Two-dimensional Wind-tunnel Investigation of Sealed 0.22-airfoil-chord Internally Balanced Ailerons of Different Contour on an NACA 65(112)- 213 Airfoil

NATIONAL AD VIS ORY CO MMITTEE

FOR AERO NA UTI CS

TECHNICAL l\OTF: No . 1099

TWO-DIMENSIONAL WIND-TUNNEL INVESTIGATION OF SEALED

0.22 - AIRFOIL-CHORD INTERNALLY BALANCED AILERONS OF DIFFERENT CONTOUR ON AN NACA 65(112)-213 AIRFOIL By Albert L . Braslow Langley Memorial Aeronautical Laboratory Langley Field, Va .

Washington July 1946 NATIONAL ADVISORY Cm-1MITTEE FOR .A.ERONAUTICS TECHNICAL NOTE NO . 1099 THO-DIMENSIONAL vffiID - T'"LJNNEJ.J llJVESTIGATION OF SEALED 0.22 - AIRFOIL - CHORD INTERNALLY BALANCED AILERONS OF DIFFERENT CONTOUR ON AN NACA 65 ( 112 ) -213 AIRFOIL By Albert L. Braslo1T A t"JO - dimensional i-lind-tunnel investigation i-/a8 made of t"TO interchan gea ble sealed. O. 22 - airfoil - chcrd internally bal~~ced ailerons on an NACA 65(112)-213 airfoil . One of the ailerons tested . was of b'ue airfoil cont our and the other vm.s modified by pa rt ly eliminatinG the cusp n ear the trailing edge .

T ests wel'e made to d.etermine the effects of the aileron co ntour modification on the section aeroQynamic characteristics of the airfoil and aileron .

The rest'.lts of the inv est igation indicated that the modifica tio n to the aileron cont o ur caused the aileron effective-: ness to increase slightly at low aileron deflections and to decrease slightly at large aileron deflect::'onsj caused the rate of chan ge of aileron section hinge -moment coefficient vlith both section angle of attack ~"1d aileron deflection to increase positivelyj caused little chan ge in the hinge -moment ~arameter for a given rate of roll at the 1m , aileron d.eflections but an increase in the hinge - moment paramete r for a given rate of roll at the high ailer on deflectionsj caused no app~eciable change in the section draB coefficient. rate of change of section lift coefficient ' ... ri th section angle of attack, and air - foil critica.l Mach n1J1l1b e rj and ca used an increase of approxi - mately 9 percent in the maximlml section lift coefficient of the airfoil vli th th e ailerons neutral . Th e application of standard rouglm ess to the leacling edge of the airfoil inc~ ~eased posi ti vely the rate of change of aileron section hin ge -moment coefficient 1-lith both section angle of attack and aileron deflection, decreased the aileron effectiveness th.roughout the aileron deflection range, and caused. a smaller chan ge in the hin ge-moment parameter for the true - contour aileron at any given rate of roll than for the modified aileron . Aileron d.eflections of -3 and 3 '\oTere f ound to have no significant effect bn the airfoil critical Mach number at the design section lift co e fficient .

- -- -_.-._ - NACA TN No. 1099 INTROiJUCTION Thickenill8 the cusped trailing ed .g e of lo,..--drag airfoils is sometimes desi",~able, mainly for structtrral reasons. Enough experi - mental data are not available at present tc. sho,·! holt to t~ic cen the cusp to best adV'antage, but a method . of thickening this part of the airfoil by straight-line fairin go 118.S been ShO'HIl during previons inve3ti ations to alter the ael'on.yn'llll.ic characteristics of some 101.;-11 'a3 airfoils, In an attenpt to keep changes in the aerod:rn~ _ ic characte~isticG at a minim'~, a comproI~tse modification w'as made to the Ct~Sp of an NA.CA 65( 112) -21 3 airloil by retaining the oriGinal <tirfoil mean line ill ile fa rinB out the upper surface to a straight lin e , The e:fect of the c ont.our IDDdification 0n the aileron effecti veness ffilcL hinge moments and on the a:l.:cfc·il clrac cha :;, 'acteristics and critical Nach l1llJ .1ber were determ in ed lro:m an investige.tion in the Langley tl.'O-dim ensio nal. lo"r-turbulence pl'eSStrre t"t.mnel of the NJl.CA 65 (11 2) - 213 airfoil equipped ~-rith t~TO interc.ha.l1geab le sealed O.22 - ai x· f o ii - chord inte ~ mally balanced ailerons; one of true airfoil c o ntour and one of the modified cont our . Te sts vrere mado ~~th th e airfoil surfaces aerodynamically S'.!!l.o ot h and vi th standarc, roughness applied t o the leadin g edge . In addition, the differential pressures acrO RS the aileron seals ,""ore o'!:Jtained for use in estimating the hinge-moment characteristics of the ailerons vcth a.11Y amotmt of s ea led internal balance.

COEFFICIENTS .AND SYMBOLS T11e coefficients and symbols used in th e presentation of resul ts are defined as follo v 's: c airfoil section lift coefficient (l/Cloc) .

l airfoil maximum. section lift coefficient Cd airfoil section drag coefficient (d/ Cloc ) seal-pressure-differonce coefficient; positive when pressure below seal is b reater thal1 pressure above seal aileron section hinge-moment coefficient based on aile~n chord. (h/Cloc ::» a aileron section hinge -m oment c o efficient based on airfoil chor(l (h /Cl cc ) NACJI. TN No . 1099 S airfoil pressure coefficient \

(rr°qO-' p)

.. l1ere 1 airfoil lift per lmit span d airfoil drag per un~t Spfu~ h aile~on hinge mooent per unit spen; positive when trailing edge of aileron tends t o deflect dm·rrnle.rd c cho1"::1 of airfoil 'Hi th ailel 'on neutral chord 0:' aileron behind. hinge ax i s

(~p V ~\

free - stream dJn~dc pressu~e

\2 0 0 I

Vo free - strea~ velocity Po free-stream density H free - stream. total pressure o p local static press ~ rre and a airf o il section angle of atta.clc, degrees o 0a aileron deflection ~dth respect to airfoil, degrees; posi ti ve when trailin g ed e is d eflec ed Clo; ·.'Inmrd chord of overhang from aileron hinge axi s tC ) middle of Cb gap seal R Reynolds number ai:cfoil critical Macl1 number Mcr

(dc \

c = a

\~(L~)Oa

NACA TN No. 1099 /

/'CPP\

P =! qO~\

ex. ' ....

\00.

'\ 0 / , 'J,

.d~~

(

aileron section effectiveness paraneter

.. d6; )C

7, incremen t of ai~foil sectio~ a~gle of attack increment of ailercn deflection aileron section effectivene3s parameter; ratio of increment of ai~:,foil section a."Ylgle of attack to increment of aileron rle flecti oC1. required. to maintain constant sec~ion lift c oe ff~cient total deli d.0 in steady roll a n aileron response paraneter increment of aileron section hinge-moment coefficient due to aileron deflecti on at const~~t section Rngle of attack increment of aileron section hinge-moment coefficient due to chan ge in section angle of attack at constant aileron deflection 6c increment of total ai1e::,' on section hinge-moment H T coefficient in steady roll 6cH T aileron section hinge -moment parameter 6ex.;;J66 a The subscripts to partial derivatives denote the variables held constant uhen the partial dsri vati ves are meas1.J.reo. . The ierivatives are meas~ecl at zero angle of attack ancl ze:i:'O aileron deflection.

NACA TN No . 1099 MODEL The model had a 24 - inch chor1. and ' ·r as cO::l.struc ted. of l aminated ma.1-J.ogany "lith the exception of the interchangeab l e ailerons, "hich 1 ,r e1.'e constructecl. of cast all;ll1inum . The two ailerons te sted , '-Thich had chorcl.s of O. ;22c and sealed intel"'Ilal balances of approximate l y 0.33ca , differed only in contour .

One was of tl'ue airfo~ . l contour (!TACA 65(112 ) -21 3 ) and the other ,.ras mociified by the -pa2."tiel elimination of the cusp near the t :rE'.E ing edge . The modification consisted of fairing out the u;rper - sUl'face c usp near the trailing edge ,·r i th a straight line fr02.!l a point O. 133c above the trailing edge tal1.gent to tile airfoil contoU1' and moctifying the lowe:' s1-lY.'face so as to retain the o::'igi nal airfoil mean line . Ordinates of the basic NAGA 65~J_J.2) - 213 airfoil section are Given in table I fu"1d the crdina-'ces fOl" th e ::ea::' 30 percent of the modifiect airfoil are 6i yen tn tal:-le II . S~cetches of t~le t\.:o ai_erons are given as fig-ure 1 . 'Rubber seale "Tere used along the c om plete span a.."1d at both en:ls of the ailerons to stop the flol-1 of air tr..rough the gaps .

For the snooth cond.ition of the model, the airfoil surfaces -..r ere sancted "lith No . 1.;.00 caTboJ:'l'l1 cl.um paper to p ror uce a1'l aero - dynamically smooth finish . For the stanclard airfoil le ad.ing - ed ge. rOIJ-8-'lm'3SS condition .• the mod.el s1-:rfaces were aerodynanically smoot:l e:-::ce !'i~ that 0 . 011 - inch carboruna,.U1l gntins vere applied to each ajrfo il surface at the l eading eeige over a sllrface length of 0.08c measured from the leadin g edge ( reference 1) .

APP .A.RATUS AIm TESTS Tests of the model vIi t~ each of the tvro ailerons were marte in the La1'lgley tvlo-c'cimensional lO",T - t1.U~ 1)~ence . pressure tQ~el . The tests incl~ ced measurements at a Re ynolds nurn ber of 8 X 10 of airfoil lift an'. d2:'ag, aileron hin ge moment, and ba~ance -p:~ess1Jre for the ae::."od;ynami c ally smooth mocl el v,rj.th various clef lec t i ons of each aileron , Airfoil lift, ailer')n ~inBe-moment) e":''1d balance-preosUl'e characte~istics Here also d e termin ed at a Reynolds m1ll1ber of 8 X 10° for the model y;tth stanr.ard r0ug,wess applied to the lead.ing edge and vii th vari ou s deflections of each c:.i leron . Iii th each ai l eron neutral, lift r-m d c_rag me9.s11.rements ·,·rere mad.e of the model Doth i n an aerodynam';'cally smooth conditi on and ,vi th standard leading- 6 6 6 edge roughness at ReynolcLs nvmbers of 2 X 10 , 6 X 10 , 8 X 10 , NACJ\. TN No. 109?

and 9 x 10 , corresponcin~ tc Mach nUDbe~s of 0.15, 0.14, 0 . 15, and 0 . 17, respectivel y . In ad/tition, airfoil surface pressures

were measurec'. f' ~o m the lead.:'.ng ed~e to o. 70c at 9. Reynol(ls

num.be . of 8 x 10 th:~ough an approximate range of section lift coefficient from - 0 . 5 to 1.0 .nth the ailercns neutral and at the design section lift coefficient of 0 . 20 v~th the ailerons deflected - 3 and JO.

Lift an(l <'l.rag meas'U'ements vere mad.e by tr_e methods briefly describe ·i in r':)fe=~ence 1. Airfoil su:cface preSSUl"eS and the presS'l}..re diffe:::ence across the aileron seals ,,,ere measured ",j.th static - pres sl1re o~ifices located along ooth airfoil slITfaces and in the chamber B . bove end below the aile~('on ba~ance plate .

Aileron hinge -moment L1eas~ements 1.;rere L1arte 1oT1 th a pressure - bellm-ls balance .

The follo11in[; fact ors were applied. to correct the tunnel data to free - air comUtions : t c = c 0 . 977 , 992c .- Cd 0 . d ::: 1 . 008<1 ' qo 1-1here the primed quantities represent the values measured in the tunnel (referenc e 1) .

RESliLTS AND DISCU SS ION The basic section lift, drag, hinge-mnment, ane. balance - pressure data are pres':mted in figures 2 to 6 fo:;.· the true - cont olIT aileron and in figxres 7 to 11 for the mvdifiei aileron .

These f.i.cm·es inclucl e data for the airfoil uiti:J. aeroc1.yn8. :mically smooth sl:.rfaces and Hi th standard roug...1mess app lied. to t.he lAftding edge . The discussion of the :lata re.fe.i."S to that obtained at a Reynolds number of 8 X 10 unless other1~8e stated.

Aileron Effectiveness The effects of the eileron contoUl' moo.ification on the

aileron section effectiveness parcmeter as and on c are

NACA TN No . 1099 shown in table III and curves of ~o against 0a at a constant c"l. of 0.20 are shov,n in fi"51ITe 12. For the airfoil in an aerod~lamically smooth condition, the effectiveness parameter ~ is sl iSl1t ly greate:.. f or the mo dified aileron than for the true- cont c1.u : a il ero n . The values of ~o for the modified and true- COil-tow: ailerons a::.ne 97 percent ancl 94 percent, respectively, of the thin- a irfoil t!leoretical effectiveness (refe re nce 2) and 17 p erce nt and. 12 percent, l' espectively, g reater than the value ( -0.4 80 ) obtained. on t he NACA 0009 airfoil section (r eferen ce 3).

Standarcl airf o il leading - edge roug11.ncss cau8ed a larger adverse effect on the effective n ess of the modified aileron than on the effectiveness of the true -c o ntour aileron.

In order to show the variation of the aileron effectiveness with lift coefficient and aileron deflection, values of the effec ti veness have been measured beb.reen definite aileron deflections at a constant section lift coefficien t. and are

desigrJ.ated (6.~0 / 6.0a ) . Values of ('/XL /6.0 \) are shown

a \. c"l.. . a c"l.

pl o tted against section lift coefficient in figure 13 for aileron - dellection limits of ±100 and ±20 . The effect iveness of the mo dified aileron is slightly greater than that of the true-contour aileron on the aeroclynaJtlically smooth airfoil when measured betHeen aileron deflections of _100 and 10 . An increase in the ailer-on -deflecticm limits to -20 and 20 causes a larger reduction in the effectiveness of the modified than of the true- contour aileron inth the result that the true - contour aileron is slightly more effective at the high aileron deflections . For the airfoil ,nth standard roughness app lied to the leading edge, the values of (6.~ o /6.0a)cl for the true-contour aileron were higher than for the modified aileron·when measured between aileron deflections of both ±lOo ~~d ±20 .

Aileron Ringe t,10ments The aileron hinge moments and balance pressures were measured whe n the a irf o il angle of attack ~o IoTas both increased and. decreased. The value s of ch and 6.p / qo were genel'ally found to be more positive for increasing than for decreasing angles of attack. The t o tal variation usually amo lli~ted to less than 0 . 006 and 0 . 06 for ch ru~d 6.p/QOJ respectively. It is felt reasonably certain that this difference in the values of ch and 6.p/Qo was ca-!;.sed by a lag in aile l'on setting as the angle of attack was changed due to the method us ed in attaching the ailerons to the presstrre-bell ows balance &~d also by friction in the control-surface and hinge-moment balance bearings . Average values NACA TN No. 1099 of the section hinge-moment coef£'icient and seal-pressure - difference coefficient ~xe used, there~ore, in the presentation of results .

eection characteristics.- The variations of aileron section hinge-moment coefficient ch and seal-:pl'essure-difference coefficient 6p/qo ~nth airfoil section engle df atta ck ~o are presented in figures 5 and 6 fOl' the truc-COI't01..1r aileron and in figu.res 10 and 11 for the modifieti aileron. The irre[;u- la r.itie8 that occur in the variation of ch ,~th ~o for the smooth airfoil correspond to the limit s of. the lov - drag range

as shOi.J11 in figures 2 ani e. Simile'.!' irrcgulari ties have been

noted iuring other t~,o-d.imensional invest.i.gations of control surfa ces (f or example, reference ~ . ) and a:ce believed to be caused . by the sudden movements in tra..'1si tion along the airfoil surfaces at t:'1e e xtremities of the Imv-drag range . Reference 4 ind.icates that no 1)nusU[~l aileron stick-force cnara.cteristics will be caus ed by the sudeten changes in the b·iO-o.:L""rJ.ensional hinge-moment coefficie::lts. The adrli tj.on clf standard. rouglmess to the airfoil leading ed.38 '3limin::l ted. the irregulari ties as sr.own in fig~'es 5(b)~ and 10(b) .

Values o~ cha' ch ' p~, and Po for both ailerons o on the smooth and rough airfoils are given in table III. The modificati.on to the aileron contoilr Ol~ ste.nde.rd. airfoil leading- edge roughness caused SIr..3.11 positive increases in bOt:'l ch ~ and c . The variation of e and 6T!/~ "lith 0a at a ho h consta.."'1t section lift coefficient of 0 . 20 is presented in fiBl.lre 12.

The basic section hinge-moment aTld balance-pressure data of

figures 5, 6, 10, and 11 may be used to estimate the section

hinge-moment charactel~istics of ailerons of similar contour and chord 'Hi th &--ry amount of sealed. int ernal balance by the method given in reference 5.

Basis for comparison.- The mean aTlgle of attack at which an aileron is ope~ating is altered by the rate of roll. The effect of the change in angle o~ attack on the aileron hinge-moment characteristics mu.st be taken into a. ccovnt for comp8.l~ison of ailerons ~rom section data. This correction is usually made by use of the constant-lift concept , jn ';'Thich the ass1ll11ption is macle that the aileron part of the ,"rinG acts at constant lift d.1..1ring steady roll . The rate of chanGe of the section hinge·· moment coefficient 1'1i th aileron deflection in steacty roll is then given by the equation

(1)

NAClI. TN No . 1099 British research, however, has indicated that the parameter Ch~ is overstressed in the constant - lift concept anct that a more accurate equation is / ( 2)

ch = ch (1 -

Orr -0\

~'There n is a response parameter clepend.ent upon the a i leron dimenslons, wing aspect ratio and taper, and. spanwise location of the aileron . A typical value of n, equal to 0.2 , is given in a British paper of limited distribution but more recent NACA data indicate that a more suitable v8~ue of n for the ailerons of' a modern fi [,h ter - t:lpe airplan e is 0 . 27, and that value has been u se d in the present analysis . Equation (2) is inadequat e for determining the three-dimensional aileron characteristics, but it may be used f or comparin g the tHO ailerons of different contour. In order to apply equation (2) t o non- linear curv es it has oen converted to 1ncrements of the total aileron section hin ge -moment coefficient in steady roll by h

(6c }J 1

n

1 - ~/-';;;a (ZICti) 6 J J

The method of ana~ ys is is the same as that used in reference 6, 6c H The hinge-moment parameter ---~, 'Thich is the ratio of the

% /60

a increment of' section hinge -moment coefficient in steady roll to the aileron effectiveness, is plotted a ga inst the equivalent change i n section angle of atta.ck !:Y:;,o required to maintain a constant section lift coefficient for various deflections of the aileron from neutral . This method of analysis takes j.nto account the aileron effectiveness and hinge moment ~~d the possible mechanical advantage between the c ontro ls and the a ilerons. The aile:con span and possible three - dim.'9nsional - flO'lT affects are not consid .e red except as indicated in equ.ation ( 3 ). Th o gmaller the value of the hing e -moment parameter f or a g iv on value of fuo, the more advantageous the combination should be for providing a 10'l'Ter control force f or a. gjven value of the win e:; -tip helix angle.

Aileron comparison.- Values of the hinge-mom ent parameter 6CHT ---- -- " ---' -'- - -.- - ---- are plotted. against /'ia j.n figure 1)+ for each aileron o

% / 60

a NACA TN ITo. 1099 on the airfoil in a smooth condition and 'Ti th stamlard r oug hness applied to the leading ed.ge. For the cmo o th airfoil , both ailero ns should provile about the srune control for~e at low aileron deflections. The tJ"D.e-contolll' aileron should pr ov ide the Im - Ter contr ol force at hiGh . iler on deflections for the airfoil in a smooth condition and. thrO"t.lgh the entil~e ran6e of deflectionA tested fOT the airfoil 'tTl th sta..'1c1ard leadhltJ-edce rousf1..ncsS.

Al though the application of st::mdal~d r01) . ghness General ly causes the value of the hingo-Doment pal'a:n ctel' to nc'~case slightly for any given value of 6a (f1 G. 14), the con~rol force for the o true-contonr n il eron ,.;ould chanGe les8 1.;i th changes in the s1Arface condition of the w~ng, as can be seen from a comparison of the val ues of 6cH'l' - for the smooth airfoil ,vi th those f or the %/6'Oa airfoil ~~th standard leading-edge rOL~~ess.

Lift The modj.ficat_on to the aileron contour or standard airfoil 1eadin s -e clge ronghnes8 had no effoct on the airfoil lift-curve slope ,.i th the aileron noutra1 as 8ho",n j.n taD1e III. The ya1ue of c is equal to O . lOl ~ for all conditi ons .

~a.

A comparison of fig1}.Tes 2 and. 7 0110'i-T8 that the aileron contour modificat:lon increases appreciably the ma...ximum section lirt coefficlent c, of the airfoil in 0. mnooth condition.

~max With the ailerons neutral tho conto1.Jr modification increases the val ue of c, from. 1. 3'i for the t~~ue- c o nto1.U~ aileron to 1 . 49 .

~max For the airfoil with standn.rd. leadinG-edge l'ouglmess, the aileron c ontour modification causes no siznificmlt change in c7, .

The reduction :i.n the v alue of c7, c aused. by standard.m~~ad.ing- max edge roughness is similar to the clecroase fOlmd fc,r other "f;ACA 65 - series airfoils of compar able thickness (ref eren ce 1) .

6 6 The effect of Re3mo1ds number bob·Teen 2 X 10 and 9 X 10 on the section lift characteristics of the ai?:'foil in the smooth and rough cor.o .. . tions is ShOHl in fiFsures 4 and 9 for the neutral position of the true-contour end. moriifiecl aHerons, respectively.

Similar effects of Reynolds number are noted for the tl.0 ailerons.

An increase in Reynolds number from 2 X 10 t o 6 X 10 causes a

large increase in maximum section lift coefficient for the smooth airfoil; how8-ver, a further increase in R0ynolCl.3 munber to 9 X 10 canses no appreciable chan3e. For the airfoi l vTi th stendard.

leading-edge r01.Jglmess, R 8ynolds n1JLlbo~ th:rough the range investigated has no significant effect on the value of c7, .

:max NACA TN No. 1099 Drag The aileron contour modification has no significant effect on the smooth airfoil section dr ag characteristics except at an aileron deflection of 20 as can be seen by comparing figure 3

for the true -cont our aileron with figure 8 for the modified

a::'leron. The values of the sect.ion draG coefficient.s for the 20 deflection of the modif ied aileron are doubtfu.l, hOi-rever, because of probable cross-floH along t.he Splli"l of the model. v11 th the e}:ception of the 20 dcflec tion, a 10.T-drag "b u cket" vras realized at all deflections of poth ailerons.

The effect of incree.8ing the Reynold.s nUl'!lher from 2 X 10

to 9 x 10 'liaS nonnal., t.hat is , the value of the minim1.1lJl section

drag coefficient ano. the range of section lift c oeff icient for

l ow' -dra G valu es decreased ,rr t.h increasing Be~molds number (figs. 4

and 9). The increase in the values of Cd caused. by struldard

airfoil leading-edge rouglmess (fi gs . L~ A!lc1. 9) is similar t.o that of other NACA 65- se' 'ies airfoils of compa:cable thickness (referen ce 1), Airfoil Pressure Distri1>ution and Cr:i.tical Mach Number The pressvre coeff icients ove r both a irf oi l surfaces from the leadin g edge t o 0.70c are prosented i!l figure 15 throu~ an approximate range of s e ction lift coefficient from -0 . 5 to 1.0 f or the airfo 1 r,rith a neutral position of both the true-contour and modified ai l erons . The varic.tion of airfoil critical Mach number Mer' estimated by v on Kc.\rman's method from the experimental surface pressures (r eference 7), with section lift coefficient is presented in figure 16. The modificat.ion to the aileTon contour had very little effect on the values of M ' Theoretical v alues cr bf MCl' for the NACA 65 ( 112)-213 airf oil seetion, calculated by the methorls of reference 1, are also pr esented in figure 16. Good agreement exists bebreen the values of lvi pr edicted from theory cr ruld from the experimental data in the raDge of secti on lift coeffi- cient for high critical Ma.ch nUl'!lbel' and lov! drag.

The chordv~se variation of airfoil ~reS8ure coeffic ient at approximately the design section lift coefficient of 0 . 20 is present ed in figure 17 for the airfoil ',ri th each aileron deflected _1 , 0 , and 3° . 'Because t.he val ue of Mcr is a direct function of the pe a k pressure on the airfoil surfa.ce, the close agreement in the peale valu es , of S for the ailcl'on d. ef lections tested 0 0

indicate a ne GleGi ble e ff ect of fu"l a il eTon deflection of -3 or 3

NACA TN rJo . 1099 on the airfoil critical tvIach number at a con3tant sectio n lift coeff:tcient of 0 . 20.

CONCLUSIONS A tv:o- dimensionel wino.-tlmnel inycstigation ~·D.s made of an NACA 65( 112) - 213 airfoil f.}quipperl ,vi th h lO int.erchangeab l e sealed O. 22-a:!.rfoil-chord internally balanced ailerons of different contoUl'. One of' the aile~Gns tested i\'8.S of true airfoi l contour and the other 1·ms lT10dificd by the pai:'tiA.l elimination of the cusp near the trailing edge. The data obtained indicate d the folloTNing c o nclusions : 1. Modif ~ cation of aileron contour caused ( a) The ailercn effectiveness to inc re ase slight l y at 101-1 aIleron deflections and to decrease slightly at high a : aeron d.eflecti o ns (b) The r1.te of change of aileron section hinse - moment coefficiont 'Hi th both section a ngle of attac:r~ and aile~on deflecti on to increase positively ( c) Little cha.."1 (; e in the hine;e-moment parameter for a giYen rate of rell at th e 10vl aileron doflections but an increase in th e hinA;e - moment par3Il1eter fOI' a gi y en rate of roll at the high aile:oon def'lect:Lcns ( d ) No appreciable change in the section . irag coeffi - cient, rate of change of sectj on lift ccefficient T /Ji th section a..'1.g1e of attack, and a irfoH critical Mach number ( e) m increase of approximately 9 percent in the maximum section lift coeff i cient of the airfoil with tr.e ailerons neutral 2 . The application of stano.are rouBl".ness to the leading edge of t he airfoi l (a) Increased p os :itiYely the rate of change of aileron section hinge-m one nt coefficient with both section angle of attack and aileron <lef1ectlon ( b ) Decreased the aileron effectiveness throughout the aileron doflectir'n range ------ ~ ~~~ -~- . .

NACA TN No. 1099 (c) Caused a smaller change in the hinge-moment parameter for the true-contour aileron at any given rate of roll than for the modified aileron

3 Aileron deflections of -3 and 3 had no significant

effect on the airfoil critical Hach number at the design section lift coefficient of 0 . 20.

Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics

Langley Field, Va , March 1, 1946

NACA TN No . 1099 BEFEREIJCES l. Abbott, Ira F., von Doenh.off, JUbert E., and Stivers , Louis S., Jr. : Summary of Airfoil Data. UACJ". ACR No. L,)C05, 194'5 .

2 . Gl s.uert , H. : Theoretical Re1at1ons}lips for an Ae rof'oil ,vith m .nGed Flap . H. f'r. lvi. No . 1095, Britiflh A.R.C . : 192 7.

3. Ar:les .• Milton E . , j·r., an<l SeaTs, Richard 1. : Determ i nation of Cont :; :o1-Surf&ce ChaJ:'actel'istics f:;:-om NACA Plain - Flap and Te.b Data. NACA P,ep. NC) . 721, 1941.

l~ . Braslm'T, Ar)ert L . : Tl-TO-Dimensional yllnd - Tu,'mel Investigation of LO'~'-Dr8-S V <)rtical -Tail, FIo:i~ .zo ntlll-Tail, and vTing Soctions Equippecl id th Sealed Inte:mc.11y Balanced. Control SlU'face3. NACA TN No . 10)+0, 1946.

5. Fioche1, J ar.k : Hill[~o M0l1lents of Se'J~erl - Inte i: nal - Bal&lce Arrangemonts for Con-c:;: ' ol Surfaces. II - Expe:::oimental I nve8t i ~ati()n of Fab:":i.c Seuls in the Pres(JDce of a Thin- Plat e O'rerhang . NJ.~J\!\J:{R No . L 5F30a, 1 945 .

6. Und .er ~Jood, Hilliam J., Branl.o-,., JUbert L. ; and Cahill, Jones F .: T'·lC-Dimens2.o nal Wind-'Il)nIle1 InYl'sti,gati cn of O. 20-Airfo:U - Cherrl Plain Ailerons or Diffel'em: Contot'!r on an NASA 651 - 210 Airfr,il Section. T-.TACA AGR N:J. L5Jt'~'r', 1945.

7. von I~£rm£'l., Th .: Comprcssi bili t;,r Eff3 ts in A0rodynami c s.

,Jou:.:' . Aero . Scj, . , v o l. 8, no. 9, July 1941, pp . 3:.7 - 3 56 .

NACA TN No. 1099

TABLE o r

ORDINATES FOR NACA 65(112)-21, AIRFOIL SECTrON

~tations and ordinates given in percent of airfoil chord] Lower Surface Upper Surface Ordinate Station Station Ordinate

0 0

-.942

1.0~2 .~83

-1.1

1.2 2 • 42

:~~

1.617 1.354- -1.~8;

1.~6

2., , 2.212 2.617 -1. 42

4.867 -2.50~

5.1;~ ;.1'~

7.6; -;.00

7.§6;

10.1;8 429

E:~22

~. 6

-4.

.86 - .092

1 4 15·1"

2. ;8

.18, 20.117 -4.592

1~.883

2 .896 25.104 -4.958

- 6.7~0

7.1 ; ;0.08; -5. 2 17

24·91~

7.433 5

3 .93 -5. R71

E •

0.042

8 -5. 25

7.56~

ez..95

45.021

-5·;50

.979 7· 53

50.

50.000 7..3;8 33

-,.1

0 6.

54.983 - .767

55. 17 58

-4.283

6. 25

5 .96 7

4 6o.0E3

65.0 6 -3.717

4 5.775

6 .9a~

-3.08;

70.0~

6~.9

E·02~

.20 -2.417

7 .942

~ 0 5.05

-1.742

0.054

~~.946

~:E'~ -1.092

85.050 .950

8 . 62 -.508

90.038 1.5G2

.700 9 .98; -.071

95·017

100.000 0

100.000 0

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA TN No. 1099 TABLE II ORDINATES FOR REAR 30 PERCENT OF MODIFIED NACA 65(112)-213 AIRFOIL SECTION [Stations and ordinates given in percent of airfoil chord] Ordinate Station Upper surface Lower surface 70.000 5.029 -3.058 4.225 -2.4 ~5. 000 0.000 3.400 -1.792 85. 000 23 2·592

-1. a

90.000 1.771 -·70 95·000 -·300 100.000

:i~~ -.133

TABLE III 0 6 SECrrON PARAMETERS MEASURED AT a = 00 AND 0a =.0 FOR R = 8 x 10 o a Surface c~ c~ ch ch P Po O () a 0 a a (1) True-contour aileron Snooth 0.104- 0.059 -0.540 -0.0038 -0.0081 0.036 0.095 Rough .029 .086 .104 .053 -.505 -.0035 -.0067 Modified aileron Smooth 0.082 0.104 0.060 -0.560 -0.0031 -0.0077 0.042 Rough .104 .051 0 -.0065 .029 -.49 -.0027 .073 "Smooth" and "Rough" refer to the airfoil with aerodynamically amooth surfaces and with standard leading-edge roughness.

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NACA TN No. 1099

Fig. 2b ; 7, f- J ~ Standard leading-edge roughness ............ -=.u~:.::r..;.J::..-.:I ...........

Figure 2 .- Concluded.

l 'XJ c.N III 1-'- Z ::t:- O ::t:- >-3 Z o i--' o \0 \0 OQ Z i 1

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I::>

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Sec t ion ang l e of a ttack , ao , de g (a) Smooth condition.

Fi gu re 5 . - Hinge-moment characteristics of a sealed 0.2 2c internally balanced aileron of true airfoil contour on an NAC A 65(112 )-21 3 airfoil section. R = 8 x 10 ; test, 'IDT 708 .

Fig. 5b NACA TN No. 1099 ...

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(b) Figure 5.- Concluded.

Fig.6a , b NACA TN No . 1099 . 1 T .~, ..

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Figure b .- Var i aolon of ~ with <10 f or a se a led 0 .2 2c in t erna lly qo balance d aileron of t rue ai r f oi l co nto ur on an NACA 65 (11 2 )-213 6; airfoil sect i on. It = 8 x 10 IDT 10S .

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, --'i-- - I--~-- 1-.1- NATIONAL ADVISORY .- -l- 1--- I- COMMITTEE FDA AERONAUTICS

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t-- r- "

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Section angle of attack, a , deg 1 a (a) Smooth condition.

Figure 10 .- Hinge-moment charac teristics of a sealed 0.22c internally bal~ced aileron of modified contour on an NACA 65(112) - 213 airfoil section. R = 8 x lC ; tests, TDT 709 and 711.

Fig. lOb NACA TN No. 1099 r-r- ,"-r- ---, e- e- rr- , .

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f- NATION AL ADVI SORY " COM M ITT EE fOIl AEA ONA ~ TIC I S I -< - -- -

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,., ~l2 8 .1 .i 2

J1tBi a

- c- - f- ~ Section ~gle of attack, (10 , des (b) Standard leading-edge roughnees.

Pigure10.- Concluded.

NACA TN No. 1099 Fig. lla, b

-I- H- --r-l o I ~- I r-I ..-+-- 1 b--ill:: IlJL

l

.-rt - H

._ L +--+-+-++-+-1---'- I I I I I I .1- fl --r- Section angle of attack , ao , \a) Smooth cond1 tion .

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~ AT ION" L ADVI SORY -'-- CO MM I TTEE f OR AE RONAU TIC S (b) S tandard leadin g -ed ge roughness.

Figure 11. - Variation of 6p with a for a sealed 0.22c internally balanced o qo aileron of modified contour on a.n NACA 65(1.2)-21, a.irfoil section .

6 ; R = 8 )( 10 tests, TIlT 709 and 711 .

Fig. 12a NACA TN No. 1099 Airt'oll surt'ace ~ - - &nooth ~ ----- Standard L.B. roughne8~ U) ~ G 'd "0 ~ ~ ., ~ >'.- ~ ~ :l ~ ..

as r--...

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\ " -< \ ..; -.16 NATIONAL ADVISORY COMMITTU FOIl AlROllAUTICS -.20 -20 -16 -12 o 8 12 16 20 -4 4 Aileron det'lection, 0a' deg (a) True-contour aileron.

Figure 12.- Variation of aerodynamic characteristics with aileron deflection at a constant seation lift coefficient of 0.20 t'or an NACA 65(112)-21~ airt'oil section equipped with two interchangeable 8saled 0.22c internally balanced ailerons of different contour. cb = O.~~ca; R = 8 x 10 • NACA TN No. 1099 Fig. 12b Alrt'oll eurt'ace ~ -- Smooth ""'r-~ - - -- Standard L.E. roughness t() ~ CD 'd ~ ~ ~ d ~ - ", - :-.~ 0 i'-...

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NATIONAL ADVISORY COMMITTEE fOIl AERONAUTICS -.20 - 20 -1 6 -12 0 8 12 16 20 -8 4 4 Aileron deflection, ~a' deg (b) Modit'ied aileron.

Figure 12. - Concluded.

Fig. 13a, b NACA TN No. 1099 .6 Smooth - 0-_

i-11 - I- -t -

r--

--

- --

---

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------

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Section 11ft coefficient, c L (a) _10° < Oa < 10°.

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m

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-

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4 1-

Starrlard

L.E. roughness - W-1

Aileron -- True contour ----Modified NATIONAL ADVISORY ~MMITTEE FOI AERONAUTICS ~ v

-.I -.2 o .2 .6 .8 1.0

t .4

section lift coefficient, c L (b) _20°< Oa <20 , Figure 1. 3 • - Vari ation of (Ano) with cL for an NACA 65(112)-213 airfoil

Wa cL

equipped with a sealed section 0.220 internally balanced aileron.

6•

10 R = 8 x

NACA TN No. 1099 Fig. 14a,b . 03 Aileron True cont ou r ~r- " 1oI0difie d " "/ .01

~

b

VI'"

~ o V

V

V

l/

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-. 02 NATlO""L AOVISORY COMMI TTEE rot .t.EAONAUTICS -. 03 -12 -8 o -4 4 fino • deg (b) Airfoil with standar d lea <.1ng- edge r oughn ess, / HT Figure 1 4. - Variation of the hinge-molDent parameter flc with equivalent change in section a angle of ~ttack required to maintain a con s t ant sec tion lift coefficient of 0. 20 for deflection

fld o/flO ~

of sealed 0. 22c internally balanced ailerons of different contours on an NACA 65(112)-213 airfoil 6 ,

section, R = 8 x 10

~.

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.277 .134 J~i

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1--- y - - -- - AERrAUTICS

-........, ==:~~:~§

ADVISORY ~ Fj -"'k 0. 22c -..,~ NATIONAL ~~~ t:::::8::~~~ surface CTMITiEE

"'-15-

chord ~ f':'-- sealed -- Lower a J J---,b:... 40 airfoil with \--" Percent .:...

I-i"'- equipped 1><-- 1 f--.

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e-contour NACA Tru 704.

an c, 503 - .~l . 7 9 , -.277 -· (8) TOT for

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AERONAUTICS ~~ FOIl ~~~ surface chord ~:__<~ ir' r--..; NATIONAL ADVISORY COMMITTEE Lower '---.,!r---- t---, , airfoil "- ,...---< -1--+-~-+--~~-+~~+--L~ -;:=::j~~ Percent "-- "----. f--.. it-""" ........, --.

~.---' ~/:~,r-':~:~.

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8 4

4. 4. 3. 3. 2. 1.

4. 2. 2 . 1.

aileron.

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.; <: ., 0 0 0 .-< ..-I ... ..-I ..: D) ..-I ... ... ... Concluded.

,

. Moditied

A l C (b) Figure15.- 0'1$2 -.309 -.537

p-' ~:r'S6 ~~:M1

-- 70 ~

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surface chord

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Fig. 16 NACA TN No. 1099

.8

r--

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ori / r-.....(i: .jJ 8 Modified ori H

C) ""

·3

r-i ori fH H ori <x: .2 .1 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

o

-.6 .2 .6 .8 1.0

-.2 o

-.4

Section lift coefficient, c L Figure 16. - Variation of predicted critical Mach nunber with low-speed section lift coefficient for an NACA 65(112)-213 airfoil section equipped with a sealed O.22c internally balanced aileron. Da: 0°.

.

, .

, .

(

NACA TN No. 1099 Fig . 17a

l..6 surf'ace Up per ---.

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r

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tit

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M oM ~

.4

.2 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS o 10 20 60 o 30 40 50 70 Percent airf'oil chord (a) True-contour aileron.

Figure 1 7. - Experimental pressure distributions f'or an N AC A 6 5(112)-213 airfoil section equipped with a sea l ed 0 . 22c internally balanced aileron. Tests, 'ID T 704 and 714.

"

Fig. 17b NACA TN No. 1099 . < 1.6 .---T---.---.---.----r---r---------------,r---,---,---~__.

Upper surt'ace cl.

°a

(deg) 0,20) .210 -) ) .204 ~ 6~--+---+_--;---~---r---r--_+-- _+--_1--~r_ --r---r---~~ o ......

H oM -0: 2.---~--+_-- +_--;---~ --~---r--_+--_+--_+--_1--~r---r-~ NATIONAL ADVISORY ( .

COMMITTEE FOI AERONAUTICS

n< . 1

o 10 20 60 50 70 Percent airfoil chord (b) Modified aileron.

Figure 17 • - Concluded.

• ..