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 the yawing moments due to an angular velocity in yaw. The model was tested at 0 degree and 5 degree pitch; 0 degree, 5 degree, and 10 degree yaw; and with split flaps covering 25, 50, 75, and 100 percent of the wing span and deflected 60 degrees. The results are given in the form of load distributions and as calculated moment coefficients. The experimental values of rolling- and yawing- moment coefficients were in fairly close agreement with theory

Rokus, F

Troller, TH

English

NASA Technical Reports Server

. 
.TECH~ICAL NOTES .-'W, f!,? 
XATIONAL ADVISORY COKXITTEE FOR AXROXAUTICS I i - .’ - qrr -I 
---I_-- 
No. 735 
- 
. 
. _- _ _.- ; 
-- 
PRESSURE-DI8TRIBUTIOM MEASWREYEHTS ON A TAPERED WING 
WITH A PARTIAL-SPAN SPLIT FLAP IN CURVED FLIGHT 
By Th. Troller and. F. Roktis 
Daniel Guggenheim Air.G=lip Institute 
l 
i7ashinston 
November 1939 
---- 
. 
II 
https://ntrs.nasa.gov/search.jsp?R=19930081496 2020-06-17T22:56:43+00:00Z
_ 
---. 
BATIONAL AD?ISORY COMUITTE: FOR.AERONAUTICS 
------ 
. TECHNICAL NOTE NO. 735 
---- 
PRESSURE-DXSTRISUTIOH XEASUREMENT'S OX A TAPERED WING 
i-1 
.- 
WITH A PARTIAL-SPAP SPLIT FLAti IN &RVZD FLIGHT 
* L. 
By Th. Troller and F. Rokus - 
SUMMARY 
Pre.ssure-d'rstribution.tests were made on the .32-foot 
whirling arm of the Daniel Guggenheim Airship Institute 
of a tapered wing to determine the rolling and the yawing 
moment6 due to an angular velocity -in yaw. The model was 
tested at 0' and 5' pitch: 0'. 5', and 10' yaw ; and with 
split flaps covering 25, 50, 75, and lOO.percent of the 
wing span and deflected 60°. The results are given in the 
form of load distributions and.as calculated moment COef- 
, ficients, The experimental values of rolling- and yaming- 
moment coefficient6 nere in fairly close agreement wfth 
theory. 
.- 
_ _ INTRCDUCTION _- - I 
An investigation of the Drespure distribution over a -. -. 
model of a Clark Y tdperod mfng;'with full-span flaps made La 
on the 32-foot whirling arm of the Airship Institute aas 
resorted in reference 1. Similar tests-are.rcported in 
this Dapor with the same type of equiDmen%-5s mfth t-he .- .- -- -y original full-span'flass converted into partial--span flaps 
covering 25, 50, and 75 percent'of the ring.span. The 
-- 
.mork was,carried out with the financial assistance of the -- 
National Advisory Committee for Aeronautics. 
. APPAeRATUS AND TESTS 
A complete description of the model, the flaps, the 
test set-up, ana the test procedure is given in referents l .- 
tier 
Each half of the full-span flap deecribcd in the ear- 
tests was cut into four equal-lengths and the result: 
ing sections mere-added to or removed-from the model to 
produce the desired flap length. - __- - --. .__._.. .> 
A fern modifications of measuring equipment were nec- 
2 B.A.C.A. TeChnicalVNote lo. -735 
essary for the partial-gpan flaps. Fqurteen additional. 
chordwise rows of pressure orif.ic.es mere inserted along 
the span of the wfng near the ends of the different flaps. 
The locations of the final 26 rows of orifices are shown 
in the sketch in fiGure 1. 
The model with the 25-, the 50-, and the 75-percont- 
span flaps deflected 60' was tested at O" and 5O pftch and 
at O", f5', and ltlOo yaw. Positive yaw denotes a displace- 
ment of the model so that the leading edge faces the center 
of rotation. For the 5' pitch condition, tests were also 
made of the model without flsps and with full-span flaps 
at all the yaw angles. For all the tests, the ratio of the 
span to the turning radius wag 0.133. 
TEST RESULTS ABD DIBCUSSION 
The pressure distribut2on over the upper and the lowor 
surfaces of the airfoil is shown by plots of ratios of-the 
static pressure p to the dynamic pressure q. correspond- 
ing to the true fliqht velocity. Typical diagrams are given 
in figure 2 for different flap lenqthg and test conditions. 
The length of the flap ig seen to affect the pressure dis- 
tribution over the entire span. The maxf@um value of P/90 
at any chord section increased as the flap was lengthened. 
The pressure diaGrams for the various wing arrangements 
mers then graphically intee;'ra.f;ed and, from the data thus 
obtained, nondimensional coefficients were computed for the 
wing as a whole. Table I shows the values of the rollin<- 
moment coefficient Cl and the yaning-noment coefficient 
% obtained by integration, The local normal-force coef- 
fici-ent c, and the local chord-force coefficient Cc are 
plotted in figures 3 an& 4, respectively. 
The coefficients are defined as follows: 
Gt = &A-- c, = 8 11 C = ---- 
q, b S q, b S cn = q, Clocal 
cc 
40 'local 
whore 
L is the rolling moment, 
N, yawing moment. 
B.A.C.A. Technical. Note No, '135 . 3 
c 
’ _ 
1 I 
_.J ’ 
I 
. 
c 
a.’ . 
n, 
C, 
b , 
. St 
CIOCR1' 
normal forts per unit span along longi- - 
tudinal axis of wing; 
longitudinal force per unit span. 
wing span. 
wing area. .. 
chord of airfoil se,ction* - 
The theoretical..rollPng-moment and yarning-moment co- 
effi.ci.ents rere computed according to the method of refer- 
ence 2. The theoretical rollipg'motient L ii 'define&by. 
and the theoretical yawing moment N is defined by 
b/2 
IN = Cn r $+ q S b + q r / QO ?+ y dy 
=: 
$/2’ . . . - . . 
where 
Czr 
rr 
V, 
q, 
%,’ 
Cd, ’ 
R, 
Y, 
is the rolling-moment coefficient obtained 
from reference 2. 
turning speed, rad.ians per second. 
flight velocity. 
dynamic pressure. 
yawing-moment coefficient obtained from 
roference.2. - -.. _ _ _ ._ .- 
loc'al profile-drag coefficient. 
turning radius at center o.f wing. -,: L 
distance along span,. - +. 
For the Clark Y wing without flap, the angle.of at- 
tack for zero lift was assumed to be -4.8o. For the con- 
. 
4 N.A.C.A. Technical Bate No. '735 
ditinn of 0' nitch this value gives CL = &8' and, for I 
the condition"of 6d pitch, 9 . 8'... 
flap deflected 60' 
The 26-percent-chord 
was assumed to be equivalent to an in- 3 
crease of the effective angle of attack equal to ha = I.$.'. . 
This effect of the flag was assumed on the basib of fiqure 
4- of reference 3. 
The profile-drag coefficient was assumed to be 0.01 
for the win4 without flap. For the wing with flap, the 
value of CD was taken from figure 4 of reference 3 as 
3.28 at a = O*. Vith cdi = 0.08 at CL = 1.3 for eoro 
enqle of attack of the reference wing with an aspect ratio 
of 6.1, a value Of -cd0 2 0.20 is obtained for the Clark 
Y win-g wit-h a 25-Dsrcent-chord split flap deflected 60'. 
Fi;ure 5 gives a comparison'between the experimental 
and the theoretical rolling-moment and yawing-moment coof- 
ficients for the O" 
and 5' pitch. 
yaw position and the positions of 0' I 
. 
1. The experimental rolling- and yarning-moment cocf- I _ 
ficients for a wing without flaps.are in close agreement 
with tile theoretical values. 
2. For a wing with a full-span or a partial-span 
flap, the experimental rolling-moment coefficients are 10 
to 15 percent smaller than the theoretical values. 
? The rolling-moment coefficients were little af- 
factad'by angl-e of yaw within the range of ttlOo with a 
- tendency to decrease as the model-was yawed in th-a posi- 
tivo direction. The yawin.%-moment coefficient0 wore lit- : 
tle affected by the angle of yaw for flap lengths up to 
50 percent of the span but shelved considerable change for 
the full-span flap with a tendency to incroaso as the 
model was yawed in the positive direction. 
Daniel Guggenheih Airship Institute, 
Akron, Ohio, Jiini1939.~ 
N.A.C.A. Technical Note No. 735 5 
REFFREXCZS 
. 
1. Troller, Th., and Rokus, F.: Pressure-Distribution 
Measuremants on a Tapered Wing with a Full-Span 
Split Flap tn Curved Flight. T.W. No. 683, 
N.A.G.R., 1939. -. 
2. Pearson, Henry A., and Jones, Robert T.: Theoretical 
Stability and Control Characteristics of Wings with 
Various Amounts of Taper and Twist. T.R. No. 635, 
X.A.C.A., 1938. 
3. Wenzinger, Carl 5.: The Fffects of Full-Span and 
Partial-Span Split Flaps on the Aerodynamic Charac- 
teristics of a Tapered Wing. T.N. No. 505, N.AZC.A., 
1934. 
. 
* & . Lc ., 
. 
TABLE; I. Moment Coefficients O.btained by Integration 
Cl %I 
No 25- 50- 75- loo- No 2% 50- 75- 100 
flap percent percent percent percent percent Qercent percent percent flap flay, fla;, flas flap flap flap flql flap 
0' -Ditch 
10 
5 
0 
-5 
-10 
0.0055 0.0075 O.Ol28 
.0072 .0079 ,Ol28 
.0090 .0087 .0138 
LOO92 .0102 ,015o 
l 0082 .0098 .0150 
r .--~7---- 
-- -- 
a0.0225 
o.coo93 
a0.00033 
0.00179 0.00373 -1. 
,00087 .00161 .00360 ao.oo832 
a.o229 a.ooo33 .00070 .00146 ;0&'67 a.OO514 
-- cm .Ooo& .OOlG? .00192 -- 
-- These coe?ficients considered unreliable 
10 1 0.0127 
5 
0 
-ii 
%013i 
".0X36 
.0130 
.ol20 
0.0127 0.0144 0.0189 
.ol20 .0150 .0196 
.0124 .0153 .0203 
.0143 .0177 .0233 
.0137 .0187 ,0267 
5" pi 
0.0284 
a.O279 
a.0282 
.a?10 
.0317 
.tch 
These coefficients considered unreli&le 
aData from reference 1. 
P.A.O;A. Technlorl Note No. 736 Flg#.l,S 
. 
, 
I $1 ' 
43 a 1 1 a34 
Ohord 
k 
3 
4 
.I 
DiItanoe 
from oenter 
Of #gall, 
a%; 
4.39 
6.03 
6.68 
2 
Dietanoe Dlstrnne 
Ohord from oenter ChOrd from oenter 
of w=% Cl m=h 
8" 
10% 10 Pa'& 
a.30 
8' 
Ezx 
14:ss 
ij: 83.33 
13 a4.m 
9 la.13 
Piwe l.- 
Bketoh of 
ting ahow- 
%r%*- 
prasaurs- 
orifioe 
ohords. 
. = :
- 
Figure 6.- Qomparleon of theoretioal rod experimental rolling-moment an3 yawing-trament oorffloientr. 
N. A.C.A. Technical Note No. 735 Fin. 
? 
-s 
, 
b ’ 
-3.0 
-2.0 
-1.01 ( j 
1 
P 
-pb 
-3 
-2 
-1 
0 Upper surfaae x Lower surface 
-2.0 Ch rd 9 cdratllI tcdt-al13l 1 1 
.R 
-1 
1 
* 
O-3 
-2 
.O 
.O 
.O 
i i i i id i i i ki i i i i i i i i I 
2a 
.- - 
- 
.(a) The 25-peroent-span flap. 
Figures 2a to c.- Pressure-distribution plots for 5' pitch, O" yaw, flaps de- 
flected 60'. . 
N.A.C.A. Technical Note No. 735 F ig. 2b - 
h 
- 0Upper surface x Lower surface 
!!I p&j i !I Ote 'aan \I , 
1.0 
$, (-J . 
t i 
i i 
i 
i 
i i 
i i 
I I I I 1 
I I
Imer suan I I I 1 
-2.0 
--- 
1 IYUJI I 0 
,l.O 
(b) The 50-peroent-span flap. 
Figure 2.- Continued. 
N.A.C.A. Technical Note No. 735 Fig.2c 
-3i01 I I 1 I I a I I I I I k I I I I lo I I I- 1 
P 
TO 
/Chdrdll 1 1 k lChdrd13 I I I\ Chcrd 5 I\Cl.ori. 8 
^ ^ I f I I Lb 
\ 9 Y 
\ 
I a, 
1.0 
$3.0 
,2.0 
,l.O 
0 
1.0 
c 
n 
0 Upper surface xLower surface 
Ch rd 9 Ch,'rd 10 Ch rd 11 , , , , , 1, ,, , , , , , 
0 i i/i i i i i Vi i 
/ H 
> - 
F I 
-1.0 
0 
1.0 
(c) The 75-percent-span flap 
Figure 2.- tinoluded. 
1 L. 
I 
5si-i i i i i i i i ii-i I 
IiiiiIiiIIIiiid_IIiiiI In 
0 Full-span flap 
x 75 percent flap 
0 50 peraent flap 
A 25 peroent flap 
0 Vithoat flap 
---ROM reference 1 
k 
!: Iiiiiiiiiiiiii-iiiiii 
d 
.: 
& 
2 
1 
~l,OO I I I I I I I I 
. 
0 0 
1.00 .Bo .60 A0 .20 0 .a0 40 .60 .a0 1.00 1.00 .Bo .60 .40 .20 0 .20 A0 1.00 ‘A 
Poeition along sing spy 
.40 .@I 
Figure 3.- Distribution of noml-form oosfficlants aloog span df riag; pitoh, 50. 
Position along tillg span 
I I 
bl 
. 
. l 4 L 
- I 
0 Full-span flap 0 50 percent flap 0 Without flap f 
x 79 parcent flap A 25 percent flap --- From refermoe 1 : 
.501 , I , , , , I I , , , , I I , I , JOI I I I I I I I I I I I I I I I I I I I 1 5 
i i i t ii i 1 i ii I i i I 
! I ! I I I I I I I I 
.40 
i ‘\J\ 
I I I I I 
I 
I I ii] j 
.lO 
0 
1.00 .a0 .60 A0 2.0 0 .20 .40 050 .a0 1.00 
Position along ring span 
outer hlf of span Inner half of aan 
I 
Flgws 4.- Disbribntio? of chord-force ooeffioisnts .3 
.’ Poaitlon almg wing .spm 
Outer half of span Inner half of span 
.long span of wing! pitch, B”. 
3 
P 
P 


Pressure-distribution measurements on a tapered wing with a partial-span split flap in curved flight

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