Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a tapered wing to determine the rolling and yawing moments due to an angular velocity in yaw. The model was tested at 0 degree and 5 degrees pitch, -1 degree and 5 degree yaw, and with a full-span flap deflected 60 degrees. The results are given in the form of span load distributions and in calculated moment coefficients. The rolling-moment coefficients are in fairly close agreement with those derived by means of a simple approximate theory even for high deflection of the full-span flap

Rokaus, F

Troller, TH

English

NASA Technical Reports Server

.. 
'fECnNICAL No 'rES 
NArIONAL ADVISORY COM~ITrEE FOR AERO PAUTICS 
No . 683 
PRESSURE-DISTRIBUTION MEAS UR E1ENTS ON A 1A?ERED WING 
WIr3 A FULL-SPAN SPLIT FLAP IN CURVED FLIG~T 
By Th. rrol l er ani F. Rokus 
Daniel Guggenheim Airs h ip I n stitute 
THIS DOCUMENT ON LOAN FIfOM THE FILES OF 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
LANGLEY ~ERONAUTICAL LABORATORY 
LANGLEY FIELD, HAMPTON, VIRGINIA 
_ft£TURN TO THE AiOVE ADDRESS. 
ft[,U[STS FOR PUaL/CATII NS SHOULD BE ADDR£'SI&t 
AS FOLLIWS 
. j a s 1 i n g ton 
January 1939 
NATIONAL ADYISQfty CO~MIITEE r o " A~R(VH,lTncs 
1724 I STRfET, f\. '/I. , 
v A<;fj :CTOri 25 D.C . 
https://ntrs.nasa.gov/search.jsp?R=19930081473 2020-06-17T23:00:46+00:00Z
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
TECHNICAL NOTE NO. 683 
PRESSURE- DISTRIBUTION MEASUREMENTS ON A TAPERED NING 
WITH A FULL - SPAN SPLIT FLAP IN CURVED FLIGHT 
By Th. Troller and F. Rokus 
SUMM.ARY 
Pressure-distribution tests were made on the 32-foot 
whirlin g arm of the Daniel Gugg enheim Airship Institute 
of a ta p ered wing to determine the rolling and yawing mo -
ments due to an an g ular velocit y in yaw . The model was 
t e sted at 0 ° and 5° p itch, _10 and 50 yaw, and with a 
full - span flap deflected 60 ° . The results are g iven in 
t n e form of s p an load distributio n s and in calculated mo-
me nt coefficients. The rollin g - mo ment coefficients are 
in fairly close a g ree ment with t h ose derived by means of 
a si mple app roxi mative t h eor y even for hi g h deflection of 
th e full-span fla p . 
INTRODUCTIO N 
An investi ga tion of the pres sure distribution over a 
t ap ered win g model, with a full-sp an s p lit f lap while in 
curvei fli ght, was made on the 3 2-foot whirlin g arm in or-
i e r to obtain the aerod yna mic mo ~en ts for such a fli g ht 
50 n dition. The followin g win g p ositions were investi gated: 
mo de l at 0° a n d 5° p itoh, eac h with _1 0 yaw and 0 0 and 6 0 0 
fl a p defleotions; model at 0° and 50 p itoh, each with 50 
yaw and 0 0 and 60 0 flap deflectio n s. Betwe e n 200 and 4 00 
pressure points d istribut8d over 12 chords along the s p an 
were me a sured for each wing p osition. For t h ese tests th e 
b l R ratio (sp an to t u rning ra d i u s ) as equal to 0 . 1 3 3 . 
APPARA TUS 
The win g mo d el used in t h ese tests was built up from 
several s e ctions of ma gnesium, wnich were bolted together , 
illac h ined to the correct size and shape, and then filed to 
2 N.1\, C; 4 . ,· Techni '~al N-ote No. 68 3 
a s mooth finish . owtng ,to ' the li m~' ted s,'1.ze of the sus p en-
sion tube, only 35 p oint s could be measured at one time. 
The win g was therefore madi h~1 16~ with the lower surface 
remo v·able so t h a:t .t h e rubber ' tubes qonnect 'ing the pressure 
orifices to the measur ing mano meter could be chan ~ ed to 
different position's'''at .the end o f : each· t ·e ,st run. The wing 
model was tapered 2 :1 with t h e 30-per6ent-~hord points on 
a straight line a·nd the. maxi mum ordina.t ,es pf all sections 
in a horizonta l plane on th e up pe r surfac e . Th e w ing mod-
el had a Clark Y profile, a maximum chord length of 8.91 
inches, and a span of 5 0 . 45 inches (fig . 1 ). 
Th e flaps used for these tests were tap ered so that 
the chord of the flap at any lo ngitudinal section was 25 
p ercent of the win g chord at the s~me section . They were 
made of 1/1 6-inch ~teel plate and were hin g ed to the wing . 
The flaps , when de,flected"w~re held in position by SDver-
al sma ll rods between the flaps ,and· .the wing at th.eir ex-
t re me ends (fig .. 2) . Pressur,e orifices were installed on 
only the outer s ~rface of the f laps. .FOT the 0 0 flap Gon-
dition , the flap was ent irel y removed and .the lower sur-
face of the wing was s mooth~d with wax . . 
The inverted wing mo d el was attached to the whirling 
arm (referen c e 1 ) as shown in fi gures 1, 2 , and 3. The 
s mall streamline tube attached to the center of the wing 
has a 2-inch chord and a m13:~imum : t h ickness of 0 . 75 inch. 
Th rough t h is streamline tube, 35 pressure tubes were led, 
con n ecti ng the various pressure ori f ices on the wing to 
the multip le manometer lo cated on ' the cen.tral shaf t of the 
wh irling arm: Here ,t h e mano meter tube~ ~~rned with the 
shaft at a radiUs of 8 . 25 inches . Light-sensitive paper 
was ins e r ted . bet wee nth e ma no m e t e r tub e s ... a n d " t h e sh aft and 
was exposed to . the ~ight of four · ~hotof loo d bulbs .during . 
the test runs . . A record of the hGli ghts of the meniscuses 
of the alcohol ,i.n ·the manor,netertuoes wa~ . thus obtained . 
For these tests it .was necessary to reduce the swirl 
in the whir li ng-arm channel (reference 1) even more than 
had been done for previous ·whir li n g -arm tests. In order 
to reduce the swirl, the thr e e scr e en baffles, which are 
p laced across the whir ling-ar m cha nnel a p proximately 1 20 0 
a part an d which h ad been used to .reduce the swirl for pre -
vious tests , we r e entire l y covered with heavy wrapping pa-
p er . A rectangular op ening of a pp r o xi mately 20 square 
feet was tftt in these sol i d baffl e s for t h e wing model to 
. pass thr ough . This opening , was of. suf f icient size to avoid 
interference with the air f low about t h e wing . 
N.A . C.A'. ' Technic~ l Note No . 683 3 
The p res sures ,on t he win g we re measured b y recotding 
the pressure f ro m 35 points at a t im e . Betwe'enea c h test 
run, t h e r ubbe r tubes co nn ecting the pressure orifices to 
the man o meter we re ' chang ed unti l the whole surface of the 
mode l h ad been surveyed . 
Ea c h record wa s obtained b y accelerating the mode l 
slowly up to fu ll speed and expos ing t he ' se~sitiz~d pape r 
for about 2-1 /2 minutes afte r t he model had st eadie d it-
self at the tes t · spoed of 45 r . p . m. The ph oto g rap h ic ma -
nometer r ecords 'Nero taken on "ozalid" paper, ' a paper of 
l ow li ght sensitivity that i ·s deve loped , afts'r exposu re, 
in ammon i a fumes . I na s mu ch as the paper is not brought 
i nto contact with mo i stur e , no s hrinkag e or warping occurs . 
Severa l tubes on t he mu l tip l e man o me t e r ' we re op ened 
directly to the at mosphe re to be us ed as zero reference 
points . I n order to ob~ain the absolute pressure at any 
po i nt , a centrifugal-force correction, due to the length 
of the column o f ai r fro m the ~an o me ter tubes to the p r~ s­
sur e orifices t~rning with the arm , had to be added to the 
readings . This correction can be c a lculated very accur -
ate l y . It is equal to the veloc it y head of the turning 
speed of the ind i vidual ori fice except for the ne g ligible 
value of an air · column f r om the center of the shaft to 
t he man bmeter. 
In order to obtain the flight speed of the mod el , 
t hat is, the speed against the air , a correction had to b e 
made for the swirl due to the motion of the arm . A hot -
wi r e unit, suitable f or mea sure ment of the smal l fluctuat -
i ng v e locity and its direct io n , was used f or the investi -
gat ion of the ' swirl . Many po ints with in the channel we re 
surveyed and the average swirl velocity , with the mode l 
o f t he tapered win g attached to the ar " was f ound to be 
0 . 8 mete r per second. This value is 1.3 6 percent of the 
mode l velocit y . Th e swirl-velocit y co mp onents . normal to 
the fli ght direct io n we r e found to be so s ma ll as to have 
p r a ctically no effect on t he p i tch and tho yaw angles of 
the airfoil. The fol Io ing table g iv~s values for the 
swirl velocit y and the effective angular change of the 
airfoil for throe p oint s wit h in the channe l just below the 
wing . 
4 R . A .C. A . Te c hn ical Note No . 68 3 
Position be lo w air f oil / Swirl ve loci tylAngular c h ange due to 
. I ( meters / sec) ~ " swir l 
I ·Pitc'h 1 Ya w · 
(de 'g ~) "_~_ 
6 inc h es beloi'T : 
i n nerti,p 
centeT 
,outer , tip 
1 . 0 
. 8 
. 7 
- 0 .;05 
o 
. 0 7 
- 0 . 03 
o 
. 07 
The angular chan g e of pitch a nd y a w du e to the sw irl com-
p oneh t is obtained from 
, -1 v s in a 
SIn 
v 
whe r e v,is avera g e s wirl ve l ocity , as me a sured on t he os~ 
cillo g ra rr: . 
a , av e ra ge a ng l e of p i tch or yaw of t he swirl ve-
l ocity , a s ~eaBured on the oscil lo g ra m. 
V , mod e l velocity , 44 , 5 me ter s p er second .. 
Fi gure 3 s hows a ph ot ogr aph of t he h ot - wire mou nting used 
in maki n ~ t h e s wi rl mea s u r e ments and fi g u re 4 shows t h ree 
t yp i ca l oscillo g r a ms of s i rl veloci ty , p i tc1 , and yaw ob -
tai n ed for one of th e p osi t ions i n v e sti ga t e d . 
The ve l ocit y o f t he l ing us ed through ou t the computa -
tions i n this r ep ort wa s taken as the velo ci ty of a p oint 
at t h e inte rs ect i on of the ce n ter line of t he span and a 
li ne p ass ing t h rough t he 30 - pe r cent po i nt of the win g 
chord . Th is velocit y wa s f ound to be 4 4 . 5 mete r s per sec -
o nd , aft e r de ducti n ) t he s wirl veloci ty , a nd as used i n 
dete r mi nat ion of q o ' which was u sed fo r t he p/q compu-
ta t ion s . 0 
TEST RES ULTS 
Typ ica l diag r ams of the p ressur e dist ri bu ti ons over 
t he uppe r and lo we r s u rf ac e s of t he airfoi l and t he f l ap 
are g iv e n i n fi gure 5 as r at i os of the static p r e ssure P 
1 
J 
N.A.C.A. Technica l Note No . 6 83 5 
to the dynamic pressure qo of t he tr ue fli gh t velocity. 
The p ressure diag ra ms for t h e various wi ng p o sit ions 
were g r a ph ically in te g rated an d from the data thus o b -
tained nondi mensional coeffici e nts we r e comp uted for the 
win g as a wh ole . 
Th e l~ca l normal - force coefficient 
c h ord-f o rc e coefficient Cc are plotted 
ing curves of fi gure 0 , 
n 
c 
c 
n 
in 
and the 1 a c'a l 
the l oad g rad -
I n fi g u re 7 are g iven t h e rollin g - mo ment coefficient CL 
and the yawin g - mo ment coeffici e nt Cn o b tained .by g r aph i-
b /2 
ca l integ r a tion J F y d y 
-b / 2 
<There 
and n , nor ma l forc e per ~nit span alon g lon g itudinal ax i s 
of \7 irJ.g . 
c , lo ng itud ina l force p er u n it span . 
clocal' c h ord of a i rf o il sect i on . 
S, total win g area . 
b, wing s p an . 
y , distance alon g span . 
Fi g ure 7 al s o g iv e s a co mpariso n of t h e expe ri men t a l 
rollin g - mome n t c oefficients with those obtained from 
l ie se l sber g er 1s t h eor y for an elliptical ~in g in curved 
flight at 0 0 yaw. (See r eference 2.) Th ese values of 
6 N,;A . C, ~ A .- ,Tec1mical Not ,e N,D, -: 6.83: : 
9i ' ' are , base~ bn' t h e" t h eoretica l moment ~ cbtainBd fro m the 
Lw b
2
' (~' +:' 1) 
, 1 6'R " 
wh e re Lw is tota l ,lift taken fr~m the te$ts . 
b , span of the wing , 1.284 met ,e.!'? -. 
R , turning radius of center of wing , 9 . 5 9 meters . 
~ , curved vortex induction faht 6~ ; 0 . 73 (reference 
2) . 
rh e theo retical values are those of an , el li ptica l win g of 
a very similar l oad g r ading in strai ght fii g ht to that of 
tho ha l f - taper ed wing used in this test . , " ,' 
Fi g ure 8 shows a co mpa rison of the theoretical load 
distributi o n a l on the s p an of the win g with the experi-
~en tal p oints as obt a in ed for the _ 10 yaw and : 50 pitch 
wi n g pos it i on (negative yaw denotes turn in g of win g so 
that oute r ti p is moved rea r ward), 
T_ e theoretical curved fl i gh t load distribution was 
c0 2puted f r om the strai ght - fli gh t g r ad i en t for a tapered 
wing ( re f ere nc e 3) by mu l ti p l y ing each s e c~i?nal value by 
(1 + §y '\ 
, R J wile r e is t h e distance of the sec -
ti~n ' frd ~ thec ~ nt~ r 6f t h e wi ng . 
Th e fo ll ow in g taol e g ives a r~sum~ of t ~ e resultant 
r ol li ng - and yawin g - moDen t coefficients arid t he ratios of 
L / Lwb and N/ Db for all the ,win e p osi t ions i nve s ti gated 
i n this test , where D is win g drag. 
I 
~~_J 
l,- __ ~ ___ _ 
N. A.C. A . r ~chn i cal Note -No. 683 7 
co nd i ti on ! 
T 
! Ra t i 0 s ','Vin g I Coeffi c ion ts I F lap 1- ·------ - t- L/Lvrb - ----- -r-:-:----Ya w F i t c h de flec t i on 1 C t Cn I lJ !Db (d e g . ) (deg .) _ ( dag .) I I 
I 
- 1 0 0 ./° 0 0076 I . 0003 3 0 . 001 99 I 0 . 0299 60 . . 02 2 9',1 . 005 1 4 . 00 1 58 , 0207 
0 
I 
. 0 1 36 . 0007 3 I . 001 66 I , 0290 -1 5 
I 60 , 0282 - . 00693 i . 0 0 1 54 . 024 1 I I I -
5 0 0 I , 0073 . 000 33 I . 00 1 99 . 0299 60 ! . 0225 . 00832 I . 001 6 2 . 0339 
I 
i 
5 5 0 . 0 13 1 . 000 73 I . 001 66 . 0290 
60_.J_ . . 027 9 , 0 10 2 9 i . 00 156 . 0359 I 
CO NCLU SI O ... S 
1 . The ro lli n t:; - woL,ent c o effi c ients .. oasurod at the 
wh irling ar~ a r c in fair l y closo a ree ~ent with t hos e d e-
r i v ed by ::leans of a sLlp l e approxi..at i ve theo r y even fo r 
h i gh d e f l ecti on of the f ul l -span flap . 
2 . The ro1lin b DO C O lts are li tt l e affocted by tho 
a ng l e o f yaw with i n the ran ~ e of the to ts . 
3 . Tho ratio L/Lwb decreases with t h o l owerin g of 
t h e f l a p b y about 1 6 pe r cent fo r tho 0 0 pitch setting of 
the wi n g and is a l J: o s t c ODsta t at 50 pitch . 
4 . The r a tio ~/Db dec r eas e s with th l owe r ing of 
the flap . 
Dan i e l Guggenheim Ai r shi p I n stitute , 
Ak r on , O'h i o , -fov e :-:.lb e r 1 93 8 . 
8 
1 . 
N.A . C . A . Technical Not e No . 083 
T roll e r, Til .: 
vol . I, no . 
REFERE NCES 
Th e ~e w Whirling Arm . Jour . Aero . Sci., 
4 , Oct. 1 9 34 , pp . 1 95 - 1 9 7 . 
2 . Wieselsber g er , C ' Zur Theori'e des Tragflfigels bei 
gekr ~m~ter Flug bahn . Z . f . a . M. M., Ed . 2, Heft . 5 , 
Oct , 1 922 , S . , 325 - 340 , 
3 , Glauert '" S .: The El ement s of Aerofoi l and Ai r screw 
Theor y . Ca mbridge Un i ve r sity Press , 1 9 26 , p . 1 55 . 
. - ~ .-------------------
1--- - - - - --- - - - - - - - - - - - - -i- - - - - - - - - - - - - - -~ -fl8.lh---___ --,,-
Flap axil K 25~ of chord length 
I 
I 
I 
I 
36" 
30% chord line is 
Note: 6 rows of pressure orifioe. are 
installed on both inner and 
outer half of span. 
str~ight : _ l~~ 
~ SO.45" ~ 
Ohord. Distance from oenter of spaa, in. 
1 2.63 
2 6.63 
3 14.88 
4 21 •. 38 
5 23.32 
6 .... a X---;r. 
~ ~ 
Figure 1.- Diagrammatio drawing showing inverted wing suspended on end of whirling arm. 
3 "x ~n , 
-6~1 
I 
I 
I 
I 
I 
I 
I 
I 
streamline tube 
to: 
.. 
o 
.. 
>-3 
CI 
o 
£ 
o 
~ 
gc 
~ 
CI 
gs 
CIl 
<Xl 
Co> 
... 
... 
~ 
.... 
N.A.C.A. Technical Note Bo. 683 Flgl. 2,3 
Figure 2.- Photograph o~ trailing edge of wing Ihowing flap deflected. 
Figure 3.- Photograph of wing on end of wh1rling arm, Ihowing Aot-wire 
mounting used for making 8wlrl meaBurementB in whirling-
arm channel. 
".------- -~- - .~--.- ---
--- ---- ---------
E9 DENOTES POINT 10FT. 8eFORE MOGEL REACHES HOT-WIflE UNIT 
20M/5 L:jjJ;J 1.5 Mis 
1.0 MiS WI,'\)' I 
P 5 M 'AS---"..l.. 
.,. lJENorE,s /NW;tA'O (rcw;t,f'O t#'Nrl,fl ""- RCrAT' / ()N) 
- De-NO rES O,-,rWA~/) . -110. _________ _ r19W i9NCLE O r ..fW/RL YELOc/TY 
P lr<:'1f A NGL.1f + f)EN"71f~ /JEN6r,s 
rigure 4.- Typioal .w1rl-meaaurement o8oillograma. 
1 
• ~ 
a 
~ 
~ 
II 
0 [ 
() 
~ 
• 0 
.. 
.. 
:;c 
CJ> 
flO 
c.o 
... 
... 
~ 
~ 
..l. 
Clo 
Outer .pan 
-2.0 Ohord 1 Ohord 2 Ohord 4 Chord 5 aherd 6 
-1.0 
°t ~~ I ,0 ,~~ :=)l 1 _" ~ Ij~ Ir;:;>x I 
- 2 .0 
-1.0 
°1 ~ >1 1(, 
1.0 
o Upper surface 
x Lowor eurfaoo 
Inner .pan 
Pressur e d1atribution along 
various ohorda 
~ I~ u __ 1 ~ 1.....- " 4 fro 
Figure 5&.- Typical pressure-d1.tr1but10n ourv •• for 50 pitch, _10 yaw, 00 flap deflection. 
~~I 
... 
~ 
;, 
!'" 
.. 
3 [ 
! 
... 
o 
.. 
co 
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CD 
'" 
... 
... 
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i' 
Outer span 
Ohord 1 Ohord 3 Ohord 4 Chord 5 Chord 8 
-1.0 
o I ! I ! I ! I I I I I I I 
L 
Qo 
1.0 
-1,0 
Inner spe.n 
o Upper eurface 
x Lower surface 
Pressure d1str1but10n along 
various chorda 
F1gure 5b . - Typ1ca1 
pressure-
distribution CUrTe. 
for 50 pitch, _10 yaw, 
600 flap defleotion. 
01 I 1 1 1 I 1 I 1 1 1 I 1 
1 .0 
'" 
~ 
o 
~ 
... g 
~ 
'" o .. 
.. 
1:' 
'" 
'" toO 
... 
.. 
~ 
gJ 
l 
~ I I { 50 p1 tch I I I I I I _ _ _ 1  yaw 
' _--c-o J J
C ~ I 1 ,/ 60' flap , , , , I ,....-- ~ - -L-..... I--:-:.~_ 'Y. 
d 2.00 1 I Ix'l \ { 5° pitch r--:--t- .... 1---" ~~5 { 00 p1tch ~ ...l 
o LV 5' ya. ~":--+--- ~ l' ya. """~ l 1 1 J 
... ~ /;' / 60°1 flap . ~r-- " soo nap 4'\ " 
" / ., , , , .• I I I 
' " --~ -" r " ~ 1.60 'I --::::~ 'ff-<>--d I -- e- 1 I t-"'~ o~ ~ L , { OoP1tch { 50 P1t ch C I~ I \\ ...l • 5 yaw , ""II' 
o '/ • 0' n ap ,I" o 1/ 60
0 
flap . -1 ya. \ \I I I J 
• 1.20 ' J~ I : I I \J I 
o /1 , " \\1 , , ~ 1-----+" l ' ~ ~," 
': r ~r-"'" -- _ ,- 0' pit,h h I. " 
.... ;:::::.- \ { 50 pi toh ,{-lOyall' '~l\. \' ! .80 k;:/ , 50 ya. " 0' flap +-- - u.: 
2 J. 0', fl.,. .. : ' - ~ ~ I I 
1 40 ~~- \{~~ ~~!Ch I I I I I I I I I I I I I I I I =rt 
Q 
o 
11 
CD 
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o 
~ 
.... 
.... 
CD 
o 
o 
CD 
Q 
,.. 
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.... 
I 
~ 
.8 
o 
';1 
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o 
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!= 
/ 00 flap I . 
01 I II I I I I I I I I I I I I I I I I I I I I I I I I I ' 1 I I 
I 
Outer half Inner half 
.40 1 I I I I I I I I I I I I I I I I I I I I I I I I I I I 
~b.. "0~HMlf£"t44J.~ ,m:h __
{ 
00 pitch 
" 10 yaw 
~ 
n 
1 1 1 1 1 1 ' I'{-~ 
" 60
0 ~~:h~r±::R1 I I 
\1, 
-+x t::Q-+= 
" 
.20 I I II I 600 flap 
'1 I I I \ l{ 00 50 pitch I I 00 p1 tcb ' I \ 50 .lO~--=tt=1 I I I I ~'{_lO and 50 yaw ffi, ClO and 50 yaw 600 
I 00 flap lOO flap I 
pitch 
Y8/ gr 
flap I I 
/ . , _L I I I . J 
0 1 I ~ ~ I I ~ I ;1 ~=&I IiI:: I • I I I i I I I Be~ 1 I I 
.60 .50 .40 .30 .20 .10 0 .10 .20 .30 .40 . 50 .SO .80 
Oistance from t of wing, meters 
.80 
Figure 6.- Distribution of normal-force coeffioi ents and longitudinal-force coeff1cient s along span of wing. 
z 
~ 
o 
~ 
"'3 
m 
() 
a 
() 
~ 
z 
o 
ri-
m 
Z 
o 
()) 
CD 
t.l 
"'J 
.... 
IN 
()) 
N.A. C. A. Tecnnical Note r,o . 6 83 Fig . ? 
I
Ul 
H ,o ~ . 020 
. 040 
ROLingIJt ~oLi f:nt.s I 
-;;:--= ~ 11-- 1_;- '-----f- - r~ - -I I ,, --, -- --I I :A- .,)'--- , - ' --1- - --L -;;:::- ', --<' - ' ~ - I ~---"", -'~---'-r-- '-~~ - I \ '0 ' 
- _ -- -- -- \ -1 rI'aw f 60° l ap 
__ I----"~ _ - \- 5° II 
,::~ 
. 030 
II 
c--
~ f-'-_ 
---::::: I- -
(.) o , r--;::::::::::: 
--,-1 yaw laO flap ~,¥ _ -
/ 5° I II ( ~ -- - I 
, I ) .--;::::::;--
I ~-,;..- __ I 
" -I 
I ' - 'i. r- --~ - - -- ~_ - - - Wi es el sberger IS coeffici ent f or 
:::::::--:.. 
-
- ellip tical wing a t 0° yaw. 
--
. 010 
o 
I , 
. 020 
I 
. 016 I 
I 
Yawtng-,6men+ co effici entQ 
___ 0----
. 008 
: ____ f-' 
- 1""'5° yaw 60° flap 
> --
~ _1 ° 11 
- , 
, 
k:--: 
L.J:--x--
r---X---
~ 
. 004 
,{ 5° yaw 0° l ap i ' -1° 11 
, 
/ 
I~ " , 
o ~.'. I 
o 
Angl e of p itch in degr ees. 
Figure 7 .- Rolling-and yawing- moment co efficients. 
j 
r 
!-I t+ :~--~~c-_1_-r-I- -- --~1 ~ I -i-ti -I I l------i 
I t---~- I! I , i J! I ------i 
t
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Pressure-distribution Measurements on a Tapered Wing with a Full-span Split Flap in Curved Flight

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