Supersonic wind tunnel test program on turbulent flow separatio

Lowson, M. V.

English

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WYtE LABORATORIES - RESEARCH STAFF 
TECHNICAL MEMORANDUM 65-4 
SUPERSONIC TURBULENT SEPARATION TESTS IN THE 
MARSHALL SPACE FLIGHT CENTER 7" TUNNEL 
Submitted on Contract NAS8-11308 
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1 .O INTRODUCTION 
Prediction of the acoustic environments beneath regions of separated flow and oscillating 
shocks i s  severely limited by the lack of experimental information on these flows. At the 
present time it i s  not possible to make realistic estimates of even the mean flow occurring in  
the separation region. In  Reference (1) i t  was shown how separated flows on flares at 
supersonic speeds could be directed into three classes as shown in Figure (1). Each class 
i s  associated with a different flow pattern, and different fluctuating pressure environments 
may be expected for each type. Once the mean flow type can be defined, further prediction 
i s  practicable on the basis of available data. However, at present there are not sufficient 
experimental results to determine the flow type from known vehicle conditions. 
Therefore, there i s  a definite requirement for further experimental work and this Memorandum 
has been prepared at the request of MSFC to specify the experiments and to give a detailed 
design for the models required. The experimental tests outlined below should provide sufficient 
information to define criteria separating one flow type from the other, and so should enable 
considerably improved acoustic environmental predictions to be made. In addition, these 
tests should have value i n  the prediction of vehicle loadings and stab’ility. 
2.0 DISCUSSION 
The types of flow can be classified with reasonable accuracy merely by identification of their 
shadowgraph flow characteristics. Therefore, as a first stage, it i s  suggested that the tests be 
f:xf:ted to a shadowgraph ffow vl-;l;n!izot;on - Recent tests at Ames have przvlded extensive 
data on the steady and fluctuating pressures caused by various supersonic flare induced turbu- 
lent separations, but unfortunately the Ames tests did not include shadowgraph information. 
The present tests could therefore be regarded as an extension of those at Ames, and wi l l  cer- 
tainly greatly assist interpretatior of that data. The present tests could also be combined with 
surface oil-flow visualization, and at a later stage the models could be instrumented for 
static pressure measurements. I t  would be desirable to make pressure fluctuation measurements, 
but development of a microphone suitable for the low intensities expected ( ~ 1 1 0  dB due to the 
low density) would cause difficulty. In view of the data already available from Ames i t  i s  
not recommended that pressure fluctuation measurement be attempted at present. The critical 
parameter for these tests i s  the Reynolds number based on undisturbed boundary layer thick- 
nesses, as shown by Kuehn in Reference (2). For values of this Reynolds number below 8 x 10 
he found a significant variation of separation length with Reynolds number. However, for 
higher values the flow seems to reach ar: assymptotic state. It i s  therefore essential to run 
these tests at a sufficiently high Reynolds number, and this puts a limitation on the minimum 
boundary layer thickness, as shown in  Table I below. 
4 
TABLE I 
Ci-iAiZACTERiSTiCS OF THE M.S.F.C. 7" WIND TUNNEL 
Minimum S (in.) Typical Tunnel Floor 
for Re 
4 Boundary Layer 0 
Mach Reynolds Number 
Number Per Inch 
8 x  10 Thickness (in.) &O 
1.54 360,353 
1.99 305,185 
0.22 
0 -262 
0.2 
0.3 
2.44 250,000 0.32 0.32 
3 .oo 185,608 0.433 0.3 
The boundary layer thicknesses given i n  Table I would require body lengths of a foot or more 
for their natural generation. For many reasons i t  i s  undesirable to artificially increase the 
thickness of the boundary layer 50 that this body length requirement rules out the possibility 
of using axisymmetric bodies in the 7" tunnel. However, i t  wi l l  be observed that the tunnel 
wall boundary layer thickness i s  very close to that desired. Thus an alternative would be 
to use a two dimensional model mounted on the tunnel wall, but this becomes impractical in 
terms of tunnel blockage at the larger step heights required for the tests. In  addition, the 
field of view of the shadowgraph in  the 7" tunnel does not include the wall, thus involving 
further complications. 
In v i e w  of the limitations discussed above i t  i s  suggested that the tests be made a semi-circular 
section cylinder-flare-cylinder half model mounted on the wall. This type model could uti l ize 
the wall boundary layer and thus satisfy the boundary layer thickness requirements. The top 
surface would be raised above the tunnel floor and thus come into the field of view of the 
shadowgraph. Also, shadowgraphs of axisymmetric bodies are generally clearer than equivalent 
two dimensional cases. Finally, the semi-circular section of the model wi l l  minimize the 
tunnel blockage occurring at any given step height. 
The wall boundary layer i s  about 0.3" in thickness 8" i n  front of the nozzle trailing edge where 
separation i s  anticipated. Previous tests have shown that below step heightboundary layer 
thickness ratios of unity the flow patterns cannot be readily interpreted because of the small 
scale. Equally, Kuehn's tests i n  Reference 2 at step heightboundary layer thickness ratios above 
8 always revealed a ''flare reattachment" case for sufficiently low flare angles. Accordingly 
flare step heights of 0.3, 0.6, 1.2 and 2.4 inches have been chosen for the present tests giving 
setp height/undisturbed boundary layer thickness ratios of 1,2,4pnd 8 respectively. 
The range of flare angle for the present tests wi l l  clearly need to cover the range between 
the just separated case of is0 and the forward facing step of 90°. Particular interest attaches 
to the lower value of flare angle since this i s  more typical of the flares on large launch 
vehicles, and also since the maior effects of flare angle are expected i n  this lower range. 
Thus, flare angles of 15,20,30,45,60 and 90 degrees are suggested for the present tests. 
The Mach number in the MSFC 7" Tunnel i s  de:ermined by fixed nozzle liners. The available 
Mach numbers lying i n  the present range of interest are 1.5, 2.0, 2.5 and 3 (approximately): 
Tests at Mach 1.5 with the largest step height w i l l  not be possible because of tunnei blockage. 
Practical considerations may also limit other projected tests. For instance, the low density 
of the Mach 3 flow may prejudice successful shadowgraph flow visualization. However, i t  i s  
known that useful shadowgraphs can be obtained at Mach 2.5 so that this problem i s  not 
expected to be severe. 
3.0 THE PROPOSED MODELS 
In order to accomplish the test requirements set out i n  the above discussion i t  i s  proposed to 
make the cylinder-flare models in two parts. Separate cylinders wi l l  be required for each 
nozzle liner, but the flares wi l l  be constructed to fit over +he cylinder in  any liner. Thus, 
the following models wi l l  be required: 
4 basic 1 "  diameter half cylinder models to fit Mach 1.5, 2, 2.5 and 3 nozzle liners. 
24 flare models (0.3, 0.6, 1.2, 2.4 stepheights) x (15', 2Oo,45O, 60°,900 Flare Angles). 
A typical model i s  shown, in  position on the tunnel wall, in  Figure (2). (The particular flare 
model depicted i s  0.6" step height, 30' flare angle). The models must be mounted on the 
removable piate shown in  Figure (2) since no hdes can be drilfed in the nozzle liner. The 
removable plate lies at an angle to the nozzle axis so that the basic half cylinders wi l l  have 
to be contoured to fit the nozzle wall and the angle of this plate. The half cylinder mates 
~ : i t h  the wa!! so that some .:are w:st be +?ken wiih +he F i t ,  particlflarly siwe i t  i s  desired 
that the wall boundary iayer fiows over the cylinder. 
Not a l l  of the nozzle contours have been received. Thus a complete definition of the actual 
dimensions to be used in the shaping of the base of the half cylinder has not been made. The 
height of the cylinder a b v e  the tunnel wall i s  fixed by the shadowgraph requirement that the 
top of the model appears i n  the field of view. Measurements by MSFC staff give the following 
dimensions for the distance of the bottom of the window above the tunnel liner (in inches) 
M =  1.5 2.0 2.5 3 .O 
Ea st 0.3 0.54 0.42 0.61 
West 0.35 0.59 0.16 0.57 
It i s  therefore suggested that the top of the half cylinder model be positioned at a height of 
0.7" above the nozzle liner datum. Details of the half cylinder model are shown in  Figure 
(3) 
On the other hand, the flare models mate with the cylinder and the only other requirement i s  
that they clear the nozzle wall. Their bottom surface has therefore been designed to clear the 
liner for the M=3 case, and in  general there wi l l  be a gap between the lower surface of the 
flare and the tunnel wall. The position of the flare shoulder has been chosen 50 that 
separation wi l l  occur approximately 8" in  front of the nozzle trailing edge. Each model 
terminates as a backward facing step 6" behind the nozzle trailing edge. The dimensions 
of the flares are given in  Table 1 1 .  (all dimensions are in inches). 
Distance Shoulder 
to rear of Model 
ri *. 
TABLE II 
DIIU'L' :SIONS OF TRUNCATED CONE/CYLI NDER "FLARE" MODELS 
Step Height 
External Radius 
Length of 
f lare from 
shoulder to 
front of Model. 
15O 
20O 
45O 
6oo 
30° 
190. 
0.3 1 0.6 
0.8 
1.120 
0.824 
0.520 
0.3 
0.173 
0 
12.75 
6.75 
1.1 
2.239 
1.649 
1.039 
0.6 
0.346 
0 
11.75 
5.75 
1.2 
1.7 
4.479 
3.297 
2.079 
1.2 
0.693 
0 
9.5 
3.5 
2.4 
2.9 
8 957 
6.594 
4.157 
2.4 
1.386 
0 
5 .O 
-1  .o 
The fixings for these models are shown in  Figures (3,4, and 5). The half cylinder model, 
Figure (3), i s  attached to the back plate by two 1/4 inch diameter counter sunk bolts with 
nuts on the reverse side of the plate. The flares, Figures (4 and 5), are attached by 1/4 
inch diameter bolts which screw in  from below the plate into tapped holes i n  the flare. It 
i s  proposed to make a l l  the models from brass and to use steel bolts. Because of the dimensions 
of the fixture required to hold the flares the rear part of the 0.3" step height flares has had to 
be widened. Thus these smallest flares have a second shoulder positioned 6" from the rear 
of the model at the nozzle trailing edge. Advantage has been taken of this to make t h i s  
second shoulder provide a seccnd test and each second shoulder i s  constructed with a step 
height and flare angle equal to that occurring ot the first, this i s  shown in  Figure (5). Since 
the second shoulder i s  about 25 step heights behind the first i t  i s  not expected to interfere 
i n  any way with the earlier flow. 
Note that Figures (3,4 and 5) cre reduced scale vzrsions of drawings available at Wyle 
Laboratories. It proved impossible to obtain reliable dimensions for the actual nozzle liners 
used at MSFC. Therefore, mould lines for the liners have been taken directly from the 
nozzle blocks. These mould lines are available at Wyle as drawing E 65349. 
4.0 PROPOSED TEST SCHEDULE 
The test schedule laid out below wi l l  only require two half cylinders and 18 flares for the 
first 33 tests. A total of 94 tests i s  suggested. Each of these tests wi l l  need to be only 
of short duration for the shadowgraph visualization. However, those tests for which o i l  flow 
visualization i s  desired wi l l  require somewhat longer run times. The proposed boundary layer 
tests wi l l  require the use of a rake already available at MSFC. 
Test Number Mach Number Step Height Flare Angles i n  Degrees 
1 1.5 None-Boundary Layer Rake 
2-7 1.5 0.3 15,20,30,45,60,90 
8-1 3 1.5 0.6 
14-19 1.5 1.2 
(in .) 
I1 I 1  
I 1  I 1  
20 
21 -26 
27-32 
33-38 
39-44 
45 
46-51 
52-57 
2.0 
2 .o 0.3 
2.0 0.6 
2 .o 1.2 
2.0 2.4 
2.5 
2.5 0.3 
2.5 0.6 
58-63 2.5 1.2 
64-69 2.5 2.4 
70 
71 -76 
77-82 
83-88 
89-94 
3.0 
3 .O 0.3 
3 .O 0.6 
3.0 1.2 
3 .O 2.4 
None-Boundary Layer Rake 
15,2O,30,&,60,90 
I t  I t  
I I  I 1  
I 1  I 1  
None-Boundary Layer Rake 
15,20,30,45,60,90 
I 1  I 1  
None-Boundary Layer Rake 
1 5,20,30,45,60,90 
I 1  I t  
I t  I 1  
REF ERE NCES 
1. Wyle Laboratories Research Staff: Quarterly Progress Report for January 
February and March, 1965. Contract NAS8-11308. 
Kuehn D.M., "Turbulent Boundary Layer Separation Induced by Flares 
on Cylinders at Zero Angle of Attack", NASA TR-Rll7. 
2. 
Figure la  
Figure l b  
Reattachment Region 
Expansion Fan 1 
Figure IC 
Reattachment Shock 
Figure 1: Types of Turbulent Supersonic Separated Flow 
+ 
r 
\ 
POINT TO BE RELATIVE TO 
I 
I 
1 
20.20 APPROX. 
FlGUR 
11 
i 
TOLERANCE 
DE: X X  f.03 X X X  ,015 
,266 DIA HOLE 
CSK.80.X S O  DIA 
HUNTSVILLE FACILITY 
CY L i NDER MODEL 'A'' 
DRAWN &s-c-ca SCALE JOB NO 51700 
FULL D65323 
I L.200 
'-NOZZLE T R A I L I N G  E D G E  E 1 3 M  3WG E65349 
I.SURFACE FINISH 9 2  / NOTES: - 
Z.-X-&'Y- DIN. TO BE OBTAINED FROM 
DWG E 6 5 3 4 9  
. 
FRONT VIEW 
ITEM I 
M I L L  1. TAP fOR 1/4-20 1 
4 PLACES .375 DEEP 
BOTTOM V I E W  
ITEY I 
* 
I -  
. 
\TO BE DETERMINED 
mon DUG ~ 6 5 ~ 4 9  
FRONT VIEW 
ITEMS 2 THRU 6 
BOTTOM VIEW 
ITEMS 2 THRUC (ALL DIM SAME AS ITEM I EXCEPT AS NOTED) 
I 
TO BE DETERMINED 
BY DWG E65343 1 &{- _ _  :_ _ _  _ _ _  _-e -i - \I - 
A* 
FIG 
FRONT V l f W  
ITEM 2 WRU6 
m u  z THRUG (ALL DIM. SAHE AS ITEN I UNLESS NOTED) 
TOLERANCE 
IEC: X X f . 0 3  X X X 2 0 l S  
- - 
I - - - - - - - -  1- 
DRAWN m - w  SCALE JOB NO 51700 
FULL D65324 
WTOM VIEW 
i -  ITEM I 
SECTION A-A 
LONT VIEW 
XQLL DRILL A TAP COR 114 -20 UNC 
4 PLACES 3 7 5  DEEP f 
80R 
NOTES: 
I. SURFACE FINISH 
W Y L E  LABORATORIES 
HUNTSVILLE FACILITIES 
 RE 5.  


Supersonic turbulent separation tests in the Marshall Space Flight Center 7 in tunnel

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