Document
N A S A TECHNICAL NOTE
NASA TN D-3725
C J
-
m s
-
-
P
e -
- -
LOAN COPY: R ,-,=, LOAN COPY: R ,-,=,
AFWL [W,E= AFWL [W,E=
FULL-SCALE WIND-TUNNEL INVESTIGATION
OF A VTOL AIRCRAFT WITH A JET-EJECTOR
SYSTEM FOR LIFT AUGMENTATION
by Jerry V. Kirk a n d David H . Hickey
Ames Research Center
Moffett Field, Gal$
N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. N O V E M B E R 1 9 6 6 ERRATA NASA Technical N o t e D - 3 7 2 5 FLTLJ;-SCAI;E WIND-TUNNEL INVESTIGATION O F A VTOL AIRCRAFT WITH A JET-EJECTOR SYSTEM FOR LIFT AUGMENTATION B y Jerry V . Kirk and D a v i d H. H i c k e y Page 29, Figure 13: -For Tc = 1.2, a l l data points are 1.00 CL high. For example, a t .< should be 2.0 rather than 3.0.
a = O
, CL
Issued 6-8-67 NASA-Langley, 1967 TECH LIBRARY KAFB, NM 0130543 NASA TN D-3725 FULL-SCALE WIND-TUNNEL INVESTIGATION O F A VTOL AIRCRAFT m T H A J E T - E J E C T O R SYSTEM FOR LIFT AUGMENTATION By J e r r y V. Kirk and David H. Hickey A m e s R e s e a r c h C e n t e r Moffett Field, Calif.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $2.00 FULL-SCALE WIND-TUNNEL INVESTIGATION OF A VTOL AIRCRAFT WITH A JET-WECTOR S Y S m FOR LIFT AUGMENTATION By Jerry V. Kirk and David H. Hickey Ames Research Center SUMMARY Aerodynamic characteristics of a VTOL aircraft incorporating a jet ejector for augmenting lift have been examined from hover up to and including wing-supported flight.
Ejector performance was measured statically and with forward speed. The maxi" static augmentation ratio measured was 1.19. The ratio of ejector thrust with forward speed to static ejector thrust increased with forward speed.
In general, the aircraft had nearly neutral longitudinal stability at angles of attack below wing stall, but above wing stall, pitch-up was severe.
Lateral and directional stability were positive. Control power for trim in transition appeared to be adequate except for recovering from post stall pitch- up. Results from tests in a full-scale wind tunnel, in flight, and in a small-scale wind tunnel generally agree well.
INTRODUCTION Many concepts have been suggested for augmenting the lift of fixed wing aircraft for vertical take-off and landing. One concept is a jet ejector system such as that used for augmenting lift on the Lockheed XV-&A.
The full-scale aerodynamic characteristics of this aircraft and its ejector system were investigated in a wind tunnel at conditions ranging from hover up to and including conventional wing-supported flight. Ejector per- formance, longitudinal characteristics, lateral-directional stability and control, and control power about all three axes were determined at various airspeeds and control settings through the transition flight regime. The results are compared with those from flight tests (ref. 1) and from a small- scale wind tunnel (ref. 2).
NOTATION A ejector exit area, sq ft b wing span, ft
-
mean aerodynamic chord, - E L b l 2 c2 dy
C D
drag c o e f f i c i e n t , -
CD qs L
rolling-moment c o e f f i c i e n t , -
C l qSb ...
L
l i f t c o e f f i c i e n t , -
cL qs M
pitching-moment c o e f f i c i e n t , -
Cm qSF I \
yawing-moment c o e f f i c i e n t , -
Cn qSb Y
side-force c o e f f i c i e n t , -
q s D drag, l b
2 r o l l i n g moment , f t -1b
t o t a l l i f t on a i r c r a f t , l b L m mass flow, ~ A V ~ , slugs/sec
pitching moment , f t -1b
M N yawing moment, f t -1b standard atmospheric pressure, 2116 lb/sq f t P O t e s t section s t a t i c pressure, Ib/sq f t PS engine t a i l - p i p e t o t a l pressure, i n . H g P t P free-stream dynamic pressure, lb/sq ft R Reynolds number S wing area, sq f t T complete e j e c t o r t h r u s t i n t h e l i f t direction, pAvj‘, l b T
t h r u s t c o e f f i c i e n t , -
T C qs a i r velocity, f t / s e c free-stream velocity, knots s i d e force, lb angle of a a t t a c k of t h e wing chord plane, deg s i d e s l i p angle, deg P
dens i t y , lb- s ec2/f t4
P PS
6 r e l a t i v e s t a t i c pressure, -
PO aileron deflection measured normal t o t h e hinge l i n e , l e f t a i l e r o n Sa down, positive, deg elevator deflection measured normal t o the hinge l i n e , t r a i l i n g edge down, p o s i t i v e , deg f l a p deflection measured normal t o t h e hinge l i n e , deg rudder deflection measured normal t o t h e hinge l i n e , t r a i l i n g edge
r i g h t , p o s i t i v e , deg
Subscripts e e j e c t o r j e j e c t o r e x i t S s t a t i c U uncorrected a variable angle of a t t a c k AIRCRAFT DESCRIPTION The Lockheed XV-4A i s a twin-engined midwing monoplane incorporating a propulsion system and a l l controls necessary f o r v e r t i c a l take-off and landing (VTOL) and f o r t r a n s i t i o n t o conventional wing-supported f l i g h t . In f i g u r e 1 i s shown mounted on t h e normal s t r u t system i n t h e t e s t section t h e a i r c r a f t of t h e 40- by 80-foot wind tunnel. Figure 2 i s a two-view drawing of t h e a i r c r a f t .
Propulsion System The exhaust from two JTl2A-3 t u r b o j e t engines, which is directed a f t i n t h e normal manner f o r conventional f l i g h t , is directed to a ducted manifold along t h e upper fuselage f o r VTOL and t r a n s i t i o n t o wing-supported f l i g h t .
The manifold has 40 elongated nozzles i n 20 rows of 2 each (see f i g . l(a)).
Each nozzle i s canted 1 0 ' a J 3 of t h e v e r t i c a l plane. The flow of exhaust gases from t h e nozzles i n t o mixing chambers i n t h e fuselage provides e j e c t o r action. Doors above and below t h e e j e c t o r s are open during VTOL and tran- s i t i o n f l i g h t .
The d i f f e r e n t propulsion configurations employed during t r a n s i t i o n from
hover t o wing supported f l i g h t studied were: Configuration I - exhaust from
both engines being diverted t o t h e e j e c t o r manifold; configuration I1 - one
engine exhausting conventionally and one exhausting through t h e e j e c t o r
manifold; configuration I11 - both engines exhausting conventionally, but
with t h e e j e c t o r i n l e t and e x i t doors open ( j u s t p r i o r t o complete conversion t o conventional f l i g h t ) .
Hover Controls Aircraft a t t i t u d e about a l l t h r e e axes is controlled with reaction jets during hover and low-speed f l i g h t . During VTOL operation, approximately 10 percent of the engine exhaust gas flows continuously to the p i t c h and yaw nozzles located i n t h e nose and t a i l of t h e a i r c r a f t . For longitudinal control, t h e division of gas flow between t h e f o r e and a f t nozzles is varied; f o r d i r e c t i o n a l control, t h e nozzles i n t h e nose and t a i l are swiveled i n opposing d i r e c t i o n s . Engine compressor bleed a i r is supplied, on demand only, t o roll-control valves on t h e upper and lower surface of each wing t i p .
Boundary-Layer Control The a i r c r a f t has blowing boundary-layer control over t h e leading edge of t h e horizontal s t a b i l i z e r and elevators to prevent t h e a i r flow from sepa- r a t i n g and to increase control power during t r a n s i t i o n f l i g h t . Engine compressor bleed a i r w a s used f o r the BLC system. The BLC w a s on f o r a l l t e s t s unless otherwise noted.
Conventional Controls The a i r c r a f t has ailerons, a rudder, and elevators f o r conventional f l i g h t . Both t h e ailerons and rudder d e f l e c t f20°. The elevators have two control limits t h a t d i f f e r with t h e f l i g h t configuration: For conventional f l i g h t t h e l i m i t s a r e +30°; during hover and t r a n s i t i o n a l f l i g h t (config- urations I and II), t h e limits a r e from 0 ' to 60°, t r a i l i n g edge d o m .
Neutral elevator f o r t h e t r a n s i t i o n a l mode i s approximately 26O, t r a i l i n g edge down.
TESTING PROCEDURE angles of a t t a c k Six-component f o r c e and moment data w e r e measured at from -le0 t o + 2 8 O . The s t a t i c performance of t h e e j e c t o r augmentation system was measured with a force balance and a pressure and temperature rake mounted beneath t h e left-hand e j e c t o r bay. Thermocouples and pressure transducers on t h e rake provided t h e data needed t o calculate t h e m a s s flow.
For most of t h e t e s t s half power w a s used because t h e l i f e of t h e e j e c t o r primary nozzle was limited at f u l l power. The data are presented i n terms of t h r u s t c o e f f i c i e n t s o t h a t t h e results can be adjusted t o other power conditions. Airspeed w a s varied from 0 t o 100 knots; a t t h e highest speed, w a s 4.7 million.
Reynolds number T e s t s a t Constant Angle of Attack Power, angle of s i d e s l i p , and longitudinal, lateral, and d i r e c t i o n a l control s e t t i n g s were varied a t a constant angle of a t t a c k . The angle of a t t a c k w a s varied with airspeed s o as t o obtain data f o r t h r u s t nearly equal t o drag. Aircraft configurations I and I1 were studied i n t h i s manner.
Ejector t h r u s t w a s measured f o r a representative combination of power s e t t i n g s and angles of a t t a c k .
Variable-Angle-of-Attack Testing Configuration variables and airspeed were held e s s e n t i a l l y constant while angle of a t t a c k w a s varied. For t h e majority of tests, t h e angle of a t t a c k w a s first set f o r zero drag w i t h longitudinal control s e t t o trim pitching moment near zero. Angle of a t t a c k w a s then varied. I n most instances t h e angle-of-attack range included t h e maximum nose-down longi- tudinal control available.
Corrections t o Data Force and moment d a t a f o r t h e conventional configuration ( j e t augmen- t a t i o n system not operating) were corrected f o r t h e e f f e c t s of wind-tunnel wall interference i n t h e following manner: No corrections w e r e applied t o t h e data f o r t h e t r a n s i t i o n configuration ( j e t augmentation system operating) since t h e e f f e c t of e j e c t o r a i r flow on wind- tunnel w a l l corrections w a s not known.
A major p a r t of t h e t e s t program w a s run with no f a i r i n g on t h e t a i l strut. Appropriate drag tare corrections have been applied t o t h e data to account f o r t h e t a i l s t r u t drag. Near t h e end of t h e program, a f a i r i n g w a s placed over t h e t a i l strut as a check on t h e v a l i d i t y of the corrections.
The r e s u l t s agree within t h e accuracy of t h e measurements.
RFSULTS Table I i s an index to t h e f i g u r e s . I n general t h e low speed config- uration I results a r e referenced to s t a t i c t h r u s t at hover.
Propulsion System Performance Unflagged symbols S t a t i c t h r u s t of t h e a i r c r a f t i s shown i n f i g u r e 3 ( a ) .
include t h e t h r u s t from t h e p i t c h reaction controls, while t h e flagged symbols a r e measurements of e j e c t o r t h r u s t only. The e f f e c t of forward speed on e j e c t o r performance (from pressure measurements) i s shown i n f i g u r e 4. The r e s u l t s i n f i g u r e 4 are calculated from pressure and temperature measurements.
Aerodynamic Characteristics The v a r i a t i o n i n a i r c r a f t longitudinal c h a r a c t e r i s t i c s with forward speed f o r two configurations and two power s e t t i n g s is shown i n f i g u r e 5 .
Also shown is a comparison of e j e c t o r t h r u s t with measured a i r c r a f t l i f t .
Figures 6 through 13 present t h e v a r i a t i o n i n a i r c r a f t longitudinal c h a r a c t e r i s t i c s with angle of attack.
Figures 14 and 15 show t h e pitching-
moment v a r i a t i o n with angle of a t t a c k f o r various longitudinal control s e t t i n g s , power s e t t i n g s , and forward speeds. These results a r e taken from f i g u r e s 9 through 13.
The v a r i a t i o n i n l a t e r a l - d i r e c t i o n a l c h a r a c t e r i s t i c s at constant angle of a t t a c k is shown i n figures 16 and 17.
Control Power Figures 18 through 21 show t h e longitudinal control power available both with and without t h e boundary-layer control operating over the hori- zontal t a i l .
Lateral control power of t h e XV-4A is shown i n f i g u r e s 22 through 24, and d i r e c t i o n a l control power i n figures 23 through 30.
DISCUSSION Ejector and Airplane Performance Zero speed.- The maximum s t a t i c t h r u s t measured w a s 6400 pounds ( f i g .
3).
This value includes t h e unaugmented t h r u s t from t h e p i t c h reaction controls.
Engine t h r u s t i n t h e conventional configuration a t t h e same engine pressure r a t i o w a s 5450 pounds. Subtracting t h e unaugmented t h r u s t gives an augmen- t a t i o n f a c t o r f o r the e j e c t o r of 1.19. Mass flow calculated f o r t h e maximum t h r u s t condition w a s 1 5 . 8 slugs p e r second corresponding t o a weight rate of The weight r a t e of flow t o t h e flow of approximately 500 pounds per second.
e j e c t o r f o r t h e two J T l 2 A - 3 gas generators i s approximately 90 pounds p e r second; therefore t h e flow augmentation r a t i o of t h e e j e c t o r w a s about 5 . 5 t o 1.
Forward speed.- The r a t i o of e j e c t o r t h r u s t t o s t a t i c e j e c t o r t h r u s t I ( f i g . 4 ) . Agreement i s good increased with forward speed f o r configuration f o r t h e three power s e t t i n g s shown. The low and high power r e s u l t s i n figure f o r configuration I c o r r e l a t e well except f o r t h e high power, high t h r u s t 5 ( a ) c o e f f i c i e n t r e s u l t s f o r l i f t . Figure 5 ( b ) shows t h e measured a i r c r a f t l i f t t o s t a t i c t h r u s t r a t i o compared with e j e c t o r t h r u s t and e j e c t o r t h r u s t plus power off wing l i f t . There i s p o s i t i v e induced l i f t f o r t h e e n t i r e t h r u s t c o e f f i - c i e n t range shown. N o l i f t droop with forward speed i s apparent. A t 50 knots (Tc = 7 f u l l s c a l e ) t h e measured l i f t t o s t a t i c t h r u s t value represents about a 15 -percent overload c a p a b i l i t y f o r STOL operation.
5 ( c ) ) a r e a l s o presented as t h e r a t i o of Configuration I1 r e s u l t s ( f i g .
forces and moments t o s t a t i c t h r u s t . Good correlation is shown f o r t h e l o w and high power s e t t i n g s .
S t a b i l i t y and Control Longitudinal s t a b i l i t y and control.- For configuration I s t a t i c longi- tudinal s t a b i l i t y throughout the t r a n s i t i o n from hover t o wing-supported f l i g h t w a s n e u t r a l t o s l i g h t l y unstable ( s e e f i g . 1 4 ) . The pitching -moment curve w a s f a i r l y l i n e a r and no pitch-up problems were apparent i n t h e angle- of-attack range below s t a l l f o r speeds up t o 70 knots. Configuration I1 has a pronounced nose-down pitching-moment v a r i a t i o n w i t h airspeed below a t h r u s t c o e f f i c i e n t of approximately 1.4 ( f i g . 5 ( c ) ) . This pitch-down is probably m a s s flow through the due t o decreased e j e c t o r effectiveness (decreased e j e c t o r due t o e j e c t o r i n l e t flow separation) and t o increased trailing-edge forward speed i s increased.
f l a p contribution t o pitching moment as The fixed incidence horizontal t a i l of t h e XV-4A is mounted above t h e v e r t i c a l t a i l ( T t y p e ) . The a i r c r a f t w a s t e s t e d far above t h e s t a l l i n g angle of a t t a c k with longitudinal controls s e t near trim and with full nose-down control (aerodynamic plus reaction) t o examine t h e longitudinal s t a b i l i t y Wind-tunnel r e s u l t s indicate t h a t a t and control i n t h e deep stall region.
low speeds (approximately 25 t o 35 knots) t h e r e is s u f f i c i e n t control t o trim . .. ..
t h e a i r c r a f t through t h e m a x i m u m angle of a t t a c k t e s t e d (28O), although pitch- up occurs between 16' and 20°, depending on forward speed ( f i g s . 1 4 ( a ) and ( e ) ) . A t higher forward speeds control i s not s u f f i c i e n t t o t r i m t h e pitching moment at t h e higher angles of attack. A t approximately 35 knots ( s c a l i n g t h e r e s u l t s t o f u l l power conditions), t h e angle of a t t a c k beyond which t h e air- c r a f t cannot be trimmed is approximately 26O ( f i g . 1 4 ( c ) ) , while a t approx- imately 70 knots t h e m a x i m u m angle of a t t a c k f o r trim is 1 8 O ( f i g . 1 4 ( d ) ) .
Control effectiveness is s i g n i f i c a n t l y decreased a t elevator deflection angles above 40° ( f i g s . 19 through 21) f o r configurations I and I1 both with and without t h e h o r i z o n t a l - t a i l boundary-layer control system operating.
Lateral s t a b i l i t y and control.- Lateral s t a b i l i t y w a s positive f o r a l l configurations and angles of a t t a c k examined. Dihedral e f f e c t at low air- speeds i s p o s i t i v e , but as forward speed i s increased, t h e dihedral e f f e c t decreases ( see f i g . 17 ) .
Lateral control power f o r configuration I is not symmetric ( f i g s . 23 and 24) because one of t h e r o l l control valves malfunctioned i n t h e r i g h t wing down position.
Directional s t a b i l i t y and control.- Directional i n s t a b i l i t y f o r config- uration I11 a t 80 angle o f a t t a c k ( f i g . 16) is probably caused by t h e flow over t h e v e r t i c a l f i n being e f f e c t i v e l y blocked by t h e wing, l a r g e fuselage- mounted engine nacelles, and t h e e j e c t o r i n l e t and exit doors as angle of a t t a c k is increased. Configuration I (see f i g . 17) had positive d i r e c t i o n a l s t a b i l i t y . Directional control power r e s u l t s show no unusual c h a r a c t e r i s t i c s f o r t h e conventional configuration and f o r configuration I ( f i g s . 25 through 3 0 ) .
Comparison of Full-Scale Wind-Tunnel Results With Flight-Test and Small-scale Results Correlation of f u l l - s c a l e wind-tunnel and f l i g h t tests.- Figure 31 presents a comparison between r e s u l t s from the 40- by 80-foot wind tunnel and f l i g h t t e s t s . F l i g h t - t e s t data were very limited and t h e r e s u l t s shown a r e from a decelerating t r a n s i t i o n . Angle of a t t a c k compares very favorably, t h e discrepancy being on t h e order of lo. Elevator deflection f o r trim i s not f u l l y under- disagrees by 5 O t o 7 O . The reason f o r t h i s discrepancy stood, but t h e r e s u l t s from wind-tunnel measurements and f l i g h t t e s t s were obtained from two d i f f e r e n t a i r c r a f t , and possible rigging differences could account f o r some of t h e discrepancy. The a i r c r a f t t e s t e d i n t h e wind tunnel had been flown conventionally and hovered but had never been flown through a t r a n s i t i o n .
Comparison between f u l l - s c a l e and small-scale (18 -percent) model r e s u l t s .- Data f o r an 18-percent scale model of t h e XV-4A ( r e f . 1) are compared w i t h f u l l - s c a l e data i n f i g u r e 32. A l l f u l l - s c a l e data were corrected t o t h e same t h r u s t c o e f f i c i e n t as t h e small-scale r e s u l t s . The e f f e c t s of reaction control were subtracted from t h e f u l l - s c a l e r e s u l t s because t h e small-scale model did not have reaction controls. The r e s u l t s agree up t o 4' angle of a t t a c k . Above 6 ' angle of a t t a c k t h e small-scale l i f t and moment r e s u l t s d i f f e r markedly from t h e f u l l - s c a l e r e s u l t s .
SUMMARY O F RESULTS A f u l l - s c a l e wind-tunnel investigation of a VTOL a i r c r a f t incorporating a j e t e j e c t o r f o r l i f t augmentation has shown: 1. The m a x i " s t a t i c thrust augmentation r a t i o is 1.19 with a flow 5.5.
augmentation r a t i o of approximately 2. The r a t i o of e j e c t o r t h r u s t with forward speed t o s t a t i c e j e c t o r t h r u s t increased with forward speed f o r configuration I.
3. Positive l i f t is induced f o r t h e e n t i r e t h r u s t c o e f f i c i e n t range i n configuration I.
4.
Longitudinal control at high angles of a t t a c k (beyond C h m ) and forward speeds above 55 knots i n configuration I w a s not s u f f i c i e n t f o r trimming t h e a i r c r a f t .
5 . The dihedral e f f e c t a t low forward speeds is p o s i t i v e but decreases as forward speed is increased.
6 . Directional i n s t a b i l i t y w a s apparent as angle of a t t a c k w a s increased i n t h e phase I11 configuration.
7. R e s u l t s f r o m f u l l - s c a l e wind-tunnel t e s t s , f l i g h t t e s t s , and s m a l l - s c a l e wind-tunnel t e s t s generally agreed favorably.
Ames Research Center National Aeronautics and Space Administration Moffett Field, C a l i f . , Aug. 16, 1966 721-01-00-08-21 1. Cook, Woodrow L., and Hickey, David H . : Comparison of Wind-Tunnel and Flight-Test Aerodynamic D a t a i n t h e Transition-Flight Speed Range f o r NASA SP-116, Paper No. 26, 1966, pp. 447 -467.
Five V/STOL A i r c r a f t .
2. Barnes, 0. G.: The VZ-10 Hummingbird 0.18 Scale Model Wind Tunnel T e s t Results. Lockheed ER-5623, April 1962.
1 0 k I a3 x P k P a , 'T3 a , d cd a , a ,
-2 d
-P
Y
0 %
-
cd W I d A-33194 (b) Three-quarter f r o n t view.
P'igure 1. - Yoncluded .
1 2 f t E = 4 . 4 3 f t -1 ~ I 5 . 5 5 f t L.
-- p 7 - 7 - T l F - ..
I f t Horizontal Vertical Wing tail tail Aspect ratio 6.0 4 . 3 0 1 . 1 9 26.44 21.93 Area 104.17
+ 3.54 f t t -
.388 .401 ,558 Taper ratio
1 u
2 . 3 3 f t -6.0 f t- m Figure 2.- General arrangement of the Lockheed XV-4A Hummingbird..
6 0 0 0 TI6 I O 0 0 ( a ) S t a t i c t h r u s t .
m 8 0 4 8 12 1 6 20 24 28 32 36 pt, -Po S t a t i c mass flow, a = oO.
(b) Figure 3.- Zero airspeed performance of t h e a i r c r a f t .
I .4 ptp -Po 0 1 6 0 24 I .2
- 1.0
Tse .a L Ejector t h r u s t as measured by pressure rake.
( a ) I .4 I .2 me
- 1 . 0
m Se .a I 0 IO 20 30 3 Tc ( b ) Ejector m a s s flow.
Figure 4.- Ejector performance with airspeed and power s e t t i n g ; a = Oo.
2.0 * a, Flagged symbols deg indicate high 0 6 I .6 power. 0 0 0 -6 I .2 L
-
T S .8 .4 .4 .2 D - 0 TS
-. 2
-. 4
15 20 (a) Configuration I.
with forward speed; 6 , = 30°,
Figure 5.- Aircraft longitudinal characteristics
zf = 400.
~ . . . . . . . .. . ,, I
I \
\
I .4 L L
- 1.2
T S or
- - -- - -
-,- -- - ,
--
T -
- 1.0
T S Ejector thrust plus power off l i f t .8 3.2 I a , de9 0 6 2.8
n o
0 -6 2.4 Flagged symbols indicatc high power L =- 2.0 I S I . 6 .4 .2 I
- D o
T s
-. 2
-.4
-. 6 . I
-. I -. 2 -. 3 0 I 2 3 T c ( e ) Configuration 11.
Figure 5 . - Concluded.
1 8
1.8 ----
r-----
0 .I .2 .3 .4 .5 a , deg C D =feet of flap position and elevator setting; V = 8 0 knots.
(a) Figure 6. - Longitudinal characteristics with power off, conventional flight configuration.
Iu I .6 I .4 I .2
I .o
.a
C , .6 .4 c .L C c -. L -.4 I 0 4 8 1 2 1 6 20 24 26 .2 .I 0 -.I -.2 -.3 -4 - . 5 0 .I .2 .3 .4 .5 .6 . 7 .8 .9 a, deg CD c m Reynolds number e f f e c t ; 6, = O o , 6f = 409 (b) Figure 6. - Concluded.
i I 0 .I .2 .3 .4 .5 .6 . 7 a , deg CD C m Figure 7.- Longitudinal characteristics in the preconversion configuration; power o f f , 6 , = O o , 6f = bo0, V = 80 knots.
-1.2 -.8 - . 4 0 . 4 CD Figure 8.- Longitudinal characteristics in the conventional flight configuration with power on; low power, 6 , = Oo, 6f = bo, V = 8 0 knots.
1 . 3 ----
I ~ rrrr
o 23.7 33 20 Low 0 18.0 3 1 30 High 0 -.I - . 2 - . 3 .I - . 3 - . 2 -.I 0 .I .2 .3 .4 M
-
D
-
Ts F T S Figure 9. - Longitudinal characteristics in transition; configuration I, 6f = 40°.
Iu w I . 4 I .3 L - 1.2 Ts I . I
I .o
-8 -4 0 -.I -.2 -.2 -.I 0 .I .2 .3 .4 .5 .6 .7 D M
-
T C J
c
Figure 10.- Longitudinal characteristics in transition; configuration I, low power, 6f = bo, V = 30 knots.
CL vs C L and C , (a) Figure 11.- Longitudinal characteristics in transition; configuration I, low power, 6f = 40'.
I 7.2 6.8
6.4 K
CL
L 6.0
5.6 k
T , sa, V , Tail strut -
5.2 b deg knots fairing 0 5.7 30 40 O f f 0 5.7 52 40 O f f
4.8 -
40 On - 0 5.7
A 3.7 I 50 O f f
b 3.7 5 0 O f f
4.4 -
t
4 . 0 - ------ I I I
-
- . 4 0 . 4 . 8 1 . 2 1 . 6 2.0 2 . 4 2 . 8 3.2 3 . 6 4
' 4.4
CD -.4 0 .4 .8 1 . 2 1 . 6 2.0 2.4 2.8 32 Q, deg C D (a) v = 50 knots.
Figure 12.- Longitudinal characteristics in transition; configuration I, high power, 6f = 40'.
(b) V = 70 knots.
Figure 12.- Concluded.
.u 8 12 1 6 20 .3 .2 *I 0 -.I -.2 - . 3 -8 -4 0 4 -1.2 -.8 -.4 0 .4 .8
a, deg C n crn
2f- Figure 13. - Longitudinal c h a r a c t e r i s t i c s i n t r a n s i t i o n ; configuration 11, 6f = bo.
i ; CL-, kLL < -e cir- .2 I TC .I 23.7 M
-
Ts C
-. I
- 3 = 2 0 knots, low power.
.2 0 9.4 52 .I E . *. 4
- M O
I I T s E
-. I
-.2 V = 30 knots, low power.
(b) V = 40 knots, low power.
( c ) Figure 14.- Longitudinal stability in transition; configuration I, 6f = bo0.
v = 50 knots, low power.
( a )
.3 T c a , , v, deg Knots 0 2.5 3 0 50 .2 ~ 0 1.4 24 67 o .a6 2 4 a5
a
. I
-. I
-.2 -.3 ( a ) LOW p o w e r .
T c %, v, deg Knots 0 1.2 2 4 100 .L -10 - 5 0 5 I O 15 20 Q , de9 (b) High p o w e r .
Figure 15.- Longitudinal stability in transition; configuration 11, Sf = bo.
%, 8fl a3 deg deg deg Configuration 0 -18 40 0 Conventional .I 2 0 -18 0 0 Conventional 0 0 4 0 0 m A 0 4 0 8 m .08 CY
.04U
O I I
-.04 L l
- . O 8 I l -.I 2 .o 2 C" -.02 -.04 .04 .02 CZ - 0 3 -8 -6 -4 -2 0 2 4 6 B ,deg Figure 16.- Lateral-directional characteristics in transition; power o f f , v = 80 knots.
1.2 . 8 . 4 CY - . 4 - . 8 .08 . 0 4 C" 0 - . 0 4 - . 0 8 Figure 17. - Lateral-directional characteristics in transition; configuration I,.
low power, 6f = bo.
.4 crn
-. 4
-. a
(a) Power o f f .
.4
\
Y c m -.4
-. 8
-1.2 -30 - 20 -I 0 0 I O 20 30 Trailing edge up T r a i l i n g edge down
a , , d e g
(b) Power on, low power.
Figure 18.- Longitudinal control power in the conventional flight configuration; v = 80, a = oO, p = Oo.
v, TC knots 0 25.4 20 - 0 9 . 4 30 Flagged sym bok indicate BLC off (a) LOW power a = 0'.
V , a,
I T , knots deg
0 - 0 12 0 18.5 20 2 0 - M T,F 50 60 a,, deg (b) High power.
Figure 19.- Longitudinal control power in transition; configuration I, Q = 40°, = 00.
( a ) LOW power.
." 0 IO 20 30 40 50 60 B e , deg ( b ) High power.
Figure 20.- Longitudinal control power i n t r a n s i t i o n ; configuration I, 6f = 40°, p = 0'.
0 50 3.7 - 67 2.2 LOW power, a = 0 ' .
(a) 0 IO 20 30 40 50 60 Ee, deg (b) High power.
Figure 21.- Longitudinal control power in transition; configuration 11, sf = 4 0 ° , p = 0 ' .
CY .04 -.04 -.08 . O 8 .04 -.04 -15 -10 -5 0 5 I O 1 5 8 0 , deg Figure 22.- The e f f e c t of s i d e s l i p and f l a p deflection on lateral control power i n t h e conventional f l i g h t configuration; power o f f , V = 80 knots, 6, = -180, a = oO.
.04 Y
-
T S
-. 04
i i i 1 i i i -.08
-.o 2
- N O
Ts b -.02 .03 .02
.o I
T,b
-.Ol -20 -I 5 -I 0 -5 0 5 IO 1 5
a0, deg
Figure 23.- Lateral control power in transition; configuration I, high power, 6f = 40 , p = 0'.
. 08
cn .06 v, knots
.08 I
0 40 .04 0 5 0 -.04
C2
- . 0 8 ./ -.I 2
L
-.I 6 d
-.20 '
- 20 -15 -10 -5 0 5 IO 15
B a g deg Figure 24.- Lateral control power in transition; configuration I, high power, 6f = bo0, p = Oo.
.04 C" -.08 .04 CI 0 -.04 -15 -10 -5 0 5 IO 15 20 &, deg Figure 23.- = f e e t of s i d e s l i p and f l a p deflection on d i r e c t i o n a l control V = 80 knots, power i n t h e conventional f l i g h t configuration; power off a = 0'.
.o 2 Q 9 deg deg - 0 0 2 5 12 N T S b -.02 -.04 .03 .02 -. 01 - - - 1 3 0 5 I O 15 S , , deg Figure 26.- Directional control power i n hover; configuration I, high power, v = 0 , 6f = bo0, = Oo.
.04 s e , P , Tc deg deg
I I
.02 0 18.4 28 0 18.4 28 -10 -.02 -.04 -20 -15 -10 -5 0 5 I O 15 20 %, deg Figure 27.- Directional control power i n t r a n s i t i o n ; configuration I, low power, V = 20 knots, a , = O o , sf = 40'.
. 0 8 Y - Ts .04 -.O 4 .04 T , deg deg Power 0 2 - o 18.0 25 0 High 0 9.6 28 0 Low - . 0 2 -.04 .04 . 0 2 -.02 -20 -15 -10 -5 0 5 IO 1 5 2 0
B,, deg
Figure 28.- Directional c o n t r o l power i n t r a n s i t i o n ; configuration I, V = 30 h o t s , a = O o , Sf = 40°.
1.2 .8 CY .4 -4 .4-.4 a e , P, Q,
.3 -.a
Tc deg deg deg Power O 5.7 40 0 -6 Low 5.7 40 -8 -6 Low .2 10.2 40 0 High 0 0 . I C" 0 . I .2 .3 -20 -15 -10 -5 0 5 I O 15 20 a r , de9 Figure 29.- Directional control power i n t r a n s i t i o n ; configuration I, V = 40 knots, 6f = 40°.
. 3 .2 a , TC deg deg Power 0 3.7 33 -8 LOW
/
- 0 6.8 3 0 0 High . I
/-
Cn
-. I
-. 2
Figure 30.- Directional control power i n t r a n s i t i o n ; configuration I, v = 50 knots, sf = bo0, p = oo.
.2 Longitudina I acceleration
-. 2
0 Flight test data 0 Wind tunnel data Q, de9 0 10 20 30 40 50 60 70 80 V, knots Figure 31.- Comparison of r e s u l t s between f u l l - s c a l e wind-tunnel and f l i g h t tests.
4 8 Figure 32. - Comparison of full-scale and small-scale (0.18) model wind-tunnel results; configuration I, Tc = 6.23, 6f = 400, it = O o , 6 , = 0'.
c “The aeronautical and space activities of the United States shall be .
conducted so as to contribute . . . to the expansion of human knowl-
edge of phenomena in the atmosphere and space. T h e Administration shall provide f o r the widest practicable and appropriate dissemination of information concerning its activities and the results tbereof.” -NATIONAL AERONAUTICS A N D SPACE ACT OF 1958
NASA SCIENTIFIC AND T E C ~ N I C A L PUBLICATIONS
* I TECHNICAL REPORTS: Scientific and technical information considered important, complete, and a lasting contribution to existing knowledge.
TECHNICAL NOTES: Information less broad in scope but nevertheless of importance as a contribution to existing knowledge.
TECHNICAL MEMORANDUMS: Information receiving limited distri- bution because of preliminary data, security classification, or other reasons.
CONTRACTOR REPORTS: Technical information generated in con- nection with a NASA contract or grant and released under NASA auspices.
TECHNICAL TRANSLATIONS: Information published in a foreign language consid5red to merit NASA distribution in English.
TECHNICAL REPRINTS: Information derived from NASA activities and initially published in the form o f journal articles.
SPECIAL PUBLICATIONS: Information derived from or of value to NASA activities but not necessarily reporting the results .of individual NASA-programmed scientific efforts. Publications include conf&e&e proceedings, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.
. # Details on the availability of these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION DIVISION NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Washington, D.C. 20546