-
!OEN Copy RM L53KO3a
NACA
RESEARCH MEMORANDUM
THE TWISTING EFFECT AT TRANSONIC SPEEDS OF SPOILER AILERONS
ON A 45 SWEPTBACK, ASPECT-RATIO-4, TAPERED WING
By Alexander D. Hammond and Jean C. Graven, Jr.
Langley Aeronautical LaboratoryLangley Field, Va.
0
I
M F-I
I Ll CLASSIFIED DOCUMENT
This material contains information affecting the National
Defense of the United States within the meaning of the espionage
laws, Title 18, USC., Secs. 793 and 794, the transmission or
revelation of which in any manner to an unauthorized person is
prohibited by law. If) NATIONAL ADVISORY COMMITTEE
ctFOR AERONAUTICS WASHINGTON January 7, 1954
N5,
.NF1DENT1AL .
-
NACA RM. L53K03a CONFIDENTIAL
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
RESEARCH MEMORANDUM
THE TWISTING EFFECT AT TRANSONIC SPEEDS OF SPOILER AILERONS
ON A 450 SWEPTBACK, ASPECT-RATIO-4 1 TAPERED WING
By Alexander D. Hammond and Jean C. Graven, Jr.
SUMMARY
An investigation was made at transonic speeds in the Langley
high-
speed 7- by 10-foot tunnel to determine the effect of the
projection of various-span inboard spoiler ailerons along the
70-percent-chord line on the twisting moment and other aerodynamic
characteristics of asolid-
steel 1150 sweptback wing. The investigation was extended
through the transonic speed range by testing in the high-velocity
fi1d over a reflec-tion plane on the side wall of the tunnel.
The twisting-moment coefficient about the 20-percent-chord line
resulting from projection of an inboard spoiler aileron of any span
(including full span) depended only on the incremental lift
produced by the spoiler. The chordwise location of the centers of
pressure resulting from projection of an inboard spoiler aileron of
any span was little affected by change in Mach number near zero
angle of attack but moved forward with increase in Mach number for
angles of attack above about 60. The chordwise location of the
centers of pressure moved toward the trailing edge of the wing with
increase in angle of attack up to approxi-mately l° and moved
forward with further increase in angle of attack at all Mach
numbers. The magnitude of the incremental lift and twisting-moment
coefficients and of the rolling-moment coefficient increased with
increase in span for spoiler ailerons having spans up to 80 percent
of the semispan.
INTRODUCTION
The use of spoiler-type ailerons on highly swept thin wings
looks attractive from both the hinge-moment and wing flexibility
points of view. Considerable work has been done in determining the
effectiveness of various types and locations of spoilers on both
unswept and swept
wings (refs. 1 t 5) . Very little work has been done in
determining the aerodynamic loads resulting from spoiler-aileron
projection, particu-larly at transonic speeds. This paper presents
the results of an
CONFIDENTIAL
-
2 CONFIDENTIAL NACA RM L53K03a
investigation, made at transonic speeds in the Langley
high-speed 7.. by 10-foot tunnel, to determine the aerodynamic
twist about the 20-percent-chord line caused by projection of
various-span inboard, spoiler ailerons. The wing used in this
investigation had a sweepback of 470 at the quarter-chord line, an
aspect ratio of 4•0 a taper ratio of 0.6, and an NACA 65Ao06
airfoil section parallel to the free airstream. The various-span,
inboard spoiler ailerons used were projected to a height of 10
percent of the local wing chord, along the 70-percent-chord line.
It should be noted that the spoiler aileron of this investigation
does not necessarily represent the best spoiler configuration;
however, this wing and spoiler configuration are representative of
configurations that might be used in high-speed flight. It is felt,
therefore, that the results of this inves-tigation give the general
trends of the variation of the twisting moment about the elastic
axis resulting from projection of various-span inboard spoiler
ailerons.
COEFFICIENTS AND SYMBOLS
CL lift coefficient, Twice lift of semispan model
qS
CD drag coefficient, Twice drag of semispan model
qS
Cm pitching-moment coefficient referred to 0.25 Twice pitching
moment of semispan model
qS
Cmt twisting-moment coefficient about the 20-percent-chord line,
Twisting moment of semispan model about 0.20c
qS
C1 rolling-moment coefficient, Rolling moment of wing with
spoiler - Rolling moment of plain wing
qSb
Cn yawing-moment coefficient, Yawing moment of wing with spoiler
- Yawing moment of plain wing
qSb
CONFIDENTIAL
-
NACA RM L53KO3a CONFIDENTIAL
q effective dynamic pressure over span of model, lb/sq ft
S twice area of semispan model, 0.127 sq ft
mean aerodynamic chord of wing, 0.1807 ft on model, -J c2dy so c
local wing chord, ft
b twice span of semispan model, 0.7071 ft
y spanwise distance from plane of symmetry, ft
P mass density of air, slugs/cu ft
V average free-stream air velocity, fps
M effective Mach number over span of model
Ma average chordwise local Mach number
MI local Mach number
R Reynolds number of model based on E
a angle of attack, deg
L increment caused by spoiler-aileron projection
A aspect ratio, b 2/S, 4.0 on model
The forces and moments on the wing are presented relative to the
axes shown in figure 1. Wing twisting moments were measured about
an axis which corresponds to the 20-percent-chord line. All other
forces and moments and the angles of attack were measured relative
to the wind axes (fig. 1) which intersect at the plane of symmetry
and the chord plane of the wing at the
25-percent-mean-aerodynamic-chord station as shown in figure 2.
MODEL AND APPARATUS
The semispan wing used in this investigation had 1150 of
sweepback of the quarter-chord line, an aspect ratio of 4.0, a
taper ratio of 0.6,
CONFIDENTIAL
-
4 CONFIDENTIAL NACA BM L53KO3a
and an MACA 65A006 airfoil section parallel to the free stream.
The wing was made of steel and was constructed to the plan-form
dimensions in figure 2.
The spoilers were made of brass and were soldered along the
70-percent-chord line. The spoilers had a projection of 10 percent
of the local wing chord and various spans from 0.20b/2 to 1.00b/2
(fig. 3).
The data were obtained in the Langley high-speed 7- by 10-foot
tunnel with the model mounted on the tunnel side wall in a manner
similiar to that of the model shown in figure 4. The model was
mounted to an elec-trical strain-gage balance through a slot in the
reflection-plane turn-table; this slot was sealed with a
sponge-rubber-wiper seal glued to the turntable to reduce the
leakage around the wing butt. The forces and moments were recorded
by means of a recording galvanometer.
TESTS
The tests were made in the Langley high-speed. 7- by 10-foot
tunnel. Typical contours showing the Mach number distribution over
the side-wall reflection plane in the vicinity of the model are
presented in figure 5. Effective test Mach numbers were obtained
from contour charts siiniliar
I/2b to those in figure 5 by the relationship M = cMa dy. So
For these tests a Mach number gradient generally less than 0.02
was obtained below a Mach number of 0.95, and the gradient
increased to 0.05 at the higher test Mach numbers.
Force and moment measurements were made for the model through an
angle-of-attack range of -5 0 to 100 measured in a plane
perpendicular to a normal axis through the quarter-chord point of
the mean aerodynamic chord. These measurements were obtained from a
Mach number of 0.61 to 1.09 . The variation of Reynolds number with
Mach number is presented in figure 6.
In view of the small size of the wing relative to the tunnel
test section, jet-boundary and blockage corrections were believed
to be insig-nificant and were not applied to the data. No
reflection-plane correc-tions were applied to the data of this
investigation.
-
NACA EM L53K0 3a CONFIDENTIAL 5
RESULTS AND DISCUSSION
The variations of the incremental lift, drag, pitching-moment,
and twisting-moment coefficients with spoiler span for various
inboard, spoiler ailerons are presented in figures 7 to 10,
respectively. The variation of the rolling-moment and yawing-moment
coefficients with spoiler span are presented in figures 11 and 12,
respectively.
The data of figures 7 to 12 indicate that for the Mach number
and angle-of-attack range investigated, an increase in spoiler span
results in an increase in the magnitude of the incremental lift,
drag, and twisting-moment coefficients and in the rolling-moment
coefficient for spoiler spans up to 0.80b/2. Spoilers having spans
greater than 0.80b/2 gave values of these coefficients about the
same or slightly less than those for the 0.80b/2 span spoilers. An
increase in spoiler span up to 0.20b/2 results in an increasingly
more negative incremental pitching-moment coefficient LCm; further
increase in spoiler span generally resulted in an increasingly more
positive 6Cm (fig. 9). The magnitude of the yawing-moment
coefficient increased with increase in spoiler span for all spoiler
spans investigated.
The variation of the incremental twisting-moment coefficient
with incremental lift coefficient are presented in figure 13 for
various-span inboard spoiler ailerons. The incremental twisting
moments presented and discussed herein represent the change in
twisting moment about an axis corresponding to the 20-percent-chord
line resulting from deflection of the various-span inboard ailerons
on a rigid wing. It can be seen from figure 13 that there is a
nearly linear variation of the. incremental twisting-moment
coefficient nCmt with incremental lift coefficient ACL
for all spoiler spans at a given angle of attack and Mach number
in the angle of attack and Mach number ranges investigated. This
indicates that the centers of pressure of the additional load
resulting from projection of the various-span inboard spoiler
ailerons lie at a constant distance from the 20-percent-chord line
and that the twisting moment about this axis is dependent only on
the incremental lift. For all practical pur-poses, therefore, the
loci of the centers of pressures of the additional load for the
various-span inboard spoiler ailerons may be considered to lie
along a constant-percent-chord line. The chord line that is
coinci-dent with the loci of the centers of pressure varies with
both angle of attack and Mach number.
Since the variation of PCmt with ACL is nearly linear, the
param-
eter gives the perpendicular distance in percent of c
between
the 20-percent-chord line and the loci of the centers of
pressure of the
DCO &A
-
6 CONFIDENTIAL NACA RN L53K03a
additional load resulting from projection of the various - span
inboard
spoiler ailerons. The variation of the parameter with angle
of
attack for various Mach numbers and with Mach number for various
angles of attack is presented in figure iA. From figure 14 it can
be seen that the loci of the centers of pressures move toward the
trailing edge with increase in angle of attack up to approximately
40 and move forward with further increase in angle of attack for
all Mach numbers. Near zero angle of attack there is little change
in the chordwise location of the centers of pressure with change in
Mach number; however, at angles of attack of 60 and above there is,
in general, a forward movement of the loci of the centers of
pressure of the various inboard spoiler ailerons with increase in
Mach number (fig. lii.).
CONCLUSIONS
An investigation was made in the Langley high-speed 7- by
10-foot tunnel to determine the effect on the twisting moment and
other aero-dynamic characteristics of various-span inboard spoiler
ailerons on a solid-steel, 450 sweptback, aspect-ratio- li-,
taper-ratio-0.6 wing. The following conclusions may be drawn from
the data:
1. At a given angle of attack and Mach number, the twisting
moment resulting from projection of an inboard spoiler aileron of
any span about the 20-percent-chord line or any other axis
approximately parallel to the 20-percent-chord line depended only
on the incremental lift produced by the spoiler.
2. The loci of the centers of pressure of the additional load
resulting from projection of an inboard spoiler aileron of any span
moved toward the trailing edge of the wing with increase in angle
of attack up to approximately 40 and moved forward with further
increase in angle of attack for all Mach numbers investigated.
3. At near zero angle of attack there was little change in the
chord-wise location of the centers of pressure of the additional
load, resulting from projection of an inboard spoiler aileron of
any span, with change in Mach number.
... In the Mach number and angle-of-attack ranges investigated,
an increase in spoiler span resulted in an increase in the
magnitude of the
CONFIDENTIAL
-
NACA EM L53K03a CONFIDENTIAL 7
incremental lift and twisting-moment coefficients and in the
rolling-moment coefficient for inboard spoiler ailerons having
spans up to 60 per-cent semispan.
Langley Aeronautical Laboratory, National Advisory Committee for
Aeronautics,
Langley Field, Va. October 19, 1953.
1. Fikes, Joseph E.: Hinge-Moment and Other Aerodynamic
Characteristics at Transonic Speeds of a Quarter-Span Spoiler on a
Tapered 150 Swept-back Wing of Aspect Ratio 3 . NACA EM L52A03,
1952.
2. Hammond, Alexander D.: Lateral-Control Investigation of
Flap-Type and Spoiler-Type Controls on a Wing With
Quarter-Chord-Line Sweep-back of 600, Aspect Ratio 2, Taper Ratio
0.6, and NACA 65AOO6 Air-foil Section - Transonic-Bump Method. NACA
EM L50E09, 1950.
3. Hammond, Alexander D."and Watson, James M.: Lateral-Control
Investi-
gation at Transonic Speeds of Retractable Spoiler and Plug-Type
Spoiler-Slot Ailerons on a Tapered 600 Sweptback Wing of Aspect
Ratio 2 - Transonic Bump Method. NACA EM L52F16, 1952.
11. Schneiter, Leslie E., and Ziff, Howard L.: Preliminary
Investigation of Spoiler Lateral Control on a 420 Sweptback Wing at
Transonic Speeds. NACA EM L719, 1947-
5 . Fischel, Jack, and Schneiter, Leslie E.: High-Speed
Wind-Tunnel Investigation of an NACA 65-210 Semispan Wing Equipped
With Plug and Retractable Ailerons and a Full-Span Slotted Flap.
NACA TN 1663, 1948.
CONFIDENTIAL
-
'1
MMI I
ki
r.1 ['1 CONFIDENTIAL NACA RM L515K03a
z
z
Figure 1.- System of axes, forces, moments, and deflection,
positivedirections denoted by arrows.
CONFIDENTIAL
-
'Jq)
63 4
(r)
It
NH
f
-Cj
0•
OQ) r -P -P0 cd
IQ)
ci)
-Pa)
z1- 0
0 r-1 U) -Pa)
1! 0
0 cci
0 r-4 ci)
rd
•t-1H rdH
rd
bO
0 q^4
14 rl C.) cc3cd r,Q
c'JUi
CONFIDENTIAL
NACA RM L53KO3a CONFIDENTIAL 9
-
U) rl 0
cd
a) H
a) H
0
U)
U)
0 •rl
(U
+
0
U) H r1 a5
4.)
a)
Ie\
10
CONFIDENTIAL I'TACA EM L53KO3a
Q)
CONFIDENTIAL
-
NACA EM L53K05a CONFIDENTIAL 11
/
- 1 -
U) '0 p1
C')
U)
U)
H rd (U U)
0 HO
'di 00
cd
IN-U)
rd
La
H cd 0
CONFIDENTIAL
-
lz
Vt
c4
e4
ui 'ouo,d U0i9/J9Juioi; euo/s/p esMuodg
- Irt
Cj
%
EIUI
m1 4b. q—Iti.
:i
U)
'-0
.0 a
('1
5_S
3
ba)
0
Id d
co (\j d
ci ct
0
0)
S.-0
+ S.
i...
S.-S.-
'I
II1 I
-
cc r.. C.) S.-
a)
r1 114
S.-
12
CONFIDENTIAL NACA 1RM L5KO3a
S.
C,..
0
r-
14, INd
W '9U0/d 1IOi//f 91
woij o,,aojs,p 9s1muodS
CONFIDENTIAL
-
(I)
+ U)
a) 4)
0 cl—I
S..
(J
r
rl
P4
rd ,-1w 0
tf\ a) zj-
a) CH 10. 0-p
00
cd -p
NACA RM L53KO3a CONFIDENTIAL 13
-
O)
i ieqwnu sp,OUA'ay
CONFIDENTIAL
-
Q(f)
a) H
0 pq
cd -p
a) •d CH 00
q-Iu) cHa) a)r-1
-p
.rq ('3
cd
a)
a)
cd
CH
a)r
r U)
ci)
00
0 r1 CU
c'3O .'-
14 CONFIDENTIAL
NACA RM L73K03a
U1 LI
-
NACA RN L53KO5a CONFIDENTIAL 15
a 4 i D 0
_LI_LU I HLLU
0 •'
0
:1• -- 5c 4.c-4
) a)ffl
oa)
c-I c—I c1) C3 0 o,-
cd to_4 cda)
riW
cd -P0
H cd
co
o 00 0 0_) "-) "-)
CONFIDENTIAL
-
16 CONFIDENTIAL NACA FM L53K03a
a 0 a o LI 4 I 0 CI 0 - - o o
Q) c c Q
V:
CH 4-1
0 ECJG)
cd
H
OHcd
bo a)
H a) 4 Ea ci
Ea
ra -pa)
a)
c: cf) o,
O
O -4 H -p cd H
a) c3
Ea
1.
pq EQ
OH
H
CONFIDENTIAL
-
NACA RM L53KO3a CONFIDENTIAL 17
,i LI o 0 i1LI D
'
rel
M9
-p
(I) •r-1
•rI 0
C4-'
a) OCJ)
0
cd to
a)
a)+ C.) .,-1 c)J
C) -)
a) ai a) •-I 0H 0
rd -P
o 0,-i
0 c •-4 + -P r4 Ca)c
•r-1 C) -,
cda) . a)
I U) a)O
a)cC;4) U) H4-) a)
CONFIDENTIAL
-
21 DOd 21 GGO
18 CONFIDENTIAL NACA RM L73K03a
ci) H
-Pc)
f-'-, 'J. LC)CI) +)ccj
ci •-1CQ
+-cH
QV) ci)tcO o oc
+H c3
co
in
oh llp
Locd
cd r-1 PA
a)
c
CONFIDENTIAL
-
0 D 0 £1 .1 0 0
Q Q Q Q Q c c
NACA BM L53K03a CONFIDENTIAL 19
v
0 pq
-cd
+) - PCD
-HO C) -H Cl)
r..
0 C)
\O
ul
cd
cd
'c: • C'JcU U
U)
CONFIDENTIAL
-
20 CONFIDENTIAL NACA RM L53K03a
A o o
0 tC)
II
0
Ii
0 (C)
0
0\1
(34
U) C)4-)H
•-I q-.IU)
'U I OHr4
- CH
+'O)U)
OG) Ea
(3J
-4 u
U) 4-) U)
"V
Ui
Ui
Nc%3J
N OO
H U) Ui
I ,p + - rdUi
00(0
Uioo o i • - r •H- PU) -) 4-Uir-1
Xi U) ri •r-lC)3
:X
P-4 (L)
(1) (') .-I
E-t pq. c) U) W .C.)
Ui C\ H+' Ui+
q)o,o c_) I. 'J C3r-1O
I I
cI
cç
CONFIDENTIAL
-
Qo
ni
Ii
th
i±
- rf I --.
NACA EM L5 K03a CONFIDENTIAL 21
C) CH i3 0
+ o ci) H
c- to o
a3 ci) H H tto C3
CJ a)
4 ) C!)
0
cd CH
C4-4
(o(o
H Ea Ua3
a)P 4)
4-1 cd
cc
( its'
CONFIDENTIAL NACA-Langley - 1-7-54 325
Page 1Page 2Page 3Page 4Page 5Page 6Page 7Page 8Page 9Page
10Page 11Page 12Page 13Page 14Page 15Page 16Page 17Page 18Page
19Page 20Page 21Page 22Page 23