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DESIGN AND TEST RESULTS ON AGENA 2000 .pdf


Original filename: DESIGN AND TEST RESULTS ON AGENA 2000.pdf
Title: Design and test results on Agena 2000 - A high performance turbopump fed 15,000 LBF thrust storable bipropellant rocket engine

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Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

DESIGN AND TEST RESULTS ON AGENA 2000
A HIGH PERFORMANCE TURBOPUMP FED 15,000 LBF THRUST
STORABLE BIPROPELLANT ROCKET ENGINE
Samuel Wiley§
Linda Andersen
Aerojet
Sacramento, CA 95813

Edward Gribben
Richard Driscoll t
Mayne MarvinJ
Atlantic Research Corporation
Niagara Falls, New York
Abstract

plus the major engine subassemblies. Also presented
are summaries of the test results from the
injector/thrust chamber water flow calibrations and
the Individual hot fire test series.

There is a need to develop a high performance, low
cost, turbopump fed storable bipropellant rocket
engine in the 12,000 to 20,000 Ibf thrust range for use
on upper stage expendable launch vehicles. This
paper describes the work performed for Lockheed
Martin Astronautics by Team Agena, an
ARC/Aerojet Joint Venture, to evolve the Agena
2000 as an ideal blend of an extremely successful and
reliable product and modern day design-to-cost
(DTC) technique. The engine design builds on the
361 successful flight heritage of Agena

Introduction
Team Agena was awarded a contract from Lockheed
Martin Astronautics to design, develop, and test a
high performance turbopump fed 15,000 Ibf thrust
storable bipropellant rocket engine as part of their
EELV family of launch vehicles. The work for the
Pre-EMD phase of the program was initiated in
January 1997 and completed in May 1998.

The Agena 2000 engine has successfully completed a
rigorous series of hot fire tests culminating in a 60
second duration prototype engine demonstration test
conducted at the Aerojet test facility in Sacramento,
CA, on January 21, 1998. Included hi the paper is a
design description of the Agena 2000 engine system,
Jui

Jun

Workhorse
Injector Test
Series

Aug.

To accomplish the required program milestones in a
compressed schedule environment, a building block
test approach was employed as shown in Figure 1.
This technique allowed testing to occur hi a logical
sequence while building hardware in parallel with the
Oct

Sept

Timothy Fischer
Lockheed Martin Astronautics
Denver, CO 80127

Nov

Jan

Dec

Reconfigure
»- Propeilant Lines to
Run Bootstrap

Gas

Generator
Test Series

1'
Revised
Assemble
Pressure-Fed
Pressureinjector to
TPA onto
Fed TCA
———> Pressure- -> TCA Assembly ->
^ Pressure/Skid Assembly
Test Series
Fed Config
Fed TCA
n

Regen
Chamber
Fabrication

^"T"™^
PressureFed GG/TPA
Test Series

Bootstrap GG/TPA
Test Series

ik

1 Turbopump
J Fab and

Reconfigure Pump
Discharge Lines to
Pressure-Fed TCA
Inlet

Figure 1. Building block test approach

* Program Manager, Senior Member AIAA
t Engineering Manager, Senior Member AIAA
t Project Engineer, Senior Member AIAA

§ Systems Engineer, Member AIAA
f Project Engineer, Member AIAA
# Project Engineer
1

Copyright @ American Institute of Aeronautics and Astronautics Inc., 1998. All rights reserved.

Prototype Engine
Test Series

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

required test activities. The first test series was with
a flightweight design aluminum injector mechanically
bolted to a carbon steel heat sink chamber. Since the
chamber was uncooled, test durations were limited to
1-3 seconds. The primary objectives of these tests
were to obtain early injector performance data and
determine the stability characteristics of the acoustic
cavity ring design. The heavy wall chamber was
designed to accommodate 6.5 and 13 grain bombs,
which were triggered to initiate an instability if
present.

Design Description

The Agena 2000 Main Axial Engine (MAE) is an
updated version of the Agena Model 8096 engine
which demonstrated a mission reliability of 99.7%
over 362 flights from 1960 to 1987. The Agena 2000
MAE, shown in Figure 2, is based on the low cost
features of the original Agena, modernized through
use of current manufacturing techniques,
incorporation of updated flight proven hardware and
components, and upgrading of performance based
upon an extensive technical data base and technology
demonstration programs. The Agena 2000 MAE
performance is compared with the production Model

Upon completion of this series, a revised aluminum
injector was welded into a flight-weight,
regeneratively-cooled thrust
, WC Actuators From RL-1Q
chamber, installed into the test
skid assembly, and tested
?ump Suction Valve From Ariane
TTA From Del
pressure-fed using facility valves
and propellant tanks. While this
"Thrust Chamber From Agena
Injector From
test series was hi progress, a
From Transtar,
Agena
turbopump assembly (TPA) with
U/R OME
a state-of-the-art inducer,
Gas
impeller and turbine rotor was
Generator
fabricated and mated with an
From Agena
existing Model 8096 Agena gear
Turbopump from
box. The skid assembly was
Agena, Transtar
designed to accommodate the
andXLR-132
installation of the TPA to the
TCA at the completion of the
pressure-fed tests so that the
Nozzle Extension From Delta
hardware conversion process was
straightforward. A gas generator
(GG) assembly, based on the
original Agena design, was also designed, fabricated
Figure 2. Agena 2000 Engine System
and tested in the August 1997 time frame.
8096 Agena performance hi Table 1. The thrust has
been reduced slightly and the chamber pressure has
The GG was installed on the TPA and tested
been increased to 718 psia to minimize weight and
pressure-fed to calibrate the GG/TPA system. The
envelope. To increase specific impulse, the nozzle
propellant lines to the GG were then reconfigured to
extension has been increased to 275:1 and the
operate the engine in the bootstrap mode. These tests
propellant
combination has been changed from HDA
were run with the pump discharge lines plumbed into
(44%
N
O
facility catch tanks, allowing the GG/TPA system to
2 4 + 56% HNO3)/UDMH to N2O4/MMH.
The
pump
impeller has also been changed from an
be hot fire tested and calibrated prior to testing the
open
to
a
shrouded
impeller design for greater
thrust chamber assembly.
efficiency.
Upon completion of the GG/TPA tests, the
turbopump discharge lines were plumbed into the
thrust chamber assembly and the planned prototype
engine tests were successfully completed.

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

Table 1. Agena Engine Performance Comparison
Agena 2000

Thrust, Ibf
Chamber Pressure, psia
Mixture Ratio
Specific Impulse, sec
Nozzle Area Ratio
Oxidizer
Fuel
TCA Mixture Ratio
GG Mixture Ratio
GG Chamber Pressure, psia
Turbine Speed, rpm

15,170
718
1.83
336
275
MON-3
MMH w/Silicon Oil additive
1.89
0.14
475
24,800

The Agena 2000 MAE is a gas generator cycle
turbopump fed engine. The MAE hydraulic
schematic is shown in Figure 3. The Agena 2000

MAE retains the original Agena design feature of an
aluminum thrust chamber with drilled coolant
channels fed by impeller pumps which are driven by a
single stage impulse turbine in a fuel rich gas
generator cycle. The turbine exhaust gases are ducted

Model 8096

17,000
545
2.69
300
45
HDA
UDMH w/Silicon Oil additive
2.94
0.15
480
24,800

back into the main flow of the nozzle, eliminating the
need for an external exhaust duct. The ducted
exhaust gases reduce the nozzle skin temperatures
and allows the use of titanium as the nozzle extension
material. The helium spin start system, with gas
supplied by the stage, provides a multiple start
capability for the engine.

. Ox Pump Seal Drain
^
SUS Mechanical interface

Ox Pump
Main and
Purge Valve

Ox Iniet

Fuel inlet
Fuel Pump Sea! Drain
Fuel Pump
Main and
Purge Valve
GG

XO

3

Turbine Start
Helium inlet ^L^^^^J^^^^J^^^—^ Valve
I
I
. Stage / Engine Hydraulic Interface

Figure 3. Agena 2000 Engine Hydraulic Schematic

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

The thrust chamber is regeneratively cooled by the
oxidizer. Another feature of both the production
Agena and Agena 2000 engines is the use of silicon
oil (SO) to reduce the heat flux to the aluminum
chamber wall. The silicon oil, dissolved in the fuel,
decomposes in the chamber and burns to form silicon
dioxide, which adheres to the wall. This thin coating
reduced the heat flux and increases thermal/structural
margins. The silicon oil, hexamethyldisilizane, is
perfectly miscible in MMH over a wide temperature
range.

injector and chamber designs are based upon heritage
from the original Agena.

TPA Bracket

The Agena 2000 MAE consists of six major

subsystems, as shown hi Figure 4. The following
sections will provide a short description of each
subsystem.
Thrust Chamber Assembly

Fuellnlet Nector

The MAE thrust chamber is shown in Figure 5 and
consists of an aluminum injector combustion
stabilized with acoustic cavities and a regeneratively
cooled deep drilled aluminum chamber. Both

Chamber

Chamber
iniet

Figure 5. Agena 2000 Thrust Chamber Assembly

Agena 2000
Engine
System

Thrust
Chamber
Assembly

•Injector

Thrust
Take Out
Assembly

Turbopump
Assembly

• Ox Pump

•Chamber

• Fuel Pump

•Covers

• Turbine
• Gear Box
8

Drive Train

0

Mounting
Bracket

Engine
Control
Component

• Interface
Plate

• Pump Suction

0

* GG Biprop Valve

Monoball
Gimbal
•TVC
Actuators

Valves

E,ect rical
Hame

sses

Altitude
Kit

* Power

• NCM

• Instrumentation

• Nozzle
Extension

• Turbine Start
Valve
• Ox Pump Seal
Purge Valve
• Filters
• Orifices

• Venturies
• Manifolds
• Seals

Figure 4. Agena 2000 Engine System Configuration

• Gaskets

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

The MAE injector, shown in Figure 6., is a grid-type
configuration that uses a high performing triplet
element for the core flow and an unlike doublet for
the barrier elements. The triplet element consists of
two canted fuel streams impinging on one axial
oxidizer stream. The barrier elements consist of one
axial oxidizer stream impinging on one canted fuel
stream. The barrier elements are located around the
periphery of the injector and provide a fuel rich
barrier along the chamber wall. The injector oxidizer
feed holes are drilled parallel to the injector face in a
cross-grid pattern with the injector orifices drilled
into the feed holes from the injector face. The fuel
orifices are drilled from the injector face into feed
holes that are drilled between the oxidizer feed holes
from the back of the injector. An injector back plate
closes off the fuel feed passages.

7. The divergent chamber section from an area ratio
of 7 to 12 is cooled with a two-pass hole pattern.

0I2J4

Core Triplets
Acoustic
Cavity

Fuel
Inlet

Barrier Doublet
Ox Cross
Drilled Feed
Holes
Fuel Holes

Orifice Plate

Barrier Doublets
Acoustic
Cavities
Figure 6. Agena 2000 Injector

Figure 7. Agena 2000 Thrust Chamber
Turbopump Assembly

The prototype turbine pump assembly (TPA), shown
in Figure 8, consists of a single stage impulse turbine
driving two centrifugal pumps through simple spur
gears. The turbine and pump shafts are supported by
ball and roller bearings, which are lubricated by a
slinger located in the gear case sump oil reservoir.
The fuel and oxidizer pump assemblies include pump
suction inducers and shrouded radial vane centrifugal
impellers. The shrouded impeller with swept back
vanes has a negative slope head-flow curve which is
beneficial for mixture ratio control and hydraulic

The acoustic cavity ring, located around the periphery
of the injector, provides for high frequency stability
and uses 15 dual tuned resonator cavities: ten deep
cavities tuned for the first tangential mode and five
shallow cavities tuned for the third tangential/first
radial modes. The cavities and webs are cooled by
the flow of oxidizer around and through the resonator
parts into the injector.
The chamber, shown in Figure 7, is fabricated from
one piece of forged 2219 aluminum and incorporates
two regeneratively cooled sections. The forward
section has two different deep drilled cooling hole
patterns: longitudinal along the barrel section of the
chamber and a helical pattern drilled at 42.5 degree
angle to the chamber centerline through the
converging-diverging throat section to an area ratio of

Figure 8. Agena 2000 Prototype TPA

stability. The one piece single stage partial admission
turbine rotor is driven by nozzles in the turbine inlet
manifold which are fed by a fuel-rich gas generator.
The turbine and impeller designs, scaled form the
Aerojet family of small, storable turbopumps are
higher efficiency units than the original Agena TPA
designs.
Figure 9 shows a comparison between the prototype
and flight versions of the TPA. The prototype turbine
pump assembly utilized an existing gear box, pump
and turbine shafts, bearings and seals, and turbine
inlet manifold from an original Model 8096 Agena
engine to reduce fabrication time and cost.
Heavyweight pump and bearing support housings
were used for the same reasons. The flight TPA will
incorporate an optimized turbine inlet housing,
lightweight pump and bearing support housings, and
new bearings, seals, shafts and gear box housing.
Oxldlzer Pump
Assembly \

Ughter Weight Pump and
Bearing support Housings

Fuel Pump
Assembly
G«arBox

New Light Weight Turbine Inlet Housing

Figure 9. Comparison of Flight and Prototype TPA

The gas generator (GG), is a direct adaptation of the
original Agena GG with minor modifications. The
GG injector consists of four core triplet elements
(two fuel streams impinging on a central oxidizer
stream) surrounded by eight fuel barrier doublets.
A baffle located at the mid-chamber enhances the
mixing of the fuel-rich combustion gases.
Thrust Takeout Assembly

The thrust takeout assembly (TTA) provides for
MAE-to-Stage mechanical attachment, thrust takeout
and pitch/yaw capability with thrust vector control
actuator attachment points. The TTA consists of a
thrust takeout plate and a monoball gimball joint.

Control Components

The Agena 2000 control components provide for the
control of propellants, purging, and spin start
capability. The major control components consist of
suction valves, GG control valves, and the spin start
helium valve.
Electrical Harnesses

The Agena 2000 MAE requires four electrical
harnesses: two that provide primary and secondary
power to the engine valves, one that provides
instrumentation power and signal returns and one for
the thrust vector control actuator power and control.
Altitude Kit

The Agena 2000 altitude kit consists of a cast Inconel
625 nozzle coolant manifold (NCM) and a titanium
nozzle extension. The NCM ducts the turbine
exhaust gases back into the TCA nozzle flow at an
area ratio of 12. The exhaust gases are accelerated
through 2-D nozzles to supersonic conditions and
injected along the nozzle wall into the main TCA
flow which cools the nozzle extension. The nozzle
extension expands the MAE combustion gases to an
area ratio of 275.
Agena 2000 Testing

All hot fire testing of the Agena 2000 engine and its
subassemblies was conducted at Aerojet's J-4 test
facility hi Sacramento, CA. Water flow testing of the
injector and TCA was conducted using the original
production Agena water flow stand at ARC's Niagara
Falls, NY facility. Hot fire testing was conducted
using facility valving and manifolding on a test skid
assembly assembled at ARC.
Injector/Thrust Chamber Waterflow Testing

The Agena 2000 injectors and thrust chamber were
cold flowed with water to verify then- hydraulic
characteristics. Injector water flows verify pressure
drops, reactant stream quality and impingement, and
validate the injector hydraulic models. Figure 10
shows good agreement between the injector AP's and
the model. In addition, the drilled orifices show
excellent stream quality and impingement
characteristics.

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

flanged version of the flight injector and a carbon
steel heat sink combustion chamber. The
uncooled chamber incorporated a bomb port and
two high frequency pressure transducer ports for
the combustion stability testing. Water-cooled
calorimeters ports were located down the length
of the chamber at the zero and 45 degree
positions around the diameter of the chamber to
obtain heat flux data. A heat exchanger was used
to pre-heat the oxidizer to temperature levels
predicted for the regeneratively cooled chamber.
Heat sink testing was conducted in June 1997 and
all test objectives were achieved. The PC/MR
survey showed the performance of the injector to
be within 1% of the predicted value of 5684
ft/sec. The bomb tests were conducted in
accordance with CPIA 655 using a RDX grain
size of 6.5. The bomb tests showed stable
operation of the injector within the operating box
and identified a region outside of the operating
box where high frequency instability was evident.
The low chamber pressure tests established the
chug stability limit for the injector at 250 to 300
psia. Preliminary heat flux data was also
obtained.

SIDE

lil

/

Ofifi

s*
£
OL

i/

»
a.

^/^

MODEL

uu

g &'

^

P gso
0

•»

+ DATA

Q

n.
0

^~^^
5

10

20

15

WATER FLOWRATE Ibm/s

PROTOTYPE CHAMBER WATER FLOW
NOZZLE SECTION: MODEL/DATA COMPARISON

Figure 10. Injector Waterflow Results

The chamber coolant passages are water flowed to
verify channel-to-channel flowrate variations are
small and to validate the flow model predictions
Figure 11 shows good agreement between test data
and the hydraulic model, and a small flowrate
variation (channel-to-channel). Chamber hydraulic
data is critical to verifying thermal behavior.
Eventually, water flow tests could be used to
reduce or eliminate hot-fire acceptance tests.

240-

DATA

220
210200
190180-

,

x

160

/

^

/

140ISO120
110-

^

/

~^
^
^
"

100
90
18

The heat sink test hardware, shown in Figure 12,
was a take-apart configuration consisting of a

s

s*

19

20

21

Injector/Heat Sink Thrust Chamber Test Series

The test objectives for the heat sink TCA test series
were as follows:

Conduct a preliminary chamber
pressure/mixture ratio survey

Investigate the high frequency stability
characteristics of the injector

Establish the chug stability margin of the
injector for start transients

Obtain silicon oil thermal/heat flux data

MODEL

s"

22

23

24

25

26

27

28

29

MASS FLOWRATE Ibm/s

PROTOTYPE CHAMBER SN 001, FLOW#2

Flow per Hole, Wtot=19.1 #/s,
S * One !icjma

M

»...s...a
*s -OneSlg iia

.......~^..

Wavg=0.309#/s<>c

v

6ne;S'igma=d.6cio#/sec

0

10

20

30

40

50

Figure 11. Chamber Waterflow Results

60

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

Bomb Port

Steel Workhorse
/ Chamber

exhaust duct and incorporated an exit orifice to
simulate the NCM inlet pressure.

Gas generator testing was conducted in
August/September 1997 and all test objectives were
achieved. The performance of the GG was
characterized over a wide range of mixture ratios and
chamber pressures. Nominal performance of the GG
(C*) was approximately 5% below the predicted
value. The C* step (transition from bipropellant to
monopropellant behavior) was found at a mixture
ratio of 0.10. The low chamber pressure test
established the chug stability limit for the injector at
200 to 225 psia.
Turbine Exhaust
Manifold

Tak@ Aoart Injector

Figure 12 Heavyweight TCA Test Configuration
Early in the test series, there was visual evidence of
thermal damage to the webs between the resonators
of the acoustic cavity ring. This limited most tests to
1 or 2 seconds. It was determined that the cause of
this thermal phenomenon was insufficient cooling of
the webs by the oxidizer flow due to a repair
incorporated into the take-apart injector after a proof
test failure. A minor redesign was incorporated into
the second injector which was fabricated for
subsequent test series to eliminate this problem.
Gas Generator Test Series
The test objectives for the gas generator test series
were as follows:

Characterize the GG performance

Verify the chug stability margin of the GG
for start transients
The GG test hardware, shown in Figure 13, consisted
of a flanged version of the flightweight GG, a turbine
inlet manifold simulator, and turbine exhaust
backpressure simulator. The GG was bolted to the
turbine manifold simulator for testing. After this test
series was completed, the GG was bolted to the
turbine inlet manifold of the prototype TPA. The
turbine manifold simulator consisted of an original
Agena turbine inlet housing modified for GG testing.
Three of six turbine inlet nozzles were plugged to
simulate the Agena 2000 turbine inlet housing nozzle
CdA. The turbine exhaust backpressure simulator
was fabricated from an original Agena turbine

Turbine
Manifold
SimuSatoir

Gas
Generator

Helium
Purge

Port

Figure 13. Gas Generator Test Configuration
Pressure-Fed TCA Test Series
The test objectives for the pressure-fed TCA test
series were as follows:

Conduct a preliminary chamber
pressure/mixture ratio survey

Obtain oxidizer temperature data from the
regeneratively cooled chamber for model
verification

Establish the chug stability margin of the
TCA for spin start transients
The pressure-fed test hardware, shown hi Figure 14,
consisted of a regeneratively cooled thrust chamber

Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

welded onto the modified injector. A flightweight
TTA was used to attach the TCA to the test stand.
Stiff links attached the TTA to the gimbal lugs on the
TCA to hold the TCA in a horizontal position for
testing.
Pressure-fed TCA testing was conducted in October
1997 and all test objectives were achieved. The
PC/MR survey showed the performance of the TCA
to be within 1.0% of the predicted value. The test

hardware showed no indication of thermal damage as
the heat sink injector had exhibited. The baseline
test at a chamber pressure of 718 psia did indicate a
combustion instability phenomena. It was determined
that the combustion instability at 718 psia could be
corrected by modifying the injector elements
(increasing orifice diameters and reducing stream
velocities). To complete the Agena 2000 test series, it
was decided to continue testing at a lower chamber
pressure (650 psia). The low chamber pressure test at
290 psia showed no indication of chug instability.

Figure 15. Pressure-Fed TCA Test Oxidizer
Temperature

GG/TPA Test Series

The test objectives for the GG/TPA test series were
as follows:

The regeneratively cooled chamber thermal



performance exceeded the design expectations. The
maximum oxidizer temperature reached 160°F, as
seen in Figure 15, and decreased to about 135°F as
the test continued, well below the predicted oxidizer
temperature of 180°F. This indicates the silicon
dioxide layer is reducing the heat flux more than
predicted. This creates the potential for improving
TCA performance by reducing the barrier flow
percentage of total flow and/or increasing the barrier
mixture ratio toward higher performance.








Verify turbopump stability and vibration
frequency distribution
Determine helium requirements for start
transient
Determine critical speed margin
Map pump performance at the design point
for 718 psia TCA chamber pressure
Map pump performance at the operating
point for 650 psia TCA chamber pressure
Demonstrate bootstrap operation
Determine the valving sequence at the
operating point for prototype engine testing

The GG/TPA test hardware, shown in Figure 16,
consisted of the prototype turbopump, the gas
generator tested previously, and a turbine exhaust
manifold with a backpressure orifice. The GG/TPA
was mounted on the TCA for ease in converting the

Regeneratively
Cooled
Chamber

Flight Thrust
Take Out
Assembly

Figure 14, Pressure-Fed TCA Test Configuration
9

Figure 16. GG/TPA Test Configuration


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