Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

SUPERSONIC AXIAL STAGE SEPARATION OF AN UNGUIDED ROCKET

Richard E. Kreeger *
U.S. Army Missile Command
Missile Research, Development and Engineering Center
Redstone Arsenal, AL

NeilR. Walker**
Nichols Research Corporation, Inc.
Hunstville, AL

Abstract D,ef = reference length = 4.50 in.

M = free stream Mach number
A primary concern during axial stage Re = free stream Reynolds number
separation of an unguided rocket is that the close a = angle of attack, deg.
proximity of the booster may destabilize the OCT = total angle of attack, deg.
payload section. A wind tunnel test was 6ra = relative motor case pitch angle, deg.
conducted to verify the payload stability and cp = roll angle, deg.
quantify the stage separation interference. It was cpm = relative motor case roll angle, deg.
assumed that the full booster section could be X/D = axial separation distance, cal.
approximated by a section of the motor two Z/D = normal separation distance, cal.
calibers in length. The simulated booster
hardware could be moved on the model support Configurations
sting to various locations axially and vertically
relative to the payload. The booster could also be DB1 = large drag brakes
rotated to different angles in the pitch plane DB2 = small drag brakes
about the model centerline. Aerodynamic data Fl = baseline payload fins
were acquired at Mach numbers 1.5 and 1.8 to F2 = alternate payload fins
determine the interference of the motor case on Ml = 1-cal. motor case section
the payload section during axial stage separation. M2 = 2-cal. motor case section
Pressure data for the forward face of the motor N1B1 = payload section nose and body
and base of the payload were measured to N4B4 = airdrop nose and body
estimate the drag of each after separation.

Nomenclature I. Introduction
Symbols A technology demonstration program
2 was conducted by the U.S. Army Missile
Aref = reference area = 15.9043 in. Command in 1995-96 to demonstrate the
CA = axial force coefficient dispense of a submunition simulant from an
CAP = forebody axial force coefficient unguided rocket. Shortly after apogee, the
CAU = uncorrected axial force coefficient payload section of the rocket deployed four grid
C, = rolling moment coefficient fins to provide post-separation stability. After fin
cm = pitching moment coefficient deployment, a linear charge cut the rocket skin
Q, = yawing moment coefficient circumferentially and staging occurred. The
CN = normal force coefficient payload continued on a ballistic trajectory until
CPB = payload base pressure coefficient the onboard sequencer initiated a gas generator
CPU = motor case head pressure coefficient device to eject the submunition. Three prototype
CY = side force coefficient test flights were conducted; the last of which
staged and successfully dispensed a
* Aerospace Engineer submunition.
** Member of the senior technical staff, Senior Initially there was concern that the close
Member AIAA proximity of the booster during staging would
This paper is declared a work of the U.S. Government and is destabilize the payload section and possibly
not subject to copyright protection in the United States
cause tumbling or coning. To mitigate this risk, a

1
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 Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

pyrotechnic piston device was added forward of shown in Figure 2. Table 2 details the locations
the motor to impart a relative separation velocity. of the various axial settings used during the stage
Passive devices, drag brakes on the booster, were separation phase of testing. Figure 3 illustrates
also considered. Because of cost and time the nomenclature for the parameters X/D, Z/D,
constraints on the program, a full Captive and 6. The axial, vertical, and angular
Trajectory System (CTS) wind tunnel test could displacements were tested relative to the sting
not be accomplished. axis and the aft end of the payload. The varied
It was then assumed that the parameters are shown in Table 3.
interference of the full rocket motor section
could be approximated by a two caliber (in
length) section of the booster. A wind tunnel test
was conducted to verify the payload stability,
stage separation interference, and relative
aerodynamic drag of the separating payload
section in proximity to the spent booster.
The wind tunnel test had multiple
objectives relating to the preliminary design of
the flight hardware and the modelling and
simulation of the upcoming flights. The major
objectives of the test are shown in Table 1.

______Table 1. Test Objectives______

• Confirm the payload fin design met stability

and axial force requirements
• Obtain data to develop a freestream
aerodynamic model of the payload section
Figure 1. Assembly Drawing of Payload
• Quantify the difference in ballistic coefficient
Section with Grid Fins
and relative drag of the separating bodies
• Size an appropriate drag brake design for the
During the second phase of the test, the
motor section
motor case with attached drag brakes was fixed
• Determine the effects of motor interference on to the metric payload section. In this manner, the
payload stability during the first 2.0 calibers of incremental axial force of the drag brakes could
the separation event be determined.

B. Test Description
II. Experimental Apparatus
The aerodynamic stability and stage
A. Test Article separation wind tunnel test was conducted at the
National Technical Systems (NTS) 4x4 Trisonic
Wind tunnel model hardware consisted of a Wind Tunnel in Rye Canyon, CA in October,
sting-mounted payload section (N1B1F1) and a 1995. The wind tunnel is a blowdown-to-
2.0-caliber representation of the rocket motor atmosphere facility with a free stream Mach
case (Ml). The payload section consisted of a number range from 0.2 to 5.0.
4.0-caliber blunted tangent ogive, a cylindrical
section 3.103 calibers in length, and four grid Table 2. Axial Stations
fins mounted in the '+' configuration. The grid
fins provided stabilization, and were not
X/D, calibers Xmodel, inches
deflected. The payload section is shown in 0.01 0.045
Figure 1. Both the payload and motor models 0.15 0.675
had a diameter of 4.50 inches and were 50.35% 0.25 1.125
scale. 0.50 2.250
The simulated motor case could be moved to 1.00 4.500
2.00 9.000
various locations and attitudes relative to the
payload section. The stage separation model is

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Table 3. Tested Parameters Table 4. General Test Description

Parameter Phase Scale Mach Motor Payload

X/D, cals 0.01,0.15,0.25,0.5,1.0.,2.0 Number
Z/D, cals 0.00,0.25,0.500 1 0.5035 0.5-2.0 None N1B1F1
A9, deg -10,-5,0,5,10 II 0.5035 1.3 Ml N1B1F1
<|>m, deg 0,45 III 0.5035 1.5, 1.8 M2 N1B1F1
Mach number 1.5,1.8 IV 0.3600 0.5-1.1 None N4B4

The model was mounted on a MICOM

sting and sting adapter using the NTS 16-inch
roll pod. The main balance (NTS MKXXXIX)
was located at model station 20.499 inches. Four
200-lb fin balances were mounted at model
station 31.031 inches, corresponding to the
centerline of the fin mounting bases. Note that
model station 0.00 was located 1.455 inches
forward of the physical model nose.
All coefficients presented in this paper are in
the body-fixed non-rolled axis system (OCT, q>)
and all dimensions are in model scale. Base
pressures were measured on the payload section,
as were pressures on the forward face of the
motor case by means of static taps.
Figure 2. Photograph of Assembled Model III. Results and Analysis

A. Grid Fin and Payload Aerodynamics

The payload fins were sized and

designed based on five main requirements. These
were an axial force requirement, approximately
one caliber static margin after staging, roll
- X--- - STING
damping characteristics, packaging within
available volume, and deployment with a
minimum of mechanical complexity.
INCOMING FlOW •
Thus, with the exception of the AV
required during staging, a high drag was actually
a desirable attribute for the payload fins, in order
to decelerate to the desired dispense velocities.
Grid fins were selected for the payload because
of their packaging, deployability, and axial force
characteristics.
Figure 3. Sign Conventions for Stage Grid fins are known to have a reduction
Separation in effectiveness in the transonic region; this was
quantified during the test. Grid fin roll damping
The test encompassed 105 hours of air-on characteristics were found to be adequate in
time and was divided into four phases. Phase I another test. Thus, the only unknown was the AV
was the grid fin downselect and payload free consideration.
stream test. Phase II was the booster drag brake The baseline grid fin design was
size selection. Phase III was the stage separation confirmed during the test. The payload with the
test. Phase IV concerned an airdrop grid fins deployed met the stability and axial
configuration which was not part of the rocket force requirements. Predictions of grid fin
launch program. A summary of the test aerodynamics were very close to the measured
description is given in Table 4. data. The free stream axial force CAP and static

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 Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

margin of the payload section as functions of separation event. For this case, the base pressure
Mach number are shown in Figure 4 and 5, on the payload and head pressure on the motor
respectively. section were of primary interest. It was found
The payload section required fins for that the net axial component of the aerodynamic
stabilization during and after separation. It was forces would initially inhibit staging. The
assumed that the booster and warhead remain at forward face of the booster, a significant
low angles of attack following skin cutting and contributor to the relative AV, would be shielded
that the payload fins would be fully deployed at by the payload section. In addition, the booster
the time of staging. section was found to alter the flowfield around
As previously mentioned, the grid fins the aft body of the payload section and to affect
were designed to be in a stowed position prior to the pressure in the interstage region differently as
launch, and to remain within the rocket moldline the booster was moved aft.
until 0.1 seconds prior to staging. This meant It was found that the payload base
that a cavity would be exposed after deploying pressure coefficient, CPB, did not yet reach
the fins. The effect of the grid fin cavity on the freestream values when the booster had moved
freestream data is also shown in Figure 4. aft 2.0 calibers at Mach 1.5 or 1.8. Figure 6
shows the payload section base pressure
coefficient CpB versus axial separation distance
X/D in calibers for an in-line separation. Of
equal importance was the pressure coefficient on
the forward face of the booster CPH, which did
not begin to increase until after 0.25 calibers of
axial separation, and had only reached ~ 0.32 by
X/D=2.0 calibers. Figure 7 shows the motor case
face pressure coefficient CPH versus axial
separation distance X/D, in calibers.
0.40 In Figure 8, the axial force coefficient
0.00 0.50 1.00 1.50 2.00 2.50 of the payload section is plotted versus X/D for
the in-line separation, with the freestream axial
Mach Number force at a=0° of the payload superimposed.
There is basically a region of impeded separation
Figure 4. Payload Axial Force versus Mach until the motor case has moved aft about 1
Number caliber, and the payload section still has not
reached its freestream axial force as far aft as 2
0.0 calibers. The in-line staging showed negligible
-Xcp impact on the static stability of the payload
(D -1.0 section, as expected.
U) -Xcg
o
-2.0 C. Payload Stability During Staging
E
o
-3.0
The presence of an offset booster at
| -4.0 Mach numbers 1.5 and 1.8 was found to cause an
upward shift in the Cm vs a curve and a
§ -5.0 downward shift in the CN vs a curve of the
payload. In some cases, the slope of the payload
Cm curve is reduced, but this effect is not large.
0.0 0.5 1.0 1.5 2.0 2.5 For small angles of attack, these effects are
Mach Number generally constant. The effect of various booster
locations on pitching moment coefficient at
Figure 5. Payload Static Margin versus Mach Mach 1.5 is shown in Figure 9. This figure also
Number shows the relative magnitude of the effects of
vertical, axial, and rotational offsets of the
B. In-Line Separation booster.
The main impact of the motor case or
Initial testing was conducted to large drag brakes appears to be on the afterbody
determine the effects of an ideal, or in-line, stage of the payload. The interference manifests as a

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 Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

drag brakes is worsened when interdigitated with

the grid fins on the payload. In this position, it is
possible for the motor case or large drag brakes
to make the payload unstable at small angles of
0.05 -____»^- - _ _ _ I _ _ _ _ _ ! _ _ _ _ . _ _ attack. This indicates that the grid fins help to
JT*~~' ———* Mach 1.8 reduce the interference of an object behind them.
0.00 r ""<«^-^i | "^--^J^ It can also be seen from Figure 9 that Z/D is the
CO -0.05
_ _ _ _ _ _ _ _ _ _ . _ _ _ . . i . ._Mach.1.EL~~"r-^~^^.
0.
largest contributor to flowfield interference on
0 -0.10 the payload section.
i
-0.15 Blach 1~.8~Ffee! Stream r From Figure 10, CN vs X/D, and Figure
11, Cm vs X/D, it can be seen that the stability
-0.20 - - - - - - Macrr tS-Free1 Stream - - - - - - - -
o
- o
- fi)
- dtT ( influence decays sooner than the axial force
n 95
interference as the booster moves aft. These
00 0.5 1.0 1.5 2.
figures also illustrate that booster interference
X/D effects are only prominent for X/D's from 0.15
to 0.5 or 1.0. Even then, only the largest Z/D
Figure 6. Payload Base Pressure versus X/D showed any large impact on the payload.
Figure 12 illustrates the effect of
moving the offset booster out of the pitch plane
with a plot of yawing moment coefficient at zero
0.35 angle of attack versus axial separation distance
for Z/D=0.5 calibers at Mach 1.5. These data
confirm that the stability interference due to a
laterally offset booster will be limited to the
plane in which the booster is offset. In addition,
the notable similarities between Figures 11 and
12 are immediately apparent; in particular, both
effects die out at around one caliber of axial
separation.
The base pressure on the payload is
0.0 0.5 1.0 1.5 2.0 much higher on the side to which the booster is
X/D offset, also tending to cause a nose-down Cm.
However, the moment arm at this location is
Figure 7. Head Pressure versus X/D small, as well. Figure 13 shows motor case head
pressure coefficient CPH versus X/D for various
values of Z/D at Mach 1.5.
: '
The fin balance data do not appear to
o./ ' • "V,^ " ~ A ' Mach Mi-Free Stream- ------ 1 vary as drastically as the main balance ACm. The
n r*
° Mach lS_Free Stream- shifts observed in fin balance data are also
0.5 - constant with total angle of attack, and are
: i Maqhl.8 consistent with causing a nose down pitching
O 0.4 - • i i
<J. O
moment increment. However, they are very
'- i i
0.2 small compared with the total increment
: i i
0.1 observed, which is nose-up.
nn .
' ! I From this, it was determined that the
0.5 1.0 1.5 2.0
grid fins on the payload remained for the most
0.0
part outside of the influence of the motor case.
X/D The incremental ACm due to a motor case offset
is shown in Figure 14 for a case with and without
Figure 8. Payload Axial Force for In-Line grid fins on the payload body. From the general
Separation trends of these curves, it can be seen that for
small angles of attack the grid fins were not
decrease in the body normal force and an appreciably affected by the booster downstream.
increase in pitching moment. The slopes change
only slightly (only a small change in static
margin). The effect of a translated booster or

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 Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

_ Free Stream
-X/D=0.15Z/D=0.0 _X/D=0.15 Z/D=0.5 THETA=+10 deg
_X/D=0.15Z/D=0.5 -X/D=0.15 210=0.25 THETA-5 deg
_X/D=0.15 Z/D=0 w/Drag Brakes -X/D=0.15 Z/D=0.5 w/Motor At 45 deg

-0.10
-0.50
0.00 1.00
-5.0 -4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0 3.0 4.0 5.0
Angle of Attack, deg

Figure 9. Payload Pitching Moment Figure 12. Effect of cpM on Payload Cn at

Coefficient versus aT at Mach 1.5 Mach 1.5

0.10

-0.40
0.00 2.00 0.00 0.50 1.00 1.50 2.00 2.50

X/D

Figure 10. Effect of Z/D on Payload Normal Figure 13. Motor CPH versus X/D at Mach 1.5
Force Coefficient at Mach 1.5, OCT=O°

0.50

0.30

0.10
E E
O O
-0.10
. Fins, Z/D=0
-0.20 -0.30 - No Fins, 7.ID=0
-Fins,Z/D=0.5
. No Fins, Z/D=0.5
-0.40 -0.50
0.00 0.50 1.00 1.50 2.00 -6.0 -3.0 0.0 3.0 6.0 9.0 12.0

X/D Angle of Attack, deg

Figure 11. Effect of Z/D on Payload Pitching Figure 14. Payload ACm Increments at
Moment Coefficient at Mach 1.5, aT=0° Mach 1.5, X/D=0.25 Calibers

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D. Drag Brakes offset of the booster at small axial separation

distances. Lateral offset of the booster had a
Various methods of imparting the required noticeable effect on payload aerodynamics,
AV to the payload/booster system were including base pressure and main balance
considered. These included active methods, such coefficients, but only small changes appeared in
as interstage jet flow, mechanical springs and a the fin balance data. In fact, to some degree,
pyrotechnic piston. Passive methods, such as interference effects were almost independent of
adding drag brakes to the booster or deflecting whether the payload fins were deployed or not.
the motor tail fins, were also considered. The small drag brakes located on the booster
Two drag brake designs were tested. were found to have a minimal effect on the
Each design consisted of four drag brakes payload.
deployed on the booster aft of the forward face, At Mach 1.5, pitch-plane interference
and interdigitated with the payload fins. The had all but disappeared at one caliber separation,
large drag brake design was 3"x3" full scale, and while effects in the axial plane (base pressure,
the smaller set was 2.25"x2". Drag increments of motor-case face pressure, wake effects) appear to
the two designs at Mach 1.5 are 0.7467 and continue well beyond the two calibers covered
0.3167, respectively, compared to a freestream by the test.
payload CAP of 0.4821 and a freestream booster
CAP of 1.530. (Booster freestream data were V. References
obtained in a later test.)
The small drag brakes had minimal 1. Hammer, R.L. and Brillhart, R.E.,
effect on the payload stability. However, as Aerodynamic Data for Several Two-Stage
shown in Figure 15, the large drag brakes Missile Configurations with Variable Stage
showed an effect on the payload section similar Separation and Without Jet Flow at Mach
to a motor case offset. Numbers 2.5, 3.5, 5.0, AEDC-TR-62-132,
In the flight tests, it was found that a Arnold Engineering and Development Center,
pyrotechnic piston could reliably impart TN, (1962).
sufficient impulsive AV. The drag brake designs
were not used. 2. Anon., Air Flow. Altitude and Acoustics
Facilities, National Technical Systems Rye
Canyon Research, Development and Test Center
Facilities Handbook, (1994)
Large Drag Brakes
Small Drag Brakes
3. Landingham, et.al., A Preliminary
Examination of an Axial Dispense Concept for
the BAT Submunition, TR-RD-SS-94-11, U.S.
E
O Army Missile Command, AL, (1994)

4. Kreeger, R.E., Pre-Test Report for the BAT-

on-a-Rocket Aerodynamic Stability and Stage
-1.25 Separation Wind Tunnel Test (Test A). LR-RD-
-10.0 -5.0 0.0 5.0 10.0 SS-95-20, U.S. Army Missile Command, AL,
(1995)
Angle of Attack, deg

Figure 15. Effect of Drag Brakes on Payload

Stability

IV. Summary of Results

Staging effects on the stability of the

payload section were smaller than expected. The
interference created biases in the normal force
and pitching moment coefficients, but did not
destabilize the payload. The major driver of the
observed stability changes was the large lateral

7
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