Flight investigation of boundary-layer and profile-drag characteristics of smooth wing sections of a P-47D airplane

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Material Information

Title:
Flight investigation of boundary-layer and profile-drag characteristics of smooth wing sections of a P-47D airplane
Alternate Title:
NACA wartime reports
Physical Description:
13, 15 p. : ill. ; 28 cm.
Language:
English
Creator:
Zalovcik, John A
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:
Bombers   ( lcsh )
Compressibility   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: A flight investigation was made of boundary-layer and profile-drag characteristics of smooth wing sections of a P-47D airplane. Measurements were made at three stations on the wing: boundary-layer measurements were made on the upper surface of the left wing in the slip-stream at 25 percent semispan; pressure-distribution measurements were made on the upper surface of the left wing at 63 percent semispan; and wake surveys were made at 63 percent semispan of the right wing. The tests were made in straight flight and in turns over a range of conditions in which airplane lift coefficients from 0.15 to 0.68, Reynolds numbers from 7.7 x 10⁶ to 19.7 x 10⁶, and Mach numbers from 0.25 to 0.69 were obtained. The results of the investigation indicated a minimum profile-drag coefficient of 0.0062 for the smooth section at 63 percent semispan. At the highest Mach number attained in the tests, the critical Mach number was exceeded by at least 0.04 with no evidence of compressibility shock losses appearing in the form of increased width of the wake or increased profile-drag coefficient. For flight conditions approaching the critical Mach number, variations in Mach number of as much as 0.17 appeared to have no effect on the profile-drag coefficient. In the slipstream, transition occurred at least as far back as 20 percent chord on the upper surface at low lift coefficients.
Bibliography:
Includes bibliographic references (p. 13).
Statement of Responsibility:
John A. Zalovcik.
General Note:
"Report no. L-86."
General Note:
"Originally issued October 1945 as Advance Confidential Report L5H11a."
General Note:
"Report date October 1945."
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 - 003613546
oclc - 71214748
sobekcm - AA00006283_00001
System ID:
AA00006283:00001

Full Text

ACR No. L5H1la


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARTIME RIE PORT
ORIGINALLY ISSUED
October 1945 as
Advance Confidential Report LSHla

FLIGHT INVESTIGATION OF BOUNDARY-LAYER AND PROFILE-
IRAG CHARACTERISTICS OF SMOOTH WiRI
SECTIONS OF A P-47D AIPLANE
By John A. Zalovcik

Langley Memorial Aeronautical Laboratory
Langley Field, Va.







NACA


WASHINGTON
NACA WARTIME REPORTS are reprints of papers originally Issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change In order to expedite general distribution.


L 86


DOCUMENTS DEPAh I lI if


4


VA-C /









































Digitized by Ihe Inlerine Archive
in 2011 Wilh lullling from
University ol Florida. George A. Smalhers Libiaries wilh 3upporl Irom LYRASIS and lhe Sloan Foundalion


hllp: www.archive.org details IllighlinvesllgalOOlang









FACA ACR i1c. L5Hlla

NATIONAL ADVISOF..Y COi.0IITTEr FOR. AERONAUTICS


ADVANCE CC:IFrDENTIAL REPORT


T.Tg-I: 'T iVS.'TGATIO OF BOUTTAR.Y-LAYER A-1 PROFILE-

DR.AG CHARACTER STIC'3 OF 3.OOT .'.. IG

.SECTi;S OF A P-47D ATRFPLA PE

By Jc.l-.r A. Zalovclk




A flight investigation was male of bounrdary-layer
and profile-dras characteristics of smooth wing sections
of a P-47D airplane. Measurements were made at three
stations on the wing: boundary-layer measurements were
made on the upper surface of the left wing in the slip-
stream at 25 percent semispan-; I'-ressure-distribution
measurements were made on the up:er surface of the left
wing at 63 percent semispan; and wake surveys were made
at 65 percent semispan of the right wing. The tests
were made in straight flight and in turns over a range
of conditions in which airplane lift coefficients
from 0.15 to 0.68, Re,,-ynolds numbers from 7.7 x 10
to 19.7 x 106, and Mach numbers from 0.25 to 0.69 were
obtained.

The results of the investigation indicated a
minirnum profile-drag coefficient of 0.0062 fcr the smooth
section at 65 percent senispan. At the highest h:ach
number attained in the tests the critical n.ich n'unber
was exceeded by at least 0.04 with no evidence of com-
pressibility shock losses appearing in the form of
increased width of the wake or increased profile-drag
coefficient. For flight conditions approaching the
critical ITach numroer, variations in I!ach number of as
much as 0.17 appeared to have no effect on t!he profile-
drag coefficient.

In the slipstream, transition occurred at least as
far back as 20 percent chord on the upper surface at low
lift coefficients.









2 '. : 'IDi; ,TIAL i:.CA ACR Nc.. L Klla


rI ".. _...'.'IO:I


I.-n r.ier tI o -' .ii coi.'-r i'on or tie -re.file-Ar..
c'.-t rac t r. st s. of vwir. sect. ns of low- 'ri d o- er
t--o:-s u.i-er ".r.-11i 1.' t c i:. itio;3, tc ts L -have been
Tmade cf u v', o F-i 1 -r a : i.e P-4tL r Pla.L. e, h3 vin
Rep'u.1l : 3-; scctio, .s a.n.. thie .'-?-'s airp'lan; -iavin
s actions .' at arie. f12'-. an ;t,CA *.C-series secti in t
the p'I .:- cf 2'"..im tr- i n L.-. -cerih, s.- l ion a- t
the tip. T'!:h ir.ve rt ation .' .t ..in, 7ectlion of the
XP-,7F ]irpla e in._:1., 3. test "o 'iS terminc t.he profile
drag of a vin.; s e r c n :t il. te Is irs rs,"1 ar.-d he
Fosi1:ion of tr'.asiti-r :.rr .: 'cn3 'rsile :.-a cl utz rlde
the .3li. tr' e-j Thj 1r.3 3.1 Ls of thIs in-:esti-aticn are
Fre-.:e ted in e '-I" n, e .

Pt? tests :.it:- the -.73D i..lrj..n, rerncut>: herein
were r-e.errall- -imiil." in ec -, to h-e o .e 3sts -it h the
V .F-.x' 1 irrl.an -:. i' -x-:t that t:e te3u?' wI; t the --17
were ext 1ei 'dez to. c ..icl j l ,' -' a;:zh nr'.;;iher' i.n
crdr-r Lto ol- t in so'. in-"ormat n* on c-.. ssib lit
effLcts at it'.:.a h n:13 rs ,.!-'cuj 'tr critic-1 val..;
'ih:- restJ ei'ere r..ade in 3tra:,is' f :it t and st-:a y turns
at various nor rnal acelerdti ons c.ve r. s e of indicated
sirsreeds fr:n 1 ) to j~5 rc.il-s per .our I t al titues
of 1 ,us00 :,.nd 2 ,O..'00 feet.





c se .3 in chord

x d'ist 3.nce r, al1n; ci'l'i f'rc.ii leadin.cg e.dge

s disA3lncl: a1cn.; sur-'c ce rfc;:r i leadin- edge

dc def'lecti ...n of tn_.riv .iz'r-e ,.cZc

y distance abov:;e sur.rce, ]-csi:-crn in wa:ke

IH free-str,-eam total r-'ez3u.re

Hy total pressure in bo!njiry- layer
71


CC "'T LE TAL









.'."A .,'R jTo. L51lla CO;'. IhT AL 3


AH loss of total pressuree in wake

pO free-stream static pressure

p local static 7Pre ssure

qc fr:-.-stream impact pressure /Ho P)

qo fr:, -stream d-r.smic pressure 1PoV)

Ty absolute t:.nr.rature in boundary laocr

T5 absolute temperature .*ust outside boundary lavFr

u velocity in .indry la-,;r

Ui velocity in b~injar" la.-:r near surface

U velocity .'ust outside bo.undar-' layer
!P 'o
P pressure coefficient \

P critical pressure coefficient, corresponding to
local velocity of sound

CL airplane lift coefficient

c:I section profile-dr a coefficient

6a aileron deflection, negative for up deflection

Vc calibra_.l airs: -e (airs, ee related to
differential pressure by scc'pted st.:.iard
adiabatic formula :..ed in calibration of
differential-pressure indicators .id equal to
true airspeed for standard sea-level conditions)

V true airsn,d

R Re.,nolds nur;.ber

'Io free-stream ;!ach nu niber

1.-, .:ch ni rrier in bound. ry ly.eer

.06 ~sl1ch num1rr ;u'st outside bo;unary layer

C.l'TPFID2:"T- II\L









CON TZJ.


*~~~~Z (N'- -t1
lit no. Lj,...~..


.'r cri ical hrech nctar

g ccel'-eration of grav'.'ity;

pO free-3tr&am density

Su3:bs ripts:

B. rih t

L left


APP.iR'.TUS


The F-'-7TD airplane is a low-vwin., sinile-engine
or.ool..e u Prc.' hitne-- R-2. ,.- l enrine and
a i'cuir-bl:. -e Purtis 3lec::.ic pro:eller (1'i. 1 The
a irrlr ne hsc ;.. g.'ors .eitht of ar ut l12,0i'. pc.L. s, a
viwin, ]L :- i : l Feet, ano a %.'inr ar ;a of 3010' square feet.
"' in incc ,ortes ,.er'ublic .-;. sirfcil sect.inns,
W'hici have r-: nresnire- ListriO ion charac tes tics similar
tr, trio.e cf the ,'-l,-A .25u-serie3 sections.

'L.re. -,inT sections were tasted (fig. 1): one on
rU. ri, t wing -fl i one on 'ivi laf: win- lCcated 3 per-
cent Ter.l spafn lfrom .'.;e plane cf s;,r L.Str.', or a'out 2 feet
outboard of '"i rlap (section nith aileron); and cone on
a-. 1: l f '.vin, l*-c.t .i. perc,'; semis3pi n from dr' e plane
of y: tr cr -.bout, 1 f.c iC:. in !he e lge of the pro-
poller 1 is. _Laclh ,f tie utcArd s-cticns lhad a chord
of -.1- i'at and a maxia .um tlicXnesss of 11 percent chord.
The iniboard: section in ut :- slipstre.mn had a chord of
3.7: feet rid a rmaximr:u.. i tfl '.--.es cf 1I.6 percent chord.
A photograph of ti.e test section ri, tihe right v.in is
shovin as f iure 2.

The upper surfa e.' of the sections en ihe left wing
ind .:he upper and lcher surfaces of the section on the
i)-..t wing were fairad by filllin tith glazing putty
anu th' en sa:dtin: smooth to reduce tihe surface waviness.
'The. surl'I'ceZ. wer.- then sprawy.. with; several cc;-ts of
white lacquer-based oiint for a protective coating and
sanded li':itly in a chordwise direction vwith No. 520
carborundum n ar.r. An indication of surface waviness
./a c, o> tin-d by means of a zurva.ure gage (fig. 5) with
legs spaced 4 percent of th:'e :iing section chord. The


COIrD:I;I!TIAL








'AC ', ACR ::n. L5Hlla C: -FIrE:'TIAL 5


wariness condition of the faired surfaces is indicated
in fi.g~re 4 by the plot of the wavir-ess index d/c
a-ain st s/c.

BP.'- nary-layer racks, each consist.inr, of one static-
..re3as.ti- tube and either one or five total-, pressure tubes
(fi,. 5), were used to determine boundary-la:--r charac-
teristics. The tubes were rladsJ of -inch brass tubin. with
1
a --inch wall thickness. 1The upstream end of the total-

pressure tube was file..: ,nd flatte-r.e3 so as to leave an
1
r,-ning 0.'0 inch .2p and inch wide and to have a
0.003-inch wall thickness. '7"e. static-pressure tube `-.ud
six orifices 0.02 inch in diameter e4i:!.y :p.ced around
1
the perirh--ry at 1- inches dwcvnstream .'lori the hemispherical

end. L.ch total-rr:saure tube of a rack 03 connected to
*.n NACA recordil multi l.! ,.nometer i,. referenced to
the static or-essure obta:c!.,d fr.i th e static-pressure
tube set i- 'ut 1/4 inch from the surface. .,ith this
arrLnemnent, the la:pact pressure was measured at various
distances above the surface .'L.e: the six-'.' u r';k w i3
used and near the surface when the two-tube rack was
used. The static rrssure Measured by the stat.c-pressure
tube was reference, to the static Tre3sure obtained '-r
means of an airZ -ed ..L-.-i mounted on a boom 1 chl-rd ahead
of the leading; e-;,- of the ri rt ":in,, ip (fig. 1).

Sir-..,'s of the wa~:e of Lhe. right v.inr section were
macs b-. rIe ns of ;e rake sov.'n In T i .ire 6 ,,-cL.n ,~d 1" per-
ent 3hord hc.lin:1 th t.r'Ailin e-e. TL.e rA-:e Consistd
of 2 1 Lotal-p'res- ure cab,..s spac?- u.5 inch ;.ud zLatic-
pressure Lu'es spaced e .I ll 3cro ';,iFe r.Z:e. The c,;tal-
pressreesur tu'es w-r We cnnecte:. te arn ACr, :.'ecordin.'
-lultille mronret-r anl- referenced ao free-stren total
pressur- in :;D.Cer L.a t-.e total-pressure loss at eac.i.
point in t-ie v.'.,ke ccul:J. ,t obtained 'Tie static nress'ire
in the alke ivcAs eas -rcd v.*1 ': t.h3 -th:r' cei,,ral st.-.tic-
pressure tubes, each r" f ;hich ws- cctnncted c the
mnincmet er, and rfe-renccd to the static rresz-re ehasurE.d
b; means of the airsrpeed head cn Ch'e boor. ja the ri.ht
wing tip. .:,Iol tufts were lccsted on !-.e upper s.r face
near the tralin.-edge region a cDut 2 feet on c-aeh side
of the center lina of T[Ie s~rutio:n t 6 perZcnt s3iisin
to det.rmnine whether an' croSS flr.' existed hijlt would
invalidate the wake surveys.


COr '1 IDIEiT AL








6 C'UIT ',- TI.L CA ACR _'o. L!illa


All pre-.ssur::, al ie'on Tositions, and normal accelera-
tions ,jnre mIeasui3 ed :)y LJACA recording instruments. An
inrlicatinr; aIceleromee :- was frovl'ed for tie pilot.





Tn cr-er to obtain fre--strea", static cr'esps'ire,
correct or on dettr-:ir.sd froin i-, airspeed calibrati n
:ere ..ne to e e staEtic rleastsre measured b': the air-
speed i.ad .-1io:nted on ,.Te boon: i-ea2 of t-he right wing
tip. Tl ese ccrrections were applied to Call :rasirements
f..r which r-iference t ft'r- 2-r.;: static pressure was
required.

2o-.:o. -- yetr .lncity; pr fi les wer. .e determined
from th-e c' :'iarv-la or r, eas -emncnts by i.;se of the
compress ible- -.'l o -e la ior.


H, 'r 1 1.. 2 T
S2/ -,1 1 '

LT

/2




cr, to a first-c .ruL a-- porr:.:i.lation,
: I




T:ie li.rnl.ne lift c oefflciernt I t wi..Ceh transition
occurred at a riven chordxvis3 position was determined
froi, a nir.t of' the ratio U1/IT a -inst aiJrplane lift
,os3ffic'enrt. T-he lift coefficient ccr.rsponding to
transition was chosen at the el--bow of the curve as the
ratio U.'TU suddenly increased f rom it. laminur level
to it. turbulent level.

The Fro"'le-dead coefficients were determined by the
irtegrating .nettod of rnierence 2; that .s, the total-
pressure loss was integrated: across the w.ke and then


C" irTPTn M T'T.L










1!.A 1.0.A No. L5:lla CO' T '.". L 7


mult.Tlil::d by factors Jspcr~ding on free-stream impact
Fpr's.ure, naxi:':1.U total-pressur'e loss, static pressure
in the ia :e, and flT .F.t Mach number.





Surveys of the wake of the smooth right wiiTr section
were made first in 3trji.;t flijit wit. level-flight
ro-e r and with the airplane *,-iir. throttled .:d then in
turns in ".,:er to cover a wil.e r;-nc of fl'ht conditions;
that is, airplai e lift coe!'icients, Reynolds numbers,
and T.ach numbers. D..rL.: he first flijh.t in turns, the
filler used to fair ite .in' z. :-r'ace cracke-; at th1.
leaiinI edce of the a:munition-L.-,r. airtment !lDor (at
11.5 percent chord). Since the crack could not be kept
smooth ind the surface unro'.l:-rn in su?.seuent flights, the
wake -.,rv\'ys were discontinued.

Boundary-la-er meds.re.m-.,ts were r.-.d: both i.'ith the
t.wo--lvbe nd~ the six-tube bou.;irry-l-."e.2 'aks on the
upper surface of the inborde section behfin_ t-s. Frc-eller
on the left wing. Li.asurenments c: static r:re3s.re and
of in'lact :-ressure next to the surface for ".:e deter-
mination of transition were made -:.:ith two-tu'b :'a:s
at '5, 10, 15, 20, ani 25 -'-cre-nt chora'c. 7essurerrents
of velocity d istribjtion t~r:.ouT- the bondiry la.-rer were
made '.-ith the six-tube racks at 15 Arnd 20 percent chord.

rsasi '.on i~ens urem .-.ts cn ttie ir~er -urfa"ce of' the
outboard sec in on d-.? left .i 'Are not' fe-.As le
because of l-:e sp~.n.si3e crac-: ac t.e ieLdin' ed- of
the A.Lr.uniticn-ccnart-ent h-or at 11.5 .-rcen i crd.
Static rer3ss3 .res, I-o.;:e'.r, r'.r3 .meas^re.d 't:. i-e
statl~ c-F-res-~ re t-ubf-s c.f :''- ociu-:dar"-lT y-3r ra'1:s at 10,
15, 20, 25, and' J.G percent chc:d on the upper surface
of tdis nin.

The teL;ts .vY.re i.,.e in straight fli-Jit (level f'iMght
and shalllo\ dives.) a at lti :.es 'f 12,000 J -C. d 2.'~,CuI feet
over a range of lndiiate_.d a rspe ..r fr.om 1-, t. ;-U riles
per hour. Tke .rpl mil1 lift coefficients 'j.t:-l.* ed in
these tests ran~ea frcm 0.15 to U.6.a; the rHe' n'ids
nu-.'ber, from 7.7 x 10-' to 1-0. .< 103'; rnd the r.lach
nuniber, from 0.25 to 0 .'6. Tests '. ls-, r.d in
nurns at an altitude of 12,C.C( feet at indicated airspeeds


CONFIDENTIAL










CC .DETIA.L IEAC. ACr. Ik L5 la


fr.'o.m 56 to 7O miles per ho'..r and at normal accelera-
tions fro.,1 -g to 41.-. The .ai ro)] Lan lift coefficients
in the turn:r ranged frrm 0. 21 to 0.5.; PeT-'ec ds numb.?r,
fr: : 1. :' 1 i0' to 1].7 x 10C'; arn >:h number, from 0.44
to 0.l..

FEI'UL S A D DI ~SC',U .T.IO'

F assure d' strip butioL *nd cl i tic? .], 1`'2zh nur,:ner.-
St.e re present s t ic-- rejs,..ri ..Ustrib.utions over
I-art of the u.?.er surface of tic left vini sections
at 25 and t. 5 p- r:-ert : ir _.- e.r n h:. .n in i .re 7.
Thie or' ti,'-3l .1hszf nu. cers of th!- twvo wing .c.ctiio.s, as
deterr'ir.3d by the von I r..r: nrti.od iref:-rence ) from
pre ssure-dli str jbuti'o r:.eas-lre.' Lr!t s at subcr tical needsed,

are '-e tted n f I)r C r .c --, .r
\/C .-- w

recresents th.e lift 'n ,ef fl lenLt tset vouli be obtained
if t'e ra.ha rI n -j'nocr were increase9 u fie nr,) o to -.cr at
the n:ryle of attack corrz.esoniitn. to CL. The flight
r:ach nuirrer nd I-l's d efle : tlon of the left aileron are
pl:tted abIove t..e cur of cr tcal rach r.r.b- r.
For t-he se-.tion at c1 rer-ent ser.isoan, the critical
I..:s nm.ber- ';ar-ed aO,.r.-Y irt tel' v line: rl-y ro.m O.6 at a
lift cceffici.nt cf 0".1' to '.', at a lift coefficiernt
of 0.SC; r the. s-ct. .on at 25 ,eroe,.nt se il s. a, th-
-. r.Li tion of critical f sah njum:'er ,cr th1 .e-Sr r-ange
of li-ft ,3neffl 1i& ts .z- from 0.-7, to C.L Although the
evali-, ti on of critics. l i. 'Y nu'rbar ir-'-rlv e' e:.:tra:r .o1,,ti on
tl.. th.e von i'ar,-,,n ,-re':-cd of statlc-presr're data cotauned
at fli,-tht ;:"ch llunmer.-s rrna-in'T fro-! 0C. to 0.50 telov
the crj.ti-al value, the results vere ir, P od a;reerent for
the entire range -f ithe extra;,oljtion. The extent of the
extrar.olation Et vr-a : c;s lift -icefficients r.-y be deter-
mined by conparin- t''he f light 1,:ach numbers at v.hich the
pressure-distri utl.;-i reas.-.rements were rrma.e wi.Lh the
critical 'l '.uch nu.-bers. 'Se fig. 8.)
A,.cor..dingr. t t'-:e results :res'.-ited in reference 4,
the critical :.'ach n : i'ers : -s determined from measurements
with ststic-przvaure tubes similar to tbosoe used in-the'
present investlzti,.tn r-ay be .as rich as 0.01 h.'lher than
wold be o.)tained fr'm neiassle'r.ents with orifices flash
with the wing surface.


CONFITDENTTAL








.AC\ .CC2 CIo. L.5 lla CrFIlET TAL 9

It should be no"'-r' that, since the left aileron was
rdeflected -pv ari fro.n 1. 50to 3.60 d.o';inr the tests
(fi.,. '), .:he cr' tical I'ach; ru.r.'bers 'at Cj, p l--it soiri-.o n
may obe so.irerat hichLr than the critic,l '..-'ch nmrIhers that
wou Id Le obtained with the ailer-or neutral. An indication
of thP railitu-d of this effect is :-'"n in reference 5,
w.'hich presents the results of tests of a r--'l of a '.ln-
sectionr vit. aileron on a P-Lb.7-3 -1sr:lane. (The [inr." sec-
tions of tnis airplane are zirilar to those of a F-'I.T,' air-
plane.) Tho results in referer,-e 5 s3.ov.ed that, at a
constant nr.le of attack Iin te rrn_ of th-' flJgbt tests,
th~e critic-l -Mach number was Ti. her by about 0.015 with
the aileron eflecteJ 1.':v.--rd 20 than with the :l-leron neutral.

3-undary -la"er h.-r-srarist ii si '- trean.- The
method of deter. lir, L -e '- :Ie' li 't i*fl eior nt, .:e-
tion ne.,nili.s n'i-ber, r'-: fli. ht ; Ic'. ru: ber c :rrc s3 nding
to transition .frrm measurements vit', a boiu-,drr:'-lay.'r
rack in a i ven position on the wn-r. surface is illustrated
in fi-zure 9 for a rsc., at 15 per-ert ch'orJ- on the u;oer
surf-ace in the slipstreams (,t 25 percent semirson). The
bro'-en lines in .s f ;Tcre L:r.i .t:, the c ncr iions for
transi ti on.
-._n: re. It s 'f0 the, bon.] alrr-! -.y:. r reisurem nts i--di-
c t t.1 .: hat I t low lift o ;ic'i..tS, 1rn nar flrw was
obtainLe, at. l.eas z s fr btL.c'.: a:.r 20 percentt ch',rd on
the '.i r 'er :surf9eS v:.'l:?h s .-.'c.at Ss far .,acl,-- a "E
be e; e te -n a .si. 1' sr win: s-tion :t s'r d t'i.z '.r -
peller- sli stream. Latin'-r f!o:-. at 20 ier-?..nt l.Scr1 1s
il lustrated ty:, troical velocity .rc f les ,' fi -.i* 13.
'[he lift -oc-i'fic .ents, e.ynolds t'.b r :.. 2 : nrs
at v:'--ch transition vi~w oOtas.:ei at 1?, 15, rad Lu 'er-
Sert chord are iver- in fi.-mre 11. A' lift refficie ts
and eirynolds nurioers lcss h t th.e ir.dicatc-d. b;.r the
curves fcr 15 and 20 percent lhord. in figure 11, 'he
flow wvas lanm.inar at t'-.-re coord:Lse positions. IAlthugh
transit-c.n Tmea.3surerents were also ale bt 5 and l'. per-
cent chcrd, these data vere. r.ot presented, "r.'_:smu.ch as
the flov, w.as always larrinar at 5 -ercenrt -h )rd 1aid
alv '.,ys turbulent at 25 rie'rcent chord.
Profile dr'a~. of vw'.-:, .s- cti on '? utsil1e s' pstrSr ..-'
Du:'in u ll the t sts the vool tuf"l: o.n L e -i r.er surface
ne.?r co percent serlisan of the riwht win~ v,.'Fre direct-.d
stral '..t bacl, and thereby ndi--tedr taat tci v..ak.e sEu.rve'ys
were not influen-d byL cress- flo".
The crofile-dr?.z f refficienit- t- f the S:.- S th =scti-n
on the right wing are presented .-n fiSure 12 fcr straight
flight and in fig:ire 1` for turns. Pli~ht :.ach number,


C3r'FIDTDTI AL








NACA AC C o'1. LS,11la


critical Mach na-ber, Reynolds number, calibrated air-
speed, &.nd Idefle-tion of te right aileron are clotted
above the orofile-drCg curves. The critical' Mach
nurber shown in figures 12 and 15 is that for the left
wina- section. Inasmuch as the richt aileron was down
(figs. 12 and 15) when the left aileron was up (fig. 8),
the critical Yach number fDr the right wing section has
been estimated on the oasis of the results of reference 5
to be of the order of 0.02 lower than the r- tical Mach
num.rber .-f tr.e left wing section. Some representative
we':e r-rofiles obtained in straight flight are shown in
figure i4.

In straight flight, the nrofile-drsg coefficient
varied from 0.0075 at a lift coefficient of 0.63 to
0.0062 st a lift coefficlent cf t.15 !fig. 12). The
minimum profile-drae *2ceffL-ient was 0.0062. Within
the accuracy rof the reasuremients, chanting from level-
flig;ht po-er to glides with engines throttled appeared
to have no effect on the orofile-rlrag coefficient.

The interoretation of the results of 'hie orofile-
drag reEasurerr-ents in t-rns (~ig. 15) complicated by
'he fact that a crack developed at the leading edge
of the airrrunition-coirmp'rtmrent door (at 11.5 percent
ch.ord) some time, dlur-.1p the fli-;ht in which these measure-
ments were rade. The tendenc;- tn vsrd lower orofile-drag
coefficients f'or "he irst series .rf turns than f',r the
other series indicated that the craEk may have developed
after t:e ii rrt series of tu.rns. At lift cn-:fficients
greater t'an O.10, the profile-ire coefficients in
turns agreed v.ith those c:ts'ned in straight flight and
thereby i.,dicated that tr-.rs-tion was probably forward
of 11.5 percent Thord t' these high lift coefficients;
at lift coefficients less than 0.'10 the profile-drag
coefficients for the eccnd !-cnd third series of turns
were s-mr.ewhat higher than those obtained in straight
flight. The minlrsum profile-drag coefficient for the
second and t-ird series of turns was 3.0066.

For flight conditions apcroaching the critical
Mach n.umT.ber, a varis.tion in ,.:ach number as large as 0.17
(fig. 12) with a relatively small variation in Reynolds
n.imbe; appeared to have no effect on profile-drag
coefficient. A similar result ;ws obtained in the tests
reported in reference 6 cn the share wing section with
transition fixed near the leading edge for smoothed and
moderately roughened surfaces. A comparison of figures 12


CONFIDENTIAL


CO rFT DEz. T AT









CC 'I DENTIAL


and 13 shows that, at a lift coefficient of 0.-7, the
s "e value of profile-.rt'-f coefficient (within the
r."erimental error) was obt:.r.eJ at ',:l. numbers vary~in
from 0.50 to 0.5); in this case, however, the var nation
in Feyolds number was ir.e- (10 x 106 to l x 13-) and
therefore is: have had an effect on the results.

At the h-'ghest Mach number attained in the tests
(0.60), the critical M ach n.- .:er vas exceed s by at
least 0.' (fig. 12) with no e',dence of co:,-ressibility
shock losses apse'ar'ng in the for rr f increased width of
the wa'- and increased rr-.'*fil:-- d_ coefiicient. This
result ; .esars to -ndicate either that irrotational flow
without shock existed to some extet at supercritical
sp ees, as s'i testedd in references 3, 7, and 8, or that
t -- effect of compression shock w;s of insufficient
rr nit7,l- to ::- reasurable b :- resent a'..:-ratus for a
small range i.f :.' i numbers above the critical value.
Mild compression shocks ":.eve been indicated ." S'.,ileren
n i.tc r ;jh obtain:-- in vi-a. t;rrels of "CA 2 5-series
atri oils. T"-se :,h.- t-. slow tl:at, upon attainment
o f local veloc ty o sound, shock first a' c s as a
series of snall shocd- waves and builds up to a well-
established shock front as the '-.nurber Is further
incr. a sed.


o ---L- 3aOnS


Thes flight investigation of bovundar7--l '-r- and
profile-dri cGharacteristics of v:.n sections of a
P-',D airplane that were sre-ially 'nr.is.-V: to ive
aero .ynmrri.cally smooth s-i-faces having :aViness rnf E-&l
.r'.aS it.:ie. indit~&ted the follo'.ir..' results:

1. 31.-rd'-ry-lay;.?r transition at least as far back
as Z0 -ercent chori i as ).c.tl; ned ci] 'e .rfacer of
a section in .he sli :stream a t IYv li- t t1: 'iV'l : -t3.

2. Tn strain h.t flir-ht (leIel fl c.ht -r shailo,'
dives) tihe crof.le-ir --ff ciernt r*:I s -cti: .uts de
the slicstres-r 'aria.-? fr.-' C. .C.0c2 at r li- o fficient
of .15 to 0.3 07 st llft ?:i]ficir.t of C.63. T:e
nminlturI profile-drja-; ,?c ff icir.t w'Bs 0.c0;c2.i

A. At the h:'~ es- t i'c!- r. -.e2 r att .id in -th tests,
the or.tical '.7ach n-urt',r ,vr- etxe1d,-ji ty .st least 0.0L


COR- T'i'T.;.L


!InCr ACR No. -L' 11. a










12 C00!7ID:'TTAL NACA AC: ::c. LHilla


wit' :? i. v1i. c-3 c -r' r'esr3 t t s' .-c! losses a-pearing
Li t-'e i: r n. f Ireged v. rith &i tt .; '.-- or I;:r as: d
*or- f I e -.'r3' c 'eff cier. t.

F" Ifl i t c n, ti : a.:proscLi t: 'r c r- it -a1
:';.ic!' n:' .ter, vsriation-, in :.?cn nir ler as lr a e ss 0.17
a.;esar:d d t. ?-'-ve nc efifc t on -rof1l e-.-! r' c o ocfficient.


Lar.. l-y ?.emnorisal Aer-.nir.t i cal Lsbor: toru-
i:'ti 2.al Advissrr'; Co:r mr .ttee i orp roni.jut ics
L le,- Field, Vs.


,!iD r7r TTALT









'.A .'" 7-o. L5H11a C :'FT D-' TIAL 13


'I E`'- r_


1. Zalovcik, John A., and S!'oog, "ic1 ri B.: Fli"ht
Investigation of BPu.-idary--i:,-y r Tr-.rs'tion and
Profile Dra,. of an E.-.erirmental Low-tr:", Wing
Inst-ill.ed on a Fihter-Tc.pa Airplone. iACA .A :
iln. L50Oea, 19115.

2. Silverstein, A., and S'S-tzoif, S.: A Sirllfi1d 1..ethod
for Deterrrinr!l' "*n: Profile Drag in Ili.:ht.
Jour. Aero. Sct., vol. 7, no. 7, .'ay 1o.0,
pp. .25-301.

5. von TI4'r.-ann, Th.: C -pr3->..-lity Efrects in Aerodynmrirs.
Jomr. Aero. Ecl., vol. 6, no. 9, J'.ly 0i-l,
*.. 557-356.

. ZaIovcl', J:.hn A., and Daur, Fred L.: Fl,-ht Investi-
cation at -11h Mach N\ubers of S-veral Met'- : of
sMeasurlr, Static Pressure on an Airplane vr'i._.
NACA '".. LJH10a, i'L.

'. z!,-'r., Arvc A.: Effect of C-.mressibility on Pressure
D stribution over an Airfoil with a Slotted Frise
Aileron. "^CA ACR Fo. LIG12, '1;L.

6. 7l-vceik, John A., and W .- Clotaire: A Fli it
InT e S Cti 1ti n of t'e E" li''e t < .f .!rf_(ce Ro1, o s
:.n pr, 1r. st ,- -1 r.' t on f -q .e': 1
".' .. .. r f Li T7 1 ...
r,, L- .-

7. sa,!a :n, C' rl: '".P .. : -,,.r e s bl- F luid .'st
a Curve E 1rf'ace. "AC.- ,. e. I" 5.

T. -arrici T. E., and i'pplan, Carl: On th, Flov. of
Cr,'cre sible Fluir the Hodo-zraph Method. I -

*TT ACL1 ; LLCe 1iL. 1 slf i tioin ch -ineDd
to Fest r ct d ,' t. 1'L_.)


C'FTP FI ;- TIAL









NACA ACR No. L5H11a Fig. 1





9.




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cbe
.iNI z0=

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0),







NACA ACR No. L5H11a


CONFIDENTIAL


Figure 2.- Smooth test panel at 63 percent semispan on
right wing of P-47D airplane.


CONFIDENTIAL


Fig. 2








NACA ACR No. LSHlla Fig. 3













to



Cd


I )Cd


at

C)


ho














-I
0d

,)
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NACA ACR No. L5H11a


CONFIDENTIAL


loixo-4


(a) At 63 percent semispan on right wing.

IYVI


6 r- Upperdr'r 2'ce
d/c
4


(b) At 63 percent semispan on left wing.


6 \ Upper surface


NATIONAL ADVISORY
COMMITTEE FOR AIERO UTICS




OE +


.10 .20 .30
,5/C


(c) AT 25 percent semlspan on left wing.
Figure 4.- Surrface-waviness Lndex of smooth surfaces of sections
on right eni left wins of P-47D airplane.


J10104


d/c.
4


CONFIDENTIAL


Fig. 4a-c


0


40 .30 .0O








NACA ACR No. L5Hlla Fig. 5a























oI




.d Cd


z 00
LaJ

z ::z
o .. 00


-









8-





IA



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NACA ACR No. L5Hlla


II


II


Fig. 5b







NACA ACR No. L5H11a


CONFIDENTIAL


Figure 6.- Wake-survey rake mounted on wing of
P-47D airplane.


CONFIDENTIAL


Fig. 6







NACA ACR No. L5Hlla


CONFIDENTIAL


(a) At 63 percent semispan.


Per
-0.97
- I. 23
-3.07
-10.26


0 ,IO .2.0 30 40
-X/C
(b) At 25 percent semispan (in slipstream).
Figure 7.- Typical pressure distributions over upper surface
of sections on left wing of P-47D airplane.


CONFIDENTIAL


Fig. 7a,b







NACA ACR No. L5H11a


d
u~ -




-4


.7 -


,6"








3 % O


Mcr


.5 i !

CONFIDENTIAL

0 2 4 6 8

CL/ o At 6 -
(a) At 63 percent semispaln.


CONFIDENTIAL


--- e--- -- ---- -- -- -
C






















NATIONAL ADVISORY
CONNITTEE FOI AERONAUTICS

3 .2 .4 .6 8

CL/ t //I-MJ.c
(b) At 25 percent semispan.


Figure 8.- Critical Mach number derived from subcritical pressure
measurements on sections of left wing of P-47D airplane. Flight
Mach number of pressure-distribution tests and deflection of
left aileron are plotted above Mcr-curves.


Fig. 8a,b







NACA ACR No. L5H11a


CONFIDENTIAL


,4---
U


0.--
0
CONFIDENTIAL

CONFIDENTIAL


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure 9.- Method of determining lift coefficient, Reynolds number,
and Mach number corresponding to transition at a given chord-
wise position. (Example shown is for 15 percent chord.
Effective-pressure center of total-pressure tube at 0.01 in.
above surface.)


O


Fig. 9







Fig. 10a,b





























0
'-4


rz



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9

|JC O T
0 orr~oiM-t
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nitltl


NACA ACR No. L5Hlla


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re






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NACA ACP No. L5Hlla Fig. 11










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NACA ACR No. L5H11a


CONFIDENTIAL


+eg 0

deg 0


20 -101


A/t1fude
(fr)

--- 12, 00
-+--24,000
-x- 12,000 EnQine throtf/ed

Z 4 000 ft


CONFIDENTIAL


cdo .0060


Figure 12.- Profile-drag coefficient of smooth section on right wing
of P-47D airplane as obtained In straight flight. Mach number,
Reynolds number, calibrated airspeed, and deflection of right
aileron are plotted above cdo-curve. Mlc-curve is from results
or left-wing tests.


Fig. 12







NACA ACR No. L5H11a


CONFIDENTIAL



deg 0o


X
---- =--xX --X r--
xx-..


-o-- Firs series of turns
- +- Second series of Turns
-x--Third series of turns
-. I- 1_ 1 I I I I


.008C


do .00o6

.0040


xx- -

SSZ1 r / -// .5ght

0 .1 .2 .3 4- .5 ,6 .7
CL


Figure 13.- Profile-drag coefficient of smooth section of right wing
on P-47D airplane as obtained in t.rns. Mach number, Reynolds
number, calibrated airspeed, and deflection of right aileron are
plotted above cdo-curves. Her-curve is from results of left-
wing tests.


4VU


30C


200


100

0


Fig. 13







Fig. 14 NACA ACR No. L5H11a








sA x.
----------.s-> ~ ^




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a








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3 1262 08104 974 3





UNIVERSITY OF FLORIDA
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