Wind Tunnel Tests of Ailerons at Various Speeds I : Ailerons of 0.20 Airfoil Chord and True Contour with 0.35 Aileron-chord Extreme Blunt Nose Balance on the NACA 66,2-216 Airfoil

Document

ACF No • . 1Fl l NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ORIGINALLY ISSUED June 1943 as Advance Confidential Report 3Fll WIND-TUNNEL TESTS OF AILERONS AT VARIOUS SPEErs I - AILERONS OF 0.2.) AIRFOIL CHORD AND TRUE CONl'OUB WI TH 0 · 35 AILERON-CHORD ~~ BLUNT NOSE BALANC E ON THE NACA 66,2-216 A:rnFOIL By w. Letko, H. G. D:mac i , and C. Freed Langley Memorial Aeronautical Laboratory Langley Field, Va .

CASE FIL ~ ~

COP

~ACA

WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to 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 change in order to expedite general distribution.

L - 431 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE CON ' FIDENTIA.;rJ REPORT WIND-TUNNEL TESTS OF AILERONS AT VARIO US SPEEDS rl ~ I - AILERONS OF 0.20 AIRFOIL CHORD AND TRUE CONTOUR WITH ~ 0.35 AILERON-CHORD EXTREME BLUNT NOSE BALANCE ON !HE NACA 66 ,2-216 AIRFOIL ' By W. ~etko, H. G.,Denaci, and C. Freed SUMMARY Hinge-moment, lift, and pres8ure-dintribution measurements were made in the two-dimensional test section of the 'NAGA stability tunnel on a blunt-nose balance-type aileron on an N ACA 66,2-216 airfoil at speeds ~p to 360 miles pe r hour corresponding to a ~ach nu~ber of 0.475.

The tests were ~ade ' primarily to determine the effect of speed on the action of this type of aileron. The balance- nose radii of the aileron were - varied fro~ b to 0.02 of the airfo~l chord and the gap width waD varied from 0.0005 to O .O~ 07 of the airfoil chord. Teats were a~so made with the gap sealed .

The variations in hinge moments and lift with Mach nu m ber. angle of attack. and 8il~rqn deflection are given in the fo rm of curves of s e ctio n hin g e-moment coefficients and section lift coefficients plotted against aileron deflection , for the various cond{tions teste~, t~gether with cross plot~ s howing the gen ~ral ~ffect of Mach number, gap width. and balance-nose radii, ' The results show that there was a considerable in- crease in the stalled ra nge of the ai~eron with increased speed. Up to the stall, the variation in hi ge - moment coefficients and lift coeffioients fo~ the sp eed range t~sted wa s small but ihe variation may b~ app rec iable when stick forces at higb speeds are considered.

INTRODUCTION Large increas~s in the size and speeds of current co mbat airplanes, in addition to high maneuverability re- quired in co mb at, have made it necessary to balance almost exactly the hinge moments of ailerons and at the same time to ma intain their effectiveness. Although m ost types of aileron balances in use today operate satisfactorily at low s p eeds, difficulty, such as overbalanc~ at high speeds; has been experienced with some existin g aile ron installa- tions. This difficulty is appa r ently caused by the large amount of bala n ce co upled ~ith the chan ges in hi ng e moment that result from Compr e ssibility effects. Consi deration of these problemG has ma de necessa r y f u rther research on so me of the currently used or r e cently p roposed balance arran gem ents.

The NA CA is ther e for e undertaking a study o f so me of the mo r e p ro ~ isi ng aileron types at hjghe r Ma ch numbers than we r e e mpl oy ed in p r eviou s develop m ents. This rep o rt d ea ls ' w ith the section ~h a rac teristics of a blunt-nose bala n ce type of a ilero n of 0 .2 0 airfoil c ~o rd with a 0 . 35 aileron balance and of t r ue contour used on an NACA 66.2~216 air f oil. T._e amount of balance, 0.35 aileron . chord. was chosen because . from the data given in reference 1, it was esti mated th a t this a m ount of balance would g ive almost c omp lete balance on an airfoil of the :ACA 230 series at a lo w an g le of a ttack.

The section lift coefficient and the section hinge - mom ent coe ff icient cha were m easured at various airspeeds u p to 300 m iles pe r hour . corresponding to a M ~c h numbe r of 0.475 . Th ese m easure m ents were ta k en thr ou gh an an g le-of-a t tack ran g e from ~ 5° to 10 and an aileron deflection range of ±20o. Th e influence of the gap width between the aileron and w ing and the influence of the radii of the projecting corners of the balance w ere investigated. The data are pres~nted in the forD of c u rves of Ct and Cha . p lotted a g ainst aileron defle~tion w ith cross p lots to show the effect of the aileron parame t ers.

SY M BOL S ]'-:" j ai r foil section lift coefficie r :t ( qc) increment of airfoil s e ction lift coe ff icient cm /4 airfoil section p~tching - m0men t coefficient about c

quarter-chord point of airfoil (~~~~)

.-i

a i1 eron sect ion h inge-momen t c oeff ic ient ' (_~L8)

ch ~ a I "':1 qCa .1 airfoil , secti~n lift airfoil section p~t chin g mom ent about quarter-chord

m C /4

point of airfoil aileron section hin g e mom ent ha c chord of basic airfoil, .in cludin g aileron c chord of aileron meaGured from hi ng e axis back to a trailin g ed ge

q dynamic p ressure (~pV2)

air velocity V " mass den sit Y,of air P angle of attack for airfoil of infi n ite , aspect r atio a o aileron . deflec , tio.n , vfi~'h re ·s' pect , to airfoil oa Mach number, ', ' .

APPARATUS AND MOD E LS The tests we r e made, in "the tw o~di ~ Bnsional test sec- tion of the stabilit y tunnel at airspeeds up to 360 mile s per hour . The test s~ction is rectan gu lar, 2.5 feet ' w ide and 6 feet high .

The mode l , of an NA CA 66 ,2-21 6 , a = 1.0 section wag mEl . de of laminat ed mahog an ' y. It COillple t ely , span n ed the test sec ' tion' and was fixe 'd into ' C'i,:r.c u lar ' elld ' dis ks that were flush w ith the tunnel wal ls. Tne en g le of att ack of the mode l ~as ch? ng ~d by rotatin g the end disks. Tables I and ' II give'th e, ordinates of the a ir fo il section and ' locations of cen ter s of balance~tio s e ' radii, respectively.

Fi g ure 1 is a , photograph of a moa, el mounted in the tunnel.

The aileron of 0.20c and 0.35c balance and o~ true a contour w as made of steel with w ooden nose pie ces having 0, O.Olc, and 0.02c balance-nose radii. (see fig. 2.)

The aileron was supported at the ends by ball bearings mounted in steel end plates attached to the airfoil.

The aileron deflection was varied and the aileron angle and hinge moments were measured by a calibrated spring torque balance and sector system. Pressure ori- fices were located alon g the midspan ' of the wing and ai- leron and the pressure distribution was recorded photo- graphically. In some cases hinge moments and lift were obtained from the pressure-distribution diagrams.

For some of ' the tes , ts the lift was also measured by an i~tegrating manometer connected to orifices in the floor and ceiling of the tunnel. The integrating manom- eter was calibrated against lift obtained by pressure distribution.

TESTS Tests wer~ made with balance-nose radii of 0, O.Ole, and O. 02 c . Wit h z e r 0 r a d i ion 1 y pre s sur e - dis t r i bu t ion tests we re made. With radii of O.Olc and 0.02~, hinge moments were measured with gap ~idths of 0.0005c, 0.0030c, 0.0055c, ' and 0.0107c and also with a 0.00550 gap sealed with a flexible sheet that extended from wall to wall.

In the test in which the 0.02c radii was used - in addi- tion to the pressure-distribution and hinge-moment : measu~ements - section lift was measured by the integrat- ing manometer.

Tests for eac h condition w ere made at !1Ve speeds which g ave Mach numbers in a ran~e between 0,195 and 0.475.

The lo w est speed corresponds to a Reynolds number of about 2,800,0 00 and the highest speed to a Reyno~ds number of about 6,700,000. Figure ~ is a plot of Reynolds number based on standard at m ospheric conditions against test Mach number. Tests were maq.e at angles of attack of -50, 00, 5°, and 10 with ~2°, ' 2°, and 7.5° added for the 0.0055c gap (open and sealed). For each angle of attack readings were ta ke n at the following ailer~n angles; , 0, ±2°, ±5°, 0 0 0 , 0 0 0 ±7 , ±10 , ±13 ±16 , ±18 , and ±20 • The . high speeds : could ~ot be attained at the large an gl es of attack with large aileron deflections because of limited tunnel power.

Pressure-distribution records were taken at Mach numbe r s of 0.195. 0.358, and 0.475 for every angle of attack tested. For each angle of attack records were made at aileron angles of 0, ±5° , _7° , ±100, and ±16°.

PRECISION Angles of attack were set to within ±O.lo and aileron angles to within ±0 . 3°. The hinge-moment coefficients . that were measured could be repeated to within ±0.003 cind the ' lif~ coeffi~ients to within ±O~Ol .

Corrections for tunnel-wall effects were not applied to the section hinge-moment coefficients. The follo w ing corrections w ere ap p lied to the section lift and section pitchin g -mo m ent coefficients and to the an g le of attack: ..

(1 [1 - ! + 2~) ] Ct c l' YC1' , (1 - 2 ~Y) c

+ =

c m I

IDc/4 c 4 a o ( 1 + Y) u'

=

where

y = ~~ (~)8

h height of tunnel

~ = 0.304 (theor e tical factor f o r NACA 66 ~2-216, a = 1

airfoil) t Ct meas u red lift co e fficien t c / ' measured pitchin g -mo m ent coefficient mc at uncorrected or g eo m etric a ng le of attack The values u sed are:

c ~ = 0.963 c 1 '

c m / = 0.986 c / + 0.006 c1 mC C

a o = 1.023 a'

Hinge moments were me'asured simult aneously by pres - sure distribution and by the spr ing torque balance for a number of varied con d ition s an~ the results are shown in figure 4. The variation in the val u es is probably due to the fact t ha t the s p ri ng balance measures the hinge moment on the entire aileron , ~ hich includes effects of boundary la yer at the tunnel wall and of gap s at the ends of the ailero n as ue ll as the effects of an y cross flow over the aileron; whereas the p ressure distribut ion give s the hinge moment at one , section of the aileron and is subject to some errors in fairing the p ressur e -di st rib ut ion diagrams .

RESULTS AND DISCUSSION In ord e r t ha t the r e sults for the tests of various model confi gu rations may be mo r e easily found table III g ives t he fi gu re nu~bers J the variations sho~n on the fig - ure, and th e corres p ondin g mo del confi g uration . Only pa rt of the data are presented for ' the 0 and O.Olc balance - nose r ad i 1.

The results show that for all co n ditions the aileron apparently stalled at an angle of defle ction that depended on the speed, the angle of attack , the gap wid th, and the balance-nose radii a nd that the hinge moments increased ra pi dly in the stalled ran be . At the transition point be - tween the stalled and unstalled ran ge the aileron was observed to oscillate between the st al led and u nstalled condition . As the speed increased the unstalled range of deflections of the aileron g ener al ly became smaller . The effect was m ost pronounced with the zero balance - nose rad i i.

Hinge Mo ment of A il e ron The ailero n section h in ge -mo m ent coefficients Cha plotted against aileron deflection Oa are given in fig - ures 5 to 7. The va lues of cha given in figure 5 are from pressure-distribution records (no other satisfactory measurements we re avail ab le); tho se given in figures 6 and 7 are from sp ring-balance measu r ements (a co mpari son of resul t s obtained by the two methods is given in fig- u re 4) . These results show that fo r a limited ran ge of aileron deflectio~s and an g les of attack the ail eron bal - ance w as fairly effective. An average value of the slope of the curve of section hinge - moment coefficient plotte d aCha a gainst aileron angle, ____ , of - 0.0057 was obtained

a 6

from values of Cha at a oa of ±5° as compared to a

0.20 chord value of -0.011 give n by unpublished dat e for a plain sealed aileron on the same wing section. Although aCha the value of of -0. 0057 is r elatively large fo r CO a the value of combat ai r p l anes . the i ncrement r eduction in cCh a iS according to 4ata reported in reference 1, about t ~6~ the same as would be obtain~d fo r a O ,35ca balance aileron on an NACA 230 -series section.

for the ran g e of Ma c h numbers M tested the most noticeable ef f ect of increasing speed on the hinge-moment c ha racteris t1cs of the ailerons uas a co nsiderable increase in the stal led ran g e of the ailercn. The ge neral t rend of the effect of M on c' in the unstalled def lection fia with ran ge is sho wn by figure B to be an increase in increase of M. So me of this trend may be du e to the change in Ee y nolds number. Ap p roxim ate valu es of Reynolds number for any value of M may be obta i ~ed from figure 3.

A c hange in the u~stalled range of the aileron is shown by figures 5 to 7 to be the princ i pal effect on the hinge-moment c harac teristics resulting from changes in ba~ance-nose radii and gap width . An increase in radii from 0 to 0.02c changed the Install ed r ange f ro m about ±4° to about ±lOo and increased the hinge - m o ~ent -coefficient slo p e. For all cases , the aileron Fi th gap sealed ha d the greatest unstalled range . An increase in t~e g3p decrea sed appreciably the un stalle1 range; the amount of change varied w ith angle of attack and the effect was u sually gre ater for th~ p ositive r ange of aileron deflections than for the negative ran ge .

Figure 9 shows that the effect of gap on in the unstalled range is usually small. The general variation of cha w ith a is sho~n in figure 10. He re again the o effects of M, balance -no se radii, and gap are small .

Lift The airfoil section lift curves, Ct (obtained with the inte g rating manometer) for a 'ba lance-nose radii of 0 . 02c with aileron neutral , are presented in f igure 11 and show that the slope of the lift curve increases with Mach number. The variation of lift-curve slope with speed for t he various gap widths is g i ~en in figu re 12 together with a cu rve showing t he theoretical variation. It is believed that closer agreement " ould have been obtained if the com - pressibility effect on tupnel-vall interference and Reynolds number effects on the airfoil characteristics had been taken into account . Fi gu re 12 also shows that the highest slopes were ob tained uith the gap s~aled. When the gap was unsealed , an i nc rease in the g ap width caused a de~rease in the slope excGpt at the highest speed tested w here an i n crease in gap resulted in an increase in the, slop , e.

Fi gu re 16 iS,a pl ot of section lift coe f ficient against aileron angle. In order to avoid confusion , faired curves have been dra wn in this figure only through the test ~oints for a Each number of approximately 0.36 . For low and medium angles of attack an , increase in the speed increases C

the value of the slope of these curves (a 1\ a~d the

dO;')6, amount ' bf increase varies ~ ith the angle of attack; At ' high a~ gl es of attack an increase in speed generally caused , ( dCt '\ a decrease ill the value of \ de - 1 • , a~

(~:.~,'; , f or a

F~gure 13 also shows that the value of (d 0 a) a.

ran g e of Sa of ±5, was highest at low and medi um an g les of ~ttack ~ith the g ap sealed and, at h~gh 'angles of attack, the slo pe was highest for the 0.0055c gap; Increases in gap usually decreased the aileron deflection a t which the stall occurred and decreased considerably the effectiveness of the ail~ron at large aileron deflect~ons . The loss tn ~ffectiveness due to the gap a t large aile ro n def l e c ti ons w as le as t at t h e high s p eeds.

The airfoil section lift and secti on pit c hing - mo me nt co ef fici ent s ob t a i ned by pressu re distribution f or ba lance- n ose r adi i of 0 a nd 0.02c are p r e s ente d in figure 14. Fi g - u r e ~ 4 ( a ) s how s. a s mi g h t b ee xp e c ted. t hat the rei s 1 itt 1 e chan g e in the lift c u r V' e (ai le ron n~utral) with chan ges in balance-nose r ad ii . Fi gu re 14(b) sh ow s the va riation of the a irfoil section lift coefficient Ct ~i th a ileron de- flecti oll for the dif fere n t radii. It is ev i dent th at the aileron with 0 . 02c r adii has a mu c h l a r ger effect i ve range than either of tThe other t~o ailer ons and that the aileron with zero radii is i nef ficient because it lo ses all its effecti venes s a t a positive aile ron deflect io n of 50 .

Variations of Airfoil section lift c oef f ici ent Ct with Mach number fo r balance -no se radii o f 0.02c are shown in figu r e 15 for three aileron deflection s . Th e gene ral tende n c y . as expec.ted, is f or the lift coe ffi c ient to in - cr ease w i th Mach number ; pa rt o f t his incre ase prob ably is due t o Re yno lds number. At an an gle of attack of 10 • however . the lift coef fici ent decr ea sed af ter a certain va l ue of Mach n u~be r was r eached as a re su lt of critical speed occurrin g o ve r the le ading edge of the ai rfoil.

I n c reasing the gap width f ro m 0 .0 005c to 0.0107c gen erall y c aused a sli ght d~crease in the value o f cL ' (see fi g . 1 6 .) The aileron with 0 .02c bala n ce-nos e radii u sed in this test is s omewhat mo re effective in pr oducing lift at an an g le o f atta c k of 1 0 than a plain se a led fl~p of 0.20c on the s ame type airfoi l a t a pp ro xima tel y the sa me Reyno ld s nu mbe r, as is illdicated by the data g i ven in r efe r ence 2 .

C ont rol-Wor ce Criteri on is a control- force cri t erio n that takes into ac c ount not . only the re - d uct io n in 6Cha -out also the po s sj.ble reduction in 6 ct (for a giv en deflection) that may be caus ed by the balanc- in g device . Xherefore . even though 6cha may be red u ced c ons iderabl y . if it is neces sar y to move the co nt rol su r - face t h rou gh a v er y large angle (d ec r easing th e stick levera ge of t h e ail erons) the p ro duct may be incr eased so m ewhat to obtain the sa me 6c& . Th e criterion a s used herei n is strict~ y val id only at the instant th at the aileron is deflected. The use of this criterion for com- puting stick forces durin g a roll w ill give an erroneous indication of these forces beoause differences in the rates of variation of hinge-moment coefficients with angle of attack (a ch/ a 0.) of the ailerons that are be ing compared are not taken i nto account.

Fi g ure 17, a p lot of 6ChaOa against 6C1 for various Mach numbers. shows that the effect of Mach number on the balance e ffe ctiveness is small except for high aileron deflections. Generally there is 8 decrease in the range of balance effectiveness w ith speed. In some cases, however, the effective ran g e increases nith small changes of sp eed but is decreased wit h furth e r increases in speed.

Fi g ure 18 shows the varia.tion of 6ChaOa with' .6c~ for the different balance - n ose radii. T he e ff ective lift- pr oducin g r ange is very s mall for the zero r ad ii and is g reatest for the 0.02c radii.

Fi g ure 1 9 sho ws the variation of 6chaOa with .6C1 for the various gap widths . . ~or small and negative angles of attack the r e s ult s w~re best r ith gap sealed but at higher an g les the results w ere best w ith a gap width ~f 0.00 55c.

A p lot to show the variation of .6ChaOa with .6c1 for various an g les · of attac k (fi g . 20) has been included as a matter of int e rest and , also, for p ossible compari- sons with other ailqrons.

The r esult s of the s e tests indicate th a t a greater a mo un t of balance than that u sed in this inv e sti gat ion is n e c G~ sar y if the . ailerons are to give satisf a ctory hinge moments for u s e on co mbat a~rplanes . The r ange of ·. balance effectiveness and th e ran ge for ~hich the aileron is effec- tive in p rod u ci ng rolling moments co ul d probably be ex- tended by an increase in the balance-nose r a dii.

For the range of speeds te~ted, increases in speed caused a considerable increase in the stalled range of the aileron and in the unstalled ran g e there w ere small in- creases in the hin g e moments. Higher speeds, ho~ever, probably would haye more effect because it is usually not until higher s peed~ are reached that t h e lift and drag characteristics of airfoils are seriously affected by com- pressi·bility.

CONCLUSIONS The results of the tests of ailerons of 0~20 airfoil chord and true contour with 0.35 aileron-chord extrem e blunt nose balance on the NACA 66 .2-21 6 airfoil indicate the following general conclusions: 1 . Increasin g the Mach number up to 0.470 generally causes a small increase of the hinge - moment and lift coef- ficients but increases the stalled range of the ailerons oonqiderably.

2. An increase of the balance-nose radii from 0 to 0.02 chord increases the ran g e for which the ' aileron is effective by acout 8° but results in increased hinge- mQment coefficients with little chan ge in lift coefficients in the unsta1led ran g e.

3. An increase of the gap wi dth ~ncreased the hinge- moment coefficients s lightly with little change in lift coefficie nt ; however. a considerable increase in the st a lled ran ge of the aileron results. The magnitude of the increase varies with the angle of attack.

4. The amount of balance tested, 0.35 aileron chord.

gave no case of co mplete balance and in some c ases the unbalance was relatively large.

Lan g ley Memo rial Aeronautical Laboratory, National Advisory Co mm i ttee for Aeronautics, L ang ley Field, Va.

REF~RENCES 1. Purser, Paul E ., and Toll, Thomas A.: Wind-Tunnel Investigation of the Characteristics of Blunt-Nose Ailero n s on a Tapered Wing. NA CA A.R.R., Feb . 1943.

2. J a cobs. Ea st man N ., Abbott, Ira H. , and Davidson , Milton: S upplement to NA CA Advance Confidential Re p ort, Preli m inary Low-Drag-Airfoil and Flap D ata from Tests at Large Reynolds Numbers and Low Turbulence. NACA, (Loose le af ), March 1942.

12 NACA

TABLE I. - ORDINATES FOR NACA ~6,2-216, a = 1.0 AIRFOIL

[Stations and ordinates i n percent of wing chor~ : Upper - sur.face Lower surface -.

-- Station ordin~te Station Ordinate

0 0 0

Ol 1.2~0 -1.1ez.

.~99

·t -

• bO

• 40

-1.3

1~128

i:~5~ 1.372

-l.b~

2.362· 2.5bO ":2.1

2.538 4.846 __ _ -2.9~2

5.154 ,.50~

7.340 .42 7.660

-R. 0

10~162

-5.140 - .106

~.8~8

6.276 1 .8 5

5 -4.930

15. 1 a

20.1 0

~.8 0

-~.56t -

7.1~

25.121 .879 - .05

~.8

2 . 00 30.100 22

.3

-6·t

-8.736

-6. 76

4 3 .924 - R 5. 076

- - 0.051

8.980 -6.838

lK,94

_- _ 45._026

9.022 -6.902

4 .97

-. - 50.000-

50.000

9.060

-6.8~4

a.875

54.975 55.02~ -6.6 a

5 .952 -6.35 -

bo.Ot 8.~2b

7. 62

65-.0 - 7- 802

6 .933

- -

- - -a·

- 70.081

6 .

919 - ,_ - • 997

b:~t6

7 .'913 -

-4.070

~5 .087

0.085 O

a. 641~

7~.915

-3, r

85.075 8 .925 -2.0 9

3.395

- -

-2.103

-BR·945 ,- -1.0 9

05~ _ 90.

95.02- .913 -.281

9 .972

100.000 a 100.000 0

L.E. radius:

1.575

TABLE II. - LOCATION OF CENTERS OF BALANCE-NOSE RADII - -

[stations and ordinates in percen~ of wing chord]

Upper surface _Lower surface Balance-nose ~---~--.

radii Station Ordinate Station Ordinate

-75.00 -

0 0

75 '.17

-4. t

~.83

1 - : 75.33'

75.~0 -2.~ .~5

' 3. 2

2 75.' 79- -1. 3'

75, 7

TABLE III. - LIST OF FIGURES NACA

--

Ba1anee- Fig- 'Variation shown Gap width nose ure r O adii {~ eha against Oa Oc 0.0055e (by pressure distribution)

6 against .01e

5a .0055e eha I" (a) .0005e (b) .0030e against .02e ( e) 5a .0055e 7 eha (d) .0107e ~ ( e ) .0055e (sealed) r(a) .Ole .0055e 8 against M eha .02e ~(b) .0055e (sealed) .Ole

against gap [~a) Varies

eha (b) .02e .01e 0.0055e ((-a) 10 against a.

eha {

o (b) .02e .0055e (sealed) .0005e .0055e 11 against a. .02e e7, o .0107e .0055e (sealed) .0005e .0055e

(g~~) _ against

12 M .02e .0107e 5a-0

{

.0055e (sealed) .0005e (b) raj .0055e against .02e 13 e7, 5 a ( c) .0107e ( d) .0055e (sealed) and eme j4 obtained by Oe pressure distribution .Ole 14 .0055e r!

(a) Variation with a.o .02e (b) Variation with 5a .0055e against M 15 C~ .02e { .0055e (sealed) gap

16 e7, against .02e Varies

{ ~chaOa against b.e 7, 17 .02e 0.0055e (showing change with 14) ~ 6Ch 5 against 6c!

a fshow!ng effect of balan~- 18 .02e .0055e nose radii) O (by pressure distribution) C against b.e],

~ 6 h!Oa

.02e Varies (show ng effect of gap) against O 6c], ~ 6cha a .02e 0.0055e (showing spread with Q.o) llACA

Flg. 1

-

NAC~ --To leading edgB-------- C_2+.00"-- ______ ...-Fig. Z r+------ Oq~Q20o~ · ---------,~ Gop A=O.0005c B = .0030c C -= .oO.5§c 0= .Oi07c Ifi fig€' OX'IS Gop sealed with fl8.(lble sheet Figure :J . -.Aileron sect/on of NACA 66,2-216 J a-I.O olrf'oll snowIng variations of' pal once-nose radii and gop.

il> 2: o ~ b;j ...,. w (J'q

I I i I

I

-

~OS l

]

'< -

: i !-~

, I i I 1-] 8

1-

1_1

+-j aj_rfoi

--1---

_ I I I i

I

W--

L- cLord

; I i

7 l .

foot 2- ere tunne a.

:b.

I I I' I !

I I ! r-- I I IL-l

r fo l i t y ' at:nosp I

I I I i

-T

stabi

-

nUIn',:,el I I

Ii, I I I I

'

i

tandard the s

--1--t_-1--+--t--+-~I'

Ma6'l of

----~--+---}- t

----L

, ; 5

I !! ii

i

-

tes ion

-i

--j-~---I--! t

-

base"\.

of sec ! I I I :

I

i-

-l

test I I

I ' I . ! I 4 !ltunber values

-~-t

-+--~--+- ds

---.L

J-t _

--~----

l f foot

----~--i--

· j--t-r

' I I I · ! S-

Iii

t----+-

--

-

--1- Reyno by

- number

I 3

I I

j__

-2 . 5-

LL

__

-ll

the

I

' I · I

i n Reynol

-t--t-

--

! I I r--i I i ;_. 3 . -

I L_+

ill

__

-----

gure

----L--t---

--i---+----t----~!l ___

_ i

r--~! ~ l !-- ~I--t---

:--I

I

:l!'

. 5 . 1 . 3 . 2 .", _

~

c i ~

, ~

pl

1-:':" fj;! Cl ~ Ixl ..;. of>.

at)

I i

1 1

I ! I I I I

I I

.-1

-~I

____

'-1

I I I I

I I I

I! I I I

I

1--'-' _ . 3

Cha

+ -+---1

, ion ,- , ___

-

·c 1-431 f---

I I , I I

L I I I I I

sect tion

±±-

-t

-1

vario

-t-

balance- I I I I ;

I I I ! r

. 2 st:-ibu

0d for

and d i

--'-"1

---=-y--

I I ! I I j I are

I Ii) ' I

I-L I--f-_ I

-[-r

,Q

"'--"'1 - peeds

-'--l

' -~ s

I- I I

I . 10 I ~ otte~

I I , I ' pre3SUl'e

- I--j'

I, . 1 l i i .

-1..._

p

-

- OJ

-

rarl pressure-iistritutio!"!

-----T

-- __

:----, I I

I ,

t I I I , I l

!! I,' I I

I! I, .

-t--~---.

V~lue, Cl.l"Git'..ui8s , nose

and

--

·

----j L 7f";

I i '

, I obtained

r I I I

I

-r

VI pq(y-i ----_--1.

.--+-

-·r

· -~+-r-

-

,y I l

1 I I--· i I I

I ! I

I I ! I ---

-~-t--'ri ..

ol - -

-r-

-

----t---t-

_ I I _.

I I i ! I sprui.g-oalanc.e

I I , I~/I COefficient

~ I I I t---1-_ I I I J 1

.~-yo T~-I ,_,---, ients - . 1 --r

-l--t

-',---'1 _ _

-1-

-

/ -

. I I I ) 1 1

i! I : ! ! I ; 1 i __

1-" I ! ,

I

oeffic - ffioment betVleen c ge

---f'J ___

I I I i I in

I I , ! ! . 1 I

I ! I

" ,'

.1- -P-- .

-

- - . 2 riso::1

--r--,

--+

--

ion

?tV!"

--~--+

-·

. I I I'

+-:- I, I I I ~- .

I I i I . i

compa q

-- -')t---- /- ect

- s hinge~moment

YJ! A

! I i .-

, l I I !

I ! " ' I ,

IlL 1 I I

I

- . 3 Aileron

r-i-- ~

I " ~

I I ! I i j~--H-- gure

. I it ! I I I-.--~-L

I 1---+---0

2r~-T-

o l

2l---

. Fi

-

- .

"'\ I .s.:1 0 m o s:: ~j m t\; r.: $-, n, (J) ·,...1 t>:, Ul I=l m o m I s:: r. o Ul :) "I ·rl .p .D . o I=l Ulrl S 8 Q) 00 ~1 .p ., I 'H 'H Q,o ~ o o .,0 ."., .,..j Ul U) >=1 o ~ Ul .£l ·rl .p rl ·rl <ll rig.S , ~ -20 -It) -12 -8

-4 o 4 8 12

16 20

Aileron O/?9/e, oQ, deq

FigVre 5. - Voriotion of olleT'O'J 5ecti()ltl lJil?()e-I'l'V'1'?~/1c coeHici eat" ('viti! olleron ongle (Oq pressore . di.:sviwtion). Babl7ce-/',IfZe ;-odi/=O; r.;op = O.OO5Sc.

"'

en

NACA ~ ,2.

~ .I ,2 .1 -:2 tJ ~ ~O

~

"- \j ::t: .

'C III

e .J

1-...

I::: IUO ~ ~ , ,2 IU Ch . ~ I ~.

-~ a

, u ~ .2 c: . \:)

1.1

- .....

~ At ./99 -.I -.2 -20 -/6 - 5 0/ +--+--+----l

o

/ ~ : i +---+---+--1 .I

-.1 ./ -/6 .8

-20 o /6 eo

Aileron on9/e, 0(1 , de O g ( 1 block. - 10/+01 Fiqvre 7. - )/oriotio/1 or aileron section l7in~e -moment coefficient with aileron anq/e. Bo./once-nose /Vdll =o.OZc.

NACA 1 1 1 1 r 1 1 ,I Fig.7b

.21--""111: 'r-+_+___l-+--++_-+-_++_+_+_____l- Anqle of attock, cC" , deq - .z

- -

~,~I~+-+--++_1__+_+_~+ 5J~_+_-+-~

./ 1--+-""_ . "" I ~:-+-"_~+St-u . L.,!-K"::'O""":"I"----+N-..--+- . --r-~-+--+-_+_-+-+- ~ . I~ ~ +---1--1---1 .1 ~ ~~ ~I'o. ~I--~ Ot--~n·_+_____l-~&-~~rr-r~~_____l-+__+__+__+__+__+___+_+_____l

o

~..fl ~t-- ~~ ~:---+--+-- __

o

-cO -/6 -Ie 4 12 16 20 Aileron onqle, 50 , deq (f black =- 10/40':) Figure 7. - Variation or aileron ~ech:on hlnqe-molllent c,QefficLent with aileron angle. B . olonce-nose rooli = O.OZc. (Gontmued) fi9.7c .2

-

I ~

o

-./ -.2

o

-./ • -IE -8 -20 4- 8

o

(f block = 10/40") Fiqure 7. - Voria! ion or aileron section hi.nQe-momel1! coefficient witIJ aileron ooqle. 8C1/ol7Ce-no~e ro(lii =0. OZc. (Continued) • NACA 1 I t l I, 1 I I Fig7d ? 1--.p:t~~"---+---+---f---1'----t--+--+-+-+--f---+- Al7qle Or attock, cCo , deq - 2 ,C,. ~ I I I • r--t--+-+~~---f---1,...---t--+--+-+-+--f---+---+--+- + - 5. / - - --~ ~ 0 0

N

1---+-- . ::--.! x 5 / +--_+__+----1 / / . ~ ~ ""'~ '~ (J /0.2 .

r\ ~~ I'o-""~

Or--+--~ ~ +-1-~~-PQ+-T~~--r--+--+-+-+-1--f--~--+--1 0

~~ "I"'-: :<h;-t--~

. Z -:114:1.

-,/

a

- ./

o

-8 -4

-co -16 -IZ o 4-

Aileron onqle, Oa , deq (1 bloCK: /0/40') Flql./re 7. -Vqriation or ai/eroa section hinge-moment coeHicient wJlh aileron on9/e : Balonce-nose rodii=O.02c. (Cont inued) Ft'q,7e ,2 , "( -./

o

-./ -./ o - , I -,2 ~ _Ie -8 8 - IE 16 ;:'0

-16 o 4

Alleron onqle, 50' deq (7 block. = /0/40")

Fiqt..lre 7. -Yoriotiol') or aileron sectiol") hin~-mo/71ent coeff i cient with

aileron o/79/e .. Bolonce-nose rodli = O.OZc. (Concluded)

F/g,80 NACA of o · ttock, oC deg Angle o , I I -5:1 "(

-- ()

---

------- 5.1 - --

---

/0.2

o

~ 7° 6..= 1- __ f--- - -f-- 1-

'-

"

1"- , 'i -:/ I- 1--

-

- -

--,

- - --

- --

-- --

-- r--

-- 1--- ~-

-

r- t---

-

--

- '-""

-

i LA ....

0 '. - t-- r' t-- - t- t-- - r-- I-- I-- -

- -

8=0 ~ .

1---

-- -

---

-

---

-- r--

- --

--

---

- t--- 1-

--

~ I--

- r-

r-- - t- r--

--

r---

'-

-

...- ~ ='"- -

-

0 =-7-

- -- -- -- I--

--

-

-- ---

-- -

-

- -

-- r--_

I---

---

1---- ~ I-- -:/ .A .A .2.

.3 :4-.5 .2- .3

.+ .05

Mach number, M

V bock:: /O/JO)

(a) Balance nose radJi = O.O/c.

Gap = . O.0055~ Gop = O.0055c (seol!3d) rigl;re 8. - Variation of aileron .5ection, hint;e-;nome"t coe.ft'icient with Mach m/II'?Der.

IYACA F/q. fJb

-

I I I I I l

Ar1qle of attock, C£" , ci.¥-

I "< -5./ ---- -- ---------- 5.1

----- -

/0.2-

o

"'--

-

-- I--

-

I---

-- 6 = 7"

-.1 ... _-

-- ---

I--

-- --

-- 1--

-- I--

- --- r--

--

- -- r--

- --

---

, ,

""

,-",- I--- 1--

I-- -- -

I-- -

- - -

-

---

-- -- r-- 1-- -

0 :: 0"

--

r-- -- -- --

---

-

r-- 7- r-_ r-- -

1- - --

-- r--

r--- ~

--

6 =-7.

~ 0 f--

.. -

--

-

- -- vv

V\r

-- - - - -

- --

---

1---

-- --

-- -- -

--:- 1-- I--

--

- -

- r-

t--

--

-:/ .2. .3 .5 .2 .3 .5 Mach number, M

(t block ~ 10/30 'J

(b) B%nce nose rodii = O.02c.

Gop =o.0055c Gap =Q005Sc (seoled) FiglLre 8. - YQriotiol'l or aileron sectiOl? hir1qe -m(Jment coefficient with A10ch numbs;. (Concluded) ~ ~

~

J;.

<,1'31 I - - j, .

t---,..

.()/O 10/30') ;:-

r------I--

/9~ ~ . .4/8 - block-

_bl..-..=-:::::b-r----1I~

.008 11= r-- (_ O l+M- - gop.

f- 5./ .a.16 = --+- 5eOledJ wi.th do _!

"- GaP - .IJ()+ .

I=:jt:=lF==t-::~:::;:::::\:f--:-==-±:f--:--=

coerriciel7t" .tt?Z O.Ole.

= chord -+-::::t=:::-:_-+t----L_~-- ::;:::-

r-~-t--r-~-+--+-~-+--+-Jl--

of

~ .'f $--"- 0 0° rodii =7° -7° 0 -mome/7t Q rroctiOl7 1-----+- hinge .

.010 f-.- -I--~ Gap, - Bolonce-nose section - .008 -I- f- (0) - 0° = .006 aileron o -+-

r-~-t~~+--L--

of cC _ -

__

.00+ ~ _ Variation .ooz.

~~~

.9. - ~~ -

~l~l=l=+=+=t=t~~~~i~i-=-t-=-t-=-t-Jl_-"-Jl_-'-'-'-'~=J.::ar----I..-..-+--

$ 6=- + o 0 0 ./ -;/ -;,/ -:1 -:-Z Figure ...

~ ...

<J~ ~ .... \.l III ~ I.J ~ ~ ~ & ~ ~ IJ IJ) ~ Qi

~ .

~ ' ~ ~ ..... -t: i-l "- ~

'" ::!1

~ ). ~ "0'

t''i5/

- ~ h)

---; 4'"- .010

--+-, 10/30

r--+_-.J := ck f---1 ./97 .417 (Confinuecl) .ros' blo M M- (7 ~ + 9QP~

{0

-- /0

_1_1~:J~~;t---I--_~

.006 '" 5.

with !_ =

'-

Gop sealed CCo , r- I -'- CO2 . coeFf'lci(?nt O.OZc.

I = chord -=:::t-- I I I Q>- ~~ 0

~-

7j

f' 7°

°i~r-~t-t-t-+-~~~-L-

I rodii Q~ 50 ~, I fraction :-~ 17I"''?!7e-moment I 0'0 ,

----- f--

Gop, __ -I-- B%nce-nose ~ect-ion .

r-- J (b) - I = .~ aileron -I- do o~ I-- .___ ,004- I--- I I Variation .002 ___ -1-- - I 9.

t=f~f=i=~t=t=t=+=~==~=t=i=J=j~=l=t=1'=

l

(j) <!r==""" ...

o

0 0 .I .oj -:2 .J -:/ () ..

-c: Figure \.) \j ~ ~ ~ ~ ~ ~ ~ ,

~ 8J

~ ~ ~ ~ ~ ..... ~

F: ' 10 'NACA . Iq!

:- .

~ -;0 ~ ...... "'-...,...

----....

f:-....

~ , ~

~= i·

" -./0 ...........

1"---, ~

'" ~ I:::::::- t::::--

--

-.::::

--

/II=Q139 c - - - ht .-I18 "<:::::: 1:::::::

~ =-

r--

,j S~ o· -- ...

.........

. .,.:::

I--

~ ~ ~ I

-.... I

~ I ',--(0) Ba/once-no:se rodii =O.Olc.

d)_ gop = O.OCJ5.s~ gop r O.OOSSe (sea" r- ~ :::::: ~ ~ 6. :-7- ' ~ t-

--

r-- ..............

r - ""-'"

~ ~ ............

~ s ~ ~

~

-4 8 -.f. 0 8

o

-+

A>I'lQIe of attack .. a:.., deQ.

~ "- -:-05 ~

I

I - ~

.. \

~

rI ~ ~ oQ= 7° -.10.

~ o ~ '':::

~

r-- :::--:::: .....

~ f:-::-...

-./5 I,

-

r- ............

-- ~ "-

M= 0./9 7 () "- ---M=.4IT -Q "1 I'---

o

r--

-

--

r--.

~ r-- Flo:: 0° ~ -:05 ~

CIt ---

~ ~ ~ t-....

-

~ -:10

r--..:: r-- ~ , , I"- r- (b) B% ' nce~no"~n:xIii= O.OZc.

""

~ -.15 90P =O.OO5"5c(se"led)_

90P = O·~"S5c

f-- .~ -c::: :::::::::: u,...

I ........

~ .05 -....::.. -.......; . ~ ~ ~ ~ D ~ ~

0 = -7

~ 0 :- " ~

-

~ , ~ :::::- ~ ~ too.

o -4 8 -~ 0

Angle o-{- at-tock JCC , dti9 O Fi9G1re 10. - Yaribtion 0-/ cn"/eron section hingQ-momenl' coeflicien"t with on9/e ~ attock.

F.' /I NACA 19, -...

/;.

\Yi) ~

r

J ~

'<t

v

& A

fA J fI

.tJ ,

/I}

P

I A

.6 ,

II

f--6ap=-.OOOSc ~

eOP,.oo"1

. I

( J

I I

v

V

~

r

-/;

/ /

~

0/.0 ¢'

f!/

If

V

A

;//;/

);

J W /

)I

I(

~ If J

j

/I J iJ!!

\ ,

V"

~ A

I'

~

r--.60p= . O/07c (jap=.0055c "-

(seQ'/ed) I

r

I II

. ~

'f

/ /

Mach nllmber) M

o

El 0 . /98 II

I

J + · ;';88 0 . 357

I I 4/7

• <i> 47~

(f

fA'

1/

Ii

If!!

ir

- -I- 0 + 8 /Z -4 0 4- 8 (1 block c 10/32 '') A n9/e or crr-toc K J c£", deg

Fi9ur~ /1.- Variarion or seer/on lirt- eoef?'lcit'nr WiTh an91e or

afteelr. 8C71C!/7ce- nos~ rtJdi/ = Q. 0;; c; 00 = 0": :r;J iJ> ~ o 1 r" f-' (\J IJQ I I I ' 1 I I I ! i : ! I , '1

, J

__ _ -1 us

! ' I I I I I I . 1

. S i c

1- ~

var r

I, I ! I I I I I

'I II; +---:--, 07c fo Gap OOj5c ."''-'005C ,)1 0055c(sealed)/ , • .

.----r---+----j --l-t-'

---1----+---

- O'

1 I I I I I

I' . 5 mb + 0 X c

~ nu

a

/

I I I ' I I ! I I I i '

!, I

il--r-l--l

+-),- j',iach

-+~--l--+-+--I '~-

. --

~

l/i/~i I 0 1 ' I i I : I I I ~L

I I .tr" I', I

. 4 with

_

___

- --1--1_

/L/~ /4

:7/.

0 !of.

I : I '+:/"L I : ! i

I I,,/! I ! I

I

, , oqI()(" _

--1' ...-/! L[

_

u pe _._ ' y ,) '!11 I I I / 1./ ! I i I ,

L

Ii' V .

T

y

sL !1.U1T"ber

-, X

+'~/)

nt

---t-----t------l-----r---+--

-L-~~:-'-

,i//-~X,!'

i -,-'- 1 I I I ' I I ! ~_ I ! 'I I I I

Mac:c.

V

t _ curve

--¥fol--

:) __

.._-v-

----

ift

I I ' I 1 ; I I

,L----+-- . 2

f l

.-

-r.:;t/ 0a

- -

of I I . I i : : i

I I 1 I

lli on i 1(2 at widths vn.riatiOn 1- i -:-

I I I I , I I I I I

!"I--1 'I' Ir,-

-r-

+---W--+--1--1

. 1 1/J ga'(J Var

--+---'---H----r---r-f-.-

. - - 00J

tt

I I I ! I ; , ! I i I

I !_

Iii I!

le'

r-

-

,!

I -

J . 1

~ I

'I I'

~l.heore~~ca, ~ =-

, ----t- gure -J

--- ---~-t---;-~+--

1 l I ,..-t- ~ I

! r ·

Ul

o

F i 13 llf-- 0 9 . . 1 2 . . 10 • o ~-· 'r V OC"l, NACA F/q,l3a 1 .1, I 1 I . 1.

~t--+- Mach n um b er ) M --+--+-----1I---+--+--+--+--+---+---+--+------,j-+ 1.6 -i r-

1---1---1---1 ° 0.198 -t .-t: J

+ .289 ±...-f--¢J- x -

-

1---1---1-- 0 .358--+--+-----j1---+--+--+--+--+---+--+--:::>"'!'1lTT~-+__+_~ X .418 i ~./

I--~~ <> .455--+-----i_~+-+--+--+--+ -b.-<fv---,;'tl'-+-+--+--+--+--+

wro /'.

1.2 /.0

< > /~ I~/~ L '

CJ lR'_o .8 o .6 .4 [

0<>8

-8 -/.0 -/'Z -20 -16 -I.e -8 -4

o 4 8 · 12 16 20

AIleron angle) 6 ) deg (0) Gap=O.0005c.

Figure /3. - Variation of s(}ctlon 11ft coeffIcient wtfh aileron anqle, Balanca-nose radII = O.OZc.

FIQ,13b NACA 1 I I I I I.

I---+--+- /YI 0 c h n tim b er, M · -+--+--+--+--+--+--+--+--+--+----il---;-i . ±~--::::r-I--I 1.6

1--1---1--- 0 O. /98 % ~rt\ L_

+ .Z88 ./ . - x t--t--+-- 0 .:3 5 7 +---+--+--+--+--+---+--+---+---7"iIo"/~"",-+--+--,)Oi+---T.!:::=---1 1.4 ('0 - 1---1---1--- 0 :~ ~ f -+--+---+--+--+---+--+--+t'~l/i_. ~T--J~)6.:'}"+K_ . ~~_+-'+~----.-',r--I ~

n"th /A/ ~ [ L

, I.;'

~ 1.0 .8 .6 : u'" .4 ..

~ ~ .Z- . ....

u ~ ~ 0 Q.I <:) ~ .

~ -2 ~ .

~ ., LiJ l..4\ 1 I J ~ VI '"'- M = :1-5ti.Jt---t-- +--+--+-+--t---+---+--+--+-- -+--+--+--; -k:~ p -/.0 v I -I.Z -8 -..,.

-16 -12

o 4 8 II? 16 20

Aileron onqle, 6 , deq (I 61ocl< = IO/4{))

(b) Gop =0.OOS5c.

Fif)l.Ire /3. - Variation of section lift coeth.'clenr with aileron anqle .

Balance-nose rod!' = 0.02 C . (Continued)

FtC/.I.Jc NACA I I I I I I f--+--+- M Qch number, M-t-_+--i-+-+--+--t-----+-~+_+__+_+___l 1.6 . ~ r--r--+--+ 0 0.195 +--+---+--:-+--11----+-+--+--+--+-+--+--+---+--+-1 ('f) + .287 f.l ~ r--r--+--+ 0 .355 -t--+--+---+--+--t--lI---+--+--t--+--+-..-..!-1.D~~----!4---I 1,"1- x I

:~j1 +--+-~-+-+~--+--r-'--~~'~:~~

t--+--+-~ 0 ~

wO o

1.2 1.0 .8 .6 .1- \J a .2 ;....

~ ~ ~ ~ Q., -.2 1--.

"

-.+

~

l' ~ -.6 . "1 -.8 -1.0 -1.2

-20 -~ -& -8 -4 o

4 8 IZ 16 eo

Aileron on9Ie,6a, de9 (z block == IO/4()") (C) Gop = O.OIOle.

FI9ure /3. - Vor/al1on 01" section 11ft coet't'lclent' wit/) aileron angle

Balance-nose radii = aOZe . (Contmued)

NACA NO.13d 1.6

-

ro III ~ , "-l 1.2- /.0 .8 ff ,..I <.)

..... '"

, It <l .......

t z

~ ~ \,) ~ .'t-.

, , -2- - ~ .

K ~ ~ -.4 -.Ii -.8 -/.O -IZ

a 4

-w -16 -IZ -8 -4 8 IZ 16 ZO

Aileron onq/Q/ co) dQg (! bloc.k = 10/40·) (d) Gop = 0.00.5:5 C SQolod.

Figure 1..3. - Voriation of s(;'ction lift COQrt=ic;gnt with ai/gron onqle

Bolonce nOSQ radii = O. OZc. (Concluded)

NASA Fi g . 14:a

I I V'

l- l-r-- -r -- o----r-+- ;r----

I I I I /11

,-- - -- -+--- ·--- ------ r- ·----I----- --1 -1-- +1----

. 6 - --- J-+--- ----l--J--+1---t---- -l!. --·--·-

I I I! d I i

,-l u

-- i---i- -- - -- l---jf'-r---~ I -

~ +- ,

:! i / 1 ~ I I I

~ Q) 'r l ()

_4 ---j---1 --~rf -- + ---- 111 - - ~

·rl (r; '1-1 ill

I -+-- --- - 71t--~-I--_:--T-- ~o

U T

+> ct ...

·rl .-1

. 2---+ ! 1-·+----- - J- --- -t-- -+-l.J .1 ]

i A J I i ! I ~

' rl +> c.> Q)

1--- -i/- - 1--1-1 r--l----' ~

U1 - o I --!------,------;:;t------ Lp -- -- 0 ~

If I 1---6-(- *1 -- I ~

m ~ -- H- --- I CDC /4 1 ! . s

1..- <1>7 1"-= - -- 1 --: 1 t- t~i-i--- :1

L

-J -- - ---'i--t---- - I +--+-+--+-- -.1 ~ ,

I )1 I r j B a l e.~ct; - no se r adi i I .§

'--~l r--- - t--- ---- -- ------ ~ ~ . o:)C ' --+---1

+ I I ; I 1 I I (J)

" I I I' I I

l Q

- . 4l2.- --f /- - -.--j - - -- - -+ ---- ---- .- -- -- -;-.. ---- .. ·-L.-··-- ·- -T----;- ---~' •• I I I . I I _ I ! __ ~ __ . _____ _ L- . - 4 0 <1 8 12 A ng l e of a t t ack , ao . deg 0 _

(a) Variati Q ~1 ii i th ('t 'o ; oa = 0

Figure 14 (a,b).- Se ctio n lif t co e fficients and s e ction pitchin g - mo me:r.. t coeffj.dent " oo t ai ne-i b~' pre s sure - d is- tributi on . M = 0 . 358 ; o ap = O. 0055c .

NACA F/g. 146 -.~ .I ~

-

N

~ '--.

"= v

.---

-~ 'J

.w"

,./

~ I::::", ""'I:

.§

CC '= 7. 7° ~ o ~ ~ -";:f::::::::-- -::).

- ~ .1

x ('(l I-- I--

-

~ ~ ............ r-- t--+

-:1 ~

r----

~ ,

r--o

~O

~ ct:- = 0°""" o ~ ~~

~9

.<J

~ +

.~ -:1

-x;--- t--- r---.

~ ---<:

~

.-/ ct

<;V

1.2

.§

/ l--y

N II) /

~

....----r

- ~

--

......

~ /

J.o u

V

~

\l /' Q '-

f

-<:I .8 \ cX.o~7.7"/ '- r",

V

'"

~ ~ .6

--

~ ............... _x f- ..".

V

-

V-

~

/

@" -+ + ~

/

V

/ . V )

~

oC = (JD

~

V

// Balance-no.se n7dii ~ I~ +

-

V

-

/- x O.Olc O.OZc

V

# .....

-4

- p-

-

~ ~ ~ ~ 0 4 8 ~ M A l/eron OI?9/e, 6 , deq Q (b) Variation Wi.th 6".

Fig~re /4. - .5ection /lfi coefficient~ a'ld seclfion pitch;n9-hMm~l?r coeffU:ients obtained by ~.$SIIre-cl~r;bl..lti.on. A1==().3S8; 9ap =O.OOSS"c. (Conclvded).

~ ~ ~ "- (" ~.

(sea/~d)

I I

.5 5.1' 2..

5./° 00 _2 '~17~ - I-

T

a;,=~o.21°

--

.005DC

--

-- r--

.4 -,

--- -

1..'131 f-- -- -

~- ~--I-.

Gap:&.OO5sJ r Gap:

-

--

-- -- --

v-

.3 0 -I-

--

-

---

-- --

f..-- -

V

-

----- --

..--- ..-- V -- - -- -

.2 v .6 .4- .2 to .8 /.e. -:2.

number 0 /D I .5 2· 77· 0° 5.

5F - -2.. M

-

- -

a;,=lor Mach

--

-

--I

--

f-_ "-- l- .4 -...

with

-

0° -

-- --

i- -- i---

number:-

--

_.-

~o - 1-

1--- ' ·3

-

- -

-

--

r- ~ Mach coefficient

-- ~ c--- f--

1--- !-- 1--- 1-- .2 'v ;irf c.

/.0 .6 .;f

.8 .z

-.2 ".4 0.02

=

'0.2 .5 /0 !< 0°

7.r

5./.

_Z,o_ -.If:r ..secrion

-- _

-

.........

a; rod;i

-

or -~

-

--

-;:r

-

--

r-::::- -

--

---

--- -- -- -

_7° - .3

---_

Yc;r/Qr-Ion Ba/ance-nose --I

~ -- -- ~

f- -

-

~- - ...--- - f-- ~

15.- 10/32") ·2 =: a .Jf .a ..6 ·2 -;2 -:4- -.6 F/gt./re black "'1 '" (I I:: \l \.l IJ I.;) (.) IU .~ ....... .0 ~ ~ ~ ~ .~ 1::: '0 N-"1CA F/g./6

--

\

v S?

C) \ / ~ c::> 1\

- "-

\I CO

I ~

: ~ \)

-1'-.."''''' ()

~-.....~, () -2

I -.....~ ....... ~

c:) . •

/ ~

- .

t--"-

II " Ii \I I '-.../

'"

If ' ~~~~ !()

~!

I

0 C) II

-

"

I.() '~I \) () I~ -.....

~o I

tOl

- ~ II

I~

~ I

/

I . ~ () - ~ I ~

1/

I II

I

- -

- ~-

I t\t 1-1

~

'J'tl ~ C?,

;' II

I

I 1 ~ ~ ......

/ / / / I

~ ~ r,o.

;+---0 00 -;....

c::: c::: . t>l C) ....

.u -+: ~ u 0..;:: C) .

Q) ¢ C) '-> u~ I I C) .....

\

I ~C) -- I ag.

. .... c:s , .c.!)

II I \ c::: . ....

C) .....

. .... 'ti

\

co 't~ II .- C) Q) ' V) Q) a \ 1..j..V) C) 0 c: <) 0 \ c: t'-...

I I()

I'.. () . C) Q) I I C) --i:: u II I C) C) c::: ~~'I " () .... C) ~c ~

I

a"-' C) C) Il:::::'~ \ o I

I

~ I .l.

e((5

Q I -.....

I

: I Q)

"- ~ I <\!

J; t).,

I

4: a

I ~

/f

/1

/1 J I

I

v ~ (;) ~ ~ 1" (';'l..IfJ/:J/.f.JEJo:; If/I f./o/;.:Jc;S ______________________________________________________ ____ _________ ~ ~ .....

)t .....

~ ~ U.

.6 .- _ _ ~,

~ ~\

~~.

\~

~

.35"7 47iC ~

ZIJ9 417 .4 ntH7Iher) O./9lJ ~ showing ~

~

• )(

• e 4

Mach ~ ;!.

.,~ .

--

A-ticl Io.e" I SID . ' = = ~ D • ' coerricienr;, a: ex; lifi- LlCl -;2 I......

~ection O.OOSSc.

= IiIII-.

V

or ", -=4 '"'0-

J

gop

(

!~ coerricienr,

~ I~ ~"'"

i ; "'e>

'\

~\ ~ ~

~~

) -·6 h-r"t-

it

o.oec

~

I~ increment-

r

j .6 radii:

yi

wiTh

:;;...... ~~ )\~ sect-ion

~ ~r r1

< A

~

n()$e ~ ~ or .4 ~ 't:J~ cr/t-erion Balo-nce .Z rorce =:;";:1 = ..

., /ncren?enr a:, cL

o

number.

-

conl-rol-

--- Mach

-:Z or

~

wit-h .

~ ~

-:4 ~

~ I~

~~

' Vario-t-/on

change 10132:) \.l -

1\

-:6 17.

f--- block Z 4 0 -z. -4

4- (1

if!

-z

., Figure

.... I

~ ~ I:) ..... "- ~ ... "- ~ <\) 0 I... ~ ~

~tJo <:J ~-4 . ~ " 1- (5

NACA Fi g . 1 8

- ~-

!

o +--t-----t----t ~.

C1l co

-1----1-_ ·

I ; I <l - 2 I ~ I

I=l I

o · rl I

I

- -- - ~ 4

--

.p ·rl H o Q) o H o 'H ~ 2 o H .p >1 o o -. 6 .

-. 4 -. 2 o . 4 . 2 Increm en t of secti on l ift coefficient , OC t Figure 18. - Va ri ation of Oc.baoa with tlc l ('by pressure distribu tion) showin g ef f ect of bal anc e-nose radii . M ::: 0. 358 ; gap::: .

0 . 0055c .

" ..... ~ ~ ~.

I~ I

I I I I

I

~

·6 I I

I I

(Sealed) showing -~~~ ~ c I I' j +--n'll

• c

·4

\

\ I .

00SSe .DOSS -0/07 .

"3 ~~~ O.OOOSe .., I

I

.1 .a

+ EI x (;) 6~ k- Io.~o ~ . I coer"ric.ient-

~

_ 0

~ I

r·}I"I.~'&t I

T-~

/lr-r z

J

.de

11 )

-:~

_I

seet"/on .

or --:4 >( ~~ I'--v..'r~ I'+--~,.

111/

"\.

\ .

coerrichy,r- ""...

M=O.417.

~ 1l;J\ I "\ ~ I I I I L

I. -:6 ; ~ ¢ increment" /irr

II

·6 wit-/} =OOZOc .

jeer/on

11ft I

11,11

4 radii or

~IIIIIIIIJ

crirerion

' '

~

I

.Z 0° = Increment" ~.o-S/ .

o

8alance-nose I · ~ con-rrol-rorce I gap or -:2 -+-- or I ~ \ . ~ 1\ k I ~ I~~ crl . I ~ -:4 '\ ")

Overview

Document

CASE FIL ~ ~

COP

~ACA

quarter-chord point of airfoil (~~~~)

a i1 eron sect ion h inge-momen t c oeff ic ient ' (_~L8)

m C /4

q dynamic p ressure (~pV2)

+ =

c m I

=

y = ~~ (~)8

~ = 0.304 (theor e tical factor f o r NACA 66 ~2-216, a = 1

c ~ = 0.963 c 1 '

a o = 1.023 a'

a 6

from values of Cha at a oa of ±5° as compared to a

the value of the slope of these curves (a 1\ a~d the

(~:.~,'; , f or a

TABLE I. - ORDINATES FOR NACA ~6,2-216, a = 1.0 AIRFOIL

0 0 0

·t -

• bO

• 40

-1.3

i:~5~ 1.372

2.362· 2.5bO ":2.1

2.538

4.846 __ _ -2.9~2

5.154

,.50~

7.340 .42 7.660

-R. 0

-5.140 - .106

~.8~8

6.276

1 .8 5

5 -4.930

15. 1 a

20.1 0

~.8 0

-~.56t -

7.1~

25.121

.879 - .05

2 . 00 30.100 22

.3

-6·t

-8.736

-6. 76

4 3 .924 - R 5. 076

- - 0.051

8.980 -6.838

lK,94

_- _ 45._026

9.022 -6.902

4 .97

-. - 50.000-

50.000

9.060

a.875

54.975

55.02~ -6.6 a

5 .952 -6.35 -

bo.Ot 8.~2b

7. 62

65-.0 - 7- 802

6 .933

- - -a·

- 70.081

919 - ,_ - • 997

b:~t6

7 .'913 -

-4.070

~5 .087

0.085 O

a. 641~

-3, r

85.075 8 .925 -2.0 9