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RESEARCH MEMORANDUIV---A TRANSONIC -WING INVESTIGATION IN THE LANGLEY 8-FOOT

HIGH-SPEED TUNNEL AT HIGH SUBSONIC MACH NUMBERS

AND AT A MACH NUMBER OF 1.2

WING-FUSELAGE CONFIGURATION HAVING A WING

OF 0 SWEEPBACK, ASPECT RATIO 4.0,

TAPER RATIO 0.6, AND NACA 65AO06

AIRFOIL SECTION

By Maurice S. Cah and Carroll R. Bryan

,! Langley Aeronautical Laboratory

Langley Field, Va.

CUS31FDSD BXUMSNT

.mtlon tffectlns the Natbml ~femaa of the Unl!d SUES withh bmLr&g of Llm tmnsmssalon or b rwehtfcn d its conlenh In anymmrarb MmSUum

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NATIONAL ADVISORY COMMI FOR AERONAUTICS

WASHINGTONMarch 6, 1951

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RM L51A02

NATIONAL ADVISORY COMMITTEE FOR

TECH LIBRARYKAFB,NM

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AERONAUTICS

RESEARCH MEMORANDUM

A TRANSONTC-WIN3 INVESTIGATION IN TEE

HIGH-SPEED TUNNEL AT HIGH SUBSONIC

&

LANGLEY 8-FCXYI

MACH NUMBERS

AND AT A MACH NUMBER OF 1.2

WING-FUSEIAGE CONFIGURATION HAVIlj2A WING

OF 0 SWEEPBACK, ASPECT RATIO 4.0,

TAPER RATIO 0.6, ANDNACA 65A06

AIRFOIL SECTION

By Maurice S. Cahn and Carroll R. Brysn

SUMMARY

As part of an NACA transonic research program, a series of wing-combinations is being tested in the Langley 8-foot high-speed

tunnel. This paper presents the results of & tivestigati;nof a wing-fuselage combination utilizing a wing of unswept quarter-chord line,aspect ratio 4, taper ratio 0.6, and an NACA 65AOC6 airfoil section.Lift, drag, and pitching-moment characteristics, downwash angles, andwake losses for various angles of attack at high subsonic Mach numbersand at a Mach number of 1.2 are presented.

Increasing the free-stresm Mach nmiber at low lift coefficientscaused.the wing-fuselage configuration to exhibit a decrease in lift-curve slope beginning at a Mach number of 0.90, a rapid decrease in themsximum lift-drag ratio at a Mach number of 0.85, a rearward movementof the aerodynamic center at a Ikch number of 0.87, and a shift in theangle of attack for zero downwash at a Mach number of 1.2. Also, atlow lift coefficients, an increase of lift-curve slope and a rearwardshift of the aeroeic center with increasing lift coefficient were -indicated at Mach numbers below 0.875. At high angles of attack, thewake 1.225 semispans behind the 25-percent mean-aer@namic-chofi stationextended at least 0.375 semispan above the wing-chord plane.

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NACA RM L51A02

INTRODUCTIOIJ -

A general research program is being conducted by tle NationalAdvisory Committeefor Aeronautics to supply :designersoftransmit ..,.aircraft with needed information on the effetitof vari&us wing-geometryparameters cm aerodynamic characteristics at.transonic.speeds.

This paper presents the results of..testson a.sting-supportedwi.ng-fuselage combination employing the unsweptwfng of a s~ries of @.ngshaving varying amounts af sweep, aspect rati_q4, taper-ratio 0.6, and an .-NACA 65Am6 airfoil section. Tests on otherwingsin thisseries are

,, 2-,snd3. :-_: _ :- .::. _..,;reported in references 1

Lift, drag, and pitching-moment data are presented for subsonicI&chnumbers from 0.60 to 0.93 sad for a sup&sonic Mach number of 1.2.

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Also presented dre point .downwashdata emd W@e 10SSeS_f.Orsever~.@i> ....heights at two spsnwise locations. :. -.

The data presented herein and in references 1, 2; and 3 are cornpaed--with data of geometrically similar configurations obtained by means ofthe transonic-bump method in the Langley 7-_by 10-fOOtiU~el, (seereference 4). 4

SYMBOLS ..

drag coefficient()

DQS

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lift coefficient L ,..-..g

()

%/4pitching-moment coefficient referred to 0.25t! -. . qsz!

mean aero@amic chord, inches

drag, pounds --

loss of total pressure in wake, pou,ndsper square foot _

p~ssure difference between upper,and lower ~omponents of .,-a yaw tube

lift, pounds.

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Mach number , . F-

.NACA RM L51A02

m/4 pitching moment about 25 percent E, inch-poundsD

()

P~ - P.P~ base-pressure coefficient

q

P. free-stresm static pressure, pounds per sqyare foot

% static p~essure at model base, pounds per square foot

q free-stream dynamic pressure, pounds per square foot()1 ~2:P

.qL local dynmnic pressure at any yaw tube

R Reynolds number based on 6

3

.

s wing area of model, square feet

v free-stresm velocity, feet per second

a angle of attack of fuselage center line, degrees .

E downwash angle, degrees

P free-stream density, slugs per cubic foot

APPARATLJSAND METHODS

Tunnel.

The investigation was conducted in the Langley 8-foot high-speed Itunnel, utilizing the plaster-lined nozzle described in reference 5.Subsonic tests were conducted with the model located in the region ofthe minimum section of the nozzle. For testing at a Mach number of 1.2,the model was moved downstream to the expanded section of the nozzle.

The minimum section of the nozzle had a constant Mach numberdistribution up to the highest point tested. In the supersonic sectionof the nozzle, the maximum Mach nmiber variation was 0.02.

Model and Support. .

. The model tested had a wing with 0 sweepback of the quarter-chordline, zero twist and dihedral, aspect ratio of k, taper ratio of 0.6, andan NACA 65AO06 airfoil section measured parallel to the model plane of.

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symmetry. The wing was machined -from14STaluminum alloy, and wasmounted on a fuselage body of revolution of fineness r~tio10.0. The --

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longitudinal position of the wing was.sucht~at the quarter-chord pointof the mean aerodynamic chord coincidet.withthe stati~n of maximum

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bcdy diameter, principal wing dimensions ar> presented with a,plfi-form -- =.drawing of the wing-fuselage combination in figure 1. The fuselageordinates and dimensions sre presented in fi&re 2. An electricalstrain-gage balance was contained within the:fuselage &nd,secured to

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the fuselage at the forward end. The rear part of the,balance comprised a sting for suppoti.ingthe model in the center of the tunnel.(referenCe 1). ~ .The sting washinged to a suppoti tube in su~h a manner that the angleof attack could bevaried by means of a remote-cortrolmechanism whiletesting was in progress. This sting-support,tubecould be made to slide

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axially on its mounts in order to move,the model fron.the subsonic testsection to the supersonic test section. Figure 3 is adiagram of this

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setup and figure & gives a general view Qf tbe model, sting, stingsupport, md.test ~ection.

Measurements

Lift, drag, and pitching-mcment datainternal-strain-gage-balticesystem. The

.weqe obtained by using the .,,, --t=setisitivityor the strain- -

gage .balanceand the scatter of-test points indicated that the accuracyof lift, drag, and pitching-moment coefficients was within +0.01, +0.001,and.+0.005, respectively, for all Mach numbeis. Downw6sh andwake-intensity measurements were made with two rakes having both yaw-pressmeand totsl-pressure tube%. The location.and geometry of these rakes aresljownin figure j. Ai@es of attack of t_he_n@el m.d of therake.were _.determined within O.1 by means of an optical system.utilizingRarallellight beamB. A description.of this device cs+nbe.foufi in reference 1.A photograph of the model and r~e setup is Shorn in fI_gure6. Thestatic pressure at the base ofthemodel wasietermined by means of astatic orifice located on the side of.the stfig in the plane of the modelbase. .-.

The yaw tubes were calibrated in.the emptytest section by measuring .

the variation of ~,-

with rake angle of-att~~ck.D~wash ~gles W&e:. .

determined with thiscalibration from the ~s measured during the -.

tests. During these measurements the static~ressure @ the -e W&-S-assumed to tieequal to the free-stream static pressure. Where the wakewas large, this is a possible source of erroti,however.,consideration ofpossible small errors in calibration, angle-of-attack measurements,scatter of tests points, and variations in local static pressure indicatedthe accuracy of the measured downwash an-gles~o be wittin +0,20 for -measurements made outside the wake md within -+0.30formeasuremetitsmadein the wake.

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Test Conditions

Data were taken at angles of attack of .20 through 14 at Machnumbers of 0.60 to 0.93and at a Mach number of 1.2 for the wing-fuselagecombination. Also, the wing-fuselage combinatioriwas tested at variousMach numbers in the above range with transition from laminar to turbulentboundary layer fixed at the 10-percent-chord line on the upper and lowersurfaces of the wing and 4 inches back of the nose on the fuselage bymeans of number 60 Carborundum grains doped to the model surface at thesepositions.

The variation of mean test Reynolds number with Mach number is shownin figure 7. Variations in atmospheric conditions caused deviationsfrom this curve, but, for any given Mach number the Reymolds number didnot differ from that shown by more than 3.5 percent.

By means of pressure measurements obtained froma series of static-pressure wall orifices, choking tendencies were obsezwed in the tunnel.No data are presented in this report where these tendencies wereevidenced.

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At a Mach number of 1.2, the location of the normal shock wasascertained by means of a portable point light source. In none of thetests for which data ere presented did the normal shock advance upstreamahead of the rakes or to the base of the model. This condition has beenshown in reference 6 to be the criteria for effects of the normal shockon the model.

CORRECTIONS

Corrections due to tunnel-induced upwash and due to model and wakeblockage and pressure gradient due to wake were calculated and appliedto the-data by using the methods of references 7, 8, and 9. The correc-tions to the dynamic pressure and to the Mach number were found to benegligible below a Mach number of 0.85 and reached a maximum of 1.4 per-cent at a Mach number of 0.93. The,msximum correction to the downwashangle at the rakes amounted to 0.2.

Base-pressurecoefficients were obtained and are presented in fig-ure 8. Comparison of these base pressures with the base pres~res from _reference 1 for the fuselage-alone configuration indicated that theaddition of the wing lowered the base-pressure coefficient by an amountapproaching 100 percent at the higher angles of attack.

No tsre corrections have been applied except in the case of stingtare; for this case, the corrections were applied tc the mexinum

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NACA RM L~lA02

lift-to-drag ratio and to drag at zero lift.._~The results of the Investi-gation of a stiilar model at low angles of at%ack (reference 10) indicatedthat this tare need not be applied to lift o~~moment values but would bean increment of 0.003 to be add@d to the dragcoefficient at all subsonicMach numbers and 0.002 at a Mach number of 1.2. As the test setup inreference 10 involved a sting-supper-l.systemsimilar tothe support .system in the present test, corresponding sti~ corrections in the twotests are assumed to be of the sameorder of kgnitude. It is alsoestimated that the sting may cause the downw5& angles to b.edecreased = ..:by as much as 1 at subsonic Mach numbers md;O.1 at a Mach numberof 1.2. In addition, the base-pressure coefficients may be increased

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by the presence of the qting by-approximateti:O.l at ali Wch numbers...

Inatimuchas these corrections were estimated-byusing data from refer-ence 10, which only consider low angles of attack, no attempt has been _made to apply them to the data except in the aforementbed cases.

Corrections in the angle of attack due to the sweep of the center-._ ,of-bending li~e and due to a pitching m~ent=in the wing were calculatedand found to be of negligible value.

RESULTS AND DISCUSSI& =

In this investigation,a wing-fuselage crjmbinationwas tested asthe basic configuration. Fuselage-alonedata.were subtracted from thedata of the Wing-$uselagecombination and the.resulting data are thewing-with-wing-fuselage-interferencedata, Basic data for the fuselagealone may be found in reference 1.

All of the following discussion pertains to transition-naturaldata unless otherwise stated.

A table of figures presenting the results follows:

Figure

Wing-fiselage force data againstMach nuM]erL; . . . . I . . . .Wing-fhselage force data against ltft coefficient , . j . . . .Wing-with-wing-fuselage-interferenceforce ~$a

agatnst lift coefficient . . . . . . .. . . .. . . . ..~ . . . .Lift-curve slope against Mach number . . . . ~. i . , -~ . . . .Zero-lift drag coefficient against Mach numbe~ . . . . ;-. . . .Maximum lift-drag ratio against Mach number . . . . . . . . . .Static-longitudinal-stabilityparameter against. . .Mach number . . . . . . . . . . . .+ . .:.. . . . . . . . .

Wske losses (wing-fuselage) . . . . . . . . ,.. . . . ~ . . . .Point downwash data against .gmgleof attack .: . . . . 0 . . . .Average rate of change of downwash angle withangle of attack against Mach number . .

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NACA RM L51OA2 7

Force and Moment Characteristics

The decrease in lift coefficient for the wing-fuselage configurationoccurred at a Mach number of approximately 0.90 for angles up to 4(fig. 9(a)). Fixing transition caused a reduction in lift coefficientat the high Mach muibers for angles of attack up to 10.

At a Mach nuuiberof 0.60, the lift-curve slope for the wing-fiselageconfiguration was 0.068.at zero lift (fig. 12). The slope then increasedto ama.ximum of O.l@L at a Mach numiberof 0.90. At a Mach nunber of 1.2,the slope was 0.08. The lift-curve slope at a lift coefficient of 0.4had the same trends. The magnitude, however, was approximately 12 percenthigher at a Mach humber of 0.60 and reached its peak value of 0.107 ata Mach number of 0.85. At a Mach number of 1.2, the lift-curve slopeat a lift coefficient of 0.4 was 11 percent lower than the lift-curveslope at zero lift coefficient. It is suspected that the increase inlift-curve slope with increasing lift coefficient at the lower speedswas associated with a separation bubble s~lar to that described inreference 11. This flow characteristic, however, decreases withincreasing Reynolds nmnber and would probably disappear at a Reynoldsnumber of approxhnately 10 million. A decrease in lift-curve slope atthe high Iif% coefficients occurs when the separated region extendsover a large part of the chord.

...

Subtracting the fuselage data frmthe wing-fuselage data hadlittle effect on the lift-curve slopes except at allifi coefficientof 0.4 where the.peak was reduced by 5 percent.

At low lift coefficients, sn abrupt drag rise occurred at a Machnumber of approximately 0.875 (fig. 9(b)). In general, fixing transitionresulted in an increase in drag coefficient, but in no instance did the transition-fixed data show any marked drag changes over the natural-transition data.,

At zero iii%, the drag coefficient for the wing-fuselage combinationremained constant at approximately 0.009 up to a Mach number of 0.875,where it increased sharply (fig. 13). At a Mach number of 1.2, thisdrag coefficient was 0.038. The ssme trends for zero lift drag wereexhibited by the wing-with-wing-fuselage-interferencedata. However,the absolute values of this coefficient were approximately ~ percentlower at subsonic speeds and approximately 26 percent lower at a Machnumber of 1.2 than the corresponding values for the wing-tiselageconfiguration.

A msximum lift-to-drag ratio of approximately 14.5 for the wing-. fuselage combination was maintainedup to a Mach number of 0.85 (fig. 14).

Above this Mach number, a rapid decrease in lift-to-drag ratio wascaused by an abrupt drag rise. At a Mach number of 1.2, this ratio

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was 5.5. The lift-to-drag ratio for the wirig-with-wi&-fuselage-interference data was approximately 73 percent higher than .thatfor thewing-fuselage configurationup toa Mqch number-of 0.85, above whichthe increase was much less pronounced. At a.Mach number of 1.2, thisincrease w~s only 22.percent.

The large favorable chsmges in drag at zero lift &nd in maxi- ---lift-to-drag ratios resulting frcvnsubtracttig the fus~lage data fromthe wing-fuselage data are largely due-to the fact that the drag of the psrt of the wing covered.by the fuselage does.notappear In the resultingconfiguration, although the coefficients are based on thetotal areaofthe wing.

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The wing-fusehige cotiiguration exhibited an abrupt decrease inpitching-moment coefficient beginning at a Mach number .of0.86 for 20 angle of attack and at lower Mach numbers for higher angles of attack.Fixing transition reduced the abruptness and inagnitudeof this variation.

acmAt zero lift, the longitudinal stability parajneter acL remained at aconstant value of 0.15 up to a &chntiber 0~0.85. (fig. 15). Above0.85 the aerodynamic center begti to move reaiward, th& mdel. becomingneutrally.stable at a Mach number of 0,905. At a Mach number of 1.2,the moment-curve slope indicated stability, being -0.@. At a liftcoefficient of 0.4, the aerodynamic center a-tsubsonicspeedswas apprOXi--mately5 percent rearward of the location for zero lift. ~s rearwardshift with increasing angle of attack @ be associated with the leading- -edge separat.i~npreviously mentioned. For.a-lift coefficient of 0.4,the aerodynamic center began a rapid rearward.movmentat a Mach numberof 0.80. This rearward mov@ment resulted in the wing-fuselage configu-ration becoming stable above a Mach number of 0.85 wher=ethe value of*m

Greached -0.08. At a Mach number of 1.2, the slope of themoment

curve was essentially the same for a lift coefficient of 0.4 as it wasfor zero lift. ..

For both lift coefficients, subtracting%he fuselage-alone momentmoved the aerodynamic center rearward approxi~tely 7 percent of themean aerodynamic chord at most speeds.

Wake and DownwashCharacteristics

At Mach nunibers of oj6 and 0.8, the wake losses for @ and 4were negligible in the region investigated, but, at a kch number of 0,93,the wake for 4 had begun to appear; thus a shock-,inducedseparationwas indicated (fig. 16). At a Mach number of.1.2, the shock had movedto the trailing edge, and wake losses were again small for 0 and 4.

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NACA RM L51A02 9

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The wske for 10 was assuned to extend at least 0.375 semispan abovethe fuselage at all speeds tested. The wake losses at the inboardstation were larger thsn those of the outboard station because of losses

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due to the fuselage.

A significant chsnge in the angle of attack for zero downwashoccurred at a Mach number of 1.2 (fig. 17) for the wing-fuselage combi- -nation as compared with this configuration at a Mach number of 0.93. Asa consequence, significant chsnges in trim of an airplane flying to aMach number of 1.2 canbe expected from this shift if the horizontal ..tail is located within the region investigated. This shift in the angleof attack for zero downwash is attributed to the fuselage inasmuch aswing-with-wing-fuselage-interferencedata did not indicate a similarchange.

The average rate of change of downwash angle with angle of attackis presented against Mach number in figure 18. ~ese ~~ues were found ..._by averaging the slopes for the two semispanstations at a location

0.375 semispsn above the wing-chord plane. This average & for theaa ,wing-fuselage combination at a Mach number of 0.80 was approximately0.5 and 0.6 for lift coefficients of O and 0.4, respectively. Wing-with-wing-fuselage-interference data below a Mach number of 0.80 exhibited

ae for a lifta zero lift%

approxtiately X2 percent lower and a h

coefficient of 0.4 approximately 15 percent higher thsn corresponding

values for the wing-fuselage configuration.a~Variation of with Mach&

number was erratic above a Mach number of 0.80.

CONCLUSIONS

The results of an investigation of a wing-fuselage combination employing a wing with unswept quarter-chord line, aspect ratio k, taperratio 0.6, and an NACA 65A06 airfoil section at high subsonic Machnumbers and at a l&ch number of 1.2 indicated the following:

1. Increasing the free stream Mach number at low lift coefficientscaused the wing-fuselage configuration,to exhibit a decrease in lift-curve slope at a Mach nuniberof 0.90, a rapid decrease in the maximumlift-to-drag ratio at-a Mach number of 0.85, snd a rearward movement

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of the aerodynamic center at a Mach number of 0.87.

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2. An increase in lift-curve slope and rearwsrd shift of aer@nsmiccenter with increasing angle of attack was indicated at low Mach numbers.

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3. me wake 1.225chord station extended

NACA RM L51A02

semispans behind the !25-percentmesn-aerodynamic-to at least 0.375 se~span above the wing-chord .

plane at high angles of attack.-.

4. At a Mach number of 1.2, the angle of attack for zero downwashwas changed by the presence of the fi.sel,age;thus significantchangesin trim going from a Mach number of 0.93 to.a Mach n~%er of 1.2 Occurred..-. -. _

Langley Aeronautical Laborato~-....

National Advisory Committee for Aerorlau~icsLangley Field, Va.

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NACA RM L51,A02 11

REFERENCES

1.

2.

3*

4..

5.

6.

7.

8.

9*

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Osborne, Robert S.: A Transonic-Wing Investigation in the Langley8-Foot High-Speed Tunnel at High Subsonic Mach Numbers and at aMach Number of 1.2. Wing-Fbselage Configuration Having a Wingof 45 Sweepback, Aspect Ratio 4.0, Taper Ratio 0.6, and NACA65A06 Airfoil Section. NACA RML50H08, 1950.

Henry, Beverly Z., Jr.: A Transonic-Wing Investigationin the Langley8-Foot High-Speed Tunnel at High Subsonic Mach Numbers and at aMach Number of 1.2. Wing-Fbselage Configuration Having a Wingof 35 Sweepback, Aspect Ratio 4.0, Taper Ratio 0.6, and NACA65AO06 Airfoil Section. NAcARML50J09, lg50.

Wocd, Raymond B. and Fleming, Frank F.: A Transonic-Wing Investigationin the Langley 8-Foot High-S~eed Tunnel at High Subsonic Mach Numbersand at a Mach Number of 1.2. Wing-Fuselage Configuration Having aWing Of 600 Sweepback, Aspect Ratio 4.0, Taper Ratio 0.6, and NACA65Ao06 AirfoLL Section. NACA RM L50J25, 1951.

Donlan, Charles J., Myers, Boyd C., II, and Mattson, Axel T.: AComparison of the Aerodynamic Characteristics at Transonic Speedsof Four Wing-l@elage Configurations as Determined from DifferentTest Techniques. NACA RML50H02, 1950.

Ritchie, Virgil S., Wright, Ray H., and Tulin, Marshall P.: h 8-FootAxisymmetrical Fixed Nozzle for Subsonic Mach Numibersup to 0.99snd for a SupersonicMach Nixnberof 1.2. NACARML50A03a, 1950.

Ritchie, Virgil,S.: Effects of Certain Flow NonUniformities on Lift,Drag, and Pitching Moment for a Trsnsonic Airplane Model Investi-gated at a Mach Number of 1.2 in a Nozzle of Circulsr Cross Section.NACARML9E20a, 1949.

Eisenstadt, ~rtram J.: Boundary-Induced Upwash for Yawed and Swept-back Wings in Closed Circular Wind Tunnels. NACATN 1265, 1947.

Goldstein, S., and Young, A. D.: The Linear Perturbation Theory ofCompressible Flow, with Applications to Wind-Tunnel Interference.R. &M. No. 1903, A.R.C., 1943.

Herriot, John G.: Blockage Corrections for Three-Dimensional-FlowClosed-Throat Wind Tunnels, with Consideration of the Effect ofCompressibility. NACA RMA7B28, 1947.

12 aEmz!z?!2!z?~ NACA RM L51A02

10. Osborne, Robert S.: High-Speed Wind-~el Investigation of the

LongitudinalStabilitysnd Control Characteristics of a ~Scale.

16Model of the D-558.2 Research Airplaneat High %bsonic MachNumbers and at a Mach Number of 1:2.

11. Nuber, Robert J., and Gottlieb, StsmleyTunnel Investigation at High ReynoldsAirfoil with High-Lift Devices. NACA

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II?ACAtiL9COk, 1949.

M. Two-D-imensionalWind-.

=bers,ofsn NACA 65AOC6~L~06, 194-8.

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WING DIMENSIONS

Airfoil section (parallel to modelplane of symnetry - NACA 65AO06

Area, sq f% ......................Aspect ratio .....................Taper ratio ......................Sweep angle, deg (25-percent

dord line .....................Incidence, deg ...................Dihedral, deg .......... ........Geometric twist, deg .............

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F = 6./25--+

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25 Ctbd -

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c 33.33

Figure 1.- Plan W@ Of model giving over-all dimensions. KU dtiensionssre in inches.

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ORDINATEX3 ~

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.0500 o~50.1000.l~oo,2000.2500.3000.3500.4000

I .00231.00298.00428.00722.01205.01613.01971.02593 03090.03465.03741.03933.04063

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L.E. radius = o,ooo5j

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Fineness ratio 10 WV ~~ ;- -,.2 -.5/4 located at D( max. )

Figure 2.- Fuselage details. All.dimensions are an inches. .-.

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0 0-. . . :, s.. ,. ., n...

P/aster liner

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Gi5wnetric minimum ~

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Figure 3.- Location of mcdel in relationto Langley 8-foottunnel teat section and support system.

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Figure 4.- General view of test

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setupshtiing model, r~es, and supportsystem .

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NACA RM L51A02 .

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.

.

- 9/

#odczl plum of

~$I I

Tok71-przssure fubas

Wing- chord plum

1Vode/ 3Ting - =5=

Figure 5.- Details of the rakes used for wake survey and downwashmeasurement. All dimensions are in inches.

.

.

NACA IIML51A02.

.

.

.

.

.- .

- --.-~i-.= -..... -- 1.-63781 ;i. -. .Figure 6.-

.Photograph of model as tested in. the La@ley 8-foot high-

speed tunnel..

...

-..

c * ,

02

4

a

2.2

2.0

1.8

1.6

1.4

6Xlo

Fs

.5 .6 .7 .8 .9 Lo 1.1 1.2

Mach number, M

Figure ~. - Variation of test Reynolds number baaed on a E of 6.w19 inches,with Mach number.

G

20 NACA RM L51A02

Transition natural

.4

,2

0

.

0

0

0

o

0

0.

Transition fixed

5 ,6 7 ,0 .9 1.0 I;i I*2Mach number, M . .

.

(d:g)

14 Xl

12 v

10 4

6V

4A

20

. .

on

-2 0

.

.

.

Fi~re 8,- Base-pressure-coefficient.vtii.ationwith Mach number. - --Unflagged symbols indicate transition natural. r-

----

NACA RM L51A02 ,

Transition natural

.-w-Q)00

. . rar sitlon fixed

T\

\ ., v

h+ w+

I

0 I0

0

0-

0 1 I II 1

{

1 , ,

I1 , ,

Io~ I II I I I I I I I I I I I I0

. I1 I I I 1

0

-9

21 .

(d%g)

14 w

12V

104

8

6

4

2

0

-2

.-.5 .6 ,7 ,8 .9 1.0 1.1 1.2

Mach number, M

(a) Lift coefficient.

Figure 9.- Variation with Mach number of the aerodynamic coefficientsthe wing-fuselage configuration with transition natural and with

P

v

A

o

o

of

transition fixed on the 10-percent-chord line of the wing. Wflawed .symbols indicate trsmsition natural.

22 NACA RM L51A02

Trarjsition. natural.

,24 Transition f i x e d

7

.22 T*. L

.20

.184 H

,

b +.16 >

.14

c!! >1

- .12E/ h

o.-0.-~w~ .10

m: 08

fL

i* d

/

.06 ii

#/

/

.04 F7A

/ rf

:02 L --El

Ar-

)./

0 .5 .6 .7 .8 .9 1.0 1.1 1.2

Mach number, M

(b) Drag coefficient.

.

.

12V

IOa

.

6V

4A

20

00

-20

.

.

Figure 9.- Continued, ... .- .-

-MmJJgx@7... .._

T

NACA RM L51A02 23.

.

..5 .6 .7 .8 .9 1.0 1.1 1.2

Mach number, M

(C) pitching-moment coefficient.

.

(d:g)

14W

I 2V

8D

6V

4A

20

on

-20

Fj_gure 9.- Concluded. .

-.

.

,,

,,I

1,

,

I14

12

10

~8

J

!6G

%4

?

2

0

-2

:2 0 0 0 0 0 0 0 0 .2 ~ .6 .8 [-o

Lift coefficient,L

(a) Angle of attack.

Figure 10. - Variation of the aerodynamic coefflclentm w~th liftcoefficient for the wing-fuselage ccmflguratlon.

, * ,,,. ,, I 1, p

Nw

11!i1

w,,, .

1 .1

z%>.

z;;

z=

IN1

!1

,.

!1 i., I

v * * , .,*

.24

.22

.20

.18

.16

n

.14

E.:G .12g

.10z&

.08

D6

.04

02

0

+ o 0 0 c1 o 0 0 024s.8Lift roafflclsnt,

CL

(b) Drag coefficient. G

Figure 10.- Continued.

-,

11 I

m

.04

E o

,:,

0

:04

I

/ Y .

/ [ .,. , I .,m-i-t--!

I.60

.70

.80

1

,85

.2 O .2 .4 ,6 .8 Lo

Lift coefficient, C,L

M

o .875

0 90

0 .93

0 1.2

:2 0 .2 ,4 .6 .8 1.0

Liit coefficient,\

(c) Pitching-mane.nt coefficient.

Figure 10. - Concluded.

, 8I I ,,, :: 1, ,,

i . .:., :,,

. * , on with M_achnunibercorrected for sting interference. -

L5M02-

,

0

.n m

o Wing-fuselage

.2

.2

c1cm

q 0

d Wing with wing-fuselageinterference

/

.5 .6. .7 .8 .9 1,0Mach number, M

Figure 15. - Static* longitudinal stability-parameter

Mach number.

1,1

variation with

1.2

uu

34 -cong@@E33w

,6

AH .4q

.2

0

(d:g)-o 4.-- &

.8

,6

AHq ,4

0

,,

NACA

.,.

-.

RM L51A02

(a) Locat$on 0.083 semispan from plane of =~etry.

..-

7 .-,

--

-fEl$Eii.-

,4. M= O.80

/

AH .

T ,2\ \

\\

o

IHEEl ----EEEE.1 .2 .3 .4 ,1 .2 .3 .4

Height obove wing-chord plune , semiepon

(b) Location 0.292 semispah from plane of symmetry..-

.

Figure 16. - l%ke losses 1.225 semispantibehind the 0.25~ location for thewing-fuselage configurations.

,-

...--. . ---

.

-,.

.. _ .-