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Copy 1, 17CONFI DENTIAL RM L56A16

rP

RE£S EAR CH M EMO RANDU M,.

FLIGHT INVEST-IGATION OF THE EFFECT OF A PROPULSVE JET

POSITIONED ACCORDING TO THE TRANSONIC AREA RULE ON

- THE DRAG COEFFICIENTS OF A SINGLE-ENGINE

DELTA-WING CONFIGURATION AT MACH

NUMBERS FROM 0.83 TO,1.36

By Joseph H.I Judd and Rl~aph A. Falanga

Langley Aeronautical LaboratoryLangley Field, Va.

CLAS EIED DOCURMNT

This material contains information affecting the National Defense of the U: ted States within the meaningof the espionage laws, Title 18, U.S.C., Secs. 793 and 794, the transmission or revelation of which in anymanner Lo an unauthorized person Is prohEbiCed by law.NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICSDEL- WAS HINGTON

Lae April 13, 1956

La.gley Field, Va

CONLFI DENTI A MENNATIONAL5 ADISR COMMITTE

NACA RM L56A16 CONFIDENTIAL

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

FLIGHT INVESTIGATION OF THE EFFECT OF A PROPULSIVE JET

POSITIONED ACCORDING TO THE TRANSONIC AREA RULE ON

THE DRAG COEFFICIENTS OF A SINGLE-ENGINE

DELTA-WING CONFIGURATION AT MACH

NUMBERS FROM 0.83 TO 1.36

By Joseph H. Judd and Ralph A. Falanga

SUMMARY

A 600 delta-wing configuration with an engine location in a podcontiguous to the underside of tie fuselage was flight tested to deter-mine the effects of the flow field of and about the propulsive jet onthe drag, lift, and longitudinal stability. A solid-propellant rocket}.j ' motor was used to simulate the sonic exhaust jet of a turbojet engine

plus afterburner and operated at a jet-exit static-pressure ratio ofapproximately 4. The jet-on Mach number varied from 0.83 to 1.36 and

Reynolds numbers varied from 6.9 X 106 to 10.4 x' 06, whereas the jet-off flight covered a Mach number range from 0,83 to 1.63 and Reynoldsnumbers from 6.9 X 106 to 17.3 X 106.

Jet-on drag coefficients were lower than jet-off drag coefficientsat transonic speeds. The maximum difference in drag coefficient, 0.0156,was attained at a Mach number of 0.99. The difference between jet-onand jet-off drag-rise coefficients to a Mach number of 1.0 can be pre-dicted approximately for this configuration by use of the transonicarea rule and inclusion of the jet in the cross-sectional-area distrib-ution for the jet-on case. Above a Mach number of 1.27, the jet-offdrag coefficients were lower than jet-on drag coefficients for thisconfiguration.

*Lift-coefficient increments of 0.045 between jet-on and jet-offflight were attained to a Mach number of 0.92.

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2 CONFIDNIAL NAcA Rm L56A1.6

INTRODUCTION

Recent investigations of the effect of the propulsive Jet on the

aerodynamic characteristics of bodies and airplane-configurations have

shown that important changes in drag and lift cejet ean occur~between jet-on and jet-off conditions. For example, positive increments! in base and boattail pressure coefficients, caused by the Jet ,expansion,*

have reduced drag coefficients between jet-on and jet-off conditions forbodies of revolution with jet exhausting from the base (refs. 1 and 2).'Also, the flow field produced by the expansion of the jet, when locatedI. in a favorable position below the wing, has been shown in references 3, 4,and 5 to. produce appreciable increments in lift from the jet-off condi-tions. It was proposed, therefore, to use the expansion of the jet toreduce the drag from the jet-off condition of an airplane configurationby the application of the concept of the transonic area rule (ref. 6) forthe jet-on condition. In this case the jet was considered a solid bodyand was used to fill the cross-sectional-area distribution and reduce itsslope at the rear of the configuration. A model of a 600 delta-winginterceptor configuration, whose single engine was located in a pod con-tiguous to the underside of the fuselage and designed according to theabove-stated principle, was flight tested by the Pilotless AircraftResearch Division of the Langley Aeronautical Laboratory as part of acurrent program on jet effects.

The jet exit was located slightly ahead of the wing trailing edgeand below the wing. The propulsive jet issuing from the sonic exhaustnozzle simulated exhaust parameters of a current turbojet plus after-burner at an altitude of 35,000 -feet and a Mach number of 1.3 by uti-lizing a solid-propellant rocket motor designed according to reference 7.

The flight test was made at the Langley Pilotless Aircraft ResearchStation at Wallops Island, Va. The Mach number range of this test wasfrom 0.85 to 1.63 for jet-off flight and 0.83 to 1.36 for jet-on flight.

The jet-off Reynolds number range varied from 6.9 x 106 to 17.3 x 106and jet-on Reynolds number range from 6.9 x 106 to 10..4 x 106.

SYMBOLS

ac distance from leading edge of mean aerodynamic chord to aero-dynamic center, percent mean aerodynamic chord, positiverearward

A cross-sectional area, sq ft

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NACA RM L56A16 CONFIDENTIAL

b wing span, ft

c wing chord, ft

wing mean aerodynamic chord, ft

C IPB - POOCp,B base pressure coefficient, qIqCD drag coefficient, DraZ

qS

CL lift coefficient, LqS

C lift-curve slope, dL per-deg

CL,T trim-lift coefficient

Cm pitching-moment coefficient, measured about model center ofgravity

Cmj pitching moment due to jet thrust about center of gravity

dCmC static-stability derivative, - per degree

C mq 0) per radian

(-2

6M C =per radian

(Cq + Cm ) longitudinal damping derivatives, per radian

moment of inertia in pitch about model center of gravity,

slug-ft2

I fuselage length, ft

M Mach number

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4 CORIILMAL xAcA iRM L56A:I6

PBengine-pod base pressure, 1b/sq ft

Pe jet-exit static pressure, ib/sq ft

pW free-stream static-pressure, ib/sq ft

P period of short-period longitudinal oscillation, sec

q dynamic pressure, ib/sq ft

r radius of equivalent body, ft

R Reynolds number, based on wing mean aerodynamic chord

S total plan form area, sq ft

t time from launch, sec

tl time required for short-period oscillation to damp to one-halfamplitude, sec

t/c wing-thickness ratio

V velocity, ft/sec

W weight of model, lb

x distance along fuselage measured from nose of engine pod, ft

distance of center of gravity from leadirng edge of wing meanaerodynamic chord, positive rearward

z distance of center of gravity from leading edge of wing meanaerodynamic chord, positive upward

M angle of attack at the center of gravity, measured fromfuselage center line, deg

aT trim angle of attack, deg

a . da/dt, radians/sec57.3

0 angle of pitch at the model center of gravity, measured fromfuselage center line, radians

: =radians/secdt'

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NACA RM L56A16 CONFIDENTIAL 5

MODEL AND APPARATUS

Model

A three-view drawing of the flight model is shown in figure 1 andthe basic geometric parameters are given in table I. The present modelwas derived from a research configuration (model 7 of ref. 8) which hada 600 delta wing mounted on a parabolic body of revolution and no hor-izontal tail. The test configuration utilized an engine installationlocated in a pod contiguous to the underside of the fuselage. To ensuredynamic lateral stability in the test model two auxiliary fins weremounted at the rear of the engine pod. The use of two fins was dictatedby the underslung booster configuration.

The basic fuselage was a parabolic body of revolution. To housethe propulsive unit an engine pod was mounted on the underside of thefuselage, 4.3 inches below the fuselage center line, To cope with theproblems of telemeter installation, a nose fairing and a cylindricalsection were mounted ahead of the engine pod. Ordinates of the fuselageand pod are given in table II. A conical boattail of 5.920 half-anglewas used on the engine pod. Figure 2, a drawing of the engine, showsthe rear of the pod and the jet and base diameters.

The,- wing used on the configuration was a 600 delta wing of solidmagnesium whose thickness ratio varied from 5 percent at the root to6 percent at the tip. The airfoil had a flat center section, 0.5c,which was located rearward of 0.3c. Leading and trailing edges werefaired to the flat center section by using NACA airfoils as shown infigure 5. The vertical fin was swept 600 at the leading edge and had ahexagonal airfoil section whose thickness 'ratio varied from 1.7 percent

at the root to 5.2 percent at the tip. Two auxiliary fins at a 450 anglebelow the wing plane were attached to the engine pod. These fins wereflat steel plates, 0.125 inch thick, with sharpened leading and trailingedges.

The basic turbojet simulator utilized in this model consisted of acombustion chamber, a flow control nozzle, and a convergent sonic-exitnozzle. A Cordite SU/K propellant grain 23.6 inches long generatedthe exhaust gases to simulate a current turbojet plus afterburner

(ref. 7). The Jet-exit diameter was 3.792 inches and the jet base diam-eter was 4.125 inches, corresponding to a jet area of 0.0786 square footand a base area of 0.0925 square foot. Figure 4 shows the relationshipof the Jet exhaust nozzle, the fuselage, fins, and wing.

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6 CONFIDENTIAL NACA RM L56A16

Booster and Equipment

An underslung booster, as shown in figure 5 and described in ref-erence 9, was used to propel the model to maximum velocity. Two lugs.,shown in figure 2, were welded to the engine pod in order to provide aforward attachment between the model and booster. The rear boosterattachment was provided with an adapter which fitted in the exhaust noz-zle of the model.

Instrumentation

A six-channel telemeter, located in the nose of the engine pod,continuously transmitted measurements of free-stream total pressure,angle of attack, longitudinal and normal acceleration, combustion cham-ber static pressure, and nozzle static pressure. The locations of thepressure orifices used to measure combustion chamber static pressure andthe exit-nozzle static pressure are shown in figure 2. The longitudinalaccelerometer was located at station 54.747 and in the wing mean chordplane, whereas the normal accelerometer was located at station 52.625and in the wing mean chord plane. Data for the flight tests wereobtained by use of telemeter, CW Doppler velocimeter, NACA modifiedSCR 584 tracking radar and rawinsonde. Model velocity, obtained withthe velocimeter, was corrected for wind velocity which was determinedfrom rawin measurements.

TEST PROCEDURE -

Tests (7

The turbojet simulator combustion chamber pressure, nozzle staticpressure, and thrust were measured in a preflight motor firing in theLangley rocket test area. Using these data, calibration curves of therocket thrust as a function of both the combustion chamber pressure andthe nozzle static pressure were obtained. The purpose of measuringthrust by two independent instruments was to provide insurance againstthe malfunctioning of a pressure cell during the flight.

The flight model was launched from a mobile launcher (fig. 5). Anunderslung, single ABL Deacon rocket motor boosted the configuration tothe peak Mach number. Jet-off data were obtained during the deceleratingflight after separation of model from the booster. Jet-on data wereobtained during firing of the turbojet simulator which was started atthe lowest test Mach number in the deceleration phase. During jet-offflight the model was disturbed in pitch by separation from the booster,

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NACA RM L56A16 CONFIDENTIAL

and later by a pulse rocket. During jet-on flight the model was dis-turbed in pitch when the turbojet simulator was started and when themodel passed a Mach number of 1.0. A time history of the angle ofattack during the flight is given in figure 6. From an examination ofthis figure, it can be seen that the angle-of-attack disturbance causedby separation of the model from the booster has been modulated and, frompast experience, indicates that a lateral disturbance probably occurredat the same time. Since the model was not instrumented for lateral dis-turbances, the pitching data obtained during this portion of the flightcannot be analyzed. Also marked on this figure is the time when thesecond pulse rocket fired, at which time several instruments failed.These were the longitudinal and normal accelerometers and the combustionchamber static pressure. Since the angle of attack was corrected to themodel center of gravity during pitching disturbances by using data fromthe accelerometers, the values of angle of attack shown in figure 6 werestopped at the time when the accelerometers failed. Data beyond-thispoint were obtained after the pulse rocket disturbance damped but werenot plotted in figure 6.

The variation of Reynolds number (based on wing mean aerodynamicchord) with Mach number for jet-on and jet-off flight is presented in

figure 7. Time histories of velocity, Mach number, and free-stream~dynamic pressure during the flight are given in figures 8 and 9.

1AnalysisLongitudinal accelerations of the model were obtained from two

sources: (1) longitudinal accelerometer and (2) differentiation of themodel velocity. Thus, when the longitudinal accelerometer failed, dragwas still obtainable. The method of obtaining jet-on and jet-off dragcoefficients is explained in reference 1.

The angle-of-attack indicator was mounted ahead of the nose and themeasured angles of att2ack were corrected to those at the model center ofgravity, according to reference 10.

The method of obtaining lift and longitudinal stability coefficientsfrom transient longitudinal disturbances is given in reference 11.During jet-on flight, the model weight, moment of inertia, and center ofgravity changed as the rocket fuel burned. The variation of these quan-tities with time is given in figures 10 and ll All data obtained duringpitching oscillations were computed using these values.

The engine-pod base pressure coefficient was computed from the exit-nozzle static pressure during jet-off flight. It was assumed that theexit-nozzle static pressure represented the magnitude of base pressureoccurring over the entire base.

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8 CONFIDENTIAL NACA BM L56A16

ACCURACY

To establish telemeter instrument accuracies, statistical data havebeen compiled on flight instrument measurements over a number of years,and on the basis of this information the maximum probable error isbelieved to be 1 percent of the full-scale calibrated range for thetelemetered measurements. These maximum probable errors in measurements,which have been used to compute the errors in base. pressure, drag, andlift coefficients for several Mach numbers, are tabulated below.

Mach C pb CDjet-off CDjet-on C.jetoffnumber C

0.93 ±0.097 *0.0067 *0.0092 *0.01761.25 ±.052 1.0036 ±.0054 +.00 71.60 ±.028 ±.0021 --. 0023

The velocity measured by the CW Doppler velocimeter is known tohave an error of less than 1 percent at supersonic speeds and less than.2 percent at subsonic speeds.. Since Mach number is.determined fromvelocity, the above-quoted errors also apply to Mach number.

The magnitude of these computed errors is large in comparison withthe magnitude of the measured coefficients. However, the longitudinalaccelerations used to compute the chord force coefficients were measureddir6ctly by telemeter and also were obtained by differentiation of themodel velocity. These values were compared and were nearly the same,,the difference being much smaller than the stated error. The measuredmotor pressures were compared with those obtained from the static testfiring and the specific impulses of the two firings were compared. Thesealso were much closer than the quoted accuracies. Because of thesechecks it is believed that the error of the jet-on drag coefficientsis no more than ±0.003, which is equal to the scatter of the dataobtained during pitching oscillations. Similarly, the error in jet-offdrag coefficients is approximately 10.002.

RESULTS AND DISCUSSION

Drag

The variation of total drag coefficient with angle of attack,obtained during pitching disturbances, is given in figure 12. Sincejet-off drag coefficients were obtained during pitching oscillations at

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NACA RM L56A16 CONFIDENTIAL 9

supersonic speeds, and jet-on drag coefficients were obtained duringpitching oscillations at transonic and sonic speeds, a direct comparisonof these data cannot be made. Although the jet-off drag coefficientsshow hysteresis as the model pitches, the minimum drag coefficient appar-ently occurs at 00 angle of attack and for the small angle-of-attackrange6 covered exhibits little variation with angle of attack. The jet-on drag coefficients show the same effect. Thus, when a comparisonbetween trim drag coefficients for jet-on and. jet-off flight is made,the effect of the difference in trim angle of attack will be neglectedinasmuch as the maximum trim angles were less than i10.

The variation of jet-on and jet-off total drag coefficients is pre-sented in figure 13, together with the base drag coefficient of theengine pod. The trim angle of attack for jet-on and jet-off flight isplotted in figure 14 and values of the jet-exit static-pressure ratio infigure 15. During the period of jet-on flight, the jet-exit static-pressure ratio is approximately 4.0 and remains relatively constant,

corresponding to flight with a current turbojet-plus-afterburner at3,000 feet altitude and a Mach number of 1.3.

The total jet-on drag coefficients are lower than the total jet-off*drag coefficients to a Mach number of 1.26 and reach a maximum difference* * of 0.0156 at M = 0.99. Above a Mach number of 1.26, the jet-off drag

coefficients are less than the jet-on drag coefficients. By subtractingthe base drag coefficient from the jet-off drag coefficient, the effectof the jet on the external drag can be determined. The differencebetween total jet-off drag coefficients (less base drag coefficients) andjet-on drag coefficients is plotted in figure 16. It should be notedthat the maximum reduction in drag coefficient occurs approximately atMach number 1.0. As was mentioned in-the introductory remarks, the jetwas located to fill out the cross-sectional-area distribution. Naturally,

after Mach number 1.0 the Mach lines from the jet sweep back at greaterangles. The influence of the jet is thus caused to affect a much smallerpart of the configuration and thereby to lower drag coefficient differ-ences between jet-on and jet-sff operation.

The variation of model cross-sectional area along the longitudinalaxis of the configuration and its equivalent body of revolution is given

in figure 17. On the side view of the configuration are several curvessh[wing the jet shape for different flight conditions. The jet bulge wasmeasured from schlieren photographs of a sonic jet operating in the 8-foottransonic wind tunnel at a Mach number of 1.0, and shadowgraph pictures ofa sonic jet operating at a Mach number of 1.4 in the preflight jet of theLangley Pilotless Aircraft Research Station at Wallops Island, Va. Themeasured values of jet diameter are close in magnitude. The value of jetcross-sectional area at a Mach number of 1.0 was used and a cylindrical jetshape was assumed after the initial bulge. The peak drag-rise coefficient

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10 CONFIDENTIAL NACA RM L56A16

was computed for the jet-on and jet-off case using the curves of refer-ence 15. The jet-off peak drag rise was 0.020 which agrees with themeasured peak drag rise of the configuration. The computed Jet-on peakdrag-rise coefficient was 0.0076. This should be expected to apply onlyat or-slightly above a Mach number of 1.0. The subsonic level of thejet-on drag coefficients was taken as the value at a Mach number of 0.85.Using this value, the drag rise to a Mach number of 1.02 (which was thesame as the Mach number for jet-off peak drag rise) was 0.0119. Thuswhile the value of drag rise to sonic speeds was higher than the estima-ted value for jet-on flight, the transonic area rule does predict apressure drag reduction between jet-off and jet-on flight.

Above a Mach number of 1.27 the jet-on drag coefficients are greaterthan the jet-off total drag coefficients. Thus the approach used toreduce the drag coefficients applied, only for the Mach number range forwhich it was intended.

The total drag coefficient for jet-off flight appeared to be high,as seen in figure 13. In order to check on these values, an attempt wasmade to estimate the drag coefficients of the configuration by additionof the drag coefficients of the components. Drag coefficients for thewing, body, and vertical tail were obtained from reference 8; boattaildrag coefficients for the engine pod were obtained from reference 15 andskin-friction drag coefficients from reference 14. The auxiliary finswere assumed to be flat plates and have turbulent skin friction over thesurface. Drag coefficients for these fins were estimated from refer-ence 14. These values are plotted in figure 18 and a considerable dif-ference is shown to exist between the measured and estimated total dragcoefficients with the estimated values being 0.0043 to 0.0068 below themeasured drag coefficients; Since the estimated values of pressure drag-,rise agree with the measured drag rise, the difference in drag level isattributed to drag caused by the interference of engine pod and fuselage.A similar drag difference caused by an unfavorable wing-fuselage juncturewas observed in reference 15; and also, the highdrag level of model 2 ofreference 9 was attributed to a similar fuselage-engine pod juncture.

Values of engine-pod base-drag coefficient presented in figure 13were obtained from a pressure measured at the wall inside the convergentsonic nozzle. Pressure coefficients for this orifice are presented infigure 19. These coefficients are considerably higher than values f6o acylindrical conical afterbody (ref. 14) with approximately the same boat-tail angle.

Lift

The variations of lift coefficient 4rith angle of attack obtainedduring pitching oscillations for jet-on and jet-off flight are given in

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NACA RM L56A16 CONFIDENTIAL ll

figure 20. Flagged symbols indicate increasing values of angle of attack,while unflagged symbols indicate decreasing angles of attack. Since thejet-on and jet-off lift coefficients were not obtained at the same Machnumbers, a direct comparison cannot be made. However, the Jet-on data,

as shown in figure 20a, indicate that at m = 0 the jet gives positiveincrements in lift coefficients, but that these decrease with increasingMach number. It is felt that this variation of incremental lift coeffi-cient at m = 0 with Mach number is a result of the disturbance causedby the jet moving rearward of the wing as free-stream Mach number isincreased. Thus, it appears that the operation of the jet increased themodel lift in the transonic speed range of the present test in a mannercomparable to that reported at supersonic speeds in reference 5.

Although the angle-of-attack range was limited, it appeared thatduring the transonic speeds of the jet-on flight the lift curve wasS-shaped. Thus, at some subsonic Mach numbers two values of lift-curveslope were obtained: (1) a value for m = 0, and (2) a value for agreater angle of attack. The variation of the slope of the liftcurve CL, with Mach number for jet-on and jet-off flight is presented

in figure 21 together with variation of lift-curve slope for a 600 delta-wing-body combination (ref. 17). The jet-on and jet-off data presentedare in agreement with reference 17, except at the lower Mach numbers.In this speed range the lift curve had a tendency to be S-shaped whichresulted in lower values of lift-curve slope hear a = 0 hence, thecause for disagreement with reference 17 at the lower Mach numbers.

Trim

The variation of trim angle of attack with Mach number is presentedin figure 14 for jet-on and jet-off flight. During jet-off flight, theconfiguration trims at positive angles of attack from transonic speedsto Mach number 1.24, whereas during Jet-on flight, the configurationtrims at a negative angle of attack at Mach numbers below 1.00. Fromthese data, it can be seen that the greatest change in trim angle ofattack between jet-on and jet-off flight occurs below M = 0.92 andthat the jet-on trim angle of attack is nearly 1.50 below the jet-offtrim angle of attack. The effect of the jet is large at Mach numbersfrom 0.83 to 0.92; this effect causes the model to trim negatively eventhough the turbojet simulator thrust was tending to trim it positively.It is felt that the radical trim change which occurred during jet-on

s flight was caused by the rearward shift of the jet effect on the wingand afterbody of the model as the Mach number increased. The trim-liftcoefficient CL,T for jet-on and jet-off flight is given in figure 22.

The plot indicates positive CLT for both jet-on and Jet-off flight

of approximately the same order of magnitude. This indicates that theflow field of the Jet produces an appreciable lift-coefficient increment

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12 CO NFIDENTIAL NACA RM L56A16

in this speed range as great as 0.045 at a Mach number of 0.85 and also,as mentioned above, an appreciable nose-down pitching moment. The thrustof the jet produces a nose-up pitching moment whose variation with timeis plotted in figure 23. The reduction in magnitude (from 0.026 to 0.0115)of the pitching-moment coefficient due to thrust is mainly due to theincrease in dynamic pressure as the speed increases, since the thrustremained relatively constant during jet-on flight.

Longitudinal Stability

The period of the short-period longitudinal oscillations is givenin figure 24. The pitching oscillations are a result of disturbing themodel in pitch by firing pulse rockets during jet-on and jet-off flightand by firing the turbojet simulator. The static stability for the modelis presented in figures 25 and 26, where the variation of the static-

stability derivative dEm and aerodynamic center with Mach number aredxL

shown, respectively. The period was used to compute Cm, and these

C M and experimental CL, were employed to compute the aerodynamic

center. The general trend of Jet-on Cma with Mach number appears nor-

mal for wing-body combinations of this type as does the aerodynamic center.The aerodynamic center moves rearward with increasing Mach number, asexpected, but the values of ac presented at M 1.00 appear to besomewhat on the high side as compared to what would be expected for thewing-body combination. Probably the auxiliary fins contributed to theincreased rearward shift of the ac.

The time required for the short-period longitudinal oscillation todamp to one-half amplitude is shown in figure 27 and the damping deriv-atives Cm + Cm, are shown in figure 28. The damping derivatives indi-

cate that the model is dynamically longitudinally stable throughout thetest Mach number range. -Also plotted in figure 27 is a theoretical curveof damping derivatives for this model at supersonic speeds, computed bymethod of reference 18. The supersonic experimental values from thistest indicate good agreement with the theoretical values.

CONCLUDING REMARKS

A flight investigation of a 600 delta-wing configuration with anengine location in a pod contiguous to the underside of the fuselage was

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NACA RM T56A16 CONFIDENTIAL 13

made to determine the effect of the propulsive jet on the drag, lift,and longitudinal stability. The jet-exhaust nozzle was located at91.62 percent of the wing root chord and 1.136 jet diameters below thewing mean chord plane. Jet-on data covered a Mach number range from 0.83

to 1.36 and Reynolds numbers from 6.9 x 106'to 10.4 x 106, whereas jet-off Mach numbers were obtained from 0..83 to 1.65 and Reynolds numbers

from 6.9 x 106 to 17.3 x 106. The jet-exit static-pressure ratio wasapproximately 4.0. The following statements summarize the results:

1. Jet-on drag coefficients were lower than jet- off drag coeffi-cients at transonic speeds. The drag-coefficient differencp reached amaximum value at 0.0156 at a Mach number of 0.99. Above a Mach numberof 1.26, the jet-off drag coefficients were lower than jet-on dragcoefficients.

2. The transonic-area-rule concept can be used to predict jeteffects on drag for this type of configuration.

3. Operation of the jet provided increases in lift coefficient-ofapproximately 0.045 at a Mach number of 0.8. The lift-coefficientincrements decreased above a Mach number of 0.92 since as Mach numberincreased the flow field induced on the wing by the jet moved rearward.

4. At Mach numbers between 0.85 and 0.92, the jet flow fieldinduced a nose-down trim angle of attack despite the nose-up moment dueto the thrust of the turbojet simulator. The difference between jet-onand jet-off trim angle was 1.500. After a Mach number of 0.92, the val-ues for jet-on and jet-off trim angle tended to converge.

Langley Aeronautical Laboratory,National Advisory Committee for Aeronautics,

Langley Field, Va., January 4, 1956.

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14 CONFIDENTIAL NACA RM L56A16

REFERENCES

1. Falanga, Ralph A.: A Free-Flight Investigation of the Effects ofSimulated Sonic Turbojet Exhaust on the Drag of a Boattail BodyWith Various Jet Sizes From Mach Number 0.87 to 1.50. NACAEM L55FOa, 1955.

2. Henry, Beverly Z., Jr., and Calm, Maurice S.: Preliminary Resultsof an Investigation at Transonic Speeds To Determine the Effectsof a Heated Propulsive Jet on the Drag Characteristics of aRelated Series of Afterbodies. NACA RM L55A24a, 1955.

5. Bressette, Walter E.: Investigation of the Jet Effects on a FlatSurface Downstream of the Exit of a Simulated Turbojet Nacelle ata Free-Stream*Mach Number of 2.02. NACA RM L54EO5a, 1954.

4. Bressette, Walter E., and Leiss, Abraham: Investigation of JetEffects on a Flat Surface Downstream of the Exit of a SimulatedTurbojet Nacelle at a Free-Stream Mach Number of 1.59. NACARM L55L15, 1955.

5. Falanga, Ralph A., and Judd, Joseph H.: Free-Flight Investigation ofthe Effect of Underwing Propulsive Jets on the Lift, Drag, andLongitudinal Stability of a Delta-Wing Configuration at Mach Num-bers From 1.23 to 1.62. NACA EM L55113, 1955.

6. Whitcomb, Richard T.: A Study of the Zero-Lift Drag-Rise Character-istics of Wing-Body Combinations Near the Speed of Sound. NACARM L52H08, 1952.

7. De Moraes, Carlos A., Hagginbotham, William K., Jr., and Falanga,Ralph A.: Design and Evaluation of a Turbojet Exhaust Simulator,Utilizing a Solid-Propellant Rocket Motor, for Use in Free-FlightAerodynamic Research Models. NACA RM L54115, 1954.

8. Sandahl, Carl A., and Stoney, William C.: Effect of Some SectionModifications and Protuberances on the Zero-Lift Drag of DeltaWings at Transonic and Supersonic Speeds. NACA EM L53L24a, 1954.

9. Judd, Joseph H.: A Free-Flight Investigation of the Drag Coefficientsof Two Single-Engine Supersonic Interceptor Configurations From MachNumber 0.8 to 1.90 to Determine the Effect of Inlet and Engine Loca-tions. NACA EM L55GO5a, 1955.

10. Mitchell, Jesse L., and Peck, Robert F.: An NACA Vane-Type Angle-of-Attack Indicator for Use at Subsonic and Supersonic Speeds. NACATN 3441, 1955. (Supersedes NACA EM L9F28a.)

CONFIDENTIAL

NACA RM L56A16 CONFIDENTIAL 15

11. Gillis, Clarence L., Peck, Robert F., and Vitale, A. James: Prelim-inary Results From a Free-Flight Investigation at Transonic andSupersonic Speeds of the Longitudinal Stability and Control Char-acteristics of an Airplane Configuration With a Thin Straight Wingof Aspect Ratio 5. NACA FM L9K25a, 1950.

12. Nelson, Robert L., and Stoney, William E., Jr.: Pressure Drag ofBodies at Mach Numbers up to 2.0. NACA RM L53I22c, 1953.

13. Patterson, R. T.: A Wind-Tunnel Investigation of the Drag of ConicalMissile Afterbodies at Mach Numbers From 0.40 to 2.47 (TED No. 1MBAD-314). Aero. Rep. 857, David Taylor Model Basin, Navy Dept.,Jan. 1954.

14. Van Driest, E. R.: Turbulent Boundary Layer in Compressible Fluids.Jour. Aero. Sci., vol. 18, no. 3, Mar. 1951, pp. 145-160, 216.

15. Welsh, Clement J., Wallskog, Harvey A., and Sandabi, Carl A.: Effectsof Some Leading-Edge Modifications, Section and Plai-Form Variations,and Vertical Position on Low-Lift Wing Drag at Transonic and Super-sonic Speeds. NACA RM L54KOl, 1955.

16. Mitcham, Grady L., Crabill, Norman L., and Stevens, Joseph E.: FlightDetermination of the Drag and Longitudinal Stability and ControlCharacteristics of -a Rocket-Powered Model of a 600 Delta-Wing Air-plane From Mach Numbers of 0.75 to 1.70. NACA PM L51104, 1951.

17. Henderson, Arthur, Jr.: Pitching-Moment Derivatives CM and

at Supersonic Speeds for a Slender-Delta-Wing and Slender-BodyCombination and Approximate Solutions for Broad-Delta-Wing andSlender-Body Combinations. NACA TN 2553, 1951.

CONFIDENTIAL

16 CONFIDENTIAL NACA RM L56A16

TABLE I

GEOMETRIC PARAMETERS OF CONFIGURATION

Fuselage and engine pod:Maximum frontal area, sq ft ... ................. . 0.340Engine pod base area, sq ft . . . . . . . . . . . . . . . . . 0.0925

Jet-exit area, sq ft ......... ...................... 0.0786

Wing:

Aspect ratio .......................... .. 2.31Taper ratio 0......................... 0Mean aerodynamic chord, ft............... 1.711Total plan form area, sq ft . ..... ... . ..... 3......80

Vertical fin:Aspect ratio (to fuselage center line) .. ........... . 0.895Taper ratio (to fuselage center line) ... .......... . . . 0.514Airfoil section ............. ........... Hexagonal airfoilArea (extended to fuselage center line) sq ft ... ........ 0.804

Auxiliary fins (for one fin):Aspect ratio (to engine pod center line) . . . . . . . . . . 1.064

Taper ratio (to engine pod center line) .. ........... . 0.504Airfoil section .. ............ ......... ... Flat plateArea (extended to engine pod center line) .......... . 0.709

CONFIDENTIAL

C

NACA RM L56A16 CONFIDENTIAL 17

TABLE II.- FUSELAGE ORDINATES

Fuselage R2

Engine pod --. 300

R1

[Letter dimensions apply only to this table and all

dimensions are given in inches]

Fuselage station R1 R2

0 01.000 .2502.000 .4803.000 .7105.000 1.l5o7.500 1.57010.000 1.95512.500 2.25215.000 2.42917.500 2.50022.625 2.500 023.015 2.500 .09723.210 2.500 .14523.600 2.500 .23924.575 2.500 .46926.525 2.500 .90228.47.5 2.500 i.29830.425 2.500 1.65834.325 2.500 2.26738.225 2.500 2.73042.125 2.500 3.04746.025 2.500 3.21849.925 2.500 3.24853.825 2.500 3.22157.725 2.500 3.16161.625 2.500 3.06965.525 2.500 2.94369.425 2.500 2.78570.063 2.500 2.75473.325 2.349 2.59477.225 2.089 2.37177.625 2.065 2.34581.125 2.11585.025 1.82687.625 -1.61589.925 1.31092.225 .83594.625 0

CONFIDENTIAL

18 COI'DENTIAL NACA IRA L56Ai6

%0

0

oo t

rqr

ILI.T I

1

I NN

NACA 1R4 L56A16 COIN'IDEITIAL 19

-H 4r ri

0o 0r 0

1-4 H)a o r

F4 0

CH)C.)& 0

$4

0.

o

00

0 a)4.3 cji00

4 IDE TIA

20 CONFIDENTIAL NA.CA. 1R L56A16

.oNj

0

0 400 trl0

-H0'

ot

0. 0 c

00 0-H- 00 c

00

.r~0

q-~ 0

CONFIERMAL

NACA 1RA L56A16 CONFORMTAL 21

Figure I.Photograph of tail section of flight model.'

e CQNFIDENTIAL

22 CONF'IDENTIAL NqACA 1R4 L56AI.

Figure 5.- Model and booster on mobile launcher. L895 1

C0NF'IDENT.AL

N4CA IR4 L56A16 CONF'IDEN~TIAL 2

Separation disturban~ce

a, dog

0

-2

Pulse rocket-h 4isturbance

3.0 3.5 14.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0

see

(a) Initial. portion of' decelerating flight.

'4

2 Second pulse rocket firing(Elements of telemeter failed

a, deg

0 A__ _ _ __ _ _ _

Disturbance at-2 Turbojet simlator Mach number 1.0

firing disturbance

12.0 12.-5 3-3.0 .13.5 14.0 14.5 15.0 15.5 16.0 165 17.0

t, see

(b) Latter portion of' decelerating flight and initial portion ofaccelerating flight.

Figure 6.- Valriation of angle of attack 'with time.

CONF'IDENTIAL

24 CONIMENTIAL NACA H L56A6

20 x 106

16-Decelerating flight

/ Ace, ierating flight

Figure 7 .- Variation of Reynolds number with Mach number.

2.0 2000 -- 1 1

M

1.8 1800- t

1.6 1600 --------

ft/sec

"l-4 11400 -- -

1.2 1200Mach nAmber- - - - - - - -

1.0 1000

.8 800

0 2 4 6 8 10 12 14 16 18 20

t, sec

Figure 8.- Variation of model velocity and Mach number with time.

CONFIIETIAL

C

NACA 1M L56A16 CONFIDENTIAL 25

4000

3500 -- -

3000 - I N2500 i ----

q, lb/sq t "

20000 0 - - "

15o00

10000

0 2 4 6 8 10 12 •14 16 18 20

Figure 9.- Variation of dynamic pressure with time.

.24---

.20

.16--

and.iz/ .1 2 - "'" -

•.08

toi

0 2 -6 8 10 12 14 16 18 20

t, see

Figure 10.- The variation of center-of-gravity position, measured fromthe leading edge of the wing mean aerodyfntic chord, with time.

CONFIDENTIAL

26 CONFIDENTIAL NAcA Fm t56A16

00

N

4'

coi

rd

;-4-

0 ccc

CONFIUMM0

NAGA RM L56A16 CONFIDENTIAL 27

.05T

M W ----- .05

.04 :_: - .0-

-+4 + 01.1 .0- j i 2261J

0 I 1 .2099~### ## og ID d..I-

(a)1+4+H -Jet-o draH cefiet .0()Jtof4rgcefiins

Fiue1.- The+ vaito f rgcefcinswt nl fatcobaie duin picin silains0lggd5Tbl nct

increasing~~~~~~~~~ M angle ofatc;ufagdsmbl niaedcesnangle o attack

.02IEITA

28 CO4FIDENqTIAL NACA IR4 L56A16

.054

.8~~ .I r.ll.2 1. 1111 1. . .

0

.8 .9 1. 1. I111 1 1 11 1. i.1 1.7

Fiur I4 Th aito of tri anl oI atac fo je-o an jetI-ofIf I II Iflightwith Mach. number.

[Ill 1C1]]Ill .1NA IllL

- -02- --

NACA IR4 L56A16 CONIPXENIAL 2

0 2- - - -4- -6 0 12 1-6 8 2

ti see

Figure 15.- The variation of jet static-pressure-ratio with time.

.02--

'D 0

02

8 09 10 11 12 13 . . . -

Fiue1,Tevraino0h ifeec ewe etofadjto

drgcefcet ihMc ubr

COCD)T

530 C0TIDNTIAL HACA IR4 L56A.6

(a),Side view of' model showing jet sizes for a, circula' jet,

.3.

(b) Equivalent body of' revolution for the configuration with and withoutthe jet.

.pJ:2,

.008'____

A/1.2 Oa e

0 .1 .2 .3 .14 .5 .6 .7 .8 .9 1.0 1.1

(c) Cross-sectional area distribution of the configuration with andwithout the jet.

Figure 17.- The effect of the jet on the cross-sectional area distributionof the configuration.

CONFIDMNIAL1

NACA R4 L56A16 CONF'IDENTIAL 3

Measured total

Component sumn

V -- Wing-Fuselage-Tall

I,;.02

.01

Eniepod Auxiliary

- - - ___ ~aae ngineBoattail~ i.Bas sknfitinfn

0-

Figure 18., The variation of the measured and estimated total-dragcoefficients and the drag coefficients of the components withMach number.

-3

2..+

0.8 .9 1.0 1.1 1.2 1.3 1-4 1~ 15 1.6 1.7

Figure 19.- The variation of the base pressure coefficient of the enginepod with Mach number.

CONFIEENTIAL

32 CONFIDENTIAL NACA 114 L56A16

.10.. . . . . .

.0 05.0

00

-2 -1 0 12-2 -1 0 0 0 0 0 1 2

a, dog

(b) Jet-o lift coefficients.

Figure~~~~~~~~~1 20-Te.aitino2lf offcetswt:ageofatcobtaine duin pichn osiltos Flge 1yl indicat

incrasig agle f atac; un~aged ymbos idicte dcresiIangleof atack

CONMI1.226A

N~ACA EM L56A16 CONFIDEN'TIAL 33

.082 1t l

Figre 1.-Th vaiatonof iftcuve lop Jwthon Ma number.

Jeto5 ae-o largr-agle

CLT

0

-. 05

-. 0.8 .9 1.0 1.1 1.2 1.3 14 1.5 1.6 1.7

M

tFigure 2.- The variation of trim-ift coicen with Mach number.

.1COillFiEtI

314 CONFIDENTIAL NACA ERM L56A16

. 03

.02

. 01

2 4 6 8 10 12 1h4 6 18 2t, sec

Figure 23.- Vaiation of' jet-on pitching-moment coefficient with time.

0.8 1 9 1 1. 11. 1111 i l 1.211 11 1 1 J i i.5 16 Je .7o

osilaton7wth=c nube f'o je-o and je-fI ight.

cm

0.8 .9 1.0 1.1 1.2 1.3 1.4 1. 1.6 17

Figure 2.- Variation of sa-t abit deri vaohesotiverwith Matchinumeosilatos it Mc mbrfor jet-on and jet-off flight.

COFIo6IA

NACA RM L56A16 CONFIDENTIAL 35

60- - - -- -.. j e t -o ff

and jet-ofJflight

0140 1.1 -1-.2 13,141 .5. 1.6 1.7

II II II I 1

IK

Figure 2.-Variation of thednai tietr dampto one-hal aumpertordewthMah ube ort-nand jet-off flight.

II I IIIIIIIIII I I I

tiIs e}11 1 Ill It i

0 --

.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 i.6 1.7

Figure 28.- Variation of ampidyngm eate with Mach number for jet-onand jet-off flight.

1111l- le eC F T

[itI II 1 II3 Jet-oII II I IIt , o1 1 --- -- , , + -I -H + I f -Hof A-"-10 '' l ' '' ' ' --------- i il

....... ~ ~ ~ _-I I II I I IIl

I i l l l l i l i l i i l l l l l i l l l I~l l II ~ l I I

and jet-off fl l ight.llllll

NACA~,, - age ilVaC N I E T A

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