Pressure Distribution Over a Rectangular Airfoil with a Partial-Span Split Flap

17 Pages·2014·12.49 MB·English

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NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
REPORT No. 571
PRESSURE DISTRIBUTION OVER A RECTANGULAR
AIRFOIL WITH A P ARTIAL-SPA N SPLIT FLAP
By CARL J. WENZINGER and THOMAS A. HARRIS
NASA FILE COpy
Loan cxr:'c~ or. last
dele stmnp ,'( ... -c~ pcvor.
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PLEASE ETC TO
REPORT DIS"' :'" "IION SEeTIO
CENrER
U~","'_EY:\- :;'.JI~
NATlor r L t -.F't.mAUTICS AIID
SPACE ADi,dNISTRATION
Lt'vlry h,ld, Virginia
1936
For sale by the Superintendent of Documents, Washington, D. C. - - - - • • _ • _ • • _ Price 10 cents
AERONAUTIC SYMBOLS
1. FUNDAMENTAL AND DERIVED UNITS
Metric English
Symbol
Unit Abbrevia- Unit Abbrevia-
tion tion
TLeimnget _h_ ______________ tl mseectoenrd _ __________________________________ ms fsoeocot n(do r( mori lheo) u_r_)_ _____________ fste.c .( o(ro rm hi.r). )
Force _________ F weight of 1 kilogram _____ kg weight of 1 pound _____ lb.
-
PoweL _______ P horsepower (metric) ______ ---------- horsepower ___________ hp.
Speed _________ V {kilometers per hour ______ k.p.h. miles per hOuL _______ m.p.h.
meters per second _______ m.p.s. feet per second ________ f.p.s.
2. GENERAL SYMBOLS
w,
Weight=mg v, Kinematic viscosity
g, Standard acceleration of gravity = 9.80665 p, Density (mass per unit volume)
m/s2 or 32.1740 ft./sec.2 Standard density of dry air, 0.12497 kg-m-4_s2 at
W 15° C. and 760 mmi or 0.002378Ib.-ft.-4-sec.2
m, Mass = -
g Specific weight of "standard" air, 1.2255 kg/m3 or
I, Moment of inertia=mk2• (Indicate axis of 0.07651 lb./cu.ft.
radius of gyration k by proper subscript.)
J.I., Coefficient of viscosity
3. AERODYNAMIC SYMBOLS
S, Area i Angle of setting of wings (relative to thrust
lD,
SID' Area of wing line)
G, Gap Angle of stabilizer setting (relative to thrust
b, Span line)
C, Chord Q, Resultant moment
b2 n, Resultant angular velocity
S' Aspect ratio
Vl
p-, Reynolds Number, where l is a linear dimension
V, True air speed
J.I.
=4p (e.g., for a model airfoil 3 in. chord, 100
q, Dynamic pressure V2 m.p.h. normal pressure at 15° C., the cor
:s
responding number is 234,000 i or for a model
L, Lift, absolute coefficient OL = of 10 cm chord, 40 m.p.s. the corresponding
=;
number is 274,000)
D, Drag, absolute coefficient OD Center-of-pressure coefficient (ratio of distance
of c.p. from leading edge to chord length)
Profile drag, absolute coefficient OD. = ~S Angle of attack
Angle of downwash
Induced drag, absolute coefficient OD, = ~S Angle of attack, infinite aspect ratio
Angle of attack, induced
Parasite drag, absolute coefficient OD = DSp
• q Angle of attack, absolute (measured from zero
lift position)
0, Cross-wind force, absolute coefficient Oc= q~ Flight-path angle
R, Resultant force
,-
--
REPORT No. 571
PRESSURE DISTRIBUTION OVER A RECTANGULAR
AIRFOIL WITH A P ARTIAL-SP AN SPLIT FLAP
By CARL J. WENZINGER and THOMAS A. HARRI
Langley Memorial Aeronautical Laboratory
78659-36 :
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
HEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.
LABORATORIES, LANGLEY FIELD, VA.
Created by act of Congrcss apj)ro\"(~d ~lar<:b 8, 1915. for tbe supervision and direction of the scientific
study of t b(' problems of flight (r. s. Code. Title 50, Sec. 151). IL' membership was increasecl io 13 by
act approved :)lal'elt _. 1 !)~!). Tbe Illember,; a re appointed h)' the 1'I"e~i(kni, and . cl"\'e ,IS sucll wit bout
comJlen~ation.
JOSEPH S. AMES, Ph. D., Clt aiI'll/ali , ('HARLES A. LINORF.ROl r, LL. D.,
BalLimore, Md. ='lew York Cit,\'.
DAno 'V. 'l'A"l."LOR, D. En"., Ficf' ('/lIIinJl(I11, \\'ILLl"\~r P. i\lACCRACKI~"', Jr., LL. D.,
WaShington, D. C. Washington. D. C.
('II.\HLES G. ABBO'I', P.c. D., Auous'rIN," 'V. ROBIJ\'S, Brigadier Genel'fll, United P.late' Army.
Secretar)', Smitbsonian Institution. Chief Materiel Division, Ail' Corps, Wri~ht Fi!'ld, ] ayton,
J,YA1.\N J. BRIGGS, Ph. D., Obio.
Di rector, National Bureau of Standard;:;. FlUGE,," L. YIDAL, C. E.,
AIl'I'IIt'R n. COOK, Real' Admiral, Unitpl] ~tatt's N:1V,Y, Dil'(>Clor of Ail' ('omml'I'C'f', Df'p:lI't nwnl \If (""llmerC\',
('bier, Bul'('uu of Ael'on'ltltief', Nu\',v ))(,P:I rt 1llf'1lt. EDWARD P. WAR:-IER. l\L , ..
\\'11.1.1'; HAY (:m·:(:G. I: ..\ .. 1\P\\' York City.
Chipf, l'niled ::>tatps \\'(',1111('1' 1:111''':111. O;:;O.\R 'YES'I'OYElt, i\I:ljol' G('n('f':ll, Unil(',1 , i:ll\,,,; A rillY,
IlAKHY l~. GUGGElNllEIM, M. A., C'llit'f of' Ail' ('orp,;, 'V,lr DL'p:trlllH'1I1.
Pori 'Y:lshingtull, Long }>;Ianl], X. Y. Olt\·ll.l.I,: 'YlnoIIT, ~t. D.,
~YJ):-IEY i\f. Kn,AuH, ('aplaill, Unilf'(] Siall''; Nav)', 1):I~'1,,", Ohio.
Bur'uu Irf .\prull:llllic:;, .\':lV.I' l)pl'arIIlH'IIt.
(:ICOHI:Io: \\'. LEWIS, ]);1'('('/01' oj .t('}'onfluti"o/ R""(,OI'('/1
.TOlIN I·'. YICTORY, "'('('re/a}'11
Ih:NRY ,1. I';. 1{~~IIl. /'In{lillce)' ill ('!t a I'{fe. /,all{liey J[em01'ial .~eroll(lliliea/ /'(100"0/01'11. Langley F'i.eld, Va.
JOIlN J. JOE. 'I'eC'hlliNI/ .i.<'{!is/Ollt ill Hllrol1e, Paris. FrailI'('
TECH:\TICA L COMMITTEES
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AIRCRAFT STRUCTURES AND MATERIALS
COOl'cti1!atiulI oj i(('sea1'('/t X('e(/s of Mi1i/(I1'!/ (lncl CiL·it .Ivic~tion
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('onsiriemtion oj /lll'entioll.~
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l'nilk<l conduct, for all agf'ncies, of ConN'(ion claSsification, compilation,
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II
---- -
--
REPORT No. 571
PRESSURE DISTRIBUTION OVER A RECTANGULAR AIRFOIL WITH A PARTIAL-SPAN
SPLIT FLAP
By CARL J. W ENZINGER and THOMAS A. H ARRIS
UMMARY and hinge-moment coefficients were about the same jor
the partial-span plit flap oj the present test as jor a
PTessuTe-distTibution tests of a OlaTk Y wing model
jull-span split flap previou ly te ted.
with a paTtial-span split flap weTe made to deteTmine
the distTibution oj aiT loads oveT both the wing and the I TRODUCTIO
liap. The model was used in conjunction with a Te A con iderable amount of aerodynamic and air-load
flection plane in the N. A. 0. A. 7- by 10joot wind data i available for imple split flaps extending along
tunnel. The 20-peTcent-choTd plit flap extended oveT the entire pan of the wing. (See reference 1 to 3.)
the inboaTd 60 peTcent oj the semispan. The tests weTe However, because of the general use on present-day
made at various flap deflections up to 45° and coveTed airplanes of the outer portion of the wing span for
a mnge oj angles of attack fTom 'ZeTO lift to appToximately aileron, most of the installations of split flap have
maximum lift JOT each flap deflection. been limited to the inner 40 to 60 percent portion of
The Tesults aTe given a aiTjoil section, 01' Tib, pTessuTe the wing span. Some aerodynamic information on
diagmms fOT the wing with flap neutml. IncTements oj partial-span split flaps is available from wind-tunnel
ai1joil section pTessuTes aTe given JOT vaTious amounts force te ts of both rectangular and tapered wing
oj flap deflection so that combined wing and flap sec equipped with uch flap (references 4 and 5). There
tion pTeSSUTe diagmms may be obtained by a simple are available, however, very fe\\T air-load and moment
addition. Oalculated coefficients oj section loads (md data suitable for the de ign of wing with partial-span
moments and oj wing loads and moments aTe also given split flaps.
fOT the wing and flap combination and JOT the flap alone. The present investigation wa made to obtain in
It was jound that deflecting a paTtial-span split flap formation, particularly of pan load distribution, that
ajfected the pressure and the section n01'1naljoTce and would be suitable for application to d sign problems
pitching-moment coefficients oveT the entiTe wing span. involving partial-span split flaps. The data were ob
The flap loads and moments weTe almost constant oveT tained from pressure-di tribution test of a wing model
the span oj the partial. . pan flap jor a given angle of with a 20-percent-chord split flap in the . A. O.
attack and flap deflection. The maximum normal-jorce 7- by lO-foot wind tunnel. The results are given as
1<---------47.0"---------->1 o.rifice
r, -------34.0." >I stations
1;. /"1'-1-'- -25.0." >I in inches
1 23.0"----->1
-12.0.·~ 0.
I r .50.0. Secft'on c-E
r-lo~~t-
1 I I 21..0000..00..
1~'-=~--d='---df. 3 _~+ _=_.=l+=~:.~ ~-J;l 5.0.00.
=.l..= - .=-----=...= != 9.0.00
12.50.0.
il ·--il illi--· 15.00.0
16.50.0.
~ t~ an.
00.0.
17.625
~. . __. ~. - ~ .. ~----.~. . ~ ~ allBB.. 675205.
-.... --.. ------.-- -------------~--~---~---~£~--"~~===l=tt-
------------ a o.rifices on lower
---,,==~~~--::+
surface of flap
c..,J,1 I ped poriion of wing
1<----36.0.": 0.60. biZ (Flap)·· I and on upper surface
1<------------60..0.·'- bI2----------->I of flap_
FIGURE I.- Diagrammatic drawing of emispan Clark Y airfoil with split flap showing pressure orifices.
1
2
HEPOHT NATIONAL ADVISOHY COMMITTEE FOH AEHO "AUTI S
F'lIinn: 2.- Sidr view of model in tunnel. l,'JGL"HE 3.-Front "jew of model in tunnel.
---------
-1.5~ .C
I
j + 1 "r ------
1------1 -35
f- -
+---
f- !--- -
I-l---l-+- Section A ---
-30 l- " c--
.. 0---
.. £----
" F-----
ttL\ " c--- I-
-25 Upper surface 0
~~ Lower " 0
11-.
11\\ 1- - - +----I--
II il\ --
-2.0 bh
~~ R l-I--
, I--'
fT-'\ '\
1, +-' " '0
-15 , \i. 7
,
,
l!q.. ,~ , \ r-~" \ ~~ . \' \
-/.0 "- , ,~ -
t-r- ./~ I
V I
, , -~,~ I
',l~
-.5 , '~-u... - - -
<::::
"" -a.
'v
o V c..--I-- - I---->--- 0 -- l_~ N-'"; "-,=~ r;:;;:
.5 rfIIw~. l~ d=6° .5 I1 / 1//V -- V~?- k-:- ~,-- b-- I~-d-- r~; -/5°
·ll- /\/'
l-t- rl
I
11
fO /.0
FU;l l{to~ 1. Pressure <iistrihulion ()\'er airroil ~e(,lions at three angles of attack; 0/=0°,
3
PRESSURE DISTRIBUTIO OVER A RECTA 'GULAR AIRFOIL
pressure diagram for the hasic wing and as increments and wa set at the variou deflection by suitable
in pre ure or load for the Dap-deflected condition . pacer block a shown in figure 2. The gap between
The charftcteri tie are given for the wing-and-flap the leading edge of the flap and the wing wa sealed
combination and for the flap alone. with plasticine for all tests.
Pres ure orifice were built into the upper and lower
MODEL A D APPARATUS urface of both the wing and flap at several ction
WING MODEL along the emispan, 166 individual pre ure orifices
being installed. The tube from the orifice were
The Clark Y wing u ed for the e test had a 20-incb
brought through the model and out at the inboard end.
chord and a 60-inch semispan (fig. 1); the portion ex
Two orifice on both the upper ul'face of the flap and
tending back to 80 percent of the wing chord wa con-
o. ..---- .;:.
-.5.-- r I
ffijF-l Ff- -r "F"" ~
r.H' 6, -15°-t-+-------1--t-+_+-+-'''+'''+-+-t-I
;;;..+-~~,
-5~_+-+~~~+-+-t-I--t-+_+~~,~~.~~~/
~-1_4_+-+-+0-304°-+-+~~~~~~df~'~~:.~
I -<>' -::I::::P ~P' \l.
o - 'c·_ t::: -:: I--:~ 1= ---i-- -o-:~.,..... o r.!l=l_ :.:.:.r==j--- ':= C.. :r'.F ..- l'- -c- J: "'"
1"2:- ~t:: ~:, ,~, .:Q::
fJ.p .".... I . T r-r:::,,,,,,,, "",
q Section A -- """.
.5 +--~-+_ t--"-r-- B ------ i--t-t',--,+-,'-:·I.-.$;',1+'I -f-i1·,1 --_Ir_-- 1
"-r-- C -- ",
+--1--t-+_+-" -r-- 0 --- - -
" - £ ---- -t--t--t-+--t- -I
.. - F -----
-/.0 .. - G ___+ -+-t-I--t--+- -I-
\\-I--Ir--t-+-+-+--t-!--IIIr--1- Between wing i - --
~), 11 .1 WingI --arnd flop ,
1~,,'r-- LUopwpeerr nsu rf"ar-c-er - 00 +I --1 +x ~ I~. ::. ~::=~t:~
-. 5 ,~ ., I I I ~.lfl'" .'
':-., :---1_ .L I . I-~ ::g,p ", ~ ~~17'
- ;/
.. -., )/
-'
Jl./
5
. '1
f'!
J I I I I ~
l7 J FIGURE 6.-Increments of pressure over airfoil sections at 6° angle of attack with the
FIG URE 5.-Ineremen ts of pressure over airfoil sections at - 6° angle of attack wi th the flap deflected various amounts.
flap deflected various amoun ts.
on the lower urface of the wing at sections E, F, and
3tructed of laminated mahogany. The trailing-edge, G (fig. 1) were u ed to obtain the pressures between the
upper portion wa con tructed of X6-inch-thick teel flap and the wing,
plate rolled to the contour and attached to the wooden MANOMETER
part of the \';-Jng with crew. Th flap, which had a Two N. A. A. multiple-tube photographic-
4-inch chord and a 36-inch emispan (20 percent c by recording manometers (de cribed in reference 6) were
60 percent b/2), was formed by hinging the inboard 60 u ed to record the point pre ures on the model. The
percen t of the metal lower surface of the rear part of manometer was connected to the orifices by mean of
the wing about its leading edge. The flap wa up rubber tubing, 0 arranged a not to affect the air flow
ported at its leading edge by small piano-type hinges over the wing.
4 REPORT ATIONAL ADVISORY COMMIT'TEE FOR AERONAUTICS
TEST ARRANGEMENT
~I The model was mounted in the J. A. . A. 7- by 10-
foot wind tUlmel (reference 7) in conjunction with a
-/.0 H -+-+-+-+-+-+-l--j--l-+-+-+-+-Hc-l-+---t-l reflection plnne at the inboard end. (ee.fig. 2 and 3).
This plane extended from the top to the bottom of the
air tream and some di tance ahead of and behind the
model and was 0 located that the model was placed
ymmetrically wi.th respect to the tunnel center line.
The tube from the pressure orifices were brought
horizontally through a rotatable section o( this plane
to the edge of the jet and were grouped together to
form a treamline shape 0 that the air flow on the
opposite side of the plane, on which the model was
located, wa not appreciably di.sturbed.
-2.0
r-r-----
TE T
+-+--1-+--+-
The static reference pre SLlre u ed to maintain the
1- -- ----r--
-1.5 !-r----f-f--IH--+-+-+-+- dynamic pre sure con tant during the te ts wa cali
brated ao-ainst dynamic-pres me mveys at the model
location with the model removed from the tunnel.
The longitudinal tatic pres me at the model location
-1Omt-lH-+-+-+-+-+-+-+--+--+-+-+--+-+-+-+-H
wa also measured and usecl to correct the point pre -
I I
m'e to the correct refer nce pre ure.
All the tests were made at a dynamic pressm 01
16.37 pounds per square foot, corresponding to an air
speed of 80 miles per hour under tandard sea-level
atmo ph eric conditions. The average Reynold Num
=.----1------- ----- --- .-....
o -- -- -- ber of the te t , ba cd on the wing chord, wa approxi
~. T T "
mately 1,220,000.
The model was tested with flap angle of 0°, 15°,
30°, and 45°. The angles of attack covered a range
.5 f--l-+-+-+-t--+--+
from approximately zero lift to 15° (approximately
I CT.-max) with each flap etting, te t points being taken
at 3° intervals. lInen the model had been fixed at
-2.511-+-1--+-+-+--+--f--l-+-t-
-+-+-+--1- ,.- a given angle o[ attack with a given Hap setting, a
JI-~r-I-I-
Il-I--I-+-+-+ Secti±on A ----I- few minutes \\'ere allowed [01' condition to become
~-t--l-r-+--+-H-:: ~ =----.--+--+--+-+-+__+_ constant; a record was then taken o[ the pressures at
t
~ ---~ the orifices hy means of the photographic manometer.
-2.0 . W :: :11--+-1-1--1--1--+
H-r-i I-r----:: i ~ =~~ rr -+--+--+--+- PRESE TATIO OF DATA
1t.-1--1--I.-. ~;j;:1·~.r~;~-:=--+-l1 _I-Between wing,-+-+-"
-1.5 Wing _t-and flop __ Section or rib pre me diagrams with the Hap neutral
t-- Upper surface 0 _ r-i-x -+-+--+--I-+-- (fig. 4) are given as ratios of point pres ure p to dynamic
1lIH-I--Lower " I D -t--I-+ -t---t--t-+--t--I pres me q for a low angle of attack (- 6°), an inter
- 1-~-+-+-~-+~~~-I-+-~-
mcdiate angle of attack (6°), and a high ano-le of
attack (15°). In addition to th section pre sure
diagram with the flap neutral, the increments of
point pres me with the flap deflected oYer the point
pres me with the flap neutral (both in terms of the
dynamic pre ure) are given (fig. 5 to 7) for all
Hap deflection and for the three previou ly mentioned
angles o[ attack. On these diagram the flap pressure
are plotted from the deflected flap chord but normal
to the wing chord. The principal advantage of the
increment diagram i that they may, by the principle
of supcrpo ition, b applied to pre sme diagram o[
any other basic wing ection that does not depart too
'------.L-'--' 1 I I I I 'Il:=..fL"-,--,--, greatly from the Clark Y ection of which the te t
FIGURE 7.- 1D cromenls or pressure oyer airroil sections al 15° anglo or attack with lho
were made.
nap deflecled "ariou" amounts,
PRESSURE DISTRIBUTIO T OVER A RECTANG LAR AIRFOIL 5
diaTghrae mdsa taa rec ogmivpeunt eda s fronmon dtihmee ninsitoengarla tecdo efpfircei enutr.e OI n /4=N-l l-O2c&/-4] w'm g pI. tCh m' g-moment COef-f-il C.le nt
c qcw w about quarter-chord point.
The coefficients for the wing-and-flap combination
include the load of the :flap projected onto the wing C"f=~-Z ] .f l ap ctl.O n hin ge-moment Oef-f1i C.l ent
chord. The coefficients are defined as follows: qCf about £lap hinge.
C"w = nqcww ] airfoil section normal-force coefficien t. O"f= qCIf~ /bI ] £laHp ahp inhgine-gme. oment coefficient about
2. OIrrr---- --1l1a25~ 0 0 debr~els H1t 7.. ~ ---- -~I S.JoIo II I I I I I v<IIl ~-- 10+ --I, ---,I- --I, --'03]00~ . .I- .,II-'- --I'I (X)1 1rr---rr1--- DItI:J II II II II I II I S9I°lo rrrr----rrrr---- l'La\)l l, ~e~)e~sl 111125 00 ---
1
1.8 I
I. 6 I I-t-
I--
1.4 I-
I.e? I 6.( 0° t-I- -- "'-"" I-- i'- ~= 15.0 l-
1.0 l- ;--I- "'- -- I-- I',
. -<>-1-<> r--- I',f..o.V -- I', c-'IA
.8 I- /'\ 1-1-- r---
~rJ- .6 I""i-o- - -- " " 1""-1--"\ 1-- ro- -- t--- "'---
.~ .4 r-r-- '-... t"""\ 1-1- ;-- --
:.:Qt:.:Jl .2 -1-,~<>-I- ~" t-I- 1- ro-
\oJ a ~ ""'\ ,~..... -I-
QJ
\.) ,
~-.2
--
I
~co
1-
- I- --r--,
g l.8 ;--
r--
."§0 I.S --- -" '\I' r- " '\
~/.4 ---, r---i-- 1"'-. " 6.( =300 '-- Ib. "- "- 6, 450 -I-
~12 - '- " 1\ I'<>. I\, r.... If
-- 1b.1"" " b-. ;-- " ;--
.~ r---- t- 1"'- l"-
" /.0 I- I'-- I--
.8 '-- 1" - 1',- i'- ,~
l- i'- l- i'- t'1>. /'\
.6 '- 1" 'l- '--- i'- I', A ' i'-I- ,,-r-~ t- " r---- - """
.4 I- -- 1" - i'-;-- 1"----'\ I-/-- "- ;"1 "--1- ;-- I-..
.2 -. 1,-~0. 1---0. r-f-- '- t-t- '- t-""
o I "- \., r-j"-I-~
'"
-.2
-.4 -- I-
~-"..
G r ED c B A " G r ED c B A
Wing sem/spon Tip Root Wmg sem/spon Tip
F,GURE .- Span Joael clislribuLion on Lhe airfoil with the flap deflected various amounts.
0". , w = qNcw 'bOw ] wing normal-force coefficient. (c. P.)w=( 0.25- Ccm".lO/4 ) X 100tu] daiinrafol ilc esnetcetri onof lopnregsi-
c = nf] flap section normal-force coefficient. sure in percentage of
nf wing chord from wing
qCf
leading edge.
ONf= Nbf ] flap normal-force coefficient.
qCI f .
m (0. P.)u=( 0.25- 00/4 ) X 100] w0g center of pres-
cmC/4 = qcWCw~ 4] aircfooeiflf icsieenctt ioanb opuit t chqiunagr-tmero-cmhoenrd t Nw wuirneg mch oprder cferonmta gwe inogf
point. leading edge.
6 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
Jl wing pitching moment about the quarter
(c. p.)r=( - Chr) X I 00, flap section center of pres- lOcl4, chord point.
cILr sure in percentage of flap
cbord from hinge. till (I,lP section binge moment per unit span
about the flap hinge.
CO. P.)r=( - ~Ir) Xl 00, flnp center of pressure in flll flap hinge moment about the flap hinge.
Nr percentage of flap chord from Jill' 'IL " moment of wing emispan load about
hinge. \\'ing root.
=( JI/'ILI' moment of [lap semi pan load about in
CO. P,)w ~~{Wt) Xl 00, lI"il~g l(ltern.l center of pl:es- board end of flap.
1at ~ sure 111 percentage of II"mg {f, dynamic pressure.
4 emispan from \\'ing root. wing chord.
fH ("w,
+-
=1_I X~ degrees LI -9' ~ r4 + __ 3° Ls I I I 60 -I- --r---- IX. deIgIre es ,.....
=t--.::-1I2S 0 I [I>7. -I6 ° I q __ =;T: o--3-I° _ o__~_ x lIf--- D 90 ----+t -- L<:Jl I I 1152 °o ff---
I I I
-60 I
I I I
-40 I I I '- /" ""
1 1 1 '\
-20 I Of = 0° 0-- 15° \
u: I I I I TT I
i I 1
20 I I I I I I I I I
I I I i I I I I -'- I I
-1-
40 ~ ~ < /7" -..,..".~
~ 60 I II . ~1'\\ I r-i- ,,~
--,-
-S 80 1I II I ~- I I '"
~
.~ 100 I I , \ , i'-f-... I
~
~ 120 -; Ib
\J I 1\
~ 140 I I \' r\ \
Q. -40 I \
-~S~:i.. -L2.E0 I I, II I1 0/=300 \ I = TJI S : '\ , 0-=1450
I I
20 i \ /" 'u I \
I ~- I
40 ,y I I ,
I
I I
60 , I I /I " i- r-.:-- ""~ I i- 1--". ... ,~
I -r--- --.I .. " ;'....
80 T I ,..-. ,
- I I I """- I r-....
100 I I ~I '"
" I 'I-
f20t c r ED c B A ~ c r ED c B A
Roof Wing semispon Tip Root Wlnq semispon Tip
FIGlJltE 9.-Airfoil secticn longitudinal centHs of pressuro against span of wing with tbe flap deQeded \'arious amounts.
+':;b')XIOO'
(C. P,)rt"t=( fla~ lateral center of pres- c" flap chord .
.L.l sure In percentage of flap b'D' wing span.
4 semispan from inboard end
b" flap span,
of flap,
where In these coefficient it is to be noted that the chord
nw IS the airfoil section normal force per unit forces on the airfoil have been neglected; i. e" the
span. longitudinal center-of-pres lire positions and the pitch
N,&) \\'ing normal forcE'. ing-mom nt coefficients were derive 1 solely from con
nIl {lap section normal force pE'r unit pan, sidrrations of the normal forces.
Nil flap normal force, The airfoil section normal-force coefficient are plot
In , airfoil section pitching moment per unit ted against the wing semispan in figure for all
wcl4
span about the qllartE'r-chord point. nap deRections and all angles of attack te ted, The