Effect of Compressibility on the Pressure and Forces Acting on a Modified NACA 65,3-019 Airfoil Having A 0.20-Chord Flap

Document

ACR No. L5G3la I I NATIONAL ADVISORY COMMITTEE FOR AERO NA UTICS I ORIGINALLY ISSUED January l.946 as Advance Confidential Report L5G31a EFFECT OF COMPRESSIBILITY ON THE PRESSURE3 AND FORCES ACTna ON A MODIFIED NACA 65,3-019 AlID'0n.

HAVING A 0.20-CHORD FLAP , , By W. F. Lindsey Langley Memorial Aeronautical Laboratory Langle Field, Va.

WASHINGTON NACA WARTIME REPORTS are reprints of pa p ers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the wax 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 - 76

NACA ACR. No. L5G3la NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE CONFIDENTIAL REPORT EFFECT OF COMPRESSIBILITY ON THE PRESSURES AND FORCES ACTI JG ON A MODIFIED NACA 65,3-019 AIHFOIL HAVING A O. 20-CHORD FLAP By W. F . Lindsey SUMMARY An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modi- fied NACA 65,3-019 airfoil having a O.20-chord flap. The investigation was made for an angle-of-attack range 0 0 extending from _2 to 12 at flap deflections from 0 to -12 • Test data were obtained for Mach mrnbers from 0.28 to app r oximately 0 . 74.

The results show that the effectiveness of the trailing-edge-type c ontro l surface rapidly decreased and approached zero as the Mach nwnber increased above the critical value.

INTRODUCTION The available information on the aerodynrunic charac- teristics of airfoils, with and without flaps, at low speeds is quite extensive and previous investigations have shown the genera l effects of c ompressibility on air- foils without flaps . For airfoils with flaps, however, the available information at high speeds is li~ited.

The earlier investigations at high speeds demon- strated the li m itations of the theoretical methods in .

extrapolating low-speed data to high speeds (reference 1).

In addition, the investigations illustrated the inappli- cability of the theoretical methods in the supercritical- speed ran ge , in whi ch pressures and forces change radi- ' cally. These radical e hanges in pressures and forces are the adverse effects of co mp ressibil ity, a knowledge of

J

CONFIDENTIAL NACA ACR No. L5G3la which is necessary in the design of high - speed airplanes.

Because of the inadequacy of th e theoretical method in applications to the supercritical region, recourse to eXDeriment is necessary to determine the adverse effects of compressibility.

Data at high speeds were req uired in connection with the design of a specific airplane; accordingly, an inves - tigation was conducted in the Langley rectangular high - speed tunne l to determine the effect of c omp ressibility on the '~)l'essures and forces acting on a modi fied NACA 65,3-019 airfoil having a 0.20 - chord flap. Pressure-distribution measurements were made at ~.;ach numbers between 0.28 and 0 . 74 for angles of attack fro m _20 to 12 and fl ap deflections from 00 to -12°.

APPARATUS AND TESTS The tests were conducted in the Langley rectangular high-speed tunnel, which is an induction-type tunnel w it ho ut return passages and has an 18-inch by 4-inch test section . The variation in the Ma ch number in the test section along the tunnel axis withou t a model installed in the tunnel is ±o .4 percent of the stream Mach number.

In a plane normal to the tunnel axis the variati on is ±O . S percent of the test-section Mach number at a streruTI Mach nll..'Tlber of 0.60 . The direction of the air flow appears to be mi salined by -0.1 wi th a po s sible variation o of to.l . No correction for misalinement has been made for the data presented herein.

ThB model completely spanned the test section along the 4 - inch dimension and was supported by large circular e nd plates which were fitted int o the tunnel walls in such a ~ay as to rotate with the mode l and to retain continuity of the surface of the tunn e l walls. The juncture between the model and the end plate was sealed .

The profile of the 5 - inch-chord model having a 0.20 - chord flap differed from the modi fied NACA 65,3-019 airfoil section in that the profile from approximately the 8 1- percent station to the trailing edge wa s formed by straight lines having an included ang le between 20 and 21°. Forty pressure orifices were installed in the

model surface in two chordwise rows t inch from and on

CONFIDENTIAL NACA ACR No. L5G31a CONFIDENTIAL either side of the model center line. The model profile and pre ssure - orifice locations are shown in figure 1.

The airfoil ordinates are given in table I.

Pressure-distribution measurements were made for a

range of Mach numbers from 0.28 to approximately 0.74 at

angles of attack from _20 to 12 and flap deflections from 0 to -12 0 (up) . Additional pressure-distribution measu re ments were mad e for positive (down) deflections of 0 0

the flap at angles of attack of 2 and 4 • These tests

we re supplemented by schlieren photographs of the flow in the supercritical speed region for a few of the low er angle-of - attack configurations. These photographs show density gradients in the flow by chan ges in light int ensity. (F or details, see reference 1.)

SYMBOLS Mach number M Ma ch numb e r at which sonic velocity was obtained locally within the flow field (as at the model surface) q dynamic pressure free-stream static pressure Po p l oca l static pressure (as at model surface) a angle of attack p - Po pressure coefficient P ( q J / critical pressu re coefficient , corresponding to 0 . 52SE· - (')) ( 10 cal M ach number of 1.0 \ -0--------

\. -

H total pressure

Pu upper-balance-cha~ber pressure coefficient

PL lower - balance - chamber pressure coefficient c section normal - force coefficient n C ONFI DENTIAL

,----

CONFIDENTIAL NACA ACR No . L5G3la s e ction pi tchin g-moment coefficient of normal force about quarter-chord location flap normal- fo rce coefficient flap hinge - morne nt coefficient (determined by the pressu re distribution fro m flap h inge axis to trailing edge) flap de fl ec tion measured with reference to the mod el chord; ne g ativ e deflection is up RE SU LT S A ND DISCUSSION Tunnel - Wa ll Effe cts The results of these tests ha ve not been corrected for c onst r i cti on o r tunnel-wall e ffects. Th e mo st impo rt ant constricti o n effect on these data in the super - critical re g ion is t he ch ange i n th e Mach numbers g iven he r ein t o hi g her effective str eam Ma ch numbers. ( See refe r en ce 2 .) The difference be t wee n the t wo Ma ch numbers in cr ease s ra p i dly as the ma ximum tunnel speed is approached.

It is f urt he r shown in ref e r ence 2 that very n ea r or at t he max imum sp ee d attaina b le f o r a g i ven model -tun nel configuration lar ge g radi e nts in ve l o city oc curr ed at the wal ls. The m axi m mn-sp eed t e st poin ts g iven her e in for ea ch angu l a r co nfiguration are therefore c on sider ed to be o f questionable value.

Pr es sur e Distribution The pressure dist r ibutions alo n g the c ho rd of the mode l a re presented in f igu res 2 t o 31, inclusive. Ea ch figu r e sh ows the ef fe ct of c ompress ibil i ty on the pressu r e d is t rib u ti on for a g i ven an g ular c onfi g uration. Th e effe c t of ang l e of attack and flap def lection can be ob tained fr om a c ompa rison of the various figures.

S ubcritical region.- The figures for the lo wer ang l es of attack a n d sma ll normal - force c oefficie nts s how that i ncreases in Ma ch n umbe r in the subcri ti c a l ranges are a cco mp anied by incr ea ses in the maximum ne ga tiv e p ressure coeff i cient , w hich is in agreement wi th theory. In the high ang l e -of-attack and large n o r ma l-force-coefficient CONFIDENTIAL NACA ACR No. L5G31a CONFIDENTIAL range the change in the maximum negative pressure coef- ficient with Mach number is approximately zero and not in agreement with theory, probably as a result of the exist- ence of separated flow .

A comparison of part (a) of figures 2 to 31 for a constant angle of attack and various flap deflections shows that the increment in load produced by a deflection of the flap is distributed approximately uniformly along the chord as could be expected f'rom low-speed tests. It

can be seen, however, at this Mach number (approx. 0.43)

that the maximum relative change in loading on the main part of the wing occurs near the leading edge. This change has an appreciable effect in increasing or relieving the pressure peaks that occur near the leading edge for some angle-of-attack conditions. (See part (a) of figs. 2

to 5 and 18 to 21.)

Supercritical region.- Although it has been shown that In the subcritical Mach number range the action of the flap in changing the loading along the chord was similar to that shown by the low-speed tests, in the supercritical range, w}en the region of supersonic flow is relatively large, the loading in and ahead of the supersonic region is a function only of angle of attack.

The extent of flow affected by the flap is limited to the region of subsonic flow behind the supersonic region and to the flow over the flap itself. (See parts (d) and (e)

of figs. 6 to 9.) The chordwise influence of flap deflec-

tion on the flow over the main part of the wing could be expected to be limited to the subsonic flow region ahead of the flap, since pressures are propagated at the speed of sound. This effect of compressibility in producing a marked change in the flow over the forward portion of the airfoil at supercritical Mach numbers is comparable to that which occurs for cambered airfoils, as evidenced by the change in the angle of zero lift. The similarity can be more clearly seen when the ca> n bered airfoil is con- sidered to be a multiple flapped airfoil.

Comparison of the pressure distributions for the flap deflected and neutral shows that, for this model, the chordwise extent of the influence of the flap decreases gradually as the speed is increased above the critical speed.

If the fundamental a. spects of the flow are con- sidered, deflections of the flap could produce changes CONFIDENTIAL

6 CONFIDENTIAL

NACA ACR No. L5G31a in the pressures in the subsonic flow region, thereby changing the shoc ~ ~ location. This effect, however, is not ap p arent in these results because of separated flow ~ and the resulting absence of the usual discontinuities in the pressure-distribution diagrams which indicate the sbocl{ location .

Schlieren Photographs Schlieren photographs of the flow in the higher Mach nl~ber range are shown in fisures 32 to 37, inclusive, for the model at angles of attack of 0 and 1+0 wi th various flap deflections. (Note compressions are white on fig. 32, black on figs. 33 to 37.) These photographs 0 0

show that for angles of attack from 0 to 4 the flow

separates at approximately the 0.60-chord location for Mach numbers near the critical values, and with increasing Mach number the separation point moves forward. The forward movement of the separation point is accompanied by a rearward movement of the shock along the separation boundary.

A comparison of the schlieren photographs with the corresponding pressure - distribution diagrams shows that the existence of separated flow has a serious effect on the pressure distribution in the vicinity of compression shocks. (See, for example, figs. 9(d) and 35(d).) It will be noted that for the condition of a well-defined shock and separated flow the pressure-distribution diagram indicates smooth compression. In addition, the location of the pressure corresponding to the critical pressure

coefficient generally occurs from 5 to 10 percent of the

chord downstream from the shock location. This phenomenon is probably the result of the existence of large static pressure gradients in the separated-flow region between the boundary and the model surface where pressures were measured. The extent of the separated flow would be reduced for an airfoil having a smaller thickness-to- chord ratio.

Critical Mach Nmnber The variation of the cri tical Mach number of each surface with flap deflection for constant angles of attack is presented in figure 38 . The large decrease in critical

Mach number that occurs at angles of attack between 6

CONFIDENTIAL NACA ACR No . L5G31a CONFIDENTIAL

and 8 for the upper surface is a result of a rapid

incre ase in the magnitude of the ne gat ive pressure coef- ficients near the leading edge. The difference between the critical Ma ch numbers for the upper and lower surfaces for corresponding conditions can be attributed to air- flow misalinement. The high e st critical Mach number for this airfoil is approxirl1ately 0.65 and is obtained for angular configurations corresponding to a normal force of approximately zero .

Compressibility Effects on Force and Moment Coefficients The normal - force, moment, flap r.ormal-force, and hinge-mom ent coefficients obtained from integrations of

pressure - distribution diagrams are presented in figures 39

and 40 . Each figure shows the variation in the coeffi- cient with Mach number at a constant angle of attack for each flap deflection . These figures are cross-plotted in figure 4 1 to s~ow the variation in the coefficients at a constant ach number.

Norma l - force coefficients. - Figure 39 shows that, with

the flap at OU:-the effect of compressibility on the normal-force coeffi cient at subcri tical ' ,'iach numbers is in accord with previous experimental and theoretical results. ThG variation for the other flap deflect ions at a fixed angle of attack, however, appears to follow, to some extent, both in dire ction and magnitude the variation for the condition of the flap at 0 , as indicated by a lesser divergence of the curves than would have been exp ected from theoretical estimations. (See parts (c),

(d), and (e), fig. 39.) A variation of this type shows

dc that the effect of c01l1pressibili ty on n is small in do the subcritical Mach numb e r range.

In the supercri ti cal I\iach number range u.bove the value at which t:Je peak normal - force coefficient occurs, the conv ergence o~ the curves for the various flap deflections at a given angle of attack indicates a rapidly dC n decreasing with increasing uach number. The effect d5 dC n

of compressibility on --- , presented in figure 4 2, is

do

in acco rd with the variations indicated in figure 39.

CONFIDENTIAL 8 CONFIDENTIAL NACA ACR l, ro. L5G3la In the supercri tical ~ach rang3 for angles of attaclc not greater than 6 it can be seen in figure 39 that the peak normal - for c e coefficient for a given angular configu- ration occurred at a Mach number of approximately 0.65, which i ndicates that this airfoil section should be rest r icted to designs wherein the maximum Mach number does not exceed 0 . 65.

A comparison of the various parts of figure 41 in

which is given the variation of th e nor m al-force coeffi - cient with angle of attack shows that as the Mach number dC n increases increases and r ea ches a maximlli~ valu e at da

a Mach number of 0 . 65 as could have been expected from

the preceding discussion.

A further ext:Lvnination of :Cigu re 41 sh ow s that

deflections of the flap p roduce a change in the angle of dC n zero nor ma l force, but have no appreciable effect on da fo r the m ore lin ea r pa rts of the curves. The effect of dC n co mp ressibility on for the more lin ea r parts of drr these curves is presented in figur e 42.

da The f l ap effectiveness - obtained fr'Jm the ratio - do' dC dC n n of to (fi g . 42) and )rese nted in figure 1. 3, do d a d ecreased wi th increased Mach m. mJ.be r and ra idly approached zero as t h e Mac h number was incr eased above the cri ti cal val ue .

Mome nt coeffici e nts. - The vari a tion of the moment coeffici en t w ith ,1 1ac11 numb er in the subcri tical ran g e as shown in figure 39 is s ma ll. In the su pe rcritical region the mome nt c o efficients generally increase wi th increasing ila ch num ber. Thi s increase is 1'ollove6. by d. rapid decrease, w hich occurs at a Mach nU l!1b er above the value at whi ch the decrease in nor ma l-force coefficient occurs .

The presslre-distribution d ia grams show that the decrease in nor ma l-force results from a gene ral d ecrease in the magnitude of the loading, and at higher Ma ch numbers the continued de cr ease in loa ding is accompanied by a change in d istrib u tion wr..i ch r e sults in a de cr ea s ed moment co e f- ficient.

C OlWI DE NTIAL NACA ACR No . L5G3 1a CONFI D ENTIAL At the supereritica l Mach numbers at which the normal forces for the various flap deflections tend to be approxi - mately the same, the direction of the change in the moment coefficieuts is similar to the direction of the change in cn . A variation of similar magnitude, howeve r , could not be expected i n these coefficients because of the large effect thnt a small change in load at the rear of the model has on the moment coefficient .

A comparis o n of the effects of angle of attack and flap deflections on the variation in moment coefficient wi th normal - force coefficien t can be seen in figure 41.

For constant flap def l ection there is a small positive increase in moment coefficient with an increase in norrnal- force coefficient . This slope re~ains approximately con- stant for M ach numbe~s to approximately 0.65 . For Mach nu..l11bers above 0.65, the varia tion depends on the flap deflect i on .

At a constant angle of attack between _2 and 6 , large changes in moment coeffici en t occur in a negative direction with an increase in the normal-force coeffi- cient. These slopes , which are apprOXimately constant to a Mac h number of 0.6, increase as the Ma ch number is further increased to 0.7.

Flap normal force. - The flap normal-force coeffi-

cients are presented in fi b ure 40 . In the supercritical

region, the variations in the coeffici ents are large and irre gu lar, probably being influenced by the effect of flow separation .

Hinge-moment coefficients. - The variation in the hin ge - molnent coefficients with fl:a ch number, also shown in

figure 40, are very similar to the v ariation in flap

norr.1al force.

At Mach numbers from 0 . 01 to 0.02 below the maximuIl1 test value, a flap deflection range is indicated in which the flap tends to become or is overbalanced. At the same Mach number and for the saDe angular configuration, in figure 39, no change occurs in the value of normal-force coefficient, and the control wou ld therefore be unresisting and ineffec.tive. Althou3h the IVlb.ch number is near the maxim1..L.l11 test value, for whi ch the data are of questionable value, the possibility of this con d ition of the control should not be overlooked .

CO li[l:<'IDENT IA L 10 CONFIDENTIAL NACA ACR No. L5G3la The hinge - moment coefficients are presented in figure 41 in the same manner as the moment coefficients.

It can be seen that for a constant flap deflection the magni t~de and direction of change of hinge-moment coeffi- cient with increase in normal-force coefficient depends on the absolute value of the flap deflection. The slope for a given flap deflection increases positively with increases in Mach number to 0.675; further increases in La ch number are accompanied by increases in the slope in the negative direction.

At a Ma ch number of 0.40 the changes in hinge moment with flap deflection at a constant angle of attack are general l y uniform over the flap deflection range. At a Ma ch number of 0 . 60 and above in the ne gat ive normal - force - coefficient range the chan ges in hinge - moment coef - 0 0 ficient for flap deflections between 0 and -4 is very small compared with the changes at l a r ge r def lections.

This effect 1s possib l y a result of the reversals in the load over ~he fl1:1p p roduced by the very thick boundary layer or separated flow. (See pressure - distribution diagrruas and schlieren photographs.)

Balance - chamber nressure - coefficient differential (PL - Pu). - The diffe~ence bet men the pressure coeffi - cients for the lower-surface and the upper-surface balance

chambers is presented in figure 44. This figure shows

the effect of Mach number, .flap de.flection, and angle of attack on the pressure - coefficient differential.

The effect of compressibilit y on the pressure - coefficient differential is generally small at Mach numoers below 0 . 65 . Tllis small effect could be expected when the magnitudes of the individual pressure coeffi- cients are sma l l and are measured a t a station in rear of the position of the maximum ne gat ive pressure coeff i cient.

The decrease in the magnitude of the pressure coefficient at the high Mach numbers is primarily a result of the ef.fects of flow separation.

The balanced hinge - mome nt coefficient for this model can be obtained by adding the hinge-~oment coefficient of fi g ure L~o to the product of the balance-chamber prE;ssure -

co ef ficient differential (PL - Pu) and a constant, the

constant depending on the length of the balance tab . A

brief comparison of figur ~s 40 a nd 44 indicates that at

CONFIDEI\TTIAL

l

CONFIDENTIAL 11 NACA ACR No . L5G31a ION speeds the ef fect of the balance would be to reduce as compared with the unbalanced condition. With increasing rtach number the effect of the balance decreases for the lower angle - of-attack range .

CONCLUDING REMARK The results of the investigation on the modified NACA 65,3 - 019 airfo i l having a O. 20 - chord flap indicate that the effectiveness of a trai l ing-edge control surface of small chord ra9idly decreases and approaches zero as the Mach number increases above the critical value.

Langley Jemorial Aeronautical Laboratory National Aevisory Cerami ttee for Aeronautics Langley Field, Va .

REFERENCES 1. Stack, John, Lindsey, W. F., and Littell, Robert E.: The Compressibility Burble and the Effect of Com- pressibility on Pre3sures and Forces Acting on an Airfoil. lACA Rep . No. 646, 1938.

2. Byrne, Robert W .: Experimental Constriction Effects in High-Speed Wind - Tunnels. NACA ACR No. L4.L07a, 1944.

CONFIDENTIAL

NACA ACR No. L5G31a 12

C ONF IDE NT IAL

TABLE I. - BASIC SECTION ORD I NATES FOR ItODIF IED

NA CA 65,3 -019 AIRFO IL SECTION

[Stations and ord inates are in pe rc ent chor d] -------- - --~ I O rdinate

I

Station

Uppe r Lowe r ~

surface sur f ac e ~

I

---r-

o 0 o

.3 1.108 - 1.108

1 1. 92 1

-1 .921

2 2.59 8 -2 .598

4. 3. -!. 620

b 4 . 437 -4.437

8 5.127

- 5·127

10 5.723

- 5·723

14 6.715 - 6·715

18 7 . 525

-7· 525

22 8 .1 92 -8 .1 92

26 8 .721

-8 ·721

30 9 .113 - 9·113

34 9 . ~71 - 9. ~71

- 9.490

38 9 · 490

00 00

39 .5 9.5 - 9. 5

LL 9 · ~_71

- 9 ·471

L~6 9 .315

- 9 .3 1 5

50 9 ·024 - 9·024

54 8 .5 97 -8 .597

58 8 . 039 -8 .039

62 7.370

-7·3 70

66 6.612 - 6.612

70 ~.791

- ?79

22 2

74 . 4 .9 - 4.92

78 \ L~.029 - L~. 029

-8 .1 28

82 I 8. 1 28

86 2.247 - 2.247

90 1.ll16 - 1.416

- .688

94 I . 683

I - .132

16 g ~ __ O .13 _ __ ---'-_ o

I I , L.E. radius :

2.139 I

1-....-- ____ _ -- - ----- ---~

CO NFI DE NTIA L

'%J ,....

I--' z > o > > o :::0 z o r c.n Q (N I--' P> oq

~

lint?

.20C

ROdiVS

.035C

AERONAUTICS.

ADVISORY

~ 1<

I\ FOR

I

Stro(ghf

NATIONAL

aXIs

C~

COMMITTEE

.Ot2

c

/flnge

profile.

Model - 1.

Figure CONFIDENTIAL

CONFIDENTIAL

/ocal/o/7s

Pressure-

-orll?ce

·...---

~

NACA ACR No. L5G31a Fig. 2a-f CONFIDENTIAL x Upper sur Tace o Lower surfOct?

- .- -- f?r -/ r---~----~--~--~----~

o h>L--+-----+---+--~~'='=-

ZI /'1= 0.618.

~ ..........

. ~ ....

-/ . ~ ~ ~ III () ~ Iv J J III , / c /'1=0.675- a: /'1= 0.703.

;::, v, .., ~ d: _2 -/ f. M=0.745.

4() 20 40 60 100 o 20 60 80 100 o 60 Perce/7f chord NATIONAL ADVISORY C OMMITT£E rOIl AERONAUTICS .

CONFIDENTIAL Figure 2.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord fla p. a = -2°; 8 = 0°.

Fig, 3a-f NACA ACR No. L5G31a CONFIDENTIAL X Upper sur/act?

o Lo wt?r sur/ace

/ l----l...----;-rCl,-::-;1-l/'-1-;-:-=-o= . ...J.. 7',5;-4"7.--L-----.J /'-1=0.6/4 -

~

'1..."

~ .~ -/ .\i ~ ~ ~ C) \.l Q) / c /'1=0.675.

"'

"'l ""I

~

CC

-2 -/ f?r I?r 11 = 0. 746.

60 /00 20 40 60 80 /00

o 20 40 80 o

NATIONAL "DVISO~Y PtPrcel7l chord COHHITTH fOIl AEIIONAUTICS.

CONFIDENTIAL Figure 3.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = -2°; 8 = -4°.

NACA ACR No. L5G31a Fig. 4a-f CONFIDENTIAL x UfJper surface

o Lower surfoce

/ a /'1=0432. /'1= 0. 6/3.

~ ., '1...

c:: ~ -I \i O s::.

"

C\ \) III ~ I ~ c 1"1= a 671. ct 1"1=0. 70/.

"J .., ~

~

_2 r----,----,----,----,----, /'1=0. 74-5.

o 2.0 40 60 80 100 0 20 40 60 80 100 Percenf chord NATIONAL ADVISORY CO MMITTEE rot AERONAUTICS.

CONfiDENTIAL Figure 4.- Pressure distribution for a modified NACA 65,3 - 019 airfoil with 0.20-chord flap. a = -2°; 8 = -8°.

F ig. 5a-f NACA ACR No. L5G31a CONFIDENTIAL X Uppe.r su r/oc t?

o L oweI' surrace ... "" -- -- f?

c r Q.

...

c::

-

<lJ "- - I . <J ~ , Ql (J Q) ~ / ;::, '-, '-, <lJ

ct

-2- -'0- - / ....

, -""'--0- .

e f? r

I ~ ____ L-,-~ ~-=~~ __ ~ __ __ e /'1 -= Q731. y 1'1~0.74S.

o 20 / 00 o 20 60 100

40 60 80 4 0 80 NATIONAL ADVISORY Perc e nf ch or d COlflffTT££ fOl AERO N AUTICS.

CONFIDENTIAL Figure 5 . - Pressure distribution f or a modi f ied NACA 65,3-0 1 9 air f 0 i 1 wi thO. 20- c h 0 r d f 1 a p. a = - 2 0; 8 = - 12 • -_. --------- NACA ACR No. L5G31a Fig. 6a-f CONFIDENTIAL X U ,P,Per s ur/ace

o Low6'r surroc6'

- / r------.-- ....... ----,---,------, (oj /'1-=0. 646.

Q '\-...

c: -/ .~ . lJ s::: ~ Ql 0.

() \; / Ql (e) /'1=0.672. (d) /'1=0.70.6.

~ ::::, ~ \U

ct

-2 ~-Q.., \ ,

//

~-~ ....-- - /

/ "r-- ~

-/?

-- - ~ ~I?r r - ~

~ /

f r NATlON~L ADVISORY

COMMITTEE rot AERONAUTICS .

I I / c = (e) /'1 0. 72 6 . (0 /'"1= 0. .743 .

o 20 40 60 80 / 00 0. 20. 60 80. /0.0.

Percenf chord CONfiDENTIA L Figure 6.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0 .2 0 -c hord flap. a = 0°; S = 0° .

---- _._-_.- ._- -- -- . -- ------- -- -- - --- ------------------ ------ ------~------~~~-- .-------------------------------------- ~- --- ~ Fig. 7a-f NACA ACR No. L5G31a CONFIDENTIAL x UjJjJer surloa> o Lower ..surhce b /"1=0650.

Q

" "i-- c:: . Cll "- .0 -/ s::

-I?r

~ III a ~ Cll '- ~ / (c) /'1= 0.677. (d) /'1= 070.9.

~ ~ q - - ~ -/

--

~-e ~ 'v...

-~ r ~

/

o / - (g) /"1 0. 733. /'1=0.746.

20 60 80 100 2.0 40 60 100

o 40 o 80

NATIONAL ADVISORY P()rcC?nl chord COMMITTEE FOR AEAONAUTICS.

CONFIDENTIAL Figure 7.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0 .20-chord flap. a z: 0°; 0 ,. -4°.

Fig. 8a -f NACA ACR No. L5G31a CONFIDENTIAL x Upper S Ulface o Lo we r surr ace /~--~- ~=V) ~H~- - - O=.~4~~~7.=,~--~ 1'1 =0. 6.2 2.

Q

.~

-/ .\.j ~ ~

~

\l ~ / .:::i '1 'I ~ Q( -2 r----.----r----.----r---~ '<l.

'0.

"'" - .....

r-- . '<>-

~

./ "',

V

- f- -

/

o f-§f---+----f---f---+---j 'f (I) H - 0.744.

/00

20 40 60 80 tOo o l?O 40 60

o

NATlON"L ADVISORY Percel7t chord COMMITTEE rOi AERONAUTICS CONFIDENTIAL Figure 8.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 0°; 8 = _8°.

NACA ACR No. L5G31a Fig. 9a-f CONFIDENTIAL X Upper surroce o Lower surroce r--~~--.--~----,- ~ r /'1=0.648 .

/1------'-""77",-J-,c-:--=~;.,.--'----l 1'1 =0 . 70..5.

-2,----,----,----,----,----, - ~ r /1 =0.743.

/(}()

;:0 o J?O 40 60 1 00

o 40 60 80 80

Percell/" chord NATIONAL ADV I SORY COHHITTEE FOI AERONAUTICS .

CONFIDENTIAL Figure 9.- Pressure distribution for a modified NACA 65,3-019 airfoil wi;th 0.20-chord flap. a = 0°; 8 :: -12°.

----------- -- ---- - NACA ACR No. L5G31a Fig. 10a-f CONf"1DEN TIAL X Upper surroce o L .Jwer surhce -I r---~----~----.----'----~ / 1-- -----' (0 /'1=0.436.

(b) /'1=0643.

d /'1= 0.705.

o ~---+-- --+---~ -----r --~

M=o.744.

40 l Oa 40

a 20 60 80 o 20 60 80 100

NATlONA.l ADVIS ORY CO MM ITTE[ rot A£AONAUTICS Percenf chord CONfiDENTIAL Figure 10.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 2°; 8 = 0°.

NACA ACR No. L5G31a Fig. lla-f CONFIDENTIAL x Upper surfoce

o Lower surface

-I r----.----,-----,----.----~ OrY--r---r------r-~~~ 1~-~~(a~)~H~=-O=.4~3~/~-~ ',b M=o.647

Q

"' ~

.QJ -I "' .I..J :-.:::.

"

U I Ql (c) M=0675 (d) /'1=0710 I....

~ ~ Ql

ct

-,2 -I cr o ~~-+----~----+---~----~ 1'1=0742 20 / 00

o 40 60 80 100 o 20 40 60 80

NIo,TlONAl ADVISORY Percenl chord ' OHHITTH fOR AUONAUTICS .

CONfiDENTIAL Figure 11.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 2°; 8 = _4°.

Fig. 12a-[ NACA ACR No. L5G31a CONfiDENTIAL x U;:;)Jf?1' .c;urfoce

o LoweI' surface

- / r-----,------.---,..-----,---, o~-_+--~--+_-=~~~~ (0) /"7=0.644.

Q ......

<:: Qj - / ....

.'V -~r ~ ~ III I;) ~ / III (c) f1=0675. (d) /"7=0.707.

"

~ ~ ~ -2

-

-/ ~;:o. .....

~

-Q~

;;/

r

/ (e) M 0.734. f) /"7=0.741.

o 20 40 60 80 100 o 40 60 80 /00

NATIONlL ADVISORY Pel'cel7f chord COMMITTEE rot AERONAUTICS CONFIDENTIAL Figure 12.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 2°; 8 = -8°.

~l

NACA ACR No. L5G31a Fig. 13a-f CONFIDENTIAL X Uppe ,- surfact?

o LOWf?r surfoce

Ik----'--,..,--'-:-~=_'-;-:;--:;-----'-------' a 1'1=0. 433. b I'1=Q648.

Q ~ ~ c:: (lJ ~/ "- ,v s::: , (lJ C) ~ / Q.l 'd /'1=0. 707.

c 1'1=0.677.

"- ~ ::s Q.l Q: _2 ___ ::::t;:=--<-~ ' "0--0--0- \ -/ cr /'1=0.742. .

o 20 40 100 o 2.0 40 60 /00

NATIONAL ADVISORY Percenf chord 00""1"£1 rot AERONAUTICS CONFIDENTIAL Figure 13.- Pres ' sure distribution for a modified NACA 65,3-019 airfoil with O.20-chord flap. a = 2 ; 8 = -12°.

j

l

Fig. 14a-f NACA ACR No. L5G31a CONI'IDENTIAL x Upper sur/oct?

o Lo wer sur/act?

- -I?,- (6) /'1=0.614.

Q i-.."\ c:: (]) -I "- \j ~ \...:: Q) ~ '-l Q) (c) /1.=0680. (d) 1'1 =0.704.

'- ~ "l "l ~ 0:: _2 -I cr ff) 11 =0.742_

o 40 60 (30 100 o 20 40 60 60 100

NATlONA.l ADVISORY Percenf c /~ord COHNITTU rot AERONAUTICS.

CONFIDENTIAL Figu re 14 . - Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20~chord flap. a '" 4°; 8 '" 0°.

L

NACA ACR No. L5G31a Fig. 15a-f CONFIDENTIAL X Upper surroce

o Lower surrocC'

(LJ) 1'1=0.644.

Q

, --f?

cr CIJ ~ ,l!-----L.------;.---:-...."...,~~=;__-'-------" ~ (c) M=O . o73 .

II>

ct

-2 r---.--,---,---,--, cr

o ~~-+----~----+---~----~

(f') 1"1 = 0.739.

40 60 80 /00 20 40 80 /00 o 20 o 60 NATIONAL ADVISORY Percent chord COMHllTE[ fot AERONAUTICS CONFIDENTIA L Figure 15.- Pressure distribution for a modified NACA 65,3-019 a.irfoil wit h 0.20-chord flap. a = 4°; 0 = -4°.

Fig. 16a-f NACA ACR No. L5G31a CONFIDENT I AL X Uppl?r sur,roce

o Lower surfacl?

1'1=0.648.

;.....

c: -9:> u <.;:: - I '---:::;""",",~----::;<~~:-r--r=-----r , Q) C> U Oi-----1'f'--+----+--+--="'<+....::::,."r---l (d) 1'1=0.706.

-2 ,----r----r--,- ---.- - - I -I I ®- --L---'-- e - ~ -L1'1.,..-,:-=-=0::-:]:!o-3 =- S. -= . ----.l-------' (f) 1'1=0.741.

r

2.0 40 60 80 100 o 20 40 60 80 100

o

NATIONAL ADVISORY Perce nf chard C0JI4HITTH fOR AERONAUTICS .

CONFIDENTIAL Figure 16. - Pressure distribution f o r a modified NACA 6 5,3-019 airfoil with O.20-chord flap. a = 4°; 8 = -8 • Fig. 17a-f NACA ACR No. L5G31a CONFIDENT I AL x . Upper sur Tace o Lo wer s ur;'oce - / r__----.----,----r-----,--___,

o ~--+--~--+--

(b) 1'1=0650 • ..

'h.

c:: . 9?

.v ~-/r------.-7"=7........,---¥r-----,--=----, "- CD o v 0~~-+--~--+-~~~~r1 CD ....

~ / t-----"----:r-c-:-} -'-:1'1;::;:;--=--;;O".6:;!:-:;79=-. -"------' (d) 1'1=0.706.

ct

-I

I1r cr

a ~r--r--r---r--+---i

r) /'1=0.740.

LO 100

o 20 40 60 /00 o 40 60

NATIONAl ADVISORY COMMlnu. fot AERONAlITtcS.

Percent chord CONFIDENTIAL Figure 17. - Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 4°; 8 = -12°.

Fig. ISa-f NACA ACR No. L5G31a r,ONFIDENTIAL x Upper s ur fa ce o L o wer surface

o I:----=;;jf---+----+-~i==o-~::i

(b) 1'1=0 . 586.

Q

'i-..

~ Q) .....

. (,) -/ ..;:: Q) ()

"

l> Ql

'"

~ / ~ (l) d: _2 r----r----,----r----,----, /~-~-(~e~)L~~=O= . ~72~4~.-L--~ CONFIDENTIAL (50

40 80 /oo o 20 40 60 80 /00

o 20

Percenf chord Fi gure IS. - Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 6°; :3 = 0°.

Fig. 19a-f NACA ACR No. L5G31a CONfiDENTIAL X Upper surface o Lower surfoce ',,- - -- p.

-/ " " c r-.----, " o / I'1=Q61.9.

Q ~ ...

c: Q) -/ "' .U ~ -..:: Q) (J <b i J....

c M=0.076. M=0.70..3.

;::, ~ III J....

Q -2 -/ -~r 60 /00 20 60 80 100

o cO 40 80 o 40

NATIONAL ADVISORY P ercenf chord COMMITTEE rOl AERONAUTICS CONfiDENTIAL Figure 19.- Pressure distribution for a modified NACA 65,3-019 airfoil with O.20-chord flap. a = 6°; 8 = -4°.

Fig. 20a-f NACA ACR No. L5G31a CONFIDENTIAL X Upper sur foct?

o Lower surlOce

-. - -!?r-,...----, (b) M=O.6~3.

Q

~~ c:: . ~ -/ . ~ , .....:: <IJ () u It) / '- (e) /'1=068.2. (d) 1'1 = .714.

~ ~ <IJ d: -z -/ cr cr o ~~-r----r----r----r---~ 1'1=0. 738.

o 20 40 60 80 /00 o 20 40 60 80 /00

NATlONA.L ADVISORY Percent chord CO,",",TTEE Fot AERONAUTICS .

CONFIDENTIAL Fi gure 20.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord nap. a = 6°; 8 = _8°.

NACA ACR No. L5G31a Fig. 21a-f CONFIDENTiAL x UppG'r sur fac e o Lower surface ' I?r----r----, Or-~-+----~----+_~~--~~ I ~---'----(,"--o--:-) ...L/"1;-::-=--;:0,.-.4':-:3:::-c :::-c.--'------' 3 (b) 1'1=0.651.

Q (d) /'1=0.704.

-2 ,....----y---y----r--y-----, -~r

o r-~-+----~----+_--~----~

1lF-----'-------'-::-=---=--=-=--=-----''-------' (e) N=o.727. (I") 1'1=0734.

20 40 60 80 /00 o 20 40 60 80 100

o

NATlOH\l ADVISORY COM Hl n[£ fot AUOHA.UTICS Percen f chord CONFIDENTIAL Figure 21,- Pressure distribution for a modified NACA 6 5,3-019 airfoil with 0.20-chord flap. a = 6°; S = -12°, F i g : 22a-f N AC A A CR No . L5G 31a CONFIDE NTIA L I?r----~I----~I----~ x U pper sur fac e

Ir-- "'_

-2 f-+----l-----+-- 0 L ow er surfac e \ I-- ~r \ .:::: -/ r----+--~ ~ -- --~ -- _r--~

~

. ~ ~ ~

1 /'

/ IJI"----'-----;- (o -:- ) ...J.../"1=-=-=---;;0==- .4 3=-6=. ----I -------I (b J /"1 =O S 8 CJ.

< -2

c:: . ~ u ~ - / , Q) o u 0 ~ -- ~ ----+----+---- ~--~ (d) i'1=O()T4.

-2 r---~---- ~ --~----~--~ "' >.. >a.-- ....

's..

- / 1----+-----'+-:",....,,="i'<::-----If---- -- ---1 - -- "Zr.

o r- --~----~----+-----If---------1

/ 1*- ----L.....- { -;- e -:- ) ....L. H -:-:- =-:: O:-:.7 ~ O::-:3=-. --'-----" f 1'1= 0. 720.

IDa o 100

o 40 60 80 20 40 60 80

NATIONAL ADVISORY Perc e nl chord CC!'!!1 ! TTH . fOR ArRONAUTICS.

CONFIDE NTI AL Figur e 2 2. - Pressur e d i stribut i o n fo r a modi fi ed NA CA 65,3-019 a irfoil wit h 0.20-c h ord flap. a = 10° ; 1) = 0°.

~----~--~-- ----------- .--- -------- -------------. --- --------------------------.----------~ Fig. 23a-f NACA ACR No. L5G31a 'CONfiDE NTI AL -3 cr I~ ace f ...

U pp~r surJc X

... , I I 'I Lower su r r oc e 0 - -2 , ,

~

I-- I1r

'I" -; ~ - I

~

. ~

~

~ 6....

~ /

/

V

/ (b) 11= 0. .5(3 6. (0) M = 0. 4..54.

~ -i?

\ \ "' i-..

c:: '-

"

. ~ - I .~ ~ ~ IU Iv IU / (d) 1'1:: 0.. 6 (31.

~ "1 "1 I\)

~

-2 -I 0 1---,"I"'f----t----i----t- - ---1 / ~_----L-----;- ( e - ) r-1--: /'1 ~ = -::C',-l, . 7 = 0= 7=.----1. - --' ) 11 = 0. . 719.

20. (]O lOa 20 40. 60. 80 100

o 40. 60. a

P~ r c e rd chord NATIONAL AOVISORY COHHJT7lf fOf .l£IKHUUTICS.

CONFIDENTIAL Fi g ure 23.- Pressure distribution for a mo d i f ied N ACA 65,3 - 0 19 airfoil wit h 0 .20-c h ord flap. a = IOU; 0 = - 4 P .

NACA ACR No. L5G31a Fig. 24a-f CONFIDENTIAL - 3

. I

X Upper sur/(.

r Lo wer sur I-+---+----!?r ---+-----1----1 - 2

~

t---...

~

~ "'-....

' ...", "e.-.,

/

V

/ (a) /'1=0. 433.

-2 ..

'10 '+-- ...

\ , c: Ib "- -/

-

"

. I.l :i:: '.;::: Ib ~ I.l ~ / :;) c 1'1=0 . 622. 1'1=0.679.

~ ~ Q - z r----,-----r----~----._--_, h-o- -<>

,..----

-- '; / V - I--- I?r.

V

.11

(f) /'1 - 0. 7 2. 0 .

60 dO /00 o 20 40 60 80 /X

o 2.0 40 Percenf chord NATIONAL ADVISORY COHMITIEE fOIl AERONAUTIC S.

CONFiDENTIAL Figure 24.- Pressure distribution for a modified NACA 65,3-019 airfoil wit h O.20-~ h ord flap. a = 10°; 8 = -8°.

Fig. 25a-f NACA ACR No. L5G31a CONFIDENTIAL x Up?er surroct?

-2 A---~--~- \ o Lower surface \ -/ f------t=~~+--_t__--+_-__l / (a) 1'-1= 0.439. (b) /'1=0557.

ct

-2 .....

\ " -;....

\ , .~ -/ .....

.v ~ ~ ~ Cl V / ~ (c) /'1=0.622. (d) /'1= 0.6.96.

'- ~ ~ ~ Q -2 Ir-"" ~ ~-o--o..

, / -/ V "0..-0.

--

--~ -...,..

V

o

[7

/ - (e) 1"1- 0.7/5. r) /'1=0.72/- 2.0 40 60 (}O /00 o 2.0 40 60 80 /00

o

NATIONAL ADVISORY Percel7f cho rd COMNITI([ fOi AERONAUTICS .

CONFIDENT! AL Figure 25. - P r essu r e dist r ibution for a modified NACA 65,3-019 airfoil with O. 20 - chord flap. a = 10°; S = -12°.

NACA ACR No. L5G31a Fig. 26a-f CONFIDENT I AL x Llj?j?er surface

o L ower surf ace

-- --- --/?

-/ I (0) 1'1 == 04 3/ 1'1 =0.620 .

O lc? --r---r---r- ~~~ (dJ fV! = 0.702.

-2 r---~-----r----,, ----.--- ~ - / 1-- -- --I---:;-==>'""f--- ~~ ---- -t---____1 -~r Olhf "-----+--t---+--t----' =-j r. 1'1 =0 . 745.

/00 lro

o cO 40 60 80 o 20 40 60 eo

NAT rONAL ADV ISORY Pr;>rC2 n t ch ord COM M ITTEE rOll AERON A UTICS .

CONFI DEN TIAL Figu r e 26.- Pressure di s tributio n f o r a mo di fi ed NACA 6 5, 3-0 1 9 a irfoil with O.20-c h ord fla p . a = 2° ; 8 = 4 ° .

Fig. 27a-f NACA ACR No. L5G31a CONFIDENTIAL x U/Jper sur loa?

o Lower surfoce I 11-----'---(-=0::0) --'-c/'1 '-:;-=""""" O ", 4 7- =;"'4 -'-- . ---'---' 3 (b) /'1:06/8.

Q (d) 1'1=0703.

(f) /'1-=0744.

o 20 4 0 60 80 • 100 o 20 40 80 100

NATIONAL ADVISORY ;:::Jer cenf chord cOJIounu fOIl AERONAUTICS.

CONF I DEN TI AL Figure 27.- Pressure distributio n for a m odified NACA 65 ,3-019 airfoil wit h 0.20-chord flap. a = 2°; 8 = 8° .

--- - - _._- -- --- -

Fig. 28a-f NACA ACR No. L5G31a CONfiDENTIAL x Upper surroct?

o Lower surroce

--f?

''"''

Q ~"\ c::: .......

, , . 10 "- - / .\) ~ ~ . ~ \\) ~ v ~ / ~ ~ V:> ~

ct

/'1=0745 .

80 /00 20 60 80 /00

o 2.0 40 60 o 40

NATlONll AOVISORY Pe rc enT chord CO MMITTEE fOR A[IK)JrUUTlCS .

CONFIDENTIAL Figure 28.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 2°; b = 12°.

Fig. 29a-f NACA ACR No. L5G31a CONF I DENTIAL X Up"oer surrocl?

o Lower surroce

--/?t-

Q

-2 '- \ ~ \ . ~ -/

..... -'?r

. \.)

~ ~ ~ 1::1 \.)

/ It.

c. 1'-1= 0647.

~ "?

ct

-2 -/ 1'-1 = 0. 74/.

(JO o /00

o 20 40 60 100 cO 40 60 80

NATIONU ADVISORY Perc{1/l1 c/;ord CO"Hlnn fOR AUON.lUTICS CONFIDENTIAL Figure 29.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap. a = 4°; b = 4°.

NACA ACR No. L5G31a Fig. 30a-f CONFIDENTIAL X Upper sur race -2 o Low er surhce " .....'""---., ____ -- .... - - rJ ' --~r -/ ~---r----.-~~----~--~ 1 P---...L-----c a;->j,-J/'1-:-:-=-O-=- . .J....,4.-::3;-4:;- • ....I-_---J 1'1= 0 . 5.90.

r

ct

-2 ,~ \

"

~ \ . QJ -/ .... ~--~-----+----~==~~;;'r . ~ . ~ \b <;) «-J IV / ~ H= 0.645". d. 1'1=0. 705.

~ c., c., ~ q: -2 -/

- \

~- ~ ~ .,..-.._---l ,., -rc,.

1'1=0- 742.

cO 40 80 /00

o 20 40 80 100 o 60

NATIONAL ADVISORY Perce/')! chord COMHITI[[ fOl A[RONAUTICS.

CONFIDENTIAL Figure 30.- Pressure distribution for a modified NACA 65,3-019 airfoil wit h 0.20-chord flap. a = 4°; 8 = So.

Fig. 31a-f NACA ACR No. L5G31a CONFIDENTIAL X Upper surt'ace o Lower su r hce

, -- --/?r

-I Q -2 ...

--

"

.~ "- -I

. 1,; I?r

" ~ ~ ~ ~ '0 ~ \..

I (c) 1'1~ 06 4J.

~ ") "> ~

ct

'"W' __ --0..

" - -{;.

Ol/---;r--+--+-----j-----"~:;c:;;f___i NATIONAl. ADVISORY COMMITTEE fot A[IONAUTIC$.

(f) H =0.741.

cO 40 60 80 100 o cO 100

o 10 60 80 Perceot c !lord CONFIDENTIAL Figure 31.- Pressure distribution for a modified NACA 65,3-019 airfoil with 0.20-chord flap.

a = 4°; 8 = 12°, z :x:- C) :x:- :x:- C) ::0 z o (]l Q CN t-J III I r '%J ~. CN ro III Cb OQ H.

airfoil P'IBLD.

0.678.

0.741.

LUOLBT M M • ABRON1UTIC8 65,3-019 (b) FOR (e) L.\BOR.UORY NACA 00.

COMMITTBB ..

AERONAUTICAL ADVISORY 0°; modified MEMORIAL = NATIONAL a a 7. LUOUY for O.?

flap.

• flow M the c) ( CONFIDENTIAL of CONFIDENTIAL 0.20-chord with photographs section 27~ Schlieren 0.644.

O. 7 • • M M 32.- d) (a) ( -.

~~ Figure -- -

I ~

, > n > > n ::0 z o r CJ1 Q c.N i-' (ll "2:J t-'. ()ol ()ol (ll ro z OQ VA.

airfoil FIELD, 744.

0.677.

= O.

= LAWOLBY - AIRONAUTIC8 b l M ( 65,3-019 elM FOR ( LABORATORY _4°.

NACA COMMITTIB c & AEROIAUTICAL ADVISORY 0°; modified MIMORIAL a NATIONAL a a LAWOLIY for flap.

•

flow 0.706.

=

M the CONFIDENTIAL (cl of CONFIDENTIAL 0.20-chord with photographs section Schlieren 651.

0 . 0.732.

= =

M 33.- M (al (dl Figure I Q P> 'XJ t--'. P> ~ z » () » » () ::0 z o r CJ1 tN I-' tN .t>- aq V4.

airfoil FIILD, 0.650.

0.742.

=

L4JGLI' = - M 41ROR4VTIC8 M Foa 65,3-019 (cl (f) L4BOalTORY COKMITTII _8°.

NACA = 4IRO"~VTIC4L ~DVI80R' 0°; MIMORI4L modified K~TIOK~L = a a LAKotlY for flap.

0.621.

0.735.

flow

= =

M M the (bl (el CONFIDENTIAL of CONFIDENTIAL 0.20-chord with photographs section Schlieren 0.708.

0.589.

=

M M 34.- (dl (al Figure .J z > o > > o ::0 z o (J1 Q ()J (ll L' t-' I 'X) 1-" ()J (J1 (ll H) aq VA.

airfoil FIILD, 743 .

0.676.

= LA.OLIY

=

- M M AIROWAUTICS FOR 65,3-019 (cl (f) LABORATORY NACA COWIITTII -120.

s

AIRO.AUTICAL ADVI80RY IORIAL 0°; ..

modified WATIOWAL = a a LAWOLlY for • flap.

0.648.

flow 0.735 :-:-

=

M M the (bl CONFIDENTIAL CONFIDENTIAL (el of 0.20-chord with photographs section 620.

Schlieren 0.700- O.

=

= M M 35.- l a (dl ( Figure ..

I z :> 0- :> :> () ::0 z o L' c..n Q (Jo:I ...... III "'J ....... (Jo:I CD (j) III OQ V~.

airfoil 0. 0.738.

FIELD.

=

M M L~KOLET ) - (b (e) ~BROK~UTIC8 65,3-019 FOR .L~80alTORT NACA _4°.

COMMITTEE c S lEROK~UTIC~L ~DVI80RT 4°; modified MIMORI1L c K~TIOK1L a a L~KOLBT for flap.

0.704.

flow = M CONFIDENTIAL the c) ( of CONFIDENTIAL 0.20-chord with photographs section 0.727.

Schlieren 0.64.4.

=

=- M M 36.- (d) (a) Figure ~ , >xJ t-'" ().l -..J Pl t-'l (J1 Q ().l f-' Pl z :x:- C) :x:- :x:- C) ~ z o r" OQ VA.

FIILD, airfoil 0.678.

0.741.

=

= M M 65,3-019 (cl (fl NACA 0°.

• &~=~~~!~T~~=~Ii!:~~~:R~&~O~!~~~~~ S 4°; J:~~~~~~L modified a a· LUOUT for I AL flap. NT 0.727.

flow 0.651.

=

F IDE M the (e l M CON CONFIDENTIAL (bl of 0.20-chord with photographs section Schlieren 0.617.

0.703.

=

M M 37.- (al (dl Figure f--. CN CXl III 0' "2;l aq z o r CJ1 Q CN ....... III z ;I> (') ;I> ;I> (') ::0

/6

deflection .

(3 flap AERONAUTICS surface of FOA

o

NATIONAL ADVISORY CO!4!4ITTEE

lower

.......-=

=I~ function

(bl

:::> a

I

-8 as /1 >

deg

attack.

o

4- /l

-z

rf)

6, of surface

-/6

/6 angle each CONFIDENTIAL CONFIDENTIAL for deflection,

race

constant a

sur

number at r/ap

o

Mach tipper (a) Critical

-8

~ 6 8

a

~ I;!.

38.- (deg) I

I I

S

.7r~

.6 .4

-/6

.6

~

ft \)

~ ~ ~ ~ Figure

~ .\) ~

~\J ~ ..c:: "- " ~

~-- ------------------------ ~---~------~

,-

Fig. 39a NACA ACR No. L5G31a

o

I CO N FIDENTIAL

~

/ c5 (de.J)

II

h(cr

D .2 I .Q --0- ~

J j

t~

:; -: ..

1DJ

I

I ~ ~

I ~-.4 I -4 -=l=....

r--+-

I

r ~ 1

Ai" -.

I

Nj, -8 }(- r-----~ ~

~1

II

'>f...

..n.

-/2

V! I

D- ~ "1

.~

j

-; I G.J NATION AL ADVISORY COMMITTEE FOR AERONAUTICS .

./ ~/ i:: -8 c: 177% -4 ID CONFIDErTIAL -./O~--~ . /~--~ . 2~--~.3~---.~4'---- .~5~---.~6-----.*7----~.8 Mach ntlm./;ffr, M Figu re 39.- Effect of compressibility on section normal-force and moment coefficients.

Fig. 3gb NACA ACR No. L5G31a

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