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,.’ c~ d “ ‘-- copy. RM L53D21.

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RESEARCH MEMORANDUII@F

MEASUREMENTS OF AERODYNAMIC CHARACTERISTICS AT TRANSONIC

SPEEDS OF AN UNSWEPT AND UNTAPERED NACA 65-009 AIRFOIL

MODEL OF ASPECT RATIO 3 WITH l/4-CHORD PLAIN

FLAP BY THE NACA WING-FLOW METHOD

●

By H&old I. Johnson

Langley Aeronautical LaboratoryLangley Field, Va.

ofb ●spiomeskm, Tltls18,v.8.C.,&c& TM—rbmlwmtblrizoip?aon IsprOhlMtOdwlw. ‘

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

. WASHINGTON

June 10, 1953

c~

TECH LIBRARYKAFB,NM

R*

NACA RM L53D21

NATIONAL ADVISORY COMMITTEE FOR AERONAWICS

Ii[lllllllllllllllllllllllllllllC11444115

RESEARCH MEMORANDUM

MEASUREMENTS OF AERODYNAMIC CHARACTERISTICS AT -SONIC

SPEEDS OF AN UNSWEPT AND UN’I’APEREDNACA 6HKW AIRFOIL

MODEL OF ASPECT RATIO 3 WITH l/4-CHORD PLAIN

FLAP BY THE NACA WING-FLOW METHOD

By Harold 1. Johnson

SUMMARY

A wing-flow investigateion was made to determine the lift, pitching-moment, and hinge-moment characteristics of an unswept ariduntaperedNACA 65-oo9 airfoil model of aspect ratio 3.01 equipped with a l/h-chordfull-span plain flap. The Mach number range was approximately 0.65 to1.10 and the corresponding Reynolds nuniberrange was approximately0.5 X106 to 0.9 Xlo 6. The effects of sealing 69 percent of the lengthof the l.1-percent-chord flap gap were investigated as were the effectson flap characteristics of adding roughness to the first 5 percent ofthe airfoil chord.

The maximum unstalled lift coefficient of the model was found tobe almost twice as great above M = 1.0 as it was below M = o.90. Acompressibility phenomenon that apparently is peculiar to fairly thickaerodynamic surfaces was found to occur near M = 0.95 at small anglesof attack and flap deflections. This phenomenon was made manifest by alarge reduction in lift-curve slope, an ab~Pt forward movement of theaerodynamic center to a position near the leading edgej ~ abrupt reversal-to a strong positive variation of hinge-moment coefficient with angle ofattack, and a reduction of flap effectiveness to aPProx*te~ zero forsmall deflections. Below M = 0.90 the hinge moments due to deflectionwith gap sealed were about equal to what would be predicted from thin-airfoil theory and above M = 1.0 the hinge moments were approximatelywhat would be expected from the concepts of two-dimensional linear super-sonic theory. The hinge-moment variations with angle of attack werevery nonlinear at subsonic speeds because of gap effects but were fairlylinear and strongly negative at supersonic speeds. The effects of sealingthe flap gap at subsonic speeds were to increase flap effectiveness,reduce the hinge moments due to deflection> and -e more linear thevariations of hinge moment with qngle of attack. At supersonic speedsthe aerodynamic characteristics were nearly the same with gap eithersealed or open.

2 NACA RM L5n21

INTRODUCTION

A wing-flow investigation was made to determine the lift, pitching-moment, and hinge-moment characteristics of an unswept and untaperedNACA 65-009 airfoil model of aspect ratio 3.01 equippGd with a l/4-chordfull-span plain flap. This investigation is closely related to thosereported in references 1 to 4 which dealt with an equivalent 35° sweptbackmodel on which full-span flaps having different kinds of aerodynamicbalance were investigated at transonic speeds. —

By present-day standards, a 9-percent-thick aerodynamic surface ofaspect ratio 3 would be considered excessively thick for most applica-tions. Like most wing-flow or tunnel-bump experiments, the Reynolds

numbers were low in the present tests (less than 0.9 x 106); in spiteof these limitations, the data are thought to be of appreciable interest.In particular, the variation of maximum lift with Mach number, the hlnge-moment measurements, and the effects of flap gap at trsmsonic speeds maybe of special interest.

%

Ch

SYMBOLS

average Mach number over model

average dynsmic pressure over model

-.

i?

total

model

model

model

model

model area

lift coefficient,Model lift

qs

chord--

mean aerodynamic chord

pitching-moment coefficient measured about axis at39.5 ??=ent mean =r@m~C chord, -Model pitching moment about 0.395E

qs~

area moment of flap about hinge line .

Model hinge momentmodel hinge-moment coefficient, =2@4.f d

.

NACA RM L53D21 3

* a

8.

c%!

% o. 395~

angle of attack ‘

flap deflection

variation of model

acLper degree, ~

variation of model

ac~per degree, —

a6

variation of model

lift coefficient with angle of attack

lift coefficient with flap deflection

pitching-moment

with angle of attack per degree,

variation of model pitching-moment

with flap deflection per degree,

coefficient about 0.397a%H~ 0.39’55

coefficient about 0.395~a%Ha~ 0.395E

variation of flap hinge-mm.ent coefficient with angle ofa%

attack per degree, —&

variation of flap hinge-moment coefficient with flap deflec-aw

tion per degree, —ab

flap relative

aspect ratio

ac~labeffectiveness, —

acklb

included trailing-edge

AEPARATUS

U(

angle of flap, deg (q = 6°)

AND TESTS

The semispan wing-flow model simulated a wing or tail surface ofaspect ratio 3.01, taper ratio 1.0, and sweepback singleof OO. The

b model was machined from solid berylliw-copper to the contour of theNACA 65-oo9 section and incorporated a l/4-chord plain flap mounted ontwo hinges. The model had a 0.040-inch-thick end plate with a diameter

.

equal to the chord affixed to its root in order that proper semispantesting conditions would be more nearly realized. A photograph of themodel with end plate attached is given in figure 1, arida drawing,including principal dimensions, is given in figure 2. The model (fig. 1)had two flush removable plates between the hinges to povide for installa.tion of a thin sheet-rubber gap seal. The length of gap sealed was69 percent of the hinge-line length for the gap-sealed condition and,for the gap-open condition, the gap width was 1.1 percknt of the airfoilchord. It should be noted that the gap was unusually large. The modelwas mounted on a strain-gage balance located inside the wing of a NorthAmerican F-51D wing-flow airplane in such a way that the pitching momentswere measured about an sxis at ‘39.5percent of the model mean aerodynamicchord.

Measurements were made of the lift, pitching moment, and hingemoment for an angle-of-attack rsnge from about -5° to 30°, a flap-deflection range of about -120 to 22°, and a Mach number range frcrnabout 0.65 to 1.10. The measurements of maximum lift were limited toM = 1.05 for reasons to be discussed subsequently. The approxw&Reynolds numbers existing during the tests are shown as .afunction ofMach number in figure 3. Some tests were made with a layer of 0.003- to0.005-inch Carborundum particles sffixed to the first S percent chord onboth upper and lower surfaces of the model. These roughness tests weremade only for the case of vsriable flap angle with the model set for0° angle of attack. No corrections were made for the effects of aero-elasticity in view of the extreme ruggedness of the model and the rela-tively low dynamic pressures encountered at the test altitude range offrom approximately 30,000 feet to 18,OOO feet. Further details concerninginstrumentation, test technique, and probable accuracies can be found inreferences 1 to 4. —

with

RESUI!TSAND DISCUSSION

Characteristics in Angle of Attack

The variations of lift, pitching-moment, end hinge-moment coefficients@e of attack at 0° flap deflection are shown in figure 4 for

increments in Mach number of 0.05 over the speed r-e-tested.

Perhaps the most striking feature shown by the lift measurements(fig. k(a)) is the exlmemely large increase in maximum lift coefficientthat occurred just prior to the attainment of sonic velocity. The valuesof C~ and of angle of attack, read either at the peaks in the lift

curves or slightly beyond the occurrence of an abrupt decrease in lift-curve slope (in cases where no definite peak existed), are plotted

NACA RM L53D21 5

against Mach number in figure 5. Comparison of these data with thoseof references 1 and 4 indicate that 35° sweptback models tested underthe same conditions gave higher msximum lift coefficients at Mach numbersbelow 0.97 and lower meximum lift coefficients above this Mach number.The low msxtium lift coefficients found in the subsonic speed range arebelieved to be due largely to the low Reynolds nu.nibersas well as to thelow aspect ratio and relatively small leading-edge radius of the65-009 airfoil section; however, it is unlikely that the large increasein maximum lift with increasing Mach number would be eliminated by anincrease in Reynolds number. Above a Mach number of 1.05 it becsmeimpossible to measure maximum lift coefficients inasmuch as the pressuresset up by the wing-flow model caused the flow field about the right wingof the wing-flow airplane to change radicaUy in an abrupt manner. Whenthis happened, the airplane was subjected to a rather violent rollingoscillation which had a frequency exactly twice that of the forcedoscillations of the wing-flow model. W&never the model reached eitherhigh positive or negative angles of attack, the model lift trace showeda sharp discontinuity and the airplsne accelerometer showed losses innormal acceleration of about lg during a 4g pull-out which representedlosses in airplane lift of the order of 8,cK)0pounds. These losses inlift coincided with the occurrence of right-wing heaviness; therefore,the model constituted an extremely effective spoiler at airplane Machnumbers approaching the maximum permissible (M = 0.75).” This phenomenonapparently establishes a limit on the ranges for which techniques suchas the wing-flow method can be used to investigate msximm lift. Thephenomenon also reemphasizes the possible injurious effects of smallprotuberances at transonic speeds.

The effect of the flap gap was to decrease the maximum lift coef-ficient by a small amount over the entire speed range. The effect ofthe gap on the lift-curve slope was very small snd inconsistent overthe speed range.

The only other lift characteristic requiring comment occurred overa small angle-of-attack range at a % 0° at M = 0.95 where the lift-curve slope suffered a decrease. Although not particularly significantin itself, as will be shown later, this very minor change in lift wasaccompanied by violent changes in hinge-moment characteristics andaerodynamic-center location.

The pitching-moment curves (fig. 4(b)) require little comment. Ingeneral, the model showed reasonably constant stability up to the initislstall and at higher angles of attack became more stable. As evidencedby the near-zero values of pitching-moment coefficient at extreme anglesof attack, the center of pressure at these singlesof attack was in theneighborhood of 40 percent mean aerodynamic chord or somewhat fartherback at Mach numbers above 0.85. Except for the sharp decrease in

6 NACA RM L53D21

stability at M = 0.95 near a = 0° which will be discussed later, #

the model becsme more stable at small angles of attack as the Mch numberwas increased from subsonic to supersonic va@es. Th~latter trend isj_of course, to be expected.

-.

The hinge-moment measurements (fig. 4(c)) show several interestingpoints. As mentioned previously, a violent change in hinge-moment char-acteristics occurred in the neighborhood of M = 0.95 at a = OO. Theonset and disappearance of this change is documented in figure 6 whichshows the hinge-moment characteristics for small increments in Machnumber between 0.90 and 1.0. The reversal in hinge moment is seen tobe most severe between Mach numbers of 0.96 end 0.97. Inasmuch as thetotal lift on the model was only slightly effected, it may be concludedthat nearly all the abrupt change in flow characteristics occurred nearthe trailing edge of the model. Further support for this conclusion wasgiven by the pitching-moment and flap-effectiveness measurements, respec-tively, which, as will be shown subsequently, indicated that the aero-dynamic center moved rapidly forward to a position near the leading edge

—

of the model and the flap effectiveness for-small deflections becameessentially zero at the ssme Mach numbers that the flap-floating tendencyreversed from with the wind to against the wind.

The flow phenomenon which caused all the foregoing undesirablecharacteristics appears to be the same as that found by several other J

investigators (for exemple, refs. 5 end 6). G&hert (ref. 5) @veS areasonable,explanationof.the phenomenon based on pressure-distributionmeasurements and Hemenover and Grshsm (ref. 6) show schlieren photographs

u

of the flow that substantiate the remaining necessary assumptions madeby Gdthert. The mechanism of the flow phenomenonmsy be described brieflyas follows:

Consider a symmetrical airfoil at zero angle of attack having al/4-chord flap at 0° deflection in a stream of, say, Mach number = 0.95.This airfoil, if of conventional shape, will have a compression shockon the upper surface and one also on the lower surface at the same chord-wise station which probab~ will be close to the hinge line. The pres-sure on both surfaces will suddenly become higher in going from ahead of,to behind the shock waves, and, if the airfoil is sufficiently thick,there will be partial flow separation starting from the base of the shockwaves. Consider now a small positive Increase in angle of attack withthe flap held at 0° deflection. On the bottom surface the shock wave willmove back and tend to become weaker and the partial separation msy bereduced. On the upper surface, however, the shock wave will move forwardand become slightly stronger and, because of the extremely critical stateof the flow equilibrium, the partial separation msy be increased to amore extensive flow separation accompanied by a scuuewhatlarger increase

—

in pressure through the upper-stiace shock wave. The forward and rear-+

ward movements of the shock waves on the upper and lower surfaces,

NACA R! L53D21 7

respectively, together with the differences in pressure rise through theshock waves on the two surfaces, leads to a higher net pressure on theflap upper surface than on the flap lower surface; this accounts for thetendency of the flap to float against the relative wind, accounts forthe decrease in lift-curve slope, and also accounts for the large for-ward movement of the aerodynamic center. Turn now to the case of flapdeflection at 0° angle of attack. Assume the flap is given a slightpositive deflection. Again, the partial separation on the lower surfacetends to be relieved. h this case, however, the abrupt turn in flow ofthe stream (which is locally supersonic for some distance shead of theshock wave) produced by the deflected flap causes the pressure to increaseon the lower surface of the flap according to expectation. This chsmge.”is in the correct direction to produce positive flap effectiveness. Onthe upper surface, however, the separation again increases, the pressurerise through the shock wave again increases, and the flow apparently doesnot expand around the corner prcduced by the deflected flap. In thepresent case these pressure changes resulted in a~roximately zero changein net lift on the aitioil and therefore the flap effectiveness becameessentially zero for small deflections. ‘l?heseseparation effects areobviously highly nonlinear because neither the large positive floatingtendency nor the zero flap effectiveness extends over very large rangesof angle of attack or flap deflection, respectively. As the Mach numberincreases to 1.0, the canpression shocks on the airfoil move back to thetrailing edge so that shock-induced separation-can no longer occur.

Inasmuch as the foregoing phenomenon is associated with boundary-layer-flow separation, the magnitude snd exact details of the aerodynamicforce chsnges would be expected to depend strongly on Reynolds number.Experience with the BelJ_X-1 research airplane (ref. 7) seems to indicatethat the basic phenomenon occurs also at full scale, at least on thehorizontal.tail of this airplane. (%thert predicted that the flow break-down would occur at higher stresm Mach nunibersfor a horizontal tail onan airplane because of the slowing up of the stream caused by the passageof the wing through the air in front of the tail. The data for theX-1 ai@ane (ref. 7) tend to bear out this prediction. lh general, thetransonic-flow breakdown under discussion is believed to occur only onairfoil surfaces of fairly large thickness ratio and, probab~, Or rela-

tively low sweepback inasmuch as no evidence was found of its existencein the investigation of thinner unswept wings in references 8 and 9 norin the several investigations of 7.4-percent-thick 35° sweptback modelsreported in references 1 to 4. Iilthis connection it should be rememberedthat on wings of large sweepback the spanwise-flow effects are veryimportant end that these effects may change the nature of the trsmsonic-flow breakdown entirely.

Returning nowto figure 4(c), it may be notedk variations with angle of attack at subsonic speeds

by the presence of the flap gap at small angles of

that the hinge-mmentwere affected greatlyattack. Opening the

.

8 NACA RM L5~l

gap caused the hinge-moment variations to become erratic and highly non-linear at small angles of attack. At supersonic speeds, however, there &

was no measurable effect of the large gap on the hinge-moment variations.The negative flap floating tendency (evidencedby the slopes of thecurves) was much stronger at supersonic speeds than at subsonic speeds in

.

accordance with expectation; also, the negative floating tendency aboveM = 1.0 was greater for this unswept model thsn for the 350 sweptbackmodels of references 1 to k.

Characteristics in Flap Deflection

The variations in lift, pitching-momnt, smd hinge-moment coeffi.cients with flap deflection are shown in figure 7 for increments in Machnumber of 0.05 over the renge tested. —

The variations of lift with flap deflection (fig. 7(a)) were reason-ably linear over the deflection range tested except at M = 0.95 wherethe flap was practically ineffective for small deflections at u % OO;at an angle of attack of 5°, the ineffective deflection range appeared

—

at6Z -5°” The reasons for the existence of these characteristicshavealready been discussed. At subsonic speeds, the addition of roughnesscaused a slight decrease in flap effectiveness, probably because of anincrease in boundary-layer thickness, snd the removal of the gap sealcaused sm appreciable loss in effectiveness, probably because of the

#

tendency for pressure equalization to occur across the gap. AboveM = 1.0, neither roughness nor leakage through the gap had very mucheffect on flap lift effectiveness.

. .

Although leakage through the gap caused an appreciable decrease inflap lift effectiveness at subsonic speeds, the data of figure 7(b)indicate that the pitching-moment-producingability of the flap wasslightly increased. These results are explainable on ~he grounds tktflow through the flap gap causes the center of pressure of th flap liftto move rearward appreciably. Another interesting point is that thepitching-moment variation with flap deflection did not go to zero forsmall flap deflections at M = 0.95 even though the lift variation didgo to zero in the gap-sealed condition. This point is largely of academicinterest in connection ,withhorizontal-tail effectiveness, however, has-much as the important change in trimming moment from a horizontal tailarises from its direct lift change rather than from any small change inpitching moment about the tail aerodynamic center due to elevator move-ment; the latter effect is often disregarded in static stability analyses.

The hinge moments due to deflection (fig. 7(c)) were always greatand, for the gap-sealed configuration, were generally of the magnitudeexpected for an unbalanced flap at either subsonic or supersonic speeds.At M = 0.95 where the flap effectiveness was essentially zero for small

d

~ID.gITmi ------ .— .. ...

NACA RM L53D21 ) 9

b deflections, the hinge moments showed some decrease but, as in the caseof pitching moments, did not reduce to zero. Over large deflection r~es~the open-gap configuration generally had the greatest hinge moments andthe gap-sealed-plus-roughnessconfiguration had the least hinge moments.Here, again, the effect of le~e thro@ the gaP is ~dicated to ~ arearward movement of the flap center of pressure which is apparently moreimportat thsnthe lift decrease due to gap leakage insofar as the hingemoments are concerned.

Aerodynamic Parameters at d .=0° and 5 = 0°

Aerodynamic parameters measured,at 0° angle of attack and 0° flapdeflection (instantaneous slopes) are presented in i?igure8. Figure 8(a)shows the airfoil snd flap absolute ldft effectiveness; figure 8(b) showsthe flap relative effectiveness; fi~ 8(c) shows the airfoil ad f~ppitching-moment parameters related to the axis about which pitchingmoments were measured; figure 8(d) shows the positions of the center ofpressure due to az@e of attack (aerodymmic center) and that due toflap deflection (c.p. due to 5), and figure 8(e) shows the hinge-mcmentparameters with respect to angle of attack and flap deflection. b fig-ure 8(a) the subsonic lift-curve slopes are compared with the theozy ofreference 10.

A word of caution is believed necessary in regard to figure 8. Thesedata strictly apply only at ve~ small angles of attack end flap deflec-

. tions. An airplsme designed for supersonic speeds might traverse thetransonic speed range at singlesof attack large enough and, possibly, withelevator deflections such that the extreme changes in aerodynamic param-eters shown by figure 8 for Mach nmbers between 0.90 and 1.00 would beavoided because, as noted PreViOUSV~ these c-es occurred o- at lowangles of attack and for smell ranges of flap deflection. Above M = 1.0and below M = 0.90 the parameters shown are reasonably representativeof characteristics over fairly large ranges of angle of attack or flapdeflection except for the hinge-moment variations with angle of attackbelow M = 0.90 which, as pointed out p=viousl.y, were highly nonltiearbecause of gap effects.

In figure 8(a), good agreement is shown ktween measured and calcu-lated subsonic lift-curve S1OP=S in spite of the low Reynolds nmbers.Also, it was found that if the flap absolute effectiveness ~a at thelowest test speeds is corrected by linear extrapolation to the case forthe gap completely sealed (conditions shown in fig. 8(a) were O and69 percent of hinge-~ len@h sesled), very good agreement is alsoobtained between measured flap effectiveness and predicted three-dimensional flap effectiveness based on incompressible thin-airfoil.theo~ (@b = 0.0327). lh figure 8(b) it maybe noted that at M> 1.0

the flap relative efiecti ely 0.22. As is well known,.

10 NACA RM L53D21

the two-dimensional linear-supersonic-theoryvalue of Z?a/&5 for al/4-chOrd fhp iS 0.25.Infigure 8(d) it is seen that the center ofpressure due to flap deflection at subsonic speeds is considerablyfarther rearward with gap open than with the gap partially sealed; thisfact has been discussed previously. At M> 1.0 the center of pressuredue to flap deflection is in close proxhnity to 87.5 percent of thechord which is the location predicted for a l/4-chord flap by the two-dimensional linear supersonic theory which neglects aspect ratio andviscosity effects. As is well Wown, the two-dimensional linear super.sonic theory predicts a uniform pressure distribution over the flap andno change in pressures ahead of the flap hinge line due to flap deflec-tion. For such a pressun? distribution, it csn easily be shown thatCha = ‘2C~5 for a l/4-chord flap. Refereticeto figuies 8(a) and 8(e)

.

sh&s that; in the present tests of the model with smooth surface atM> 1.0, %5 was approximately equal to 0.013 and ‘%5 was approxi-

mately equal to -0.027so that the relation %5 . -2CL5 was almost

exactly satisfied. Therefore, it is implied by the present tests thatthe pressure distribution on the flap due to flap deflection becsmeessentially uniform soon after a Mach number of 1.0 was exceeded. Fi&lly,figure 8(d) shows that the rearward transonic aerodynamic-center shiftwas, neglecting the abrupt forward movement at small angles of attackbetween M = 0.9 W 1.0, about 16 percent of the mean aerodynamic chord.This value is about tk same as that found from the tests reported in sreferences 1 to 4 of 35°sweptback models of the ssme aspect ratio andtaper ratio.

.

CONCLUSIONS

Wing-flow tests at Mach nunibersbetween 0.65and1.10 of an unsweptand untapered NACA 65-009 airfoil model of aspect ratio 3.01 having al/4-chord full-span plain flap indicated the following conclusions:

1. The mudmum unstalled lift coefficient was almost twice as greatabove M = 1.0 as it was below M = 0.90.

2. A compressibilityphenomenon apparently peculiar to fairly thickaerodynamic surfaces was found in the region of M = 0.95for smallangles of attack and flap deflections. Evidences of this phenomenonwere a large reduction in lift-curve slope, an abrupt forward movementof the aerodynamic center to a position near the leading edge, u abruptreversal to a strong positive variation of hinge-moment coefficient withangle of attack, and a reduction of flap effectiveness to essentiallyzero for small flap deflections.

NACA FM L53D21

:

I-1

● 3. The hinge-moment varia~ion with flap deflection with gap sealedhad large negative vsl.uesof about the magnitude predicted by thin-airfoil theory at speeds below M = 0.90 and of about the magnitude

. expected from the concepts of two-dimensional linear supersonic theoryat Mach numbers above 1.0. The hinge-moment variations with angle ofattack were very nonlinear at subsonic speeds because of gap effectsbut were fairly linear and had large negative values at M> 1.0.

4..The effects of sealing 69 percent of the length of the l.l-percent-chord flap gap were to increase the flap lift effectiveness appreciably,to move the center of pressure due to flap deflection forward appreciably(with a consequent reduction in hinge moment due to deflection), and toincrease the linearity of the hinge-moment variations with angle ofattack, at speeds below M = O.~; at supersonic speeds, sealing the flapgap had little effect on any of the measured aerodynamic parameters.Sealing the gap increased the msximum lift coefficient slightly at allspeeds.

5. The additionof roughness to the first 5 percent of the airfoilchord on both upper and lower surfaces generally reduced slightly boththe flap effectiveness and hinge moments due to deflection.

Langley Aeronautical Laboratory,National Advisory Comittee for Aeronautics,

Langley Fie”ld,Va.

12 NACA RM L53D21

.REFERENCES

1. Johnson, Harold I.: Measurements of Aerodynamic Characteristics of

1 Chord Plain Flapa 3’5°Sweptback NACA 65-009 Airfoil Model With ~-

by the NACAWing-Flow Method. NACA RML7F13, 1947.

.—

2. Johnson, Harold I., end Brown, B. Porter:Characteristics of a 35° Sweptback NACA

~- Chord Horn-Balanced Flap by the NACA

RM L9B23a, 1949.

3. Johnson, Harold I., and Browm, B. Porter:Characteristics of a 35° Sweptback NACA

Measurements of Aerodynamic65-w9 Airfoil Model With

Wing-FIOW Method. NACA

Measurements of Aerodynamic65-009 Airfoil Model with

~-Chord Bevelled-Trailing-EdgeFlap aud Trim Tab by the NACA Wing-4Flow Method. NACA RM L9KL1.,1950.

4. Johnson, Harold 1., and Goodmsm, Harold R.: Measurements of Aero-dynamic Characteristics of a 35° Sweptback NACA 65-009 Airfoil

Model With ~-Chord Flap Having a 31-Percent-Flep-ChordOverhang #

Balance by the NACA Wing-Flow Method. NACA RM L50H09, 1950.

5. G6thert, B.: Control Effectiveness at High Subsonic Speeds. Reps.and Translations No. 72, British M.O.S. (A) VWcenrode, Feb. 1947.

6. Hemenover, Albert D., and Grahsm, Donald J.: Influence of AirfoilTrailing-Edge Angle and Trailing-Edge-ThicknessVariation on theEffectiveness of a Plain Flap at High Subsonic Mach Nunibers. NACARM A51C12a, 1951.

7. Drake, Hubert M., and Carden, JohnR.: Elevator-StabilizerEffective-ness and Trim of the X-1 Airplane to a Mach Number of 1.06. NACARML50G20, 1950.

8. Rathert, George A., Jr., Hanson, CarlM., snd Roll-s,L. Stewart:tivestigation of a Thin Straight Wing of Aspect Ratio 4by theNACA Wing-Flow Method.- Lift and Pitching-Moment Characteristicsof the Wing Alone. NACARMA8L20, 1949.

9. Goodson, Kenneth W., and Morrison, Wi~iam D., Jr.: AerodynamicCharacteristics of a Wing With Unswept Quarter-Chord Line, AspectRatio 4, Taper Ratio 0.6, and NACA 65AO06 Airfoil Section. Transonic- , -Bump Method. NAcARML9H22, 19490

NACA RM L53D21 13

10. DeYoung, John, and Harper, Charles W.: Theoretical Symmetric SpsmLoading at Subsonic Speeds for Wings Eaving Arbitrary Plan Form.NACARep. $Z1, 1948.

14

. .

.-. .

..

. .. . ..@ii5id-

-..q)n~,.l

Figure l.- Photograph of unswept NACA 65-009 wing-flow model with l/4-chordfull-span plain flap.

NACA M L55D21

A

.

NACA RM L53D21 15

3.27”I

i.2$+5- -,8z~

-

~4.C.=3.2’lU

P---

— .75C++5C

A

JECTION/VACA 65-009

‘~+ .035” 6AP /

—

--6”

– WING SURFACE

Figure 2.- Drawing of unswept semispan airfoil model with l/h-chord full-span plain flap. Model area, 15.&)sq in.; flap area, 3.88sq in.;aspect ratio, 3.01.

-. ..-.

1.0 x 1,06

.8

.6

/

.4

.2

I

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

Mach number

Flm 3.- Varlation d Reyndfb number with Mach nuder.

. , b ,

18 ~ II 1

Pm

1 1, . r

.:

M-+.M .m ,W., m.m, m.a I.mha 1.10 l.lti

(a) Vsxiation of ~ with U.

Fi~ b.- Aerdynamic characteristics due to angle of attack throughouttest Mach nmiber rage. 5 = OO.

18

?

.65

.70

.75

.80

.85

.90

.95

1.00

1.05

1.10

1.15

NACA RM L5~l

‘mJ.1 I

I Igapsealed,

— — .Ollcgap

R [ I Io. / ‘

a, deg

(b) Variation of C& with c%.

Figure 4.- Continued.

.

.

—.. .-—

.

.

.

.

NACA RM

M&

.65

.-70

.-75

.&l

.85

.W

.95

1m

1.05

1.10

1.15

L53D21.

Ch4.08

.04

0

0

0

0

0

0

0

0

0

0

0

-.04

-.oa

-.12

-.16

-.20

-8 -4 0 48 12 16 a3 24 2a 22=, deg

19

(c) Variation of ~ tith a.

Figure 4.- Concluded.

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12

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.2

=&=—

0 1.6 .7 .8 .9 1.0 1.1

Mach number

Figure 5.- Variation with Mach number of approximate lift coefficientand corresponding angle of attack at which flow b~s.kdown occurred(approximatems.xmusabk lift coefficient).

.—

●

�

..-

-_

.

.

.

●

.

.

NACA RM L53D21 21

:.s07

.925

.W3

Sm

.940

.940

.9s3

.956

.957

.907

.9e8

.970

.972

.078

.%3

.9.S4

=,&g

Figure 6.-Variations of hinge-moment coefficient with angle of attackfor Mach nmibers between 0.9 and 1.0. 5 = OO.

—

.8

1111111 1111111111111 L’ “

I I I “ {v% +Dlpw -

c

I 4 a 12 16 m5,&r

(a) Variation of ~ with ~. M. 0.65, 0.70, o.~, @ o.ti.

Figure 7.- Aerodynemic characteristics due ta flap deflection throughoutMach number range.

I , , , ,

.8

.0

4

.2

0

-a

-.4

CL

.

-.6

.0

.4

.s

o

-.s

-.4

% w

(a) Continued. M “ O.b) 0.9) 0.95) - l.~.

Figure 7.- Continued.

.0

,6

.4

.2

0

-.2

-.4

-.8

CL.0

.4

.2

0

-2

-.4 _,~-12 % -4 0 4 a 12 18 al -12 -e -4 0 4 a 12 16 al.

8,dq

(a) Cancluded. M = 1.05, 1.10, andl.lk.

Figure 7.- Continued.,

1 “1’4i, ),, ) ,.. I

I *

.2maled

.Ollc gap

M =- .65sealed + roughress

~\ 1

M= .711

.1\

\

a = 6.00\ ., =% & \ U.-5 -

\ . @\

<

T .‘ ‘J p

o\

\ ‘\

\ ~~ ..> 0

.50: -~

-.1‘.50

.%

.2

cm

M = .75 M -.80

~\

.1 \ .

\ a = 4.80 ‘% \\ a - 6.00

\+

o\ \/

-.1\

-al -lo 0 10 z) -m -lo 0 10 z)a, deg

(b) Variatim Of ~ with 5. M= 0.65, 0.70, 0.75, ad 0.80.

Figure 7.- Continued.

.2

\ . —1 sealed + rougheas

.1

\

\

.0

\ ~..10 1

-.1 .40 < \%+ 00

.2

cm

a = 5.2P

2P\

> \<

.1 M = .96 u -1.03

\

o

Y

i-.1

-xl -lo 0 10 al -m -lo 0 10 208, deg

(b) Continued. M = 0.85, O.$N, 0.95, and 1.oo.

Figure 7.- con-kin-.

II

rom

,i. ● ✌

✌❞✎

w

,,, IL

.;, ,,.

I . , , ,

.2sealed

- .OIIC gap

— em.le.d+ i-oughm?sa

\ M -1.06

.1> \ M -1.10.

3 /a - &OO

1.40J

o\

-.1

cm .2

.1\

M - 1.14/

m - 8.(XI‘)\

$& “\

Io , \

\

\

-.1\ - -

-Z1 -lo 0 10 m -22 -lo 0 10 m& deg

(b) Concluded. M=l.05, 1.10, andl,14.

Figure 7.- Continued.

.-

28 NACA F?ML53D21

.4

.3— a-5 o

qM -.65 M -.75

\ \\\

.2 . sealed \—

N ‘

— .O1lCgap

.+0 sealad+ rmgtmeas &

8,0.1 \ \

\\

\\\\\

\\\\

o\

\

-.1

\ \

-.

\

0..\

-.3\’ \

\,\L~ ,?> ,’;

\

Ch \~a~o-

\\ M -.70\

.2M - .Fxl

/,

+

.1 \ yL, \

\\

o\ \ \

\ ,b

\\ \\ ‘\

-. 1‘~, k,

\\\.

\\\

\

\\\,

h ‘\/

-. 2\\

\’ ,. - 520+ \

/ 7 \\. so a - b.d ‘\ \

-.Y \\\

.OP’\$ ‘~h\~4‘:

0

!@v‘Jo “

-.4.40 \

=E= -

- .6-al -lo 0 10 20 -z -10 0- 10 al

e,deg

(c) Variation of ~ with b. M = 0.65, 0.70, 0.75, and 0.80.

*

—~.—

—.

—

.

9

—.

==._

s

Figure 7.- Continued.

..“.-xG ~.-.

mm m L53D21 29

.

.

.

.

●

✎

.4

\.3

\ \ M -.85 M -.95

i\ \ y \

.2 S9akl— — .O1lCgap

sealed+ mugtuws\!

\ \\

.1 $

\

Q\\

{\i

-.1\\\

\ ).

\\ \

\)\

\Y\

-.2‘\ \ \\\ \

\’\\\ \

-.3 A~ \ \\+

. - 4.70 J, t--. c1t

c~ \t, 4\ \ \ , \,\

\1, N

\~ \d \

~o \ \

.2 M- .K1 M = I.m \

\\\’ \

! ,

\, \

.1 \‘ ‘\

\)

o \\ \

L’ --i ,

-.1~1

\

1 ~L \\

\-.2

~\ \ \ \’\

Y\\\

A ,’y\\ ~E1 0\

\

\ ,.\ ‘,a -4.8°

-.3I

\ $

‘~ ,1-.10\ ‘J

\

-.4\,

\ _ db/\ ‘*\

a . 5.P

~ .60/

-.6 -m-lo 0 10 m -m -lo 0 10 al

5,&g

(c) Continued. M = 0.85, 0.90, 0.95, and 1.00.

Figure 7.- Continued.

30 NACA RM L55D21

.4

\\\\

.3\

i M- 1.14,

\ ! M -1.05.

.2 \\\ \\

eaded— — — .Ollcgap

\sealed+ rmgluwss \\

y

.1!

\\\ \

o \\

\ \

\ \-.1

\\

\

,\

\ \

\-.2

\ ‘i\\

\ \

\

\

-,3 $ \

\ ‘,

\

L 7Pc~ \ \ \,.

\ \\\

.!’J’\- ~,.

,30.2 M- 1.10

\

%

- tl.m”a

\..1 \ -a .5.

\

o \\

\ \,-.1

\ ~1\

-.!2\\

\\~,\

-.3MI. w

\ fj,

_g 0

-.4 v-

\- -’ - E.@’

-.5 -m-lo 0 10 20 -m -10 0 .lil z)

5.da?

(c) Concluded. M=l.05, 1.10,

Figure 7.- Concluded.

md 1.14.

.

.

.

.

.

.

.

.

.

NACA RM L5~l

.08

.07

.06

CL

a .05

,04

.0:

.02

CL8

.01

(

su@&e GaQ

—0 Smooth Sealed—–~ Smooth Open—-+ Rough Sealed -

/

Theory\

31

Lp. .\\/’ -m1’ ‘-El

i

—..—

E- -El-“\

-~.-a--, -,\

‘Q\\;

.6

1 I 1 I ,x I I I

.7 .8 .9 1.0 1.1 1.2

(a)

Figure 8.-Sunmary

Mach number

Variation of C~ - %5 ‘ith ‘=

of aerodynamic characteristics at 0° angle of attacksmd 0° flap deflection.

I

NACA RM L53D21

.

.—.

.5

.4

.3

i%

.2

.1

0

—0 Sealed––+ Open

.- . .._-

Ek.-El.

‘R,‘m,

M

I I I I I.6 .7 .8

v.9 1.0 - 1.1 1,2 —

(b)

Mach number —

Variation of

Figure 8.-

&@ with M.

Continued.

.

.

.

.

●

.

NACA RM L53D21

slu&u2QGaD

~ Smooth Sealed–---H Smooth Open—-+ Rough Sealed

.02

r

33

cm=.395E

I “’~

~- —-. —--.01 ‘\ “.

or. . !, Cm

c “-_ / a“39~-‘5.395Z’

.——. __-n

o,~r degree

-.01 I I I I I I I 1

.6 .7 .8 .9 1.0 1.1 1.2

(c) Variation of NO 3956.

Figure 8..

Mach nunkr ,

‘d ~~o.3975with M.

Continued.

.

.

34

100

80

20

c

-2C

surfaQ2@u2

—0 Smooth Sealed‘– ‘%1 Smooth Open— --+ Rough Sealed

NACA RM L53D21

.

.

.—

r!)

~t I I I I I I

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

.

Mach number

.=a-

(d) Position of aerodynamic center and center of pressure due to flapdeflection plotted against Mach number.

Figure 8:- Centinued. .

.

NAC!ARM L53D21 35

.03

.02

.01

Cha o

-.01

-.02

.

Surface W

—~ Smooth Sealed----~ Smooth Open–-+ Rough Sealed

-.01

-.02

-.03

(e) Variation of

Figure

%--x

\

\

\

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

Mach number

8.-Concluded.

NACA-L8ngley -6-10-53 -S25