Document
TECHNICAL MEMORANDUM
x-127
CONTROL OF A RESEARCH-TYPE AlRPLANE CONFIGURATION
By David E . Fetterman, Jr.
Langley Re search Center Fie1 Lanql -ey GPO PRICE OTS PRICE( Hard copy L
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NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON Cctober 1959
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0 < i ’ i A S A CI; O d T M X O R A D N U M B E R 1 IC-TEGORYI . l A .
b NATIONAL AERONAUTICS AND SPACE ADMINISTRATION TECHNICAL MEMORANDUM x-127 EFFECTS OF SIMULATED ROCKET-JET EXHAUST ON STABILITY AND CONFIGURATION CONTROL OF A RESEARCH-TYPE AIRPLANE AT A MACH NUMBER OF 6.86* L By David E. Fetterman, Jr.
SUMMARY
38/35
An investigation has been undertaken in the Langley 11-inch hypersonic tunnel at a free-stream Mach number of 6.86 to determine the jet-interference effects at high jet-static-pressure ratios on the stability and control of a research-type airplane configuration.
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Compressed-air tests with a jet exhausting from the base of t e fuse-
lage were conducted over a Reynolds number range of 0.57 X 10 2 to
I
3.95 x lo6, based on fuselage length, and over a jet-static-pressure-
ratio range of 0 to 1460. The results of these tests indicated that the operation of the jet induced a sizeable separated-flow region over the vertical- and horizontal-tail surfaces which could be approximately duplicated at low angles of attack b y use of metal jet-boundary simu- lators. The results of force tests, during which these metal jet- boundary simulators were used, indicated that this separated-flow region caused a large reduction in the longitudinal stability and con- trol and a smaller reduction in the lateral and directional stability and control. By extending the divergent section of the nozzle and thus reducing the jet-static-pressure ratio, these losses were diminished.
INTRODUCTION Previous investigations have shown that a jet exhausting from the base of the fuselage may alter the base drag, the afterbody pressure distribution, and also the aerodynamic characteristics of the test
configurations. (For example, see refs. 1 to 5.) High-performance
and high-altitude aircraft will be subjected to these effects during the lower altitude portions of their trajectories. As the low ambient
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Title, Unclassified.
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pressures associated with very high altitude are approached, however, jet-static-pressure ratios greatly exceeding those considered in pre- vious investigations will be encountered. In order to determine the jet-interference effects which may occur at these high jet-static- pressure ratios and high Mach numbers, the investigation described in this paper was undertaken in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86.
SYMBOLS L jet-exit plane area 1 *j jet throat area At b wing span Draq drag coefficient, CD qs Lift
lift coefficient, -
CL @ Rolling moment rolling-moment coefficient, qSb pitching-moment coefficient, moment reference 0.20E, Cm Pitching moment qSc' Yawing moment yawing-moment coefficient, Cn qSb
-
C wing mean aerodynamic chord base diameter of fuselage jet-exit diameter fuselage length jet-exit Mach number free-stream Mach number jet static pressure jet total pressure ....... ...............
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I free-stream static pressure P , dynamic pressure R Reynolds number, based on fuselage length axial distance from base of fuselage xb Z vertical distance from bottom of fuselage L Z lower vertical-tail span from bottom of fuselage measured at t tail trailing edge-fuselage intersections (see fig.
8) U angle of attack angle of sideslip
e
clx distance from fuselage base to jet exit horizontal-tail deflection, positive to produce positive ' h CL differential horizontal deflection, positive to produce &h positive Cl vertical-tail deflection, positive to produce negative ' V Cn jet specific-heat ratio Y j initial jet-boundary slope ' j The following stability parameters are referred to the body axis system: rate of change of yawing-moment coefficient with sideslip
CnP angle, acn/ap
rate of change of rolling-moment coefficient with sideslip
czP angle, &z/ap
rate of change of yawin -moment coefficlent with vertical- cn8V tail deflection, aCn 7 3 8 , rate of change of rolling-moment coefficient with vertical- % V tail deflection,
??€ 1 /asv
............... . . 0.. 0 .
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I rate of change of yawing-moment coefficient with differential ‘n6 h horizontal-tail deflection, a C n / & , 1 rate of change of rolling-moment coefficient with differential
‘Z6h ‘
horizontal-tail deflection, &,/ash 1 APPARATUS Wind Tunnel L The tests were conducted in the Langley 11-inch hypersonic tunnel is equipped with a single-step, two-dimensional nozzle constructed which of Invar. Tunnel operation is of the intermittent type and a running time of about 80 seconds is possible. The nozzle was designed by the of characteristics with a correction made for boundary layer and method operates at an average Mach number of 6.86. The variation in Mach num- 10 seconds of running time is about 1 percent.
ber after the first During these tests, the stagnation temperature was maintained at about 675O F by means of a variable-frequency, electrical heater equipped Y with Nichrome tube resistance elements. This high temperature is neces- In order to elimi- sary to avoid air liquefaction in the test section.
nate the effects of water condensation, the absolute humidity of the air was kept less than 1.87 x 10-5 pounds of water vapor per pound of air for all tests.
Further details of the Langley 11-inch tunnel facility may be found in reference 6.
%lance and Force Model Support Force and moment measurements were made through the use of a six- component-strain-gageforce balance, the design of which allows four components to be located internally in the model. The other two
components - axial force and rolling moment - are mounted externally
at the rear of the balance and are shielded from the air flow during the test runs. The model and balance were mounted in the test section on a movable support strut which could be rotated through an angle-of- attack range. During each test, the period of essentially constant Mach number flow was long enough to permit testing through the angle- of-attack range. Angles of sideslip were obtained by offsetting the model and balance support to the desired sideslip angle prior to each run. Thus, the data were obtained at an essentially constant sideslip angle over an angle-of-attack range.
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Models Force-test model.- A three-view drawing of t h e m o d e l used f o r t h e f o r c e test of t h i s i n v e s t i g a t i o n i s presented i n f i g u r e 1, and the geometric c h a r a c t e r i s t i c s are given i n t a b l e I. The model c o n s i s t e d of a s l i g h t l y b o a t t a i l e d , ogive-cylinder combination w i t h s i d e f a i r i n g s , a t r a p e z o i d a l wing, swept h o r i z o n t a l - t a i l s u r f a c e s , and swept lower and upper v e r t i c a l - t a i l surfaces.
L The h o r i z o n t a l - t a i l panels a r e d e f l e c t e d together f o r p i t c h con- t r o l and d i f f e r e n t i a l l y f o r roll c o n t r o l . The inboard p o r t i o n s of the v e r t i c a l t a i l are f i x e d and support the speed brakes, d e t a i l s of which a r e shown i n f i g u r e 2. The outboard p o r t i o n s of t h e v e r t i c a l t a i l are d e f l e c t e d f o r d i r e c t i o n a l c o n t r o l .
Compressed-air-test model.- A sketch showing the c o n s t r u c t i o n d e t a i l s and p e r t i n e n t dimensions of the model used i n t h e compressed- a i r t e s t t o o b t a i n t h e e x t e n t of the j e t - i n t e r f e r e n c e flow f i e l d and t h e jet-boundary shapes is presented i n figure 3 . The model was made of s t a i n l e s s steel and the e x t e r i o r s u r f a c e s were smooth. The f u s e - lage, which is i d e n t i c a l t o that of t h e f o r c e - t e s t model previously described, i s supported on t h e r i g h t s i d e by an 8.75-percent-thick sup- r' p o r t s t r u t from the t u n n e l s i d e w a l l . This s t r u t contained copper air- supply tubes and the jet-stagnation-pressure tube. The s t r u t w a s extended from the l e f t s i d e of t h e fuselage far enough t o permit a symmetrical f l o w f i e l d on each s i d e of the f u s e l a g e . Dry a i r a t 2000 l b / s q i n . from a storage tank w a s used as t h e jet exhaust gas and w a s piped through a t h r o t t l i n g valve i n t o the air-supply tubes a t approximately atmospheric temperature. The j e t s t a g n a t i o n p r e s s u r e s were measured i n a small o f f s e t chamber ahead of t h e j e t - s t a g n a t i o n - By employing t h i s arrangement p r e s s u r e chamber as shown i n f i g u r e 3 .
l i t t l e , if any, e f f e c t of the flow v e l o c i t y i n t h e j e t - s t a g n a t i o n - p r e s s u r e chamber was obtained i n the recorded s t a g n a t i o n p r e s s u r e s .
Nozzles.- The majority of t h e compressed-air t e s t s were made by u s i n g t h e nozzle A c o n f i g u r a t i o n shown i n f i g u r e 4. This nozzle i s a supersonic, convergent-divergent, conical nozzle having a semidivergence angle of 20'. A l i m i t e d number of tests were a l s o made w i t h nozzles B and C , a l s o shown i n f i g u r e 4, which were e s s e n t i a l l y i d e n t i c a l t o t h e s e nozzles were nozzle A except t h a t t h e divergent s e c t i o n s of extended so that t h e i r e x i t planes were a t d i f f e r e n t d i s t a n c e s from t h e base of the f u s e l a g e . These nozzles were used t o determine the e f f e c t s of nozzle extensions. A l l t h r e e nozzles were designed on the basis of i n a l t e r i n g t h e design of t h e a r e a r a t i o only f o r However, y j = 1.25.
nozzles f o r use w i t h a i r ( y j = 1.4) so t h a t t h e c o r r e c t values of Mj, dj/db, and &/db were maintained, t h e r e s u l t i n g t h r o a t diameters of c I the nozzles were not constant. Pertinent information concerning these
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nozzles is given in the following table: B
3 - 93 .785 .393 10.07
L C 4.21 * 927 *592 12.90 I+ A typical installation of nozzle A in the compressed-air-test model is shown in figure 3.
Nozzle A was calibrated by obtaining total-head-pressure measure- ments across the nozzle exit plane. The average Mach number obtained from the measured ratio of total head to stagnation pressures varied Nozzles B .
not more t,han fO.O1 from the average design jet Mach number.
and C could not be calibrated with atmospheric jet-exit pressure because sufficiently high jet-supply pressure was not available to start the nozzles completely. I TESTS A l l tests were conducted at a free-stream Mach number of 6.86. An off axis, single-pass, two-mirror, schlieren system utilizing a mercury- vapor light source was used during all tests. Schlieren photographs were recorded on standard panchromatic film exposed for approximately 11150 sec.
Compressed-Air Tests The compressed-air tests with nozzle A in place were made at tunnel stagnation pressures of 5, 10, 17.5, and 34 atmospheres. These stagna- tion pressures in combination with the free-stream Mach number of 6.86 and stagnation temperature of 6 7 5 ' resulted in Reynolds numbers, based on fuselage length, of 0.57 x l o 6 , 1.20 x 10 6 , 2.05 X lo6, and 3.95 X 10 6 , respectively. Because of the limited capacity of the nozzle high- pressure air supply and the accompanying losses in the air-supply piping, the maximum jet-sta1.ic-pressureratios obtainable varied with different test Reynolds numbers. These maximum jet-static-pressure ratios and also ....... ...............
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t h e test a n g l e s of a t t a c k for t h e various nozzle and Reynolds number combinations are given i n t h e following t a b l e : Nozzle R f o r y j = 1.4 L A 0.57 X 10 6 1460
I 1.20
0 , +2, +4
I I
L B
2.05 x lo6 170
C
2.05 X lo6 13 9
Six-component f o r c e and moment d a t a a t Reynolds numbers of 1.20 X 10 6 and 2.03 X 10 6 were obtained with a l l c o n t r o l s undeflected f o r an angle-of-attack range of +bo and s i d e s l i p a n g l e s of 0 ' and 4 ' .
A t t h e model was a l s o t e s t e d with t h e v e r t i c a l t a i l s d e f l e c t e d p = 0 ' - 5 O and a l s o a d i f f e r e n t i a l h o r i z o n t a l - + , a i l d e f l e c t i o n of 1 0 ' .
The a n g l e s of a t t a c k were s e t using a lens prism imbedded i n t h e model sur- f a c e t o r e f l e c t and focus a spot from a l i g h t source o n t o a p r e v i o u s l y c a l i b r a t e d screen. By using t h i s method t h e t r u e a n g l e s of a t t a c k were of balance d e f l e c t i o n under load.
obtained d i r e c t l y , i r r e s p e c t i v e P r e c i s i o n of Data The probable u n c e r t a i n t i e s in the f o r c e and moment c o e f f i c i e n t s due t o balance r e p e a t a b i l i t y have been estimated and a r e presented as follows:
CL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k O . 0 0 5
CD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . kO.001
C,. . . . . . . . . . . . . . . . . . . . . . . . . . . . . +0.003
C , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +O.OOO'j
C z kO.0003 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The accuracy of a and p w a s within t h e l i m i t s of f 0 . 1 0 ' .
JET SIMULATION TECHNIQUE Simulation of Initial Jet-Boundary Slope A typical variation of jet-static-pressure ratios pj/pm with altitude for the case of yj = 1.25, which is intended to simulate hot gases being expelled from the nozzles, is presented in figure 5 for the three nozzles under consideration. In calculating these curves a jet- stagnation-chamber pressure of 600 lb/sq in. was assumed. Indicated in L
figure 5 at an altitude of about 158,000 feet are the jet-static-pressure
ratios which were simulated during the force tests of this investigation.
For nozzle A , pj/pm = 170; and for nozzle C, pj/pm = 420; for nozzle B, Pj/Pm = 105.
Since air (yj = 1.4) instead of a hot gas (yj = 1.25) was used as the exhaust medium, it was necessary to test at equivalent air-jet static-pressure ratios so that the initial jet-boundary slope could be reasonably duplicated. The effect of specific-heat ratio of the initial boundary slope of the jet from nozzle A is shown in figure 6 for specific- * heat ratios of 1.25 and 1.40. In calculating these curves, the flow over the afterbody was assumed to be attached and two-dimensional-oblique- shock and expansion relations were used in the manner suggested in w references 7 and 8. A s seen from figure 6, increasing the specific-heat ratio from 1.25 to 1.40 causes a considerable reduction in the initial jet-boundary slope; therefore, in order to duplicate the jet-boundary slope which would occur under hot-jet gas conditions, a higher equivalent air-jet static-pressure ratio is required. For example, in order to duplicate the jet-boundary slope at a hot gas-jet static-pressure ratio of 420, it was necessary to use a n equivalent air-jet static-pressure ratio of 1,200 during the compressed-air tests.
I t should be noted that in computing the jet-static-pressure ratios presented herein, a straightforward two-dimensional approach, based on the perfect gas law, was used to calculate the nozzle exit pressures; however, because of the high jet stagnation pressures encountered at the higher values of p j / p , , some departure from the perfect gas assumption does exist. Furthermore, because of the large area ratios for the extended nozzles, some air liquefaction probably occurred within nozzles B and C. However, since this investigation was intended to determine only the general trends of jet-interference effects at high jet-static- pressure ratios, no attempt was made to correct the jet-static-pressure ratios for these effects.
Simulation of Jet-Interference Flow Field Schlieren photographs of the flow field produced by the compressed- air jet indicated that the pressure rise produced by the jet-exit shock was sufficient to induce a separated-flow region which extended over and forward from the rearward end of the fuselage. This jet-induced separated-flow region may be seen in figure 7(a) in which a schlieren photograph of the flow field produced by the air jet from nozzle A exhausting into a Mach number 6.86 air stream is presented. Because of the orientation of the schlieren knife edge, the separated-flow- L region boundary can be seen directly only over the lower surface of the model; however, the location over the upper surface is indicated by the discontinuity of the jet boundary and the point of origin of the jet- exit shock. It should be noted that the jet boundary is asymmetrical, and the larger separated-flow region occurs over the lower portion of the fuselage. The reason for this behavior will be discussed in a subsequent section.
In view of the rather large area covered by this jet-induced separated-flow region, it was believed that significant jet-exhaust effects would result from the presence of this separated-flow region over a large portion of the fuselage and the conventionally located tail-control surfaces. The jet-simulation technique employed consisted of inducing these separated-flow regions over the rearward end of the force-test model. This was accomplished by mounting metal jet-boundary simulators on the windshield of the force balance just aft of the model.
It was necessary, however, to neglect the jet-boundary asymmetry, since the attainment of the shape and the fabrication of asymmetrical jet- boundary simulators was impractical. Instead, a symmetrical jet boundary was approximated through the use of axisymmetric, metal jet-boundary simulators which were machined to the lower-surface jet-boundary shapes at a = 0 ' obtained from schlieren photographs of the compressed-air tests. When it was necessary, these metal fairings were then modified by a reduction in the length of the jet-boundary simulator to produce an angle approximately the same lower-surface separated-flow regions at of attack of 0 ' as those obtained from the compressed-air tests. In figure 7(b) the separated-flow region produced by one of these jet- boundary simulators is shown. This jet-boundary simulator has been modified to induce the same lower-surface separated-flow region as that induced by the air jet (fig. 7(a)) and is the only one of the three jet- boundary simulators tested which required modification.
A better comparison of the lower-surface separated-flow regions induced by the air jet and the Jet-br-indary simulator is shown in figure 8. In this figure the extent of the separated-flow regions are shown in terms of the parameter z/zt in which z is the height of the separated-flow region and zt is the height of the lower vertical tail.
1 0 A t a = 0 ' it i s seen from t h e figure t h a t t h e jet-boundary simulator V induces a v e r y n e a r l y i d e n t i c a l separated-flow r e g i o n t o t h a t induced by t h e a i r j e t . A t angles of a t t a c k , however, t h e a c t u a l jet-boundary shape changes so t h a t a decrease occurs i n t h e e x t e n t of t h e separated-flow r e g i o n on t h e high-pressure s i d e of t h e c o n f i g u r a t i o n w i t h i n c r e a s i n g angle of a t t a c k . A s mentioned previously, it was i m p r a c t i c a l t o o b t a i n t h e t r u e jet-boundary shape and t o c o n s t r u c t asymmetrical m e t a l j e t - boundary simulators f o r each angle of a t t a c k ; t h e r e f o r e , the zero angle- of -attack jet-boundary simulators were used throughout t h e t e s t angle- o f - a t t a c k range. Comparisons of t h e lower-surface separated-flow r e g i o n s L induced a t 2O and 4 ' a n g l e of a t t a c k by t h e air-jet and jet-boundary simulator a r e a l s o shown i n f i g u r e 8. These comparisons i n d i c a t e t h a t 4 a t angles of a t t a c k , t h e jet-boundary simulator induces a p r o g r e s s i v e l y l a r g e r separated-flow region t h a n does t h e a i r j e t . A r e v e r s a l of t h e s e 2 t r e n d s would be expected t o occur on t h e low-pressure s i d e of t h e con- f i g u r a t i o n . I n view of t h e s e r e s u l t s then, t h e angle-of-attack range 0' of t h e f o r c e t e s t s w a s l i m i t e d t o +4 .
RESULTS AND DISCUSSION Compressed-Air T e s t s Representative s c h l i e r e n photographs which show t h e flow f i e l d a t a = 0 ' produced by t h e a i r j e t a t v a r i o u s values of p p, and Reynolds
jl
number are presented i n f i g u r e 9 . The p i c t u r e s f o r t h e j e t o f f a r e included t o show t h e l o c a t i o n of t h e shock waves and wake from the model- I n t h e photographs f o r p j p, > 0, t h e jet-induced support s t r u t .
I
separated-flow regions are i n d i c a t e d , and t h e s e , again, can be seen d i r e c t l y only over t h e lower s u r f a c e of t h e f u s e l a g e . It should be noted p . p,, t h e j e t boundary i s asymmetrical and t h e t h a t f o r a l l values of J / l a r g e r separated-flow regions occur over t h e lower s u r f a c e of t h e f u s e - lage. Tests with t h e canopy removed ( s e e f i g . 9(b)) i n d i c a t e t h a t t h i s behavior was due t o t h e asymmetrical flow f i e l d about t h e f u s e l a g e caused by t h e presence of t h e canopy.
The e f f e c t s of equivalent air-jet s t a t i c - p r e s s u r e r a t i o and Reynolds number on t h e e x t e n t of t h i s jet-induced separated-flow r e g i o n over t h e lower surface of t h e f u s e l a g e at a = 0" are shown i n f i g u r e 10. The parameter z / z t i s again used t o i n d i c a t e t h e e x t e n t of t h e separated- flow region; however, i n t h i s f i g u r e , z i s t h e h e i g h t of t h e separated- a t t h e base of t h e fuselage.
flow region A t t h e lower Reynolds numbers t h e separated-flow region i n c r e a s e s r a p i d l y w i t h j e t - s t a t i c - p r e s s u r e r a t i o , and f o r a Reynolds number of 0.57 x lo6 and j e t - s t a t i c - p r e s s u r e r a t i o s g r e a t e r t h a n 1,000 t h e separated-flow r e g i o n covers t h e e n t i r e lower v e r t i c a l t a i l . A t a c o n s t a n t j e t - s t @ t i c - p r e s s u r e r a t i o , however, am m a m am am a am. a am. ma a m m m a m m e m a m m m m a m a m m m a m a m a . e m m m m a m a a a - ame m m . a m m ama m a m m m a m am am.
L.
. s the e x t e n t of t h e separated-f low region d e c r e a s e s r a p i d l y w i t h i n c r e a s i n g Reynolds number. The values of z / z t f o r pj/p, = 0 are r e a l l y t h e boundary-layer t h i c k n e s s e s at t h e v a r i o u s Reynolds numbers w i t h t h e j e t o f f .
A Reynolds number of 2.05 x lo6 w a s chosen f o r t h e nozzle B and nozzle C compressed-air tests and a l s o f o r t h e ensuing f o r c e tests. For nozzle A, t h e e q u i v a l e n t a i r - j e t - s t a t i c - p r e s s u r e r a t i o of 1,200 a t a Reynolds number of 2.05 x LO6 corresponds t o t h e hot-gas-jet v a l u e of 420. (See f i g . 6.) Since t h e jet-air supply p r e s s u r e was i n s u f f i c i e n t L t o permit t e s t i n g a t t h i s equivalent j e t - s t a t i c - p r e s s u r e r a t i o and Reynolds number combination, e x t r a p o l a t i o n s of t h e a v a i l a b l e data, i n d i - c a t e d by t h e dashed l i n e s , were made by using t h e lower Reynolds number v a r i a t i o n as a guide t o determine t h e separated-flow c o n d i t i o n s t h a t would A. S i m i l a r e x t r a p o l a t i o n s of t h e available d a t a were e x i s t f o r nozzle a l s o necessary f o r nozzle B and nozzle C. These e x t r a p o l a t e d s e p a r a t e d - flow c o n d i t i o n s f o r t h e v a r i o u s nozzles are i n d i c a t e d by t h e s o l i d sym- b o l s . Since t h e flow from nozzles B and C i s underexpanded t o a l e s s e r degree t h a n t h a t f r Q m nozzle A, t h e e x t e n t of t h e separated-flow r e g i o n s for t h e s e nozzles i s less t h a n for nozzle A.
Since a determination of t h e separated-flow r e g i o n induced by nozzle A could not be obtained experimentally a t t h e r e q u i r e d e q u i v a l e n t j e t - s t a t i c - p r e s s u r e - r a t i o and Reynolds number combination it was neces- s a r y t o use the a v a i l a b l e experimental data a t p p, = 528 and at a j l reduced Reynolds number of 1.20 x lo6, which figure 10 shows c l o s e l y approximated t h e d e s i r e d separated-flow c o n d i t i o n s . The s c h l i e r e n photo- graph from t h i s test shown i n f i g u r e 7(a) w a s used f o r d e f i n i n g t h e j e t - boundary-simulator shape f o r nozzle A used d u r i n g t h e f o r c e tests.
Force Tests The e f f e c t s produced by t h i s simulated- jet-exhaust technique on t h e l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e t e s t c o n f i g u r a t i o n a r e shown i n f i g u r e 1 1 i n which t h e v a r i a t i o n s of CD and a w i t h CL are p r e s e n t e d f o r t h e j e t o f f and w i t h t h e jet-boundary s i m u l a t o r s f o r nozzle A, nozzle B, and nozzle C i n place. These d a t a show t h a t t h e presence of
tl:e j e t exhaust reduced both t h e l i f t - c u r v e s l o p e C b and t h e minimum
drag c o e f f i c i e n t from t h e j e t - o f f values. This r e d u c t i o n i n C k can be a t t r i b u t e d t o the loss i n the l i f t c o n t r i b u t i o n s of t h o s e p o r t i o n s of t h e h o r i z o n t a l - t a i l s u r f a c e s and f u s e l a g e which are submerged w i t h i n t h e low-energy flow of jet-induced separated-flow regions. The l a r g e s t r e d u c t i o n i n C b , about 20 percent, occurred f o r nozzle A, which would c be expected s i n c e as noted i n f i g u r e 10 t h e jet boundary from t h i s nozzle ............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . .
........................
induced t h e l a r g e s t separated-flow region. The r e d u c t i o n i n minimum d r a g c o e f f i c i e n t , also about 20 p e r c e n t , was t h e sme f o r a l l three noz- z l e s t e s t e d and was probably due t o a n i n c r e a s e i n base and a f t e r b o d y p r e s s u r e s over t h o s e p r e s e n t w i t h t h e j e t o f f .
The l o n g i t u d i n a l s t a b i l i t y and c o n t r o l r e s u l t s are shown i n f i g - ure 1 2 i n which t h e v a r i a t i o n i n pitching-moment c o e f f i c i e n t w i t h lift c o e f f i c i e n t i s presented f o r h o r i z o n t a l - t a i l d e f l e c t i o n s of 0 ' and - 2 0 ' .
I n considering first t h e curves f o r with t h e j e t o f f , t h e con- 6h = 0 ' f i g u r a t i o n i s seen t o be s t a t i c a l l y s t a b l e with a s t a t i c margin of about L 13 percent of t h e mean aerodynamic chord. Under t h e i n f l u e n c e of t h e separated-flow r e g i o n from t h e j e t simulator f o r nozzle A, however, t h e i s s t a t i c a l l y u n s t a b l e over a small p o s i t i v e and negative c o n f i g u r a t i o n l i f t - c o e f f i c i e n t range. Because of t h e smaller separated-flow r e g i o n s induced with t h e extended nozzles B and C, less l o s s e s i n l o n g i t u d i n a l s t a b i l i t y occur.
The e f f e c t s of t h e simulated-jet exhaust on t h e l o n g i t u d i n a l con- t r o l power of t h e h o r i z o n t a l t a i l are i n d i c a t e d b y t h e d i f f e r e n c e i n t h e curves f o r 6h = 0 ' and 61., = -20°. With t h e j e t o f f , some l o s s Y i n c o n t r o l power occurs at negative l i f t c o e f f i c i e n t s because of wing- wake impingement on t h e l L o r i z o n t a l t a i l s . T h i s behavior was p r e v i o u s l y observed i n tLe i n v e s t i g a t i o n r e p o r t e d i n r e f e r e n c e 9. The combination of t h e wing wake and jet-induced separated-flow r e g i o n from the j e t simulator f o r nozzle A, lAowever, causes a l a r g e r e d u c t i o n i n c o n t r o l power so t h a t a t negative l i f t c o e f f i c i e n t s , t h e h o r i z o n t a l t a i l becomes almost i n e f f e c t i v e . With t h e extended nozzles B and C, t h e c o n t r o l power a t negative l i f t c o e f f i c i e n t s i s reduced o n l y about 15 p e r c e n t below t h a t w i t h t h e j e t o f f .
A s pointed out p r e v i o u s l y and as i n d i c a t e d i n f i g u r e s 1 1 and 12, t h e t e s t s w i t h t h e nozzle A jet-boundary s i m u l a t o r i n p l a c e were conducted a t a reduced Reynolds number of 1.20 x lo6. I n o r d e r t o determine what e f f e c t tLis r e d u c t i o n i n Reynolds number had on t h e foregoing curves f o r nozzle A, t h e model w i t h t h e j e t o f f w a s a l s o tested a t t h i s reduced Reynolds number. The r e s u l t s of t h i s t e s t are compared with t h e j e t - o f f data at a Reynolds number of 2.05 x LO6 i n f i g u r e 13. T1.ese comparisons i n d i c a t e no Reynolds number e f f e c t on t h e l o n g i t u d i n a l c h a r a c t e r i s t i c s when 61, = Oo. For = -20°, however, t h e lower Reynolds number r e s u l t s sllow a l o s s i n c o n t r o l power; n e v e r t h e l e s s , t h i s l o s s i s not n e a r l y as l a r g e as that i n d i c a t e d by t h e curves f o r nozzle A i n f i g u r e 12. I n view of t h e s e r e s u l t s , then, t h e d e v i a t i o n s from t h e j e t - o f f c h a r a c t e r - i s t i c s i n d i c a t e d by t h e curves f o r nozzle A i n f i g u r e s 1 1 and 12 can reasonably be a t t r i b u t e d t o t h e i n f l u e n c e of t h e j e t -induced separated- flow recion.
Contrary t o t h i s t r e n d , l i t t l e , i f any, changes occurred between t h e j e t - o f f and jet-on conditions i n t h e l i f t and pitching-moment char- a c t e r i s t i c s of t h e c o n f j g u r a t i o n w i t h t h e speed brakes d e f l e c t e d 3 5 O .
shown i n f i g u r e 14. The minimum drag c o e f f i c i e n t under These r e s u l t s are t h e j e t - o n conditions, however, was again reduced about 20 p e r c e n t below t h e j e t - o f f value. The jet-on r e s u l t s are presented only f o r nozzle A s i n c e almost i d e n t i c a l r e s u l t s were obtained w i t h nozzles B and C. Since t h e c h a r a c t e r i s t i c s of speed brakes of t h e t y p e used on t h e c o n f i g u r a t i o n are affected t o a l a r g e e x t e n t b y Reynolds number, both t h e j e t - o f f and jet-on data of t h i s f i g u r e are presented f o r a Reynolds number of L 1.20 x 106.
For t h e d i r e c t i o n a l s t a b i l i t y and c o n t r o l r e s u l t s p r e s e n t e d i n f i g - due t o t h e u r e 15, l o s s e s are seen t o occur i n both Cn and C %v P By extending t h e e f f e c t s of t h e simulated-jet exhaust from nozzle A .
nozzle d i v e r g e n t s e c t i o n (nozzles B and C ) t h e s e l o s s e s are diminished so t h a t no change i n and o n l y about a 6-percent r e d u c t i o n i n %v CnP occurs.
The lateral s t a b i l i t y and c o n t r o l r e s u l t s are shown i n figure 16 which a l s o i n d i c a t e s an unfavorable c o n t r i b u t i o n t o t h e e f f e c t i v e d i h e d r a l parameter a t p o s i t i v e v a l u e s of a and a r e d u c t i o n i n t h e r o l l i n g -
c%
due t o t h e e f f e c t s of t h e simulated j e t t a i l - c o n t r o l parameter C l g h , Smaller r e d u c t i o n s i n t h e s e parameters were a g a i n obtained from nozzle A.
w i t h nozzles B and C .
The curves presented i n f i g u r e s 15 and 16 summarize t h e s i g n i f i c a n t simulated-jet-exhaust e f f e c t s on t h e lateral and d i r e c t i o n a l s t a b i l i t y and c o n t r o l obtained during t h i s i n v e s t i g a t i o n . The model w a s a l s o tested
a t t h e reduced Reynolds number o f 1.20 x l o 6 and a l s o w i t h t h e speed
b r a k e s d e f l e c t e d 3 5 O ; however, t h e s e data showed no n o t i c e a b l e change i n t h e lateral and d i r e c t i o n a l s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s between t h e simulated-jet-on and j e t - o f f conditions.
F u l l - s c a l e Considerations I n view o'f t h e foregoing r e s u l t s , one question which might n a t u r a l l y arise i s whether these simulated-jet-exhaust e f f e c t s are t r u l y represen- t a t i v e of t h o s e which may be encountered during an a c t u a l f l i g h t . I n it must be noted t h a t during t h i s i n v e s t i g a t i o n , s c h l i e r e n answering, photographs i n d i c a t e d t h a t t h e boundary l a y e r was laminar over t h e f u l l l e n g t h of t h e fuselage and v a r i o u s attempts t o induce t r a n s i t i o n a r t i - f i c i a l l y were unsuccessful. O n a f u l l - s c a l e v e h i c l e , however, boundary- l a y e r t r a n s i t i o n may very l i k e l y occur ahead of t h e t a i l surfaces; t h u s m m e m m emm m m m m m m m m m e m m m m m m m e m m m m m m m m m m m m m m a m m m m m .
# t h e jet-induced separated-flow r e g i o n s and consequently t h e j e t - i n t e r f e r e n c e effects would b e expected t o b e smaller. Therefore, although t h e use of t h i s simulated-jet-exhaust technique may not p r e d i c t t h e e x a c t magnitude of t h e s e e f f e c t s , it i s b e l i e v e d t h a t t h e s e r e s u l t s are u s e f u l f o r i n d i c a t i n g t r e n d s and p o i n t i n g o u t problem areas which should be con- s i d e r e d i n t h e design of h i g h - a l t i t u d e , high-performance aircraft.
CONCLUDING REMARKS L 4 An i n v e s t i g a t i o n has been conducted i n t h e Langley 11-inch hyper- s o n i c t u n n e l a t a free-stream Mach number of 6.86 t o determine t h e j e t - 1 i n t e r f e r e n c e e f f e c t s a t high j e t - s t a t i c - p r e s s u r e r a t i o s on t h e s t a b i l i t y 2 and c o n t r o l of a research-type a i r p l a n e c o n f i g u r a t i o n .
The r e s u l t s of t h i s i n v e s t i g a t i o n i n d i c a t e d t h a t t h e j e t - e x h a u s t a considerable separated-flow r e g i o n i n t h e v i c i n i t y boundary induced of t h e t a i l surfaces, t h e e x t e n t of which increased w i t h j e t - s t a t i c - p r e s s u r e r a t i o and decreased with an i n c r e a s e i n Reynolds number.
A j e t - i n t e r f e r e n c e - f i e l d - s i m u l a t i o n technique w a s developed which approximated t h e jet-induced separated-flow r e g i o n over t h e t e s t con- f i g u r a t i o n at low angles of a t t a c k . The r e s u l t s obtained w i t h t h i s simu- lated jet-exhaust technique i n d i c a t e d t h a t t h e separated-flow r e g i o n caused a l a r g e r e d u c t i o n i n t h e l o n g i t u d i n a l s t a b i l i t y and c o n t r o l and a smaller r e d u c t i o n i n t h e lateral and d i r e c t i o n a l s t a b i l i t y and c o n t r o l .
By extending t h e divergent s e c t i o n of t h e nozzle, and t h u s reducing t h e j e t - s t a t i c - p r e s s u r e r a t i o , t h e s e l o s s e s were diminished.
Langley Research Center, National Aeronautics and Space Administration, Langley F i e l d , V a . , J u l y 27, 1959.
a . .a* a . 0 0 0 U . . . . a a . . a 0 .
REFERENCES 1. Cortright, Edgar M., Jr., and Schroeder, Albert H.: I n v e s t i g a t i o n a t Mach Number 1.91 of Side and Base Pressure D i s t r i b u t i o n Over Conical B o a t t a i l s Without and With J e t Flow Issuing From Base. NACA RM ~ 5 1 ~ 2 6 , 1951.
2. Cortright, Edgar M., Jr., and Kochendorfer, Fred D.: Jet E f f e c t s on Flow Over Afterbodies i n Supersonic Stream. NACA RM E53H25, 1953.
L 4 3. Love, Eugene S.: Aerodynamic Investigation of a Parabolic Body of 1 Revolution a t Mach Number of 1.92 and Some E f f e c t s of an Annular NACA TN 3709, 1956.
Supersonic Jet Exhausting From the Base.
(Supersedes NACA RM LgKOg. ) I n v e s t i g a t i o n at 4. Bromm, August F., Jr., and O'Donnell, Robert M.: Supersonic Speeds of t h e E f f e c t of Jet Mach Number and Divergence' t h e Nozzle Upon t h e Pressure of the Base Annulus of a Body Angle of of Revolution. NACA RM ~54116,1954.
5. Grigsby, C a r l E.: A n Investigation of t h e E f f e c t s of Jet Exhaust and Reynolds Number Upon t h e Flow Over t h e V e r t i c a l S t a b i l i z e r and Rudder of t h e Douglas D-558-11 Research Airplane a t Mach Numbers of 1.62, 1.93, and 2.41. NACA RM L54E03, 1954.
6. McLellan, Charles H., W i l l i a m s , Thomas W . , and Beckwith, Ivan E.: Investigation of t h e Flow Through a Single-Stage Two-Dimensional NACA TN 2223, Nozzle i n t h e Langley 11-Inch Hypersonic Tunnel.
7. Love, Eugene S.: I n i t i a l I n c l i n a t i o n of t h e Mixing Boundary Separating An Exhausting Supersonic J e t From a Supersonic Ambient Stream. NACA RM ~ 5 5 ~ 1 4 , 1956.
8. Love, Eugene S., Woodling, Mildred J., and Lee, Louise P.: Boundaries of Supersonic hisymmetric Free Jets. NACA RM ~ 5 6 ~ 1 8 , 1956.
9. Penland, Jim A., Fetterman, David E., Jr., and Ridyard, Berbert W.: S t a t i c Longitudinal arid Lateral S t a b i l i t y and Control Characteris- t i c s of an Airplane Configuration Having a Wing of Trapezoidal Plan Form With Various T a i l A i r f o i l Sections and T a i l Arrangements a t a Mach Number of 6.86. NACA RM L55F17, 1955.
...... . .e.- ..
e . e . .. e- .....
e e e. e e.
e. . e * e e e e. e. e e e.. e. e.. e.
TABLE I GEOMETRIC CKARACTERISTICS OF MODEL Wing :
. . . . . . . . . 11.520
Area. t o t a l . sq i n . . . . . . . . . . . . .
. . . . . . . . . 6.050
Area. exposed. s q i n . . . . . . . . . . . .
. . . . . . . . . 5.366
Span. i n . . . . . . . . . . . . . . . . . .
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 2.500
. . . . . . . . . 3.578
Root chord. i n . . . . . . . . . . . . . . .
. . . . . . . . . 2.640
Root chord. exposed. i n . . . . . . . . . . .
. . . . . . . . . 0.716
Tip chord. i n . . . . . . . . . . . . . . . .
. . . . . . . . . 2.465
Mean aerodynamic chord. i n . . . . . . . . .
Sweepback angles .
Leading edge. deg . . . . . . . . . . . . . . . . . . . . . 36.75
25-percent element. deg . . . . . . . . . . . . . . . . . . 25.64
T r a i l i n g edge. deg . . . . . . . . . . . . . . . . . . . . . -17.74
. . . . . . . . . 0.200
T a p e r r a t i o . . . . . . . . . . . . . . . .
. . . . . . . . . 0 . 0 0
DilAedralangle. deg . . . . . . . . . . . .
. . . . . . . . . 0.00
Incidence angle. deg . . . . . . . . . . . .
A i r f o i l s e c t i o n ( p a r a l l e l t o fuselage c e n t e r
l i n e ) . . . . . . . . . . . . . . . . . . . NACA 66005 (modified)
Horizontal t a i l :
Area. t o t a l . sq i n . . . . . . . . . . . . . . . . . . . . . . 6.643
Area. exposed. sq i n . . . . . . . . . . . . . . . . . . . . . 2.878
Span. i n . . . . . . . . . . . . . . . . . . . . . . . . .
4.339
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 2.833
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.206
. . 1.658
Root chord. exposed. i n . . . . . . . . . . . . . . . . . .
. . 0.506
Tip chord. i n . . . . . . . . . . . . . . . . . . . . . . .
. . 1.184
Mean aerodynamic chord. exposed. i n . . . . . . . . . . . .
Sweepback angles -
. . 50.58
Leading edge. deg . . . . . . . . . . . . . . . . . . .
. . 45.00
??-percent element. deg . . . . . . . . . . . . . . . .
. . 19.28
T r a i l i n g edge. deg . . . . . . . . . . . . . . . . . . .
. -15.000
Diliedral. deg . . . . . . . . . . . . . . . . . . . . . .
A i r f o i l s e c t i o n ( p a r a l l e l t o f u s e l a g e c e n t e r (modified)
l i n e ) . . . . . . . . . . . . . . . . . . . NACA 66005
Upper v e r t i c a l t a i l :
Area. exposed. sq i n . . . . . . . . . . . . . . . . . . . . . 2.356
Span. exposed. in . . . . . . . . . . . . . . . . . . . . . . . 1.099
Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . 0.516
T a p e r r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.741
. . . . . . . .
0 0 0 0 0 . 0 0 0 - m b - 0 0 . 0 0 0 TABLE I.- Concluded GEOMETRIC CHARACTERISTICS O F MODEL
. . . . . . 2.450
Root chord. i n . . . . . . . . . . . . . . . . . .
. . 1.815
Tip chord. i n . . . . . . . . . . . . . . . . . . . . . . .
Mean aerodynamic chord. i n . . . . . . . . . . . . . . . . . . 2.148
Sweepback angles .
. . . . . . 30.000
L Leading edge. deg . . . . . . . . . . . . . . .
25-perceiit element. deg . . . . . . . . . . . . . . . . . . 23.413
T r a i l i n g edge. deg . . . . . . . . . . . . . . . . . . . . . 0.000
f u l l wedge
A i r f o i l section ( p a r a l l e l t o fuselage center l i n e ) . . loo
Leading-edge radius. i n . . . . . . . . . . . . . . . . . . . . 0.010
. . . . . . 1.323
Area. c o n t r o l surface. i n . . . . . . . . . . . . .
. . 2.248
Root chord. c o n t r o l surface. i n . . . . . . . . . . . . . .
Mean aerodynamic chord. c o n t r o l surface. i n . . . . . . . . . . 2.039
Lower v e r t i c a l t a i l :
Area. s q i n . . . . . . . . . . . . . . . . . . . . . . . . . . 1.982
Span. exposed. i n . . . . . . . . . . . . . . . . . . . . . . . 0.920'
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.429
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0 783
Root chord. i n . . . . . . . . . . . . . . . . . . . . . . . . 2.450
Tip chord. in . . . . . . . . . . . . . . . . . . . . . . . . . 1.919
Mean aerodynamic chord. i n . . . . . . . . . . . . . . . . . . 2.200
Sweepback angles -
Leading edge. deg . . . . . . . . . . . . . . . . . . . . . 30.000
25-percent element. deg . . . . . . . . . . . . . . . . . . 23.413
T r a i l i n g edge. deg . . . . . . . . . . . . . . . . . . . . . 0.000
A i r f o i l s e c t i o n ( p a r a l l e l t o fuselage c e n t e r l i n e ) . . 1 0 ' f u l l wedge
Leading-edge radius. i n . . . . . . . . . . . . . . . . . . . . 0.010
Area. c o n t r o l surface. i n . . . . . . . . . . . . . . . . . . . 1.149
Root chord. c o n t r o l surface. i n . . . . . . . . . . . . . . . . 2.248
Mean aerodynamic chord. control surface. i n . . . . . . . . . . 2.093
Fuselage :
Length. i n . . . . . . . . . . . . . . . . . . . . . . . . . . 11.76
Maximum diameter. i n . . . . . . . . . . . . . . . . . . . . . 1.12
Maximum width (including s i d e f a i r i n g s j . i n . . . . . . . . . . 1.76
Fineness r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 10.50
Base diameter . . . . . . . . . . . . . . . . . . . . . . . . 0.960
J
t
IO k
I 0 0 N rr) W - c
!
L W ’c ? !
t c W
z
L
e. e.. e e e e e e e e 0 e e e . . e .
e . 0 . e .
e . . . e 0 e. e e. e e e . . e . e e . e . 19 e. e.. e.
e 0 e e.. e.
TOP VIEW _ -
- -- -- -
.350
S I D E VIEW
All dimensions are i n i n c h e s .
F i g u r e 2 . - Details of t h e speed brakes.
.. ... . ... . .. .. . . . 0 . . ..
.-- k d
I
%I
1 7 In
rn rl L FI Ill rri e. e.. e e e. e. e.. e e.. e.
e . . . . e 4 . . e e . e . . e e . . . . e . . e .
e . . e e . e .
e. 0 . . ~ 0 . e e.. e.
t
L
Nozzle A
2 oo
. 7 0 8 4 20° Nozzle B Nozzle C Figure 4.- Details of the nozzle configurations.
All dimensions are in inches.
100 120 140 Altitude, fi Figure 5 . - E f f e c t of altitude on jet-static-pressure ratio.
= 600 l b / s q in.; rj = 1.25.
’t,j .. ... ... ... i - . .. ..... ..
.... . ... ..
... . . .. ..
.. ... .. .. . .e d 0 . ... ..
-% d a3 Q m N m 0 rl Q Q f m JP.8 hl N 0 d Q -f m hl 1 1 I 0 r( A k t , I I
I 'I
f 0 0 '9 '9 N l N" "1 N " I co al 0 . 0.. . 0.. . 0 . 0.. 0 .
e . .
2 6 : : 0 : : 0 : 0 . . .
0 . 0 . 0 0 . .
0 . 0.. . . . ... 0 .
j e t off Pj/Pm = 938 (a) Nozzle A; R = 0.57 x 10 .
L-59-5008 Figure 9.- Schlieren photographs of the flow f i e l d s produced by the a i r j e t a t various Values of p p, and Reynolds number. a = 0'; M, = 6.86. j/ c t x Jet off Pj/P, = 475 Pj/P, = 195 no canopy, pj/pa = Y O (b) Nozzle A ; R = 1.20 X 10 .
L-59-5009 Figure 9. - Continued .
0 . ... . 0.. . .e 0 . . . . ... 0 .
Nozzle A; Jet off Nozzle A j p /p = 470 t J a Nozzle A i pj/poo = 203 Nozzle A; pj/p,, = 100 Nozzle B j pJ/pm = 170 Nozzle Cj pj/p, = 139 L- 59 -5010 ( c ) R = 2.05 x io .
F i g u r e 9.- Concluded.
NOZZLE R
rt off 2.05 x 10 6
I 1.20 2.05 '2.05 C
c
Figure 11.- E f f e c t of simulated j e t exhaust on l i f t and drag c h a r a c t e r - i s t i c s .
M , = 6.86; € i h = 0'.
p a m m m m m m me e a e a m m m e m m m m m e m m m m a m a ma m m m a m mee me d E u
s
.10 CD Figure 13.- Effect of Reynolds number on longitudinal characteristics M , = 6.86.
of model with jet off.
5A * Cm Jet off Nozzle A .10 CD -05 .05 CL Figure 14.- E f f e c t of simulated j e t exhaust on l o n g i t u d i n a l c h a r a c t e r - istics of t h e o d e 1 w i t h speed brakes d e f l e c t e d 35'. M , = 6.86; .
R = 1.20 X 10 .
. 0 2 NOZZLE R Jet o f f 2.05 X 10 A 1.20 B 2.05 C 2.05 Cn 6 V - .002 -4 -2 0 2 4 a, del3 .
Figure 15.- Effect of simulated jet exhaust on directional stability and control. M , = 6.86.
e. e.. e 0 e .e e. e e.. e e.. e .
* * e e . . . e * e e . . e e .
. e * * e e . e .
e.. e.
c:
- .001
2.05 X 10"
0 J e t off
OA 1.20
.001 -.001
. 001
c z 6 v 0 -.001 -4 -2 0 2 4 a, deg Figure 16.- E f f e c t of simulated j e t exhaust on lateral s t a b i l i t y and c o n t r o l . M , = 6.86.
NASA - Langley Field, Va.
L-412