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AIR DOCUMENTS DIVISION

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HEADQUARTERS AIR MATERIEL COMMAND

WRIGHT FIELD, DAYTON, OHIO

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FROM ANY LITIGATION WHICH MAY

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ARB Ho. L5B17

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

AT! N0.W6 WARTIME REPORT

ORIGINALLY ISSUED

April 191*5 as Advance Restricted Report L5B17

WIND-TUNKEL INVESTIGATION OF A RECTANGULAR

NACA S212 AIRFOIL WITH SEMISPAM AILERONS AND WITH

NONPERFORATED, BALANCED DOUBLE SPLIT FLAPS

FOR USE AS AERODYNAMIC BRAKES

By Thomas A. Toll and Margaret F. Ivey

Langley Memorial Aeronautical Laborat Langley Field, Ta.

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WASHINGTON

NACA WARTIME REPORTS are reprints of papers original:y Issued tu provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without cHfflgeTlf^rder to expedite general distribution.

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TEC fife;

NACA ARR Ho. L5B17

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

ADVANCE RESTRICTED REPORT

WIND-TUNNEL INVESTIGATION OP A RECTANGULAR

NACA 2212 AIRFOIL WITH SEMISPAN AILERONS AND WITH

NONPERFORATED, BALANCED DOUBLE SPLIT FLAPS

FOR USE AS AERODYNAMIC BRAKES

By Thomas A. Toll and Margaret F. Ivey

SUMMARY

Tests have been made in the Langley 7- by 10-foot tunnel to determine the applicability of nonperforated, balanced double sollt flaps for use as aerodynamic brakes. Information was desired on the braking power of the flaps as well as on the effectiveness and the stability of a conventional trailing-edge aileron located immediately behind the flaps.

A rectangular 10- by 60-lnch wing model of NACA 2212 airfoil section was used for the tests. Results were obtained for flat-plate flaps with no wing cut-outs and for flaps having Clark Y sections with cut-outs made in the win,<5 to simulate the space left open by the deflected flaps. The flap deflections, the chordwlse location, and the gaps between the flaps and the airfoil contour wero varied over wiäo ranges Ln order to determine the optimum configuration. Tu addition to the force tests, an investigation was mado to determine i?ny buffttintü tendencies of the aileron. Silk tufts and a flexible torque rod were used for these tests.

The drag was only slightly lower for the model having alrfoil-aecfclon flaps an'l v;ing cut-outs than for the model having flat -pitte flaps and no cut-outs in the wing; for both arrangements the drag was higher than that obtained in previous tests of an NACA 23012 fcirfoll with full-span, 0.20-alrf'.>il-ühord. perforated double split flaps. The aileron effectiveness was lov; in either case, except when the flap gaps were equal tc- about 20 percent of the wing chord and v;hen the noses of the flaps «ore at least SO percent cf the chord fr.iir. the leading edge of the wing.

f

I

NACA ARR No. L5B17

Although the entire model showed some tendency to shake, tufts Indicated that the air flow over "the aileron generally was smooth. Tests of the aileron attached to a flexible torque rod indicated almost no tendency for the aileron to shake; however, when the flap gans were 15 Dercent of the wing chord or less, the aileron acted as though it were overbalanced and usually tended to float against the stops for either positive or negative deflections.

INTRODUCTION

I The nresent investigation was made because certain

unpublished data had Indicated that satisfactory drag and lateral control characteristics had been obtained on an airplane with balanced double 3pllt flaps mounted ahead of a conventional aileron. Tests of balunced single split flaps on the lower surface of a wing hud previously been made by the NACA (reference 1), and certain flap locations were found at which the aileron was as effec- tive with flap deflected a3 with flap retracted. Tests of perforated double split flaos having no t;aps between the flaps und the airfoil contour (references 2 to 5) showed that such flaps produced desirable lift, drag, and pitchlng-moment characteristics for use as dive brakes and that the drag increment increased as the flaps were moved forward on the wing. The tests reported in refer- ence 2, however, showed that almost no effectiveness could be expected from an aileron located behind these flaps.

The present tests were made with a model configura- tion similar to that of references 2 and \ but having two flaps, similar to the flap of reference 1, symmetri- cally disposed abov« and below the wing. It was desired to determine if there were any flap locations at which sufficient lateral control as .veil as satisfactory drag characteristics could be obtained simultaneously.

APPARATVS AND TESTS

Vodel

The wing model was built of mahogany to the NACA 2212 profile. The model was of rectangular plan form; the span

«*•>

NACA ARR NO. L5B17

I

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was 60 Inches and the chord, 10 inches. Semispan ailerons having chords equal to 18.5 percent of the wing chord (0.18'jc) were provided. The ailerons were not balanced and had small gaps at their leading edges.

Two 3ets of flaps were used with the model. Both sets were full span, were nonperforated, and had chords of 2. inches. One set was made of flut steel plate

( —-in. thick) e \l6 )

and had rounded leading edges. Each flap

of this set wa3 attached to the wing by eight fittings along tho 3pan. The fittings were adjustable to allow variations of flap deflections, chordwlae locations, and gaps between the flaps and the wing. The wing had no cut-outs to simulate the space left by the flaps when deflected. Photographs of the modal mounted in the tunnel are given as figures 1 and 2. The second set of flaps was constructed of steel plate and wood to the Clark y section (fig. 5). Cut-outs In the wing were made to simulate the space ltft by the flaps when deflected. Each flap was attached to the wing by six fittings, which rested on narrow bridges left across the wing cut-outs.

The dimensions of the model and the flap locations and deflections tested are given in figures U and tj.

Tests

The dynamic pressure maintained for all tests was 16.37 pounds per square foot, which corresponds to a velocity nf about 80 miles per hour under standard sea- level conditions and to a test Reynolds number of 609,000 based on the chord of the model wing (10 in.). The effec- tive Reynolds number, based on a turbulence factor of 1.6 for the Langley '(- by 10-foot tunnel, was about 975,000.

The tests consisted principally of the determination of the lift, drag, and pitching-moment characteristics of the model with the ailerons neutral and of the rolling- and y»twing-moment characteristics of the aodel with the right aileron at various fixed deflections. * few tests wer* :nadä to determine the aileron hinge-moment coeffi- cients 2nd to Investigate the flow conditions in the vicinity of the aileron.

NACA ARR No. L5B17

Tests of the model with no wing out-outa and with flat-plate flaps were made with the flaps at a number of chordwise locations, gaps, and deflections. Only a few configurations of the model with airfoil-section flaps and with wing cut-outs were tested. These tests were made principally to check the validity of the assumption that the wing cut-outs and the flap section would have little effect on the results when the flaps are at high deflections. •

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RESULTS AND DISCUSSION

Symbols

In the presentation of the results, the following symbols are used:

CL lift coefficient (L/qS)

CD drag coefficient (D/qS)

Cm»/!, pltching-moment coefficient,about quarter-

chord point of airfoil [ •''^ ) \ qsc/

Cn aileron hinge-moment coefficient (H/qbaca2)

Cj' rolling-moment coefficient (L^iSb)

Cn' yawing-moment coefficient (N'/lSb)

where

L lift

D drag

H aileron hingo moment

Mc/L pitching moment about quarter-chord point of ' t,lrfoll

L1 rolling m'jmi.'iit ubiut wind axis in plane of symmetry of modol

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fiAC/. ARfi Wo. L5B1?

q

P

V

c

ca

S

b

a

yawing moment about, wind axis Iii.'pl&jie .of symmetry of model

Of. U

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dynamic pressure of free air

density " . . .. ° V.

velocity ° •.'••..••••"•• 'iiocf..-.,^, •;--

• r<>"y ••' wing chord c " °\- .... . o^ to ° O ' f • »•' •

aileron chord <3° o„ ••"

wing area / •"'

wing span * .".'Vs, v. ; ""-••

span of aileron "

angle of attack • ••' ••_. .-•'•'_-t

aileron deflection _; :' •••'

upper-surface split-flap deflection meaäürea..-). from wing chord line " .. ::••:.• -

°'• '•'•' ••'.£••?•' lower-surface split-flap deflection mea^u^'ed;".'..:;,^

from wing chord line '„,, ";••..•'•'•'."/.•

Gap 13 defined as the distance, measured fl&vßßjii*.W";'-"'' dlcular to the wine chord line, between the trues;'ij'pf''>:.^i'" contour and the portion of the flap nearest XHe.^lJisfVdi'i' contour. (Soe figs. J+ and 5.) „ ° !';'.-v',;:"; *$:'.•.'

Chordwise location is definod «s. the" dt^'t^c^v " measured parallel to the wing chord llnjs; frofcVö^tf-.^ß . wing leading edgo and the tangent - pej'pcJiöi.cu>l^-$$-:;;«h*-.. wing chord line - to the oortlon of the „flap. tie*be;atj>F^'..:•'••''•• airfoil contour. (See figs, k and 5.) '' ','•'•! '•'-.,/'• V;%i>,;-:-v.':?&:

Aileron effectiveness is defined as the incrfcmeiit",:.''.v!-.-"'';'.•'.•'. of rolling-moment coefficient between curves corresponding- -.".".'•• ';'• to two fixed aileron deflections.

0

I

Corrections

Ko corrections were applied for the effects of support-strut interference. Thu standard jtt-boundary i

1JACA ARR No. L5317

corrections, which were applied to all the force-test „data, arei

. Aa*8jrCL 57-3

AC» = B|CL2

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where to. Is In degrees, 8 is the jet-boundary ocorreotlon fs-ctor, and C is the cro3»-sectional area of the jet (69.59 sq ft), A value of e = 0.112 for the cloaed-throat wind tunnel wus used in correcting the results. Ho corrections were apolied to the pitching-, yawing-, rolling-, or hinge-moment crmfriclcnta: those corrections ar^» all cuall because of the relatively small size of the model.

Wing without Flaps

Testa wore made of tho model without flaps In order to provide u ba3is upon which to compare the tests of the model with flaps. The results of these tests are given In figures 6 to G. The almost linear variation of lift coefficient v,lth anjle of attach UI3. 6), the large and almost constant increment of rolling-moment coefficient between aileron deflections of ±20^ (fig. 7), and the approximately constant negative slope of the hinge-moment curves (flg. H) should bo noted.

King with Plat-Plate Plups

The model nas tested with two symmetrically located flat-plate flap3 at a number of chordwlse locations, gaps, and deflections. The results of the to&ts are given in rigures 0 to 20. The effect of flap deflection (flaps located at 0.'50c and with 0.05c f.;aps) is j:iven in figure 9. A comparison of this fl^ur-; with figure 6 indicates that, at zero an?jle of att'ici:, increments of Ora^ coefficient of 0.1;2 anu 0.!TiJ8 aru produced by tho flaps whon deflected 30° and 60° , respectively. Conparablu values of thü drag increment caused by i"ull-;äpan, 0.20c, perforated doublt split flans at the- same chordwiso

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»ACA ARR No. L5B17

location on an NAGA 23012 airfoil (fig. 3 of reference 2) are O.llj. and O.33. The irregularities in the curves of Cjj against a for the model with flaps deflected (fig. 9) are °f interest. The effectiveness of the ailerons is very low - at times, even negative - for this configuration (fig. 10).

When the flaps are deflected 30°, the irregularities in the curve of C^ against a are less pronounced when the gaps are 0.10c (fig. 11) than when the gups are 0.05c (fig. 9). The aileron effectiveness is greater when the gaps are 0.10c (fig. 12) than when the gaps are 0.05c (fig. 10). Increasing the flap deflection to 60° results in large irregularities in the curves of CL against a (fig. 13) as well as in reductions in the lift-curve slopes, particularly when the flaps are located far for- ward. The aileron effectiveness (fig. lk.) is generally lower and more irregular when flaps are deflected 60D

than when flaps are deflected 30° (fig. 12). Tests were made with aileron deflections of ±10° as well as 0° and ±20° for the condition of the flaps located at 0.80c (fig. ll(.(c)).. in order to determine if preater effective- ness might be obtained at the smaller aileron deflections. The effectiveness seems to increase almost linearly with deflection for low angles of attack but is about the same for 6a = ±10° as for 6a = ±20° at high angles of attack.

The characteristics of the model with the flaps deflected 60° and with gaps of 0.15c are given in fig- ures 15 and 16. The irregularities in the lift curves increase in magnitude as the flaps are moved forward (fig. 15). The aileron effectiveness decreases a3 the flaps are moved forward (fig. 16).

With the flaps located at 0.80c and with gaps of 0.20c, tests were made with the flaps deflected 60°, 90°, and 120° (figs. 1? and Iß). The lift curves for the conditions of flaps deflected 60° and 120° are char- acterised by flat regions near zero angle of attack (fig. 17). When the flaps were deflected 90°, un irregu- larity occurred, which was similar to those noted previ- ously. The maximum values of the lift-curve slopes for these conditions are only about one-half the value of the lift-curve slope for the model without flaps (fig. 6). The aileron effectiveness is relatively high

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NACA ARR No. L5B17

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(»bout 80 percent of the effectiveness when no flaps are attached) and does not seem to he anoreclably affected by the flap deflection (fig. 18).

Tests were made with flap chordwlse locations of 0.90c, gaps of 0.20c, and deflections of 6o° and 120°. The results are given In figures 19 and 20. The condition of flaps deflected 60° seems to be the most favorable of all the configurations that have been discussed. The lift curve (fig. 19) Is almost linear and the value of Its slope for angles of attach greater than 2° Is about 80 percent of the value of the lift-curve slope of the model without flaps (fig. 6). The ailerons are as effective as when no flaps are attached.

Tests were made with one flap located at 0.80c, with a 0.10c gap, and with a deflection of 60° (figs. 21 and 22). For the negative angle-of-attack range with the flap placed below the airfoil and for the positive angle-of-attack range with the flap placed above the airfoil, the effectiveness of the aileron for ±20° deflection is about the same as the effectiveness when no flans are attached. When the flap is below the airfoil, the effectiveness of the aileron deflected 20° decreases as the angle of attack is increased above -2° (fig. 22(a)). When the flan 13 above the airfoil, the effectiveness of the aileron deflected -20° decreases as the angle of attack is decreased below -2° (fig. 22(b)).

Wing with Alrfoll-Sectlon Flaps

The results of tests of the model with Clark Y alrfoll-sectlon flaps are given in figures 23 to 37- The lift, drag, and pitchlng-moment characteristics of the model with flaps deflected 30° and at chordwlse locations of 0.60c and 0.70c are given in figure 23(a) for flap gaps of 0.05c and in figure <J3(D) for flap gaps of 0.10c. A comparison of the curves for the 0.70c location of figure 23(b) with the corresponding curves of figure 11 reveals that the airfoil-section flaps and wing cut-outs result In slight decreases in the drag coefficients. A similar effect through most of the angle-of-attack range may be noted by comparing figures 27, 29, and 31 with figures 15, 15, and 17, respectively. Part of the reduction in drag coefficient Is probably a result of the fact that fewer fittings were used to attach the airfoil-section flaps to the wing than were

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f HACA ARR No. L5B17

used to attach the flat-plate flaps to the wing. The aileron effectiveness generally Is slightly higher for the model having airfoil-section flaps and wing cut-outs than for the model having flat-plate flaps and no cut-outs in the wing; this fuct can be noted by comparing figures 2.8, 30, and 32 with figures ll).(a) and ll;(b), lo(b) and 16(c), and l^(a), resDectlvely.

The variation of the rolling-moment coefficient with aileron deflection was determined for the model with the flaps located at 0.70c and with gaps of 0.13c and 0.20c (fig. 33). At an angle of attack of 0° the rolling- moment coefficient varied almost linearly with aileron deflection, but at an angle of attack of 12.1° the variation with negative deflections was irregular when the gaps were 0.15c.

Aileron hinge moments were measured for a number of model configurations and are presented in figures 3U and 35. When the flan gaps were 0.15c or less, the aileron seemed to be overbalanced and usually tended to float against the 3toDS for either positive or negative deflections. With the flaps located at ü.70c or at 0.90x, the overbalance wus eliminated by increasing the gaps to 0.20c. At an tangle of attack of 0° and at small aileron deflections, the slope dCh/ööa was still considerably less negative, however, than when no flaps .were attached to the model (fig. 8).

Because the model had a tendency to shake when the flaps were .deflected 60° or more, an investigation was made to determine if this shake were accompanied by a buffeting tendency of the aileron. No such tendency was noted when the aileron was restrainp-S only by the flexible torque rod used for the hinge-moment measurements. The investi- gation was extended by observing 3ilk tufts mounted from masts attached to the aileron ^t its mid span, «iidchord location. The directions and the stability of the various tufts are indicated in figure 3° f°r several model configurations. The tufts on and near the surfaces of the alloron were almost invariably smooth and were pointed In the downstream direction. Aileron buffeting therefore do«s not seem to be a st-rious problem for an airplane with balanced double split flaps.

A summary of the effects of gap and of chordwlse location of the two seta of flaps (each set ut deflec- tions of 60°) on the aileron effectiveness relative to

1

10 NACA ARR No. L5B17

I

that of the plain wing and on the drag coefficients is presented In figure 37- The aileron effectiveness increases as the gaps are increased and as the flaps are moved rearward. The drag increases as the gaps are Increased and as the flaps are moved forward. The varia- tion in drag is probably caused by the increased depth of the wake aa the flaps are moved forward while constant gaps are maintained between the flaps and the surfaces of the wing and also by the higher local velocities occurring at the forward oortlons of the wing. Refer- ence 5 showed that the Increment of drag caused by perfo- rated double split flaps was more than doubled when the flaps were moved from the wing trailing edge to the 0.30c location. From the results of the tests reported herein, however, the 0.30c location would be expected to result in little or no effectiveness of ailerons located back of the flaps, even though the gaps were large. Because the reduction in drag as the flap3 are moved rearward of the 0.6üc location is not very great and because the rearward flap locations result in improve- ments in the other wing and aileron characteristics, it seems desirable to locate balanced double 3pllt flaps at about 0.30c or farther rearward. Gaps of about 0.20c are necessary to obtain satisfactory wing lift, aileron- effectiveness, and aileron hlngu-moment characteristics.

CONCLUSIONS

Prom the results of tests of full-span, nonperfo- rated, balanced split flaps on a rectangular NACA 2212 airfoil, the following conclusions may be drawn:

1. The effectiveness of a conventional aileron behind balanoed double split flaps was generally low but increased as the flans were moved rearward and as the gaps between the flap3 anc' the eirfoll surfaces were increased.

2. The dray of the model increased as the flaps were moved forward and as the flap gaps were increased.

3. There was usually an irregularity in the curve of lift coefficient Lguinat angle of attack for the model with balanced double split flips deflected. The magnitude of the irregularity increased as the flaps

ilAfiA AKH i«o. L5317 11

F

i

were moved forward, as the flap gape were decreased, and as the flap deflections approached 90°.

k- The slope of the curve of lift coefficient against angle of attack generally decreased as the flaps were moved forward and as the flap gaps were Increased.

5. An aileron back of a balanced single split flap with a small flap gap may be us effective through a large part of the angle-of-attack rang«? as an aileron on a wing having no flaps.

6. The effectiveness of the aileron on the model having airfoil-section flaps and wing cut-outs was generally slightly higher than the effectiveness of the aileron on the model having flat-plate flaps and no wing cut-outs.

7- The drag of the model having airfoil-section flaps and wing cut-outs was generally slightly lower than the drag of the model having flat-plate flaps and no wing cut-outs.

3. Although the model with balanced double split flaps showed some tendency to shake, the aileron was usually oteady and the air flow was smooth on and near the surface of the aileron.

9. Plain ailerons back of balanced double split flaps acted as though they were highly overbalanced when the flap gaps were 15 percent of the wing chord or les3.

10. From a consideration of lift, drag, aileron- effectiveness, and aileron hinge-moment characteristics, a satisfactory practical configuration probably could be obtained with balanced double split flaps located at 80 percent of the wing chord and with flap gaps of 20 percent of the wing chord.

11. The crag of this sndol was higher than the dray of an NACA «13012 airfoil «1th full-span, 0.20-airfoil- chord, perforated double split flaps at the- same chord- wise location.

(-

Langley Memorial Aeronautical Laboratory Rational Advisory Committee for Aeronautics

Langley Field, Va.

12 NACA ARR No. L5B17

REFERENCES

I

1. Rogallo, ?. V... and Lowry, John G.- Yilnd-Tunnel Investigation of a Plain Aileron and a Balanced Aileron on a Tapered Winf.; with Full-Span Duplex Flaps. NACA ARR, July 19U2.

2. Purser, Paul E., and Turner, Thomas R.j «find-Tunnel Investigation oi Perforated split Flaps for Use a3 Dive Brakes on a Hectanßular JlACA 23012 Airfoil. NACA ACR, July 19is-l.

3. Purser, Paul E., and Turner, Thomas R.: «ind-Tunnel Investigation of Perforated Split Flaps for Use as Dive Brakes on a Taperud NACA 23012 Airfoil. NACA ÄP.K, Nov. I9I4.I.

Ij.. Purser, Paul E., and Turner, Thomas R.! Aerodynamic Clif.ractc-risliics and Flap Lomis of Perforated Double Sollt Flap3 on a Rectangular KACA 2*012 Airfoil. HAOA ARI".. J.'ui. V)l3-

5. Blenkush, Fhilip G. , Kermos, Raymond F., und Landls, Merle ii.j Effect of Dive Brakes on Airfoil and Airplane Characteristics. Jour. A<TO. Sei., vol. 11, no. 3, July X0Uk> PP- 25U-260.

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X Stmlspan aileron

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^ Hinge axis

•Plane of symmetry

-§•30"

fa)Plan form of wing.

S.Ow*M

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Figure d.- Dimensions and flap configurations of the model with balanced double split flaps having flat- plate sections. Wing airfoil section, HACA 2212.

Flg. Ba,b,e NACA ARR No. L5B17

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X _1_

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60 60

ft) Fhp deflections for various c/iordmst hcahonsand gaps-

Figured - Dimensions onet f 'tap configurations or the modti tvith ba/anced (fotfb/e split flaps having Clark Y sections.

Wing airfoil section* MAC A 22.12..

W :„

\

NACA ABR Ho. LBB1?

U 1"£

?! ?1

5*

-.6 1.0

@ |

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of the t/ACA IIU *">3 ""^ei "° f/"?S' °"

-.^j/fc"' '.• ,'V '.>- • •

n». ? NACA ARR No. L5B17

\

•12 -3 -4 0 4 8 12 16 20 Angie of attack, a, deg

figure 7.-R.olling-and yawing-moment characteristics of the right semispan aileron on the AfACA ZZ/Z wing mode/. A/a f/apa.

y'rt ;^;#'-.;; "W-

NAC* «If NO. LBB1,

'ig. 6

\

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'*. (I. ' ':• -'v- 4-

"r '

»1». 9 NACA ARR Mo. L5B17

\

^r~w

-6 -4 -2 0 .2 .4 .6 ß 1.0 1.8 Lift coefficient, CL

F/gureSrUftsCtrqgiarMf pitching-mornent characteristics of the MACA 21/2 w/ny model equipped with balanced double split Flaps having ftnt-plaie sections- Chordwise loctrtioihQBQc,#ap3,005c; Sa,0'.

(-

NACA ARR NO. L5B17

<& f Fly. 10a

\

fc

I

<§

2^j3QL W5c W5c

.02 /v,

-zo 0

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0 V3

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^ 0

i* M M< M M ^ H N u •H fe *H S^ ^ 5*! i—'

ITTEE m mou WTtC

fi<7 (deq)

A-2Ö O 0 p 20

5 ft .0/ *^ i 0

- -oi f % -12 -8-4 0 4 8 1? 16 20

Angle of at tack, ex., deq (a)6fuanc/6fL,7so".

figure /O- f?of//ng-on(/yawinjj-momen+ characteristics of the right 3emispan aileron on the A/ACA 22.12. wing model equipped with balanced doubl« split flaps having flot-p/ote sections. Chore/wise locet/an, OSOc, gaps,Q05c

Pig. lob

.1

8 o

I 1

£

NACA ARR No. LBB17

/SL60'

M ^

01 r |

,20 0

E ^

•P A -4 l-J

\ r-t t*-d

.01 'ä •"4 ̂

JOB

A

t= £ incE FM UMMU WTKI

A -20 o fl O 20

-IB -8-4 0 4 8 IB 16 BO Angle of attack, a. tdeg m 6^antt6^,60'.

Figure HX-Concluded.

.01 h* 0 |l

-.01 s; u

I-

^

I

NACA ARR NO. LSB17 Fig. 11

© ]

Chonlwse location

(fraction ofc) D 0.70 O .60

-.4 -£ 0 2 A Lift coefficient, CL

Figure lir Lift, drag, ana'pitching-moment characteristics of the NACA 22l£ w/ng moats/ equipped with balanced doud/e jo/it f/qa? hairing flat-pkrte sections. Gops,0J0ct 6a, 0) &. and 6fL, JO".

Flg. 12a NACA ARR No. L5B17

\

{ 1 1

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^r-^f^-zr \ \JOc

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NkriOHM. «WBCNn commit m aunwwrics

-12 -8 -A 0 4 a 12 16 20 Angle of attack , a, cleg

(a)Chordwise location, 0.70c. Figure /Er Rolling-and' yawinjr-moment characteristics of the right

sern/span aileron on the HflCfi ZZI2. wing model equipped with balanced double split flaps having flat-plate sections. Gaps, 0.10c-, 4j/ and <^, 30°.

f-

i

NACA ARR No. L5B17 Fig. 12b

€ f

fcrr -80c-

, , l/Oc

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C/OChorjw/se location ,060 c Figure /Z.- Conc/ut/ec/.

n-2.0 o 0 D 20

o ?£

Pig. 13 NACA ARR No. L5B17

I

-4 -.2 0 .2 .4 .6 .8 IJ> Lift coefficient, £

F/ffure 13.-Lift,dray, and'p/tching-moment ctoracter/st/es of the NACA ZZiZ w//y model equipped with balanced doable sp/if flaps having f/a?-plbfe sect/0/73. Gaps, O.IOcs 6a, 0] 6fuand 6fL)6(f.

NACA ARR No. L5B17 Pig. 14a 1

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I

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(dog) A -20 o tf D ZO

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1 0 t \

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NATKMM. uvaoav cammt m anaurrict

-ml -18 -8-4 0 4 8 18 16 20

Angle of attack, a, deg (a)Chordwist /oca//on,0L60c.

figure 14-Rolt/ng-andyawiry-moment characteristics of the right semispon af/eron on the A'flC/l ZZIZ wing model eqmppei with bo/anced douö/e sp/it f/aps haying fiatplate sectians Gaps, 0.10c\ 6fu and SfL, 60°.

Figf. 14b NACA ARR NO. L5B17

\

!

a

A -60'

H&L IrJQc

-70c- XJOc

(deg)

on JÜL . . j—.

ni \_ ui =T , V n

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-.0/ £ «

-a -5 -4 O 4 8 ia 16 20 Angle of attack, ac, deg

WChorstwtee location,0.70c. Figure Z4. -Continued-

(

\

NACA ARR NO. L5B17

•3» OZ

s I I

jof

:0I

& 1 Fig. 14c

g'joe

$

-IB -8-4 0 4 6 12 16 80 Z4 Angle of attack, a, deg

(c)Chorcfw$e location,üdOc, Figure f4. - Concluded.

\

NACA ARR Mo- L5B17

\

Fig. 16

location {fraction ofc) A 0.60 D .70 O .80 x .90

6fu and \ , 6° •

<

I

NACA ARR No. L5B17

$ /5c

Pig. 16a

^- • U»4 3b^ r/5e

•60c IV^-60" 6o

fdegt

ft» o 0 1 .oe 4»

S -J k tf* 1 n 20 f 1 *s ^

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ooMm run • Ml HUVT ICt 0/|§

r

-12 -8 ^4 0 4 6 18 16 20 Angle of attack ,<x, deg

folChordme location, 0.60c. figure 16-Rolling-and yawing-moment characteristics

of the right semispan aileron on the NACA ZZIZ wing ffadel equipped with balanced double split flops having flat-plate sections. Gaps, 0.15c, 8f and 6f ,60°.

?ig. 16b NACA ARR No. L5B17

\

.03

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1^

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-/£ -8 -4 0 4 8 /Z 16 20 Angle of attack,oc,deg

fb) Chord wise location, fl 70c. figure /&- Continued.

.01

0

-01

\<5

•5^

NACA ARR No. L5B17 fig. 16c 'f

I

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;* .03

c E 1

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p?33fifcM>^::te 5Ö -?0\

1 i -a * 8 -IS -8-4048/8

Angle of attack, or, deg (ciChordwm location, 0.80c.

Figaro 16.- Continued.

16 80 24

Fig. 16d NACA ARR No. L5B17

I

£

-8 -4 0 4 8 12 Angle of attack, oc,deg

(d)Chordmse location, 0.30c. Figure 16.-Concluded.

•»">

NACA ARR NO. L5B17

*fc>

Fig. 17

f

t

I ;t •4

v'

1 ^ ./

& •*•"• P Ö

^1 -./

20

16

12

t a * * V 4 ^ ^ S 0 \

^ -4

-8

-/2

-i4 -.2 0 .2 A •> Lift coefficient ,6^ Figure 17- Lift, drag, and pikhing-moment characteristics of

the /i/ICfl 2ZI2 wing model equipped with balanced double split flaps having floi-plote sections. Chordwise location, 0.80c, gaps, 0.20c. 6Qi0

o.

Pig. 18a NACA ARR No. L5B17

\

-3-4 0 4 8 IS Angle of attack,a., deg (o)6fu and &^ , 60'.

F/guro IS-Rolling-and yawing-moment characteristics of the right tern/span aileron on the MAC A ZZI2. wing model equipped with balanced double splii flaps having flat- plate sections. Chord wise location, 0.80c, gaps, 0.£0c.

NACA ARR No. L5B17 Pig. 18b ]

•vT

| .08

u

S o

5 -.oe

•8. 03

-.04

b V$0#

^Öc £T *<* ~"]L^~^

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-.01 £ 5 -/£ -Ö -4 0 4 «S It 16 BO

Angle of at tack, or, deg fb)6fu and 6fL, 30'.

Figun I8.~ Continued.

f-

Fig. 18c NAOA ARR No. L5B17

\ c

<S

\^\>eo' .20c

-80c- .20c

LMISOI

-8-4 0 4 8/2 16 iO Angle of attack , or, deg

(cJ 6f and 6^ , IZO '. Figure 18-Concluded.

f-

NACA ARR No. L5B17

@ f Kig. 19

I

6fuand6fL

q , IZO

-.4 -£024 Lift coefficient, CL

figure 13,-Lift, drag, and pitching-moment characteristics of the NACR ZZIZ wing model equipped with balanced double split flops having flat-plate sections. Chordwise location, 0.90c. yaps, 0.20c. 6a, 0'

f ,

Fig. 20a NACA ARR No. L5B17

\

NATIONAL JtfMMJMY

-12 -8-4 0 4 8 12 16 20 Angle of attack, ot , cleg

fa)öfo and fy 60'. figure 20- Rolling-and yawing-momenf characteristics of

the right •samispan aileron on the NftCft 2212 wing model equipped with balanced double split flops hovirj flat-plate sections. Chordwse location,090c^ps,02 Oc.

NACA ARR NO. L5B17

g f Fig. 20b

\

I

\20c

^ M:

:90C- .20c

rj*\m,f\

•MTIOMl WVUOUT cstwrra m waumu

-12 -8-4 0 4 8 18 16 SO Angle of attack, a, deg

fb)6ffJ and ^,/ZO' Figur a 20- Concluded.

f-

I

Rig. 21 NACA ARR Mo. L5B17

Configuration A Configuration B Configuration C a oft

figure Z/.- Lift, drag, and pitchina-moment choracier/sf/cs of ttta HflCfl £2I£ mng model equipped with bohnced single and double split flops honing flat-platt sect/ons. Chordmst /ocotion, OJDOC, gaps, 0.10c, Sg,0^&ond&f , 60'.

NACA ARR No. L5B17

9 f riff. Z2a

\

V?

I

s 8 -01 IS £

-4040/26 Angle of attack, cc , deg

(a) Upper ftp removed; 6$, 60". Figure 2trRdling-andyaiy//}g-moment characteristics of the right Jem/span aileron on the NACA 2212 wing model equipped Hith

balanced single <split flap having flat-plate-section • Chordtvise hcation, QSOci gap, 0.10c,

Flg. 22b NACA ARR NO. L5B17

\

.£ -ß -4 0 4 8 * IS & Angle of attack, oc, deg ^ NATMWL «V»«

(b) Lower flap removed; bf^G? • Figure Or Concluded.

NACA ARR No. L5B17 ^ f

Fig. 83a

\

Chonkmsa location

(fractionate) A 0.60

.70

NATKMUU. WVtSOAV

•4 -2 0 .2 4 .6 .8 Lift coeffic/'ent,Ci

(o) Gaps, 0.0fc. Figure 2d-Lift, drop, andpitching-rnoment characteristics of the NACA £&d wing model equipped with balanced double split fkpa hairing Clark Ysections, do, 0} 6*ond6fL, <30e.

Piff. 23t NACA ARR No. L5B17

l

ChorävisQ location

(fraction ofd A 0.60

-ff -4

Figure £3r Concluded.

.70- Dl..

.^ 0 .Z A .6 .8 Iß l/ft coefficient, C± mmuL ^^

fb) Gaps, aßc.

NACA ARR NO. L5B17

B* rig. 24

1

\

rö«T

^ V fc«V 30'

Jn <&M

is —1

<s \-ßl -1 )

X V -20

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(MV nwr «

6«

x-/<3 O 0

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-12 -8 26 -4 0 4 8 i2 16 Angle of attack, a, dbg

Figure £4.-Mling-and t/awing~moment characteristics of the right semispon aihron on the NfiCfl &2I2, wing model equipped with balanced double split flaps having Clark y sections, Chordwisa location, 0.60c; paps, 0,IOc}

6fu and &fL, 30°.

\-

rig. 25 NACA ARR No. L5B17

\

Chbfiwis» location

(fraction ofd a 0.60 a .70

NATKMM. Wm

.64^

•te" .60 %

.66 %

\- .51 t> §

IA -.2 0 Z .4 6 .8 Lift coefficient, Q

Figure £5-Lift, drag, and pitchmq-moment charachritUcs of the NfiCfi Z2I2 wng model equipped with balanced double split flaps having Clark V sections. Gaps,0.05c. 6a,0i

6fu and 6fL, 60°.

NACA ARR NO. L5B17 fig. 26

\

JUS Go

(dey) -20 -to

/K .cu t

/2 3*

St-

t-' f.

20 —

— -- - -JB

—

— —•

<_ *.

A b*b H M M H P M r* |H H &e«r_- r1

W» •-- — ..._

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1 ! COWWTIU FOt MMUUTK1

«to fag)

A-20 X-/Ö o 0 O iO • 20

-12 2d -8 -4 0 4 8 12 IS Angle of attack , ac, <feg

Figure 2&-Rolling-ond yamng-moment characteristics of the right semispan aileron on the A/ACfl 2212 wing model equipped with balanced double split flaps having Clark V sections. Chordw/se location, 0.70cs

gaps, 0.05c•, 6fu and 6fL, 60°.

\

Fig. 27

vl

&.;

«5:

I

NACA ARR No. L5B17

1 „ CAonfwise J//?6A,=S0"

Chordmt location

{fraction ofcJ A 0.60 a .70

-4 -2 0 .2 A .6 lift coefficient, CL

Figure ZZ-Lift, drag, and pitching-mom ant characteristics of the NACR 2212 wing model equipped with balanced doub/e split flaps having Clark y sections. Gaps, 0.10cs

6a , 0\ Bfu and^ , 60*.

(-

i*4f • ••:*'-•

NACA ARR NO. L5B17 Tig. 28a f

\

is

>

t

-/2 -8 -4 0 4 8 /2 if 2d A/ty/t of attack , ec, def

(a) Chordtviss location, 0.60c. figure 18.-Ro/ling-ond yowing-mornent characteristics of

the right semispan aileron on the NACA Z.Z 12 wing model equipped with balanced double split flaps having Clark V sections. Gaps, O.IOc} 6fu and \ , 60'.

tit- 28b NACA ARR No. L5B17

I

d

c

-70c- —i^ y* =60°

'. ft* "" \ ""~ mm, t

04 •~

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V*-& =6»*

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IUTKMU «WIM» cowrm na «MM IT«

» is fl

-/2 -8 -4 0 4 8 /2 /6 20 Aty/e #f attack, a, Off

fblChordwise locotion, 0.70c. rigvre £8.- Concfuded.

°W

\

f NACA ARR No. L5B17 Pig. 29

Chordmse location

(fractionate) U 0.70 3 .80

-.2 0 .2 .4 .6 Lift coefficient, C^

Figure 2.9-Lift, drag, and pitching-moment characteristics of the NflCfi ZZIZ wing model equipped with bo/onced doubfe spfif flaps having Clark V sections. Gaps,ÜISc, 6a,0°s öfu and 6FL,60°.

"•£* I

Pi;. 30a NACA ARR No. LSB17

-70c- i^'60'

-/2 -8-4 0 4 8/2 Mg/e ef attack, a: , deg

to) Chordww location, 0. 70c. Figaro 30.- Rolling- and yawing-moment characteristics

of the right semispan aileron on the NACfi BZI2 ning modal equipped with balanced doo/>/e split Plops having Clark Y sac lions. Caps, 0.15c. 6f and Sf , 60*.

4-

NACA ARR NO. L6B17

Wl Flff. 30b

I

JSL

-12 -8 -4 0 4 8/2/6 20 Angle of attack ,&, deg

(iJ Cnordwise location, 0.80c. Figure 30.- Concluded.

>»••>»•

i

f '

Pig. 51 NACA ARR No. L5B17

\

> *

H ./

^^" 0

-1

-2 0 .1 A .6 .8 Lift coefficient, CL

figure 3/.-Lift, draff , aod p/fchiof-moment character/sties of/he A/ACA 2Z/2 wiop mode/ equipped with balanced double sptit f/aps hewing Clark V sections. Gaps, 0.20c. 6a, 0° , <5~fyand<SfL , 60°.

t

I

NACA ARR No. L5B17

\£j*«r

^^•60-

J rig. 32

-4 0 4 6 tZ Angle of attack, a, dag

tvpitre32-Qol/ing-ondyaw/np-mooient characteristics of the r/ybt aem/sfion a//eran on the NACA 12.12 winy model equipped with balanced doub/a split f/a/>a havi/w CArr/r Y sections. Chordvtise location, OSOc, pops, 0.20c i dfo and 6fL, 60°.

rig. 33a NACA ARR No. L5B17.

.70c

<^m fa .15c

.15c

\

I

8 1

!

cc fdegJ

o 0 ( \ \ A s

ft

r

t\ K F k I ¥ k X, i

.04

JD3

.02

.01

0

-JOt

-M

-.03

-.04 -30 -10 -10 0 10 10 30

Aileron deflection, 6a, deg (a) Gaps, O-/6c.

figure d^rRoMng-moment characteristics of the rightaemiapa/t o/feronanthe A/ACA 22/2 wing model equipped with bo/aneed doub/e sp//t flaps having Clark Y sections. Chord- m'se location, 0.70c, 6^ and 6fL , 60*.

{

NACA ARR No. L5B17

.70c -I A

tig. 33b

\ .04

V .03

1 t .01

% % .01

} 0

s -JOI k .*: ^ -.01 ^

-.03

-.04

ds,- cc

sfc <dsqj

^- ° 0

YV r A J2I

«k uiiiwBy«

-30 -10 -10 0 10 ZO 30 fliltron frflection, ög, dag

(Ü Gop5,O.20c. ffyure ZZrConc/uefeef.

?*•••,.*'•••••?.*••'-?-?

rig. 34a NACA ARR MO. L5B17

\

./ Unstable 1 i leg) regions -^ J 0

0 \ >J ^

,. SL IP.? 7 7 T \ <3= 1

*r _ / < F= =« r* >' Chordmse location, 0.60c%gaps,0.05c

it c

oeff

a ^

-•^ • ( deg)

1 " * Jl >^ sj J -A (<_ 0

l?R fe s* *r i f -*i

\ /

' -/ ^ ;— t>

Chordwise locatton,0.6Oc-,gopst0.l0c C $

a (dea) / ./

A -~J ! • o t* 0

0 i ' Iff.

fe *' 1

-./ —( ym

Chordwise location. 0.70c :aaDS. ÜKk

-30 -20 -10 0 10 80 30 Aileron def/ection, 6a, deg wrKmi_mMKr

fa) 6f =<f/r m30*. mmaammmmm Figure MrH/nge-moment characteristics of Me riphf

semi span a/feron on the A/ACA Z2./1 winy model equipped with ba/anced dot/Me spfrff/opa hoviny Cfark Y sections.

-. • -fi

NACA ARR NO. L5B17 I.

fig. 34b

J

\

L Chordmse y~stat/on

-^=P^ \fu.60'

Gap

1

-30 -E0 -10 0 10 20 30 Aileron deflection, Sa, deg

(b)6fu-8fL-6Q°. ftyure 34r Concluded

Flg. 35a NACA ARR No. L5B17

\

J <&

¥ 2

i .1 > <o 0

^ § -.1

^ J

* Z $

.1

0

Mos/oAh region

,-H i or kfegJ

0 7

- I&M

^S k i k N

t \

* . i

Gaps, 0.15c

M K 11 k

i N L HATiewu. Mvnoav »•na m IIWMIT

it (dag)

• IZ.I

Gaps ,0.20c

Lox K \ ̂—

-.z -30 -Z0 ~/0 0 W 20 30

Aileron deflection, 6a, deg (o) Chordw/se location, OJOc.

f/gure35.-Hinge-momonf characteristics of the right semi- span aileron on the A/ACfl 2212 wing model equipped with balanced double split flaps having Clark V sections. fy, and 6?L, 60?

\

NACA ARR No. L5B17 Fig. 35b 9 I

Mc

^z:m Gap

* 6f,'60

T Unstable region k \ \ -^ ^£ K fh a.

(degJ t- IZ.Z

1 Y^ L A U »,

W Gaps, 0.15c N\ \ —

, t z A] « v a:

töte? \ si /_. s^ _s •T K >J^

V. K^ i j / T -Ö-- K t •UTNNM. AMHMY

comma m amum

<4U Sfr

Gfl/«, #<J0C Is 4* r& •s S

i

v£ J

s

I J

* •/

'J

.JO -to -io o io IO Jb Aileron dafltctwn, 6g. deg

(b) Chordw/se location, 0.80c. Figure J5- Concluded.

{

rig. 36» NACA ARR No. L5B17

\

ruff <x,o'

r-*- ~1'

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Gaps,aosc Gaps.aioe Chordwise location, 0.60c a

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figure JG.- Tuff study or flow conditions above and below the right semispon aileron an the NACA £212 w/ng model equipped w/th bo/anced douö/e split flqts having Clark Y sections. Tuftslocated at aileron midspan; S indicates smooth flow; R indicates rough flout

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Figure 37- Effect of chordwtic local/on and jaos on the drag coefficients and the aileron effectiveness of the NACA 2.2.12. Wing model equipped with bdoncet* doable split flaps having flat plate or Cbrk V sections. a,0''cfy and c£, 60°

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TITLE:Wind-TüTinel Investigation of a Rectangular NAC A 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes AUTHORS): Toll, Thomas A.; Ivey, Margaret ORIGINATING AGENCY: National Advisory Committee for Aeronautics. Washlngtpn. D. C. PUBLISHED BY: (Same) JMS f , , , , , &YL

ATI-6476

(None) OIIO AOOtCY NO

ARR-L5B17 PUflliMMO AOfNCT NO

I April '45 COUNTBY U.S.

HUMTtATMNI I photos, diagr, graphs I mmk I U.S. | Eng. I 41 I photos, dia

ABSTRACT:

Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corres- ponding cutouts made in wing were tested for various flap deflections, chordwise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80% wing chord behind leading edge, airflow was smooth and buffeting negligible.

^Aw^/7/^/t ^^^ DISTRIBUTION: Request copies of this report only from Originating Agem

/V/ DIVISION: Aerodynamics^ SECTION: Wings and Airfoil» <«)- J rj

ATI SHEET NO.: R-2-6-51

£encv_ SUBJECT HEADINGS: Airfoil theory (06200); Flaps (37450); Control surfaces - Aerodynamics (25600); Airfoils - Drag (08200)

Document» Divition, Intelligence Department Air Materiel Commond

AIR TECHNICAL INDEX Wrifto-fotteraen Air Force I Dayton, Ohio

AÜG

RESTRICTED TITLE: Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semisp-n Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes AUTHOR«): Toll. Thomas A.; Ivey. Margaret ORIGINATING AGENCY: National Advisory Committee for Aeronautics, Washington. D. C. PUBLISHED BY: (Same)

070-6476

(None) OCJO. AOCNCV MO.

ARR-L5B17 Hrirwin» AOSCT w.

OATO

April '45 DOC a a. Restr. U.S. E"g-

Iuuna nosj photos, dlagr. graphs

Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corres- ponding cutouts made in wing were tested for various flap deflections, chordwlse locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80% wing chord behind leading edge, airflow was smooth and buffeting negligible.

DISTRIBUTION: Request copies of this report only from Originating Agency DIVISION: Aerodynamics (2) SECTION: Wings and Airfoils (6)

ATI SHEET NO.: R-2-6 -51

SUBJECT HEADIN6S: Airfoil theory (06200); Flaps (37450); Control surfaces - Aerodynamics (25600); Airfoils - Drag (08200)

Air DocumoflH Division, bitolllnonco Ooporrmont Air Materiel CoEoraond

Aid VtCKNICAL I.392X RESTRICTED

Kfriflht-Pottonon Air Forco I Dsyton, Ohio