Wind tunnel investigation of longitudinal aerodynamic characteristics of a powered four-duct-propeller VTOL model in transition

NASA TECHNICAL NOTE

N A S A T N D-3193

- - e . . - -

( % , .I

LOAN 'CbPY: RE1

AFWL (WLll K I R T W D AFB,

WIND-TUNNEL INVESTIGATION OF

LONGITUDINAL AERODYNAMIC

CHARACTERISTICS OF A

POWERED FOUR-DUCT-PROPELLER

VTOL MODEL I N TRANSITION

by Kenneth P. Spreemann

Langley Research Center

Langltvl Stution, Humpton, N A T I O N A L A E R O N A U T I C S A N D S P A C E TECH LIBRARY KAFB, N M NASA TN D-3192 WIND-TUNNEL INVESTIGATION OF LONGITUDINAL AERODYNAMIC CHARACTERISTICS O F A POWERED FOUR-DUCT-PROPELLER IN TRANSITION VTOL MODEL By Kenneth P. Spreemann Langley Research Center Langley Station, Hampton, Va.

NATIONAL AERONAUT ICs AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $1.55 WIND-TUNNEL INVESTIGATION OF LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A POWERED FOUR-DUCT-PROPELLER VTOL MODEL IN TRANSITION By Kenneth P. Spreemann Langley Research Center SUMMARY An investigation of the aerodynamic characteristics of a powered four-duct- propeller model of a VTOL airplane has been conducted in the 17-foot test sec- tion of the Langley 300-MPH 7- by 10-foot tunnel. The model was tested through an angle-of-attack and transition-speed range from hover to normal forward flight. Duct deflection angles from -5O to 90' were investigated at appropriate speeds and power conditions.

The results showed that the first configuration tested was longitudinally unstable at high thrust coefficients in the cruise condition. Longitudinal sta- bility was achieved by diminishing the destabilizing effects of the front pair of ducts by reducing the size of the front fairings between the ducts and the fuselage. Duct lip stall, which was aggravated by the low Reynolds numbers Of the tests, was delayed to higher angles of attack by increasing the lower lip radius of the duct. Delaying the stall on the lower duct lip also helped to Trim and control requirements in the delay stalling on the upper duct surface.

transition-speed range might best be attained by a combination of differential thrust and deflection of vanes within the ducts.

INTRODUCTION An investigation to study the longitudinal aerodynamic and control charac- teristics of a four-duct-propeller VTOL transport airplane configuration has been conducted in the 17-foot test section of the Langley 3OO-MFH 7- by 10-foot tunnel by using a l/?-scale model. The four ducts were arranged in pairs at the front and rear of the model. This arrangement, as pointed out in refer- ence 1, provides good pitch and roll control in hovering and transition by means of differential thrust of the appropriate pairs of ducted propellers. In the transition-speed range this type of aircraft can experience longitudinal and lateral instability associated with flow into the ducts and interference between ducts.

The present investigation was undertaken to study the longitudinal and lateral stability characteristics from hover through transition to normal for- ward flight. Included in the investigation were various types of modifications used in attempts to reduce some of the stability problems encountered, particu- larly in the low-transition-speed range. This paper will present only the lon- gitudinal characteristics and the effects of the appropriate modifications.

COEFFICIENTS AND SYMBOLS The force and moment coefficients used in this paper are based on the dynamic pressure in the tunnel free stream. The positive sense of forces, 1 for the complete model and in fig- moments, and angles is indicated in figure ure 2 for the isolated duct-propeller assembly. Moments of the complete model are referred to the assumed center-of-gravity location indicated in figure 3 .

Moments of the isolated duct-propeller assembly are referred to the individual propeller plane.

Measurements for this investigation were taken in the U.S. Customary System of Units. Equivalent values are indicated herein parenthetically in the Inter- national System (SI) in the interest of promoting use of this system in future Details concerning the use of SI, together with physical con- NASA reports.

stants and conversion factors, are given in reference 2 .

lift coefficient, FL -

CL C I S

drag coefficient, FD -

CD qs

pitching-moment coefficient, MY -

Cm SSD lift coefficient of duct, FL,D CL,D gs FD,D

drag coefficient of duct, -

CD , D ss MY,D

pitching-moment coefficient of duct, -

‘ m , D SSD

propeller normal-force coefficient, FN,P -

CN,P

propeller pitching-moment coefficient, My,P -

cm, P SsD

propeller power coefficient, - 2ZQP

CP pnb5

propeller thrust coefficient, - TT

CT C I S propeller thrust coefficient based on free-stream dynamic pressure, cT, 1

- TP

ss I

. 111.1

propeller t h r u s t coefficient based on r o t a t i o n a l speed of propeller, 'T, 2 TD pn2D4 l i f t force of complete model, l b f (N) FL drag force of complete model, l b f ( N ) F D moment about Y-axis of complete model, f t - l b f (m-N) MY l i f t force of duct alone, l b f (N) FL, D drag force of duct alone, Ibf (N) FD , D propeller normal force, l b f ( N ) FN,P propeller pitching moment, f t - l b f (m-N) MY,P moment of isolated duct with respect t o duct reference axis, MY,D f t - l b (m-N)

free-stream dynamic pressure, 5 pf, l b f / f t 2 (N/m2)

mass density of a i r i n f r e e stream, slugs/cu f t (kg/m3) P D propeller diameter, 16.8 in. (0.427 m) propeller shaft torque, f t - l b f (m-N) Q b wing span, i n . ( m ) C l o c a l wing chord, in. (m)

-

C S wing reference area, 9.0 sq f t (0.835 m2) spanwise distance from plane o f symmetry, in. (m) Y T t h r u s t , l b f ( N ) t o t a l propeller t h r u s t , thrust plus drag (propeller o f f ) a t a = Oo, TT l b f ( N ) t h r u s t of each propeller, l b f (N) TP n propeller r o t a t i o n a l speed, rps or r p m U model angle of a t t a c k with respect t o fuselage reference l i n e , deg isolated-duct angle of attack, deg model p i t c h angle with respect t o fuselage reference l i n e i n hovering, deg duct deflection angle, deg vane o r elevon deflection angle, deg horizontal-tail, o r s t a b i l i z e r , incidence angle, deg thickness of front duct fairing, in. (m) free-stream velocity, f t / s e c (m/sec) velocity of full-scale airplane, knots lift force of complete model a t zero =le of attack i n the hovering condition out of ground effects, 1bf (N) height of model above ground board, f t (m) Subscripts: F front R r e a r Model components: front duct f a i r i n g between duct and fuselage DF H t horizontal tail, or s t a b i l i z e r horizontal tail, or s t a b i l i z e r , designated a s number 1 ( f i g . 6) H t , l v e r t i c a l t a i l

vt

MODEL A N D APPARATUS A drawing of the basic model with pertinent dimensions i s given i n f i g - ure 3 . The model w a s mounted on a sting-supported six-component strain-gage balance for measurements of the complete model forces and moments. Thrust w a s

provided by four ducted propellers - two forward mounted close inboard and two

rearward mounted on a wing a t the r e a r of t h e model. Also mounted on the rear wing were four nacelles simulating engine nacelles of a full-scale airplane f o r powering t h e four propellers through shafting and gearing. However, f o r t h i s tunnel model each propeller was powered by a separate e l e c t r i c motor. The duct center bodies were, therefore, larger than they would be on an actual airplane.

The motors were mounted on strain-gage balances t o measure the forces and moments of the propellers.

The two right-hand motors w e r e instrumented t o measure normal force and pitching moment; t h e two left-hand motors were instrumented t o measure side Each p a i r of ducts was mounted on a shaft with actua- force and yawing moment.

t o r s attached so t h a t the f r o n t ducts could be rotated through an angle range independent of t h e r e a r ducts. The various horizontal stabilizers, control surfaces, and modifications used on t h e model a r e shown i n figures 4 t o 8. The wing incidence angle w a s held constant a t 3 O throughout the investigation.

TESTS AND CORRECTIONS The tests were conducted i n the l7-foot t e s t section of the Langley 3OO-MPH 7- by 10-foot tunnel. The arrangement and calibration of t h i s t e s t The tunnel velocity and the section are given i n the appendix of reference 3 .

model propeller speed were held constant throughout the angle-of-attack range Except f o r the few runs noted i n the figures, the vari- f o r a p a r t i c u l a r run.

ous t h r u s t coefficients were obtained by varying the tunnel velocity from one The t h r u s t coef- run t o t h e next while maintaining a constant propeller speed.

, f i c i e n t CT w a s based on the t o t a l thrust % of the propellers, which was obtained by taking the difference between the longitudinal force with the pro- p e l l e r operating and the longitudinal force with the propeller removed at zero angle of attack and zero duct deflection angle. The thrust coefficients of each propeller, CT,l and C T , ~ , were based on the measured thrust of each propeller a t each angle of attack.

Corrections t o t h e free-stream velocity t o account f o r blockage were e s t i - mated by the method of reference 4, were found t o be negligible, and thus were not applied. The jet boundary corrections were estimated by a method f o r rec- The tangular tunnels (ref. 5 ) and by a method for square wind tunnels.

resulting corrections were applied t o the cruise configuration data ( f i g s . 10 t o 17) a s follows: a = % + 0.322% where %> CD,m, and Cm,m a r e t h e measured values i n the tunnel.

Because of t h e uncertainty of applicability of model corrections t o a tan- dem configuration of t h i s type a t t h e high-lift coefficients encountered i n transition, no corrections w e r e applied t o the t r a n s i t i o n and duct-alone data (figs. 18 t o 27).

The Reynolds numbers for the range of thrust coefficients used based on wing chord or duct chord and free-streamvelocityare given in the following table:

Reynolds number based on -

I lb/ft2 N /m2

Wing chord Duct chord 478.8 9.8 x 105 4.8 x 105

1 .8 478.8 4.8

9.8 8.16 4.0 336 0 168.0 2.85 5=8 62.2 1 . 7 2 3.5 20.2 2 . 0 1.34 .66 9 . 1 PFU3SENTATION OF FU3SULTS The results of the investigation are presented in the following figures: Figure Hovering characteristics . . . . . . . . . . . . . . . . . . . . . . . .

Cruise configuration characteristics . . . . . . . . . . . . . . . . . .

10-17 Characteristics in transition:

Duct stall characteristics . . . . . . . . . . . . . . . . . . . . . . 1 8

19-20 Trim and stability in transition . . . . . . . . . . . . . . . . . . .

21-24 Control effectiveness . . . . . . . . . . . . . . . . . . . . . . . .

Isolated-duct characteristics: Large-vane effectiveness . . . . . . . . . . . . . . . . . . . . . . .

Small-vane effectiveness . . . . . . . . . . . . . . . . . . . . . . .

Large vane on and off . . . . . . . . . . . . . . . . . . . . . . . .

2 7 2 8 Summary of duct lip stall . . . . . . . . . . . . . . . . . . . . . . .

Summary of effects of thrust coefficients . . . . . . . . . . . . . . .

2 9 Summary of control effectiveness . . . . . . . . . . . . . . . . . . . .

Flow patterns indicated by tuft studies at various thrust coefficients, duct deflection angles, and angles of attack are presented as parts of fig- ure 19. Combinations of duct deflection angles and model angle of attack are given only where significant changes in stall characteristics occurred. For some tests, propeller data were recorded and are presented as parts of the appropriate figures. For the sumnary figures 2 9 and 30, the model was treated as a 0.20-scale model of a 15 000-pound (66 720-newton) airplane.

DISCUSSION Hovering Characteristics The hovering characteristics of the model within ground effect (h/D = 0.595) are given i n figure 9 , which indicates appreciable changes i n the center-of -pressure location, M y /(FL),~D, with changes i n angle of attack and a s l i g h t favorable effect of the ground on l i f t near zero and nose-up attitudes.

Cruise Configuration Characteristics The first configuration tested, indicated by the c i r c l e s i n figure 10, w a s found t o have very low longitudinal s t a b i l i t y with power off, CT = 0, and became slightly unstable with power on, CT = 0.8; consequently, a number of modifications were t r i e d t o a l l e v i a t e t h i s condition before the t r a n s i t i o n studies were made. Increasing the size of the outboard s t a b i l i z e r helped, but greater increases i n s t a b i l i t y were obtained by reducing t h e s i z e of the front duct f a i r i n g s or removing them entirely.

(See f i g s . 10 and 11.)

Since the smallest duct fairing (number 3 shown i n f i g . 5 ) gave a more stable configuration than the other fairings t r i e d , it was, therefore, used f o r the r e s t of the investigation.

The increased s t a b i l i t y provided by the smallest f a i r i n g can be attributed t o the reduction i n l i f t carry-over on the front ducts and fuselage caused by the f a i r i n g geometry. The e f f e c t s of s t a b i l i z e r incidence and the e f f e c t s of removing the s t a b i l i z e r a r e shown i n figure 12.

A large improvement i n l i n e a r i t y of the pitching-moment curves and i n sta- b i l i t y a t r e l a t i v e l y high t h r u s t coefficients was achieved by reducing the Although deflection of both the front and rear ducts 3° as shown i n figure 13.

duct deflection angles of Oo on the front and -5' on the r e a r give a good cruise configuration, it must be realized t h a t the a i r c r a f t w i l l have t o pass through the higher duct angles while going through transition.

From figure 14, it i s seen t h a t with power o f f , removing the propeller had Also, with suf- l i t t l e e f f e c t on the aerodynamic characteristics of the model.

f i c i e n t power t o provide a zero drag condition a t zero angle of attack, the longitudinal s t a b i l i t y of the model w a s v i r t u a l l y t h e sitme as with power off.

Data f o r t e s t s showing the e f f e c t s of various model components on the aerody- namic characteristics of the model a r e given i n figure 15.

In reference 6, it w a s shown that rotating t h e propeller against the t i p vortex resulted i n a slight reduction i n power required and pitching moment i n the t r a n s i t i o n speed range for an unshrouded propeller located a t the w i n g tip.

However, with the ducted f a n of t h i s investigation, direction of propeller rotation had no e f f e c t on l i f t (on which power required i s a d i r e c t f'unction) and pitching moment ( f i g . 16).

A few tests with the engine nacelles removed indicate that the nacelles produced a slightly destabilizing contribution to the pitching moment at a thrust coefficient of 0 . 8 and that in general the effects of the nacelles were very small (fig. 1 7 ) .

Characteristics in Transition Duct stall characteristics.- W t studies of the original configuration at a duct incidence of 13O and greater indicated separation of the flow from the duct lower lip at combinations of power and speed corresponding to steady level flight (Thrust = Drag). Similar separation problems were encountered in the

study reported in reference 7 . In that investigation, separation occurred much

earlier on the model than on the corresponding full-scale configuration. In

reference 7, it was found that nearly full-scale conditions could be simulated

by increasing the lower lip radius of the duct.

Consequently, a number of sim- ilar modifications were made in an effort to delay the lower duct lip separa- tion until higher angles of attack were reached. A sketch of the most effec- tive duct lip modification which was used throughout most of the investigation is shown in figure 4 . Data for the other duct lip modifications, therefore, are not presented in this paper.

The enlarged lower duct lip delayed stalling in the duct as shown by the power-off lift curves in figure 18(a) and power-on stall boundaries in fig- ure 28 (determined from tuft studies on the isolated duct). Moreover, with the enlarged lower duct lip, stalling in the duct was delayed to higher duct angles than stalling on the upper duct surface (fig. 2 8 ) . Also, delaying flow separa- tion on the lower duct lip helped to delay the stall on the upper duct surface above thrust coefficients of 2 . 0 .

Increasing the lower lip radius altered the camber of the lower duct pro- file and resulted in a change in pitching-moment characteristics with power off (fig. 1 8 ( a ) ) , as might be expected. However, with power on (fig. 1 8 ( b ) ) , the effect is reduced, because some of the moment obtained is due to the flow being turned into the duct. With power on, the pitching moment is less affected by duct profile.

Trim and stability in transition.- The transition characteristics are sum- marized in figure 29 where-thedata from figures 19 and 20 have been used to calculate the-characteristics of a 15 000-pound (66 720-newton) airplane by assuming the model to be l/5 scale. In the lower transition speed range, 20 to 50 knots, the data indicate that the airplane would be unstable and have large out-of-trim pitching moments. A similar analysis of the data from reference 8 indicated that the four-duct-propeller configuration of that investigation would have almost twice the out-of-trim moment compared with that of the assumed airplane of this investigation. The lower nose-up moments of the pres- ent configuration are probably due to the cambered wing and 3 O incidence com- pared with the symmetrical airfoil section and 0 ' wing incidence used on the configuration of reference 8 .

The maximum moments of the airplane of this investigation would require about f2000 pounds ( 2 8 8 9 0 newtons) of differential thrust from front to rear ducts t o trh. (See t h r u s t required and moments i n f i g . 30.) A combination of both di.fferentia1 thrust and vane control may be a b e t t e r means of providing trim f o r the airplane i n t h i s speed range.

Control effectiveness.- Figures 2 1 t o 24 show the control effectiveness of the v & e s a h the ducts a t various thrust coefficients, which correspond t o the range of very low t r a n s i t i o n speeds t o full forward flight speeds. The

pitching-moment curves show that vane deflections of -loo t o +loo provide suf-

f i c i e n t control t o trim the m o d e l i n t h e higher t r a n s i t i o n speed range, where values of CT ranged from 0.40 t o 0.80. However, i n t h e m e d i u m and very low CT were 2.1 and above and the higher duct t r a n s i t i o n range, where values of deflection angles were required, vane deflection w a s inadequate and thus dif- f e r e n t i a l thrust between t h e front and r e a r ducts i n addition t o vane control w i l l be required (figs. 2 1 t o 24).

A s previously mentioned, it was noted from t u f t studies t h a t various p a r t s of t h e ducts s t a l l e d at d i f f e r e n t duct deflection angles. It w a s also observed t h a t the upper vane surface behind the upward rotation side of the propeller s t a l l e d a t about the same duct angles of attack as the duct upper surface, whereas the section behind the downward side was delayed 1 5 ' t o 2 0 ' higher due t o t w i s t i n t h e propeller slipstream. Early s t a l l i n g on p a r t of the duct vanes, of course, would be detrimental t o t h e control effectiveness of these surfaces.

A wind-tunnel investigation of a powered four-duct-propeller model of a VTOL a i r c r a f t indicated the following conclusions: 1. The first configuration t e s t e d w a s longitudinally unstable a t high thrust coefficients i n the cruise condition with the center of gravity midway between the centers of rotation of the ducts. Longitudinal s t a b i l i t y w a s achieved by diminishing the destabilizing e f f e c t s of the front p a i r of ducts.

This w a s accomplished by reducing the size of the f a i r i n g s between the ducts and the fuselage, and thus, the l i f t carry-over t o the fuselage was reduced.

Reduction of duct incidence i n the cruise configuration also greatly imgroved the longitudinal s t a b i l i t y .

2. Duct. l i p stall, which was aggravated by the low Reynolds nwbers of the tests, w a s delayed t o higher angles of attack by increasing the duct l i p radius.

Delaying the s t a l l on t h e lower duct l i p also helped t o delay stalling on the upper duct surface.

3 . The l a r g e s t nose-up moments i n t r a n s i t i o n were encountered a t a duct deflection angle of about 60° (approxima,tely 40 knots on a full-scale airplane).

A t t h i s condition, differential t h r u s t between the front and rear p a i r s of D i f f e r e n t i a l duct deflec- ducts will be required t o provide t r i m and control.

t i o n and deflection of control surfaces within the ducts have very l i t t l e Ill I Ill1 I1 I jl I1 l 1 l l 1 l 1 I 1 I 1 1 1 I IIIII I 1 I 1 I Ill I1 l111ll effectiveness at high duct deflection angles and low speeds, but provide power- f'ul control at low duct deflection angles and high speeds.

Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., October 7 , 1965.

1. Newsom, William A . , Jr. : Aerodynamic Characteristics of Four-Duct Tandem NASA TN D-1481, 1963.

VTOL-Aircraft Configurations.

2 . Mechtly, E . A . : The International System o f Units - Physical Constants and

Conversion Factors. NASA SP-7012, 1 9 6 4 .

3 . Kuhn, Richard E . ; and Hayes, William C., Jr.: Wind-Tunnel Investigation of Longitudinal Aerodynamic Characteristics of Three Propeller-Driven VTOL Configurations in the Transition Speed Range, Including Effects of Ground Proximity. NASA TN D-55, 1 9 6 0 .

4 . Herriot, John G.: Blockage Corrections for Three-Dimensional-FlowClosed-

Throat Wind Tunnels, With Consideration of the Effect of Compressibility.

NACA Rept . 995, 1950. ( Supersedes NACA RM ~7~28. )

5. Gillis, Clarence L . ; Polhamus, Edward C.; and Gray, Joseph L . , Jr.: Charts for Determining Jet-Boundary Corrections for Complete Models in 10-Foot Closed Rectangular Wind Tunnels. NACA W R L-123, 1945. [izerly NACA ARR L5G31.)

Investigation of a Semispan Tilting-Propeller Con- 6 . Spreemann, Kenneth I ? . : figuration and afects of Ratio of Wing Chord to Propeller Diameter on Several Small-Chord Tilting-Wing Configurations at Transition Speeds.

NASA TN D-1815, 1963.

7 . Goodson, Kenneth W . ; and Grunwald, Kalman J . : Aerodynamic Characteristics

of a Powered Semispan Tilting-Shrouded-Propeller VTOL Model in Hovering NASA TN D-981, 1962. and Transition Flight.

Transition Characteristics 8 . Davenport, =win E . ; and Spreemam, Kenneth P . : of a Aircraft Powered by Four Ducted Tandem Propellers. NASA Tm D-2254, 1 9 6 4 .

h Ground board la) Hovering tests.

Figure 1.- Conventions used t o define positive sense of forces, moments, and angles for the complete model.

(b) Tunnel tests.

Figure 1.- Concluded.

Figure 2 - Conventions used to define positive sense of forces, moments, and angles for the isolated duct with propeller.

P w

/-Tc-- Outboard siabifizer no2

(Modi fled)

600 - \/ Propefler 3

- 5 4 0 (f37/

( f 52)

- - 980( 249)

> (2.3401 I e- 2870 - -2870- (7291 .-. I740 - >- I 7 4 0 ( 729) - - - (442) (4421 I --/460-- (3711 Figure 3.- Drawing of complete model. Dimensions are given first i n inches and parenthetically i n meters.

18.60 42 (I063 1g73' 105081 ~, zero Dimensions are given first in inches and parenthetically in meters.

Figure 4.- Drawing of isolated duct with duct lip modifications.

r UI

I

fD

t

F a i r i n g no. 2 ~~ J.

2.50 L063.5..

F a i r i n g s 2 a n d 3

\ I Fuse /age

\ \_1

- ; ~ 1 I

I fD

- -1

Vane _L__ . - - - - . -----

-1 Duct -!I

F a i r i n g no./

t -

4.00 f. /0/6) $ - / 3 . 2 0 ~ - 4 (329) F a i r i n g no. / Figure 5.- Details of fairing between duct and fuselage. Dimensions are given f i r s t i n inches and parenthetically i n meters.

-- 4.60 -c-

No. 3 ( N A C A 0015 -64 section} Duct , , , 9 / / t-

- -4 5.05

Duct p i v o t - f.1283) No. 2 f N A C A 64 A 415 section} D u c t N o . / ( N A C A 64 A 415 section) Figure 6 - Details of rear-mounted outboard stabilizers. Dimensions are given f i r s t in inches and parenthetically in meters.

fuse. sta.

88.20 (224)

j W i n g t r o i / i n g edge

1 1 0 6 3 5 )

T

b 2 1 6 ) S e c t i o n A - A Figure 7.- Drawing of tail-fuselage fairing. Dimensions are given f i r s t i n inches and parenthetically i n meters.

Vane p i v o t \-- __ F r o n t l e f t v o n e

--r\-- \ 5.4 7

1 . 1 3 9 ) I

1-

/L54 (293) I

.I

L 4.20

0066) < 18.66 L474) Rear / e f t vane Figure 8.- Details of vanes (elevons) used i n ducts. Dimensions are given f i r s t in inches and parenthetically i n meters.

-5 5

- /O /O I5

Figure 9.- Aerodynamic characteristics i n hovering ~ D , J = 900; ~ D , R = 900; 6 v , ~ = Oo,

-2 I

I

I

I I

-4 OF H 7 I 0 / -. 6 I 0 2 Off 2.5 2.C 1.: 1c - 1 1 5 30 la) CT = 0; propellers windmilling.

Figure 10.- Effects of outboard stabilizer and f r o n t duct f a i r i n g o n t h e aerodynamic characteristics of t h e model at h40 t h r u s t COeffiCients. 6D,F = 5O; 6D,R = @; 6 v , ~ = 0 ' ; ~ V , R = 0 ' ; it = 0 ' ; Vt Off.

.21

/

Off

0 .2

1 . 0 (a) Concluded.

Figure 10.- Continued.

.5 I Cm -. 5 I ; ! ,

- /.I

Hr

I 5.0 0 I G 3.0 2.0 /

f.'

-LG

- 2.C

20 25 30 / 5 - 5 0 5

- /O

(b) CT = 0 . 8 .

Figure 10.- Continued.

HT

I

0 1

I. 0 0 2

I i i i i i U I i i i i i i i i

0 1 Off

- 1D 0.5 0 .5

(b) Concluded.

Figure 10.- Concluded.

Tail- fus.

fairing b/D 4 Y Off ,067 O f f o r On O f f 0 2 ,067 O f f On ,067 0 3 On On A 3 .067 O n On b 3 ,097 .2

i ' l

e , -2 - 4 -.6 3.6 2 5 2.0 L 5 c, LO

I

I - .5 I

I I

- 1.0 -10 - 5 5 IO 15 20 25 30 (a1 CT = 0; propellers windmilling.

Figure 11.- Effects of f r o n t duct fairings, distance of f r o n t duct from fuselage, and tail-fuselage f a i r i n g o n the aerodynamic characteristics of the model at various t h r u s t coefficients. ~ D , F = 50; ~ D , R = 00; %,F = W; 6V,R = 00; Ht,l on; it = 00.

Tai/- fus.

DF b / D f a i r i n g o / .067 O f f

2 .067

O f f On .067 Off Off A .067 On On h .097 On I I

i

t

1 1 I

f

‘ !

I I

+

I I

i r

i

-I

i

I

I

ffl

H t

0 .2 .6 .8 I. 0 CD (a) Concluded.

Figure 1 1 . - Continued.

.26 Tail-fus.

f a i r i n g v, t,/D O f f .067 Off On Off .067 On Off On .067 On On .097 On .5 Cm

-i

-.5 -/.G 5 C 4.c 3.6 2 C , I t I 1 ‘ I

;, 1

CL I. c -/,C -2.C 0 -5 5 l0 1 5 20 25 (b) CT = 0.8.

Figure 11.- Continued.

vr

o / .067

O f f

O f f

0 2 .067

O f f On

0 3 .067

Off

On

A 3 .067 On

On

0 3 .097 On

On

- LO -. 5 0

,5 LO

co

(b) Concluded.

Figure 11.- Continued.

..28 ?ai/- fus, t , / ~ fairing .067 Off Off .067 Off On Off .067 On .067 On On 5.G 4.0 3.0 2.0 1.0 - /.o - 5 5 /O /5 20 25 30 Figure 11.- Continued.

Tui/- fus.

fairing

4 t,/D

v, 0 /

.067 Off

Off .067 O f f On 3 .067 Off On A 3 .067 On On !

f

-2.5 -

2.0 -1.5 -LO -.5

co

(c) Concluded.

Figure 11.- Concluded.

I . 4 Cm -. 2 -4 - . 6 o / 0 1 0 O f f 3 . 0 2.5 l.5 CL 1 . 6 I , , . , , , , I , , , I / < , I " & .d C -5

i l i

- LL -0 -5 0 5 /5 30 20 25 (a) CT = 0; propellers windmilling.

Figure 12- Effects of outboard stabilizer and stabilizer incidence on the aerodynamic characteristics of the model at two thrust coefficients. ~ D , F = 50; ~ D , R = go; ~ V , F = 00; 6V,R = 00; Vt on; tail-fuselage fairing on.

-- - . . . - . . .

3.0

2.5

2.0

/. 5

Hf

1 . 0

/

/ -10

-

Off

.5

0 1

-. 5

-10

0 .2 4 .8

m CD (a) Concluded.

Figure 12- Continued.

.5 Cm -. 5

- /.o

Ht 0 / 0 0 / 40 - 0 Off

4. c

3.c

b

2 .c

I

- /.I

-2.

/ 5

/O

- 5 0

- /o

Figure 12- Continued.

0 / 0

a / - /O

0 O f f

i I I I I I T

I

I I T I

I

I T I I I I T I

I

I T I I

I

I

I I

t

I

I

I r

I

I I r I I

I

r

I

I

-LO -.5 0

co

(b) Concluded.

Figure 12- Concluded.

B i

i

5.0 0 0 5./7 -525 0 0 -5.25 - 1 0 -5 /O 25 30 (a) CT = 0; propellers windmilling.

Figure 13.- Effects of duct deflection o n t h e aerodynamic characteristics of t h e model at various t h r u s t coefficients. 6 v , ~ = Oo; ~ V , R = Oo; Ht,l on; it = 00; Vt on.

. ..

3.0

2 . 0

/. 5

- 525

-525

0 .2 .8 L O /.2

(a) Concluded.

Figure 13.- Continued.

Y H

;r

I

I

- .5

-LO -525 5/7 5.0 -5.25 4. L 3.L

C

- 1 . 1

- 2.0

- 5 0

- /o

Figure 13.- Continued.

5.0 0

5. /7

-525

0 -5.25

40 -. 5 0 .5

(b) Concluded.

Figure 13.- Continued.

-525 I 6.1 5l 4.L 3.1 c , 2.L /.1 -/.I - 211 - /O - 5 0 5 15 20 25 30 Figure 13,- Continued.

5.0

4.0

3.0

2.0

I. 0

-525

i

It.

-10

Ij

- 2.0 ULlll

! I I

-25 -2.0 - L 5 40

-. 5

(c) Concluded.

Figure 13.- Concluded.

.2 cm

- .4

! 0

- . 6

?pef fer O f f

- .8

idmiffing 0 5000 rpm 2 . 5 2.G f.5 c, f.c .5

i l l !

- LL

-fO -5 0 5 f0 20 25 30 Q, de9 Figure 14.- Effect of propellers and sufficient t h r u s t for zero drag o n t h e aerodynamic characteristics of the model. ~ D , F = 5O; ~ D , R = Oo; h , ~ = OO; ~ V , R = 00; Ht,l on; it = Oo; V t on.

2.5

2.0

0 . 2 4 .6 .8

/. 2

CD Figure 14.- Concluded.

L 5 .5 -.5 Prop e / / e Windmilling 3. G Windmillin4 O f f 2.5 2 . 6 /. 5 -.5 25 30 / 5 20 - /O - 5 0 5 Figure 15.- Aerodynamic characteristics of t h e major components of t h e model. Power off (includes m u t u a l interference effects); BV,F = 00; B V , R = Oo; Ht,l on; it = Oo.

0 5

0 On Windmilling

0 5

0 f f On Windml'llhq

0 O f f O f f O f f O f f

0 .2

. 4

. 6 .8

LO I. 2 c, Figure 15.- Concluded.

I

.5

i i

cm

-.5

1 r r

- /.c

Propeller ro totion //nboord/ 0 Upward 0 Do wnword 4.C 3.c

2 . c

CL /.a

- /.c

-26 - 1 0 -5 0 5 /O 1 5 20 25 30 Figure 16.- Effects of direction of propeller rotation o n t h e aerodynamic characteristics of t h e model.

CT = 0 . 8 ; 6D.F = 0 ' ; 6D,R = -5.25O; bV,F = 0 ' ; 6V.R = 0 ' ; Ht.1 on; it = Oo; Vt on.

Propeller ro f a f ion

flnboardj

Upward

Downward

4 . 0

3.0

2.0

C L LO

- 1 . 0

-2.0

40 -. 5

CD

Figure 16.- Concluded.

.. . . . . .

.2 c, 52 - . 4 , -.6 3.c 2 . C /1: c, / 5 20 25 5 / O - 5 0 (a) CT = 0; propellers windmilling.

Figure 17.- Effects of engine nacelles on t h e aerodynamic characteristics of t h e model at two t h r u s t coefficients.

6D,F = 5O; 6 D , R z o o ; 6 v , ~ = 0 ' ; ~ V , R z o o ; Ht.1 On; it = 0 ' ; V t O n .

1-1 1 Engine

II n ace//es

I I

1 0 O n

7 I I 0 O f f

0 -2

. 4

.6 .8

CD (a) Concluded.

Figure 17.- Continued.

I

I

Engine n acelfes 0 On

I I

, / 2.0

- LC

-2.c 25 30 IO l5 - 5

- IO

(b) CT = 0 . 8 .

Figure 17.- Continued.

i i

T I 1 1

I I I I R 1 n I I T r r i I I 7 I r i t 1: I I

! I I

[I i r

T 1 ' I

I

i T1

I T

f T I

1 I I I I 1 I I 1 . 1 . 1

Engin

nace//c

On

Off

i I

I

f

r

T

f

I

t I

r

I L

- LO 0.5

.5

(b) Concluded.

Figure 17.- Concluded.

-.2 cm - . 4

i " I I1

-.6 Duct lips 0 Enlarged 0 Original I 2.5 2.G 1.5 .5 - .5 -5 0

- lo

(a) CT = 0; propellers windmilling.

Figure 18.- Effects of change in size of h e r duct lip o n t h e aerodynamic characteristics of t h e model at two t h r u s t coefficients.

6D,F = 15O; 6D,R = 15O; 6 v , ~ = 0 ' ; ~ V , R Zoo; Ht.1 On; it = 0 ' ; Vt O n .

2.0

1 1 5

Duct lips

o Enlarged

I. 0

t l O r i g i n o l

.5

-.5

1 1 4

.8 L O

.2 . 6

(a) Concluded.

Figure 18- Continued.

.5

c, -lG D u c t //;os 2.0 l0

-10 -5 0 5 25 30

Figure 18- Continued.

D u c t /ips

0 E n l a r g e d

O r i g i n a l

i I I I I I X

4.0

I k

I I

I I I I I I

I

I I I

I

I I I I

LO

I

t

I I I

- 1

0 .5 L O

-10

(b) Concluded.

Figure 18.- Concluded.

b, F , deq - /O 0 /O /5 20 (a) CT = 0.4.

Figure 19.- Effects of duct deflection angle on t h e aerodynamic characteristics of t h e model at various t h r u s t coefficients.

6v F = Oo; 6~ R = 00; H t 1 on; it = Oo; V t on; modified duct lips; tail-fuselage f a i r i n g on.

4.0

il

‘ I

3.5

l i

3.c

I / I /

2 5 I I CL 2.c 30 $

i i

1.5 L O

‘I

.5

i

0 ’ - 4 -2 0 .2 4 .6 .8 /. 0 /.2 1.4 (a) Continued.

Figure 19.- Continued.

I J i (a) Concluded.

Figure 19.- Continued.

I

i I

i L

I

i

-

85, F deg

-

/ 5

/5

,

i

I

i

I

' I I I I

I

I

I

I

I

I

i

I

5 !

J

I

-10 - 5

/O

Figure 19.- Continued.

i f

I

t

I

T

I I I

r

I

45 #

- 1c

-.5 0

.5 LO /. 5 2.0

C D

(b) Continued.

Figure 19.- Continued.

I (b) Concluded.

Figure 19.- Continued.

1 1 1

10 I

Cm

- /.o

0 45 8 . 0 I A 60 60 IO I / _- 6 C 5 . 0

4. c

C 2.c

I

0 5 io / 5 20

- /o - 5

(d CT = 2.1.

Figure 19.- Continued.

3.0

2.0

1.0

4 5 - /.O - .5 0 /.5

.5 1.0

(c) Continued.

Figure 19.- Continued.

\ (c) Concluded.

Figure 19.- Continued.

i

f

A t l .

!

- 1 0

-5 (d) CT = 7 . 0 .

Figure 19.- Continued.

/4.q

i T

I

i I

I

I

I

I L

I I I I I

I

I

I

K i

q

rr

/3.a I

I

f

I I

I I I

I I I I

I

I

I

I

I

I I

+

m 0 1 I

I

I

I

i

i I

I

I

I

I

I

I I 1

I

I

i

I I 1 I

8 . 0 1

I

I

I I I

CL

I

s , I I

6 0 1

I

L

I

I I

i l l

t I

4.0 I I

I I I

I

i i l

I I I I I I I

!

I l l

f-

2.0; I I

I

I

I t I 1

'

I I I I I 1- 1

0 '

4.0

-40 - 2.0 0 2.0 6.0

(d) Continued.

Figure 19.- Continued.

(d) Concluded.

Figure 19.- Continued.

-~ ._.. ,._,. . ...... . ... ._.... .,.... .. ... ....,. .. . . . ~ ,.

/O

ii

- /O -5

0 5 l0 /5 20 25 (e) CT = 25.

Figure 19.- Continued.

L L L IO / 5 20

-15 - /O - 5 0 5

(e) Concluded.

Figure 19.- Continued.

/ 5

p

Cm /O

I

I

sD, F P deg A 75 G 6L 5c 4c CL 3c k

c

-5 0 5 /O /5 20 25

- 10

Figure 19.- Continued.

i

I

I I I I

I

60 I

I

I

I

I I

I

f

I

I

I I

I

I

I I I

/O

-20 - 1 0

/O

CD

( f ) Concluded.

Figure 19.- Concluded.

I

I

0 5 0 .2 I O 5 0 I 5 I O L A / 5 I 5 c, -.2

- . 4

-. 6 -. 6 2.c c - 5 0 /5 20 25 30 (a) CT = 0; propellers windmilling.

Figure 20.- Effects of differential duct deflection angle o n t h e aerodynamic characteristics of t h e model at two t h r u s t coefficients.

6 v , ~ = g o ; 6V,R = Oo; Ht,l on; it = 0 ' ; Vt on; modified duct lips; tail-fuselage f a i r i n g on.

/O

/5

1 0

/5

/5

i

i

c I

;i

I 1

I I

I I

I

I I

I

I

I

I

0 .2

.4

.6

.8

(a) Concluded.

Figure 20.- Continued.

l0 / 5 A /5 .5 Cm -5 -1.0 5 . 0 4 C 2 c C

- 2 . 1

25 30 I5 20 5 IO (b) CT = 0 . 8 .

Figure 20.- Continued.

0 5

0 l0

0 /5

A /5

. . . . . . .

I 1

40 1

r i

T

1 I

t I

I t

r

T

I

I

I

20 c T

I I

CL

r

f

c

1.0

I

t

I

t

c

I

i

- L O

t

t

I

f

-20

40 -. 5 0 .5 1.5

(b) Concluded.

Figure 20.- Concluded.

~ D . F ,de9 ,de9 ,deg 0 5 0 0 0 5 -10 IO 0 5 IO -10 A 0 0 - IO n 5 0 - IO .6 .2 Cm 0 -. 2 - 4 -. 6 2 . 2 ZL L5 I. L CL .5 -. 5 -LL 4: 3 - 5 0 5 IO I5 20 2 5 30 a .deg (a) CT = 0; propellers windmilling.

Figure 21.- Effects of vane deflection and duct deflection o n t h e aerodynamic characteristics of t h e model a t various t h r u s t coefficients.

~ D , R = Oo; Ht,l on; it = 00; V t on; modified duct lips; tail-fuselage f a i r i n g on.

7;

- 10

/O

-/O

- /O

- 10

0 .2 . 6

-8 l.0 CD ( a ) Concluded.

Figure 21.- Continued.

1.0 .5

;t

cm

1 j

-. 5 -LG sv, F -/.5 0 0

- lo l0

10 -10 4.6 0 -/O 0 -10

I /

3.C

- LC

r l l

i

- 2.6

- /O -5 0 5 20 25 30

Figure 21.- Continued.

.. . ..... . .. . . . .

& * R I ~ e g

- /O 10

/O - /O

A

0 - /O

b

0 - /O

. .

I

! I

. .o

-. 5 0

(b) Continued.

Figure 21.- Continued.

.... . . . . ... .- . . - . . - . . - ._ S Y , R , deg 0 5 0 0 0 5 - lo l0 0 5 l0 - lo L O 0 - IO Propeller / P r o p e l l e r 3 . 5

cr, f

.2G .IC CP .05 C k P O -. 05 .02 %P O 1 0 2 -IO IO 20 30 -10 0 (b) Concluded.

Figure 2L- Continued.

- /o /O 0 -10 0 -10 -10 - 5 0 20 30 ( C ) CT = 2.1.

Figure 21.- Continued.

I

IO -10 -10 -10

5.0

4.0

3.0

CL 2.0 1.0

- 1.0

-2.0

- 1.5

-20

-2.5

(c) Continued.

Figure 21.- Continued.

/O - /O - / 0

- 10

Propeller 3 P r o p e l l e r .5 cr, f .20 cp .lo .05 c/v# - 0 : 05 .02 cm,P O 7 02 0 20 30 - /O - 1 0 0 20 (c) Concluded.

Figure 21.- Concluded.

-10 -5 0 5 IO I5 20 25 a,deg (a) CT = 0.4.

Figure 22.- Effects of vane deflection and differential duct deflection angle o n t h e aerodynamic characteristics of t h e model.

Ht,l on; it = Oo; Vt on; modified duct lips; tail-fuselage f a i r i n g on.

4.0 3.5 2.0 / 5 /. 5 / 5 / 5 .5 LO /. 2 .2 4 .6 .8 -.4 -. 2 0 CD (a) Continued.

Figure 22- Continued.

.3c .PL c< I .I C , 2 L . I C .IO . 0 5 1 0 : .oi -.Oi ? 0 IO 20 30 -IO 0 IO PO 30 ff, de9 0 , deg (a) Concluded.

Figure 22- Continued.

.. . _. . . . - - ._ 30 0 0 25 35 0 0

30 - 10 1 0

A 30 30 -20 .5 Cm

-. 5

- L O

5.0 4.0 3.0 CL LO L

-/O -u

3 5 /O 20 25 Figure 22- Continued.

a 6

0 0

30 30

0 35 0

1 0

0 30 30 - 10

30 -20 20

5.0 I

I

3.0

2.0

1.0

.5 /. 0

-. 5

(b) Concluded.

Figure 22.- Continued.

Propeller speed, rpm J v , ~ , d e g Fronf Rear 7000 7000 45 45 0 0 7000 7000 40 50 0 7000 7000 0 45 45 40 A 7000 7000 45 -20 6500 7500 b ~ 45 0 LO -LO ZO 6.0 5.0 4.0 3.0 2.0 d 0 /O 5 /5 25 - /O -5 0 (c) CT = 2.1.

Figure 22- Continued.

0 45

45 0 0 7000 7000

1 3 40

50 0 0

7000 7000

0 45 -10 /O 7000 7000

A 45 45

-20 20 7000 7000

n 45 45

0 0 6500

-LO -. 5 0 .5

/.O l.5 2.0

2.5 (c) Concluded.

Figure 22- Continued.

0 60 60 0 0 0 55 65 0 0 0 60 60 -10 10 A 60 60 -20 20 Cm 2.0 /4.0 12.0 1 0 . 0 6.0 4.0 5 /O / 5 20 25 - /O - 5 0 Q I deq (d) CT = 7.0.

Figure 22- Continued.

0 60

60 0

55 65

0 60

60 - 1 0

/O

A 60 60

-20

/4. c

i

l0.C

8.C

CL

2.0

- 4.0 - 2.0

2.0 4.0 6.0

(d) Concluded.

Figure 22- Concluded.

Propeller speed, rPm Front Rear 7000 7000 0 0 0 7000 7000 0 -lo IO 20 7000 7000 A -20 6500 7500 h 0 0 I5 /O Cm C L /5 IO I5 20 - 5 0 5 IO -IO Figure 23.- Effects of vane deflection and differential t h r u s t o n t h e aerodynamic characteristics of the model.

CT = 25; 68 F = 750; 6~ R =750; Ht 1 on; it = 00; Vt on; modified duct lips; tail-fuselage f a i r i n g on.

?repeller speed, rPm f r o n t Rear

7000 7000

7000 7000

A

-20

n

6500 7500

i i

I

E I

cL

I

/ 5

I

I

- L - 5 0 5 / 5 Figure 23.- Concluded.

Propeller speed, rPm Front Rear 6500 7500 D C , 30 /O 0 - / O 15 20 25 -10 -5 0 5 Figure 24.- Effects of differential t h r u s t on the aerodynamic characteristics of t h e model. CT = 60; 6D,F = 90°; ~ D , R = 900; Ht,l on; it = 00; Vt on; modified duct lips; tail-fuselage f a i r i n g on.

Propeller speed,

rpm

F r o n t Rear

h 6500 7500

n

7000 7000

0 20 30 40

Figure 24.- Concluded.

0 0 IO 0 I5 A 20 b 25 n 30 b -10 0 -20 O -25 n ~ 70 80 90 IO 20 40 5 0 60 (a) CT = 0; propellers windmilling.

Figure 25.- Effects of vane deflection on t h e aerodynamic characteristics of t h e isolated ducted propeller w i t h large vane at various t h r u s t coefficients and w i t h modified duct lips.

.3 .2 cm, D ./ l l -10 3 -20 o -25 .6 0 -30 .4 cL,D .2 - .2 .4 .3 .2 CD, D .I -. / 40 50 70 80 90

- I O 0 IO 20 30

Figure 25.- Continued.

I I . ... ....-... .... .._. ... -.. 1 - . . . - 1 1 . . . I

I S I / , dep 0 l0 0 1 5

A a

/ / n 25 i !

n -10 / I -20 i 0 - 25 n - 30 20 30 40 50 60 70 80 80, deg (b) Concluded.

Figure 25.- Continued.

r . L C 0 f 5 A 20 b 25 -./ 0 30 0 -/o 0 -20 .8 0 -25 -30 . 6 .4 CL, D .2 0 .5 -2 4 .3 .i

-_ /

-. 2 - 1 0 0 20 30 40 50 60 70 80 90 Figure 25.- Continued.

0 I5 A b 25 n n -IO 0 -20 c -25 n -30 50 60 70 80 20 30 40 - /O 0 /O (c) Concluded.

Figure 25.- Continued.

0 1 0 20 30 50 70 80 90 40 60 Figure 25.- Continued.

n 25 n 30 0 -IO 0 -20 0 -25 -30 40 50 IO 20 30 deg (d) Concluded.

Figure 25.- Continued.

. 8

I i

i

G ? , D -6 . 4 .2 3.0 O l 2.5 2.0 CL,D 1.5 /. 5 I. 0 .5 L O a , , deg 0 0

-

0 IO 0 I5 A 20 b 25 -5 b -IO 0 -20 V -25 -LO 0 -30 -/.5 - 2.0 90 IO0 1 1 0 /O 20 30 40 50 60 70 80 =D, deg (e) CT = 7.0.

Figure 25.- Continued.

, LO .2 0 .IO .20 m .IO 0 IO 1 5 -10

- 20

-25 .IO -30 .05 -.05 90 I O 0 110 70 80 30 40 50 f0 20 Q DI deg (e) Concluded.

Figure 25.- Continued.

I

2.0 I. 5 Cm,o LO .5 a , , de9 0 0 0 IO 0 I5 A b 0 30 n -IO -30 1 CD> D -2 - 4

i

-6 50 60 80 1 0 0 30 40 Figure 25.- Continued.

6.0 4 . 0 Cr,/ 2.0 - 20 -25 2 0 -30 .lo CP .IO em. P -.lo -.20 90 loo 120 60 70 80 30 40 50 aD,deg ( f ) Concluded.

Figure 25.- Continued.

1 8 /6 0 1 4 0 /5 A /2 n a -10 /O -20 -25 n -30 - 1 1 1 , 40 50 60 70 80 90 100 /IO Figure 25.- Continued.

/5

I 1 ‘ I l l ‘ # I

IO Spdep 0 0 Cr,l 0 l0 0 15 A n 25 0 30 -lo 0 -20 0 -25 1 -30 . /o CP 1.0 ‘N,p .5 Cm,p -20

- 40

ll0 120 90 loo 70 80 50 60 30 40 Figure 25.- Concluded.

.- . . . . . . . .. .. .. - . _.

./a . 0,: cm,D

I

-.05 .3 .i ./ CL, D

I

I

-.I -.2 .2G ./ 5 CD, D ./C .05

! 7

3 0 20 60 70 ’0 9( (a) CT = 0; propellers windmilling.

Figure 26.- Effects of vane deflection o n t h e aerodynamic characteristics of t h e isolated ducted propeller w i t h small vane at various t h r u s t coefficients and w i t h modified duct lips.

./ em, D -./ 0 0 /o D -lo .B .6 CL,D .2 .3 .2 co, D ./ -./ -. 2 70 80 20 30 40 50 60 - /O 0 (b) CT = 0 . 4 .

Figure 26.- Continued.

0 0 0 /o n -/o ~ I I

i

I I ~ !

j

i

I i - /O 0 20 30 40 50 60 80 ( c ) CT = 0.8.

Figure 26.- Continued.

/.2 .8 A . 4 .2 cD, D 40 50 60 70 BO 90 20 30 (d) Cy = 2.1.

Figure 26.- Continued.

I n 20 30 40 50 60 70 80 90 100 //o deg (e) CT = 7.0.

Figure 26.- Continued.

1. 5 .5 CD, D -2 -4 -6 50 60 70 80 90 30 40 QD I deg ( f ) CT = 25.

Figure 26.- Continued.

i cm, D I L 20 /8 /6 /4 0 0 0 /o 0 -/o /2 /5 CL,D /O 8 5 6 CD, D -5 -10 0 - /5 30 40 50 60 70 80 90 (g) CT = 60.

Figure 26.- Concluded.

. /o

I l l

.05

ii

Cm, D

i i

-.05 'one Sv,deq o Large 0 ff .3 .P -. f .30 . PO .IO .05 40 50 60 70 80 90 - /O 0 lo 20 30 00,deg (a) CT = 0; propellers windmilling.

Figure 27.- Effects of vane o n and off o n t h e aerodynamic characteristics of t h e isolated ducted propeller at various t h r u s t coefficients and w i t h modified duct lips.

-.

vane o L a r g e I 3 O f f - /O 20 30 40 50 60 OD, deg (b) CT = 0.4.

Figure 2 7 . - Continued.

.2 ./ cm,D vane &,de9 -. / o ~ a r g e 0 O f f .8 .6 CL,D .4 .2 . 4 .3 .2 cD,D .I -. / -. 2 60 70 80 -10 . 20 30 40 50 ( c ) CT = 0.8.

Figure 27.- Continued.

I

. 3 .i cm, D ./ /.2 c L O Vonc Large 0 ff . 6 .2 CD,D 0 -. 2 -.4 -.6 20 30 40 50 60 70 80 90 Figure 27.- Continued.

. 8

I 1 I

. 6

I I i

.2

i i I

3.0 0

! I

vane Large Off l.5 2.0 D L O /.5 .5 /.o 0 .5 -. 5 -LO -L5 - 2.0 40 50 60 70 80 90 /o 20 30 Figure 27.- Continued.

c

2.0 1.5 Cm,D LO .5 Vane Large 0 CL, D - Off

' 1 i

I C co, D -2 , - 4

I

-6 50 60 70 80 I IO 30 40 Figure 27.- Continued.

C / l8 I 6 Vone o Large 0 I2 - Off lo cL,D /O D -5 - 10 70 80 90 50 60 30 40 (g) CT = 60.

Figure 27.- Concluded.

I

S f d i e d

T = l

Uns f o / / e d I 1 ./

.2 4 . 6 .8 / 2 4 6 810 20 30

G

Figure 28.- Summary of isolated-ducted-propeller stall t h r o u g h t h e t h r u s t coefficient range.

, Steady l e v e l f l i g h t _ _ _ _ _ _ Upper- f ron t-duct s t a l l -- U p per- rear- due t s t a I / 54,200 40,000 37, IO0 20, 000 Mr I MY, m-N f t-lb 0 0 -20,000 /355 IO00 f t - l b O deg - /355

- IO00

IO0 S D I deg I20 I40 20 40 60 80 100 V, knots Figure 29.- Pitching moment and duct deflection angles required t h r o u g h the transition speed range for an assumed 15 000-pound (66 720-newton) airplane. a = Oo; T h r u s t = Drag; ~ D , F = ~ D , R .

I' 66,720 /s,ooo

t

/o,ooo 4 4,480 r, 7; N / b 5,000 22,240

t

54,200 40,000 i 20,000 2 7,100

I

M Y , MY* -20,000 -27, /00 m -N f fJb - 54,200

- 40,000

- 8/, 300 -60,000 - /08.400

- 80,000

0 20 40 60 80 /oo /20 /40

Figure 30.- Pitching moment and thrust required through the transition speed range for an assumed 15 OM)-pound (66 720-newton) airplane. a = 00; Thrust = Drag.

NASA-Langley, 1966 L-4599 “The aeronairtical and space dctivities of the United States shall be conducted so as t o contribzite . . , to the expansion of hiimati knowl- edge of phenomena in the atmosphere and space. The Administration shall provide f o r the widest practicable and appropride dissemination of information concerning its actillities and the resitlts thereof.” -NATIONAL AERONAUTICS AND SPACE ACT OF 1958

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