IN THE

REPORT No. 336

TESTS OF LARGE AIRFOILS

PROPELLER RESEARCH TUNNEL, INCLUDING

.—

TWO WITH CORRUGATED SURFACES

By I)ONALD H. WOOD

Langley Memorial Aeronautical Laboratory

.--—

m4397-+~9 761

REPORT No. 336

TESTS OF LARGE AIRFOILS IN THE PROPELLER RESEARCH TUNNEL,

INCLUDING TWO WITH CORRUGATED SURFACES

By DONALD H. TVOOD

SUMMA13Y

This report gires the resuh of the tests of 8eren 1?lIy Ifi foot airfoils (Olark Y, smooth and corrugated, QWingen 398, N. A. (7.A. .W8, and N. A. C. A. 84). The testgtceremade in the Propeller

Research Tunnd of ihe National Aduiaory (70mnaitteejor Aeronautics at R8ynald8 Numb-s up

to i!,OOOIOOO.Tiii Clark Y airfo27waa testedw“ththree degreesof eurface smoothness.

l%?efect of small variations of 8moothnessof an airfo27is slwwn to te negligible. Corrugati~

the surface causes a $atteni~ of the lift curoe at the burble point and an in.creoxein drag at small

$ying angles.

INTRODUCTION

At the annual conference of the h’ational Advisory Committee for Aeronautics with aircraft

manufacturers held at Langley Field, Vs., in May, 1928, COLV. E. Clark and others mentioned

the lack of test data on corrugated wings and suggested that tests be made in the Committee’s

Propeller Research Tunnel. Here a comparatively high Reynolds number may be secured

due to the large size of the models that can be used. It also seemed deairabIe to secure data on

sdme representative wing sections with a view to the possible comparison with existing data from

..—-.

other tunnels.

In the Propeller Research Tunnel with its 20-foot dimuetar throat, airfoils of 2-foot chord

and 12-foot span may be tested up to velocities of 100 M. P. H. This oondition gives a Reynolds

Number of about 2,000,000, which corresponds quite closely with that attained in the Variable

Density Tunnel at 10 atmospheres pressure.

Four airfoils (Clark Y, GWingen 398, N. A. C. A. M-o, and N. A. C. A. 84) were selected

for the present teats. The CIark Y was tested with three degrees of surface smoothness. In

.—

addition, two corrugated metal covered Clark Y airfoik, one having CIark Y section at the top

of the corrugations and the other Clark Y section under the metal coverhig, were tested.

Thus, eight separate tests were made at speeds of approximately 80 and 100 M. P. H. The

average ReynoIds hTumberswere 1,575,000 and 1,940,000, respectively.

METHODS AND APPARATUS

SUPPORTS

The Propeller Research TunneI, where this investigation was conducted, has been described

in Reference 1. The regular tunnel equipment was employed so far se possibIe. Referring

to Figure 1, the sirfofl to be tested is supported on two heavy, braced bars and fitted to pivot

about a point within the airfoil slightIy above the chord at the quarter point. A ‘(sting”

attached to the center of the airfoil is carried back to a vertical tube to which it is pivoted. A

rack and pinion operated by a crank serves to raise and 10wer this tube, thereby changing the

angle of attack of the airfoil. These members are boIted to the floating frame of the balance.

The Iift and drag forces may then be read on the pIatform soaks on the floor below.

To reduce the tare drag of the system all supporting members were surrounded with faking

attached to the fixed frame of the balance. To reduce interference with the airfoil the fairings

were not carried up to the wing, the last 2 feet of the supports being streamlined instead. The

effectiveness of this arrangement is indicated by the fact that the tare drag was only 3 pounds at

100 M. P. H. at most angl~ of attack. This was about 50 per cent of the gross mirimum dr~. -r-—— 763
.

764 REPORT NATIONAL ADVISORY COMMI’lTEEFOR AERONAUTICS ——. —

In measuring this tare drag the set-up was mowed so that--the wing was supported independently of the supports and sting; hence, the drag me~sured was that of the supports alone

in the presence of the airfoil. This was accomplished by supporting the sting from a tube

connecting the front supports within the wing, but not touching it. The wing was then sup--

ported by wires and pipes arranged at distances from the supports. This arrangement is,shown

in Figure 2. The angle of attack couId then’be easily changed by simply moving the wing and

turning the regular crank to bring the sting parallel underneath. Readings on all the b~ancg+ .= ~.. ___

were taken at several angles.and air velocities so that the proper corrections could be made to

the lift and drag readinga. !A small correction to therb~~ce read”ti~ was also ne=~, due

to the different distribution of the weight at the several angke of attack. ~

CONSTRUCTION OF AIRFOIIA

Since the airfoik were to be of 12-foot span and 2-foot chord, the standard ordinates of the ... .

airfoil sections in per cent of the chord were reduced to inches on a 2-foot chord. The.model

FIWJEE1.–Arrangementforwingtests

maker was gi~en these ordinatea to the nearest hundredth of an inch and was aeked to work

within xOOor % inch. Tables I–VI give the standard, specified, and measured ordinates. The

measured values are the average of three measurements at the center span and h~fway from

the center to each tip. It will be noted that there me differences of as much as XOOinch from

the specified ordinates. They occur at- the leading edge wh~e the surface is well rounded,

There is a considerably larger deviation for the thick corrugated airfoil which is accounted for

by the difEculty of construction.

The leading and trailing edges were of lam@ated_wood glued and formed to tamplates. At

about the mid-chord a 3-inch wide beam was placed. These three mexnbe~ were spaced by

solid ribs at 12-inch intervals along the span. The space between leading and trailing,edges

on the tip and bottom surfaces was originally covered with xc-inch 3-ply plywood. After the

tit airfoil was completed examination showed considerable bowing and buckling of this thin

covering. It was decided, however, before discarding this construction, to make a test, thereby
TEISTS OF I.AR(XO AIRFOILS IN TEE PROPELLER RESEMCH ‘IWNN_EL 765

determining the effect of these small variations of su@ace contour. The plywood was then

removed and )&inch sheet aluminum substituted and a test made. The whole airfoil was then

painted with two coats of brusbing lacquer, sanding between coats. This gave a uniform smooth

surface, although not as smooth as the bright sheet metal. All screw holes and cracks were

filkd with Lithargeand glycerin before painting.

—-

In Figure 3 are views of some of the airfoils.

The corrugated airfoils, one of which iailhstrated at be bottom of Figure 3, were constructed

in the same manner as the plywood airfoiI, the corrugat~d metal covering being screw@ to its

surface. The metal sheet was of jf~-inch thick s,Iuminum with the corrugations rolled on a

grooved wood form. The dimensions of these corrugations (fig. 4) were found by scaling down

the average of several standard wings. Simw it is impractical to run the corrugations cornpletdy around the leading edge, the sheet was left flat there, the corrugations starting a slight

..—

distance back on the top _andbottom surfaces. In order to bring the corrugations into a scalloped edge at the rear, they were displaced one-half pitch on the top and bottom surfaces. This

Fmcrm 2.-Arrangement fm taredragtest

and the leading edge construction necessitated slight departur= from the basic Clark Y section.

These are indicated in F~ure 4.

TESTS

After mounting the airfoil a cover plate was screwed over the pivokfitting opening, Ieaving

only enough gap to allow the support to cIear as the angle of attack was changed. Angles of

attack indicated by a pointer on the moving rear support were checked against an inclinometer

held on top of the sting just behind the airfoil.

Each test was run attunnel air speeds of approximately 80 and 100 M. P. H. Air speeds

were computed from the readinga of a manometer connected to plates set in the walls of the

tunnel passageacalibrated against Pitot tubes suspended in the air stream at the position of the

airfoil. Two readings were taken of front lift, rear lift, drag, and manometer at each angle of

attack at each speed, one when the angk.s were succesaively increased from —9° to + 35°, and

the second when decreased from + 35° to – 9°. This was done simply to secure two independent

readings at each setting.

..-

.—

——. -

.-
766 REPORT NATIONAL -ADVISORY COMMIiTEE FOR AERONAUTICS

RESULTS

The results are..given in the form of tables and curves of the absolute nondimensional

co&cient9 CL, ~DI ~MI (& From the observed readings these coefficients are computed in

..—

the usual manner from the equations

.1

..- —

c. _Iift

qs

~ Drag c. @

g= D@amic prtieaure. —

S= &ea of airfoil.

Gfd,= M6ment~,

c = Chord of airfoil. @c

‘%14 (7P =.25——(?L

-—.——

.:

:-.

--.,.

,.

-—

...,..

+.-

..:.,. ;:=.—

: , .~—. ,.k .:. - --- —

...-”, —— -— --- ---- ,

.,. -

I

l-. ,,

-T.y....+.,—=..-. .-r ,.-ya-. ...... . ----- . =. ,

.

. .- ‘,.”.1.

FIGURE8.—AMoW

The results have been corrected for boundary interferace.in accordance with the method . . . .

given in References 2 and 3. For the open jet-the interference amounts to an added downwash

or an increase in the induced drag and induced angle of attack. The ~rrected test POiIItSas

plotted, therefore, corrmpond to lower drag values and lower angles of attack than were measured. Since the airfoils were rectangular, the corrected resultsapply to rectangular wings rather

than to elliptical.
‘IZQSTS OF LARGE AIRFOILS I-N THE! PROPELIJ9R RESIIARCH ~L 767

The resultsaregiven in the form of curyes (b. ?T12) of CL, -CD, W, ~d I% %wt ~de

of attack, and apply directly to rectangular wings of aspect ratio 6 in free air. Numerical values

are given in Tablea VII-XIV. T

In view of the estab%hed rules for aspect ratio correction, another type of diagram has

come into quite extensive gee, especially in England. b this diagram, Figures 13-18, profle

.

drag Cm, &ad and angle of attack ~o for i~t~ aspect ratio me plot~ %ainst ~t Coefficients

By tiply tiding the indu~d hag ~d indu~ ~gle of attack ~m~onding tO ~Y given

aspect ratio to the values from the curves, the coefhcients for that aspect ratio may be determined, thus eliminating the double computation reqtied when converting from aspect ratio

6 to another aspect ratio. Only one curve is given for the Clark Y airfoils, that for the metal

——

covered and painted, as this is comparable with the otier airfoils of the wries aud there were

negligible dMerencw” in the ratifi for the sevm~ s~aces. For handy use the n~tic~

.-

values taken from the ftied cuvss me givenin l“ablmXV-n. The induced drag for the

loading corrwponding to tie pmtictiar wing shape sho@d, of course, be used in deriving the

coefficients fo~ any hit% aspect ratio.

Some of the charactetitim of the airfoils

are quite closely related, and, accordingly, the

results for the C1ark Y with various surfaces

have been replotted in Figure 19. A set of

points from a t-tin the old Variable Density

Tunnel corrected for tunnel wall interference

hae been added for comparison. To compare

the two corrugated wings with the smooth

wing, F~e 20 is given. To aid in the selection of an airfoil for any given speed range,

Figures 21 and 22 give $-%abst \r ~

MAs.

or the speed ratio.

DISCUSSION

The reason for testing at two speeds was

to determine the Presence of w~e effect. The

7-6- ‘------

—

‘——-----l

Sfm&rd &k-Y crdhfes, oh fqo of cwrqofed

.——

meid m fhh akfol W md bottcvn of cwruqokd

mefa’ on fhdr orti”l”

--

‘*k’

FIGUEE4.-c0lrugated IlirfOas. cm Y, basicSedion

““-‘-

differences were ~ slight that only one curve has been ~avzn through the pointi. The scattering

of the po:mtsis,therefore, more u indication of tie precision of tie ~s~. The sm~ forc~ at

low angles of attack limit the precision of the minimum drag coefficient to *10 per cent.

on examination of the CUIW*a few s-g Poin~ @ be no~d. ~rne of tie CIWVW~OW .-~

breaks at the high angles (25° to 300). It was noted during the tests that these breaks occurred

at a higher angle when the angle of attack was being increased than when it was being decreased.

The angles were not changed rapidly so the phenomena can not be charged to oscillation of the

airfoil. There is probably mme effect at these high angles, producing a condition which makes

the flow tend to continue in a given way even though the new mgle of attack dictates a chmge.

These portions of the curves are mainly useful in discussion of rotmy instability.

—

The N. A. C. A. M-6 shows consistently higher maximum lift at 100 M. P. H. than at 80.

Experiments in the Variable Density Tunnel have shown that there is a variation of lift with

Reynolds h’umber which may be quite rapid at certain values. It may be that for this airfoil

the lift does increase rapidIy at these Reynolds Numbers. The small center of pressure movement cor&ms other tests on this airfoiI.

The effect of the different surfac= on the charactaistica of the Clark Y airfoil is shown in

Figure 19. Apparently reaaonably small devkitions from the true smooth surface have SIight

.—

effect on the aerodynrunic characteristics of this airfoil. While the range of surfam smoothness

was not Iarge, the unpainted plywood was certainly rougher than the doped fabric of a wing

-—

as used on airplanes.

..
768 REPORT NATIONAL ADVISORY COMMI!lTE10 FOR AERONAUTICS

c’

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Frmm 6.-Olark Y, metal coverad, rmpalrrtad. Asptnt ratio 6. W ah

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FIOUFIX7.–Clark Y, metal mvarad, painted. Aspwt ratio 6. Frw rdr

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FrGuBx8.-Clark Y, mrnrgatad, A. Aspect ratio 6. Frw alr

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‘IT)STS OF LARGE) AIRFOILSIN THE PROPEIJZIR RESEAROH TUNNEL 769

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FrmIzx 9.-Cbrk Y, mnngated, B. Aspect ration 6. Rea ak

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CL CD

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FIGURE11.—N.A. C. A. M-f3mete.I cover@ Patnted. Aspeot ratio 6. Fr= afr

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FIGVM W-N. A. C. A. 84jmetal mvared,pafnted.Aspwtratio6. Free eir

●

.—L -

--

-.

.—

.—
.

770 REPORT NATIONAL ADVISORY

CL

FIOUEE18.-Clark Y, metal cmvererl painted. Imldte awed ratio.

Frae ah

.0s 20°

.04 /6°

.03 12“

CQ

.02 8“

do

.01 4“

o o“

-.10 -4”

f--!!cfi

-200 .2 ,4 .6. .8 i.O /.2 14-8”

CL

FIGURE14.—ClarkY, eorrugats& A. Infinite 88P rat!o, Frm ak

.05 r?a

.04 IF

,03 12”

c+

.02 8“

&o

.01 4“

o 0°

-./0 -4°

C;otktiit o -.2 .4 .t? .8 /(3 /!2 14-p

c.

COMMITTEE FOR AERONAUTICS

Fmurw W.-CWtthwen 99S. hfkdte eapectratio, Free Elr

..05-- 20°

.04 /6’

.0.3 12”

CDO

.02 8’

%

.!0.01 4’

0“

-Jo -4”

c%

-.200- .2 .4 .6 .8 ~~ j~ ~4-8e

c=”

FIGUE? 17.—N. A. C. A. M-6. hftn[~ aapectratio. Frw alr

.05~ , , “, ~ ,“

I I J- 1.

20”

.02 ITQ’

1

80 AhwI

/00 ]“1

L/Tlllll

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c%

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01

0 o“

-.10 -4”

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t

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.4 ‘.6 .8 /.0 L2 .44

FIOUBElS.–Clark Y, eorrngatrd, R IrUlnite aapwt ratio. Free air

L“

FIOURE18.-N, A. C. A. 81. Inflxrlteeapeetret&. Free W
TESTS OF LARGE ~OHJS IN THE PROPELIJUR EJISEMtCH TUNNEL 771

Examination of the corrugated Clark Y airfoil in comparison with the smooth Clark Y

(fig. 20) revesls a marked flattening of the lift curve for the corrugated sections at the burble

point and a lower negative slope beyond the burble. Throughout the normal flying range the

slope of the lift curve is unaffected. At any given amglethe corrugated surface airfoil shows a

Ill 111111111 11111111 C. c.

[--”------” Wywuad Cm/wed

UeH covereducbhtid. i I f I 2.2056

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Fmu’bxIO.-Effect d dh%mnt sarfema on ohemcWetfcs of Olark Y. Plywood Fmun !Z1.-E~t Of

cover@ metel covered, nnpelnted; metal WV* peintsd

corrq#om inmparisonwith aemcdlr CIerkY. Conngeta A; mnnga , B: metal CO- nnatd, MIIJOth

.G28 B

.024 _024

.020 .020

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— ..

~*=+. qp=+

~GU’Ez Z1.~omperieon onrfes. Profile

-m. mqgog”~’ a= bad .,

FIGWEE Z+C!omperlson onrves. Profile dreg for constant grrse load and

S-@. Dithrent llhfOfk

slightly higher lift with considerably greater drag so that the Lift/Drag ratio is inferior to that

of a smooth ahfoil, although ordy slightly so at 6° and above.

A general flattening of the lift curve is to be noted for all of the airfoils near the burble in

contrast to the sharp breaka often found from low-scale tests.

.-——

—

-L—.

.-

.

-.

..—.

.
772 RII)PORTNATIONAL ADVISORY COW!ITEE FOR AERONAUTICS

The variable density tunnel test points (fig. 19) indicate a slightly greater slope and a higher

maximum lift. The minimum drag is also higher. A comparison with atmospheric tunnel

tests on the same airfoil indicates the same scale effect as that predicted by the tests in the

-.

variable density tunnel. The agreementis quite@ line-with what would be expected in different

tunnels.

-.

Figures 21 and 22 have been prepared @ place the selection or comparison on a common

basis. It maybe said, in general, that the best airfoil_@ any given speed is the one which has

the lowest profile drag, The actual total drag will bg_greater by the amount of the induced

drag which depends on the effective aspect ratio. It is desirable also to have a high maximum

lift to reduce the required wing area.

‘yplO’*(&)~-t(%)= fi’we ‘ecme a ‘

convenient diagram for comparing the sections on the basis of total profile drag for constant gross

(& Do

load and stalling speed. Since ~= o—~ ~the flying condition corresponding to any given

m.

speed ratio is indicated on the curves. The values of speed ratio for these conditions are taken

from Reference 4. For general use the Clark Y appeam the best, while the N. A. C?.A. M+ has

.

the advantage at very high speeds; furthermore, the small center of pressure travel of the M-6

is also of value. The corrugated sections are inferior under all conditions, Too muc,h dependence should not be placed on these diagrams, however; because the partimdar application may

alter the relative position. Tests of a complete model should be the final criterion, Lacking

other data, however, the comparison on this basis will be quite useful.

CONCLUSION .

In the present tests Reynolds Numbers of 2,000,000 were attained by using large models.

This is about 60 per cent of normal full scale.

1. The effect of smallvariations in the surface of an airfoil on the aerodynamic ohsracterietim

is shown to be neg@ible.

2. Corrugating the surfaceof an airfoil flattens out the lift curve at the burble point with a

small increase of lift; but causes a reduction in effectiveness (~/D) throughout the normal IIying

range due to the increase of drag. Pressure di@ributjon_ tests wo~d probably indicate the

nature of the holding off of the drop of the lift curve at the burble.

3. A general flattening of the lift curve at the burble is noted for all the airfoils tested rather

than the sudden break found in low-scale tests.

-.

4. The results appear to be in good agreement yith those from other tests at the same

Reynolds Numbers.

LANGLEY .MEMOIUAL AERONA~ICA~ LABORATORY,

NATIONAL “ADVISORYCOMMITTEE FOR AEROPTAUTICS,

LANGLEY FIELD,VA,, May %’4,1929,

REFERENCES

1. Weiok, Fred E., and Wood, DonaldH.: The Twenty-Fret Propeller R%arch Tunnel of the National Advfaory

Committee fo~ Aeronautics, N. A. C. A. TechnicalReport No. 300, “1923.

2. Prandtl, L.: Applications of Modern Hydrodynsmiw to Aeronautioa. N. A. C. A. Teohnical Report No.

116, 1926.

3. Glauert, H.: The Elements of Aerofoil and Airmrew Theory. Cambridge Univemfty Pram, 1926.

4.DkM, Walter S.: Engineering Aerodynamics. The Ronald Press, 1928.
!EIZSTS OF LARGE AIRFOILS IN !CKE PROPEIXZ!R RESEARCH 9?13NNDL

TABLE I

Ordinatea of Olark Y airfoil metal covered and painted

r

Dfstance

from l.mdl?&%%:t

2-FOOT CHORD, 12-FOOl’ 6PAN

UpllerSurfma I

Iawer mrfam

Standard

!n#la13:

of chord

::

i6s

kE

220

z 67

2.7a

281

2.74

2.62

;%

izd

.67

.20

.02

L22

.66

.s3

.ce

0.84

.46

::

.145

.10

.04

.O1

.Co

-w

.m

.m

.M1

.fm

.00

.0)

.52

Ordinates of Clark Y airfoiI corrugated metal

~ARK Y, &) BBUGATFJ), A

2-FOOT CHORD, 12FOOT 8PAN

..— —CL:

.22

.16

.10

.02

.fm

.m

.m

.m.01 -.01.lm.@) .01 .01

o

L 26

25

5

7.6

10

E

20

40

:

70

E

1%

o.%

. l-z

2.16

2.29

2.67

2.82

‘:. ~

2.78

2.46

202

1:.5J

.64

.12

.-.

I

II:

1 71

204

227

%

!L87

296

Z.sl

Zdi

:2J

L40

.iQ

.43

.04

-——--.

L 51

1.74

209

222

260

274

292

a.01

292

272

24U

L96

L48

.s3

.49

.20

0.. —-—

:6J

:= .23

.14 . 1s

J?J . la.07 .01 &

.co

.m :~ .Co

-m .lm .al .W

.m flz

:: .K.11

cLa4

.57

:g

.=

.21

.14

.I1

.11

. Il.

. Il.

.11

.11

.11

.::

.11

773

——0.66

.40

.24

..27

:%

.14

.14

.13

.10

.10

::

.14

.16

.a

-.

. . . .

:-

.-

-. ——.

-. .-

,..

:..-

.-

.-

“.x

.=. .

.-

..-—

..—..——

—.—

.“---

-.

.=. _

...—

,-’.

:.-

., ..=

.—. e

----

. ...

--

,..4

. . ..-_

..:.

--

i.——= -
.

774 REPORT NATIONAL ADVISORY COMMITTEE FOR KORONAUTICS ...

TABLE III

Ordinates of Clark Y airfoil corrugated metal

CLARK Y, CORFtUQATED, B

2-FOOT CHORD, 12-FOOT SPAN

Uppersurfara I Lower Surfaca

.—.

.—

I

./ .

TOP Of00MI@OM Bottron of wrrlrgf@or19 T~p of “tmrrrgeti;iia Bottom of mrmgat!om

, ...

.-

.

-..

Dfster.ux

momlead.

~f *9

0?%’?

—ifdrf SJ&ia: Meaeured

, ordln@a

foo-#oJ fgo-pot foo::cn.lht

inch iqhw. fnehee

:md~

fxz::mot

inches

,.=.

-/ I

..-..s .-. .

L33

1.M

k%

3.24

L 49

2.00

2,76

2.66

2.4d

2.is

“;l.

.67

.87

..13

...... ..-

0.50

.w

.!23

.91

JJ

.O11

.W

.03

.:!!

::

:%

.Od

0.E4

.da

::

.:3

.16

. la.11 .11

.11

.11

,11

.ll

.11

.11

.11

—. --.-,

0.47

.46

.ao

.m

::

.14

.18

.18

. la.14

. lb

.16

.17

.17

. 1s

,.

...

,.

, ..

;.

,7

.

.

.- .._-==—.

.

- -. . ..

0rdinat6s of G6ttfngen 398 airfoil, metal oovered and paintad

%FOOT OHORD, 12-FOOT SPAN

Upper eurface IIAwer Surfw I

Dfetanoe

fromleed-

~ti @Ot

Or chord

{

M&m&l

fgor~t

fnohea

Standerd

rdlnate in

p4r cent

Of ohord

a.74

k;

“$!!

IL 26

;? tiJ

1a77

la. 84

“Ma

10,M

S.46

;;

i4a

0.90

$

.08

.00 .01

:8!?

.06

.00

.07

.07

.M

.03

.01

.@)

+-._

-.

.

.-.—---

0::

,la

JrJ

.02

:8

.06 ,-

.ai

.00

- .-

.
OF LARGE AIRFOILS IN THE PROPELLER RESEARCH TUNNEL 775

TABLE V

Ordinates of N. A. C. A. M+ airfo~ metal covered and painted

2-FOOT CHORD, 12rFOOT KPA.N

-..

Upp?reork?e

Dfetance

hm leOdf2jw&’&&

Lower eurfaco

.-

--

6tanderd :~:g

p~t~ for 2-foot

of chord chord fn

, fnehes

ap9clEg

fmmooht

fnoha

:=

htirf+jmht

fnches

——

——. —

–0. 33

-.51

::

-.

-. 84

-. .sJ

—.

-.96

zti

-ba2

-. 63

—.2S

:3

Stmderd

mituete h

w cent

Ofchord

.

..—,.,.-

.-:.

---

.. . .-

0

..._

L25

2.6

6.0

7.5

10

:

30

4JI

54

20

70

30

m

1%

o

-H403

4.94

6.71

&32

W

8.06

7.26

t:

3.03

L66

.a3

.26

0

.47

.67

i:

L37

L64

;:

iE

L46

L 10

.73

.87

.21

.03

-!..76

-a ‘al

-z n

-a 03

-K 24

-3.47

–3. 62

-8.79

–3. ‘al

-a 94

-a e2

+g

-1.77

–L 0S

-. 20

o

--42

-.=

:E

-.

-. E

Zz

-. H

-. 95

-.

-. E

-.

-. g

—.

-. 00

-.. --—.— 0.52

.72

LIM

L!Z3

L41

L 66

L81

LQ6

k$

L43

L 11

.75

.42

.34

.09

.—. —

.-

,-— ..-=

.-—”

I

f

I

—

,>.

. . ...

... —

.—

- --

- -- I ..— —

Ordinates of N. A. C. A. 84 airfoil, metal covered and painted

2-FOOT CHORD, U-FOOT SPAN

upper eurfaca I LOw.?rSurfece

Dfstenoe

from lead

‘%zd

.-

. . ..._- ._ ..— ----

——. .—-.

.. -—.

.—-——..

.——

.

-:— - ...-.

--=

-.

I 1

%60

.96

.41

:$

.00

:%

.00

.O1

:%

.00

.00

:%

.Im

260 ~:

L35

L46

!% L 87

2.17

2? 2.40

IL 6 2.76

3.06

E? 8.36

IL 11 3.ss

M 60

U. 81 M

laa :4J

La L06

2.41 .6a

.30 .07

..---— -

L23 ,

L49

L91 ;

2.19

2.42 I

Z2E

3.06 ‘

3.35

8.2.7

8.22

292 ~

244

LS6 :

LCS

.61

.11 I

:MJ

.10

.02

.01

.m

%

.m

.m

.m

:%

.00

:%

.Cnl

..---.. — o.2a

.14

.05

.03

.01

.01

.01

.01

.03

.01

.01

.01

.O1

.m

:%

.

:---

-.

-. _%:

___

. . ...

.-..

-. -:
776 REPORT NATIONAI.JADVISORY COMMITTEl?lFOR AERONAUTICS

TABLE VII

CLARK Y, PLYWOOD COVERED

TABLE IX

CLARK Y, MPJTAL COVERED AND PAINTED

Span, 12 . -. feet. Chord, 2 feet. #mea, 24 square feet,

Reynolds No. 1,940,000

Span, 12 feet. Chord 2 feet. Area 24 square feet:

Reyno~de No. 1,940,0(!0 .U

—

ASPEO’P RATIO & FREE Am A6PEOT RATIO 6, FREE AIR “

CL c,

+

ao46

-. m

-.616

-L 01

:%

.862 .8s .813

.Sm

%J

.276

..2%1

. ..WI

:P?

:%’

, .948

.U4

-

C@

“-a Om

-.061

-.0s2

-.062

-. w

-.062

-.061

-:om

-.049

-.042

-.044

-.026

-. @l

-.022

-,036

-.044

–. 8s2

–. w

–. 076

-.0‘? 5

a

-;”

-8

-6

-4

-2

:“

4

6

;

:

16.

,:

24

:

‘E

M

.. .-

,,.m’ .- CD t/D CL :

–: ~

-.061

.076

.214

. am

.W1

.M2

.783

.9!23

;%?J

L 280

L 221

1.128

L Ml

.078

.W12

.277

:E

.Sm

.222

a CD

-.

,.

-------

.--.....

.......-

7.78

R%

20.E?)

la 60

16.10

1428

12.76

11.21

9.69

7.19

M

8.!77

2.66

2.81

‘2.w

1.s

L 62

0.0160

.Olm

.Olm

.mii

.0118

.0162

.0242

.02m

.0492

.m56

:!%

;g

:E

;;

:E

.678

.620

........i“.

.........

...-----

1; E

.2L22

m.@

lam

IL 90

M Ca

E?.69

U4

1?&!

i 81

::

iu

L 86

L 70

1.56

?;

-am -a ml

-.179 -.082

-L 034 -. m2

1:y; -,060

-.078

.461 -.075

. 2Q4 -.072

.867 -.002

.884 -. Om

.Sls -. m

-.8’56

g

-.046

; fi6

.267 -.063

.811 -.089

.626 -. Cal

,84 -. m

.26s -,068

.266 -.102

.267 -.101

. a62

:%

:%J

-0.246

-.184

-.020

.070

:%J

.627

.781

.217

1.041

L 149

L 215

L 214

1.161

::

.ml

. s@7

.269

.6%0

.

.

... . . ,,.

.

,“ ,

“>,. ...,

. -::. .- , -a

., .*.,.

— -

TABLE X

CLARK Y, METAL ‘COVIBRED, UNPAINTED

CLARK Y, CORRUGATED METAL, A

Span 12 feet. Chord, 2 feet. Area 24 square feet.

Reynolds No. l,940,0d0 3pan, 12 feet. Chord, 2 feet. Area, 24 square feet.

Reynolds No. 1,940,000

ASPE!CIT RATIO 6, FREE ALR ASPECT RATIO & FREE AIR

.-.— .-. ..,”.

.. .._

CL CD’ L/D c’, cu.,{ CL CD LJD CX,]4

:—.,.

.,.

a omo

. Olm

.0124

. owl

.0142

. Olm

.0274

. b27b,

g

.1281

.1610

.2010

.2646

.8190

.2760

.4!xn

.4620

%%

.671

.682

-Ci “~b

-.0$8

-. ml

I. 814

.644

.464

.277

.860

.227

.am

:%!

0“276

,.281

. ml

.W

.m2

.882

.838

.840

:E

.222

.2m

-i 234

-.. m4

-.072

:%?

.868

.406

.632

,7e7

i=

L 162

+.%

LIM

L070

L~

;978

.m4

.m2

:E

.846

.826

–a 070

-. W3

“~~

-.066

-. w

-. WJ8

;: K

-,026

-. Ozs

-.022

“: g2

-.660

;.

-. (S6

-. m

-. W9

-. 0s4

-.076

-.032

-

-0.246

~.

,:f%

.8m

g

i!%

LL42

L 238

;g

i 217

L 163

LW3

. ‘me

:n

;~

;~

:603

.464

.286

.860

;;

.276

.266

.274

%!

.ao

.822

.2m

.8d6

:8%!

.676

.872

-0.074

-:074

-.078

-. U72

-,070

:%J

-.068

-.,(B2

:%J

-.0?4

-.047

~.

-. w

;#

-. n2

-,110

-.124

--------

--------

........

7.16

18.88

2$

la 00

16.92

1421

12.90

lL 62

9.12d

7.2a

6.46

L 16

2.60

2.76

!L84

2.08

L82

:%

L40

--------

...*.-.

......-.

6.17

18.82

17.w

17.61

uLs8

g!

lL 18

9.78

7.76

6.20

4.86

a.m

&10

$L

:;

1.64

L60

:“

,.- +“

. .

-.. .
TESTS OF LARGE AIRFOILS IN TKE PROPEI&EE RESEARCH TUNNQL 777

TABLE .XI

CIARK Y, CORRUGATED ~lETAL, B

TABLE XIII

N. A. C. A. M-6, METAL (JOVERED AND PAI~D

Span, 12 feet. Chord, 2 feet. Area, 24 square feet

Reynolds No. 1,940,000

Span, 12 feet. Chord, 2 feet. Area, 24 square feet.

Rej-nolds No. 1,940,000

ASPECT RATIO 6. FREE AIR ASPECT RATIO 6. FREE AIR

CX44 -1 a Cl/ CD &!D c, CL c. ...

.==....—.-

–;

-3

-6

-4

-z

0

z

-a 261

–. m

-.06s

. Oio

:8%

.495

A#

.910

:%

?%!

LZ5S

L247

L 215

L L%

LC6S

L 017

:R

.m

.24s

-a m

—-0i6

—./

L%!

:%.3M

.361

.323

. 2J.2

:E

.294

:%

.3UJ

.386

. WI

.850

:E

.Zo

-3?4

.373

-o. w

~.O&

–. 036

–. Cfa

–. Cn35

–. 05.5

–. CB4

–. 061

–ma

–.M9

–. U5

–. 043

–. 044

:-

–. 092

-.Km

-. ma

–.Uo

–. no

-.ma

–.m

–.m

–Q HO

–. 471

–.337

–.m

–:Ll&

:E

.4if3

.61s

.7ss

.693

H!

L Q34

‘ :GS&

.9S

.791

.7io

.m

.756

.737

.7m

U&

.mz

:%

.147

.m

.Za3

:%!

.826

-Zu

:E.23s

::

:%

.Zia

:%!

.244

ILCh&

:%%

.006

.00?

:5!

m&

. OIS

.0a3

:&?

.018

–:K’!

–.019

–.022

–.021

–.017

: W&

.W4

.—

.—

----

. . ..-——.

.—

. -..—

.-

. -.

——

.-—

.-

. . .

-.--, —i-

----—-.

—-

GtktV_INGEX398, hlETAL COT-EKED xrm PAINTED

Span, 12 feet. Chord, 2 feet. Area, 24 square feet.

ReynoIds No. 1,940,000

N. A. C. .& 84, METAL COKERED AND PUNTED

;pan, 12 feet. Chord, Z feet. Area, 24 square feet.

Reynolds No. 1,940,000

A8PECT RATIO 6, FREE AIR ASPECT RATIO & FREE AIR

. .’ --: .u—

..-. _ _

“,—. .——

CL

-a 1%3

–. M

–. m

.In

.Zi2

:%

.72s

.s71

Lm5

LIZ

l%

L347

L 82#

L !2?31

L%

LISO

L 149

L lM

:g

.923

.909

CD a c,

–:x ‘

‘-3L M

:E

.43s

.3Q7

.36Z

.340

.8L9

.301

.295 ;

.293

.396

:%!

.320

.331 !

.342

%!

.376

.?S5 )

.3ss ~

45

252

: F4

.4s3

.03

;%J

. 31i’

.803

:%

.2ss

.Zaz

.m

.813

. 82s

:%

.350

.369

. 8n

.-

. - -k

. . . . . .

.-—---

---—--

IL 60

la 01

IQ.Zi

2?!

14 w’

12.81

IL 92

xa 69

::

5.16

432

3.64

~a

121

L s

LS3

L63

.-

- -—. .

- +“

..-

. . ..

--L.:. -.

----

-—

—-“.

... .

..—— -.-,

--- — —

-

..—

. .

- .-.. ___
778 REPORT NATIONALI ADVISORY COMMI~E FOR AERONAUTICS \

TABLE XV

CLARK Y, METAL .COVKIREDAND PAINTnD

Infinite aspect ratio characteristics. Computed from

A. R.= 6 tesiw rectangular loading

TABLE XVII —

CLARK Y, COEFtUQATED METAG F1

bfinito as ct ratio oharacterkdics.

~ R.=6 tests rectangular lo%~~u~d “om

cDf-aam Cx,l

q-a.ecuCL

PLO’M?ED AS FIGURE IS

.:

CW-O. 056.2CL*

Ui=k 694CL

PLOTTED AS FIGURE 15

—.

:.. .=.. -

.+. .- .,.- .-

,“ ~ —

CD. a. cM.f4 cL CD. a,

i ml

.Oow

. m7

. owl

.0101

.0121

.0146

.0179

.0207

.0!?49

. owl

-i 76

-4..26

-a 76

-1.27

ii?

am

47s

6.26

;;”

li40

:7s

-427

-277

-1. re

.X1

k:

4.6s

6.22

7.77

9.44

-o. 0s1

.076

.214

.aEa

.501

.043

.7s3

. em

L OHI

L 176

120S

-Q m2

-.6s0

–. 076

–. 076

-. mz

-.069

“-. W

-. 06a

-.053

:.

-0.086

.070

.212

.a64

.406

. SW

.77s

.910

L MO

L 166

12S7

L259

0.0116

.0104

. Oloa

.0116

. Olaa

. o16!a

.0174

.0203

:%%

. c414

.0726

-clOoii

-.036

-.C86

-.Mb

-;006

-.004

-.ml

-.056

-.046

-.045

-.046

-.044

, , ..<. .._—

. -.

.—.

..

TABLE XVI

CLARK Y, CORRUGATn;pM~;LjLA. CLARK Y, INSIDE

CMTTINGIIN39S,METAL COVIJRED AND PAIN-I%D

Infinite aspect ratio chmmterietios.

A.R.=6 testsrectangular

Cb/-a (W Cr)

W-a 612cL

Computed from

loading

Infinite aspeot ratio oharactaristics. Computed from

A. R.= 6 tests rectangular, loading

CDJ-O. 0669cL:

w-a 676 CL

PLQTTED AS FIGURE 14 PLOTTED AS EIGURE 16

—— ... ..+ ,.

.

...

,.

=.-’ ___

I

CL CD.

: II?;”

.Olm

. Olar

. O*

.011$1

.0171

.0190

.02m

. 02%4

.04W

.072s

cJlo/4

-0.07.3 -.072 -.070

–.m

-.623

-.IM2

-.MS

-.fx12

–. 044

-. 0a2

–. 022

-. W1

Cr,

-0.m

. la

,.Z2

.422

.67s

.728

.271

LLKB

$2

L 847

cm

-

0,0103

. Oma

.0110

.0120

b

;%

. o16s

.ouJl

.o.!a5

.0a02

.04N

.Ona

a.

-i. 76

+J

-L 22

i:

8.%

476

6.so

7.87

9.66

1149

-Il. 067

.W

. 1s7

. aas

.W

.620

.761

i%!

L 14S

l.ma

L 260

-; 99

-447

-z 97

-L61

7$

::

6.66

7.69

1!%

.——

-0.,094

-. Oaa

-.093

-. C&

-. cm

:. ~

-. 53s

-.057

-. !266

-. cm

-.002

,.-. -.:_

. .

----

I

.
TESTS OF URGE AIRFOILS IN THE PROPELLER RESEAROH TUNNEL 779

TABLE XIX

N. A. C. A. M-6, hfETAL @PERIiID AND PAINTED

Infinite aspect ratio characteriatica. Gomputed from

A. R.= 6 teds rectangular loading

t%f.a &i64 cL1

ar~ W CL

PLOTTED AS FIGURE 17

~

CL

-0.067

.rma

.Xn

.ilm

.476 i -m

.76s

I %!

CD.r

a,

& W76 -; 76

.00$7

.0771 ‘i;

.m 281

.Cn63 4a3

. Ol!zl 6.84

. Olm

.OIEs kc

.0233 +aa

. ma

-1cLkf4

amo .037 ,

g I

.015

.Ols

.OM I

.0.23;

.024i

TABLE XX

N. A. C. A. 84, MEITALCOVEREDAND PAINTEID

In6nite aspect ratio charaobistica. Computed from

A. R.= 6 teats rectangular loading

cDfd ~ CL%

t7f=&~ CL

PLOTTED AS RIGUBE 18

--

CL. CD.

-a 070 a 0123

.0s0 .0114

.0111

.:g .OLU

::%

.59s ‘

:~ 1 ::p;

I

.01.ss

iE? .~6

LlE6 .0307

L Z78 .0417

L6!Z4 , . 06Ls

1639 ~ . 09m

a.

-~. 76

-&Ill

-467

-s. 17

-L 66

-.16

L40

294

462

k:

9.40

%%

----“J---

..-.

..- _

.

0