 11.90 10.60 10.60

13.70 13.80

Wing L i l CL @ S.L.

@ Max. Speed

0.300

0.185 0.185 0.173 0.148 0.211 0.192 0.184 0.191 0.165 0.159 0.229 0.338

0.280

0.271 0.271

0.291

0.258 0.212 0.207 0.173 0.259 0.278 0.218 0.308 0.246 0.218 0.172 0.203 0.205 0.178 0.200 0.200 0.171 0.165

0.282

0.228

0.241 0.159 0.205 0.248

0.148 0.118

:oef f .

CL @ 6000 f t

0.359 0.335 0.221 0.221 0.207 0.177 0.252 0.230 0.220 0.228 0.197 0.190 0.274 0.404

0.324 0.324 0.337 0.348 0.273 0.308 0.253 0.248

0.288 0.190 0.245 0.297

0.177 0.141

17

I Wing Designation Aspect

Ra t io

PIPER AIRCRAFT CORP.

PA-28-161 Warrior 2 PA-28-181 Archer 2 PA-28RT-201T Arrow 4 PA-28-236 Dakota PA-31-325 Navajo C / R PA-31-350 Ch ie f t a in PA-31P-350 Mojave PA- 32- 30 1 Sara toga PA-32R-301 Saratoga SP PA-32R-301T Saratoga SP PA-32-301T Turbo Saratoga PA-34-220T Seneca 3 PA-38-112 Tomahawk 2 PA-46-310P Malibu PA-60-602P Aeros t a r PA-60-700P Aeros t a r

SPORT & HOME-BUILT

Tay lo r Mini-Imp Tay lo r Bul l e t Swearingen SX-300 Fa1 co Windex 1100 Moni Moni Tri-Gear P r e s c o t t Pusher Whisper S t a r - L i t e Si1 houe t t e Lanca i r (Lancer) 200 G1 a s a i r FT- 180 Sparrow Hawk MK I1

7.20 7.20 7.37 7.37 7.23 7.23 8.35 7.35 7.35 7.35 7.35 7.25 9.27

10.56 7.56 7.56

8.25 8.25

6.40 16.20 10.00 10.00 7.60 7.10 8.10

12.90 7.27 6.65 8 .25 7.12 9.00 9.00

a. 32

Wing Loading

PSF

14.30 15.00 17.00 17.60 28.30 30.50 30.30 20.20 20.20 20.20 20.20 22.70 13.40 23.40 33.70 35.40

10.75 11.78 33.60 16.81 6.34 6.66 7.46

21.60 22.20

9.33 16.80 20.90

7.14 15.60 6.78 8.93

8.77

Power Loading 1 b/HP

15.20 14.10 14.50 12.70 10.00 10.00 10.28 12.00 12.00 12.00 12.00 10.80 14.90 13.20 10.34

9.02

13.30 23.50

8.00 11.30 23.00 22.70 25.45 13.30 11.25 12.00 15.90 12.75 9.44

20.00 13.60 14.60 12.50

Wing L i CL @ S.L.

@ Max. Speed

0.260 0.258 0.168 0.247 0.168 0.178 0.159 0.258 0.232 0.187 0.217 0.177 0.319 0.123 0.160 0.147

0.185 0.200 0.168 0.143

Coeff.

CL @ 6000 f t

0.311 w 0.308 0.201 0.295 0.201 0.213 0.190 0.308 0.278

0.301 0.170 0.166 0.270 0.208 “.a/ 0.127 0.125

18

aiaa- aerodynamic design of low speed aircraft

Documents

t e s t

t t e r t h i s

f t ranges

f fo t

f t propeller speed

f t fuselage fineness

tha t f o r conventional

hp f o r