Vogler, Raymond D

English

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SECURITY INFORMATION
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RESEARCH MEMORANDUM
WIND -TUNNEL INVESTIGATION AT HIGH SUBSONIC SPEEDS OF
A SPOILER-SLOT-DEFLECTOR COMBINATION ON AN
NACA 65AO06 WING WITH QUARTER-CHORD
LINE SWE.PT BACK 32.6°
By Rapmn.d D. Vogler
LangleyAeronauticalIkboratory
hl@~Y Field,Vas
cIAmnmLlDocuMEm
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON
May 29,1953
RECEIPT SIC!! ”-uRE
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https://ntrs.nasa.gov/search.jsp?R=19930087571 2020-06-17T09:25:38+00:00Z
TECH LIBRARY KAFB, NM
NACA RM L53D17
NATIGNAL ADVISORYC!OWIIT’I’JRFOR AERONAUTICS
.
,
RESEARCH MEM3UMDUM .
WIND-TUNNELINVESTIGATIONAT HIGHSUBSONICSpEIX)SOF
A SPOILEI1-SLOT-~lIWcTORC MBINATIOIJON Mi
NACA 65AO06 WING WITH WARTER-CHORD
LINE SWEPI BACK 32.60
By Raymond D. Vogler
SUMM4RY
An investigation was made in the Langley high-speed 7- by 10-foot
tunnel through a Mach number range from 0.4 to 0.91to determine the
lateral control characteristics of a wing-fuselage combination with the
wing qurter-chord line swept back 32.60, an aspect ratio of 4, a taper
.
ratio of 0.6,and an NACA 6~AO06 section. One wing panel was equipped
with a 50-percent-semispan inboard spoiler-slot-deflector combination
located between the 55-and 70-percent-chord lines.
.
*
As was previously found at low speed, the loss in rolling-moment
effectiveness of an unvented spoiler at high wing angles of attack is
materially reduced by the incorporation of a slot and deflector at Mach
numbers up to 0.91. The optimum ratio of deflector to spoiler projec-
tion for best results varied with angle of attack, but a ratio of three-
fourths to one gave appreciable roQing-moment effectiveness through the
angle-of-attack range from O” to 20”.
INTRODUCTION
The spoiler used as a lateral-control device has been the subject
of considerable investigation at low and high speeds, and on both swept
and unswept wings. Recent investigations of spoilers used as lateral-
control devices have shown that on thin wings with small leading-edge
radii the unvented spoiler loses effectiveness rapM1.y as the angle of
attack is increased above about 8° (ref. 1). References 2 and 3 have
? shown that this loss in effectiveness at the higher angles of attack
could be reduced by using a slot in the wing behind the spoiler that
allowed the air to flow through the wing from the lower to the upper
.
surface when the spoiler was deflected.
=
2 NACA RML53D17
The purpose of this investigation was to determine whether a slot .4
plus a deflector is as effective at high subsonic speeds as it was at
low speeds (refs. 2 and 4), and if so, to determine the ratio of deflec-
tor projection to spoiler projection for most benefici-alresults. The .
Investigation was conducted in the Langley high-speed 7’-by 10-foot
tunnel through a Mach number range from 0.4 to 0.91 ~d at an angle-of-
attack range from 0° to 20° except when limited by tunnel operating con-
ditions. Rolling and yawing moments were obtained wit_hspoiler alone,
with slot plus deflector, and with spoiler-slot-deflectorcombination.
Lift, drag, and pitching moments of the model were also obtained for
.—
the spoiler-slot-deflectorcombination.
.-
SYMBOLS AND COEFFICIENTS
The forces and moments measured on the model are presented about
an orthogonal system of axes. The longitudinal axis was parallel to
the free-stream air flow and the vertical axis was in the vertical plane
of symmetry. The origin of the axes was in the plane of symmet~ at a
longitudinal position corresponding to the projection_of the quarter-
chord point of the mean aerodynamic chord (fig. 1). ~
.—
CL lift coefficient (Lift/qS) –~
Cal drag coefficient (Drag/qS)
.
cm pitching-moment coefficient (Pitching mome@/qS=)
cl rolling-moment coefficient resulting from spoiler and/or
deflector projection (Rolling moment/qSb)
Cn yawing-moment coefficient resulting from spoiler and/or
deflector projection (Yawingmoment/qSb)
dynamic pressure) pounds per square foot ()
PV2
~
-T
P mass density of air, slugs/cu ft —
v free-stream air velocity, fps
s wing area, 2.25sq ft .1-
b wing span, 3.0 ft
c local chord, ft
E mean aerodynamic chord of wing, 0.765ft
v
m#&i:&!WF’J&
NACA RML53D17
h M Mach number
R Reynolds nuniber
a angle of attack, deg
5= spoiler projection, negative when projected from upper surface
of wing, percent chord
bd deflector projection, positive when
surface of wing, percent chord
APPARATUS AND MODEG
projected from lower
A drawing of the model and pertinent information are given in fig-
ure 1. The solid aluminum-alloy wing had I?ACA65AO06 airfoil sections
parallel to the fuselage center line, a quarter-chord line swept back
32.6°, an aspect ratio of 4, anda taper ratio ofO.6. The lateral-
control devices investigated included a spoiler, a slot-deflector com-
bination, and a spoiler-slot-deflector combination (fig. 1). The slot
consisted of an opening through the right wing between the 55- and
70-percent-chord lines extending spanwise from stationO.139b/2 to
h station 0.639b/2. In order to provide more strength in the wing, two
chordwise ribs of metal were allowed to remain. For the plain-wing con-
figuration, this slot was covered on the upper and lower surfaces of the
. wing with a l/16-inch steel cover plate. Spoiler projections were
obtained by raising the rear edge of the upper cover plate and bending
the plate along the 55-percent-chord line. Deflector projections were
obtained by lowering the forward edge of the plate on the lower surface
of the wing and bending the plate along the 70-percent-chord line. The
plates in the unprotected position were flush with the wing surfaces as
were the edges attached to the wing in the projected positions.
The model was n.muntedon a sting-t~e support system in the Langley
high-speed 7- by 10-foot tunnel. The sting was supported by a vertical
strut downstream from the test section. The support system allowed the
angle of attack of the model to be varied by rotating the model and sting
in the vertical plane about an axis through the quarter-chord point of
the wing mesn aerodynamic chord. The forces and moments on the model
were measured by means of electrical strain gages mounted inside the
aluminum fuselage. The fuselage ordinates are presented in reference 1.
TFSTs
.
.
Data were obtained at Mach nunibersof O.M, 0.80, 0.85, and 0.91,
but since no unusual Mach n re apparent for Mach numbers
4 wiiw?wwiiri NACA RML53D17
between 0.40 and 0.85, the basic data have been presented for only those z
two representative lkch numbers. However, rolling-moment coefficients
are presented for additional Mach nunibers. The angle-of-attack range
—
was from 0° to 20° except for the higher Mach numbers where the maximum .
angle of attack was limited by tunnel choking conditions. The spoiler-
alone was tested through a projection range from O to_=O.125c and the
deflector alone from O to O.1OC. Various combinations of eq=l and
unequal simultaneous projections of spoiler and deflector with slot open
.—
were
mean
tested.
The variation of test Reynolds number with Mach number based on the
aerodynamic chord is given in figure 2.
CORRECTIONS
I
The test data have been corrected for jet-boundary effects by the
method given in reference 5. Blockage corrections based on the model
with plain wing as determined from reference 6 to account for the con-
striction effects of the model on the tunnel free-stri%m flow were
applied to the data. T@e drag has been corrected by an increment
obtained by adjusting the”pressure at the base of the~uselage to eq~l
free-stream static pressure. No corrections for wi~” bending or twisting
have been applied since corrections as calculated from static loads on a-
the wing before the slot was cut were found to be small for the bending
-..
and negligible for the twisting of wing.
._
RESULTS AND DISCUSSION
The variation of lateral control characteristicswith angle of attack
for various projections of the spoiler alone is given_in figure 3. The
rolling moments produced by the spoiler in this investigation are con-
siderably smaller than those given in reference 1 for-the same spoiler”
span and projection. This difference in effectivenes-sas recoided In
the two investigations is caused mainly by the manner in which the pro-
jection was obtained as proved by unpublished results--oflater tests.
Maximum spoiler effectiveness is obtained when the spoiler is approxl-
.-
mately perpendicular to the wing surface. In this investigation, pro-
jection was obtained by raising the trailiti edge of the spoiler in an
operation very similar to the operation of a split flap. Except for
this difference in size of the rolling-moment coefficient the spoiler
in this investigation had characteristics very similar to the one in
reference 1. The rolling-moment coefficient decreased rapidly above En ““ . =
angle of attack of 8°, becoming zero or slightly ne~tive at 160 and above.
-.
.$!-..” %: :.. “*
NACA RML53D17 5
A recent low-speed (M x 0.3) investigation on a
6-percent-thickwing showed that some of the spoiler
ness at high @es of attack could be eliminated b.v
45° sweptback,
loss in effective-
means of a slot
and deflec;or (;ef. 2). Figure 4 shows the effect &n rolling-moment
coefficients of a slot and deflector without a spoiler. The figure
clearly shows the slot-deflector combination to be generally more effec-
tive at the higher angles of attack than at the lower angles except for
small projections of the deflector. The results obtained by combining
the spoiler with the slot and deflector are shown in figure 5. Various
ratios of deflector projection to spoiler projection were investigated.
Some of these projection ratios were more effective than others, but
almost all of them showed some effectiveness throughout the angle-of-
attack range.
The yawing-moment characteristicswere very similar for spoiler
alone, slot-deflector combination, and combination of spoiler, slot,
and deflector (figs. 3 to 5). They were positive or favorable at the
lower angles of attack and usually small and unfavorable at the higher
angles of attack.
Because the tests of the spoiler-slot-deflector combination are
considered the ones of more importance, the lift, drag, and pitching-
moment data of those tests along with the plain wing data are presented
in figure 6. The larger projections of the spoiler and deflector pro-
duced increments of positive pitching moment, but none of the projections
caused any significant change in stability as measured by the slope of
the pitching-moment curve, except the maximum projection at the higher
Wch number. The lift data indicate that the loss in lift is a function
of the amount of spoiler-deflector projection, the ratio of spoiler
projection to deflector projection, and the angle of attack. The drag
coefficients generally increased with increased projection of spoiler
and deflector. The drag coefficient of the spoiler-slot-deflector com-
bination is much greater than that of the spoiler alone at low angles
of attack (fig. 7), but at angles of attack of 8° and above, the drag
coefficients of the two configurations are essentially the same, although
the spoiler-slot-deflector combination produces much greater rolling-
moment coefficients.
Figure 8 shows the effect on rolling-moment coefficient of varying
the deflector projection while holding the spoiler at a given projection.
The given projections of the spoiler are -0.05c and -O.1OC. The figure
slmws that for angles of attack up to 8°, the maximum rolling-moment
coefficients are obtained by a deflector projection equal to one-half
the spoiler projection, but for angles of attack above 8° best results
are obtained with deflector projection about equal to spoiler projection.
Since there is considerable “Uss in effectiveness at zero angle of attack
for a ratio of deflector projection to spoiler projection of 1.0 to 1.0,
a ratio of 0.75 to 1.0 is probably the most advantageous throughout the
angle-of-attack range.
-
6
,.- ‘. “
NACA RM L53D17
The rolling-moment data of figure 9 show that the rolling effec-
tiveness of the spoiler-slot-deflectorcombination with the spoiler and
deflector projections numerically equal is nonlinear over the projection
rsmge but generally increases with increase in projection.
CONCLUSIONS
A wind-tunnel investigation was made through a Mach number range
from O.~ to 0.91 to determine the lateral control characteristics of
a wing-fuselage combination equipped with a flap-type spoiler hinged
at the 55-percent-chord line; a deflector hinged at the 70=percent-
chord line, and a slot in the wing between the tw hinge lines. The
span of the controls was 50percent of the wing semispan and the con-
trol was located inboard on the right wing. As a result of the inves-
tigation the following conclusions are made:..
1. As was previously found at low speeds, a spoiler-slot-deflector
combination is effective in producing rolling mmnents over a greater
angle-of-attack range than an unvented spoiler alone at Wch numbers _
up to 0.91.
2. The optimum ratio of deflector projection to spoiler projection
for rolling-moment effectiveness varies with angle of attack, but a
ratio of 0.75 to 1.0 gives appreciable effectiveness through the mgle-”
of-attack range from 0° to 20°.
3.A deflector and slot without a spoiler is also effective in
producing rolling moments at high angles of.attack.
.
—
—
—
.
1!
.
Langl~ Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va.
.
.
NACA RML53D17 7
REFERENas
1. Vogler, Raymond D.: Wind-Tunnel Investigation at High Subsonic
Speeds of Spoilers of Large Projection on EUINACA 65AO06 Wing
With Quarter-Chord Line Swept Back 32.60. NACA RML5U1O, 1952.
2. Watson, Janes M.: Low-Speed Lateral-Control Investigation of a
Flap-Type Spoiler Aileron With and Without a Deflector and Slot
on a 6-Percent-Thick, Tapered, 45°Sweptback Wing of Aspect
Ratio 4. NACARML52G1O, 1952.
3. %mnondj Alexander D., and Watson, James M.: Later@l-Control
Investigation at Trsnsonic Speeds of Retractable Spoiler andPlug-
Type Spoiler-Slot Ailerons on a Tapered 60° Sweptback Wing of
Aspect Ratio 2. Transonic-Bump Method. NACARML5m_6, 1952.
4. Wenzinger, Carl J., and Rogallo, Francis M.: Wind-Tunnel Investiga-
tion of Spoiler, Deflector, and Slot Lateral-Control Devices on
Wings With Full-span Split and Slotted Flaps. NACA Rep. 706, 1941.
5. Gill-is,Chrence L., Polhsmus, Edward C., and Gray, Joseph L., Jr.:
Charts for Determining Jet-Boundary Corrections for Complete
Models in 7- by 10-Foot Closed Rectangular Wind Tunnels. NACA
WR L-123, 1945. (Formerly NACA ARR L5G31.)
6. Herriot, JohnG.: Blockage Corrections for Three-Dimensional-Flow
Closed-Throat Wind Tunnels, With Consideration of the Effect of
Compressibility. NACA Rep. 995,1950. (Supersedes NACARMA7B28. )
41Ek%ami@im
8Wing Data
Am
Aspect fVtiO
ToPer mtio
NACA RM L53D17
45?2 >
32#sqin
4?0 /77-om.,1.
.6
Section A!ACA 65AO06
Spon 36.Oin
Root chord l/25in
Tip chord 6.75in +
Quarter -cfwd
sweepbuck 326°
~ 30.0
A II dimensions in inches
‘u
Sections A-A
_ Ploin wing j
8*=0, &=o
—
- Spoiler pm]ected,
8~”o
“-Spoiler off, ”
~eflector projected
Spoiler and deflector
projected
.
.
.-
.—
.
.
%
, , ,
. v *
4
o
x /06
—
.
/ ‘
/ ‘/
I
I 1
.4 .5 .6 l7 .8 .9 10
Mach number, M
!2
I,,
Figure 2.- Vmiation of Reynolds number with Wch number i
10 MACA RM L53D17
l
.04
.02
o
-.02
.02
C“ o
-.02
–J’s ad
@ercenf chgd)
o 2.5
0
; L
A /0 :
n /2.5 o
-4 0 4 8 /2 /6 20 24
Q, deg
(a) M=O.40.
Figure 3.- Variation of lateral control characteristicswith angle of
attack for various projections of the spoiler alone with slot closed.
.
.
.
NACA RM L53D17
.
11
.
.
.04
.(22
C[
o
-.02
.02
–& ad
(Percentchord)
O 2.5 0
a5 o
0 Z5 o
A /0 o
Q /2.5 o
702
-4 0 4
(b)
Figure
8 /2 /6 20
G?, d?g
M= 0.85.
3.-Concluded.
K?
*--
NACA RM L53D17
.04
.02
q
o
-.02
o off 2.5
•1 off 5
: a; 7.5
/0
.02 “
Cn o
-.02 —
-4 0 4 8 /2 /6 20 24
a, deg
(a) M =0.~.
Figure h.-,Variation of lateral control characteristicswith angle of
attack for various pro~ections of the deflector with_slot open end
spoiler off.
—
.
l
NACA RM L53D17 13
.04
.02
o
.02
0 off 2.5
q off 5
0 off Z5
A off 10
.
02
c~ o
-.02
-4 0 4 8 /2 /6 20
d?,deg
(b) M=o.85.
Figure 4.- Concluded.
NACARM L53D17
.04
.02
0
-.02
-.02
c“ o
-m
05 23
055
A5 Z5
A 25 2;5
-4 0 4 8 /2 /6 20 24
CY,deg
l
.
.
.
(a) M= O.40.
Figure 5.- Variation of lateral control characteristicswith angle of
attack for various projections of the spoiler-slot-deflectorcombination.
NACARM L53D17 15
8
.
.04
.02
0
.02
c“ O
-.02
f?“
-8S
(percent
o 10
q /0
o /0
A /0
n 25
8~
chord)
2.5
5
7.5
/0
Z5
-4 0 4
(a) M =
Figure
8 /2 /6 20 24
a, a’eg
0.40. Concluded.
5.- Continued.
16
,04
.02
o
-.02
.02
c~ o
-.02
r
NACA RM L53D17
.
.
-& ad
(percent chord)
05 2.5
05 5
A5 Z5
A 25 2.5
-4 0 4 8 /2 /6 20
a, deg
(b) M =0.@.
Figure 5.-Continued.
P
.
l
.
c
3Y
lUCARM L53D17 17
.
.
Cz
.06
.04
.02
+ ad
(pefcent chore’)
.02
c~ 0
-.02
T
I
-4 Q 4 8 /2 16 20
.
a, deg
(b) M = 0.85. Concluded.
Figure 5.- Concluded.
NACARM L53D17
./
c. o
-./
20
/6
/2
a)~g 8
4
0
-4-
05
Q5 7.5
Vo
n /0
&
4
.3
.2 CD
,/
o
-4 -.2 0 .2 4 .6 .8 /.0
(a) M= O.kO.
Figure 6.’-Effect of spoiler and deflector projections with slot open
on the aerodynamic characteristics in pitch.
cm
NACARM L53D17
-
.
.
.’/
C* o
-./
20
/6
/2
ff,deg8
4
0
-4
n5 75
VD
.3
.2
CD
./
o
(b) M = O.@.
Fi~ 6.-Concluded.
lWA RM L53D17
M
— 40
-——
.85
*
-—
0 4 8 /2 /6
a ~a’eg
Figure 7.-Ratio of drag snd roUing-moment coefficients of spoiler-slot-
deflector combination to that of spoiler alone at vagious wing angles
of attack. 6s = -10 percent chord, 8~ = 7.5 percent chord.
.
.
lUC!ARM L53D17 21
.02
-.02
.04
c“ .02
(?
.
M= .40
M= .80
hf=.85
M’.9/ - n
o 2 4 6 8
.04
.02 Cz
c?
o (de@
00
q 4
A8
o /2
V /6
n 20
.04
8d, @?fce~fCM
(a) b. = -5 percent chord.
Figure 8.-Effect on the rolling-moment coefficient of various ratios of
.
deflector pro~ection to spoiler pro~ection for the spoiler-slot-
deflector combination at various wing angles of attack.
22 NACA RM L53D17
.04
.02
0
-.02
.04
Cz .0.2
0
1 I I 1 I I I I I I I
0 2 4 6 8
(b) b~ = -10
Figure 8.-
percent chord.
Continued.
/0
—
-
.
-d
s.
I?ACARM L53D17 23
.06
.04
c’
.02
.06
.04
.02
6’
0 2 4 6 8 10
~~, percent chod
(b) b~ = -10 percent chord. Concluded.
Figure 8.-Concluded.
24 l’UICARML53D17
.04
.02
0
. fw
—.W&.
?04
M=B5
\
o
-.02
.
.
-/0 -8 -6 -4 -2 0
~~,percent Chrd
Fi&ue 9.-Variation of rolling-moment coefficient with conibinedprojec-
tions of deflector and spoiler in 1.0 to 1.0 ~tio for the spoiler- l
slot-deflector conibination. 5s = ‘bd.
.
NAcA-Lan@ey .6-2Q-tj3 -ss


Wind-tunnel investigation at high subsonic speeds of a spoiler-slot deflector combination on an NACA 65A006 wing with quarter-chord line swept back 32.6 degrees

Wind-tunnel investigation at high subsonic speeds of a spoiler-slot deflector combination on an NACA 65A006 wing with quarter-chord line swept back 32.6 degrees

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