Dynamic stability of slender cones subjected to high oscillation amplitude - measurement in wind tunnels by free flight technique

Dayman, B., Jr.

English

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FREE-FLIGHT CONE DYNAMIC STABILITY TESTING 
AT 
HIGH AMPLITUDES OF O~SCILLATION 
GPO PRICE $ 
CFSTI PRICE(S) $ 
Hard copy (HC) a 
Microfiche (MF) - 
ff 653 July 65 
ITHRUI 
N65 - 294 13 
L ICODCI 
To Be Presented at the 22nd 
Semi-Annus1 STA Meeting i n  
Brussels, Belgium, September 1964 
, 
Jet Propulsion Laboratory 
California Institute of Technology 
Pasadene, California 
https://ntrs.nasa.gov/search.jsp?R=19650019812 2020-03-17T02:00:46+00:00Z
INTRO WCTION 
This paper presents a summary of information from several published 
reports and tests recently completed which deal with the measurement of 
dynamic stability of cones at high amplitudes of oscillation by use of wind 
tunnel free-flight techniques. 
are very briefly mentioned as they have been described in adequate detail 
elsewhere. However, considerable damping data are included in order to 
fully document the value and consistency o f  the rerults. 
The techniques of testing and data reduction 
T- 
The cone models (0.5 t o  1.0 inch d im. )  are constructed of th in  p las t ic  
or metal (aluminum or  magnesium) shells and are ballasted with either lead 
or gold cores. 
The models are launched in to  free-fl ight t ra jec tor ies  across the view- 
ing area of the  wind tunnel by ei ther  of two ways. They can be supported on 
a ver t ica l  wire (at any desired i n i t i a l  angle of attack) at the upstream 
edge of the viewing window and released in to  flight by rupturing the wire 
at a notch within t h e  model. 
by a pneumatic launch tube located downstream of the window, giving tra- 
jector ies  both upstream and downstream across the window. 
are used fo r  supporting the model on the launcher. 
'at tack (e 40 deg) they can be supported on a wedge within the  model base. 
For high angles of attack (up t o  120 deg) the models can be supported i n  
Models can be propelled aginst the airstream 
Two basic methods 
Up t o  moderate angles of 
a cradle and are launched i n  a m e r  similar t o  that used by a shot- 
put ter  at a t rack meet. 
The model motion is recorded on high-speed (2OOO-5000 frames/sec) 
half-frame 35 mm motion picture film using conventional high-speed cameras. 
Back-light (e i ther  si lhouette o r  schlieren) is used t o  outline the  model. 
Although steady l i gh t  can be used, i n  order t o  e l imina te  model motion during 
f i l m  exposure (60 p sec or longer), use is made of multi-flash ( ~ - 1 O O O  
flashes) short duration (2 p sec) strobe light, synchronized t o  the camera. 
Even though the model motion I s  normally confined t o  the ve r t i ca l  plane 
and, consequently, requires only a single camera, a second one is used t o  
record the motion i n  the  horizontal plane. 
-3- 
Each model has many cycles of osci l la t ion during i ts  viewing-window 
tradectory ( w i t h  a large number of high quali ty pictures being obtained fo r  
each cycle) and the motion is appreciably damped. 
typical angle-of-attack history for a cone during a dynamic s tabi l i ty  test .  
I n  Figure 2 the osci l la t ion amplitude of another run is plotted aginst  the 
re la t ive  airstream distance in the manner used f o r  data reduction. 
Figure 1 presents a 
For further details on model design and construction along with addi- 
t ional  information on the launching equipment and data acquisition, 
References 1-3 w i l l  be quite helpful. 
cribed i n  Reference 4. 
Data reduction methods are des- 
-4- 
EXAMPLES OF DYNAMIC STABILITY DATA 
Even though t h e  sole  purpose of the test  may be for  the measurement 
of model dynamic s tab i l i ty ,  the model drag and pitching moment slope are 
generally obtained during the data reduction. The effect  of osci l la t ion 
amplitude on drag is shown i n  Figure 3. Data from a recent test  (wire- 
release, August 1964) are compared t o  the extensive data of Reference 5 
where t e s t s  (gun-launch) were run during April 1964. The comparison is 
q u i t e  satisfactory fo r  flat as w e l l  as hemlspherical afterbodies on the  
sharp-nose, 10-deg half -angle cones . 
Figure 4 shows the cone pitching moment slope as a f'unction of Mach 
number. 
af'terbody does appreciably affect the pitching moment. 
Reference 6 were used fo r  the theoretical  curve. 
Contrary t o  the drag comparison, the  effect of the  hemispherical 
The cone tables of 
An example of a large amount of dynamic s t a b i l i t y  data (taken from 
Reference 5) is shown i n  Figure 5. 
are good. 
damping coefficient at various amplitudes of osci l la t ion.  The increase of 
damping w i t h  osci l la t ion amplitude f o r  the theoret ical  case is quite  small 
even though the  increase is appreciable for  Newtonian damping at increasing 
The data consistency and repeatabil i ty 
Also shown i n  t h i s  figure I s  the Newtonian solution for  the 
loca l  angles of attack. Because the Newtonian solution is nat sensit ive 
t o  osci l la t ion amplitude, the Newtonian case w i l l  be considered only at  
zero amplitude i n  the succeeding figures. 
The effect  of cone center-of-gravity location on the damping coeffi-  
c ien t  I s  shown i n  Figure 6. The experimental data behave in  the manner 
predicted by the various theories -- Tobak (Reference 7) and Newtonian. 
-5- 
Figure 7 presents cone damping through a Mach number range (2 < M < 6) 
as a function of oscil lation amplitude. 
with Tobak is adequate but not ideal since the experimental and theoretical  
curves are not parallel. 
A t  zero amplitude the comparison 
Qynamic s t a b i l i t y  data from Reference 5 are transferred t o  the condi- 
t ions of t h e  August 1964 free-flight test for  presentation in  Figure 8. 
Here again the recent data compare favorably with the previous data. 
hemispherical afterbody on the  cone models does not a f fec t  the damping, 
with or without boundary layer t r i p .  
t o  give turbulent cone boundary layer and wake at  zero angles of attack. 
The 
The t r i p  has been shown (Reference 5) 
Limited tests were performed a t  M = 6 in  order t o  compare the damping 
In Figure 9 the  comparison of trends of cones with different apex angles. 
with Newtonian is shown t o  be quite good. 
A t  M = 2 and 4 several flights were made w i t h  flat-based cones blunted 
t o  a nose radius t o  base radius ra t io  of 0.2. 
s ignif icant ly  decrease the damping, even more so than predicted by Newtonian 
theory (see Figure 10) . The location of the model center of gravity, i n  
respect t o  the base, was the same for  the blunted and sharp cones. 
This amount of blunting did 
-6- 
SUMMARY 
Enough dynamic stability data for slender cone models have been pre- 
sented in order to demonstrate the usef'ulness of this aspect of free-flight 
testing in conventional wind tunnels. Data were obtained for cone angles 
of oscillation up to 30 deg amplitudes. Additional wire-release and gun- 
launch tests have been performed on a series of cone models up to initial 
oscillation amplitudes of 90 deg at M = 4, 6, 8. Preliminary data analysis 
indicates test results as good as those presented in this paper. 
-7- 
REFERENCES 
1. Daymn, Bain Jr., "Simplified Free-Flight Testing i n  a Conventional 
Wind Tunnel", Technical Report No. 32-346, Jet Propulsion Laboratory, 
Pasadena, California, October 1962. 
2. Holway, H. P., J. G. Herrera, and B. Dayman, Jr., "A Pneumatic Model 
Launcher for  Free-Flight Testing i n  a Conventional Wind Tunnel", 
Technical Memorandum No . 33-177, Jet Propulsion Laboratory, Pasadena, 
California, March 1964. 
3 Dayman, Bain Jr , "Definitive Interference-Free Experimental Studies 
of Vehicle Motion", AIAA Preprint 64-476, (Presented at the 1st AIAA 
Annual Meeting, Washington, D. C ., June-July 1964). 
4. Jaffe, Peter, "Obtaining Free-Flight Dynamic Damping of an Axially 
Symmetric Body (At A l l  Angles of Attack) i n  a Conventional Wind Tunnel", 
Technical Report No. 32-544, Jet Propulsion Laboratory, Pasadena, 
California, January 1964. 
5. Jaffe, Peter and Robert H. Prislin,  "Effect of Boundary Layer Transi- 
t ion  on Dynamic Stabi l i ty  over Large Amplitudes of Oscillation", 
A I M  Preprint 64-427, (Presented at  the 1st AIAA Annual Meeting, washington, 
D. C., June-July 1964). 
6. Sims, Joseph L., "Tables fo r  Supersonic Flow Around Right Circular 
Cones at Small Angle of Attack", NASA SP-3007, Washington, D. C., 
1964 
7 Tobak, Murray and Wil l iam R .  Wehrend, "Stabil i ty Derivatives of Cones 
at Supersonic Speeds", NACA Technical Note 3788, September 1956. 
NOMENCLATURE 
2 A model reference area, nd /4 
B.L. boundary layer 
Total drag drag coefficient; c = 
cD D SA 
‘m 
Pitching moment 
SAd 
pitching moment coefficient; Cm = 
pitching moment slope per radian 
cy 
dynamic damping-in-pitch coefficient per radian m. + c  
Q cy 
c* 
d 
f 
HS 
L 
M 
Q 
r 
r 
B 
N 
V 
x C 8  
x 
(assumed constant over cycle) 
model base diameter (reference length) 
frequency of oscillation (cycles/sec) 
hemispherical 
model length (for  blunted models, length taken as that for 
sharp-nose model) 
freestream Mach number 
f reeet ream dynamic pressure 
m o d e l  base radius 
radius of nose bluntness 
freestream velocity 
distance of center of gravity aft  of nose (for  blunted models 
from sharp nose before blunt lng) 
model position relative t o  air flow 
freestream Reynolds number based on model base diameter 
NOMENCLATURE (CONT.) 
cyenv 
angle of attack 
oscillation amplitude 
effective (average) oscillation amplitude over a complete trajectory i 
= 262 - 2  env 
root-mean-squared angle-of-attack during a complete trajectory 
cone half-apex angle 
model angular velocity 
'3 a 
ln 
m 
0 
c- 
l-i 
0 
1 .  
I 
0 , .  . . . . . .  
$: ~ d d d d d d d d d 0 O 0 0 0 0 0 0  
I .  
-121 I I 
loo ZOO 
' M - 4.54  
X,JL = 0.48 
60 cpr 
I 4600 homer 'sec 
I I 
) 400 $00 600 
FRAME NUMBER 
Fig. 1 Example of Gold-Core Cone Gun-Launch 
Fig. 2 Oscillation Envelope as a Function of Model Travel Relative 
to A i r  Flaw 
am 
a- 
ceo 
I D 0  
0.m 
0.10 
u = I O  deg 
rN/rB=O I 
I 
Jaffe-Prislin Free-Flight 
0 JPLmm-FLICHp 
(August 1964 Test ) 
sw Taw D I A . ~ .  
0 0  
A 0.00s 
0 0.010 
D 0.015 
0 0.020 
I I I 
Y = 3.02 Y =2.01 I .  
I I I I 
a00 400 eo loo -* 100 
-1.4 
- 1.2 
z a 
E 
2 -1.0 
\ 
-0.8 
I I I 1 0- 
0 m I5 
I I 
10 I 5  10 
Fig. 3 Effect of Oscillation Amplitude on Cone Drag 
Q = 106 
rN/rg = o 
-0.6 
0 I 2 3 4 5 6 7 
MACH NUMBER 
Fig. 4 Cone Effective Pitching Moment Slope 
+ 
0 
E" 
-0.5 
O >  
IV I d  LV 
-2.5 
-1.5 
-1.0 
rN/rB = O  A 0.005 
0 0.0 IO 
0 0.0 I5 FLAT BASE 
0 0.020 
I 1 
- - 
82, deg2 
Fig. 5 Free-Flight Cone Dynamic Stabi l i ty  (Jaffe-Prislin) 
;" 
E" 
0 + 
0 
rN /r8 = 0 
u=10 deg 
-FLAT BASE 
1 I 
0.6 
OB I 
0.4 0.5 
L (FROM NOSE) 
X C d  
Fig. 6 Effect of Center-of-Gravity Location on Cone Dynamic Stabi l i ty  
Fig. 7 Effect of Mach Number on Cone Dynamic S tab i l i ty  
f -1.5 
- I s o  
-0.5 
JPL 'FREE- FLIGHT I 
0 -HEMISPHERICAL BASE. 
OPEN - M=4 - (FLAT BASE) SOLID-M=2 
TAG-WITH E L  TRIP 
SHADED CURVES A -FLAT BASE 
JAFFE-PRISLIN AT M=2,4  
I I I J 
Fig. 8 Effect of Oscillation Amplitude on Cone Dynamic S tab i l i ty  
c 
-6  
- 4  
.U 
E 
0 + 
+ 
-1.0 
- 0.8 
- 0.6 
-0.4 
Fig. 9 Effect of Apex Angle on Cone Dynamic S tab i l i ty  
0 0.05 0. IO 0. I5 0.20 0.25 0.30 
"1'0 
Fig. 10 Effect of Nase Bluntness on Cone Dynamic S tab i l i ty  


Free-flight Cone Dynamic Stability Testing at High Amplitudes of Oscillation

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