Document
NASA TECHNICAL
NOTE 133 -
N A S A T N D - 2
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u
M M c ev I m z c/I z
LARGE-SCALE WIND-TUNNEL TESTS
OF A N AIRPLANE M O D E L
W I T H A N U N S W E P T TILT W I N G
OF ASPECT RATIO 5 . 5 , A N D
W I T H VARIOUS STALL C O N T R O L DEVICES
by J u m e s A , W e i b e r g u n d D e m o J . G i u l i u n e t t i
A w e s Reseurch C e n t e r
M o f f e t t F i e l d , C a l i f o r n i a
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 . F E B R U A R Y 1 9 6 4
1 - - -
LARGE-SCALE WIND-TUNNEL TESTS O F AN AIRPLANE MODEL WITH AN UNSWEPT TILT WING OF ASPECT RATIO 5.5, AND WITH VARIOUS STALL CONTROL DEVICES By James A. Weiberg and D e m o J. Giulianetti 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 Office of Technical Services, Department of C o m m e r c e ,
Washington, D.C. 20230 -- Price $1.25
LARGE-SCALE WIND-TUNNEL TESTS O F AN AIRPLANE MODEL WITH AN UNSWEPT TILT WING O F ASPECT RATIO 5-37 AND WITH VARIOUS STALL CONTROL DEVICES By James A. Weiberg and Demo J. G i u l i a n e t t i SUMMARY Results are presented of t h e e f f e c t s of slats, f l a p s , wing-fuselage ramp f a i r i n g , and propeller r o t a t i o n on t h e flow separation, b u f f e t , and descent c h a r a c t e r i s t i c s of a tilt-wing deflected-slipstream VTOL model. The r e s u l t s indicated t h a t wing stall and r e s u l t i n g b u f f e t i n descending f l i g h t could be delayed approximately 15 knots by a slat, BLC nose f l a p , o r increased t r a i l i n g - edge f l a p effectiveness.
INTRODUCTION The t e s t s reported i n reference 1 indicated t h a t there would be a problem of a t t a i n i n g descent r a t e s of t h e order of 300 f e e t per minute a t low speed without buffeting. The buffeting w a s produced by air-flow separation from t h e center section of t h e t i l t e d wing which w a s not immersed i n t h e propeller slipstream.
To determine the extent t o which t h i s separation could be contained or delayed, t e s t s were made of the model with various leading-edge f l a p s and slats and a f a i r i n g between t h e t i l t e d wing and fuselage. To determine t h e magnitude of t h e buffeting, the o s c i l l a t o r y force a t t h e model support points and accelerations i n t h e model s t r u c t u r e were measured. The r e s u l t s of these tests are presented herein.
NOTAT I ON b wing span, f t c wing chord p a r a l l e l t o plane of symmetry, f t
c - mean aerodynamic chord ' S =! r'2~2 dy, f t
measured drag CD drag coefficient including t h r u s t , %S _ _ l i f t
CL l i f t coefficient, -
%S r o l l i n g moment C 2 rolling-moment c o e f f i c i e n t , % S b p-itching moment1 pitching-moment c o e f f i c i e n t , Cm goes .- - - .
=Moments are presented about t h e center shown i n f i g u r e 2 ( d ) .
yawing moment yawing-moment coefficient, Cn 90oSb side-force coefficient, side force CY %os wj
jet momentum coefficient , - VJ
g G acceleration of gravity, 32.2 ft/sec2 g angle of stabilizer relative to fuselage reference line, deg it propeller advance ratio, J nD propeller angular velocity, rps n propeller diameter, ft D free-stream static pressure, lb/sq ft p , total pressure in flap duct, lb/sq ft
pd '
free-stream dynamic pressure , lb/sq ft
4 , gas constant for air, 1713 ft2/sec2 R R wing area, sq ft S duct temperature, R Td thrust thrust coefficient,
TC '
%S
jet velocity for isentropic expansion Jm-1 - , fps
Y - 1 free-stream velocity, fps W weight rate of air flow through nozzle, lb/sec j spanwise distance, perpendicular to plane of symmetry, ft Y a angle of attack of fuselage reference line, deg sideslip angle, deg P flap deflection measured in plane normal to flap hinge line, deg wing tilt measured from a wing-down position having 8.3' incidence of the root chord with respect to the fuselage reference line, deg ratio of specific heats (1.4 for air), and glide angle, deg Y Subscripts f trailing-edge f l a p n BLC nose f l a p MODEL AND APPARATUS The model f o r these t e s t s i s shown i n figure 1 and w a s t h e same model used i n reference 1.
The wing could be t i l t e d 20°, 30°, and 50' from a wing-down position a t which t h e incidence of t h e root chord w a s 8 . 3 O with respect t o t h e fuselage reference l i n e . Except where noted on t h e figures, a l l t h e t e s t s w e r e made with t h e v e r t i c a l t a i l off and t h e horizontal t a i l a t 14O incidence t o t h e fuselage reference l i n e . The geometry of the model and d e t a i l s of t h e stall control devices are shown i n t a b l e I and i n f i g u r e 2. The stall control devices a leading-edge slat ( f i g .
consisted of 2 ( b ) ) , a BLC nose f l a p ( f i g . 2 ( b ) ) , and a ramp-type f a i r i n g between t h e fuselage and t h e leading-edge of t h e t i l t e d wing ( f i g . 2 ( c ) ) . This l a t t e r ramp-type f a i r i n g w i l l be r e f e r r e d t o as t h e ramp. For some t e s t s , only the portion of the ramp a f t of the wing leading edge w a s t e s t e d .
This portion of t h e ramp w i l l be referred t o as t h e aft ramp. The BLC nose f l a p completely spanned t h e wing. The various spanwise extents of slats t e s t e d a r e shown i n figure 2 ( b ) . The full-span slat extended over t h e fuselage with t h e wing t i l t e d and ended at t h e side of t h e fuselage with t h e wing down.
The blowing boundary-layer-control system used on t h e f l a p s i s described i n reference 2. The height of t h e j e t nozzle w a s 0.060 inch on t h e trailing-edge f l a p s and 0.080 inch on t h e leading-edge f l a p s . The momentum flow c o e f f i c i e n t , Cp, used on t h e f l a p s w a s above t h e c r i t i c a l value defined as t h e minimum f l o w c o e f f i c i e n t above which l i t t l e or no change i n l i f t occurred. For t h e t r a i l i n g - edge f l a p s t h i s value w a s determined from the d a t a presented i n reference 2. For t h e BLC nose f l a p , t h e c r i t i c a l w a s determined from t h e data i n figure 3.
C,, The model w a s equipped with 4 three-bladed propellers. The geometric character- i s t i c s of these propellers a r e given i n reference 3. The blade angle a t 0.75 blade radius w a s U.5'. The majority o f t h e d a t a w a s obtained with t h e propellers r o t a t i n g i n t h e d i r e c t i o n shown i n f i g u r e 2(a) and, unless noted, t h e data pre- sented w e r e with t h i s r o t a t i o n .
Tests were made at free-stream v e l o c i t i e s from 0 t o 93 f p s (9, = 10 and a Reynolds number of 2.8 million based on t h e wing mean aerodynamic chord of 5.18 f e e t ) .
The data presented i n t h e f i g u r e s include t h e d i r e c t propeller forces as w e l l as t h e aerodynamic forces. The propeller t h r u s t c h a r a c t e r i s t i c s f o r Oo wing tilt are given i n reference 1 and are f e l t t o be s u f f i c i e n t l y accurate f o r all t h e tilt angles t e s t e d based on t h e d a t a i n reference 4.
Moments are presented about t h e center shown i n figure 2(d).
Tunnel-wall corrections w e r e not applied t o any of t h e data.
I
Measurements of t h e stall b u f f e t i n t e n s i t y w e r e made using t h e output from strain-gage-type load c e l l s mounted between t h e model and t h e support s t r u t s and from accelerometers located i n t h e model s t r u c t u r e . These data were recorded on an oscillograph.
REXULTS AND DISCUSSION The basic l i f t , drag, and pitching-moment c h a r a c t e r i s t i c s of t h e model are presented i n f i g u r e s 4 t o 9.
Data a r e presented f o r f l a p deflections of O o , 3 0 ' , 50'7 and 80° , wing tilt angles O f 0 ' , 2 0 ' , 30° , and 50' , and t h r u s t coef- f i c i e n t s T c l from 0 t o 4. The corresponding f i g u r e s f o r these configurations a r e l i s t e d i n t h e following t a b l e together with t h e device used t o control wing stall.
S t a l l control device Figure None None Full-span slat Full-span slat Full-span slat None N l - s p a n slat None Wing-fuselage ramp f a i r i n g Center slat Outboard slat with ramp Full-span slat Full-span slat with ramp BLC nose f l a p with ramp BLC nose f l a p with aft ramp 7(h) BLC nose f l a p Full-span slat Effect of S t a l l Control Devices on S t a l l and Buffet Tuft studies showed t h a t with t h e wing t i l t e d , flow separation originated on t h e wing center section over t h e fuselage and on an area between t h e nacelles on each wing panel ( f i g . 10). From these areas t h e separation spread spanwise with increasing angle of a t t a c k . The flow separation i n these areas w a s delayed t o higher angles of a t t a c k by e i t h e r the full-span slat o r the BLC nose f l a p which extended over t h e e n t i r e wing span (including t h a t p a r t o f t h e t i l t e d wing above t h e fuselage).
I n addition, t h e full-span slat and t h e BLC nose f l a p increased maximum l i f t and t h e angle-of-attack f o r maximum l i f t (e.g., see f i g s . 3(a) and (b); 6(a) and A slat over only t h e center section between t h e (b); 7(a) , (e) , ( f ) , (g)).
inboard nacelles d i d not noticeably alter t h e flow separation i n t h i s region nor provide any increase i n l i f t ( c f . f i g s . 7(a) and ( c ) ) . With t h e ramp f a i r i n g flow between t h e fuselage and t h e t i l t e d wing but with no leading-edge device, separation s t a r t e d i n t h e area between t h e nacelles on each wing panel and pro- gressed spanwise u n t i l it had spread onto t h e r e a r portion of t h e a f t ramp. The a f t ramp alone w a s as e f f e c t i v e as t h e complete ramp i n delaying the center sec- t i o n stall ( c f . f i g s . 7(g) and ( h ) ) . The ramp did not a l t e r the steady force c h a r a c t e r i s t i c s but did delay t h e onset of t a i l buffet due t o flow separation from the wing.
The flow separation and stall w e r e accompanied by buffeting of t h e model.
Measurements were made of t h e magnitude of t h e fluctuating forces involved.
These fluctuations were random i n magnitude as shown by t h e t y p i c a l oscillograph t r a c e s i n figure 11. The frequency of t h e o s c i l l a t i o n s corresponded t o t h e n a t u r a l frequency of t h e model s t r u c t u r e . The maximum peak-to-peak fluctuations were measured f r o m t r a c e s such as those i n figure 11. The v a r i a t i o n s with angle of a t t a c k o f half of t h e peak-to-peak value of these fluctuations f o r some of t h e configurations t e s t e d a r e shown i n figures 1 2 and 13. The stall buffet intensi- t i e s shown were measured with an accelerometer on t h e t i p of t h e horizontal t a i l and f r o m a load c e l l on one of t h e model support s t r u t s . The accelerations were measured perpendicular t o t h e h o r i z o n t a l - t a i l plane and t h e o s c i l l a t i n g forces, perpendicular t o t h e wing reference plane. These data a r e f o r a free-stream dynamic pressure of 10 p s f . Fromthese data, the angle o f a t t a c k f o r t h e start o f buffet r i s e w a s obtained and is shown i n t h e basic force data of figures 4 t o 9 by a shaded symbol. For the p l a i n wing, flow separation accompanied by buffet r i s e u s m y preceded C L ~ ~ , but w a s not necessarily accompanied by an abrupt change i n force c h a r a c t e r i s t i c s . I n general, t h e slat, BLC nose f l a p , and ramp configurations delayed flow separation and accompanying b u f f e t t o higher angles of a t t a c k so t h a t t h e buffet r i s e more nearly coincided with CL .
m X Effect o f S t a l l Control Devices on Operational Characteristics Flight t e s t s of a tilt-wing airplane ( r e f . 5 ) have shown t h a t t h e wing stall and separation ( r e f . 6) which occurs i n a d e s c e n t - o r a decelerating conversion leads t o buffeting and e r r a t i c motions with general d i f f i c u l t y i n handling t h e a i r c r a f t . For t h e present model, the e f f e c t s of t h e various stall control devices on t h e descent c h a r a c t e r i s t i c s i n the t r a n s i t i o n w e r e estimated. Glide angles were computed from t h e data i n f i g u r e s 4 t o 9 f o r an airplane having a wing loading of 50 p s f . Results are presented i n figures 14 f o r two configura- t i o n s showing t h e v a r i a t i o n of g l i d e angle throughout an angle of a t t a c k and t h r u s t coefficient range. Boundaries of stall b u f f e t rise and Ci.5' angle of a t t a c k a r e shown on these f i g u r e s . The stall boundary defines, f o r a given vel- o c i t y , t h e glide angle above which a rise i n the buffet magnitude w a s indicated from t h e measured fluctuating forces. This boundary w a s determined using data similar t o t h a t i n figures 1 2 and 13. Boundaries f o r stall buffet rise f o r t h e various configurations t e s t e d a r e compared i n f i g u r e s 13 t o 18. A s an indication of t h e descent rates a t t a i n e d with these g l i d e angles, a curve showing t h e g l i d e angle f o r 500 fpm descent is included on t h e f i g u r e s . Flap deflection, wing- fuselage ramp fairing, slats, and BLC nose f l a p s all increased t h e value o f A l l but f l a p deflection provided t h i s buffet-free g l i d e angle at a given speed.
increased g l i d e angle as a result of delay i n angle of a t t a c k before buffet.
Change i n propeller r o t a t i o n from r i g h t hand (used i n r e f . 1) t o t h a t shown in f i g u r e 2, reduced the center section stall with a subsequent increase in t h e Some d a t a were a l s o obtained with t h e g l i d e angle before b u f f e t (see f i g . 1 7 ) .
inboard propellers interchanged t o give down-going blades between the propellers.
This configuration a l l e v i a t e d flow separation between the nacelles on each wing panel but worsened t h e center section stall. This interchange of propeller r o t a t i o n did not a l t e r the steady force c h a r a c t e r i s t i c s but d i d reduce t h e g l i d e angle before b u f f e t .
Although most of these stall control devices provided sizable improvements 500 fpm could not be i n t h e descent c a p a b i l i t y , descent r a t e s g r e a t e r than a t speeds below 4 0 knots without b u f f e t even with the best configuration.
obtained The b e s t configuration investigated w a s blowing BLC trailing-edge f l a p s deflected 30' w i t h full-span leading-edge slats o r BLC nose f l a p , counterrotating propel- lers and wing-fuselage ramp f a i r i n g .
Lateral-Directional C h a r a c t e r i s t i c s The c h a r a c t e r i s t i c s of the model i n s i d e s l i p a r e shown i n figure 1 9 f o r 0 ' wing tilt and f l a p s Oo and 5 0 ' with the v e r t i c a l t a i l on, and 4 ' h o r i z o n t a l - t a i l incidence. N o l a t e r a l - d i r e c t i o n a l data were obtained on t h i s model with a l l t h e propellers r o t a t i n g i n the same d i r e c t i o n . However, comparisons with data from reference 3 on the aspect r a t i o 1 0 wing (from which the present model w a s made by cutting o f f t h e t i p s ) , indicate t h a t , a s expected, propeller r o t a t i o n has a l a r g e e f f e c t on 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 . The counterrotating propel- l e r arrangement of t h e present t e s t s eliminated o r g r e a t l y reduced the l a r g e CnP and C z P ; t h e l a r g e v a r i a t i o n of e f f e c t s of power on CY with angle of a t t a c k ( f i g . 20) was a l s o e s s e n t i a l l y eliminated. This l a t t e r l a r g e v a r i a t i o n of side force with angle of a t t a c k w a s shown t o impair t h e a i r p l a n e ' s handling q u a l i t i e s ( r e f . 7) but it is not necessarily t y p i c a l of multiengined propellered airplanes ( r e f . 8 ) .
CONDLUDING RENARKS The wind-tunnel t e s t s showed t h a t maximum l i f t of t h e tilt-wing deflected slipstream VTOL model w a s l i m i t e d by separation f r o m t h e center section of t h e wing outside t h e slipstream and from an area between the nacelles.
High-lift trailing-edge f l a p s and powerful leading-edge stall control devices delayed t h i s separation so t h a t some descent c a p a b i l i t y could be obtained without b u f f e t i n the t r a n s i t i o n speed range.
It w a s shown t h a t propeller r o t a t i o n had a l a r g e influ- ence on wing flow separation and on l a t e r a l and 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 .
Ames Research Center National Aeronautics and Space Administration Moffett Field, C a l i f . , Oct. 28, 1963.
b 1. Weiberg, James A . , and Holzhauser, C u r t A.: Large-Scale Wind-Tunnel Tests of an Airplane Model With an Unswept, T i l t Wing of Aspect Ratio 5.3, and With Four Propellers and Blowing Flaps.
N A S A TN D-1034, 1961.
2. G r i f f i n , Roy N . , Jr., Holzhauser, C u r t A., and Weiberg, James A.: Large- Scale Wind-Tunnel Tests o f an A i r p l a n e Model With an Unswept, Aspect- Ratio-10 Wing, Two Propellers, and Blowing Flaps.
MEMO l2-3-58A, N A S A 1958 ' 3. Weiberg, James A . , and Page, V. Robert: Large-Scale Wind-Tunnel Tests of an Airplane Model With an Unswept, Aspect-Ratio-10 Wing, Four-Propellers, and Blowing Flaps. N A S A TN D-25, 1959.
4.
Yaggy, P a d F., and Ragallo, Vernon L. : A Wind-Tunnel Investigation of Three Propellers Through and Angle-of-Attack Range from Oo to 8 5 O , N A S A TN D-318, 1960.
5. Pegg, Robert J.: Summary of Flight-Test R e s u l t s of t h e VZ-2 Tilt-Wing A i r c r a f t . NASA TN D-989, 1962.
6.
Mitchell, Robert G.: Full-scale Wind-Tunnel Test of t h e VZ-2 VTOL Airplane with P a r t i c u l a r Reference t o t h e Wing S t a l l Phenomena. N A S A TN D-2013, 7. Innis, Robert C . , and Quigley, Hervey C.: A Flight Examination of Operating Problems of V/STOL Aircraft i n STOL-Type Landing and Approach.
NASA TN D-862, 1961.
8.
Quigley, Hervey C . , and Innis, Robert C.: Handling Qualities and Opera- t i o n a l Problems of a Large-Four-Propeller STOL Transport Airplane.
NASA TN D-1647, 1963.
TABLE 1 . - GENERAL GEOMETRIC DIMENSIONS O F THE MODEL Horizontal V e r t i c a l Dimension Wing surface surface
-
30.6 Area, sq f t 145 .O 56.5 16.03 7 9 1 - 9 Span, f t 28-33 4.68 5.18 E , f t 3 -30 1.69 Aspect rat i o 5 -54 4-33 Taper r a t i o 9 45 .55 2.2O 0 Geometric t w i s t , deg (washout ) Dihedral from reference 0.8 plane, deg
- -
Incidence from reference 8 - 3
plane, 0 ' tilt , deg
NACA 0012 Section p r o f i l e (constant) NACA 23017 NACA 0012 6.07 4.61 5.88 Root chord, f t 4.18 2.54 2.65 Tip chord, f t Sweep of leading edge, deg 2 12 24 - - - - 1 8 . 0 3 ~ T a i l length, f t Distance from 0.25 E of wing t o 0.23 i5 of horizontal t a i l , Oo wing tilt.
I
A-28169 Figure 1.- The model with slats and wing-fuselage ramp f a i r i n g mounted i n the wind tunnel; 6f = 50°, 6 , = 30°.
P
\ / 2.92
I .27c Dimensions in feet
r - , 3 3.0 4 1
4.77 - I - I I '-Wing pivot (a) General dimensions.
Figure 2.- Geometry o f the model.
I
T
I Center Outboard Full span S l a t span n o t a t i o n Blowing nozzle detail Leading-edge slat detail Blowing nozzle Blowing flap (b) Details of f l a p s and slats.
Figure 2. - Continued.
Ramp fairing
i
tation 0 33.0 in lape ee' ( c ) Wing-fuselage ramp fairing.
Figure 2. - Continued .
STA 12.08
,$,OoTilt { W L 2.43
pivot { W L STA 14.31 1.61 Moment center Fuselage W L O 1 4 .- 1 3 1 2
I I I 1 1 ! . 1 ! I I I
t Y- L Q) c -= 2 I
0 1 0 30 40 50 60
Wing tilt, deg (d) Moment center v a r i a t i o n with wing tilt.
Figure 2. - Concluded.
I
C L 4 0 .02 .04 .06 .08 . I 2 . I 4 .I 6 . I 0 Figure 3.- Variation of lift coefficient with BLC nose flap momentum flow coefficient; 6 , = 30°, 6f = ~ o O , CPf = 0.063.
I
-
,4 -3 -2 -I 0 I 2 -30 -20 -10 0 I O 20 30 a CD Figure 4.- Aerodynamic characteristics of the model; 6 , = 0 , 6f = Oo, it = 1 4 ' .
CL I
-
4 -3 -2 -I 0 I 2 -30 -20 -10 0 I O 20 30 2 I 0 - I -2 C O d Cm (a) 6f = 30°, Cpf = 0 Figure 3 . - Aerodynamic characteristics of the model; 6, = Oo, it = 1 4 ' .
2 I 0 "-I -2 Cm = 0 , full-span s l a t .
Figure 5 . - Continued .
I -4 -3 -2 - I 0 I 2 -30 -20 -10 0 I O 20 30 2 I 0 - I -2 CD C Y . Crn ( c > 6f = 3oo, Cpf = 0.063, f a l - s p a n slat.
Figure 2. - Continued.
I I h , ,‘
E - , ’
I -4 -3 -2 -I 0 I 2 -30 -20 -10 0 I O 20 30 2 I 0 -I -2 C D d Cm (a) 6f = 80°, cPf = o ,088, full-span slat.
Figure 3 . - Concluded.
1 0 CL -4 -3 -2 -I 0 I 2 (a) 6f = 30°, CPf = 0 Figure 6.- Aerodynamic characteristics of the model; 6 , = 20°, it = 1 4 ' .
C ' L I - 4 -3 -2 - I 0 I 2 -30 -20 -10 0 I O 20 30 CD (b) 6f = 30 9 Cpf = 0, full-span slat.
Figure 6. - Concluded.
--
CL 5 I c, -30 -20 -10 0 IO 20 - 30
-
.4 -3 -2 -I 0 I 2 2 I 0 - I a CD Cm (a) 6f = 30°, Cpf = 0 Figure 7.- Aerodynamic characteristics of the model; 6, = 30°, it = 1 4 ' .
--===e-
CL
a
I -4 -3 -2 - I 0 I 2 -30 -20 -10 0 I O 20 30 2 I 0 - I C D a Cm (b) 6f = 30°, Cpf = 0, ramp.
Figure 7. - Continued .
a CL 5 I -4 -3 -2 -I 0 I 2 -30 -20 -10 0 IO 20 30 2 I 0 - I C D a C m ( e ) 6f = 30°, CPf = 0, center slat.
Figure 7.- Continued.
I O ...
2. I 3.9 CL 5 I -30 -20 -10 0 IO 20 30 0 - I 2 I -4 -3 -2 -I 0 1 2 C Y C m CD ’ CVf = 0, ramp and outboard s l a t .
(d) 6f = 30’ Continued.
Figure 7.- - - Td 0. I a I .o . . . . . _ _ .
~- --- - - . . . . . .............. __- -- ............ - . . . . . . .
+ L + - CL I -4 -3 -2 -I 0 I 2 -30 -20 -10 0 1 0 20 30 2 1 0 - I CD U Cm ( e ) 6f = 30°, Cpf = 0, full-span slat.
Figure 7.- Continued.
I -4 -3 -2 -I 0 I 2 -30 -20 -IO 0 IO 20 30 2 I 0 -I a CL Cm ( f ) 6f = 30°, CPf = 0, ramp and full-span slat.
Figure 7. - Continued .
7 . . . . . . _ _ I -4 -3 -2 -I 0 I 2 CY C D (g) 6f = 30°, CPf = 0, ramp and BLC nose flap, CPn = 0.070.
Figure Continued.
7.- CL 4 I 2 I 0 - I -30 -20 -10 0 I O 20 30 -3 -2 -I 0 a Cm C D 6f = 30°, ciq = 0, a f t ramp and BLC nose f l a p , CcLn = 0.070.
Concluded.
Figure 7.- W I 1 0 I CL I 2 -30 -20 -10 0 IO 20 30 2 I 0 -I -3 -2 -I 0 I d CD Cm Figure 8.- Aerodynamic characteristics of the model; S, = 30°, 6f = 5 0 ° , Cpf = 0.063, BLC nose flap, = 0.074, it = 1 4 ' .
ccln w - -30 -20 -IO 0 I O 20 30 0 I 2 -4 -3 -2 - I Q Cm CD Figure 9.- Aerodynamic characteristics of the model; 6 , = TO0, 6f = 30°, Cpf = 0, full-span slat, it = l b O .
w t - '
.. . .... . Rough f l o w
S t a l l e d Q = - 1 6 O a = - i 2 O Figure 10.- Wing flow separation patterns; S, = 30°, 6f = 30°, CPf = 0, Tc' = 1.1 (force d a t a i n f i g . 7 ( a ) ) .
Figure 11.- Examples of oscillograph records of buffet loads on t h e model.
I w c
f .8
Plain wing Wing- fuselage ramp fa iring
f . 6
Full span slat
* .4
* .2
f200
-
* * I 5 0 a CL
- * I O 0
E L
f50
-32 -24 -16 -8 0 8 -32 -24 -16 - 8 0 8 16
d d
T d =O. I
T d = 1.0 Figure 12.- Variation of fluctuating loads with angle of attack; 6 , = 30 , 6f = 30°, CClf = 0 .
9.0
-. .
Plain wing : L E + - -
BLC nose flap Cp.. = .074
* . 6
with fore and a f t ramp-
k.4
f.2
f 2 0 0
f150
2 100
* 50
-
-32 -24 -16 -8 0 8
-32 -24 -16 -8 0 8
a a
Ti= 0. I Ti= 1 . 0
cClf = 0.065. of attack; 6w = 30°, 6f = 30°, Figure 13.- Variation of fluctuating loads with angle b -10 0 0.1 iffet 0 1.0 undary 0 2.0
a 3 . 9
-30 -40 (a) 6 , = 00, 6f = 30°, cCLf = 0, p l a i n wing.
t c W 0 u) -20 W D -30 0 20 40 60 80 100 120 140 160 180 200 V knots (b) 6 , = 30°, 6f = 50°, CPf = 0.065, BLC nose f l a p , Ccln = 0.074.
Figure 14.- Variation of g l i d e angle with forward velocity.
1 . 1 I l l
, . I . ! . . . . ! ! . ! . I
descent -10 t c Q) 0 -20 v) Q) a -30 -40 (a) P l a i n wing.
-30 -40 0 20 40 60 80 100 120 140 160 180 200 V knots (b) Full-span s l a t .
Figure 15.- Effect of wing tilt on the stall b u f f e t boundary; 6f = 30°, CCLf = 0.
I.:
I i t
.n
E 2 0
1 0
50 .065
80 30 .088 O t
500 f pm
t descent L -10 si -20 (a) P l a i n wing h -10 t c -30 V Knots (b) Full-span slat.
Figure 16.- Effect of trailing-edge f l a p deflection on t h e s t a l l buffet boundary; 6 , = oo.
... . . . . .. ....... . ., , , .., . 1 . . - 1 . - - ..I.. I I, I I I , I I I 111 I I ~ '
-30
- 40
-10 -40 0 20 40 60 80 1 0 0 120 140 160 180 200 V Knots (b) 6-f = 5oo, CPf = 0.063, BLC nose f l a p , CPn = 0.074.
Figure 17.- Effect of propeller r o t a t i o n on the s t a l l b u f f e t boundary; 6 , = 30'.
I O
n
1 1 / I
/ I
E
.- e
7-r I /
m
I
0 Plain wing fairing 0 Wing-fuselage ramp b -10
. --
0 Full t c a
3 u
-20 ln a c3
I[
-30
I
-40 (a) 6f = 30°, CPf = 0 IO -10 t c a -20 l n a n -30 -40 180 200 0 20 40 60 80 100 120 140 160 V knots (b) 6f = 20°, CClf = 0.062, BLC nose f l a p , CPn = 0.074.
r.
6 , = 30".
s t a l l control devices on the buffet boundary; Effect of Figure 18.- - - - - . .
....
.O 6 . O 4 .o 2 Cn -.o 2 -.04 t l l l .o 2 cz 0 70 2 .2 -.4
-ill
- . 6 -4 0 4 a 1 2 4 8 -a -12 -a -4
a
B (a) 6f = O 0.065 Figure 19.- Lateral and directional characteristics in sideslip; 6 , = O o , a = oO, it = lkO, vertical tail on.
- -
prop rotat ion
viewed from front 1
20 24
a
Figure 20.- Effect o f propeller r o t a t i o n on t h e side-force variation with angle of attack; SW = O o , 6f = 50°, Cpf = 0.065.
NASA-Langley, 1964 A-752