(AMBR) Engine

46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
25 - 28 July 2010, Nashville, TN
AIAA 2010-6883
Performance Results for the Advanced Materials
Bipropellant Rocket (AMBR) Engine
Scott Henderson1
Carl Stechman , Kim Wierenga3 and Scott Miller4
Aerojet, Redmond, Washington, 98073
2
Larry Liou5
NASA Glenn Research Center, Cleveland, Ohio, 44135
Leslie Alexander6
NASA Marshall Space Flight Center, Huntsville, Alabama, 35816
John Dankanich7
Gray Research, Cleveland, Ohio, 44135
Aerojet Approval Log Number 2010-020
Performance Characterization of the Advanced Materials Bipropellant Rocket (AMBR)
engine was completed at Aerojet’s Redmond, Washington test facility in the summer of 2009.
This project was funded by the NASA In-Space Propulsion Technology (ISPT) Project
Office. The primary goal of this project was to maximize the specific impulse of a pressure
fed, apogee class earth storable bi-propellant engine using nitrogen tetroxide (MON-3)
oxidizer and hydrazine fuel. The secondary goal of the project was to take greater
advantage of the high temperature capabilities of iridium/rhenium material used for the
combustion chamber and nozzle. The first round of hot fire testing of the AMBR engine
occurred in October of 2008; during which, the engine demonstrated a maximum specific
impulse of 333.5 seconds at a mixture ratio of 1.1 and a thrust of 151-lbf (672 N). This
operating thrust level at this mixture ratio was close to the thermal stability (loss of fuel film
cooling) thrust limit of the engine. A second round of testing, including random vibration,
shock and extended hot fire testing, was added to the program with the goal of bringing the
AMBR engine to a Technology Readiness Level (TRL) of 6. Following successful random
vibration and shock testing, the AMBR engine went through a second hot fire test series in
the summer of 2009 to document the performance (thrust and mixture ratio) operating
margins and to demonstrate a long duration burn. During this testing, the engine
demonstrated a specific impulse of 333 seconds at a mixture ratio of 1.1 and a thrust of 141lbf (627 N) (i.e. the AMBR engine demonstrated high performance with margin). This
performance also occurred during a successful 2,700 second long burn. This program also
successfully demonstrated the secondary goal of hot fire operation at a 4000°F (2200°C)
combustion chamber temperature.
1
Project Engineering Specialist
Technical Principal, AIAA Senior Member
3
Manager of Programs
4
Director of Program, AIAA Associate Fellow
5
Project Area Manager, NASA ISPT Project Office, AIAA Senior Member
6
Technology Area Manager, Chemical and Thermal Propulsion Project, TD05, AIAA Member
7
Lead Systems Engineer, NASA ISPT Project, AIAA Senior Member
1
American Institute of Aeronautics and Astronautics
2
Copyright © 2010 by Aerojet. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
Nomenclature
AMBR
AR
C*
CVD
El-Form
FFC
GEO
hg
HPLAE
HiPATTM
Ir/Re
Isp
kg
lbf
lbm
LAE
MMH
MON-3
MR, O/F
N
N2H4
NTO
Pc
psia
sec
= Advanced Materials Bipropellant Rocket
= Area Ratio
= Characteristic Velocity of rocket combustion products
= Chemical Vapor Deposition
= Electroform process for fabrication of iridium lined rhenium material systems
= Fuel Film Cooling
= Geosynchronous Earth Orbit
= Convection coefficient
= High Performance Liquid Apogee Engine
= High Performance Apogee Thruster
= Iridium lined Rhenium material system
= Specific Impulse- lbf-sec/lbm
= Kilogram
= pounds force
= pounds mass
= Liquid Apogee Engine
= Monomethylhydrazine, N2H3CH3
= Mixed Oxides of Nitrogen, nitrogen tetroxide and 3% NO by mass in solution
= Mixture ratio
= Newton
= Hydrazine, N2H4
= Nitrogen Tetroxide, N2O4
= Chamber pressure
= Pounds per square inch absolute
= Seconds
Introduction
NASA sponsored the Advanced Materials Bipropellant Rocket (AMBR) program1 with the primary goal of
increasing the specific impulse of small (100 lbf - 200 lbf (445-890 N)) thrust class pressure-fed earth storable
bipropellant rocket engines to 330 seconds with nitrogen tetroxide (MON-3) and monomethylhydrazine propellants
and 335 seconds with MON-3 and hydrazine propellants. At the start of this program the state of the art storable
rocket engines in this thrust class delivered 323 Isp seconds for MMH and 328
seconds Isp for hydrazine fuel.
The AMBR program’s performance goals were primarily achieved by
tailoring the operational conditions that maximize the benefits of the flightproven iridium/rhenium (Ir/Re) high temperature capability combustion chamber
technology. This material system has the capability to withstand steady-state
wall temperatures exceeding 4000°F (2200°C) compared to the common usage
Ir/Re combustion chamber on the HiPATtm engine2 at less than 3200°F
(1760°C). In order to increase engine performance, changes to the operational
conditions included increases in chamber pressure and the mixture ratio
compared to the current state of the art, both resulting in higher chamber
temperature. In addition, the engine design includes modifications for increased
injector performance and chamber/nozzle efficiency over the current state of the
art.
Development testing of a high performance bipropellant rocket engine using
nitrogen tetroxide (MON-3) and hydrazine (N2H4) propellants was conducted at
Aerojet’s Redmond test facility. Multiple unlike impinging oxidizer and fuel
doublets were used for the core injector elements and a dedicated set of fuel
injection orifices along the chamber wall was used for cooling the combustion
chamber to injector interface. The enhanced performance was aided by the
incorporation of an Aerojet patented “step” assembly3.
2
American Institute of Aeronautics and Astronautics
Figure 1. AMBR prototype
rocket engine- after test
The program was divided into three phases: Base period4, Option 15 and Option 26,7,8,9. This report will describe
the approach and final results from the Option 2 period which was the characterization of the performance and
operational characteristics of the prototype engine shown in Fig. 1. This engine was flight-like in all aspects.
Engine Evolution and Description
This section describes the evolution and basic design of the NASA NRA
AMBR engine. The AMBR design is based on the Aerojet model R-4D15(DM) HiPATtm (Fig 2) with the goals being higher specific impulse, longer
life, and lower cost manufacturing methods. The design was created using a
team of engineers with expertise in design, material and processes,
manufacturing, engine performance, thermal and structural analysis. Items
considered include thermal capability, strength at temperature, oxidization,
galvanic compatibility, part tolerances, method of fabrication, cost, lessons
learned, and what is needed to provide the proper operational criteria.
The AMBR engine is comprised of individual subassemblies which are,
except for the valve/injector interface, all welded.
Two co-axial solenoid valves
Injector assembly with mount flange
Aerojet patented “Step” assembly3
Ir/Re combustion chamber and exit nozzle
C-103 niobium alloy exit nozzle extension
Titanium exit nozzle extension
Figure 2 Aerojet model R-4DThe two solenoid valves are production configurations with no modifications
15DM HiPATtm rocket engine
tm
that are used on the HiPAT and C-103 niobium R-4D-11 100lbf (445 N)
engines. These valves were originally used on the R-4D configurations for the Apollo RCS on the service module
and lunar modules reaction control system. The valves are attached to the injector assembly mount flange using a
collar that fits over the top of the valve and 4 bolts that are screwed into the injector assembly mount flange. A leak
tight seal between the valves and the injector assembly mount flange is accomplished using redundant soft seals.
The injector assembly incorporates the manifolds transferring the propellant between the valves and the final
injector orifices and the final injector orifices that include the fuel film cooling. The injector was modified based on
the results of copper chamber tests accomplished in the Option 1 phase. Multiple unlike impinging oxidizer and fuel
orifices were used for the core injector elements and a dedicated set of fuel injection orifices along the chamber wall
was used for fuel film cooling the step assembly and a portion of the combustion chamber. Thermal standoffs”
integrated into the assembly are used to provide thermal isolation of the valves from the injector during soak back
periods after engine firings.
The Aerojet patented step assembly is designed specifically for
the operating conditions (mass flow and mixture ratio) and
propellants (MON-3/hydrazine) and integrated into the Ir/Re
combustion chamber prior to attachment to the injector. The precombustion chamber has three functions. First, it acts as a
mechanical joint transition between the main combustion chamber
and the injector assembly. Second, it provides a means for
protecting the iridium layer on the internal surface of the rhenium
combustion chamber from reacting with the partially unburned
propellants and combustion products in the vicinity of the injector
assembly. Third, the sudden area expansion at the end of the step
assembly enhances the mixing of the partially reacted main core of
combustion gases with the fuel film used for cooling.
Figure 3. El-Formtm Combustion chamber
as delivered from PPI
3
American Institute of Aeronautics and Astronautics
The iridium lined rhenium combustion chamber and initial portion of the exit nozzle (Fig. 3) was fabricated
using the El-Formtm process. This technology involves electroplating using an anode/cathode submerged in an
electrolyte solution. The anode material, such as rhenium, is deposited onto a mandrel of the desired shape. After
deposition is completed, the cathode and mandrel are removed. The applied material is of net or nearly net shape
and material properties can be tailored by varying the composition of the applied material layers that prevent crystal
growth during chamber heating. The process has a potential capability of reducing fabrication costs compared to
other existing processes for the Ir/Re material system and permits multi-component processing.
Fabrication of the chamber assembly was accomplished by Plasma Processes, Inc. (PPI) of Huntsville,
Alabama. Aerojet defined the design details of the final chamber configuration and worked with PPI to ensure the
processes were in place to fabricate the chamber. Fabrication of the chamber was completed and it was delivered to
Aerojet in September 2008.
The exit nozzle extension attached to the exit of the nozzle portion of the iridium/rhenium combustion chamber
consists of an R-512E silicide coated C-103 niobium alloy “conical” section and a titanium nozzle extension. The
design of these assemblies is identical, except for slight dimensional variations, to the HiPATtm design.
Option 1 Phase A Testing-Prototype Engine8
The development engine is shown in Fig. 4 prior to testing . Initial hot fire testing of the testing of the engine
demonstrated that the nominal baseline design thrust level of 200 lbf (890 N) at 400 (27.5 bar) psi inlet did not
have adequate thermal margin with respect to fuel film cooling. Simultaneously
the system studies showed that the baseline inlet pressure of 400 psi (27.5 bar)
was not practical in the near term from a propellant tank design pressure aspect.
The test program documented that the engine when operated at an inlet
pressure of 300 psia (20.7 bar) or less with a corresponding thrust of 140-150 lbf
(620-670 N) would exhibit thermally stable operation and a specific impulse of
>333 seconds. Fig. 5 shows the demonstrated specific impulse of the engine as
function of thrust over the 1.05 to 1.15 design mixture ratio while Fig. 6 shows
the measured combustion chamber temperature over the same range.
The engine was subsequently operated at higher mixture ratios to
demonstrate the operating temperature capability of the material system since the
lower mixture ratio testing at 1.05-1.15 operated at cooler temperatures. At a
mixture ratio of approximately 1.4 the temperature of the engine as measured by
the pyrometer on the outside surface was approximately 4000°F(2200°C) with
the inside temperature estimated to be, based on the thermal model, in excess of
4200°F (2300°C). Tests were also accomplished at lower inlet pressures and
wider mixture ratios to define the low frequency combustion stability margin
(chugging).
Table I summarizes the primary accomplishments of this test phase. Fig. 7
Figure 4. AMBR engine
shows the range of thrust levels and mixture ratios tested along with the
prior to test program
demonstrated specific impulse. Specific impulse levels approaching 335
seconds were encountered at the higher thrust levels (200 lbf) but the lack of adequate fuel film cooling prevented
attainment of steady state operation. Fig. 8 summarizes and documents the area of thermally stable operation.
There was no indication of any abnormal combustion aberration such as a 1st tangential instability or any other
combustion abnormality. In general the maximum thrust capability of the engine with respect to steady state
operation is in the range of 140-150 lbf (620 to 670 N).
4
American Institute of Aeronautics and Astronautics
AMBR Performance
340
Specific Impulse - Sec
Mixture Ratio Range = 1.05 to 1.15
330
320
310
Isp = 7.164E-06F3 - 3.626E-03F2 + 6.187E-01F + 2.974E+02
300
50
100
150
200
250
Engine Thrust - Lbf
Figure 5. AMBR specific impulse versus thrust
Chamber Temperature versus Thrust @ MR=1.1
Chamber Temperature - °F
3900
TCH2
3800
3700
3600
3500
3400
3300
3200
3100
3000
0
50
100
150
200
250
Thrust - Lbf
Figure 6. AMBR Chamber temperature versus thrust
Table I Phase A test summary
Operating parameter
Total firings
Total firing time
Maximum sustained external chamber temperature
Maximum chamber pressure
Minimum chamber pressure (no chug)
Specific impulse
Thrust range
Mixture ratio range
Engine mass
Demonstrated value
48
4397 seconds
3925°F (4025°F Transient)
289 psia (~20 bar)
99 psia (6.8 bar)
333.5 seconds (1.1 MR and 150 lbf); 400:1 Ae/At
73-214 lbf (320 N to 935 N)
0.9 to 1.4
10.8 lbm (4.9 Kg)
5
American Institute of Aeronautics and Astronautics
Figure 7 - AMBR thrust, mixture ratio and specific impulse as a function of inlet pressure and test
points
Figure 8. AMBR operational limits
6
American Institute of Aeronautics and Astronautics
Option 2 Phase A Testing-Prototype Engine9
The Option 2 Period was primarily composed of tasks which would bring the AMBR engine to a technology
readiness level (TRL) of 6. These included:
Random vibration testing
Shock testing
Additional performance characterization and long duration hot fire testing
Aerojet successfully conducted random vibration testing of the AMBR engine in December of 2008. The test used
the HiPATTM qualification level vibration spectrum which was individually applied to all three axes. The resulting
test data showed close agreement with the pre-test stress predictions. Fig. 9 shows the random vibration level that
was incorporated in the test and Fig. 11 shows the engine installed in the vibration fixture.
Figure 9 AMBR random vibration level
Figure 10 AMBR engine installed in vibration fixture
The AMBR engine did not have a
customer defined engine shock
specification so the engine structure
was designed around the existing
HiPATTM 100 lbf (445 N) flight engine
design structure. This lack of a specific
requirement also required an analysis to
be performed on the engine to
determine the shock levels the design is
capable of handling. This analysis was
used to define a shock spectrum that
would then be applied during the test.
This shock acceleration spectrum level
shown in Fig. 11 was to be applied
individually to each of the three global
axes.
The actual shock test was
performed by the Jet Propulsion
Laboratory in Pasadena, CA.
Figure 11. AMBR Shock level spectrum
7
American Institute of Aeronautics and Astronautics
Fig. 12 shows the AMBR engine installed on the JPL shock test
beam. After the test was performed the engine was examined for
any deformation or other damage. Dimensional measurements did
not reveal any permanent deformation of chamber/nozzle.
The end goal of this program is to reach a TRL-6 level of
development for the AMBR engine. This includes increasing the
total burn time on the engine as much as possible given the
propellant budget. During this additional firing of the engine, the
performance map established in Option 1 was expanded to cover
more off-design conditions. The testing was concluded by
demonstrating a long duration burn, of relevant length to an actual
application, to demonstrate long term engine thermal stability.
Following functional testing, the engine was delivered to the test lab
for instrumentation installation and finally installation in the test
cell. The engine was installed in the test cell in the same manner as
for Option 1, and as shown in Fig. 13.
The initial hot fire test series included operation at low mixture
ratios and low thrust levels. Before the long duration test a series of
tests were performed over a range of mixture ratios (0.6 to 1.23)
and thrust levels (86 lbf (382 N) to 173 lbf (770N)).
Figure 12. AMBR engine installed in
JPL test fixture
The long duration test demonstrated that the engine
was capable of meeting mission durability and that
the engine exhibited thermal characteristics that were
similar to the existing HiPATTM which included low
soakout temperatures after engine firing termination.
Fig. 14 and 15 show the engine temperatures during
and after the 2700 second firing duration.
The injector/mount plate temperatures during the
steady state operating mode of 200° to 280°F (98°C to
138°C) and the valve temperature after engine firing
termination and soak back of 180°F (82°C) are well
below the component material and operation limits.
The performance of the AMBR engine was
characterized as a function of the thrust level at a
constant thrust (see Fig. 5) and as a function of
Figure 13. AMBR engine installed in Aerojet test cell
mixture ratio over the thrust range of 130 to 214 lbf
since the specific impulse was relatively constant at
any mixture ratio for the data that was acquired. Fig. 16 shows the specific impulse as a function of the mixture
ratio.
The Option 2 Period of the AMBR program successfully brought the engine to TRL 6 with
the completion of the random vibration test, shock test and long duration hot fire test. The
random vibration test successfully executed a qualification level test using the HiPATTM vibration
spectrum with no deformation, as predicted, to the engine. The shock test applied a derived
spectrum based on the capabilities of the AMBR design.. Finally, the long duration test demonstrated the high
performance of the engine at a reduced operating condition (i.e. at an operating condition that has margin). The
cumulative hot fire testing on the AMBR engine achieved the following milestones:
• 89 engine starts
• 9,138 seconds of firing time
• 3,925°F (2160°C) maximum sustained chamber temperature
8
American Institute of Aeronautics and Astronautics
• 289 psia (20 bar) maximum chamber pressure
• 99 psia(6.8 bar) minimum chamber pressure
• 333.5 seconds maximum specific impulse
151lbf (672 N) at a 1.1 mixture ratio
• 2,700 seconds maximum single burn duration at 141lbf (627 N) and 1.1 mixture ratio
Figure 14. AMBR mount flange (tf1, 2 and 3) and valve (Tvo and Tvf) temperatures during and after
2700 second firing test
Figure 15. AMBR chamber( Tch) , Titanium/C-103 nozzle weld joint (Tn1) and titanium nozzle exit
(Te1) temperatures during and after 2700 second firing test- Note: Initial variation in Tch due to
pyrometer alignment adjustment
9
American Institute of Aeronautics and Astronautics
Figure 16 AMBR engine specific impulse as a function of the mixture ratio
Summary
Summary and Conclusions
The objective of the NASA AMBR program performed by Aerojet, Redmond, WA was to maximize the specific
impulse of a pressure fed, apogee class earth storable bi-propellant engine using nitrogen tetroxide (MON-3)
oxidizer and hydrazine fuel. A specific impulse of 333.5 seconds was demonstrated. The secondary goal of the
project was to take greater advantage of the high temperature capabilities of iridium/rhenium material used for the
combustion chamber and nozzle though operation at higher temperatures. The durability of the engine design using
the iridium lined rhenium El-FormTM combustion chamber and nozzle was demonstrated by operation of the engine
over a large range of propellant mixture ratios and thrust levels for more than 9000 seconds. In addition the engine
was subjected to typical qualification random vibration and shock levels with no damage. The thermal limits of
operation were characterized. Both low frequency (chugging) onsets and high frequency (1st tangential) stability
was demonstrated.
References
1
Scott Miller, Scott Henderson, et al., “Performance Optimization of Storable Bipropellant Engines to Fully
Exploit Advanced Material Technologies,” 2006 NASA Science and Technology Conference, July 2007.
2
Wu, P.K., Woll, P., Stechman. C., McLemore, B., Neiderman, J. and Crone, C. “Qualification Test of a 2nd
Generation High Performance Apogee Thruster” AIAA 2001-3253. 37th AIAA Joint Propulsion Conference, July
2001
3
Stechman, R.C., Woll, P.E., Neiderman, J.M., and Jensen, J.J., Kaiser Marquardt, Van Nuys, CA, U.S. Patent for a
"High Performance Rocket Engine Having a Stepped Expansion Combustion Chamber and Method of Making
the Same," U.S. Patent Number 6,397,580, June 2002
4
Portz, R., Henderson, S. Krismer, D., Lu, F., Wilson K., Miller, S., ”Advanced Chemical Propulsion System
Study,”AIAA-2006-032, 43nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Cincinnati,
Ohio, July 2007.
5
Henderson, S., Wilson, K., et al., ”Performance Increase Verification for a Bipropellant Rocket Engine”, AIAA2008-4844, 44th Joint Propulsion Conference and Exhibit, Hartford, CT, July 2008.
6
Liou, L., “Advanced Chemical Propulsion for Science Missions”, IEEE Aerospace Conference, Big Sky, MT,
March 2008.
10
American Institute of Aeronautics and Astronautics
Larry, et al., “Recent Development in NASA In-Space Chemical Propulsion”, 55th JANNAF Propulsion
Conference
8
Henderson, Scott, “Period Final Report Advanced Material Bi-propellant Rocket (AMBR), Option 1” Aerojet
report 2009-R-3322, dated October 22, 2009.
9
Henderson, Scott, “Period Final Report Advanced Material Bi-propellant Rocket (AMBR), Option 2” Aerojet
report 2009-R-3347, dated November 17, 2009.
7Liou.
11
American Institute of Aeronautics and Astronautics