Daniel Guggenheim Airship Inatitute
Akron, Ohio
PRESSURE DISTRIBUTIOI~ • SUR 'El.TS ON A TAPERED /ING
tlTH A PARTIAL SPAN SPLIT PLAP II CURVED FLIGHT
In Fulfillment of First Half or Contract NAw616
Introduction
An inveati~ation of the pressure distribution over a Cl rk Y tapered
inc model ith rtial .span split flap a mo.de on the 32foot irling
arm in ordor to determine the rolling and yawing coment coeffioient for
the follorinn oonditiona: I!lodel nt o0 and s0 pitoh, enoh with o' :t s0
, and
±4 0 y w, e oh 1th a 26 °/o, 50 °/o, and 75 °/o span flap dofleotod ~o0 •
For the 6° pitoh oondition tests ere also de on the model ithout i l~pa
and th n full span split flap for all the yaw angles. The result are
given in tho form of spo.n load distributions and in oalouloted moment ooefficients.
For these teat tho ratio b/R (spon to turning nldius) s equal
to 0.133.
 pp:ratua
The tapered ing model for these tests s tho s ms model as uaed for
~he teata described in Ref. 1. However, in order to adapt this modal for the
present tests a few modificationo were nece11ary. Twelve additional chordviee
rows of pressure orifioes, bringing the totnl number or rows to 26, were 1ns~~r£
along the epo.n or the wing. The ne;: rows were loo ted near the transition
pointD of the different flap lengths. Figuro 1 is aketoh of tho wing,
shouing the ohord1 alonG which the pre &ure orificea were located.
The flaps usod for these tests woro made from the fUll paD flap alao
described in Ref. 1. Both the inner and outer halves of this full span flap
ere cut into four equal lengths, each being 25 °/o of half tho total epan.

Daniel Guggenheim Airship Institute
Akron, Ohio
2
These quarter seotiona wero then meroly ndded to or removed from the model to
produoe the desired flap lengtha25 O/o, 60 °/o, or 75 °/o of the 1pe.n.
A complete deaoription of the teet setup and the teat prooedure oan be
found in Re!'. 1.
Test Result•
Typioal diagrams or the pressure diatributione over the uppor and lower
aurtaooa or tho airfoil for different flap lengths are given in Figures 2a,
2b, and 2c aa ratios of the statio pressure P to the dynemio proesure q0
of the true flight velocity.
The preaaure diagro.ms for the various wing positions were Braphioally
inteerAted and from the d ta thus obtained nondimeneioncl ooeffioienta were
computed for the wing as a whole.
The looal normal foro& ooefficient Cn and the local ohord force ooetficiont
C0 for 5o pitoh and o0 , ~5°, +10° yaw are plotted in tho load grading
curves of Figures 3 and 3b, ~ere:
n
In Table I is given a resume of the rolling moment coefficient ~1 .
and the yawine moment coefficient Cn obtained by graphical integration
• ) '2 f ~ dy where
 /2
\; • L
7 q0 b S
en • ?~
q0 b S
and n a nornal foroe per unit span along longitudiDal axis of vring
0 • loDgitudinal foroe par unit span
olooal • ohord or airfoil eeotion
I
Daniel Guggenheim Airship Institute
Akron. Ohio
b ., tt1. ng 1 pan
y ., dietanee along span.
Figure 4 zive~ a comparison bo een the exporimontal and thooretioal
rolling and ye.wing moment ooerfici nt for the o0 nd s0 pitoh an~ o0 yaw
positionc of the modol. The theoretical oooffioionte were computed aooording
to Ref. 2.
The theoretical rolling mo~onta nre defined bye
L n • rb . q.S·b
.t 2 v
where: L ~ rolling moment
r ~ turning 1peed in ndians per seo
b • s n of' wing • 1.284 m
S 10 winB re o .215 m2
V u flight v looity
q • velocity head
G rolling moment coefficient o tained from Fig. 11, Ror. 2, for a
wing taper of 0.5 and an aGpeot r tio of 1.1. For the Clor~ Y
section or the vinb r.i.thout fl p, the anglA:!br zero lift 1
~D :11'il8d to be 4.8 • For the condit~v ~0 pitch thie siveas
d. · 4.8°; for the co11dition s0 pitohi oL  9.8°. Tho €:>0 defleoted
l~J 1f 25°/oohord Wll6 Q awned to be qlV le L wO an ircre 88~
L:!ol 9.6° or the effeotive angle of attao1' oJ. .
'lho theoretic 1 yawing n:omen ~ f2
def· (e"l by )'"
"I·' C .,. , s b J •"O + J J.iD .• •4'' + t" j •  r ·V .. o R
0
Y·dy
here: H
• 2
c l'A i1ng moment
yo.wing aomont coefficient obtained from Figs . 12 ~nd 13 of Ref. 2
waa assumed to be O.Ol for the wing without flap. For the wing
~~ flap it wus taken from Fig . 4, Ref. 3 , that Cd• 0.26 at
°" • o0 • i th cd1 0.08 at CL • 1.3 for 0° angle of' ttaok ot
t;,e r ferenoe wing 1th an aepeot ratio or 6.1, t his eivea
odo 0.20 tor the Clerk Y ng with a 26 °/o chord aplit flap
deflected so0 •

Conolusi.ons
Daniel Guggenheim AirGh1p lnet i t ute
Al.Ton, Ohio
l. Tho rolling moment coefficients for both o0 nd s0 pitoh are in olo e
agroeDent th the theoretical onoa given in Bet . 2. exoept for tho 100 °/o
Spln flap, ere t.~e mea_ured coefficionta aro lnrger.
2. The oxperinental yQwing moment coef!'ioionts for o0 pitch are in very
cloee agreement •. i th the theoretical once r;ivel'l in Ref. 2, whilo thoae fort
6° pit~h nro in cloe gree~ent ror ~lap l enstho up to 50 °/o or the spa~
but are somewbat larger at 75 ~lo nnd 100 °/o 1 n flap longtha.
s. The roll1n~ moment ooeff1cients are little affected by the an~le of ya ,
thin the rcnge teGted of !: 10°, mth n tondenoy to dooreaee as the model
ie ~iaued in the poeitivo dirootion. Tho ye.wing moment coefficients re
little sffeoted for flep lengths up to 50 °/o of the span, but show oona1derable
obange for the full span flap, ith the tendency to inorcaee ne the model
is yawed in the poeitive direction. (Positive denotes turuin of tho model
an t hc.t tha leading ed e f eec tmard the oonter of' rotatio.:i}.
~.. Th .... oxtont oJ' the flap ffecte tht. pressure distribut ion over the entire
paD, onu ing n i~cro ee in the xiimlm value of the P/q for any chord
eeotion as the .flnp length is i1loreased •
• R.
Jun , 1939
REF RENCES
Daniel Guggenheim Airship Institute
Akron, Ohio
1. Troller, Th. and Rokus. F.: Preseure Distribution ee.aurementa on
Tapered DS with a FullSpan Split rlap in Curved lieht.
l .• A.C .A. Technical liote IJo. 683, 1939.
2. Pearson , Henry A. and Jones, Robert T. : Theoretic 1 Stability and
ontrol Chnraoteristios or ing with Various Amounts of Taper and
Tw st, h . A.C.A. Teohnioel Report o. 636, 1938.
s. cnEingur, Carl J.: ~he r.:rr ot of FullSpnn and ertial S n Split
.1 pa o the crodynD.Itl.c Ch otorictics of a Taporod ing ,
.A.C.L. Teohnio 1 c ¥o to. 505, 1934.
FIGURES
Daniel Guggenheim Airship Inetitute
Akron, Ohio
1. Sketch or win showing location of preaaure orifioe chords.
2a. Pressure dietribution ourvea for s0 pitch, OO yaw, 26 0/o span flap
d fleeted ao0 •
I
ressure distribution curves for 5° pitoh, 00 yaw, 50 °/o span flap
defleoted 60°.
2b.
2c. ree ur distribution curves for 6° pitch, o0 yaw, 75 °/o spnn flap
d fleeted 00°.
3a. Curves how1ng distribution of normll force ooeffioionte nlonr; •pan
or ng.
3b. Curves eho n distribution of longitudinal foroe coefficients along
span of wing.
4. Curve bowing oomparison of theoretiaal ad cxpor ent l rolling and
yawing moment cooffioienta.
•  . .. t ...
• I I
~ ! _l __.__._,
______ l_J__
~     ____._ _ !
"" ~
~ J e+~+++1~~++~+i_.
r _I ~· j 1,;f T<flV
'::IC.U ~ t; I ii A'.+...;! O~;...__~~...ill"L..;.<LaoLT~....._,.......J 4L.I ........
r
r ROlLJN6~HOH~N1:
 I I I
'
,
 4 
The t heoretical rollingmoment and yawingmoment coefficients were
computed according to the method of ref erenoe 2. The theoretical rolling
moment L is defined by
L• j !:kqsb
r 2V
and the theoretical yawing moment N
N = ''nr ¥v q S b + qj t/2
t/2
is defined by
R + y y dy
R
vmere
t is the rollingmoment coefficient obtained
.i: is the turning speed, radians per second.
v is the flight velooity.
q is the dynamio preeauro.
Cn 1• 6 the ya ingmoment coefficient obtained r
Odo is the looal profiledra~ coefficient.
R l• S the turning radius o.t oenter of vr\11g.
y iE the diste.noe alonu spti n.
from reference 2.
•
from reference 2.
For the C la1~k Y wing i thout f'le.p, the an le of ttaok for zero lift 1
0 assumed to be 4.8 • 1'or the condition of o0 pitch, thie value 0
ive~ °" ~ 4.8
and, for tho ondition of 6° pitch, 9.8°. The 25 percent chord flap defleoted
600 :as ssumed to be equivalent to an increase of the effeoti vc an, le oi' et tack
equal t.:> U • 14°. Thie effect of the f lap ms aseumed on the basis of figure
4 of ref erenoc 3.
 5 
The profiledra coefficient was esumed to be 0.01 for the ~dng without
flap. For the win vtith fl p. th value of Cn . 1 t ken from figure 4 of refernee
5 0 .28 1 cj._e o0 • 1ith Cd1 . o.os at CL . 1. 3 for' ro angle or
att ck of the reforenoe nng with an aspect ratio of 6.1. a value of Odo• 0.20
i obtained for the Cl rk Y wing with 25peroent ohord split flap deflected
60°.
Figure 5 giveo a oomparieon between the experimental and the theoretical
rollingmoment and ya~~ngmomont ooeffioicnt1 for the o0 yaw position and the
position of o0 and 6° pitoh.    .  ~·  .  '  . 
  .  . . .    '. ,    ._  .  .  ~  . .   ·  . . . . 
   ..   . .  . .   :  ...  . •
CO}CLUSIONS
1. The exper imental r olling und yawin moment ooefficiente for a wing
• without f lape r in close a&roement with the theoretical valuee.
2. For wing with a ful l span or u par tial epan flap the experimental
roll ing moment ooeffioiente are 10 to 15 percent smaller than the theoretical
values .
3. The rolling moment coefficients wer e little ffe cted by anble of
0 yaw wi thi.n the ran e of + 10 l th a tendency to decrease the model was
~a ed in the poeitive direction. The ynwing moment coefficients ;ere little
affected y the an61e of y w for fl p len the up to 50 pare nt of th s an
but showed o neider ble ch nge for the fullspan f l ap with a tendeno~ to ir.cr nae
as the model s y wed in the poeitive di rection.
Daniel u genhei Air hip Institute,
kron, Ohio, June. 1939.
•