REPORT No. 499

WIND-TUNNELCONTROL

RESEARCH COMPARING LATERALDEVICES, PARTICULARLY AT

HIGH ANGLES OF ATTACK

XH-UPPER-SURFACE AILERONS ON WINGS WITH SPLIT FLAPS

By FREDE. WEICKand CARL J. WENZINGEE

SUMMARY

This report coveTsthe twel!h of a se%s oj systemattite8t8being conductedby the ~dio?ud Advisory Committeejor Aeronuuiti to comparedi~ereni tied controldevictxm“th partimdar ri$mm.w to their e$ectwene+wat highangles of attack. The pres.mi tw!a were ma& in the 7-by 10#oot wind tunnel with two sizes of upper-surfaceaileronx on rectangulux Clark Y wing modeik equippedwith fwiL8pan split jlaps. The upper+urface aileron-swereford jrom the upper potiti of the qii.t trai.liqt?d@21Oj tti Wi7198. The ttxts showed the e~ect oj theupper-surface aikron$ and oj the splti jhzp8 on the gen-eral performance characteristicsof iYi4wing8, and on thelateral COntrollalnliiyand 8tabiMy charactenktics. Therewdt8are comparedwiih tlwszjor plain wings wriihordi-nu~ aiikrorwoj similur 8ize3.

Wdh@ps @d, the upper-eurfme aibrorw with up-om!y movementgave rolliq mama at angb oj aitackbehw the 8td thd were reamnubly close to an assumedsatisfacto~ valwe. The yawi~ moments (wind a.xa)were positive (javorable) wiih @rge aileron dq%tiombti, ai all except the lowed angles of attack, they weslightly negalwe (adverse] with small dq%cti.cww. Thecontroljorces were mwch greuhmthan tho8e oj ordinu~ai.lerow oj timi.lur siza havi~ convmti.onulnwvemmt.Wtih t?wjilzp8 dejlectedfor maximum lift, t?wupper-swr-

fa.ce aikrons gave contTolmomentscmwidedly belowthevalue amumedto be sdi.sfactoq. The magni.tud.aoj tbpositive (favorabb) yawing momenh were emal& thanthQ8e with j?Izp8 &T(d ad 9WgdVt? ((Z&MTSe)onesoccurred with smull aileron o?q?ectiow & all angles ofaitazk. Above the 8taU,j%ps neutral or dejlected, bothsiza oj upper-eurjace aiLeron8indicated poor contTol.

The autorotaiw chura.a%-isticsof the wings with the@p8 okjkzted were 8om4?whuikxsjaaorabl.e thun wiih theji’aps retracted.

INTRODUCTION

A series of systematic wind-tunnel investigations,one of which is covered by this report, is being made bythe National Advisory Committee for Aeronautics inorder to compare various lateral control devices. The

various devices are given the same routine tests toshow their relative merits in regard to lateral controlla-bility and their effect on the lateral stability and onairplane performance. They are being tested tit onrectangular Clark Y wings of aspect ratio 6, followedby wings with diiferent plan forms, wings with high-lift devices, and also on wings with variations thataffect the lateral stability. The ii-at report of thisseries (reference 1, part I) denls with three sizes of ordi-nary ailerons, one of which is a medium-sized ailerontaken from the average of a number of conventionalairplanea and used as the standard of comparisonthroughout the entire investigation. other work &athaa been done in this series is reported in reference 1,parts II to XL

The present report covers an investigation of “upper-surface” ailerons, which appear to be one of the sim-plest devices for lateral control of a wing that obtainshigh lift by means of split flaps along the entire trailingedge. Upper-surface ailerons are formed from theupper portion of the split trailing edge of the wing,which is hinged and deflected upward for contiol.The split flaps increase both the lift and the drag of thewing, enabling slower speeds and steeper glides. Refer-ences 2, 3, and 4 give aerodynamic characteristics ofwings equipped with such flaps.

APPARATUS AND TESTS “ ‘

Models,-The model wings tested were equippedwith medium-sized and with long narrow upper-surfaceailerons, together with full-span split flaps havingmedium and narrow chords, respectively. The mainportion of each of the two wing models was made oflaminated mahogany and the split trailing-edge por-tion was made of aluminum alloy. The wings had theClark Y proiile and were rectangular in plan form witha chord of 10 inches and a span of 60 inches.

The narrowwhord upper-surface ailerons were 15percent of the wing chord wide and 60 percent of thewing semispan long. This wing model was fitted witha full-span split flap also 15 percent of the wing chordwide. (See fig. 1.)

463

4 - . ..-. .-—. . ~.— 2 . . . . . —. A. LL.-. —.. . . . .. ——— ..- -—. ––—

464 REPORT NATIONAL ADVISOIZYCOMMPITED FOR AI!IRONAUTIOS

The medium-chord upper+urface ailerons were 25percent of the wing chord wide and 40 percent of thewing semisprm long. A full-span split flap 25 percentof the wing chord wide was used in conjunction withthem ailerons.

Both the ailerons and the flaps were mounted on thewings in such a manner that they could either be lockedriggdly at any desired deflection or allowed to rotatefreely about their respective hinge axes. The gapsbetween the ailerons or flaps and the wing were made assmall as practicable, and then sealed with a light grease.

Wmd tunqeL-All the present tests were made inthe N.A.C.A. 7- by lo-foot open-jet wind tunneL Inthis tunnel the model is supported in such a mannerthat the forces and moments at the quarter-chord pointof the mid-section of the model are measured directlyin coefficient form. For the testing of the wings inrotation, the standard force-teat tripod is replaced bya special mounting that permits the model to rotate

I If A

~--’” ‘/2 ,;:

i I “(l

about the longitudinal wind axis pawing through themid-span quarter-chord point. This apparatus ismounted on the balance, and rolling-moment coefficientscan be read directly during forced-rotation teats. Acomplete description of the abov~mcmtioned equip-ment k given in reference 5.

Tests,-The tests were conducted in accordmce withthe standard procedure, and at the dynamic pressureand Reynolds Number employed throughout the entireserie9 of investigations on lateral control (reference 1).The dynamic pressure ~aa 16.37 pounds per squarefoot, corresponding to an air speed of 80 miles per hourat standard density, and the average Reynolds Numberwas 609,000, based on the wing chord of 10 inches.

The regular force tests were made with several flapdeflections and at a sticient number of angles of at-tack to determine the maximum lift coefficient, theminimum drag coeliicient, and the drag coefficient atCL= 0.70, which is used to give a rate-of-climb crite-rion. The force tests -were also made with a stickgtnumber of aileron deflections, with flaps both neutral

and deflected various amounts, to give data for theaileron rolling- and yawing-moment coefficients. Be-cause of the large effect of yaw on the lateral stibility,tests were made not only at 0° yaw, but also at an angleof yaw of 20° , which represents the conditions in nfairly severe sideslip.

Hinge moments of the ailerons were meaaured bymeans of the calibrated twist of a long slender torquerod extending along the hinge axis from the aileron tothe balance frame outside the air stream. These mo-ments were obtained for various aileron deflectionswith the flaps both neutral and deflected diiTerentamount&

Free-autorotation teds were made to determine theangle of attack above which autorotation wna self-starting with ailerons neutral. Forced-rotation teatswere also made in which the rolling moment while rollingwas measured at the rotdional velocity corrwponding~ p’b

~V= 0.05, the highest value likely to be obtained in

gusty air, and at anglm of yaw of both 0° and –20”.The accuracy of the results presented in this report

is the same as that obtained in part I of the series. Itis considered satisfactory at all anglea of attack exceptin the burbled region between 20° and 25°, where therolling, yawing, and hinge moments are relativelyunreliable due to the critical, and often unsymmetrical,condition of the burbled air flow around the wing,

RESULTS

Coefficients,-The force-test results are givcm inthe form of absolute coefficients of lift and drag nnd ofthe rolling and yawing momentm

c.. !$

u: rofi:;;ment~%,- yawing moment

!zb~

where i3 is the total wing area, 6 is the wing span, and qis the dynamic pressure. These coei3icients are ob-tained directly from the balance and refer to the wind(or tunnel) axes. The results as given am not correctedfor tunnel-wall effect.

The rewdta of the hinge-moment te9ta are given~bout the aileron hinge axis by:

c.= hinge momentqcs

where c is the wing chord. A positive sign of 0“ de-motes a moment tending to make the trailing edge ofthe aileron move downward, and a negative sign indi-~atea the reverse. A positive sign is given to the

WRQ’D+?UNN13L RESEARCH CO=ARING LATl!lRAL C3NTROL DEWWES Mi5

downward deflection of ailerona horn neutral, and anegative sign to the upward deflection.

The results of the forced-rotation tests are given,also about wind axes, by a coefficient representing therolling moment due to rolling:

CA+

where x is the rolling moment m-ured while thewing is rolling, and the other factom have the usualsignificrmce. This coeilicient is used to indicate one ofthe critical lateral-stability characteristics of a wingwhen it is subjected to a rolling velocity equal to themaximum likely to be encountered in controlled ilight invery gusty air. This rolling velocity may be expressed

‘b%

in terms of the wingspan as ~ E 0.06, where Vis tie air

speed at the center section of the wing and p’ is theangular velocity in roll about the wind axis.

The results of all the teats, in terms of the foregoingcoefficients, are given in table I to VIH and in &urea2t09,

DISCUSSION IN TERMS OF CRITERIONS

. For a comparison of the different lateral controlarmngements, the rtxndts of the tests are discussed interms of criterions, which are explained in detail inpart I of reference 1 and briefly in the following par-agraphs. In a few casea it has seemed advisable,as the rcsuh of flight teds, to modify the originalform of the criterion, and where this has been donethe changes are noted. By use of the criterions a com-parison of the effect of the diflerent control deviceson the generrd performance, the lateral controllability,and the lateral stability may be made.

The ailerons used in the present tests are comparedwith each other by means of the criterions, under theconditions with flaps neutral and with flaps deflectedin table IX. In addition, values are included hornpart I for the standard (medium-s@ed) and the longnarrow ordinary ailerons on plain rectangular wings.

GENERALPERPOEMANCE

(.4mnoNt3 NIJumfi)

Wing area required for desired landing speed,—Thevalue of the maximum lift coefficient is used as a cri-terion of the wing area required for the desired landingspeed, or ccnveraely for the landing speed obtainedwith a given wing area. The value of the maximumlift coefficient was practically the same for both wingstested with flaps neutral as for the wings with theordinary ailerons. The maximum lift coefficient wasincrensed from 1.27 to 2.05 with the 15 percent c flapdown 60°, and from 1.26 ta 2.09 with the 25 percent cflap down 46°, (See @s. 2 and 3.) These values are

about what would be expected from the results of pre-vious tests with split flaps (reference 2).

Speed range.-The ratio OLJOD=,R is a convenient

figure of merit for comparison of the relative speedrange obtained with various wings. The value of thespeed-range ratio was slightly greater for the wings&tad with flaps neutril than for the wings with ordi-nary ailerons, the difl!erences probably being due toslight variations in the models within the accuracy ofconstruction. With the 15 percent c flap down 60°

HLYTIHY+++13i: ______f-t

~ 60JP I I I

u /~jQ - r5r~- ,4‘/1’; o— 0° down

“---- %“ --u--l. +——s 80

2.00

t-h-l-i-’’o’ititi

, ,,t

, tI /v /1/ I

0

‘.496 -8 0 8 /6 24 32 40Angle of a{tack, degrees, d

~GUEB 2L-Lff& diw, andcantiOfpm%mefcmwingWfthOdscftm-spnOplft68Pand0.16cbyO.@)bfa nppmarfaca dkons.

the value was increased about 61 percent, and with the25 -percent c flap down 45° the increase was about 66percent.

13ate of climb,-In order to establish a suitable cri-terion for the tiect of the wing and the lateral controldevices on the rate of climb of an airplane, the perform-ance curves of a number of typca and sizes of airplaneswere calculated and the relation of the maximum rataof climb to the lift and drag curves was studied. Thisinvestigation showed that the L/D at CL= 0.70 gave aconsistently reliable figure of merit for this purpose.

,.—.. .—-—

466 RDPORT NATION+- ADVISOIZY

The numerical T&le of this criterion was about thesame for the *._with flaps neutml m for the wingswith ordin~ ailerons. The values were greatlyreduced, however, with the split flaps deflected formaximum lift, and they were lees for all flap deflec-tions &ted than for the flap-neutral condition.

LATERALCON’YROLLAB~Y

(CONTROLS FULLY DEFLECTED)

Rolling criterion,-’l%e rolling criterion upon whichthe control effectiveness of each of the aileron arrange-

6A.o“

IG

2.00 IA./-:~t

/‘‘d ;!, I I I I I I1.60 /,+,,’/ r

II T f

I ./l#-[ A 1 / I 1

3 r.20 I I i I ,’+’.’I, /

I I I I I I I I I I I I I I-% -8

Io 8 16 24 32 40

Angle of a+ fOck, degree”s,o! –

FIGURE3.—Llf&@, andcantiofpmsmmfor wing with 026 c fmllspan split napand 025 Cby 0.40 b/ZUPPW*C3 alle.roxm

ments is judged is a figure of merit that is designed tobe proportional ta the initial acceleration of the wingtip, following a deflection of the ailerons from neutral,regardless of the air speed or of” the plan form of thewing. Expressed in coeEcient form for a rectangukwmonoplane wing, the criterion as used up to themxent has been

where Cl is the rolling-moment coeilicient about thebody axis due h. the ailerons. It appears de&able at

COM&UTTElil FOR AERONAUTICS

this time, as the result of numerous flight observations(reference 6) obtained since the criterion was firstestablished, to alter the form of RO slightly so that inthis report

RC’+L

where Cl’ is the rolling-moment coefficient about them“nd amk due to the ailerons, and onIy changes appre-ciably in value from 01 at high anglea of attack. Thegeneral form of RC, which is applicable to rmy wingplan form, may be found in part I of the series.

The numerical value of the criterion that is assumedto represent satisfactory control conditions is appro.si-mately 0.075, the value given by the standard ordinaryailerons with the assumed maximum deflection of+25° at an angle of attack of 10°. (See part I, refer-ence 1.) As a result of some recent flight tests (ref-erence 7), it appears that a somewhat lower value ofRC’, 0.040 to 0.050, might be satisfactory underordinary flight conditions. Under other conditions,particularly when controlled flight is attempted atslow speeds in extremely gusty air, it is possible thateven the value of O.O75 might not be high enough forentirely satisfactory control. Further flight informa-tion would be of distinct value in clearing up presentuncertainty as to what constitute satisfactory control.

The ailerons are compared by means of the criterionsgiven in table IX for four representative angles of~ttack: 0°, 10°, 20°, and 30°. The 0° angle repre-sents the high-speed and cruisiig attitudes; a = 10°represents the highest angle of attack at which satis-kwtory control with ordinary ailerons is obtained onplain wings; a= 20° is the condition of greateat lateralinstability for the Clark Y wing, and is probably aboutthe greatest angle of attack obtainable in a steady#ide with most present-day airplanes; and finally,z= 30° is given only for a comparison with controlsFor possible future types of nirplanes. The com-pxisons are based on an up-only deflection of 70°,the highest likely to be used, but which gave rLsome-mhat lower rolling-moment coefficient nt an angle of~ttack of 10° than the standard ailerons with an3qual up-anddown deflection of 26°.

At a= 0°, flaps neutral, both eizca of ailerons gavevalues of RC~ greatly in excess of that considerednecessary. With flaps deflected for maximum lift,the values were reduced to slightly below that assumedM satisfactory.

At a = 10°, flaps neutral, both sizes of upper-surfaceii.lerone gave somewhat less than the assumed sr&-;actory value of RC1. With flaps deflected for maxi-mmmlift, the values of RC’ given by the upper-surface

tierons were about 60 percent of the assumed satis-

!actcny values. It should be noted that, with flapslow-n, better rolling control could be obtained byinflecting the opposite aileron down in addition tohe up-aileron (@e. 4 and 5). An equal up-and-downm a differential motion of the ailerons could be used.

WIND-lWNN’EL m!HOxftcm COM2=G LA!I!13R4L CON’rROL DFIVICES 467

IiI

I v 1 1 I 1 1 I I I I I I I I l\

I!\: foIll

I I

““H+‘dL(J -300

v -500+-70”(Signsrefer

I IJJ—LL—I I . l\ IL ffij~~:;

_l

FmcmE &—EoUlng-and yawing-moment cMNJenk of 02J c uppx+nrfam alkon with split tip nenti& and with dap down

468 RDPOET NATIONAL ADVISORY CO~ FOB AERONAUTICS

At a= 20°, flaps neutral, the values of RC’ givenby the upper-surface ailerons were lW than half of

the assumed satisfactory value, but were slightlyhigher than the vahms given by the ordinary ailerons.With flaps deflected for masimum lift the values ofRC’ were so low as to make the ailerons ineffectiveas a source of rolling moments.

At a= 30°, flaps neutral or down, the values of RC’given by the upper-surface ailerons were practically

zero. The long narrow ordinary ailmons with equal

up-and-down deflection, used on the plain rectangular

wing, gave a higher value of RC’ than any of the otherarrangements. (See also part I.)

Lateral contiol with sideslip.-If a wing is yawedappreciably, a rolling moment is set up that tends toraise the forward tip. The magnitude of this rollingmoment is always greater at very high angles of attackthan the available rolling moment due to ordinaryailerons. The highest angle of attack at which theaileron can balance the rolling moment due to 20°yaw has been tabulated for all the ailerons tested, asa criterion of control with sideslip. As previouslymentioned, 20° yaw represents the conditions in afairly severe sideslip. The upper-surface ailerons(flaps neutral) gave rollkg control against the effectof 20° sideslip up to a.qlcs of attick 10 or 2° lowerthan for the ordinary ailerons of similar sizes. Withflaps deflected for maximum lift, the argle of attackat which the upper-surface ailerons gave controlagainst the sideslip was 2° lower than when the flapswere in the neutral position.

Yawing moment due to ailerons.-The magnitudeand oven the direction of the ymving moment de&ablefrom +lerons have not been definitely determined upto the present time. It w-as thought in the past,particularly with reference to acrobatic flying andprobably also with reference to most ordinary ma-neuvem, that to the pilot the maneuvem would seemas if they occurred about the airplane, or body, axes.For a highly maneuverable or acrobatic airplane,therefore, it was thought that complete independenceof the three aerodynamic controls about the bodyaxea would probably be a ddrable feature. Recentflight tests made in an investigation of several lateralcontrol devices (reference 6) indicate that the yawingaction of the ailerons as observed by the pilot is thatwhich would be expected from the yawing momentsoccurring about the wind axtx, not those about thebody axea. It is hoped that a continuation of thisinvestigation, in which some of the most promisingailerons and spoilers developed in the series of wind-tunnel tests on lateral control devices are being teatedin flight, will give sticient information on yawingmomenti to settle the que9tion as to the amount ofyawing moment dmirable for various flying conditions.The indication is, at the present time, that zero or

very small yawing moments about the wind axes are

de&able for acrobatic flying and possibly for flying in

general, but that yawing moments of such a sense

that they tend to retard the low wing in IL turn defi-

nitely improve the lateral control at anglea of attack

above the stall. From the remdts of the above-mcm-

tioned flight tests, it is believed desirable in the present

report tQ give the yawing-moment coefficient in tho

criterion table about the wind axes (0.’), rather than

about the body axes (Cn) as in previous reporh of

this series. The yawing moments are often negative

with respect to the wind axes but at the same ti&e

positive with respect to the body axes. The signs of

the ymving-moment coefficient as given in the tablea

and figuw are in agreement with the N. A.C.A no-menclature in which yawing moments tending toproduce clockwise rotation are regarded as positim.The concept of positive yawing moments as momeutsthat aid the roll (generally termed “favorable”) nndnegative moments as those that oppose tho roll (gm-eraUy termed “adverse”) is also used throughoutexcept as regards the aileron when deflected down-ward. The aileron being at the right wing tip thentends to produce roll in a counterclockwise directionand the coefficients therefore have signs opposite tothose of the up-aileron at the same tip.

At anglw of attack below the stall both sizes of uppor-surface aileron9 (flaps neutral) gave positive (favor-able) yawing moments with large aileron deflectionsbut at medium and high angles of attack they gave verysmall negative (adverse) moments with small deflec-tions. Just above the stall the yawing moments werenegative (adverse) even with large deflections. Themcharacteristics are definitely better than those of cor-responding sizes of ordinary ailerons. With flopsdown for maximum lift, the magnitude of the positive(favorable) yawing moments were smaller than thosewith flaps neutral, and negative (adveme) yowingmoments occurred with small aileron deflections atpractically all angles of attack.

LATZEALSTASILITY

(AILERONSNEUTRAL)

Angle of attack above which autorotation is self-starting,-This criterion is a measure of the range ofangles of attack above which autorotation will startfrom an initial condition of practically zero rai% ofrotation. With the split flaps neutral the limitingangle of attack was the same as for the wings withoutflaps, but with the split flaps deflected for maximumlift the limiting angle was reduced 3° to 4°.

Stability against rolling oaused by gusts.-Tostflights have shown that in severe gusts a rolling veloc-

ity may be attained such that ‘lb =0.06. Con-2V

sequently, the rolling moment of a wing due to rolling

fVIND~L RESEARCH COlbAItRiG LME@AL COltl.’liOL Di)ViCl!l$ 469

at this due of@~V gives a measure of one factor

aflecting lateral-stability characteristics in rough air.

In the present case, the angle of attack at which this

rolling moment becomes zero is used as a more severe

criterion than the previously mentioned angle at

which autorotation is self-starting, to indicate the

practical upper limit of the nsefd angle-of-attack

range. As in the case of the angle of attack above

which autorptation was self-starting, the angle of

p’binstability while rotating with — =0.05 was the same

2Vfor the wings with split flrLps neutral as for the wings

without flaps. With flaps deflected for maximum liftrtt0° yaw (fig. 6), the angle of attack for initial instabil-ity wna 4° lower than for the wings with flaps neutral.

$,.%7.~

0!

$~

c

?# .:.

~oL~..$$

~ngle ~f a{ fock, de-grees,d

FK3UXEO.-Rollfng-mom@ csdklent due ~ mlb at ~-o.~ fa wfw wfth

0.16c fnlkpm spilt 136Pnentral,and wftb fip down. W yaw.

With 20° yaw, the wingg with split flaps neutral,like the wings with ordinary ailerons, had an angle ofattack for initial instability 6° or 7° lower than thatwith 0° yaw. With the wings with split flaps deflectedfor maximum lift, the angle for initial instability wasshifted to negative values so that the wings showed adistinct tendency, at all normal angles of attack, toincrease an initial rate of rotation in roll when thedirection of motion of the roll and the yaw were of thesame sign. (See fig. 7.) This characteristic might bempectcd to impair the lateral stability of airplanesequipped with split flaps.

The precading criterion shovm the critical rangebelow which the stability is such that any rolling. isdamped out, and above which instability exists. Theremaining lateral-stability criterion, maximum G,indicates the degree of the maximum instability. Allthe rotation teats showed somewhat unsymmetrical

conditions in the two directions of rotation, and themaximum value of CXfound with any angle of attackin either direction of rotation is used as the criterion.At 0° yaw, the wings with split flaps neutral had thewme mtium tendency to autorotate as the wingswith ordinary ailerons but, with split flaps down formaximum lift, this tendency was increased somewhat.

The maximum autorotational moment at 20° yaw isof importance for the condition in which the airplane isdridded and the forward wingtip is rolled upward or therear tip downward by means of a gust. This autorota-

Angle of affack, degrees, d

Fmwm7.-RoUing-moment cnatTJ.entdue to rollfng at~O.OS for wfng vdtb

0.15CfUU4P3nsplitfillIMIItd,and with fhp down. –W YP.W.

tional moment, which is large for the wings havingsplit flaps neutral and for the wings with ordinaryailerons, increased slightly with the narrow-ohordflaps deflected for maximum lift and decreased slightlyfor the medium-chord flaps.

CONTEOL FORCE BEQUIBED

The hinge-moment coefficients Ior the two sizes ofupper-surface ailerons are plotted in iigures 8 and 9 forboth the flap-neutral and flapdeflected conditions. Acontrol-f orce criterion, with which the various lateralcontrol devices are compared in regard to the control-stick force required to attain the asanmed maximum

. .. .. —...——. —— —-

470 REIPORT NMPIONAL ADVISORY CO~ FOR AJ3RONAUTICS

deflections, is based on a control-stick movement of+ 25° and is independent of air speed. This critarionis

I

ment are about three times m great as those of ordi-nary ailerons of corrwponding sizes with equal up-anddown movement (split flaps neutral). Comparedwith the ordinary ailerons having an up-only move-ment of 70°, however, the values of CF for the upper-

C3*.$....:L:“u

1:“;$$-.

FmIJRE&—Hinge-moment c@30fents of 0.16c nppw+arfam aikon wfth splitflap nential, ond with:lhp down.

where F is the force applied at the end of the controllever of length 1, and &/25 is the gmr ratio betweenthe aileron and the control lever.

Values of CF are given in the table of criterions(table IX) for the two sizes of upper-surface ailerons

surface ailerons are about the same. (See part I,reference 1.) These values are much too high forpractical operation, and an investigation of methods forreducing the hinge moments of upper-surface aileronsis now under way in the 7- by 10-foot wind tunnel.

.

@..

2<.$-%.Q)oc1u.c

;

&&b

s

-.Aileron de flecfion,degrees, 6A

FIGURE9.—Hfngem0meatuwllidenta of 02S c npparamffme afleron wftb splft tip neutral, and with ffap down

with split flaps both neutral and deflected for maxi- One possible method of reducing the control forcemum lift, and for the two corresponding sizes of ordi- might be to rig the upper-surface ailerons up o smallnary ailerons. At a=OO and CZ=lOO, the values of amount when neutral and to provide them with an

CF for the upper-surfam ailerons with up-only move- ordimq differential movement, although this might

‘iKEND-7XJNN13LRESEARCH COMPKRING LATERAL CONTROL DEIVICES 471

cause a small increase of minimum drag. A pre-

liminmy investigation indicated that this method

was not very promising and other arrangements

which appear more satisfactory are being investigated

in the wind tunnel.

With the split flaps deflected for maximum lift,the values of OF for the upper-surface ailerons ata==OO rmd 10° are reduced to nearly the same ssthose of the ordinary ailerons with equal up-and-downmovement, on account of the reduced speed at thesame angle of attack.

It will be noted that for approximately the samerolling control the values of CT’ are considerablysmaller for tho long narrow ailerons than for themedium ailerons.

CONCLUSIONS

1. With the split flaps neutral, the upper-surfaceailerons gave values of the rolling criterion R(?’reasonably close to the assumed satisfactory value at

anglea of attack below the stall. With the flaps

deflected for maximum lift, the rolling control was

considerably below the assumed satisfactory value,

but might be sticient under ordinsxy flight condi-

tions. Above the stall, little or no rollhg control

was indicated with the flaps either neutral or down.2. At angles of attack below the stall both sizes

of upper-surface ailerons (flops neutral) gave positive(f~vorable) yawing moments with large deflectionsbut with small deflections at all except the lowestangles of attack they gave small negative (adveme)yawing moments. Just above the stall the yawingmoments were negative (adverse) even with largedeflections. With the flaps deflected for maximumlift the magnitudm of the positive (favorable) yaw-ing moments were smaller than those with flapsneutral, and negative (adverae) ones occurred withsmall aileron deflections at all angles of attack.

3, The control forces required to operate upper-surface ailerons with up-only deflection would be toogreat for practical use.

4. The autorotationrd tendencies of both wingswere somewhat greater with the flaps deflected thanwith them retracted. With the flaps deflected andthe wings yawed, a tendency to rotate in one direc-tion was shown throughout the entire usable angle-of-attack range, a characteristic that might be ex-pected to result in some impairment of the lateralstability of airplanes equipped with split flaps.

IJANGLHY MEMORIAL AERONAUTICAL LABORATORY,NATIONAL ADVISORY CO~~ITTEE FOR AERONAUTICS,

LANGLEY rrELD, VA., June 8, 1934.J01+*31

REFERENCES

l. —” -------- “--”-–-

2.

3.

4.

5.

6.

7.

Wind-”1’nnnel Jiesearoh Uomparing Lat8ral Ucmtrol Devices,Particularly at High Angles of Attaok.

I. Ordinary Ailerons on Rectangular wings. T.RNo. 419, N. JLC.A., 1932, by Fred E. Weiokand Carl J. Wenzinger.

II. Slotted Ailerons and Friae Ailerons. T.IL No.422, N.A.C.A., 1932, by Fred E. Weick andand Richard W. Noyea.

III. Ordinary Ailerons Rfgged Up 10° When Neutral.T.R No. 423, N.A.C.A., 1932, by Fred E.Weick and Carl J. Wenzinger.

IV. Floating Tip Ailerons on Rectangular Wings.T.lt No. 424, N.A.C.A., 1932, by Fred E.Weick and Thomas A. Harris.

V. Spoilers and Aileronson Rectangular Wings.T. R.. No. 439, N. A. C.& 1932, by Fred E.Weick and Joseph A. Shortal.

VL Skewed AUerona on Rectangular Wings. T. Il.No. 444, N. A. C. A., 1932, by Fred E. Weiokand Thomas A. Harris.

VII. Handley Page Tip and Full-Span Slots withAilerons and Spoilem. T.N. No. 443, N. A. C. A.,1933, by Fred E. Weick and Carl J. Wenzinger.

VIII. Straight and Skewed Ailerons on Wings withRounded Tips. T.N. No. 445, N. A. C. A., 1933,by Fred E. Weick and Joseph A. Shortal.

IX. Tapered Wings with Ordinary Ailerons. T.N.No. 449, N. A. C. A., 1933, by Fred E. Weick andCarl J. Wenzinger.

X. Various Control Devices on a Wing with a l?kedAuxiliary AirfoiL T.N. No. 451, N. A. C. A.,1933, by Fred E. Weick and Richard W. Noyes.

XI. Various Floating Tip Ailerona on Both Rectangu-lar and Tapered Wings. T.N. No. 458, N. A,-C. A., 1933, by Fred E. Weick and Thornas A.Harris.

Weick, Fred E., and Harria, Thomas A.: The AerodynamicCharacteristic-s of a Model Wing Having a Split FlapDeflected Downward and Moved to the Rear. T.N.No. 422, N. A. C. A., 1932.

Wenzinger, Carl J.: The Effect of Partial-Span Split Flapson the Aerodynamic Charaoteristhx of a Cfark Y Wing.T.N. No. 472, N. A. C. A., 1933.

Wenzinger, Carl J.: Wind-Tunnel Measnmmanta of AirLoada on Split Flaps. T.N. No. 498, N. A. C. A., 193A

Harrfa, Thomas A.: The 7-by 10-Foot Wmd Tunnel of theNational Advisory Commit% for Aeronautfca. T.ILNo. 412, N. A. C. A., 1931.

Weick, F. E., Sou16, H. A., and Gough, M. N.: A Flight In-vestigation of the Lateral Control Characteristics ofShort Wide Ailerons and Various Spoilers with DifferentAmounts of Wing Dihedral. T.FL No. 494, N. A. C. A.,1934.

Souh5, Hartley A., and Wetmore, J. W.: The Effect of ‘Slotsand Flaps on Lateral Control of a Low-Wing Monoplaneaa Determined in Flight. T.N. No. 478, N. A.C.A. ,1933.

472

TABLE

REPORT NATIONAII ADVISORY CO~ FOR AERONAUTICS

I.—FORCE TESTS. CLARK Y WING WITH 0.15c BY 0.60 ; UPPER-SURFACE AILERON

FULL-SPAN SPLIT FLAP

(Valnmaregivenfor oneallenmat @t wingtip)

R.N.= @M,MO. VdodtY-@3111.P.h. yaw-~

AND 0.16c

I a –l& –lr –P -4” –?? 0“ & lW Is+ 14” le I& Ir 16” w * 26° w 4~

I l~Al AJLERON AND FLAP NEUTRAL I

CL @ .–-–.. .–.—- fMu.& :l& 0.190 ag 0.7@3CD v -.—– --—–

L lW ------- Lm.016

l.. l.% l.. Llb2.Ml .0x3 –.- . . . . . M

:l& ~ll& :~ C!g am. !a17 .729

II RIGHT ~ERON UP-FLAP NEUTRAL

[ f~~ g :~~ :;j$j:~j]:~;f:!::--’-‘ ‘0.M7 ----- .—.-. am ------- am –-------—– –O-cm ------- ------- . . . . . . . a(

.Cm ------- –.0)1 .-.-.. –––.-.- .Coo ----- -:2.- ----- -:[”” --

c{cm’c{c,’cl’CL’Cf%%cm’

gq 1.__.1 a~

“M001 . . . . . . -. m

-.001–. E ::::::: .m

.at5 ------- -. m-. w .-..-.. .am

.01s . . . . . . . -. Oos-. m ------- .Wo

.023 .-.-.. –. mal.013 ---— .Um ---—- ----

l“ I I I I II.Ual .. . . . . . . . . . . . . . . . . . . . –. m ------- a

1 I

10.537

.luo

.a@-: %

.Ow

0:~

,Ow.Coil.W1.OY1

I RIGHT AILERON UP—FLAP DOWN W

T 111111

CL $ –:= -CDCf w ----.–.CL’ ?$ ;____Cf —----c%’ ?w -------

I

.0.111.am.046.011.079.m H

am ------------ am LOll L419 LS49 LMl.Wo -------------- .Wa .104 .107 .223.6W -------------- .643 ------- -Ml --:!!. .U.3S..m ------------ .8X2 ------- –. m -------- –. WI.070 –.-.. ------- ; ~; ~::-. .075 ------- .W7.017 ––—–--...

L4.S5.M5

. . . . . mL462 ------ L413. m .------ .3M

. -- —----------------

:% am o.cw.470

.% . . . . . . . –:%!

K1l.._-.l .%

am-: %!-: %-.001

LM6.Tzo,001

-.001.ml

-.001

, ---. --———.—..—.. .-—---—--------------------------

–: Cl& 1.-...1-.0

---1 m 1-----.-1-.ml -------------------------- -m 1-------1-.0

I RIGHT AILERON UP-FLAP DOWN $Y’

t t 11 1 I I t I I I I I I Iw -a m aas9

.mls _L!!?- .0413iP -—–.- ;C&m .-.—-.m ---—– .02s

a492 tl-------6..4L2f0 l.. l,, l.. ~~ L6S4 .-..-.. L461.079 –—–----–-

1.348 L6S3 L098. 16s .24 ------- .375 .411 .6@ .652

.642 ------------ .044 ------- .W .040 -.–---. --.-.. –..–.–.----- . . . ..–

.W . . . . . ..–- . . . ..m . . . . . . . . m

I

00.073 ------- ------ .077 l–:–-l -cm.016 –—-- ------ .

hll

I I I I I I I I I I I

,:----.---::1::::3 .Wll L---l –.mj –.W .-.-–– --------------------------- –:K& ~:..:.: :1

.075 -----------------------------------.m l–—–l .033 –.m ..–.-.. .-..–.---,---- –—.. .-.-.– –. m .-:.:. .000

l– RIGHT IJLERON UP-FLAP DO~ ~

0.655--------r-.... l-g L410 L767 LM L676 l.g l.. ;:::--- +%7 L lW.U41 ----- ------- .x@ .3.50.OQ --—– ----- .046 ..:?- .0i4 .041 -2?!. -------- ------ --:1 -------

.4567

.ml.m4 ----.-- --.--– .W --:--- –: g –. W9 ------ ------ –.–– ----- ------- –: g.075 ------- ------- .079 ------- .076 –--–-- ------ ------ ------- -------.014 ------ ------- .m –-–.. .W –. m ------- .-.—– ------- ------- . . . . . . . –. ml

l.. L 110.609

------- .ml. . . . . . . .ml. . . . . . . .ml. . . . . . . –. au-1

1:$12# a673.040.Oal

-: % _: ~.Ow,Oal -,001L

CLCDCfc.’Cfcm’

I I

I RIGHT AILERON NEUTRAL-FLAP DOJYN6F

CL w -ag :g am -.--– ..-.–- : ~4 L~ LmCD r

z Olz 2cw :g LS43----- L434 l.. l.. l.% 0.m a.s41.163------- –-.–- .270 .3E3 .407 .Uo .440 ------- .469 .W3 .869

I RIGHT AILERON UP-FLAP DO~ @

c{g;c.’c{C&

CL’Cfcm’CfCL’ 1-—----.—..—-—-.——.—----.—--—.—--.-—-—-—----——-----—-——

am.m.Ols.m.a?a.m.w?.m.lm.Ol!z.076.019 11

am -.--—–—--.m --–.-_--—.OM .–-. ------.cm –—.- .-.-.–.Ct31.--. ---––—.m —---------.644 –—– —.-.aa –––-..-.–.W_5 ---------.W7 -.--–-.-.–-.076 .–-. ------.O11 –—.- .–-..-

aan–. all

. 0L5–. all

.ml-. W.2

–:%.@37S&

.(3331—-..—--------.—————.--.—--.—--—-----------.--.—-—--————

aoo7–. Ooa

.015–: C&

–. m

–: E

–: g

–. m

am 0.all.mo -, ml.Ow _: g

–. ml-.001 .am-.001 -, W1

-: E -: %

-: % -:3#am -: g

-. ml

am-. m

.014–. au

–:%

–: g

–. w.076

–. WIm.-.—.-.——-.-.—------.-.---..——--------------.—.-——.---.—--——————---——--——.——-----.-.--.—-.-.—..-.-------—--—.--.—.-............---—--———.-.—---.—----—-—.——.———-.-..—.-..-.—.-.—.-.———-.—-.—---—----._--.---..—_.-..-.....---—----------------—--—----------------acre . -..-.. -a 001–. ml .-..-.. .ml–. 001 .. . . . . . –. all–: ~ ;:.::-. .Oal

-. CQl–. 062 -:..: -. ml

.001 .-..-.. -.001-: ~ .:..:: - -. ml

.W?–. cm :.:-: :. all

.am . . . ..-. .cm

.m . . . . . . . –, 001

RIGHT AUERON DOWN-FLAP DOWN W

c{

I I

tr ------ -am -aW7 ------ ----– –a m --.-–- -ag -a an ---–-– ––.-.-–—-- –.-– -.-— am -c%’ P .—.–- -.031 .m ----- ------ .CIxl -------

-... -a cm

Cf M’ -—---.all ---–-– --––----—– –—– –...- .Wo ----- -. ml

–. 017 –. m –—-- –-.– –. 021 ------- –: ~ –. 010 -----–- –––– —-- ----- ----- .m ----- –. axc,’ w .–.–.. –. ml . ml –-.– –—– .W2 --.-–. .W4 --.–— .–.–-- --- –—– ..-.— .W1 ------ -. ml

\g aam

-, m -: %.Om -. ml

WlND+lWNN33L RESEARCH COldTAFUNG LATERAL CONTROL DEWICES 473

TABLE 11.-FORCE TESTS. CLARK Y WING, WITH 0.15c BY 0.60 ; UPPER-NJRFACI.I AILERON AND 0.15c

FULL-SPAN SPLIT FLAP

(ValuMaregivanforone olkmn at rfght whg tip)R.N..~,ow. Velo&y-80 m.p.h. Yaw=–ZY

a –16” –lW –6” -4” –P r P I@ l% 14” 16° I& 1P lW w 2P w &y @

,

8A AILERON AND FLAP NEUTRAL

cL w .... ... . .. ....- :~ II% :;: U& CLg :$& : yi7 :% y ;OJ l.. : ;8J 1.190cD @ ...-..-. ... ..-.of m .............-.–:g –:Gl&–:g –:CC& –:g –:g –:g –::~fi–:yg –:g –:~ –:%cm’ @ ...-.-.....---- .017

RIGHT AILERON UP—FLAP NEUTRAL

7@ . . . . . . . . –.... - 0.M’4 ------- ------- am ------- am -------- -------- 0.cm ------- -- . . . . . ------- am -...._ afea a OIS -aMz%’ 7W . . ..-.– -..... - .017 . . . . . . . ------- .011 ----–– .001 . . . ----- -------- –.om ----- ------- .-..-.. –. m ------- -.014 –. m –. ~

AILERON NEUTRAL-FIAP DOWN W

CL w -: ~ WJ 0.002 ------ ------- LW3 l.. 1:O& l.. l.~ l.. 1.818 --.---- LWO :% L lWcD o“ . la –.-.. - ..-..– .IEa .393 .-.---- .m

L160 l.. M&l

Ct w -: ~ –: &4 –. 017 ------- ------- –. 018 –: g –: g –: $2.2 –: :g –: ~j –. 048 ------- –. Ma –. la3 –: E –: Mc’.’ ‘P .0s3 ..--... ------- .W4

-: &i –. m.CQo ------- .019 .035 .041 .049 .O.51

RIGHT AILERON UF-FLAP DOWN O(P

cl’c,’ .016 ------- -------

-0. ml–. ml

TABLE III-ROTATION TESTS. CLARK Y WING WITH 0.16c FULL-SPAN SPLIT FLAY

c~b@Wnmf0*”mti0n8t~$-am([tl=&%%tin

AileronmntraRM.=m,~. valDdty-80 nLp.11

a –lW –P W’ Iw H’ 14” 10” w W w w w w 403 IFLAP NEUTRAL-YAW-V

‘m ‘-------7m m ‘-m-ml

(+) Ihhtto. (IJ~)------------––– q ------- -------- -0. ~ -------- -Q. 018 -0.015 -0. m am ----–– 0.024 am am ao~o -am

(-) EoMIon (mu. ‘edmhW ------ CL -------- -------- –. m -------- –. 017 –. 016 –. m

FLAP DOWN W—YAW=W

(+) Rotation (okkwfse)---------- CA -0. m -------- -0.027 -am -0. ml a 018 0.049 a 019 ----–– am .-.---.- amo a WI -a 001

(-) Rotation (munkokkwfw)----- Cx -0.031 -------- –. m –. Ols –. 015 .fr22 .021 .016 -------- .m .------ .m . ml ml

FLAP NEUTRAL-YAW=-2V

(+) Rotation (cJI@3YM-------- CL -------- -------- -0.023 –-.-... -a m

r H

W –o.m -a 007 -------- -0.072 -am -aC87 071 064

(-) RoWfon (wuntmdmkwim)----- CL .-..–.- -------- –. 017 -------- .W4 .011 .026 .044 ------.- .C82 .C&5 .W7 .077 .04u

FLAP D OWN W—YAW-–W

(+) llo’otl”n (dmkwfta)--------- c,

.Cy .W1 .Onl .02 .m .097 ..1 .m.101 .023

-a 043 -0.0-46 -0.046 -ao44 –so-44 -0. oM -0.078 -am –a CQ4-o. Qa5 –0.0J4 -awa -ao72 -ao6s

(-) Rotatfon (mu. ~)------ C, –.~ 0

474 REPORT NATIONAIJ ADVISORY CO~ FOR AERONAUTICS

TABLE IV.-HING=N1OLIENT COEFFICIliX?TS. CL.412K Y wING ‘iVITEt 0.16aBY 0.130: tJPPER+URFACE ~LERON

AND 0.15c FULL-SPAN SPLIT FLAP

(C. 19givem[oroneailaronat *t wfngtip)

R.N.. w,am. VeltitY.@) m.p.h Yaw-O’

I.-1aA 1P P @ –5” –W’ –m -w –m –m

bl FLAP NEUTRAL .

w .—. -. —.-.. .-. -.. -. —.-KP . . . . . . . . . . . . --..---—

I

{;B :m

w —---------- —---------- –. m –. m

amam

–. ml

–. ml

–. ml

–. ml

FLAP DOWN W

–o.Om4 am am am a Wlo–. m .ml–. Wx4 –. m –: E –: E–. m –. am

–: E–. m –. m

o 0–. m

o1°

0

a CII14.mll.aO1

o0

.m

–. ml

a 0321. W19.OmJ.m.m

.m

–. ml

o.ml

Io.mto

.m .CuM

.0319 I .Wz2.mf2.mllo

.m . al12

.m .Wlo

o –. ml

0:g:

–. C&’o

0.0s21 CLcKr& 0.UQ7.0)18 .fxB5.CO?3 . Wlo .W14. ml .am

o.m

o 0

TABLE V.—FORCE TESTS. CLARK Y WING WITH 0.2.% BY 0.4~ UPPER-SURFACE AILERON AND 0.26u l?ULL-

SPAN SPLIT FLAP

(vafnm am gfvmfor me afkon at light Wfng tip)

12.N..oW,~. VeJodty-&l m.pk Yaw-cP

a –W –MY I —5” -4” –v w b“ Iv I& I 14”m 1Plr 18” w 2P W w 4P

8A AILERON AND FLAP NEUTRAL

CL w —CD

\ ~7 l.. :% 11~ O.soa.7X

RIGHT AILERON UP—FLAP NEUTRAi

G{. ] $ 1:=:::: ~=::. a% -----– -–-–- ag ~z::: arm -–--.– -------- CLooal.-.-.--l.-_--~.....-l Il,lyo I.... JCL Oql l-a 016------- ----—

KF --_.--- --––—.am ------ –: 3 ::::: _;:: _-:-

------- —.-— .012 ------ . —-—— .aKF ------- — ------- ----—w -–----- ._ L:- ------- ——— .aM ------ :0?2 .—–– -----– :0w ______ .-—– -.—. — --.—. .W2 ----- –: g –—– .-.–— –. a* .-.._ .- .__.- _- —- -——w --––.. -—---- -.——- ------w . ------- ----

:% -:::: X&3 .~:: +:;z –: E :::::: .:::: ::::::: –: &j 1:::::::1 z;--—--- ---

w ..–---- --–----—- -— .a32 -..... - . . . . . . . .-...-.

--—. — -..—— . on -----–m .—---- -------- ----- -----

.C03 .—.-.. -—_– .Uml -–.-- --.-..- .-..---.W8 ----— .On ------ --—--~m. m . ------- —---

.W -–... - . ----- . . . . . . .. (u+ ----- _____ .019 -----– .U!3 –.----- --–-.-- .MQ ------ ------- . . . . ..- . m ------- –. ml , .W1

RIGHT AILERON UP-FLAP DOWN F

t

–. lm -.010–: K .:::::: -.004 . Mli

.ot2 . . . . . . . .Qos -.cnl

0s1 .Om

CL%

,am#cm.mo.Cw.Olm.Cm.O111. ml.Wo.W

cL

1 II

w ..––.. -&COO anl -.-.--- ------ a 701 :c9J L402 LSSJJcD v -.–.--- .W.9

L 691 1-~ L W ‘ . . . . . . L 418 1::1 1:~ 1:g CLtaJ a~

cl’.am .----.- ------ .078 .I&9 .213 .X3

m .—.-.. .OM ------- ---.--- -------.290 .--..-- .3Z

.039 -.- . . . . .0s –-----. .010 -1?-------- ------ . . . . . . .w --_.-.-

.G20 . . . . . . .. of2 ------- -–.-.. -----

.027 –. Mil -: z.m ------- –. ml ------- –.alt -------- ------------- ------- –. m ------- -. m ;g . ml

%’ 7W .-–.-.. .073 ------ ------ _____G’ m .--.--–

.074 ------- .0s5 ._-----.031 _-.--- ----- ----

.0s3 . -------------- . . . . . . ------- .Oto . . . ..– .012 .Wo.016 ------- .QM ------- . W . ------ .--–– . . . . . . ------- –. ml . . . . . . . -. m -. ml .m

RIGHT AILERON UP-FLAP DOWN ?&

cL I

4

v -a% M&I afire .-_.._ ----- a973 L344 L713 L843 L~CD .117 .—-— ---––

1

LW . . . . . . . . . . . . . . L402 L@XI LGW l.% :g C$.E#

cl’ G .-..––. W .217 .=

.047 ---—– ----- -----.2-M . ..-. -..-. –.. .416

c.’.042 -----.- .0t4 ._:?- _l?- -------- ------------- .–..-.

m .-.-–:Z ..::!!.

.010 —.–-. ------ -..-.–.Wo

.ml -—--- –. @It -------- -------- -------- ------- ------ ------- –: ~ . . . . . . . –: E ;c#. ml

$ ------- . a31 --..--– —--- .–.-.. .077 --------. ml

z<.a37 -------- ------- -.... -.. ------ ------ -.- . . . . . . . ..-. .W

----—. .0%3 -------- ------- ----- .014 --------. ml

.Wr2 -—–– .-.---.- -------- ------------- ------- .ml . . . . ..- -. m -. ml -. ml

RIGHT AILERON NEUTRAL-FLAP DOWN 45°

cL IllW-agUJgaem ---------- <;~ : ho LE23 LW :@& l.% l.. l.. L 176CD v .am –.––---–– .3s7

: ;OJ L 100.47

L 1!25.024

cl% 0.816. 7U .724 #w

T7TND-’ITJNNEL RESEARCH COMPARING LATERAL CONTROL DEVICES 475

TABLE V.—FORCE TESTS. CLARK Y WING WITH 0.25c BY 0.4($ UPPER-SURFACE AILERON AND ().%c FULL-

SPAN SPLIT FLAP—Co;tinued

a —16° —lo” -6” 4“ -& w b“ l!?’ @ 14” Is” lef ‘ v w m 22’= 25” W w

RIGHT ~ERON UP—FLAP DOWN W

Ctcm’cl’c.’Cfc“’Cfc“’01’c.’cl’c.’

.... ----

.-......

........

.-....-.

.-....-.

.-......

.-......---------------.................... ----

.. --.... ... ---- ......

..------ . ...... -----1--------..--------------------.......---------------..-—---......---------------...----.M93.013. WI.U23.033.Oa

“)11.mo-..:-----------------.070---------------------.017. . . . . . . . . . . . . . . . . . . . . ..CC32. . . . . . . . . . . . . . . . . . . ..-.025 ----------------------

O.cm . ------.Cm -------.014 . . . . . . ..ml -..-...

–: E

--l.cm _--:.M6 -------.m .-.--.-.07a .L_-.015 -..... -

2 -------- ------- ------ ------- —.a) --------

1 111’

.013 -------- ------- -...-..---.--- .ad -------- –. mt -------- ------- . . . ----------- -. all -------

I . ------ .027 --.----- .am -------. ------- -------------- . U31 .------ :E .031 .(. . . . . . .

9 1..----..1.--. ---1-.-.-..1 .0t3 -------–. W5-------- –. an -------- .-.-... ------ ------- -ml ------- –. ml

. yl .--.----.W1 –.(

.040 -------- ------- ....-.------.- .COt -....-. .* .flll .000‘-3 . . ..--.. ------- . . . . . . . . . . . . . . –.a-

1 -------- ------- ------ . ------ .0101-.--1 . aQ I

! l--------l--–---l-------l-------l o-oa 0)7 -------- a m ‘ I I I I–:~; --------–.~–.c-

a cm–. m

.O-JJ

~c& _a~

.all :mo

.Om –. ml

%’!

xl::;%’, ,w -----------------------------

.CrM -----------------------------

! l--------l-------l-------l-------l -“ !2

021--.4 .Lnml

CR -------. u16 -------.032 -------

I–. m–:%

:Cim .Cti. m :g.m.0)2 .W1.ml .m

RIGHT AILERON DOWN—FLAP DOWN 46”

Ct

1

n“ -------- -o. Wt9 -------- . . . ---- . . . . . . . -0. Im9 ------- –. m7 --––.. -a m . . . . . . . . . . . . . . . ------- . . . . ..- am ------- 0.ml a~ clc&c.’ b“ . ..-..-. .m .-.----- ------- -------cl’

. Ixll -------1.5” . -------

.031 -------- .001 -------- ------ -..-.-- ------- .WI ------ –. 001-. alt -.- . . . . . ------- ------- –. m4 ------- –. am -------- –. 011 --------------- ------- .-–...

c.’ 16” .. . . . . . . .m ..-.. -.. . ------ -------.ml -------

.0)1 . ------ .0)7 -------- .OM --------------- ------- ------- . W1 ------- –: E.ml.ml

.m

.m

TABLE VL—FORCE TESTS. CLARK Y WING WITH 0.25c BY 0.40 ~ UPPER-SURFACE AILERON AND 0.25c

FULLSPAN SPLIT FLAP 4

(values we gimn feronedlenm atright wing tip)

R.N. =IWI,OII. Veldt y-Sl m.p.h. Yaw- –~

- t ‘8.4 AILERON AND FLAP NEUTRAL

0“ . . . . . . . . .-.- . . . . aom a~ a o% C&& :$ 0.W7 I.@&%

L 116$ .-..-... -------- .Om

L 14S L 1.53 L 173 L 181 1.17W am aw.6 0.’%2!4Cf

I-:$$ –: ~ –: E

.119. ------- . . .. —-

C.’-: ~ -: ~ -. m

w . ..-.-.. ------- .m–: g -: HO -: g –: ~ –; # –: ~ -: ~ –:: –: ;;;

.Ixe-: %

.033 .am

a m3

–: E.M3

RIGHT AIZERON UP—FLAP NEUTRAL

cl’ w .. .. .. . . .. . ..-. ao54 -------------- awa ------- am-a ----...- .- . . .._ ao76 -------------- ------- 0.074 acu9 _____ aau –~gcm’ m . . . . . . . . -------- .021 -------------- .017 ------- .007 -------- ...----- .W1 -------------- --.---- .am -.m7 –..--.. -.012

AILERON NEUTEtA&FLAP D OWN 4P

aL @ +. ;~ 0:~ am ------- ------- : p2 I. 3?s L 042cD w .16$ ... .. .. -.- .. ..

L 74S L849 LW7 L833 L 543 L 2Q5 L220 1.Ea 1.lz L036 am

cl’ w –. 012 -.016 -.ola ..-.-.. ------- –. 019 -: % –: E -: %?c“’ w .0)7 .m

I–: %’J -: #4 –: g –; g –: g –: g -; g –: E -: % –: ~~

.02s -.... -. ------- .Om .0a3 .010 .011 .049 .m

RIGHT AILERON UP—FLAP DOWN 46°

cl’ w -------- aon ---------------------- 0-070 ------- ao77 -------- son -------- ------- aa33 ------ awo ------ am a 010 -awc.’ w . . ------ .027 ---------------------- .Om ------- m -------- .Oot -------- ------ .ml ------- -. m ------ –. cot –. w -.012

476 RDPORT NATIONAJ ADvIEORY CO~ FOB AERONAUTICS

TABLE V1l—ROTA’IT.ON TESTS. CLARK Y WING WITH 0.2S0 FULW3PAN SPLIT FLAP

Cl is givenforformalmtatfonat*~[~] ‘titi~&~&im

Aflemnnentrd

R.N. -@Eylb3. Veloeity-W M.p.k

a —10” –’9 v 10” H? 14” Iv IT I& w n“ * 24” w w 4(Y

1

IFLAP NEUTR&YAW-V

---1- -1= ~1=-1= ~ ~ ~ = ‘------ m ‘------- m -W1 =

(+) RO~~OU (d~) c, ------– -.–-.--.-am ------- -0.019 -0.016 -o. m ------ &w O.aw . . . . . . . . 0.04s -------- Owl -a m -o. m

(_ay @=@-C, ----

FLAl? DOW 4&-YAW-iY

(+) Rotation (Ckkwfse) Cx

i

-a m —------ -0.027 -&oin –am

H

a 0L5 0.047 soul ao14 0.W4 -------- ..-- . . . . am . . . . . . . . 0 -a ml

‘-JO&%;--.––on (mnnter-

C1 –. 031 ------- –. Ola –. 016 –. 016 .021 . as7 .0a3 .011 .m7 -------- -------- .UM -------- .m .001

FLAP NEuTRAL-YAW= –W

d= =1----- ’014 ‘------- ml i“, ‘------ ~ m ------ ~ - > ~ =

(+) Ro~~On (d-) Cl -------------- -am ------- -0. m -aoio -ao.is –-.-.. -am -a 007 -------- -a 07s . . ..-.-. -aag2 -aom -0. 06s

FLAP DO- W—YAW-–W

(+) Rotation (dmlmrke ) cl -0.0’42–a WO–a 040 –a 044 -a044 -a 047 -a 076 ------- -o. aw -0. m -0.078 -a 076 -a m . . ..-.-. -a W -a O&l

(-Jay (mImter- C—.-— 1 –. at7 .W2 .037 .011 .011 .Cr20 .m ------ .a34 .am .CrW .m .m . . . . . . . . .071 ,065

TABLE VIII.-HINGEhfOMENT COEFFICIENTS. CLARK Y WING WITH 0.26c BY 0.40 ~ UPPER-SURFACE

AILERON AND 0.25c FULL-SPAN SPLIT FLAP.

(C. is givenfor one atkon at rfght wing tip)

RX.. W,@33 velocity-al mph Yaw=w

8* W 6“ w I –P –m –w -wI

–30=’ –W I –7P

I

al FLAP NEUTRAL

w ..--– . . . . . . ..-.-------- am a CU24 aam

W’ . . . . . . . . . . . . ------------ .0)10 .C4m .m

m’ ..--. -..-.– --------.--- .OxM . (W3 .0330

w ---------- .-__. ----- –. m .m . Wlow ---------- ---------- –. 0m4 ------------ –. m ow ..–.- . . . . . . ----------- –. m –. m –. m –. mm –: E

FLAP DO~ W

amn.C041.m40.m

–: z

amta

.M4s

.Wi’a

.W29

.Wfl–. ml

am am a ml1$ –. Im3

a 0014 a CO: am am

T

a w a ml Cl@&l–. m –:%

.Im13–: E

.0)19 .C#9 . ml

!$o

–. Cu!5 –. m –. Oms.0m7

–. ml.al14

–:%.WM :E .ml

?# –. ml –. ml –. mm –. ml.0314 .-..6------- . . . . ..Gi.

–. ml –:% –: M o

WIND-’IWNNDL RESEARCH COMPARING LATERAL CONTROL DFWK!ES 477

TABLE 1X.—CRITERIONS SHOWING RELATIVE MERITS OF AILERONS

Ailerm.. 0.15s by 0.60b/2 Aikrom O& by 0.40 bfi I

‘W& U!J#&hifnary Ordfmry

Snbjed Crfterlon ?@down469)

cm8eu-

up.#Y

Flap#w

up#y

maaleu’

up-OJYstandardw up,mdn.

MndardZb” up,26”b

uIM#Y

2(I371SS6

%s2

:E.014.an

I&

CL015*—.ml

.mzJ_. r@

.m4—.~

o0

I@

–%

aos7

:E.010.012

0

WIDE am or rnfnfnmm9p+d- %usped ranga.--_.-_... ____ %#$cD.f*

PAim--------------------------

Rate of time ------------------ L4D 8t cL=0.70

{

Rc a-w-------------------------------------Latmal cnntmRablUty_.. ---- gz ::~:------------------------------------

— — ------------------------------ .RC u= W--— ------------------------------

Lateral contil wftb aldeMp-. - M&mW~;Aat whlob8fIemnawfRbalancaC{ dtu

L n48&3M.4

. lm

.bw

–: Elm

Zamw. o~~

.070.0!3.W

o16”

L270~;

:0%.m.a?a

w

.---- .-— .–o. m

-----------. Ols

-----------.023

--------- .-.015

lx’

1711”

O.OLS

i%.m.m.W7

LZ5781.616.9

. 18s

.f@a

–:%la”

a 0190.m

&_. ~. ml

.—.~b.m

.------—-Ir

K

O.ws.W3.&&.016. Ols. ml

a Olao.W.?b—.~

oe—- ~

. . mo

w

armb—. ~

----------<—. @-j7

ob—. ~

.-- —-----.—. ~~

M“

:~

o.M.101.fm.ms.m

o

. . . . ...

bk%%eri%%cf*w). 1momonts Cij a-@--------------------------------------

f~) Eve ... . . . ..-. -... -----amble------------------ cmJ a=lw -------------------------------------

c.A~ U=w-------------------------------------

----------0.036

---------J—. 0]4

..------–. 019

l\c.Afa=w------------------------------------- ---------–. 023

18°

Yaw-@ ..------. -.--; -----------------------Yaw= W------------------------------

if&xfrnumunstable C’aat @b/2 V-O.M~aw.lY -------------------- ------------

btdd

Control

atabillty (#L-OO)_

am:Ml.m.m.MQ

am.an.027.fm.m

o

Yaw= T-----------------------------------

-..J(:$ ::~------------------------- . . .. —------------ ------------------ .force reqnfre”i.-.

P

CF a=w--------------------------------------CF a=w-— --------------------------------