Scale-effect tests in a turbulent tunnel of the NACA 65₃-418, a = 1.0 airfoil section with 0.20-airfoil-chord split flap

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Title:
Scale-effect tests in a turbulent tunnel of the NACA 65₃-418, a = 1.0 airfoil section with 0.20-airfoil-chord split flap
Alternate Title:
NACA wartime reports
Physical Description:
10, 9 p. : ill. ; 28 cm.
Language:
English
Creator:
Tucker, Warren A
Wallace, Arthur R
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Reynolds number   ( lcsh )
Aerofoils   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: The effect of Reynolds number on the aerodynamic characteristics of a low-drag airfoil section tested under conditions of relatively high stream turbulence was determined by tests in the LMAL 7- by 10-foot tunnel of the NACA 65₃-418, a = 1.0 airfoil section with a split flap having a chord 20 percent of the airfoil chord. The Reynolds number ranged from 0.19 to 2.99 x 10⁶; the Mach number attained was never greater than 0.10. The data are presented as curves of section angle of attack, section profile-drag coefficient, and section pitching-moment coefficient against section lift coefficient for various flap deflections. The maximum lift coefficient increased with Reynolds number. Deflecting the flap added an increment of maximum lift coefficient that seemed to be almost constant at all Reynolds numbers. The slope of the section lift curve with flap deflected showed no consistent variation with Reynolds number, although the slope of the section lift curve for the plain airfoil increased up to a Reynolds number of about 1.0 x 10 ⁶ and then remained nearly constant up to a Reynolds number of about 3.0 x 10⁶, the limit of the tests. For flap deflections about 15°, the slope of the section lift curve decreased with increase in flap deflection. The section drag coefficient with flap deflected remained almost constant with Reynolds number of about 0.8 x 10⁶ and then remained nearly constant to a Reynolds number of about 3.0 x 10⁶.
Statement of Responsibility:
by Warren A. Tucker and Arthur R. Wallace.
General Note:
"Report no. L-128."
General Note:
"Originally issued September 1944 as Advance Confidential Report L4I22."
General Note:
"Report date September 1944."
General Note:
"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 previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003613066
oclc - 71198376
System ID:
AA00009431:00001


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Full Text
f lPC> L I


ACR No. I4122


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


WARTIME REPORT
ORIGINALLY ISSUED
September 1944 as
Advance Confidential Report L4I22

SCALE-EFFECT TESTS IN A TURBMOLE TUNNEL OF
THE NACA 65 -418, a = 1.0 AIRFOIL SECTION
WITH 0.20-AIRFOIL-CHORD SPLIT FLAP
By Warren A. Tucker and Arthur R. Wallace


Langley Memorial Aeronautical
Langley Field, Va.


Laboratory


WASHINGTON


NACA WARTIME REPORTS are reprints of papers originally Issued tu 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 128
DOCUMENTS DEPARTMENT


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Digitized by Ihe Iniernei Archive
in 2011 wilh funding from
University of Florida, George A. Smathers Libraries with support Irom LYRASIS and the Sloan Foundation


hltp: www.archive.org dela s scaleeffectlesisOOlang








IACA ACR !'-. rIT,22


I:ATI,'TAL ADVISo'PY COrT.'TITTTE FOE /r.E i.,LUTICS


ADrUA C C'V;TIDE,'TIAL rEPORT'


COALr-EFFECT TE7.S? 1T A f"' '.LTYT T'TJT'EL Or

T'E i'Ah A 6cz-.!l8, 1. A .IR L SI CTIf

TF 0.-ATpRFD L-:HC.7, SPLIT FLAF

Ey warren n .. Tucker and Arthur P. 7'allace





The effect '-f .-n l. s nurbe.r n t-h? er n d',ma'nrd..
chsara ',i-:'i c7 7- 1,' --' 3, *:." r I rf', 1 -.tiln te ted
under rd i tons ol r: lit iveli hi -h .tr. c. turbulere
'as det:'rr-in=d by, test? i- r:: li LAL 7- by 10-foot tunnel
of th.e 6ACA c6z-Ll, a = 1.0 ai0rf'il sr-.cti n ith h
spli t flc- hav"i'in chord 20 r=rnt of" th. rfol 1
chord. The c' E ,ri)1,s nl i be- r r-r. 7=d fro'. 1-1 to
2. c x 10 the .'ch. riurrber r tt irin-d '.'va n v-?v-r !r- after
than C.10. 'The 'Idata ar- r r Ent:d s ,curdves ;f section
ar--.-: .f atta-ck, s-ct .n ro"il;--rs o f fi :i nt nLInd
secticn1 J t r n -r:rren t co- f'i ieri. r. 'ns t section
ILft coef i i .! n rt for v riojs i c fl .ti .-...

The rsxim um lift co ff,'i -ent in cre- s-d *-:Lth De-r.oDld
numbib r. r,eflectin.r th? i 9.' x-I' -.J. sn inr.?-m:-ent of
mYlYxi -.u"1 lift coefifcien t that se"- e. to be ialost cs n-
star.t a i1 ll Pyn los nu.'1b, Pr The sli of 'the c.cti on
lift crv- j r th flr *.efl.Ecte- d zho'.d 1no -:.n sten-
varie.tinn with 'eynol's nurrber, 3lth,' i-h the s c:e of
thL sectionn Hli't Q urve f r to- rl'n 1 rifr, 1l in c r-e -d
up to a Peynoldls nunb.-r of sDort 1.C Y 101- sad then
rer.aine-I' n' r-'1- c ,t ur. to n -,-nold. number )f
about 5.0 V 10', the lr-.t tof the t'1ts. For flau
deflect n s a'-Doe 150', thE sol.op o the sector. lift
curve decreased -ith inzresz-c in flr, defl-t!cor..

The section dr4g zoeffI.iir.t with fl.i-: deflected
remained almost constant with reynol'v- number, lthouh
the sect.r.n .:lr coefficient for the 0lIin 91rfoil .
decreased up to a Rernolds nu-Ibe r of about C.o > 10
and then rc-rrained nearly constian-t to a eynold's number
of about 5.0 x 10 .








2 COFFIDErTIAL Il.CA ACP N:'. Ld-d


The .itchlng-mo:eLt-:n-ffieent -lope with .tlao
deflected was erratic, Lur the *1itc-i.n-mrrment-coefficient
slooe for the )isin ai.rf-il tecLmzme slightly more nezatlve
with increasing Feynolds number.


I ITYDUCTT OI


SCall': effect on low.-ie airfoils has Teclsirl; been
det.-n inejrd at 9eynolds inumbrs taove 7.0 x 1iC in the
I'ACA two-,imenrional 1ovw.'-tibul-nce rr-e ssure tuiinel
ide s gnated TDT). Tests were recently made in the TDT,
in th.. :,A'A two-dimrinsional low-turo.fcA:nc: tunnel, and
in the- [E.MAL 7- by 10-foot tunnel to determine scale and
turbulence effects Dn the lift arind drug hnaract.:r. stice
of a tyriclo low-drag airfcil ,sctirn over a wide range
of f3 ynolds numb c r 'refer nce 1).

The object of the present investigation was to find
the effect of Reynolds nur-ber on the aerodynamic
characteristics of & typlic-l low-dra.; f la secti.,-n tested under condlti-ons of relatively hiEh stream
tu-rbulencr. The FACA (- -!-18, a = 1.0 airf'lil section
equi!-. ped '-L tr a split flap having a ,chord 20 percent of
the cirtoil hznrd 10.20c) .-as tested in the L.'AL 7- by
10-foot tunnel over a range of ynold3 number from 0.19
to 2. J)' 10.


ODF Lr AID TESTS

Model


Two models of 7-foot "pan i] th chords of 1 foot and
L fet were tested. rioth models vnrc built of laminated
wood -ith sul.table stecl reirnforcerrnts and wer,? shaped
to the .ACA (-5 %.i profile. Ordirnites for this section
were derived by the T-thods of refer:nc, 2 .nd' ar-e given
in table I Both models ,:we-re artfullyy finished and
were :-olished just be fore testing.

A 0.G.Oc s.lIt fla waa tested on -&3h model. The
flas_- were mad? of sh-eet steel a:nd were forrm'ed to the
airfoeil contour.

The a'.rfoll section ,,it.-, the flare is shown in figure 1.

CONFIDENTIAL









ITACA AC? No. LL122


Tests


The mrIdels were mounted vertically in the tunnel
so that the test section was spanned completely e;..-'.t
for a small clearance at each end. The models were ri--Idly
attached to the bnlanse frame by torque tubes exten-ing
through the tunnel walls. The angle of attack was set
by rotating the torque tubes by means of a calibrated
electric drive. This installation is thought to approxi-
mate closely two-dimensional flow, thus :'-;,'nig it possible
to determine the section characteristics of the models
being tested. This setup is descrit1i In refere.,ce 5.

E*ch model was tested at dynarric pressures of 1.02,
.,C g.21, and 16.57 pounds per square foot, which
correspond to tunnel airsp eds of a:-r'..mately 20, 4)0,
60, and 0o miles -er hour, respectively. These air-
seeds correspon.l to test Reynnlds numbers of 0.19,
0.57, 0.56, and 0.75 x 106, re.:-ctivel:.-, for the rrmoel
of 1-foot ch-rd and 0.75, 1.50, 2.2b., ,rd 2.09 x 106,
rFrpectively, for the rodel of 4l-foot chord. The
turb,.ulnce f ?,tor of the LMAL 7- b; 10-foot tunnel is 1.6.
Although the data are presented for various test Reynolds
nur'bers, the correspond-ir9 effective F.:,.'olds numbers can
be obtained by multiplying the test Reynolds numbers by
the turbulence factor. Th Li'- -st ':.',Y number reached
was 0.10, so that no effect of i..ch number on maximum
lift coefficient is thought to be present (reference b.).

At each tunnel.alrsped, each model was tested both
qs a ..,Irln arfo: nd vith the fle.. itt L d-h and-
deflect-d It' 305 *':IC, 60 rTh. fl-, ; fl-,t n v.'t r=
set. b- ree ns of t..:r r t 9n.,d ".. Pr.: che :V,':d n.'t.- r .h
t st. The lap w~as su'f'fici nt ly brs :-.,1 s.o tl-.at n:'
oerceptitle dEfl-ctlcn occurred Lnder loCd.

Balance rc 'i ..s wEre 'lus d t,:, r-,.as.ur& lift, rr';g,
and pitching moment, e -ccpt i'or thc drha of i-. Ikl..in
airfoil. Because of the in.:-nsitivity of th,- tunr.el
b-lance -st,--, articu]arl,- at lov; s..j ds, th" 1 Jc ag of
the plain 9'rf.i1 r'as o trainer fro n:,'ak -survcy t.;.ts.

rTh- angle of attack ranged fromrr a negative angle
through the !.t&ll for Pach trt. Tn -iot. cases,
reading: were taken at 2 intervals, ..'ith 1 inmrenmr-nt
near the steal .


CO )N'TDE I'TTI AL


COIIFIDENITIAL










NACA ACR No. L4122


PRTSEINTATTON OF RESULTS

Coefficients and Symbols


The tr-st results are presented in the form of
standard nondiriensi nal section coefficients. The coef-
ficients and symbols used Lare defined as follows;

section lift .-,oeff ..ent f(7/q l

cdo section profi.le-drag7 ccafficient (d,/qc)

cc,,/. section : ':it chin.r- .roi:'-- nt/ .c7e ffic .'.ent about quarter-
mou chordd r-oint (rn/4c5

c max maxirr.um section lirt coet ffi ie ;it

where

7. section lift

d, ssectfion profilee drogj

m section pitchinL moment about quarter-chord T int

q free-str a. d'iynamic pressure 2-p

S asirfnil chard, fe-et

V sirs edj, feet p2er svcon-

p mass density of air, slu,.w rer cubic font

and

S Peynrolfs n7-umber IV/,i)

M !"ach number ('V/a)

a speed of sound (l12') f-:s)

4 viscosity of air, pound-se-.onjs per square foot
a ancle of attack for infinite aspect ratio


CONFIDENTIAL


CONFIDENT TIAL










FACA ACR No. L4122


86 flap deflection, measured from flap-retracted
posi t!on

dc7/daQ slope of lift curve for infinite aspect ratio


Pre cision

Accuracy o" test results.- "h: experimental errors
in the results ,rcsen.ted herein are believed to be within
the 11mits Indicated in the following table.

Limit of accuracy
SChord c' c- C .t 2, = O.L.
(fr ) M-ax A'th -o

0.19 1O 1 to.10 10,-.r- i0.O 5

. .6 i .o06 +.' -.'D 07
.75 +.Ci4 15 +. 00
75 I I.co .,15 t.oC'L
1.5: L 1.05, .12 t.. "2 .
2. L .O.L .COQ ,h)
2.-- iL .03 .CO;6 ,L0 6


The .verae errors ar-. nruch F:ralier. '.ith flap d flected,
errors may be ss mush as three tires the values given.
The angle of attack and fla deflection v;rc held v.ithin
the following limits of a&c.ra&y.

a., degrees . . 0.
65 degrees . . et _

.'ind-tur.nel ccrrcctions. T! lift coeflficiern.ts
are correctEd fo. tuni 1 i nterf3re nce ef fect ( refere-ce 5$.
The drag c -ffi ciernt for the :laln- airfol, which v'-e-re
obtained fror wake-survey tests, were corrected or
blocking as In reference 1. I'o corretlcnJ to the drag
and pItching-rromrnt coeffi2ients have bden 9etter'--ned for
tvo-di-ensicnal force tests in th.e L"AL 7- by 10-fcot
tunnel.


CONFIDENTIAL


CONFIDENTIAL










L ACA ACR No. L4122


DISCUSSION


The curves of section angle of attack, section
profile-drpa coefficient, and section pi tching-moment
coefficient against section lift coefficient, for the
various Treynolds numbers investigated, are presented
in figure 2.

Lift.- The angle of attack at the nmiximui lift
coefficient seem2 to increase progressively with
Peynolds ni.uiber. There is no sale effect on the angle
cf attack for zero lift, although there Is an unexplained
difference between the an l ls for zero lift of the models
of 1-foot and b-foot h-ord. As 2hoG-,n by the curves cf
maximum section lift coi'ffili.ent against Reynolds number
(fig. 5), the scale effect on c- is of the usual
form; that is, c7, In reases -..ith increasing R.
Moreover, deflectin.- the split flap adds an almost con-
stant increir nt of czmax through the Peynolds nLumber
range. This effect is usual for a split flap (refer-
ence 5). The scale effect on the slo,.e of the lift
curve within the low-drag range is giv'En in figure L.
The slore of the lift curve for the rlsin airfoil
increases up to a Feynolds number of about 1.0 wx 10
and then remains almost constant up to a Reynolds number
of about 5.0 x 10 6, the li-'it of the tests. 7ith fla'
deflected, the slope is erratic but approximately con-
stant with Reynolds number. For flag deflections above
150, the slope of the lift curve d:areases with increase
in flap deflection.

DrSa. The effect of Peynnoldz number on the section
profil7-Trsg coefficient is sho'rn in figure 5T. The dragp
coefficients for the plain airf1il were obtained from
wal.-e-survey tests: the dra.f coefficients f:r the air-foil
with flcp deflected were obtained from fore tests. All
drags were taken at the angle of attack corresponding to
the design lift coefficieit (C-.hJ of the plain airfoil;
this value corresponds to an angle of atta-Ik of about 10.

Tor the plain airfoil, the drag decreases sharply
with increasing Reynolds number below a Peynold3 number
of about 0.3 v 10K. Above this Reynolds number, the
drag remains nearly constant.
CONFIDENTIAL


CONFIDENTIAL










NACA ACR No. L4122


For the airfoil with flap deflected, the results show
no consistent variation of section profile-drag coefficient
with Reynolds number. In fact, it may be concluded from
thec.e results that the section profile-drag coefficient
with flap deflected is, to a f*rst approximation, inde-
pendent of veynolds number.

Pitching moment.- The somewhat irregular curves of
section ritching-moment coefficient at the lowest
Reynolds numbers appear to be caused by the inaccuracy
of the tunnel balance system at the low speeds. This
inaccuracy is also shown by the lasge difference between
the orlgInal ,nd check tets at R = 0.19 x 106 (fig. 2(a))

Accuracy at ? nTilds ru.rbcrs h.:7her than 0.19 x 106 is
much better, as A-iwn by the table in the section
entit]:-i "}. :-ision." Th-e slope of the pitchini-mroment-
coefficient curve of the plain airfoil becomes slightly
more r.esative with increase in 'e:,'-~lds number (fig. 6).
The pitc;ii ,-moment-coe -icient slo'e for the airfoil with
flap de]-'.'ct: ed vLri a-d a lth lift co-lfici'-nt in sE lch way
that presentation of the slopes was not practicable.


CONC LU2'_ O'S


Scale-effect tests of the i'ACA 6' -1)l,' a = 1.0 air-
foil section v.' th a s lr it filsn h.avin:' a h:-rd 2 per-ent
of the a.i i-'. l l.l crd }, 1' h en r':_- le irn t.. L.,'.L 7- by 10-
foot tunnel, Th- t.r.'!~- n :."u.: Co, -.r'. e,-i fic:,. 1. c, to
2. 9 x 1 '; t ie I".ch nurr.mb r attained l3s n v-''r -rET atrr
than C..10., Tr-m these t3sts, the f lowing conclusions
have been drarwn:-

1. The rraximurr. lift ioefficiEnt 'ncrased with
reynolds number. Leflectii the flap addedd ,n ii.crement
of m.a:- ; 'v..m lift ?cefficient thit seamed to be almost
constant at all PeynJods n,.irnbe.s,

2. Th, slorc of thr. section lift curve with flip
deflected showed Pno consistent ve.rini,.:n w'th Pe~.nolds
number, althouTh the slorpe of thn section lift ?urve for
the plain airfoil increased uc to P Reynolds number of
about 1.0 x C30 and tiren remained nearly constant up to
a Reynolds number of about ..0 > 1C ,"the lirit of the
tests.
CONFIDENTIAL


CONFIDEHTIAL










OIACA ACR Fo. L1422


3. Fcr flip deflections .bcve 110, the slcpe of the
section lift '.urve decreaseld iith increaFe in flap
d?flecti on.

!. The section rrofile-dra zoe)fficient with flap
deflected remained almost constant with L-,eynolds number,
although the set!or,. prof'l e-drsg coefficient for the
plain airf' il decreased up to a Re.ynolds number of
about 0.0 x 10c and then re'nalned nearly constant to
a FeynolAs number of about f.t' 3 1'.

.. The Rlo:.? of the pitchig-ncmwnt-zoefficient
,urve of the pls!n r-irfilr b:- me sl-ihtly more ng~ntive
with increase in o:-rnold' nu-n'er. :he it hing-mo.i nt-
coeffi-cient slope f.r th aisIrfoil with flap deflected
varied with li ft cneffic-ient in such a way thvat presen-
tation of the slo;:es .a- not ;. r- a.ti.able.


Lan&gley Iero ri n1 A ron ut c 1 Labtr"t '.r
National Advisory i.Committec }-.I- Aeronautics
Laneley Field, Va.


CONFIDENTIAL


CONFIDENTTIAL









NACA ACR No. Ll.122


REFERENCES

1. Quinn, John H., Jr., and Tucker, W'arren A.: Scale and
Turbulence Effects on the Lift and Drag Character-
istics of the i;ACA 653-418, a = 1.0 Airfoil Section.
NACA ACR No. I!.Hl, 1924.

2. Jacobs, Eastman N., Abbott, Tra H., and Davidson,
rilton: Preliminary Low-Drag-Airfoil and Flap
Data from Tests at Lrrge Feynolds Numbers and Low
Turbulence, and Supplement. NACA ACR, March 191+2.

5. Wenzinger, Carl J., and Harris, Thomas A.: 1Wind-
Tunnel Investigation of an N.A.C.A. 25012 Airfoil with
Various Arrangerents of Slotted Flaps. NACA Rep.
Yo. 66.4, 1(5; .

4. Stack, John, Fedziuk, Henry A., and Cleary, Harold E.:
Preliminary Investigation of the Effect of Compres-
sibility on the i.Iaximum Lift Coefficient. UACA
ACR, Feb. 19h5.

5. Jacobs, Eastman [P., and Sherman, Albert: Airfoil
Section Characteristi-s as Affected by Variations
of the Reynolds Furrber. NACA Rep. i!o. 536, 1957.


CONFIDENTIAL


CONFIDENTIAL







NACA ACR ITo. L4122


TABLE I
OFLI'!.TuS OF NACA 653-418, a = 1.0 ATIFOIL SETTION

([Stations and ordinates in percent airfoil chord]

T1;-per surface Lower surface
Station Ordinate Station Ordinate

0 0 0 0
.28 1.L2 .72 -1.22
.50 1.73 i.o0 -1.5
.9 2.21 1.5 -1.7
2.1 ".10 2. -2.56
L. 6L U,. 15.56 -5.22
7.12 5. 7. 8 -5.7
a.62 6. 10.5i 8 -4.41
14. '.. 7.0 15.56 -525
19.67 9.0o 20.533 -5.Lc
24.72 9.91 25.8 -6.355
20.77 10.54 30.23 -6.65
3L.85 10.98. 55.18 -6..2
9 11.1 _:.D. 12 -6.06
.94 CQ 11. 5.06 -6.71
no 10.77 50 -6.56
55.05 0l.20 54 .95 -5.-
60.09 941 59.91 -.12
5.15 8.45 69.8 -5.
70.15 7.57 69.85 -3.
75.15 6.18 74.85 -2.6o
80.15 14 95 -1.74
90.09 2.535 '. -.2
o c.' i. 1. -*.95 .9D
100 0 1:. j o

L.E.. radius: I.C.6
Zloe of radio, s through enc :,f c ar..; C'. 16



IY.TIOA.L ADVISOEY
COI.'iTTEE FOR AERO'_AAUTICS


CONFIDENTIAL


COITFIDENTIAL






NACA ACR No. L4122 Fig. 1


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