Low-speed characteristics of a variable-sweep supersonic transport model with a blended engine fuselage and engine-mounted tails

?

&OW -SPEED CHARACTERISTICS OF A VARIABLE-SWEEP

SUPERSONIC TRANSPORT MODEL WITH A BLENDED ENGINE FUSELAGE AND ENGINE-MOUNTED TAILS By Vernard E. Lockwood and Wilson E. Thompson Langley Research Center Va.

Langley Station, Hampton, NOTICE T h i s document should not be returned after i t hos satisfied your requirements. It may be disposed of i n occordanca with your locol security regula- tions or the appropriate provisions of the Industrial Security Manual for Safe-Guarding C l a s s i f i e d lnformat i on.

NATIONAL AERONAUTICS AND SPACE ADMlN ISTRATION .

0 . 0.0 0 0 0 0 0 0 0 0 0 0 . 0 0 0 0 0 0 0 0 . 0 . 0 .

0 0 0 . 0 . 0 0 .

LOW-SPEED CFLARACTEHISTICS OF A VARIABLE-SWEEP SUPERSONIC TRANSPORT MODEL W I T H A BLENDED ENGINE FUSELAGE AND ENGINE-MOUNTED TAILS* By Vernard E. Lockwood and Wilson E. Thompson Langley Research Center SUMMARY An investigation has been made at low speeds to determine the longi and lateral stability and pitch control characteristics of a model of a proposed

supersonic transport airplane configuration designated SCAT 14. The geometric

.__ var.iables _ _ _ st-~died were r r i n o kslrg p m e l sweep, wing twist,; and variations in sweep and deflection of the fixed area between the fuselage and the movable wing. Lat,eral stability data were obtained for one model configuration t'nrough a range of

leading-edge sweep angles from 13.5' to 7 7 ' . The investigation was made in the

Langley 3OO-MPH 7- by 10-foot tunnel at a Mach number of 0.18 which corresponds

to a Reynolds number per foot of 1.24 X 10 .

The results showed the basic model w a s longitudinally stable throughout the angle-of-attack and sweep range investigated for a moment reference located at the wing pivot station; however, a reduction in stability occurred at angles of

attack greater than 80. Varying the wing angle fron: l3.5O to 75' gave a rear-

ward shift in the aerodynamic center of about 5.5 percent of the fuselage length

and reduced the untrimmed maximum lift-drag ratio from 14.3 to 6.6. A change

from the TO0 wing-fuselage flaps to the 7y0 flaps resulted in pitch-up for wing

sweep angles of 25' and 45O, a reduction in aerodynamic-center variation with sweep, and lower values of maximum lift-drag ratio.

Deflection of the TO0 wing-fuselage flap gave stabilizing moments at high angles of attack and also resulted in reduced aerodynamic-center variation with

panel sweep. - Varying the wing sweep angle from 13.5O to 7 5 O delayed the ons

of directional instability from angles of attack of 1 8 ' to 23.5'.

INTRODUCTION This paper contains the results of a low-speed investigation of a supersonic The configuration had a variable-sweep wing transport model designated SCAT 14.

with an outboard pivot and in this respect is similar to the model reported on in

*

Title, Unclassified.

reference 1. However, t h e aft fuselage design and powerplant geometry d i f f e r m a t e r i a l l y from t h a t of t h e reference model. The present configuration had a blended engine-fuselage arrangement with t h e intake duct located beneath t h e fuselage. The h o r i z o n t a l t a i l s were mounted on the s i d e s of t h e engines.

The i n v e s t i g a t i o n w a s concerned primarily with t h e e f f e c t of wing geometry on t h e longitudinal c h a r a c t e r i s t i c s . The geometric variables studied were wing panel sweep, wing t w i s t , and v a r i a t i o n s i n sweep and d e f l e c t i o n of t h e f i x e d area between t h e movable wing and t h e fuselage. Tail-off d a t a were a l s o obtained t o aid i n mderstanfiing t h e c m p l c t e m d e l resillts. h t c z - a l s t a b i l i t y data were obtained f o r one model configuration through a range of leading-edge sweep angles from 1 3 . 5 ' t o 75'.

The i n v e s t i g a t i o n w a s made i n t h e Langley 300-MPH 7- by 10-foot tunnel at a Mach number of 0.18 which corresponds t o a Reynolds number of 1,240,000 per f o o t .

SYMBOLS The force and moment d a t a contained herein a r e r e f e r r e d t o t h e a x i s system shown i n figure 1. The reference dimensions used i n reducing the d a t a based on the 75' swept wing are area, 7.00 square f e e t ; chord, 31.35 inches; and span, 38.25 inches.

The moment reference point i s located at t h e wing pivot, s t a t i o n (fuselage s t a t i o n 50.00) unless otherwise specified.

b wing span, i n .

Drag

drag c o e f f i c i e n t , -

CD q s L i f t

l i f t c o e f f i c i e n t , -

cL q s increment i n l i f t c o e f f i c i e n t &L C slope of l i f t curve a t CL = 0 LU Rolling moment rolling-moment c o e f f i c i e n t , qSb -1 . I . . . 0 . 0 . . ... . ... 0 .

increment in C 2 due to addition of vertical tail M 2

P P

Pitching moment pitching-moment coefficient, Cm PScref increment in pitching-moment coefficient E m 'AC m h horizontal control effectiveness parameter ml slope of pitching-moment-coefficient curve at CL = 0 * %/dCL Yawing moment yawing-moment coefficient, Cn qSb increment in Cn due to addition of vertical tail E n

B P

Side force side-force coefficient, CY qs E increment in C due to addition of vertical tail

yB yP

' r e f reference chord, i n . ' * L/D lift -drag ratio maximum lift-drag ratio (L/D) max 9 dynamic pressure, lb/sq ft S reference wing area, sq ft x, Y coordinates of wing-fuselage flap, in.

lower surface ordinate of wing-fuselage flap, in.

Yl upper surface ordinate of wing-fuselage flap, in.

YU a ar~g’~e of attack of fuselage reference line, deg w P angle of sideslip, deg increment in sideslip angle between 4 3 p = f 5 ’ , corrected for balance, and strut deflection, deg horizontal-tail dihedral (positive up), deg rh horizontal-tail deflection, deg ‘ h .

wing-fuselage flap deflection (positive when leading edge is down), %F deg A wing leading-edge sweep angle, deg Con t’ igurat i on de s i gnations : F fuselage E horizontal tail S sharp leading edge for wing-fuselage flap r round leading edge for wing-fuselage flap

v vertical tail

wing with planar lower surface w1 wing with linear twist, (2’ nose-down at tip) w2 WF wing-fuselage flap (see fig. 2) MODEL The model configuration features a variable-sweep wing with an outboard pivot location, a four-engine side-by-side arrangement which blends into the .

fuselage a t t h e rear of t h e model from an i n l e t located beneath the fuselage, and horizontal surfaces mounted from the sides of the engine ducts. A three- view drawing of t h e model i s presented i n figure 2 ( a ) and photographs of t h e model mounted i n t h e Langley 300-MPH 7- by 10-foot tunnel are shown i n figure 3.

Various model dimensions are given i n tables I and 11.

Fuselage cross sections drawn t o model scale a r e presented i n f i g u r e 4. It should be noted t h a t t h e sections i n the v i c i n i t y of t h e intake duct a r e solid, no provisions being made f o r i n t e r n a l flow.

one untwisted and Two s e t s of wings were used i n t h e investigation, W 1 Wing 2 w a s twisted l i n e a r l y about t h e 50-percent chord l i n e one twisted W2.

Both wings had i d e n t i c a l a i r f o i l sections ,from spanwise s t a t i o n 20.21 t o 40.00.

'which were developed from an NACA 65~006 section. The ordinates of t h i s section were sheared upward t o provide a f l a t bottom except i n the immediate v i c i n i t y of t h e leading edge where t h e nose sections were rourided t o provide a radius equal t o 0.007 chord.

The model was provided with replaceable f i l l e t s between the fuselage and t h e movable wing which served t o provide changes i n planform, deflection, and leading-edge contour. The f i l l e t s or wing-fuselage f l a p s are shown i n f i g - ures 2 ( b ) and 2 ( c ) and are described by the leading-edge sweep, t h e leading-edge contour ( s = sharp; r = round), and t h e deflection of f l a p i n a plane perpendic- WF = 7OosO0).

ular t o t h e hinge l i n e as shown i n f i g u r e 2(a) ( f o r example, When deflected, Only t h e f l a p with the TO0 sharp leading edge w a s deflected.

the break i n t h e upper surface of t h e f l a p was f a i r e d over t o provide a smooth The horizontal t a i l i s shown i n f i g - t r a n s i t i o n between adjoining surfaces.

ure 2 ( a ) ; a d d i t i o n a l dimensions a r e given i n t a b l e I. The t a i l dihedral angle w a s zero except where noted otherwise.

TESTS AND CORRECTIONS 300-MPH 7- by 10-foot tunnel with The i n v e s t i g a t i o n was made i n the Langley t h e model s t r u t supported from the f l o o r of the tunnel as shown i n f i g u r e 3.

Forces and moments were measured by an i n t e r n a l l y mounted six-component s t r a i n - gage balance attached t o t h e support strut. To insure a turbulent boundary- l a y e r t r a n s i t i o n s t r i p s approximately 1/8 inch wide of No. 100 carborundum grains

-

w e r e attached t o t h e model surfaces a t t h e 7-percent chord s t a t i o n .

The i n v e s t i g a t i o n was made at a dynamic pressure of 45.6 pounds per square * foot which corresponds t o a Mach number of 0.18 and a Reynolds number per foot All configurations were investigated through a range of angle of of 1.24 X 10 .

a t t a c k at 0 ' s i d e s l i p , and selected configurations were ais0 investigated a t a The drag data were corrected t o correspond t o a pressure s i d e s l i p angle of +5O.

a t t h e base of t h e engine nacelles equal t o free-stream s t a t i c pressure.

The jet-boundary corrections calculated f o r the drag and angle of a t t a c k by t h e method of reference 2 a r e as follows: ~ .............. .....

: : 0: : m : : . . * p . i r c s ; ; 0. ...

' I . . . . . . ...

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

. . . . .

' D = 'D,measured + ( 0 . 0 1 1 4 C : )

The jet-boundary corrections to the pitching-moment data were found to be negli- gible. The data were also corrected for wind-tunnel blockage hy the method ?re- sented in reference 3 . The angles of attack and sideslip were corrected for % deflection of the balance and sting under load. The effect of the support strut on the model characteristics is unknown but because of the thinness of this strut it is thought that the corrections to the data would be small.

PRESENTATION OF DATA The data obtained in the wind tunnel of the subject model are presented in the following figures: Figure Longitudinal characteristics: Effect of wing sweep and horizontal tail.

wl; WF = 700r00 . . . . . . . . . . . . . . . . . . . . . . . . . 3 to 6

w , ; WF = 70°r00 . . . . . . . . . . . . . . . . . . . . . . . . . 7

L

Effect of wing-fuselage flap geometry. W2; Fh = 0 ' . . . . . . . . 8

Effect of horizontal tail a t ; various wing sweeps.

W2; WF = 75°~00 . . . . . . . . . . . . . . . . . . . . . . . . . 9

Effect of wing sweep. W2; WF = 70°s00 . . . . . . . . . . . . . . . 10

Effect of wing-fuselage flap deflection at various wing sweeps.

w2; WF = 7oos . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Effect of horizontal tail at various wing-sweep and wing-fuselage

flap deflections. w2; WF = 7 0 % . . . . . . . . . . . . . . . . 12 to 14

Effect of wing sweep on the increment in pitching-moment coeffi-

cient due to addition of horizontal tail aC, and horizontal

control effectiveness q/Afjh. wl; WF = 70°r00 . . . . . . . . .

Effect of model geometry on variation of aC dCL, CL,, and (L/D)- ml with sweep.

Effect of horizontal tail . . . . . . . . . . . . . . . . . . . .

1 6

Effect of wing twist . . . . . . . . . . . . . . . . . . . . . . .

Effect of wing-fuselage flap geometry . . . . . . . . . . . . . .

Effect of wing-fuselage flap deflection . . . . . . . . . . . . . 19

Figure Effect of wing-fuselage f l a p geometry and wing sweep on t h e increment i n pitching-moment c o e f f i c i e n t due t o addition of

horizontal tail. FW2V . . . . . . . . . . . . . . . . . . . . . . 20

Increment i n pitching-moment and l i f t coefficients due t o

deflection of wing-fuselage f l a p . W2; WF = TO0 s2go . . . . . . 21

Lateral c h a r a c t e r i s t i c s : Effect of wing sweep and v e r t i c a l t a i l on lateral s t a b i l i t y c h a r a c t e r i s t i c s . . . . . . . . . . . . . . . . . . . . . . . . . 2 2 t o 2 4 DISCUSS I O N Longitudinal S t a b i l i t y The results of investigations of the basic m o d e l a r e discussed and compari- sons are made of the e f f e c t of changes i n model geometry on the aerodynamic c h a r a c t e r i s t i c s . The moment c o e f f i c i e n t data are r e f e r r e d t o t h e wing pivot location except where s t a t e d otherwise.

%sic model.- The data of f i g u r e 5(a) show t h e basic model (FWIHV; WF = 70°r00) i s longitudinally s t a b l e throughout the angle of a t t a c k and sweep range investigated f o r the moment reference located at t h e wing pivot s t a t i o n .

For intermediate sweep angles (25' t o 55') a reduction of s t a b i l i t y occurred at The reduced s t a b i l i t y o r itch-up c h a r a c t e r i s t i c angles of a t t a c k above 8 ' .

i s a l s o noted i n t h e t a i l - o f f data (FWIV) of f i g u r e 57b). Tuft studies made during t h e investigation show that the pitch-up i s t h e result of separated flow over t h e outboard panel. The flow separation begins at an angle of a t t a c k of

about 8O, and, except f o r a short span next t o the wing pivot, complete separa-

t i o n occurs on the wing panel at an angle of a t t a c k of 12'. The pitch-up tend- encies of t h e wing-fuselage combination a r e compensated t o a large extent by the presence of a low horizontal t a i l as shown i n figure 15. The increment i n

pitching-moment coefficient aC, due t o the horizontal t a i l increases with

Also presented angle of a t t a c k and i s generally independent of wing sweep angle.

determined from f i g -

i s t h e h o r i z o n t a l c o n t r o l effectiveness parameter aC, /ash

we 6. It i s noted t h a t t h e control effectiveness which generally decreases - between a = 0 ' and a = 12' results i n l e s s s t a b i l i t y with t h e horizontal t a i l d e f l e c t e d a t angles of a t t a c k g r e a t e r than 12'.

The v a r i a t i o n of basic model s t a b i l i t y with wing sweep f o r low l i f t c o e f f i - c i e n t s i s shown i n figure 16. This model, l i k e other variable sweep configura- t i o n s having a \ r e l a t i v e l y l a r g e fixed area ahead of t h e moment reference, shows a reduction i n s t a b i l i t y at high sweep angles. (For example, see r e f . 1.) The maXimUm change i n t h e aerodynamic-center location occurs between t h e l3.5O and

t h e 6 5 O wing sweep angle and amounts t o about 1 8 . 5 percent of t h e reference

chord. Between wing sweep angles of 13.5' and 7 5 O t h e aerodynamic-center

variation i s reduced t o

15.7 percent of the reference chord o r about 5.5 percent

of t h e fuselage length.

I The summary of low angle-of-attack c h a r a c t e r i s t i c s ( f i g . 1 6 ) shows the . - usual reduction i n l i f t - c u r v e slope C and untrimmed (L/D)- with sweep; L a

the values of (L/D)max varied from 14.3 with A = 1 3 . 5 ' t o 6.6 with A = 7 5 '

~ f o r configuration FWIHV.

Wing 2.- A comparison of f i g u r e 6 with figure 7 shows no s i g n i f i c a n t d i f f e r - ence between wing 1 and wing 2 i n the v a r i a t i o n of

C , with CL; i n f a c t , t h e -

small amount of t w i s t ( 2 ' nose down on W2) gave almost i d e n t i c a l values of &!,/aC, as i s shown i n figure 17.

Some s m a l l differences were noted, however,, f o r the l i f t - c u r v e slope and (L/D)m, between the t w o wings. I c L a I I Wing-fuselage f l a p geometq.- A comparison of the aerodynamic characteris- t i c s of t h r e e wing-fuselage f l a p v a r i a t i o n s i s presented i n f i g u r e 8 f o r con- I figuration FW2HV. The pitching-moment c h a r a c t e r i s t i c s were not a f f e c t e d sigriifi- { cant& by the leading-edge radius of the TO0 wing-fuselage f l a p but were a f f e c t e d by t h e increased area and sweep of t h e l a r g e r f l a p , 1 A t low angles W F ' = 750s0°.

I of a t t a c k reduced s t a b i l i t y margins occurred f o r each wing sweep angle and a n o v e r a l l reduction i n uerodynwiic-center v a r i a t i o n between 25' and 7 3 ' wing sweep.

(See f i g . 18.) A t the high angles of a t t a c k pitch-up tendencies were indicated f o r wing sweeps of 2 5 ' and 4>O, and a reduced s t a b i l i t y l e v e l w a s indicated f o r a wing sweep of 7 > O .

( S e e f i g . 8. ) For a more d i r e c t comparison of t h e high- l i f t p i t c h c h a r a c t e r i s t i c s of the two f l a p configurations, the values of' C, f o r WF = 7 5 ° s 0 0 were adjusted t o give t h e same l e v e l of l o w - l i f t s t a b i l i t y as f l a p W F = 70°s00 had a t A = 25'. These adjusted values of C, a r e indicatea by the flagged symbols i n figure 8; the pitch-up tendencies, although reduced, a r e s t i l l apparent i n t h e t r a n s f e r r e d data. Most of t h i s pitch-up tendency i r e s u l t s d i r e c t l y from the added l i f t , only a s m a l l amount being a t t r i b u t e d t o reduced t a i l effectiveness. (See f i g . 20.)

C A change i n t h e wing-fuselage f l a p a l s o r e s u l t e d i n a small change i n La w a s and (L/DImax as i s shown i n f i g u r e 18. The l i f t - c u r v e slope

ckt

increased and values of were reduced f o r t h e range of sweep angles (I,/D),= investigated.

.

Wing-fusel%e f l a p deflection.- The d a t a of f i g w e 11 show t h a t d e f l e c t i o n of the wing-fuselage f l a p was an e f f e c t i v e device f o r increasing model s t a b i l i t y at high angles of a t t a c k .

It i s noted from figures 11 and 1 4 t h a t 290 d e f l e c t i o n

(WF = 70°s290) reduces pitch-up tendencies and gives

of the wing-fuselage f l a p variations Of C, with CL which a r e f a i r l y l i n e a r over t h e design sweep range A = 2 3 O t o 7 5 ' . Increments i n pitching-moment and lift c o e f f i c i e n t s due t o t h e deflection of the f l a p ( f i g . 21) show t h a t t h e s t a b i l i z i n g tendency is due i n p a r t t o the l o s s of l i f t over the flapped area and i n p a r t t o t h e a d d i t i v e lift of the horizontal t a i l . A t high angles of a t t a c k t h e l o s s i n lift i s generally less with t h e horizontal t a i l on than with the horizontal t a i l off; a t low angles of a t t a c k p o s i t i v e increments i n l i f t were obtained with the t a i l on.

The low-lift c h a r a c t e r i s t i c s of t h e configuration a r e summarized i n f i g -

ure 19. These data generally show reductions i n aCm/&, and C . Increases

LU a r e indicated f o r some gxffigG&tions where the f l a p was deflected.

i n (L/D),, T a i l dihedral.- A comparison of t h e longitudinal c h a r a c t e r i s t i c s of the r h = -28.3' with those models with Fh = 0 ' i s shown i n f i g - model with .ures 9 and 13. No s i g n i f i c a n t change i n the p i t c h c h a r a c t e r i s t i c s i s noted f o r t h e negative t a i l dihedral.

# L a t e r a l S t a b i l i t y Directional s t a b i l i t y parameters a r e presented i n f i g u r e 22 f o r model con- f i g u r a t i o n FW$V and WF = 70°s00. Increasing t h e wing sweep angle delayed t h e onset of d i r e c t i o n a l i n s t a b i l i t y from a = 1 8 O at A = l 3 . 5 O t o a = 2 3 . 5 O a t A = 75'. This e f f e c t i s due primarily t o an increase i n v e r t i c a l - t a i l con- t r i b u t i o n a r i s i n g from a favorable e f f e c t of sweep on the wing-induced sidewash at t h e v e r t i c a l t a i l as s h a m i n f i g u r e 24. Wing sweep a l s o has a favorable e f f e c t on the s t a b i l i t y of the wing-body combination as shown by the increasing values of C , f o r increasing values of wing sweep i n f i g u r e 23.

P

The e f f e c t i v e dihedral parameter C given i n figures 22 and 23 varied a manner generally s i m i l a r t o t h a t f o r other variable-sweep configurations i n of t h i s type.

For wings of low sweep, t h e e f f e c t i v e dihedral increased up t o wing stall angle and then decreased; t h i s decrease was followed by increasing dihedral e f f e c t as the angle of a t t a c k w a s increased f u r t h e r . The wing of high

sweep A = 75'

gave increasing dihedral e f f e c t up t o an angle of a t t a c k of 24'.

SUMMARY O F RESULTS Results of a preliminary low-speed s t a b i l i t y and control investigation on

a model of a supersonic transport a i r p l a n e configuration designated SCAT 14

are summarized as follows: 1. The b a s i c model w a s longitudinally s t a b l e throughout the angle of a t t a c k and sweep range investigated f o r a moment reference located at t h e wing pivot

s t a t i o n ; however, f o r intermediate sweep angles (25' t o 55') a reduction of

s t a b i l i t y occurred at angles of a t t a c k greater than 8'.

Varying the wing sweep

angle from 13.5' t o 75' gave a rearward s h i f t i n the aerodynamic center of about

5.5 percent of t h e fuselage length and reduced t h e untrimmed l i f t - d r a g r a t i o from 14.3 t o 6.6.

2. A change from t h e TO0 wing-fuselage f l a p s t o the 75' f l a p s which increased t h e wing l i f t i n g area adjacent t o the fuselage r e s u l t e d i n pitch-up

for wing sweep angles of 25O and 45O, a reduction in aerodynamic-center variation

with sweep, and lower values of maximum lift-drag ratio.

3 . Deflecting the 70' wing-fuselage flap 2 9 ' gave stabilizing moments at

high angles of attack over the sweep range invest-igated.

Deflection of the flap small also reduced the aerodynamic-center variation'with sweep and provided increases in maximum lift-drag ratio.

4. Varying the wing-sweep angle from 13.5O to 75O delays the onset of

directional instability from angles of attack of 180 to 2 3 . 5 ' .

Langley Research Center, National Aeronautics and Space Administration, I Langley Station, Hmpton, Va., August 12, 1964.

1. Lockwood, Vernard E . ; McKinney, Linwood W. ; and Lamar, John E. : Low- Speed Aerodynamic Characteristics of a Supersonic Transport Model With a NASA T M X-979, 1964. High-Aspect-Ratio Variable-Sweep Warped Wing.

2 . Gillis, Clarence L.; Polhamus, Edward C.; and Gray, Joseph L., Jr.: Charts

for Determining Jet-Boundary Corrections for Complete Models in 7- by

10-Foot Closed Rectangular Wind Tunnels. NACA WR L-123, 1945. (Formerly

NACA ARR L3G3l. )

3. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-

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

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

c e em e m e e e e e e e em e m m o m e e e m m e e m e m e e m m e m e e m . e m e m e e e mea em

TABU I . MODEL DIMENSIONS

Reference:

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.00

Area. s q f t

. Span. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.25

Chord. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.36

Fuselage:

Length. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 9 . 0 0

Base area of engine. sq ft . . . . . . . . . . . . . . . . . . . . . 0.1365

Horizontal tail:

Leading-edge sweep. deg . . . . . . . . . . . . . . . . . . . . . . . 60.0

Trailing-edge sweep. deg . . . . . . . . . . . . . . . . . . . . . . 28.6

Root chord. in . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.70

Tip chord. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.20

span (panel). in . . . . . . . . . . . . . . . . . . . . . . . . . . 8 . 0 0

Span (overall). in . . . . . . . . . . . . . . . . . . . . . . . . . 2 8 . 0 0

Exposed area (total). sq ft . . . . . . . . . . . . . . . . . . . . . 0.970

Vertical tail:

. . . . . . . . . . . . . . . . . . . . . . . 70.0

Leading-edge sweep. deg

Trailing-edge sweep. deg . . . . . . . . . . . . . . . . . . . . . . 42.0

Root chord. in . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.52

Tip chord. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.60

span. in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.68

Ekposed area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . 1.000

WF = 7 0 ~ ~ 0 ~ . WF = 70°r00:

Area. s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.99

WF = 75°s00:

Area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.714

TABLE 11.- WING AREAS AND ASPECT RATIOS 'WITH CONSIDERATION FOR VARIOUS WING-FUSELAGE FLAPS [Wing areas based on extension of leading and t r a i l i n g edges t o plane of symmetry)

I I I

Without WF f l a p

70' W F f l a p I 75' W F f l a p I

b, - i n .

Aspect Aspect Area, Area,

r a t i o sq f t Area, sq f t 1 Aspect r a t i o sq f t r a t i o

81.96 6.898 6.242 13.5 5 053 6.763 7.474 99 232 78.46 8.491 6.704 5.801 6.377 7.369 5 035 7.424 5.201 7.202 73.44 6.760

5.540 35 5.@+5

7.484 66.68 5.420 4.125 4.570 5.697 6.757 4.078 3.114 3.410 58.34 7.590 55 5.796 6 932 48.66 2.102 2.291 7.823 65 6.513 2 9 525 7 179 38.18 1.208 1.313 8.378 75 8 199 1.235 7 707 .

0 0 0 0 0 0 8 0 0 0 0 0 0 0 0 8 0 0 0 0 0 0 0 0 0 0 8 8 0 0 8 8 0 0 0 0 8 0 0 0 8 0 0 0 0 8 0 0 0 h\; Reference dimensions Leading-edge sweep 75.

Area ZOOsqf?

i i

Chord 3L36fc.

Span 3 8 2 5 i n .

I

I l l

I

1 \ L-'J-

-LF-

8 W f /\ Section A-A 33.50 /)ota?ed clochwise 5Z5* I (a) Three-view drawing.

Figure 2 . - Drawing of model tested. A l l linear dimensions are i n inches.

.

7 0 ' WING-FUSELAGE FLAP ORDINATES Section B-B Section C-C I I X, YU I y1 J Y Z 2 in.

in. in. in.

', 0 0 0 v 0 1.000 ,200 -.010 1.000 .600 -.016 ~~ 2.000 ,374

-.030 F 2.000

- .034 1.200 3.000 .520 -.050 3.000 .372 - .060 2.000 4.000 ,660 -.080 4.000 - .088 3.000 5.000 .774 -.110 5.000 .550 -.112 3.460 6.000 ,874 -.136 5.630 .600 - .130 7.000 .970 -.la 8.140 1.060 -.210 Model center line - Y

~ 1 Station 36.14

-i

-

3.35

I - Section A-A

Radius = 0.105- 2.00 2.30

4 . 0 0 I i

Radius = 0.065 -

I

6.88 Section C-C Centroid of a r e a ~ Radius = 0.045 WF = 7OosO0 Total a r e a = 56.00 sq in.

W F = 70°r00

* 18.91 - 7

(b) Details of TO0 wing-fuselage flap. W F = 7OosO0 and W F = 70°r00.

Figure 2. - Cont inued .

0 . m.. . 0.. 0 . 0 . . . 0.. 0 .

0 . 0 . 0 . . . a - - - - - - -

Q . - e - % a ., I .

7 5 ' WING-FUSELAtiE FLAP ORDINATES

I Section A-A Section B-B

~ x, Yu, 81 2 YU, X , in. in.

l!,. i n . in.

t 0 c I l.U(J0 ,070 -.C14 u i - . k ' l I ,142 -.om 2.000 ,212 - ,042 3.000 ,284 - ,026 4.000 ,354 - ,070 3.000 ,426 - a x 4 m o o ,496 - . O Y ~ 7.000 .!I68 -.11% x.ono .ti38 -.126 9.000 ,710 - . ~ n 10.000 :/ec - ,114 11.000 .H>O - . l 6 b 12.000 ,922 -.152 ,392 -.l,jti i.ctin - 210 I

i

.I-

( c ) Details of 75' wing-fuselage flap, WF = 75OsOo.

Figure 2.- Concluded.

I 20 inches Figure 4.- Model fuselage cross sections.

Reference - - l i n e - :

+--- -I

------I

I

I 20 inches

I I 10 inches

I

Reference t- t - l i n e

9 - s t a . 82

Eta. 80 sta. 84 Figure 4.- Concluded.

./ I. 0 c , -./ .9 -. 2 .8 .7 .6 .5 CD .4 .3 .2 . / 20 0 / 6 I 2 deg 8 *, - 4 -8 - 4 -.2 0 .2 I .6 .8 l.0 1 . 2 14 1 6 1.8 CL (a) w1m.

= Figu r e 5.- Ef 70°r00.

6h = 0 ' ; fect of w i n g sweep on aerodynamic characteristics in pitch.

I - ....... ...............

. 0 . 0 . 0 .

.........

. . . . . . . .

..........

(a) Concluded.

Figure 5.- Continued.

.e 0 .

0 .

e .

0 .

- -

.f c , -.f -.2 -.3 . I f6 I 2 a,deg 8 - 4 -8 - 4 -.2 0 .2 4 . 6 .8 f.0 f.2 f,4 /.6 f.8 CL (b) m1V.

Figure 5.- Continued.

L 'D (b) Concluied.

Figure 5.- Concluded.

-.4 -.2 0 .2 4 .6 .8 1.0 1 . 2 f.4 f.6 1.8 CL (a) A = 1 3 . 5 ' .

Figure 6.- Effect of horizontal tail on aerodynamic characteristics in pitch- FWlAV; WF = 7OorO0.

a (b) A = 25'.

~ i g u r e 6.- Continued.

0 . 0.. * 0.. . 0 .

0 . . e 0 .

.

.. 0 . . 0 . . .

.. 0 . 0 . . . .

0 . 0 . 0 . . .. 0 . . . 0.. 0 . ... 0 .

.3 2 4 f6 f 2 o,deg 8 -4 -8 ( c ) A = 55'.

Figure 6.- Continued.

I. 0 .9 .8 .7 .6 .5 . 4 .3 .2 .I a '4 (d) A = 45'.

Figure 6 . - Continued.

e . e. e.. e . e e.. e . e e. e *:w: *e: e . . e * e . . e e . .

e . e . e . . e e . .

e. e.. e e e e. e. e e .e. e. e.. e.

Cm C

b

,deg ( e ) A = 55'.

Figure 6.- Continued.

a. a. . a * a - 0 . a.

a . * . a .

e . . *.. . a . e * a . a .

.* de. a. ..a . - a . .a ..a a.

.f f.0 c , -.f .9 -. 2 .8 .3 . 7 . 6 .5 cD .4 .3 .2 . f f6 f2 a,deq 8 -4 - R

-

-4 -.2 0 .2 4 . 6 .8 f.0 f.2 f.4 f.6 f.8 CL (f) A = 6 5 O .

Figure 6.- Continued.

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

0 . .. .. ... 0 . 0 . ... .. ... . * . - * - ;-: .. ... . . . .. 0 . 0 . ... .. ..

./ I. 0 c , -.I .9 -. 2 .8 .7 .6 .5 .4 .3 28 .2 24 . / /6 I 2 0,deg 8 -4 -8 - 4 -2 0 .2 I . 6 .8 l.0 /.2 1.4 /.6 1.8 CL (g) A = 75'.

Figure 6.- Concluded.

I. 0 .9 .8 .7 .6 .5 .4 .3 .2 .I ff -4 -.2 0 .2 4 . 6 .8 1.0 1.2 l.4 l.6 1.8 CL (a) A = 1 3 . 5 ' .

Figure 7.- Effect of tail on aerodynamic characteristics in pitch. F W P ; WF = 70°r00.

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

........ .. 0 . 0 . .* 1 : - io; . . . . . . .

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

- 4 -.2 0 .2 I . 6 .8 1.0 1.2 l.4 l.6 1.8 C L (b) A = 2 5 ' .

Figure 7.- Continued.

.f f.0 r) -.f .9 - . 2 .8 .7 . 6 .5 .4 .3 .2 . f f6 f 2 a '4 8 -4 - R

-

-4 -.2 0 .2 4 .6 .8 f.0 f.2 /.4 L 6 /.8 CL ( c ) A = 55'.

Figure 7.- Continued.

............ 0 .

.. 0 . 0 . . 0 .

......... 0 .

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

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

a,de ./ I. 0 -.I .9 -. 2 .8 -.3 .7 .6 .5 cD . 4 .3 .2 .I 20 0 / 2 ' 8 -4 -8 ( e ) A = 55'.

Figure 7.- Continued.

0 . 0.. . 0.. . 0 . 0 . . . 1 ... 1 .

.I -.I -.2 1 6 a ' 8 -4 -8 -9 -.2 0 .2 4 . 6 .8 1.0 1.2 /.4 f.6 1 s CL (f) A = 65O.

Figure 7.- Continued.

m a am ma. e ama ma a m a m m a a b a a a m m m a m e a - e e m m a m m m m a m e m a m e a am m e a em em a a a m me .f I. 0 -. f .9 -. 2 .8 :3 .7 .6 .5 .4 .3 .2 24 . I f6 f2 II, de '4 8 -4 -8

-

-4 -.2 0 .2 4 . 6 . t ? 1.0 1 . 2 l.4 l.6 l.8 CL (9) A = 75'.

Figure 7 . - Concluded.

? 39

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

. . . . . . .

..........

- - . . . . . . . . . . . . . . . .

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

I. . . .

w .I , -.I Cm -. 2 / 2 rr,de '4 8 - 4 -4 -.2 0 .2 I . 6 .8 1.0 1.2 /.4 /.6 1 . 8 CL ( a ) A = 2 5 O .

Figure 8. - Effect of wing-fuselage f l a p geometry on aerodynamic c h a r a c t e r i s t i c s i n pit.ch.

N2W; 6h = 0'. Flagged symbols i n d i c a t e moment reference at 0.Ojcrer ahcad of wing pivot station.

om 0 . 0 eo. om o m o m o m o m . m o o o m

i ' : m a m - m a

.I c , -.I -. 2 - 3 I6 de *, -4 - 8 CL (b) A = 45'.

Figure 8 . - Continued.

0 . 0 . 0.. 0 . . 0.. 0 . . 0 : . - : ... * . : . . 0 . . 0 . . 0 . .

0 . 0 . 0 . .

0 . 0.. . . . 0 . 0 . . . ... 0 . 0.. 0 .

I. 0 c, -.I .9 .3 .7 .6 .5 .4 .3 2 4 . I I 2 u,deg 8 -4 -8 - 4 -.2 0 .2 4 .6 .8 1.0 1 . 2 f.4 f.6 1.8 ' L (c) A = 75'.

Figure 8.- Concluded.

I6 f 2 a,de 9 8 -4 (a) A = 2 5 O .

Figure 9.- Effect of horizontal tail on aerodynamic characteristics in pitch.

Fw2Rv; W = 7’j0so0.

- 43

.I

e, -.f

-.2 -3 1 2 a 1,deg 8 -4 -8

- 4 -.2 0 .2 4 . 6 .8 1.0 1.2 /.4 /.6 /.a

CL (b) A = 4 5 ' .

Figure 9.- Continued.

.I -_I Cm -.2 -.3 I6 1 2 -4 -8 -4 -.2 0 .2 4 . 6 .8 1.0 1.2 / 4 /.6 1.8 CL ( e ) A = 7 5 O .

Figure 9. - Concluded.

./ 1 . 0 c , -.I -.2 B -.3 .6 .5 .4 .3 .2 . / /6 / 2 cl,deg 8 -4 -8 - 4 -.2 0 .2 4 . 6 .8 l.0 /.2 /.4 /.6 1 . 8 CL (a) m2m.

Figure 10.- Effect of wing sweep on aerodynamic characteristics in pitch. 4, = 0'; WF = 70°s00.

L -.6 -.4 -.2 0 .2 4 . 6 .8 LO 1 2 L4 l6 CL (a) Concluded.

Figure 10.- Continued.

.I 1.0 -.I .9 -.2 -.3 .5 .4 28 2 . I 20 0 1 8 -4 .

" -.6 - 4 1 2 0 .2 .4 . 6 .8 LO 1 . 2 /14 1.6 CL ( b ) Concluded.

Figure 10.- Concluded.

u,deg 8 - 4 -8 - 4 -2 0 .2 4 .6 .8 1.0 1.2 f.4 f.6 1.8 CL (a) A = 25'.

Figure 11.- Effect of wing-fuselage flap deflection (WF = 70's) on aerodynamic characteristics in pitch. PW..HV; 6h = Oo.

.2 .I Cm -_ f -. 2 2G I6 a r, de '4 8 -4 -8 .2 .I I. 0 Cm -.I .9 .8 .7 .6 .5 cD .4 .3 .I I 6 I 2 a,deg 8 - 4 -8 - 4 -.2 0 .2 4 .6 .8 f.0 1.2 l.4 l.6 /.8 CL ( c ) A = 75'.

Figure 11. - Concluded.

0 . 0.. 0 . . 0.. . 0.. 0 .

0 . . . 0 . .. 0 .

. . a . . . 0 . . 0 . . . a

... . . 0.0 . 0 . . 0 . 0 .

0 . . a . 0 . ... . . 0. 0 . . . a 0.. 0 .

./ 1.0 -. / .9 -.2 .8 -.3 .7 .6 .5 .4 .3 .2 24 . / I 6 / 2 r 8 -4 - 0 -4 -.2 0 .2 4 . 6 .8 l.0 1.2 /.4 /.6 1.8 CL (a) A = 25'.

Figure 12.- Effect of horizontal tail on aerodynamic characteristics in pitch.

FW2HV; W F = 7OosO0.

e 0.. e.

e.. e 0.0 e e.

0 .

* e . .

e . . e 0 . e e e . . .

0 . e . . e . . e e . .

e . e . .

. e .e e.. e e e. e. e e.. e. e.. .e .I -_ f f6 I 2 7 8 -4 -8 - I -.2 0 .2 4 . 6 .8 f.0 f.2 /.4 /.6 f.8 CL (b) A = 45'.

Figure 12. - Continued.

.I I. 0 -.I .9 -. 2 .8 .7 .6 .5 cD . 4 .3 .2 24 . I I6 I 2 7 8 -4 -8 - -4 -.2 0 .2 4 . 6 .8 1.0 1 . 2 /.4 L 6 1 . 8 CL ( c ) A = 75'.

Figure 12. - Concluded.

.9 .5 CD .3 .2 ./ - 4 -.Z 0 .Z 4 .6 .8 L O 1 2 /.4 L 6 L 8 CL (a) A = 25'.

= 70°S150.

Figure 13.- Effect of horizontal tail on aerodynamic characteristics in pitch.

FW Hv;

56 I

./

e, -./

/ 6 / 2 a,de 1 8 -4 -8 -4 -.2 0 .2 4 .6 .8 1 . 0 1.2 L 4 L 6 1 . 8 CL (b) A = 45'.

Figure 15.- Continued.

.I I. 0 c, -.f -. 2 .7 .6 5 cD .4 .3 28 .2 24 .I 20 0 I6 I 2 deg 8 -4 - R ( c ) A = 75'.

Figure 13.- Concluded.

.2 .I f.0 ern -. f .9 72 .8 .3 .7 .6 .5 cD .4 .3 .2 .I I6 I 2 a,deg 8 -4 -8 -4 -.2 0 .2 I . 6 .6 f.0 1.2 /.4 /6 1.8 CL ( a ) A = 1 3 . 5 ' .

Figure 14.- Effect of horizontal t a i l on aerodynamic characteristics in pitch. FW2HV; W F = 70°s290 . e . e . ..

. e e . . . . . e . e . 0 . 0

.. .. 0 . . . . a 0 . . 0 . 0

.. . .

. e ... . . e .

...

e . e . . e .

.. ..

.. . e . . . e

.2 1.0 Cm -. / -. 2 .7 .6 5 cD .4 .3 .2 2 4 . / /6 / 2 deg 8 -4 -8 - 4 -.2 0 .2 4 . 6 .8 l.0 /.2 /.4 /.6 1 . 8 ' L (b) A = 25'.

Figure 14.- Continued.

.2 .I crn -_I -.2 . 3 I6 I 2 cr,deg 8 -4 -8 ( c ) A = 35'.

Figure 14. - Continued.

.2 .I 1.0 Crn -. / -.2 8 .7 .4 .3 . / 1 2 a,deg 8 -4 -8 -4 -.2 0 .2 4 .6 .8 l.0 /.2 /.4 /.6 f.8 CL (d) A = 45'.

Figure 14.- Continued.

62 -

.I c D , dc ( e ) A = 55'.

Figure 14. - Continued.

1.0 .9 crn .8 .7 .6 .5 cD .4 .3 . I a,dc - 4 -.2 0 .2 4 .6 .8 1.0 1.2 l.4 l.6 1.8 CL (f) A = 6 5 O .

Figure 14.- Continued.

... ..... : - i:. . :.. : . *

.... .....

... .r . . . .

.............. a. ........

.2 ./ 1.0 Cm -. / .9 .8 .7 .6 .5 .4 .3 .2 . / /6 / 2 deg 8 -4 -8 (g) A = 75'.

Figure 14.- Concluded.

b

%

h

“0 lo 20 30 40 50 60 70 80

A , deg & r n Figure 16.- Effects of h o r i z o n t d t a i l on l i f t parameters -, Cw and ( L / D ) - .

a c , y v ; UT = 700ra0.

0 IO 20 30 40 50 60 70 80

A tdeq

Figure 17.- Effect of wing twist on parameters -, ac, CLa, and ( L / D ) , -

acL m; % = oo; WF = 70°r00.

0 1 0 20 30 40 50 60 70 80

A ,w

ac, Figure 18.- Effect of wing-fuselage flap geometry on parameters CW and (L/D)-.

-, &L F w p ; €jh = 00.

0 . e.. . 0.. * I . - ..e *: : 0 . : 0 . 0 . 0 .

. . 0 . . e. .

0 . .

0 . . * e . .

0 . 0.. . e . 0 . e. . e.. e. 0 . . ..

"0 /O 20 30 40 50 60 7 0 80

A,deg ac,

-

Figure 19.- Effect af wing-fuselage flap deflection on parameters & L ' Fw2W, % = 0 ' ; WF = 7 0 ' s .

0 0 e88 0 0 8 0 e 8 0 0 0 0 0 0 0 8 8 0 8 0 0 0 0 8 0 0 0 0 8 0 0 0 0 - 0 0 . 0 0 0 0 0 8 0 - 8 8 0 8 8 0 0 0 0 8 0 0 0 0 0 0 0 8 0 8 0 0 0 8

.04

A cm

-. 04

- .08

-. /2

-.I6

-8 - 4 0 4 8 /2 /6 20 24 28

0 , deg

Figure 2 0 . - Effect of wing sweep and wing-fuselage flap geometry on the increment in pitching-moment coefficient due to addition of the horizontal tail hc;, for configuration FW2V.

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

. . . . . . .

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

aJ +

9-

I I

i d

Q

I I' 0 0 e . . 0 0 - 0 0 0.0 0 0.0 0 .

0 0 0 0 . 0 . 0 .

0 0 0 . 0 0 . *.. 0 .

0 0 0 0 . 0 . 0 0 e . 0 0 0 0 0 0.0 .. 0 0 0 0 . 0 0 e 0.0 0 .

Q , deq Figure 22.- EEfect of wing sweep on lateral stability characteristics in pitch.

= -ioo; W-F = 70°s00.

FW~HV; - 4 0 4 8 I2 I6 2D 24 28 a , d w (a) A = 1 3 . 5 ' .

Figure 23.- Effect of vertical tail on lateral stability characteristics in pitch.

FW H; ah = -ioo; WF = 700~00.

.V , L -4 0 4 8 I2 I6 20 24 28 o , d w ( c ) A = 45'.

Figure 23.- Continued.

.

-4 0 4 8 I2 I6 20 24 28 a ,deg (a) A = 65O.

Figure 23. - Continued.

( e ) A = 75'.

Figure 23.- Concluded.

.ooz

.oo/

A

-.oo/

.om

.LWZ A -.002 -.004 .004 .002 d c -.002 -.004 -.008 0 4 8 /2 I6 20 24 28 Q .deg Figure 24.- E f f e c t of wing sweep on v e r t i c a l - t a i l contribution. FW2H; fjh = - 1 0 ' ; WF = 70°r00.

“Tbe aeronutical and space activities of the United States sball be

condwted so as to contribute . . . to the expansion of human knowl-

edge of phenomena in the atmosphere and space. The Administration Jbdj provide for the widest practicable and appropriate dissemination

of information concerning its actiuities and the resrtlts thereof .”

-NATIONAL AERONAUTICS AND SPACE ACT OF 1958 , , \

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS

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 considered to merit NASA distribution in English.

NASA activities TECHNICAL REPRINTS: Information derived from and initially published in the form of journal articles.

or of value to SPECIAL PUBLICATIONS Information derived from NASA activities but not necessarily reporting the results of individual NASA-programmed scientific efforts. Publications include conference 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. PO546 i i