Low ASR Wing With Flap

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N CC tR E S E A R C H M E M O R A pyrN D U M

CopyR M A 5 O E 1

THE TRANSONIC CHARACTERISTICS OF A LOW-ASPECT-RATIOTRIANGULAR WING WITH A CONSTANT-CHORD FLAP AS

DETERMINED BY WING-FLOW TESTS, INCLUDINGCORRELATION WITH LARGE-SCALE TESTS

By George A. Rathert, Jr., L. Stewart Rolls,and Carl M. HansonAmes Aeronautical LaboratoryMoffett Field, Calif.

SW O U aAEZ WILDRestriction/Classification Cancelled


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NACA RM A50E10NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

R E S E R C H I v I E M O R I \ I D U M

THE TRANSONIC CHARACTERISTICS OF A LOW-ASPECT-RATIOTRIANGULAR WING WITH A CONSTANT-CHQRD FLAP AS

DEITRMLNED Y WING-FLOW TESTS INCLUDINGCORRELATION WITH LARGE-SCALE TESTS

yGeorge A. Rathert Jr. L. Stewart Rollsand Carl M. Hanson

SUMMARY

Wing-flow--method measurements are presented from 0.7 to 1.15 Machnumber of the longitudinal-stability and -control characteristics of atriangular wing of aspect ratio 2 equipped with a constant-chord flap.Comparisons are made with wind-tunnel and free-flight-test results athigher Reynolds numbers to determine to what extent the data are affectedby test technique and scale.

Changes in the lift-curve slope and aerodynamic-center positionthrough the transonic speed range were moderate and gradual. The agree-ment with results from wind-tunnel and free-flight tests was sufficientlygood to insure that the aerodynamic characteristics presented for-thisplan form are qualitatively valid..

The flap retained positive effectiveness and nearly linear lift and

Restriction/Classification Cancelled


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2ONFIDENTIALACA RM A50E10INTRODUCTION

Low aspect ratio, large amounts of leadingedge sweep, and extremetaper are planform characteristics increasingly associated with liftingor control surfaces for missiles and highly maneuverable aircraft wheredesign speeds are moderately supersonic and operation in the transonicspeed range is required. Information on this type of- plan form is beingfurnished through a coordinated study of a particular triangular wing ofaspect ratio 2 by several facilities of the Ames Aeronautical Laboratory.Results at low speeds have been reported in references 1, 2, and 3 a tmoderate to high subsonic speeds in references 1 and 5 and at supersonicspeeds, 1.09 to 1 53 Mach number, in references and 6 he character-istics of a constantchord flap fitted to the basic wing are presented atlow speed and full scale in reference 1 and up to high subsonic speeds inreference 7

This report adds the transonic longitudinalstability and controlcharacteristics indicated by wingflowmethod measurements extending con-tinuously through the Mach number range 0.17 to 1 15 Some hingemomentdata and the effectiveness of the constantchord flap are included.

The validity of the wingflow--method measurements has been Investi-gated by comparing them with results from the abovementioned references.Comparisons also have been made with results from investigations of modelsof a complete fighter airplane which differed in detail from the subjectconfiguration but had a comparable triangular wing and constantchordflap. Large and small--scale tests of this configuration at transonicspeeds are presented in references 8 and 9 respectively.

COEFFICIENTS AND SYMBOLS

The following coefficients and symbols are used in this report:


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NACA RM A50E10ONFIDENTIALC ate of change of hingemoment coefficient with flap angle, perd e g r e e

CLate of change of lift coefficient with an gle of attack, per degreeC ate of change of lift coefficient with flap angle, per degree

C ate of change of pitchingmoment coefficient with angle of attack,per degree

C ate of change of pitchingmoment coefficient with flap angle, per5e g r e each number ( )

Peynolds number P V Z F )S ing area of the semispan model , square feet

Virspeed, feet per secondapeed of sound, feet per secondb ting semispan, feet


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CONFIDENTIALACA RM A50E10qynamic pressure .PV2) , pounds per square footypanwise location, feet

ngle of attack, degrees

6ngle of the flap relative to the wing chord line corrected fordistortion under aerodynamic load), degreesflap angle uncorrected for distortion effects), degrees

air viscosity, slugs per footsecond

Pass density of air, slugs per cubic footTEST EQJJIPINT

Flow Field

The data were obtained by mounting semispan models in a region ofaccelerated air flow over a builtup test station on an airplane wing.The test Mach numbers presented are those determined without the modelin place but at a location corresponding to the model centroid of area.Due to pressure gradients in the flow field, the Mach number decreasedapproximately 0.01 per inch spanwise from the root to the tip of the model.Chordwise, the Mach number gradient was zero at low speeds and increaseduniformly to a value of 0.01 per inch, increasing from apex to trailingedge, at a Mach number of 1 15 The ratio of the displacement thickness


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NACA RM A50E10ONFIDENTIALThe inboard end of the flap was clamped rigidly in a cylinder which

could be rotated to vary the flap angle. Spanw1e support was providedby the two hinges shown in figure 1. The average gap between the flap andthe wing was 0.007 inch or 0.006 flapchord length . Figure 2 is a photo-graph showing the flap deflected.

The model with maximum thickness at 20percent chord and no flap wasalso tested with two modifications. As a change in the surface condition,the first 20 percent of the chord was coated with a mixture of lampblackand lacquer containing particles 0.002 inch high; also the original ridgeline was rounded sufficiently to reduce the section thickness ratio from5 0 to 4 8 percent.

Balances

Two types of recording balance were used. The lift and pitchingmoment characteristics were measured on a threecomponent electricstrain-gage force balance which was rotated at a rate of 1 0per second to varythe model angle of attack. This equipment has been described in refer-ence 10.

The flap hingemoment data were determined separately with a secondbalance upon which the flap was oscillated at a rate of 40 per secondwhile the flap angle and hinge moment were recorded. The forward portionof the model wing was fixed to the teststation surface at approximately0 0angle of attack with respect to the local air flow. The hinge momentwas recorded by photographing a dial gage attached to a beam restrainingthe flap in torsion.

DATA CORRECTIONS AND ACCURACY

Corrections derived from static calibrations were applied to compen-


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6ONFIDENTIALACA PM A50E10The accuracy of the physical measurements, Including the effects dis-

cussed, is ,summarized by the following table presenting the uncertainty Inrepresentative values of the test data at a lift coefficient of 0.50 forthe highest and lowest dynamic pressures. The listed uncertainties aredue principally to factors such as strain-gage zero drifts which remainconstant within one oscillation of the model; therefore, slope parameterssuch as L and m are considerably more accurate than the tabulatedabsolute values or Intercepts.

QuantityAbsolute value and uncertainty

M 0 . 1 4 7 M 1 . 1 5q = 1 72 lb /sq ft q = 697 lb/sq ftMach number, M 0.47 0.01 1.15 0 . 0 2Angle of attack,egrees 11.5 0.4 10.1 0 14Lift coefficientL 0.500 0.035 0.500 0 . 0 1 1 4 -Pitching-moment coefficientm -O.062 0.006 -0.112 0 . 0 0 1Flap angleegrees 7.8 0.3 7.1 0 . 3Hinge-moment coefficienth -0.050 0.018 -0.161 0.069Computed for a,.= 00 and= 80.

TESTS A1 D RESULTS

The lift and pitching-moment forces on the complete models wererecorded. from 4o +160 angle of attack at various constant Mach numbersfrom 0.47 to 1.15. The corresponding test Reynolds numbers, based onmodel mean aerodynamic chord, varied from 920,000 to 1,6 90,000 as shownin figure 3. The following configurations were tested:

a Maximum thickness at 20-percent chord., sharp ridge line,polished surface, no flap


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N C RNA5OE ONFIDENTI LThe lift and. pitching-moment characteristics of the basic wing(configuration (a)) at various Mach numbers are presented. in figure ii..For the other configurations the data are presented only in the form ofcross plots which illustrate the significant effects of the various modi-fications. The effects of the airfoil-section and surface-conditionmodifications, configurations (b) throu gh (d.), are shown at three repre-sentative Mach numbers in figure 5.

The flap-effectiveness measurements (fig. 6 re presented in theform of cross plots showing the variations of lift and pitching-momentcoefficients with flap angle at an angle of attack at 0 0 . Symbols arepresented in the figure in order to show the number of points used todetermine the fairing and the amount of scatter present. Figure 7 s h o w sat three representative Mach numbers the small effect of angle of attackup to 0 and the effect of sealing the gap between the flap and the wing .

The flap hinge moments were recorded from 50o 150lap angle atan angle of attack of 0 0 . These data are presented in figure 8 h eamount of displacement of the curves corresponding to increasing anddecreasing flap angle is an indication of the amount of friction andhysteresis.

DISCUSSION

Data obtained by the wing-flow method can be used with added confi-dence if their validity is substantiated by comparison with results frommore conventional test techniques. In the subsequent discussion, agree-ment between the wing-flow data and data obtained at relatively largescale from wind-tunnel and free-flight tests of comparable models will beused to establish which of the wing-flow results can be applied usefullyto full-scale configurations and which would not be valid. In the lattercase, the differences which will be noted may be due either to deficien-cies of the test method .or to aerodynamic scale.


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8ONFIDENTIALACA RM A5 E1The secondary comparisons are much more general and inc lude data bothfor full-span, sting-mounted, w ind-tunnel models of the same wing and for complete triangular-wing fighter-airplane configuration. The latter had slightly greater aspect ratio, 2.31 compared to 2.00; a slightly greaterthickness ratio, 0 06 compared. to 0 05; the same percent-area flap; buthad a rounded leading-edge airfoil section and a large fuselage and ver-tical tail. Data from free-flight rocket-powered models and a sting-mounted wind-tunnel model of this configuration have been used primarilyto furnish comparisons of control characteristics to higher Macfr numbersthan are available from the tests of reference 7.

Triangular-Wing Characteristics

Basic wing.- Results obtained directly from the wing-flow investiga-tion of the basic wing, figure 4 show no lift-coefficient or moment-coefficient irregularities or force breaks within the ranges of Mach num-ber and angle of attack investigated. The lift- curve slope and aerodynamic-center location are presented in figure 9 along with the correspondingAmes 12-foot pressure wind-tunnel data for Mach numbers up to 0.95 fromreference 4 In the wing-flow tests, the lift-curve slope increasessmoothly to a maximum of 123 percent of the low-speed value, then decreasesat supersonic speeds. The pitching-moment-curve slopes indicate gradualrearward movement of the aerodynamic center from 0 38 at 0.50 Mach numberto 0.47 at 1.15 Mach number.

The effects of Mach number on the lift-c urve slope and aerodynamic-center position and the increase in lift-curve slope at high lift coeffi-cients indicated by the wing-flow and the wind-tunnel tests are similar,although at a given angle of attack the wing-f low model develops less liftand has a more forward center of pressure. The differences, 6 percent ofCL and 0.02 are relatively small and unaffect ed by Mach number. This

acomparison is supported by the more extensive but less exact sec ondarycorrelation with experimental results and theory show n in figure 10. Gen-


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NACA RM A50E10ONFIDENTIALsurface would provide a considerable amount of effectiveness at all speedswith relatively moderate hinge-moment changes. These characteristics areparticularly important where the size and weight of the control actuatorsor the total amount of energy available are paramount considerations.

Airfoil-section modifications.- Two different locations of maximumthickness, 20-percent and 50-percent chord, were tested to determine ifvarying the location to allow for a th icker nose section and a largerleading-edge radius or to improve the supersonic drag characteristics(references 14 and 5) would affect the transonic stability characteristics.Figures and 12 Indicate little effect except a slightly greater movementrearward of the aerodynamic center with increasing Mach number in the caseof the model with the 50-percent-chord location of maximum thickness.

The ridge-line and surface-condition modifications were intended, togive some additional insight into the effects of low Reynolds numbers onthis plan-form and airfoil-section combination. Viscous effects couldchange the surface-pressure distributions at this scale by affecting thepressure peak over the physically sharp ridge-line, transition, and. flow-separation characteristics. It was hoped that these effects might besimulated by the modifications. The data in figure 12 show, however, thatneither rounding the ridge line nor adding surface roughness had any sig-nificant qualitative effect. Rounding the ridge line did decrease thelift-curve slope about 7 percent, contrary to the results of reference Ii.where a similar amount of rounding had no effect at the larger scale.

Flap Characteristics

Effectiveness.- The flap characteristics in figures and 7 a r enearly linear functions of flap angle and angle of attack at all Mach num-bers, even at flap angles of 10 and 20 The comparison in figure 7 betweenthe wing-flow results and the Ames 12-foot pressure wind-tunnel data fromreference 7 indicates that at the larger scale the flap produced propor-


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10ONFIDENTIALCA RM A 5 E 1Although thespeeds, the trendsThe Ames 12footin effectiveness tflow data show nofrom 0.70 to 1.05difference is greasoftening effect

lifteffectiveness values agree fairly well at l owwith increasing subsonic Mach number are contradictory.

ressure windtunnel tests indicate a continuous increase95 Mach n umber, the limi t' of the tes ts. The win gincrease; instead the flap loses effectiveness gradually

Mach number, the lose being about 30 percent. Theter than could reasonably be attributed simply to theof the Mach number gradients in the flow field.

When the secondary comparisons with controleffectiveness data fromr efer enc es 8 a n d 9 are considered (fig. 14), the seine differences betweensmall and largescale results are noted. The 'two sets of smallscalelifteffec tiv en ess data agr ee quali tativ ely as d o the two s ets of lar gescale data, despite differences in test technique in each case, indicatingthat the lift effectiveness is definitely-subject to a considerable scalee f f e c t . F r o m t h e p r e v i o u s c o n s i d e r a t i o n o f f i g u r e 7, it appears that theeffect at low Reynolds number shows up primarily as a reduction in theeffectiveness at small flap angles, suggesting that at small scale theflap may be operating in a separated flow region increasing in thicknesswith increasing Mach number. Both the magnitude of the overall loss' ineffectiveness at transonic speeds and the values at supersonic speeds arerelatively unaffected.

The effects of Mach number on the center of pressure of the additionalloading due to the d eflection of the flap are summarized in figure 15.Here the d iffer enc e between the lar gescale windtunn el tes ts and the win gflow data is comparatively small (0.04), indicating that the distributionof the additional load due to the flap on the wingflow model is relativelyunaffected by the small scale even though the total amount is re duced.The shift of the center of pressure of the additional loading back towardthe center of area of the flap at supersonic speeds is s hown.

The significance of the effectiveness measurements in terms of theabili ty of the c ons tan tc hord flap to balance a hypotheti cal allwin g air-plane of this plan form through the transonic speed range is demonstrated


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NACA PM A50E10ONFIDENTIALscale. The general qualitative trend. of the change shown at transonicspeeds by the wing-flow tests is in agree ment with the results from therocket-powered models. As pointed out in reference 7 this increase inCh is a serious control-design problem and figure 17 indicates large-scale tests would be essential for quantitative information on a particu-lar configuration.

CONCLUSIONS

Wing-flow-method measurements from 0.117 to 1.15 Mach number atReynolds numbers from 920 000 to 1 690 000 of the characteritics o f anaspect ratio 2 triangular wing with a constant-chord flap and comparisonswith larger-scale wind-tunnel and free-flight data have resulted in thefollowing conclusions:

Changes in lift-curve slope and aerodynamic-center position attransonic speeds were moderate and gradual. Agreement with the larger-scale tests was sufficiently good to indicate that these characteristicsare qualitatively valid.

2 The constant-chord flap retained positive effectiveness and nearlylinear lift and pitching-moment characteristics throughout the Mach numberand flap-angle ranges. In the wind-tunnel and free-flight tests the flapwas proportionally more effective at small flap angles resulting in non-linear characteristics.

3 Approximately 30 percent of the flap lift effectiveness was lost.at transonic speeds; however the change was gradual and computationsshowed that a hypothetical all-wing airpl ane could be balanced from 0.50to 1 15 Mach number without excessive or abrupt control movements. Thevariation of lift effectiveness with Mach numbe r at high subsonic speedswas opposite to that noted in the wind-tunnel tests possibly due tooperation of the flap in an increasingly t hick separated flow region.The magnitude of the over-all loss at trans onic speeds and the values atsupersonic speeds showed relatively satisfactory agreement.


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12ONFIDENTIALACA RM A5OE10REFERENCES

Anderson Adrien E.: An Investigation at Low Speed of a Large-ScaleTriangular Wing of Aspect Ratio Two. - I. Characteristics of aWing Having a Double-Wedge Airfoil Section with Maximum Thicknessat 20 -Percent Chord. NACA RM A7F06 1947.

2 Anderson Adrien E.: An Investigation at Low Speed of a Large-ScaleTriangular Wing of Aspect Ratio Two. - II. The Effect of AirfoilSection Modifications and the Determination of the Wake Downwash.NACA RM A7H28 191 7.

3 Rose Leonard N.: Low-Speed Investigation of a Small Triangular Wingof Aspect Ratio 2.0. I - The Effect of Combiation with a Body ofRevolution and Height Above a Ground Plane. NACA RN A7K03 1948.

4 Edwards George G. and Stephenson Jack D.: Tests of a TriangularWing of Aspect Ratio 2 in the Ames 12-Foot Pressure Wind Tunnel.I - The Effect of R eynolds Number and Mach Number on the Aerodyna-mic Characteristics of the Wing with Flap Undeflected. NACA RNA7K05 1948.

5. Berggren Robert E. and Summers James L.: Aerodynamic Character-istics at Subsonic and Su personic Mach Ni.imbers of a Thin TriangularWing of Aspect Ratio 2. I - Maximum Thickness at 20 Percent ofthe Chord. NACA RM A81 l6 1948.

6 incenti Walter G. Nielsen Jack N. and Matteson Frederick H.:Investigation of Wing Characteristics at a M ach Number of 1 5 3I - Triangular Wings of Aspect Ratio 2. NACA RM A7I1O 1947 .

7 Stephenson Jack D. and Amuedo Arthur R.: Tests of a TriangularWing of Aspect Ratio 2 in the Ames 12-Foot Pressure Wind Tunnel.II - The Effectiveness and Hinge Moments of a Cnstant-ChordPlain Flap. NACA RM A8E03 191 18 .


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NACA P.M A5OE1OONFIDENTIAL3ii. DeYoung John: Theoretical Additional Span Loading Characteristics

of Wings with Arbitrary Sweep Aspect Ratio and Taper Ratio.NACA T N 191 19471 2. Sprieter John P.. : Aerodynamic Properties of Slender WingBody

Comb inations at Subsonic Transonic and Supersonic Speeds.NACA TN 1662 1948.

13. Puckett A. E. and Stewart H. J.: Aerodynamic Performance of DeltaWings at Supersonic Speeds. Jour. Aero. Sci.. vol. 14 n o . 1 0Oct. 1947 pp. 567 578.14. Puckett Allen E.: Supersonic Wave Drag of Thin Airfoils. Jour.

Aero. S3i. Vol .13 no.9 Sept. 1946 pp. 475 - 4 8 1 4 .


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Low ASR Wing With Flap