Czarnecki, K R

Lord, Douglas, R

English

NASA Technical Reports Server

F. 4 CA //• 
RESEARCH MEMORANDUM 
LOADS ASSOCIATED WITH SPOILERS AT SUPERSONIC SPEEDS

By Douglas R. Lord and K. R. Czarnecki 
Langley Aeronautical Laboratory

Langley Field, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON
June 14, 1957 
Declassified September 1,1959
https://ntrs.nasa.gov/search.jsp?R=19930088820 2020-06-17T08:40:46+00:00Z
NACA PM L55E12a 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
LOADS ASSOCIATED WITH SPOILERS AT SUPERSONIC SPEEDS 
By Douglas R. Lord. and K. R. Czarnecki 
Sjii4 
The available information concerning the loads associated with 
spoilers at supersonic speeds has been reviewed and the effect of some 
of the more important variables in the problem has been considered. 
Although a large proportion of the data now available are fundamental in 
nature, they lend considerable basic knowledge to the study of spoiler 
loadings and permit some estimations to be made. 
This paper presents typical data available for various spoiler 
installations and presents an approximation method for estimating the 
loadings caused by an unswept spoiler. Some preliminary data and discus-
sion are also presented for spoilers yawed to the main flow. 
INTRODUCTION 
At the present time there is available only a limited quantity of 
experimenta l  data concerning the loads associated with spoilers in super-
sonic flow. During the past few years, however, several tests of a fun-
damental nature have been made which give some insight into the loadings 
caused by spoilers and enable a better understanding of the flow phe-
nomena involved. The majority of the data obtained to date are for 
spoiler installations in what might be termed idealized conditions, and 
the application of these results to three-dimensional lifting wings with 
their attendant spanwise variations will undoubtedly introduce new com-
plications. The present results are, however, a vital first step in 
understanding the flow characteristics and in developing methods for 
predicting the spoiler loads for an actual installation. The purpose 
of this paper is to present typical data from some of the most recent 
tests and to discuss the conclusions which have been reached to date. 
All data presented are for turbulent boundary layers. 
SCVIBOLS 
M	 stream Mach number 
MI	 local Mach number
2	 N&CA RM L55E12a 
p	 stream static pressure 
P1	 local static pressure 
q	 stream dynamic pressure
- 
P	 pressure coefficient, 
p	
p 
effective angle of flow separation ahead of spoiler 
1	 horizontal distance from top of spoiler to point of flow 
separation 
h	 spoiler height 
A	 sweep angle 
D	 spoiler section drag 
cd	 spoiler section drag coefficient, D/qh 
H	 spoiler section moment about base of spoiler 
Ch	 spoiler section binge-moment coefficient, H/c1h2 
MODELS AND TESTS 
Figure 1 illustrates the testing techniques which have been used 
in studying this problem. Although some of the tests were initiated as 
part of shock—boundary-layer interaction programs, they are inherently 
suited for studying the loadings ahead of spoilers. On each of the 
sketches shown, the horizontal lines above and below the diagram define 
the location of the tunnel walls. In the upper left sketch of the figure, 
the two-dimensional step technique, used both at the David Taylor Model 
Basin (ref. i) and at Princeton University (ref. 2), is shown. Orifices 
ahead of and on the front face of the spoiler were used to determine the 
loadings. 
The collar-on-a-tube technique, illustrated in the top middle sketch 
of figure 1, was employed in tests in a blowdown jet of the Langley Gas 
Dynamics Branch (ref. 3) . Here again, a single row of orifices was used 
to obtain pressures along the tube and on the front face of the collar.
NPCA RM L55E12a	 3 
Tests were made in the Langley 9-inch supersonic tunnel (ref. 4) on 
a two-dimensional airfoil (as shown in the upper right sketch of fig. 1) 
in which spoilers of various heights were mounted at three chord.wise loca-
tions on the airfoil upper surface. A single row of orifices along the 
surface measured the pressures both ahead of and behind the spoiler. 
The three techniques shown so far have been used primarily to study 
the effect of spoilers placed normal to the flow. In order to study the 
loadings caused by a spoiler in the yawed condition such as would be 
encountered on a swept-wing spoiler installation, with the simplifying 
condition of uniform flow ahead of the spoiler, the technique shown in 
the lower left sketch of figure 1 was used in the Langley ii-- by 4-foot 
supersonic pressure tunnel. Spoilers of varying height, span, and deflec-
tion angle were mounted on a turntable in a flat boundary-layer bypass 
plate. The turntable was instrumented with approximately 260 orifices 
located so that, as the turntable was rotated to obtain various sweep 
angles, rows of orifices were always so alined as to give detailed pres-
sure distributions in the streamwise direction. Pressures were obtained 
ahead of and behind the spoilers as well as on the front and rear faces 
of the spoiler itself. 
In the lower right-hand corner of figure 1 is a plan view of the 
three-dimensional semispan-wing model which was tested in the 1 by 4-foot 
supersonic pressure tunnel with various spoiler installations. A typical 
location for a full-span spoiler is shown, in addition to the five rows 
of orifices located across the wing span. Some of the variables inves-
tigated in these tests were spoiler height, span, chordwise location, 
and sweep.
DISCUSSION 
Unswept Spoilers 
In figure 2 the typical loadings caused by an unswept spoiler in 
supersonic flow are illustrated.. In the left part of the figure the 
pressure distributions ahead of and behind three spoilers have been super-
imposed. Two of the spoilers were vertical spoilers of different heights 
and the third was a flap-type spoiler deflected 45 0 to the surface. As 
has been previously demonstrated at subsonic speeds, when the loadings 
are plotted as a function of distance from the top of the spoiler in terms 
of the actual spoiler height above the surface, the loadings are nearly 
identical. Further investigations have shown that this remains true with 
reductions in spoiler height until the height becomes of the same order 
of magnitude as the boundary-layer thickness.
N&CA RM L55E12a 
Since it was found that, at this Mach number, the spoiler caused the 
flow to separate approximately	 spoiler heights ahead of the spoiler 
top, a simple approximation to the loading can be obtained, as shown, 
by following these rules: First, drawing a line from the separation 
point to the top of the spoiler determines the angle through which the 
flow must turn and, therefore, the average pressure in the forward dead-
air region can be calculated.. Further, from examining the pressure dis-
tributions
.
, it appears that at the spoiler top the flow expands through 
an angle approximately three times the 'value of the separation angle 
just determined. By calculation then, it is possible to get the pres-
sure immediately downstream of the spoiler. It would ordinarily be 
anticipated that this pressure would remain constant until the flow 
impinged on the surface and again had to be turned to realine with the 
stream. Experience has shown, however, that the compression takes place 
in a gradual manner and is approximately completed at a point downstream 
of the spoiler the same distance as the separation occurred upstream. 
A straight line connecting the last computed pressure point with the 
proper distance along the axis in the downstream direction, therefore, 
completed the approximation. 
On the right-hand side of figure 2 are shown the loadings on the 
front and rear faces of the three spoilers depicted. The vertical spoilers 
exhibit marked pressure increases on the front face near the bottom and 
top of the spoiler which indicate stagnation of the local flow at these 
points. The 150 spoiler has its highest pressure at about 80 percent of 
its height, followed by a rapid expansion. These variations are caused 
by the circulatory flow in the dead-air region ahead of the spoiler. On 
the rear faces of the spoilers, there appears to be little effect of 
spoiler height or deflection angle, and indications of any circulation 
are lacking. At the present time, no technique has been obtained for 
estimating the distribution of pressures along the front face of a spoiler; 
however, the uniform pressure on the rear face may be approximated by 
using the pressure obtained just after the expansion of the main flow 
at the spoiler top. 
If the experimental pressures are known on the wing surface ininiedi- 
ately ahead of and behind the spoiler, a good approximation of the aver-
age loads on the spoiler can be obtained by assuming that these pressures 
apply uniformly over the adjacent spoiler faces. 
Since, in the discussion of the technique used in estimating the 
loadings caused by a spoiler, it was necessary to first know the extent 
of the separated region, the obvious question which follows is how to 
determine this distance. In figure 3 the separation distance from the 
spoiler in terms of the height is plotted against Mach number. All the 
data on this figure were determined from the tests discussed in figure 1. 
The two-dimensional results are shown as symbols, whereas the results from 
the only available three-dimensional tests are shown as a shaded area.
NACA PM L55E12a	 5 
The data for any one family of test points in which there was a 
Mach number variation indicate a decreasing trend with Mach number. 
From consideration of the scatter of the available results, it appears 
that the assumption of a constant separation distance of 5 spoiler 
heights as shown by the dashed line would be satisfactory for estimating 
loadings in the Mach number range from 1.6 to 3.0. 
Effects of Sweep 
Up to now, only unswept spoilers have been discussed. Figure 4 
illustrates the changes in upstream influence of a spoiler caused by 
increasing the spoiler sweep from 0 0 to 600. In considering the effects 
of spoiler sweep, a new phenomenon is involved: The flow not only must 
be deflected by a shock from the surface to surmount the spoiler height, 
but a new shock is necessary to turn the flow along the surface - thus 
allowing the flow to move parallel to the face of the spoiler. 
It can be seen from figure 4 that, for an unswept spoiler, there is 
relatively little effect of the spoiler tips on the upstream influence 
of the spoiler within its spanwise boundaries. As the spoiler is swept, 
the curve of initial disturbance assumes the position of a detached shock 
about the upstream tip of the spoiler. 
This interaction is illustrated better in figure 5, in which the 
effect of 450 sweep on the pressure distributions in streaniwise rows at 
two stations along the spoiler span is shown. The upper diagrams show 
the loadings ahead of, behind, and on the spoiler at station 1; whereas 
the lower diagrams show the same variations at a station considerably 
closer to the spoiler tip. At 0 0 sweep, the loadings are almost identical 
at the two stations shown. When the sweep is increased to 450 2 the com-
pression ahead of the spoiler occurs in two steps and is separated by 
an expansion region. At station 1, the change in sweep from 0 0 to )o 
increases considerably the upstream influence of the spoiler. At sta-
tion 2, nearer the tip, the initial compression occurs much closer to 
the spoiler than it does at station 1; however, the peak of the expansion 
region appears to be about the same distance from the spoiler at both 
stations. The final compression ahead of the spoiler is much greater 
at station 2 and is also illustrated by the higher pressure along the 
front face of the spoiler at this atation. The pressures in the region 
downstream of the spoiler are generally more negative at station . 2 for 
this sweep condition. 
Since the variation in loading along the span has been shown, it 
follows that the integrated lift and pitching moment caused by the spoiler 
will also vary spanwise. Because of the limited number of stations across 
the span and the relatively low spoiler span-to-height ratio for these
6	 NACA RM L55E12a 
tests, it is impossible to give a complete picture of the spanwise varia-
tions of lift and pitching moment caused by a swept spoiler. Indications 
are, however, that when a spoiler having a sweep of 150 or greater is 
used, there is a strong tendency for reversal in lift and reduction in 
pitching moment within approximately 10 spoiler heights of the upstream 
tip. These indications have been borne out by correlations obtained 
between data from these tests and data from three-dimensional tests in 
the Langley 16-foot transonic tunnel (ref. 7) on a 450 swept-wing—spoiler 
combination in which the local wind velocities ahead of the spoiler were 
supersonic. 
In accordance with the discussion of the effects of spanwise loca-
tion on the changes with sweep, figure 6 shows the variations in spoiler 
section drag and flap-type spoiler section hinge-moment coefficients 
along the span for various angles of sweep. At 0 0 sweep, the drag and 
hinge moments are constant along the span insofar as was investigated. 
As the spoiler is swept to 60 0 , the drag and hinge moment first fall off 
on the downstream portion of the spoiler and then fall off all along the 
span.
CONCLUDING REMARKS 
Considerable information of a fundamental nature is now available 
on the status of research on loadings caused by spoilers at supersonic 
speeds. Estimations can be made very simply to determine the loadings 
caused by unswept spoilers in uniform flow fields. The effects of sweep, 
though understood somewhat, are still too complex to permit any simple 
approximation techniques to be demonstrated. It is anticipated that 
further analysis of the available information should clarify this problem; 
however, detailed loading investigations of spoiler installations on 
three-dimensional lifting wings will be needed for a complete solution 
to the problem. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Langley Field, Va., April 21, 1955.
NCA RM L55E12a	 7
REFERENCES 
1. Patterson, R. T.: The Characteristics of Trailing-Edge Spoilers. 
Part II - The Effects of Gap, Flap Deflection Angle, Thickness, 
and Sweep Angle on the Aerodynamic Characteristics of Two-Dimensional 
Spoilers, and the Pressure Distribution Near the Tip of a Partial-
Span Trailing-Edge Spoiler, at a Mach Number of 1.86 - TED No. TMB 
DE
-3109. Aero Rep. 827, David W. Taylor Model Basin, Navy Dept., 
Dec. 1952. 
2. Bogdonoff, S. M., and Kepler, C. E.: Separation of a Supersonic 
Turbulent Boundary Layer. Rep. 249 (Contract No. N6-onr-270), 
Princeton Univ., Dept. Aero. Eng., Jan. 1954. 
3. Lange, Roy H.: Present Status of Information Relative to the Prediction 
of Shock-Induced Boundary-Layer Separation. NACA TN 3065, 1954. 
4. Mueller, James N.: Investigation of Spoilers at a Mach Number of 1.93 
To Determine the Effects of Height and Chordwise Location on the 
Section Aerodynamic Characteristics of a Two-Dimensional Wing. 
NACA EM L5201, 1953- 
5. Hallissy, Joseph M., Jr., West, F. E., Jr., and Liner, George: Effects 
of Spoiler Ailerons on the Aerodynamic Load Distribution Over a 
Sweptback Wing at Mach Numbers From 0.60 to 1.03. NACA EM L54Cl7a, 
1954.
,r.SPOILER M"I.61 
M2.OI
ru NkCA RM L55E12a 
TEST CONFIGURATIONS 
,—COLLAR 
MI.86	 M-3.03 
*M3
(IM ML93,	 -SPOILER 
STEP 
DAVID TAYLOR MB 35 LANGLEY 
*PRINCETON GAS DYN. LAB.	 GAS DYN. JET	 LANGLEY 9" SST 
LANGLEY 4 SPT	 LANGLEY 4 SPT 
Figure 1 
SPOILER LOADING
M 1.61 
°
-o-REAR 
FACE 
C 0 
P / c APPROX.	 / P 
.2 10 THEORY—' .2 
.4 . - 4 
.6 .6 '/ FRONT 
'tFACE 
.8 ----UPSTREAM— —DOWNSTREAM--- 8 
-4'	 -2 0	 2	 4 6	 8 0 50	 100 
DISTANCE IN SPOILER HEIGHTS PERCENT SPOILER 
HEIGHT
Figure 2
NkCA EM L55E12a
EFFECT OF MACH NUMBER ON SEPARATION
A=O'
TWO-DIM. TESTS 12-M1
h	 0 DAVID TAYLOR MB 3"x5" 
T	 a PRINCETON GAS DYN. LAB. 
G LANGLEY GAS DYN. JET 
LANGLEY 9" SST 
8	 4 LANGLEY 4 SPT 
I 
I
THREE-DIM. TESTS 
LANGLEY 4 SPT 
O	 I	 I	 I	 I 1.2	 1.6	 2.0	 294	 2.8	 3.2
M1 
Figure 3 
UPSTREAM INFLUENCE OF SPOILER 
M"I.61 
I,	 1• 
SPOILER	 SPOILER	 SPO$LE\ 
A=O*	 A"30°	 A"60° 
Figure 4
10 NACA RM L55E12a 
EFFECT OF SWEEP ON SPOILER LOADING 
SPAN/HEIGHTI4 MI.6I 
-	
rOIL/: I 
------2
	
Ar45°
	
REAR 
o A=O* 	
FACE 
:
	
0 
P	 STATION I	 P 
.4	
FRONT 
FACE 
REAR 
.8-
-.4zr C Ar 4r
_7^" 
P 
.4F	
A^^ STATION  
8L —UPSTREAM_I DOWNSTREAM —'-

DISTANCE IN SPOILER HEIGHTS 
Figure 5 
SPOILER DRAG AND HINGE MOMENT
SPAN/HEIGHT 14; MI.6I 
Z44L
SPOILER 
Cd 
.2-	 Ch
A, DEG 
0 
.8-	 - CA
----------30 
.4 -
60 
C	 I	 I	 I	 I	 I 
.6	 A, DEG 
0 
.4 
Ch	 30 
.2
	
__60 
I	 I	 I	 I	 I	 I 
0	 20	 40	 60	 80	 100
PERCENT SPOILER SPAN 
Figure 6
NACA -Langley


Loads associated with spoilers at supersonic speeds

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