Hammond, Alexander D

Keffer, Barbara M

English

NASA Technical Reports Server

c 
C-IAC COPY RM L53C23 
RESEARCH MEMORANDUM 
EFFECT AT HIGH SUBSONIC SPEEDS OF A 
FLAP-TYPE AILERON ON THE CHORDWISE PRESSURE DISTRIBUTION 
NEAR, MIDSEMISPAN OF A TAPERED 35* SWEPTBACK WING OF ASPECT 
https://ntrs.nasa.gov/search.jsp?R=19930087716 2020-06-17T12:07:31+00:00Z
NACA RM L53C23 
I 
I c 
NATIONAL ADVISORY COMMmTEE FOR AERONAUTICS 
T€E EFFECT AT HIGH SUBSONIC SPEEDS OF A 
FLAP-"PE AILERON ON TEE CHORDWISE FRFSSURE DISI 'RDUTION 
NEAR "ISPAN OF A TAPERED so SWEFTBACK WING OF ASPEZT 
By  Alexander D . Hammond and Barbara M. Keffer 
An investigation was made in the  Langley  high-speed 7- by  10-foot 
tunnel  through a Mach number range  from 0.60 to 0.95 to determine the 
effects on the  chordwise  pressure  distributions f deflecting a flap- 
type  aileron 011 a swept wing. The semispan 35O sweptback w i n g  had 811 
NACA 65A006  airfoil  section, an aspect  ratio of 4, and a taper  ratio 
of 0.6. - 
The results of the  investigation are presented  as  curves of chord- 
wise  pressure  distributions near the midspan of the flaptype aileron 
(0.46-wing-semispan  station). 
The use of  flap-type  ailerons  for lateral control has been  the 
subJect of a large number  of  investigations  at  both low and high speeds. 
These  investigations have for  the  most part dealt with  force  measure- 
ments to determine the  effects of various flap-type  ailerons on the 
characteristics of both  swept and unswept wings. There is, however, 
very  little infomation concerning  the aerodynamic loads on swept winp 
equipped  with  flap-type  ailerons at high subsonic epee&. In order  to 
obtain sane information on such loads and on the  flow  in the vicinity 
of an aileron on swept wings, choritwise  pressure  measurements  have  been 
made  at one spazrwise location on the upper and lower  surfaces  of aso sweptback  semispan wing equipped  with a one-half-semispan  sealed 
aileron having a chord of 20-percent  local wing chord  in  the  Langley 
high-speed 7- by 10-foot  tunnel. 
2 NACA R h 1  L53C23 
For  these  tests  the  flap-type  aileron was deflected  through a 
deflection  range from -30° to 30'. The tests were made at  Mach  numbers 
from 0.60 to 0.95 for an angle-of  -attack  range from Oo to ZOO except 
where  the  angle-of-attack rmge was  limited by tunnel  choking  at  the 
high Mach nmbers. 
M 
P 
P 
PZ 
Pa 
9 
U 
C 
- 
S 
b 
C 
X 
Y 
A 
6 
I 
Mach  number 
pressure  coefficient, 
PI - P 
9 
free-stream  static  pressure, lb/sq ft 
local  static  pressure, lb/sq ft 
atmospheric.  pressure 
free-stream  aynamic pressure, lb/sq ft 
angle  of  attack, deg 
~ ~ 
mean aerodynamic chord of wing, - ZSb/Z czdy, 1.020 ft 
s o  
twice  wing area of semispan model, 4.00 sq ft 
wing span, ft 
local  chord, f% 
chordwise  coordinate of the  pressure  orifice  measured in a plane 
parallel  to plane of synnnetry, ft 
spanwise  distance frm plane of symmetry, ft 
angle  of  sweep of the  quarter-chord  line,  deg 
deflection of aileron, deg 
NACA RM W3C23 
APPARATUS AND MODEL 
3 
The model used i n  this investigation was a semispan sweptback-wing 
model mounted ver t ical ly  i n  the Langley high-speed 7- by l0-foot tunnel 
with the ceiling  serving 88 a reflection  plane. 
The geometric characteristics and d3mensions of the wing are shown 
in figure 1. The w i n g  was made of steel and had 3 5 O  sweepback of the 
quarter-chord line, an aspect ratio of 4, a taper r a t i o  of 0.6, and had 
no twist or dihedral. The wing had NACA 65A006 airfoi l  sect ions paral le l  
to  f ree  stream. 
The pressure  orifices were located on the upper and lower surfaces 
a t  the 46-percent-semispan station. The chordwise positions of the 
or i f ices  are l i s t e d   i n  table I. The difference in the chordrise posi- 
tions between the positive 6 and the negative 6 resulted frm the 
testing technique used Ln obtaining the data. 
TESTS 
A l l  the tes ta  were made in the Langley high-speed 7- by 10-foot 
tunnel. The  data presented in this report are representative of a flap- 
type aileron  deflected fm -30° t o  30° through a Mach number range from 
0.60 t o  0.95 at angles of attack frcun Oo t o  20'. Hcwever, since the 
model was symnetrical, it was found convenient t o  f ix the aileron at  a 
given positive  aileron  deflection and t e s t  through the positive and nega- 
tive angle-of-attack range; this procedure explains the differences in 
the chordvise ordinates given i n  table I for the positive and negative 
aileron deflections. The Reynolds nmber varied frm about 3.1 x IO6 
at  M = 0.60 t o  about 4.0 x 106 at M = 0.95 when based on the wing 
man aeroaynamic chord. 
PRESENTaTION OF DATA 
In  order t o  make these data available, the pressure distributions 
(plots of P against chord) are  being presented without integration or 
discussion. It is f e l t ,  however, that the pressure data, as presented, 
w i l l  be useful i n  the prediction of the local chordwise wing loadings 
and in  the prediction of scme of the aileron loads for aileron configu- 
rations similar t o  those of this investigation. 
4 NACA RM L53C23 
The chordwise pressure distributions as obtained i n  t h i s  investi- - 
gation on a 9 sweptback tapered wing equipped with a flap-type  aileron 
are presented i n  figures 2 t o  14. The so l id  vertical line at  80-percent- 
chord station on the figures indicates the position of the  aileron hinge 
l ine .  The dotted portions of the pressure plots were portions of the 
curve for which there were  an insufficient number of orifices  to  deter-  
mine definitely the exact fairing for the pressure curve. The angle of 
attack shown on the  figures has not been corrected  for  air-etrem mls- 
allnement and t u n n e l - w a l l  effects.  
The following discussion describes the procedure used i n  obtaining 
the data f o r  the figures as presented. Fach or i f ice  on the upper and 
lower surfaces of the Xing was connected t o  a manmeter tube the loca- 
t ion  of which on an arbitrary unit length of the mancmeter board corre- 
sponded to the location of the orifice on the wing chord wlth 6 = 0. 
Photographs were made of the manometer board and the chordwise pressure 
distributions were fa i red on blueprints macle frm the negatives. These 
blueprints were l a t e r  bleached and used for  final figuree. It should 
be noted, therefore, that the vertical  scale for these plots is a 
function of the tunnel Mach nMiber and that the  horizontal  scale is, 
of course, the same for all figures. 
Langley Aeronautical Laboratory, 
National Advfsory CmmLttee for Aeronautics, 
Langley Meld, V a  . 
NACA FU4 L53C23 I__. '5 
TABU I. - u 3 c A T I O N  OF PRESSURE ORDICES 
Chordwise position of 
or i f ices  on upper surface 
O r i f  ice 
1 
2 
3 
4 
5 
6 
7 
8 
9 
10 
I l  
1 2  
13 
14 
L5 
16 
17 
18 
19 
20 
2 1  
22 
Percent chord 
0 
1 
4 
a 
15 
2 5 1  
3 
452 
3 
55- 2 
3 
63 
67 
69 
72 
73 
75 
81 
83 
874 
9 9  
93$ 
9 9  
9 5  
Negative 6 
0 
2 
6 
10 
211, 
3 
3 a  
3 
412 
3 
5& 
5 s  
3 
3 
65 
68 
71 
74 
77 
82 
851 
3 
ssb 
3 
913 1
9 9  
97L 
99L 
3 
3 
Choraise  position of 
orif ices  on lower surface 
~~ ~ ~~ 
k i f   i c e  
1 
2 
3 
4 
5 
6 
7 
8 
9 
10 
u 
12 
13  
14 
15 
16 
17 
18 
19 
20 
2 1  
Percent chord 
Positive 6 Negative 6 
6 NACA RM L53C23 
Comp/ete Made/ C&rncferistks 
A M  4 sq ft 
A s m t  mtio 4 
Tuwr m f io 0.6 
NACA 65AOO6 uirfoi/ section 
\ 
Spunwise /mofion of pmssum 
I orifices 
Figure 1.- Geometric character is t ics  of the 35O sweptback w i n g  equipped 
with a flap-type aileron. A l l  dimensions are i n  inches unless other- 
wise noted. 
. 
7 
-.5 
P 
0 
.5 
-.5 
P 
0 
.5' 
(a)  M = 0.60. 
Figure 2.- Chordwise pressure distribution over a 35O sweptback w i n g ,  a t  
the 46-percent-semispan s ta t ion ,  equipped with a flap-type aileron. 
6 = -30~. 
8 
-LC 
-.5 
P 
0 
.5 
I.( 
9 
-1.c 
-3 
P 
0 
5 
I.C 
4 
(a) M = 0.60. Concluded. 
Figure 2.- Continued. 
d r i  
. 
NACA RM L53C.23 
Mach = 0.80 
-.5 
P 
0 
1.0 0 . / 2 . 3 1 . 5 . 6 . 7 . 8 9  I 
f 
9 
(b) M = 0.80. 
Figure 2.- Continued. 
10 
-1.3 
- 1.C 
-5 
P 
0 
.s 
LC 
NACA RM L5X.23 
(b) M = 0.80. Concluded. 
Figure 2.- ContFnued. 
NACA RM L53C23 11 
a-8 b = - m  
- 1.: 
- / I  
-.d 
P 
c 
.s 
Id 
b 
I 
I , ,  
I 
O . I P J . 4 . 5 § . 7 . 8 9 I O  
8 
( c )  M = 0.90. 
Figure 2. - Continued. 
!k E 
(d) M = 0.95. 
Figure 2.- Concluded. 
NACA RM L53C23 
P 
uppsr surfoca 0 
SUffOC8 A 
P 
(a) M = 0.60. 
Figure 3 . -  Chordwise pressure distribution over a 35O sweptback wing, at 
the 46- ercent-semispan s ta t ion ,  equipped with a flap-type aileron. 
6 = -2s . g 
14 
't I 
-I.! 
- 1.l 
-3 
P 
0 
.5 
0 
P 
(a) M = 0.60. Concluded. 
Figure 3 . -  Continued. 
NACA RM L53C23 
\ L 
Mach = 0.80 
U m r  surfoce o 
L w r  surfam A 
-.5 
P 
0 
(b) M = 0.80. 
Figure 3 . -  Conthued. 
, 1 1 1 1 1 1 1  
I P . 3  # . d  § . 7 . 8 . 9  f 0  
# 
4.2 
4.c 
-3 
P 
0 
.d 
I. 1 
NACA RM L'j3C23 
Uppsr surfoce o 
Larav surfoce A 
44 
(b) M = 0.80. Concluded. 
Figure 3.- Continued. 
? 
3K NACA RM L53C23 
ff-r surfoce o 
Low? surface A 
P 
( c )  M = 0.90. 
Figure 3 . -  Continued. 
t 
"I 
P 
Figure 3.- Concluded. 
. ._ 
-.5 
P 
0 
-.5 
P 
0 
Mmh = 0.60 
0.0 8*-m 
(a) M = 0.60. 
Figure 4.- Chardwise pressure distribution over a 35O sweptback wing, at 
the 46-percent-semispan station, equipped with  a fhp-type aileron. 
6 F -2OO. 
20 NACA RM ~5x23 
1. 
6 O J  .P .3 5 6 I 9 
c 
* 
ID 
P 
(a) M = 0.60. Concluded. 
Figure 4.- Contfnued. 
21 
. 
- t.L 
-3 
P 
t 
I \  
(b) M = 0.80. 
Figme 4.- Continued. 
0 
22 
- P.C 
4.: 
-11 
-3 
P 
0 
.J 
1.c 
-1.5 
-1.C 
-.5 
P 
0 
.5 
I.( 
(b) M = 0.80. Concluded. 
Figure 4. - Continued. - 
W C A  RM L33C23 
P 
( c )  M = 0.9. 
Figure 4.- Continued. 
1 -5 P 0 
(a) M E 0.95. 
Figure 4.-  Concluded. 
iK NACA RM L53C23 
P 
0 
Q.8 &-I5 
I .  , 5 ~  0 .I 2 .3 .4 5 6 .7 .8 9 ID 
9 
I 
P 
Mach = 0.60 
U m r  rurfoec o 
Lowar surfuce A 
(a) M = 0.60. 
Figure 5.- Chordwise  pressure  distribution  over a 350 sweptback wing, at 
the  46-percent-semispas station, equipped with a flap-type aileron. 
6 = -15'. 
26 
I 
P Mach= 0.60 ad2 8.45 
- 1.7 
-bC 
55 
P 
C 
.d 
. 
(a) M = 0.60. Concluded. 
Figure 5.- Continued. 
NACA RM L53C23 
P 
Mmh = 0.80 
Q.0 &-I5 
O J . P . 3 6 5 6  
b 
.7 .8 9 IO 
Upper surfocc 0 
L m r  surfac8 A 
I I  
(b) M = 0.80. 
Figure 5.  - Continued. 
28 - 
1. 
O . f P J . 4 S B 7 . 8 9  
?re 
Uppsr surface Q 
Louw surface A 
P 
lJACA RM L53C23 
I 
(b) M = 0.80. Concluded. 
Figure 5.- Continued. 
NACA RM L53C23 
P 
P 
- II 
-3 
P 
C 
.d 
IJ! 
M = 0.9. 
Figure 5.- Continued. 
WCA RM L53C23 
P 
(a) M = 0.95. 
Figure 5. - Concluded. 
P 
-3 -5 
P P 
0 0 
Mazh 0.60 
0.0 &-IO I 
Figure 
the 
6 =  
(a) M = 0.60. 
6.- Chordwise pressure distribution over a 35O sweptback w i n g ,  a t  
4-6-percent-semispan station, equipped with 8 flap-type aileron. 
-100. 
32 NACA RM L5X23 
P 
O J P J A b . 6 . 7 8 . S f . O  
5 
-.5 
P 
0 
.5 
(a) M = 0.60. Concluded. 
Figure 6 . -  Continued. 
5K 33 
. 
'9 
\ 
\ 
\ 
,9' 
f 
- 1.0 
P 
Mach = 0.80 
ff .4 8.-10 
(b) M = 0.80. 
Figure 6 . -  Continued. 
P 
(b) M = 0.80. Concluded. 
Figure 6.  - Continued. 
-I 
. 
35 
P P 
P 
I 31 
4 . C  
-5 
P 
0 
5 
LC 
( c )  M = 0.90. 
Figure 6.- Continued. 
NACA RM ~ 5 x 2 3  
. 
. 
P 
(d) M = 0.95. 
Figure 6 , -  Concluded. 
NACA RM L53C2j 37 
(a) M = 0.60. 
Figure 7.- Chordwise pressure distribution mer a 35O sweptback w i n g ,  at 
the 46-percent-semfspan station, equipped with a flap-type aileron. 
6 = -50. 
NACA RM L53C23 
- f.5 
- 1.0 
-.5 
P 
0 
.5 Much 9 0.60 ~ = k y )  89-5 
P 
U-r surface o 
Loner rurlocs A 
- f.0 
(a) M = 0.60. Concluded. 
Figure 7. - Continued. 
33 
P 
40 1 
-3 
P 
0 
-3 
P 
0 
1.6 
NACA RM L53C23 
J 
(b) M = 0.80. Concluded. 
Figure 7.- Continued. 
' I  I 
NACA RM L53C23 0 41 
U w r  sur foce 0 
Loner surfore A 
P 
M a h  = 0.90 
a.0 8.4 
O J P J  
5 
( c )  M = 0.90. 
Figure 7.- Continued. 
42 
-p.ol 
-3 
P 
0 
M&= 095 
0.0 I" 
.5 . 
TI.! 
- fL  
-d 
P 
0 
.5 
fa 
(d) M = 0.95, 
Figure 7.- Concluded. 
NACA RM L93C23 43 
Upper surfucs o 
L owe? surfoce C 
(a) M = 0.60. 
Figure 8. - Chordwise pressure distribution over a 35' sweptback wing, at 
the 46-percent-semispan station, equipped with a flap-type  aileron. 
6 = 0. 
44 NACA RM L53C23 
-1.5 
Mach = 0.60 
8-12 B=O 
. C  
-1.3 
-1.C 
-3 
P 
0 
.5 
7 
I.( 
7 
(a) M = 0.60. Concluded. 
Figure 8.- Continued. 
. 
45 
I j 3  0 .I P . 3  4 . 5  6 7 . 8 9  
b 
0 
Upper  surface o 
Lower surface a 
(b) M = 0.80. 
Figure 8.- Continued. 
3 
46 NACA RM L53C23 
J 2 J I 4  B f 8  .9 IO 
Upper rurfoce o 
Lowsf  8 U f f U C 8  A 
(b) M = 0.80. Concluded. 
Figure 8.  - Continued. 
L 
-4 
P 
0 
47 
U m r  surface o 
Lacar surfoce 
( c )  M = 0.9. 
Figure 8.- Continued. 
Upper wrfocr 0 
Lower s u r f e r  A 
-15 
-Id 
-3 
P 
0 
3 
K 
(c) M = 0.9. Concluded. 
Pigwe 8.- Continued. 
NACA RM ~ 5 x 2 3  " - b 
P 
-5 
0 
49 
U m r  surfoce o 
Lamr sur foce A 
f 
(a) M = 0.95. 
Figure 8.- Concluded. 
50 
Upper surfoce 0 
Lower surface A 
0 
(a) M = 0.60. 
35' sweptback w f n g ,  at  
EL flap-type aileron. 
Figure 
the 
6 =  
9.- Chordwise pressure distribution over a 
46-percent-semispan station, equipped with 
50. 
NACA RM L53C23 
P 
Upper surfoce o 
Lowe? Surface A 
Mod = 0.60 
O I . P . 3 * 5 6 . 7 B S 1 & 7  
% 
(a) M = 0.60. Concluded. 
Figure 9. - Continued. 
NACA RM L53C23 
U 
Upper surface 0 
Lover surfuce 
(b) M = 0.80. 
Figure 9.- Continued. 
NACA RM L53C23 53 
P 
9 
Upper surfocs 0 
Lower surface A 
(b) M = 0.80. Concluded. 
Figure 9. - Contfnued. 
54 NACA RM L53C23 
Mod, -0.90 
u -0- 8-50 
7 
4 t 
h 
(c )  M = 0.90. 
Figure 9. - Continued. 
I 
Upper m f a c r  o 
Lower mrfacs A 
4.5 i 
( c )  M E 0.90. Concluded. 
Figure 9.- Continued. 
NACA RM L53C23 
Upper surfaco o 
Lower surfaco A 
Figure 9.- Concluded. 
'K NACA RM L53C23 
" I 
P 
Mach - 0.60 
O I P J l . S . 6 . 7 8 J f . O  
% 
Figure 
the 
s =  
Upper surface 0 
Lower surfac. o 
(a) M = 0.60. 
NACA RM L53C23 
P 
0 
.5 
I I 
(a) M = 0.60. Concluded. 
Figure 10.- Continued. 
59 
P t 
P 
0 
Uppar surface o 
Lowor surfaco A 
(b) M = 0.80. 
Figure 10.- Continued. 
0 
d 
upp.r surface 0 
Lover surface A 
(b) M = 0.80. Concluded. 
Figure 10. - Continued. 
Upper aurfac. o 
Lowor surface A 
- I S I S  
.P 0 
M = 0.w. 
Figure 10. - Continued. - 
. .  . . .  
" .  . 
*per w r f o c e  o 
Lower surface A 
(C) M = 0.90. Concluded. 
Figure 10.- Continued. 
. .. 
Upper surface Q 
Lower rurfoca A 
I 
(a) M = 0.94. 
Figure 10.- Concluded. 
64 NACA RM ~ 5 3 ~ 2 3  
- 1  
Figure 
the 
6 =  
(a) M = 0.60. 
11.- Chordwise  pressure  distribution  over a 35' sweptback wing, at 
46-percent-semispan station, equipped with a flap-type  aileron. 
15'. 
K EACA RM L53C23 
L 
I 
Upper surface 0 
Lower surface A 
(a) M = 0.60. Concluded. 
Figure 11.- ContFnued. 
" 
Upper surfoco o 
Lowor surface A 
(b) M = 0.80. 
Figure U.- Continued. 
\ 
.. 
Upper surface o 
Lower surface c 
1 
(b) M = 0.80. Concluded. 
Figure U.- Continued. 
( c )  M = 0.90. 
Figure 11. - Continued. 
. .  
Upper surtocr o 
Lower  sorfoce A 
(c) M = 0.9. Concluled. 
Figure ll.- Continued. 
I 
. .  
.... . 
Figure U.- Concluded. 
. .  
NACA RM L53C23 
W p e r  surface o 
Lower surface a 
M = 0.60. 
Figure 12. - Chordwise  pressure distribution over a 35' sweptback wing, 
at  the  46-percent-semispan station, equipped with a flap-type aileron. 
6 = 20°. 
NACA RM ~ 5 x 2 3  
-.5 
P 
Mt?dr-0.60 
# =PO.B-PO. 
0 
.5 
-/r- 
. .  
*. 5 
0 
.5 
Upper rurface o 
Lower surfoce A 
(a) M = 0.60. Conclded. 
Figure 12.- Continued. 
NACA RM L53C23 73 
Upper ourfooe o 
Lover rurfoce A 
74 NACA RM L53C23 
e 
I d  
(b) M = 0.80. 
Upper surface o 
Lower  surface A 
Concluded. 
Figure 12.- Continued. 
. .  
I 
Upper surface o 
Louer surface A 
c 
Figure 12.- Continued. 
. .  . 
I 
Upper surfaces o 
Lower surface A 
( C) M = 0.90. Comluded* 
PFgure 12.- Continued. 
Upper m r h c e  Q 
Lower surface A 
(a) M = 0.94. 
Figure 12.- Concluded. 
78 NACA RM L53C23 .
\ i 
Much = 0.60 
P 0-8. 8-25. 
-. 5 
P 
0 
Upper surface o 
Lower surfuce A 
P 
0 
(a) M = 0.60. 
Upper surfuce o 
Lower surfuce C 
(a) M = 0.60. Concluded. 
Figure 13.- Continued. 
80 
P 
0 
P 
0 
I 
Figure 13.- Continued. 
M = 0.80. 
NACA RM L53C23 
I I 
81 
Upper rurfoce o 
Lower surfucs A 
(b) M = 0.80. Concluded. 
Figure 13. -  Continued. 
NACA RM ~5x23 02 
a. L
4.5 
4.c 
-3 
P 
0 
/A 
I 
P 
0 
Upper rurfoce o 
Lower surface 6 
( c )  M = 0.90. 
Figure 13.- Continued. 
lower Nrfasm A 
Leper wrface 0 
. .  . .. . ,  
Upper  surfoce o 
Lower  surfoce A 
I I 
Upper surfoca o 
Lower surfuce A 
(a) M = 0.60. Concluded. 
Figure 14.- Continued. 
. 
86 
&per surfoce o 
Lower surfaca A 
(b) M = 0.80. 
Figure 14. - ContFnued. 
NACA RM L53C23 
4 
P 
9 
Upper surfoce o 
Lower surfocs A 
(b) M = 0.80. Concluded. 
Figure 14. - Continued. 
0 
88 NACA RM L5X23 
( c )  M = 0.9. 
Figure 14.- Continued. 
. .  . . .  . .  
w 
Lower surface A 
upper aurfocs D 
.- I I 
(c) M = 0.90. Concluded. 
Fllgure 14.- Concluded. 


The effect at high subsonic speeds of a flap-type aileron on the chordwise pressure distribution near midsemispan of a tapered 35 degree sweptback wing of aspect ratio 4 having NACA 65A006 airfoil section

The effect at high subsonic speeds of a flap-type aileron on the chordwise pressure distribution near midsemispan of a tapered 35 degree sweptback wing of aspect ratio 4 having NACA 65A006 airfoil section

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