AE8900-Slater-Steffan-Report

Cycle Analysis and Trade Studies of the RS-25
Space Shuttle Main Engine Powerhead
AE-8900-MAV Special Problems
Presented by: Steﬀan Slater
Advisor: Prof. Dimitri Mavris
April 10, 2015
Grade received:
Advisor’s signature:
The Guggenheim School of Aerospace Engineering
Georgia Institute of Technology
Honor Code Statement
I certify that I have abided by the honor code of the Georgia Institute of Technology and followed the
collaboration guidelines as speciﬁed in the project description for this assignment.
Signed:

AE 8900-MAV Spring 2015 Steffan Slater
Cycle Analysis and Trade Studies of the RS-25 Space
Shuttle Main Engine Powerhead
Executive Summary
As new advanced technologies such as additive manufacturing have become more mature, interest in in-
corporating these technologies into new aerospace components and systems has steadily increased. NASA,
Aerojet-Rocketdyne, and SpaceX are all pursuing the integration of additive manufacturing into new rocket
engine programs. Simultaneously, manufacturing considerations are being integrated into the design pro-
cess earlier than the detailed design stage as was historically done. Both of these trends necessitate both
physics-based subsystem modeling as well as a unified sizing and synthesis environment integrating several
subsystem analyses. For a spacecraft or launch vehicle, rocket engine cycle analysis is a critical element of
this environment.
The Space Shuttle Main Engine (SSME) is a large (500,000 pound thrust class) staged combustion engine
using liquid oxygen and liquid hydrogen propellants. The SSME was developed for use on the Space Shuttle
and was designed to be reusable for up to 100 flights. With the Space Shuttle program concluded, 16 engines
currently remain in inventory. These engines are slated for use on the upcoming Space Launch System (SLS),
a new heavy-lift vehicle under development by NASA Marshall Space Flight Center. Each expendable SLS
core will use 4 Space Shuttle Main Engines in concert with two boosters to lift large payloads to space.
The remaining SSME inventory is only sufficient for the first four SLS flights, so in order to support later
flights, Aerojet-Rocketdyne intends to restart production of the SSME as a new expendable variant. There is
significant interest in taking advantage of new manufacturing techniques to improve the affordability of this
expendable engine, as well as in making other design or operational changes to the SSME to take advantage
of the expendable nature of the SLS, trading reusability for performance, mass, or cost. In particular, NASA
is particularly interested in affordability for this variant, with mass penalties considered an acceptable price
to pay for better affordability. This project will produce a cycle model of the SSME and utilize it to perform
trade studies to identify potential design and operational changes for use on the expendable variant of the
SSME, with emphasis on the engine powerhead (pumps, turbines, preburners, pipes, and valves).
There are a variety of options available for tools to perform this analysis. Low fidelity tools such as
spreadsheets require the least information but can only perform simple analyses and often rely on many
simplifying assumptions. High fidelity tools model most or all of the components in the engine and perform
a true thermodynamic cycle analysis, but they require more information and tend to take more computational
time. These tools are often implemented as computer programs or languages such as Numerical Propulsion
System Simulation, which is heavily used in gas turbine cycle analysis. Extremely high fidelity tools such
as computational fluid dynamics and heat transfer models provide the highest accuracy and model full 3-D
flow fields, but require the most information, such as full geometry and boundary conditions. These tools
also have long execution times for non-trivial models. For this application, high-fidelity tools are most
appropriate, as the impacts of technological advancement and manufacturing processes are only reliably
visible when individual components are modeled and rapid execution time is a necessity for design space
exploration. One of the major concerns with such tools is the availability of input data. The SSME is one of
the most studied engines ever made, with a large quantity of published data available in the public domain,
so that is not a concern for this application. For the specific high-fidelity tool, the ROCket Engine Transient
Simulation (ROCETS) tool from NASA Marshall Space Flight Center was selected and obtained through a
software agreement, as it has been applied to rocket engines before and has existing models and literature
available from NASA.
The ROCETS SSME model was benchmarked against published performance data. Multiple data points
were available for both the Full and Nominal Power Levels, so both settings were benchmarked against
the available performance parameters. The benchmarking indicated a match on high-level performance
parameters within 2.5%. Detailed performance data at the component level was also available for the
Nominal Power Level and was compared against model predictions, showing more spread than high-level
i

parameters like thrust, but still generally good agreement.
Trade studies on the SSME were performed at the operating conditions experienced on the SLS, not on
the Space Shuttle. This entails operating at sea level rather than vacuum as well as higher inlet pressures due
to the greater height of the vehicle. Three trade studies were performed. The effect of increased chamber
pressure was examined, showing that higher pressures increase both thrust and specific impulse. A 10%
increase in chamber pressure provided a 12% increase in thrust and a 2% increase in specific impulse. In
order to recover the thrust lost by moving to sea level from vacuum, a chamber pressure boost of over 17% is
required. The second trade study examined the effect of oxidizer-to-fuel ratio, showing that a decreased ratio
improved specific impulse and decreased chamber temperatures at the cost of some thrust. The third trade
study examined a design change, the reduction of the pump diameter. This would save weight by shrinking
the pump and housing, but also decreases the pressure rise across the pump at a given rotational speed.
This trade study showed that, as expected, shrinking the pump increased the rotational speed necessary to
maintain the same chamber pressure.
While not enough is publicly known about the SSME or SLS to make serious design recommendations,
some potential options can be seen. If an increase in engine performance is desired, increasing the chamber
pressure will improve both thrust and specific impulse. As long as the pressure is not so high as to rupture the
chamber during normal operation, this can be seen as trading reusability (the need to be able to withstand
numerous such pressure loading cycles) for performance. Lowering the oxidizer-to-fuel ratio also provides
a modest gain in specific impulse, assuming the engine can operate at a lower value. Oxidizer-to-fuel ratio
is also closely tied with the design of the vehicle, as a change in this ratio impacts the amount of each
propellant that must be carried, affecting tank sizes and masses. Finally, trimming the pump diameter may
provide a way to reduce mass and cost without sacrificing performance. The higher rotational speeds may be
acceptable on an expendable engine, depending on the stresses and rotordynamic margin currently present
in the pump system. While no one solution is going to address every concern, especially without the detailed
engineering data behind the engine, the demonstrated ability to make and analyze these trades builds a
framework for integration into a system with the capability for thorough examination of the effects of these
changes on the entire vehicle as well as providing guidance for more detailed engineering analysis.
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AE 8900-MAV Spring 2015

Cycle Analysis and Trade Studies of the RS-25 Space
Shuttle Main Engine Powerhead
Steffan Slater∗
steffan.slater@gatech.edu
ROCETS, rocket engine, cycle analysis, SSME, RS-25, SLS
New technological advancements and an increased emphasis on affordability12
are driv-
ing the integration of technologies like additive manufacturing into new rocket engine pro-
grams3
.2
Since these technologies are new, lacking historical data, and have impacts across
multiple subsystems, a physics-based parametric analysis approach is required.6
A physics-
based, high-fidelity modeling system for liquid rocket engines, ROCETS, is discussed and
a model of the Space Shuttle Main Engine (SSME) developed. The model is benchmarked
against published SSME data,9
with high-level performance parameters matching pub-
lished data within 2.5%. The model is applied to the SSME as used on the Space Launch
System,13
with corresponding changes in operating conditions, and trade studies on opera-
tional and design changes are performed. Both performance-increasing and mass-reducing
changes are identified, and potential application of each to the Space Launch System are
presented.
Nomenclature
ψ Head Coefficient
ρ Density
g Acceleration Due to Gravity
h Vertical Height
N Rotational Speed
p Pressure
r Radius
FPRB Fuel Preburner
HPFP High-Pressure Fuel Pump
HPFT High-Pressure Fuel Turbine
HPOP High-Pressure Oxygen Pump
HPOT High-Pressure Oxygen Turbine
LPFP Low-Pressure Fuel Pump
LPFT Low-Pressure Fuel Turbine
LPOP Low-Pressure Oxygen Pump
LPOT Low-Pressure Oxygen Turbine
MSFC Marshall Space Flight Center
NPSS Numerical Propulsion System Simulation
O/F Oxidizer-to-Fuel Ratio
OPRB Oxygen Preburner
ROCETS Rocket Engine Transient Simulation
SLS Space Launch System
SSME Space Shuttle Main Engine
∗Graduate Research Assistant, Aerospace Systems Design Laboratory
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I. Background and Motivation
In recent years, disruptive technologies such as additive manufacturing have increased greatly in technolog-
ical maturity. These technologies promise improvements in performance, cost, and speed of manufacturing
over conventional methods, and companies are eager to take advantage of them. Within the aerospace indus-
try, additive manufacturing in particular has received a great deal of attention from both NASA and industry.
NASA’s Marshall Space Flight Center (MSFC) has tested additively manufactured parts for use on the J-2X
engine,1
and Space Exploration Technologies (SpaceX) has qualified additively manufactured valves for use
on its Merlin 1D engine.2
While these programs have been primarily experimental and research programs,
extensive work is going into bringing additively manufactured parts to the production line. SpaceX appears
to be taking the lead on this front, with its upcoming SuperDraco engines having an additively manufactured
combustion chamber,2
but other companies are not far behind.3
In addition to integrating new technologies, aerospace manufacturers are beginning to change the way
manufacturing and design interact. Traditionally, there are three phases of design, conceptual, preliminary,
and detailed, and manufacturing considerations are not considered until the detailed design stage.4
This can
lead to problems and expensive design changes when manufacturing considerations are finally incorporated.
By considering manufacturing earlier in the design process, at the conceptual or preliminary stage, potential
issues can be addressed early on and at lower cost. This idea of manufacturing-influenced design can decrease
both the cost and duration of the design process, and is an active research area4
.5
The conventional design process relies heavily on historical data and regressions to inform analysis and
decision-making. For advanced designs incorporating new technologies or manufacturing considerations, this
paradigm is insufficient. The historical data simply does not exist for the designer to use. In order to
analyze these kinds of designs, physics-based tools which allow for the integration of these new technologies
are necessary.6
Since these physics-based tools are generally created for discipline-specific analyses, an
environment which can synthesize these tools and size the system is also required. These disciplinary tools
also exist at varying levels of fidelity, and determining the appropriate fidelity level depends on the specific
problem, so a sizing and synthesis environment will often include similar tools at different fidelity levels for
the different stages of design.
Figure 1. Close-up of a Space Shuttle Main
Engine Test8
This sizing and synthesis approach has been successfully ap-
plied in the past to aircraft problems6
.4
These environments
have been used to analyze the impacts of technology infusion
and to enable manufacturing-influenced design for air vehicles.
This approach has been applied far less to the design of space
vehicles. Application to space vehicles will allow the design
and analysis of advanced concepts for challenging future space
missions, such as a manned mission to Mars.7
The Aerospace
Systems Design Laboratory at Georgia Tech is currently in-
vestigating this area in collaboration with MSFC. Just as gas
turbine engine performance analysis is a critical element of in-
tegrated aircraft sizing and synthesis environments,6
rocket en-
gine performance analysis will be a critical part of a spacecraft
sizing and synthesis environment.
The RS-25 rocket engine, better known as the Space Shut-
tle Main Engine (SSME) is a large staged combustion engine
developed for the Space Shuttle program in the 1970s. Using
liquid oxygen and hydrogen as propellants, the SSME is rated
for 470,800 pounds of vacuum thrust with 452 seconds of spe-
cific impulse. Three of these engines were used (along with solid
rocket boosters) to launch the Space Shuttle into orbit. The
SSME has a wide throttle range from 67% to 109% of rated
thrust, and is unique among large engines in being designed
for reusability.9
The engine was designed to be reusable for up
to 100 flights.10
Although the last flight of the Space Shuttle
occurred in July 2011,11
16 functional Space Shuttle Main Engines remain in NASA inventory.12
The Space Launch System (SLS) currently under development by MSFC is a large expendable heavy-lift
launch vehicle which will make use of the RS-25 on the core stage.13
Four RS-25 engines will be required
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for each SLS vehicle, meaning the 16 engines currently in inventory are only sufficient for the first four SLS
launches. For subsequent launches, Aerojet Rocketdyne intends to restart production of the RS-25. The
manufacturer would like to integrate new manufacturing techniques developed since the engine was designed
and produced in order to reduce cost.12
Furthermore, since the SSME was designed for reusability but will
be used as an expendable engine on the SLS, there is the potential for design or operational changes to be
made in order to increase performance or decrease cost at the expense of reusability.
This paper discusses the use of a high-fidelity physics-based rocket engine cycle analysis model to demon-
strate the capability to perform rocket engine performance analyses. This analysis will be applied to the
RS-25 as a baseline proof-of-concept, making use of the extensive available literature on RS-25 performance.
This paper will focus on the engine powerhead, comprising the pumps, preburners, valves, and propellant
lines, with another researcher focusing on the combustion chamber and nozzle. The model will be used to
conduct simple trade studies, primarily to prove the capability as a basis for future work, but also with an
eye toward trading reusability for performance or cost for application to the SLS.
II. Approach and Methodology
The technical approach and methodology to be employed will be described, beginning with a brief de-
scription of cycle analysis. The selection of a tool for performance analysis, as well as the benchmarking
and trade study processes will be described.
A. Cycle Analysis
Strictly speaking, a thermodynamic cycle is a closed series of thermodynamic processes such that the system
returns to its initial state.14
In engineering, however, it is common to use the term to refer to processes
which are not closed, with the working fluid entering the system, being acted upon, and then leaving rather
than returning to its initial state and being reused.15
These cycles are frequently used to represent real-world
systems, such as the Brayton cycle used to model gas turbine engines. For more advanced analyses, a cycle
model is built from a variety of components such as compressors, combustion chambers, and turbines, each
of which has a specified impact on the properties of the fluid on which it acts. The cycle model for the
SSME, showing propellant flows and component operational parameters, is shown in Figure 2. The process
of linking these components together to represent a real system at a specific operating condition is called
cycle analysis. The object of cycle analysis is to determine high-level performance metrics of the system.
For a rocket engine, these metrics would be thrust and specific impulse, as well as turbomachinery operating
conditions and local propellant temperatures and pressures. This is used to predict engine performance at
different operating conditions and match experimental data. Cycle models can also be used in the other
direction, to determine the requisite propellant flow rates in order to achieve a specified thrust. This can be
used to design a flow schedule to match a desired thrust profile.
B. Tool Selection
Rocket engine performance is a crucial and well-studied problem in the aerospace field, and as such, a variety
of tools utilizing different approaches exist for analyzing engines. These different categories of tools fill differ-
ent niches and the correct tool choice for one problem might not be the correct choice for another. Generally
speaking the tool options can be split into three groups based on fidelity, though there is differentiation
within each group.
Low fidelity tools do not provide high analysis accuracy but generally run quickly and require less infor-
mation about the design. These kinds of tools are often implemented in a spreadsheet form or as a simple
computer program. For rocket engine analysis, the simplest version of this kind of tool would be a chemical
equilibrium analysis, essentially analyzing only the combustion chamber. More complex tools expand to
include more geometric information and components, such as the nozzle, pumps, and preburners, but the
analysis remains relatively simple.
High fidelity tools add significantly more geometric and operating information and generally require
more execution time. They can often analyze a wider range of designs, and provide more accurate results.
Compared to the low fidelity tools, they require more user expertise in order to provide meaningful input
and interpret the output. These kinds of tools frequently break the engine down into a number of more
simple components which are then connected, such as pumps, turbines, and nozzles. Unlike the low fidelity
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Figure 2. SSME Propellant Flow and Operational State, 104.5% Rated Power Level9
tools, the connections between the components are modeled, and the analysis for each component is more
sophisticated.
Extremely high fidelity tools include computational fluid dynamics and similar tools. These codes require
a great deal of user knowledge and setup in order to function well, and take the most computational time.
They do, however, provide extremely accurate results for nearly any design imaginable and across a wide
range of operating conditions. These tools are difficult to use in a parametric manner and require significant
human interaction, so they are generally reserved for the detailed design stage and are not integrated into a
sizing and synthesis environment.
For this application, it was decided that a high fidelity tool was the appropriate choice. The effects of
new manufacturing technologies can only be captured reliably at the component level, excluding the low
fidelity tools. The need for rapid execution in order to perform trade studies and the long-term goal of
integration into an integrated environment eliminates extremely high fidelity tools. One concern with high
fidelity tools is the need for a large number of inputs. The RS-25 is one of the most widely studied rocket
engines ever made, and the majority of the literature on it is in the public domain, providing an unusually
large amount of freely available technical data. This means the need for detailed inputs is not a concern for
this application.
In selecting a high fidelity tool, there were three leading options: a new custom-built tool, the Nu-
merical Propulsion System Simulation (NPSS), and NASA MSFC’s ROCket Engine Transient Simulation
(ROCETS).16
Building a custom tool provides the most control over inputs and outputs as well as underlying
physics, but requires a prohibitively large amount of development time and resources, as well as potentially
hindering integration with other tools. NPSS is a C++ based object-oriented tool developed by NASA
Glenn Research Center in collaboration with industry.17
NPSS is an industry standard, especially for gas
turbine engines, and is widely used for engine cycle analysis.18
It has been used for rocket engine perfor-
mance analysis in the past1920
and there is a great deal of existing literature. ROCETS is an internal MSFC
tool which the Aerospace Systems Design Laboratory acquired through a software licensing agreement. It
is conceptually similar to NPSS, connecting predefined components to create an engine.16
ROCETS has
more existing components for rocket engine analysis, while NPSS is a more general purpose tool. Due to its
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rocket-specific nature, ROCETS was selected for this performance analysis problem.
C. Benchmarking
A critical part of developing a model of a physical system is benchmarking that model. The benchmarking
process consists of using the model to predict system performance at a given operating condition, then
comparing the predicted performance to the actual performance at that point. The model can then be tuned
for a better match to the data, effectively ”anchoring” the model. For many propulsion systems, performance
data is tightly controlled in order to protect proprietary or export-controlled information. The SSME is an
excellent choice for benchmarking activities due to the abundance of published performance data9
.21
The
SSME model created was compared to the actual performance data at the operating conditions for which
high-quality data is available. The results of the benchmarking activities are described in Section III,
Subsection B.
D. Trade Studies
Once a model has been benchmarked, it can be used for a variety of purposes. In this case, the developed
SSME model will be used to perform trade studies around the design point in order to explore the design
space and identify opportunities to trade reusability for cost or performance improvements. These trade
studies consist of identifying metrics of interest and variables affecting those metrics, and systematically
perturbing those variables in order to observe the impact to the metrics. In order to thoroughly explore
the design space, a design of experiments can be used to select the perturbed variable values. The resulting
engine performance metrics can then be examined to identify trends and trade-offs which can be made in
the design. The design space exploration aspect of trade studies also provides a useful starting point for
design space exploitation, where the variables are changed in order to optimize the design relative to some
metric or combination of metrics. Design optimization is beyond the scope of this project. The results of
the conducted trade studies are discussed in Section III, Subsection C.
III. Results
With the goals of the modeling effort clearly defined, the methodology described in II was applied
to produce a functional SSME cycle model. This model was benchmarked against published SSME
performance data, and the model was used to evaluate the impact of design and operational changes to
engine performance, with an emphasis on the powerhead.
A. Modeling
A highly detailed cycle model of the SSME was provided as part of the ROCETS distribution, which was
modified for use on this project. The granularity provided by the model is extremely good, with every valve,
pipe, and volume included. On the timescale of this project, it was not feasible to build such a high-fidelity
model from scratch. A homebuilt lower-fidelity model encompassing only the major engine components was
feasible within the schedule constraints of the project and would have likely provided sufficient accuracy to
determine performance trends. This option was discarded in favor of the existing model, as the emphasis
of the project was on building familiarity with the tool, demonstrating capabilities, and performing trade
studies, not on engine cycle modeling.
The model consists of a variety of fundamental components connected to each other. Each component
has one or more inlets and one or more outlets, each of which is connected to another component. In
addition, some components can have other connections, such as a heat flux or a mechanical linkage. A
notional component showing these connection is shown in Figure 3. For example, the pump side of a simple
turbopump is connected to an upstream pipe (inlet), a downstream pipe (outlet) and a shaft from the turbine
(mechanical linkage). A flow splitter has a single inlet and two outlets, while a mixer has two inlets and one
outlet. All the components are connected to each other until the boundary conditions are reach, which in
this case are the liquid oxygen and liquid hydrogen tanks on the upstream side, and the ambient atmospheric
conditions on the downstream side.
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• Pump
• Turbine
• Pogo Suppressor
• Preburner
• Combustion Chamber
• Nozzle
• Pipe
• Valve
• Fluid Volume
Component
Fluid Inlet
Heat Flux
Shaft
Fluid Outlet
Figure 3. Notional Component Inputs and Outputs
In addition to the physical layout of the engine and the connections between components, the operation
of the engine is also controlled by the boundary conditions and the control logic. For this SSME model, the
boundary conditions are the tank and ambient conditions. The control logic is in two parts for the SSME
- the engine controller and the solver control. The engine controller is a module in the engine model which
controls a variety of valve positions based on current operational state. This models the engine control logic
on the real SSME. The solver control is user-defined and defines what the engine is ”run to.” These settings
are similar to the controls a pilot might have in an aircraft, namely throttle setting and mixture ratio. They
model engine operational settings which might be changed during a single engine firing, typically chamber
pressure and oxidizer-to-fuel ratio (O/F).
B. Benchmarking
Figure 4. SSME Performance Parameters9
The created SSME model was benchmarked against published
SSME performance data. The most complete performance
data, including propellant flow rates, component rotational
speeds, and localized temperatures and pressures, was avail-
able for the Nominal Power Level condition, 104.5% of rated
power. Detailed data about the engine state at that power
level is shown in Figure 2. High level performance parame-
ters for several power settings are shown in Figure 4. Bench-
marking cases were run at both the Nominal Power Level and
the Full Power Level (109% of rated power). Both thrust and
chamber pressure data are available for these thrust settings,
so the model was run twice, targeting both chamber pressure
and gross thrust, in order to determine the accuracy of the
results. Since the published data is in vacuum,9
the engine
model was operated at near-vacuum conditions, approximately
84 km (52.7 mi) altitude.22
The oxidizer-to-fuel ratio was fixed
for the benchmarking exercises at the nominal value of 6.032.9
The results of these two benchmarking analyses are summa-
rized in Tables 1 and 2. The calculated thrust values can be
seen to be slightly low when running to chamber pressure, and
the chamber pressure value is high by approximately the same amount when running to match thrust.
Specific impulse was consistently predicted to be approximately 1% lower than published.
Since more detailed performance data is available for the Nominal Power Level setting,9
that data was
also compared to the model predictions. These parameters are more detailed and are generally at the com-
ponent level, including the Low-Pressure Fuel Pump (LPFP), Low-Pressure Fuel Turbine (LPFT), High-
Pressure Fuel Pump (HPFP), High-Pressure Fuel Turbine (HPFT), Low-Pressure Oxygen Pump (LPOP),
Low-Pressure Oxygen Turbine (LPOT), High-Pressure Oxygen Pump (HPOP), High-Pressure Oxygen Tur-
bine (HPOT), Fuel Preburner (FPRB), Oxygen Preburner (OPRB), the main combustion chamber, and the
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Table 1. SSME Nominal Power Level Benchmark Results
Performance Parameter Published Value9 Match Pressure Match Thrust
Value Error Value Error
Gross Thrust (lbf) 491,900 479,400 -2.5% 491,900 0%
Chamber Pressure (psia) 2,871 2,871 0% 2,946 2.6%
Specific Impulse (s) 452 447 -1.1% 447 -1.1%
Table 2. SSME Full Power Level Benchmark Results
Gross Thrust (lbf) 512,900 502,300 -2.1% 512,900 0%
Chamber Pressure (psia) 3,008 3,008 0% 3,071 2.1%
Specific Impulse (s) 452 447 -1.1% 448 -0.9%
nozzle. These comparisons are shown in Table 3, in which it is clear that even at a highly detailed level, the
ROCETS engine model generally captures the actual engine behavior very closely. With a few exceptions,
the error for both Nominal Power Level runs was below 5% for each parameter examined. Importantly,
the parameters most closely tied with overall engine performance – chamber pressure, chamber tempera-
ture, and propellant flow rates – are extremely close to the published values. Considering these results, the
benchmarking effort was considered to be successful.
C. Trade Studies
The trade studies performed using the SSME model focused primarily on application to the Space Launch
System vehicle. Two major categories of studies were considered: operational changes and design changes.
Operational changes refer to not modifying the engine design, but operating it differently in order to extract
more performance at the cost of reusability or component life. Design changes require physically modifying
engine components in order to improve performance, reduce cost, or improve manufacturability. The trade
studies in this project were primarily operational changes, with some design changes applied to the power-
head. Another researcher conducted further trade studies on the combustion chamber and nozzle using the
same engine model.
The RS-25 engine application on SLS is significantly different from that on the Space Shuttle. In par-
ticular, the RS-25 was used on the Space Shuttle as an engine for use all the way to orbit, while on SLS it
will operate exclusively in the atmosphere as a first-stage engine. Additionally, the SLS vehicle is taller than
the Space Shuttle, which impacts the inlet conditions to the engine.23
Finally, as discussed previously, the
engine will be used in an expendable rather than a reusable configuration. The first two differences mean
that the baseline configuration for the SLS, around which trade studies will be performed, is not the same as
the baseline for the Space Shuttle. This baseline must be established before attempting any trade studies.
To account for the fact that on the SLS the SSME will operate exclusively in atmosphere, the SLS
baseline was set at sea level conditions (14.7 psi) rather than vacuum. To account for the difference in
vehicle heights, the change in inlet pressures was calculated. The pressure at the engine inlet consists of
both the tank pressure and the hydrostatic pressure due to the weight of the propellant. The engine inlet
pressure is described by Equations 1 and 2.
pengine = ptank + phydro (1)
phydro = ρgh (2)
The precise tank pressures are not easily available, though some approximate ranges are available,24
but what is known are the engine inlet conditions for the SSME as used on the Space Shuttle, which are
100 psi for the liquid oxygen and 30 psi for the liquid hydrogen.25
Using these values, and assuming that
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Table 3. SSME Nominal Power Level Detailed Benchmark Results
Fuel Flow Rate (lbm/s) 155 155 0% 159 2.6%
LPFP Speed (rpm) 15,519 14,994 -3.4% 15,238 -1.8%
LPFP Discharge Pressure (psia) 298 264 -11.4% 271 -9.1%
LPFT Pressure Ratio (none) 1.331 1.336 0.4% 1.340 0.7%
HPFP Speed (rpm) 34,311 34,125 -0.5% 34,758 1.3%
HPFP Discharge Pressure (psia) 5,956 6,097 2.4% 6,312 6.0%
HPFT Pressure Ratio (none) 1.551 1.502 -3.2% 1.512 -2.5%
FPRB Pressure (psia) 4,793 4,871 1.6% 5,042 5.2%
FPRB Temperature (◦
R) 1,770 1,726 -2.5% 1,756 -0.8%
Oxygen Flow Rate (lbm/s) 934 920 -1.5% 943 1.0%
LPOP Speed (rpm) 5,018 5,170 3.0% 5,253 4.7%
LPOP Discharge Pressure (psia) 421 406 -3.6% 411 -2.4%
LPOT Pressure Ratio (none) 9.029 9.466 4.8% 9.656 6.9%
HPOP Speed (rpm) 22,250 22,700 2.0% 23,141 4.0%
HPOP Discharge Pressure (psia) 4,025 4,043 0.4% 4,174 3.7%
HPOT Pressure Ratio (none) 1.553 1.519 -2.2% 1.529 -1.5%
OPRB Pressure (psia) 4,812 4,864 1.1% 5,032 4.4%
OPRB Temperature (◦
R) 1,331 1,465 10.1% 1,483 11.4%
Main Chamber Temperature (◦
R) 6,460 6,549 1.4% 6,553 1.4%
Nozzle Mass Flow Rate (lbm/s) 1,085 1,072 -1.2% 1,099 1.3%
the SLS will use similar tank pressures to the Space Shuttle, allows calculation of the SLS inlet pressures
by calculating the change in hydrostatic pressure. The top of the Space Shuttle liquid oxygen tank was
approximately 154 feet above the engines and the top of the liquid hydrogen tank was approximately 97
feet above the engines.24
The SLS vehicle is taller, with the top of the liquid oxygen tank 178 feet above
the engines and the top of the liquid hydrogen tank 128 feet above the engines26
.23
Using Equation 2, this
gives a change in hydrostatic pressure of approximately 12 psi for the liquid oxygen inlet and 1 psi for the
liquid hydrogen. Therefore, the inlet pressures for the SLS baseline will be 112 psi for the liquid oxygen
and 31 psi for the liquid hydrogen. The SSME model was run at these conditions, maintaining the Space
Shuttle Full Power Level chamber pressure (3,008 psi) and oxidizer-to-fuel ratio (6.032) to establish the SLS
baseline. The results are summarized in Table 4. From this table, the impact of both the shift to sea level
as well as the higher inlet pressures can be observed. The sea level condition causes an 18% decrease in
specific impulse and thrust due to the backpressure and flow overexpansion. The increase in engine inlet
pressure causes a decrease in pump rotational speed as the higher inlet pressure means the propellant does
not require as much of a pressure increase from the pump before entering the combustion chamber, allowing
the pumps to run more slowly.
Table 4. Space Shuttle and SLS Baseline Engines
Parameter Space Shuttle Baseline Space Shuttle Sea Level SLS Baseline
Gross Thrust (lbf) 502,300 412,600 412,600
Specific Impulse (s) 447 368 368
LPOP Speed (rpm) 5,323 5,323 5,315
HPOP Speed (rpm) 23,508 23,508 23,479
With the SLS baseline established, trade studies were performed around that point. First, the effect of
increasing chamber pressure was examined. Notionally an expendable SSME might be able to support a
8 of 12

higher chamber pressure as there is no need for the engine to withstand multiple start/shutdown cycles. Two
pressure increase cases were examined, one with a 10% increase to main chamber pressure, and one with an
increase in engine thrust to recover the loss due to atmospheric backpressure, increasing the sea level thrust
to 500,000 lbf. These changes are focused on performance enhancement, not on weight or cost reduction.
The impact of these changes is shown in Table 5.
Table 5. SLS Chamber Pressure Performance Impacts
Performance Parameter SLS Baseline
10% Pressure Boost 500klbf Thrust
Value Change Value Change
Gross Thrust (lbf) 412,600 461,300 11.8% 500,000 21.2%
Specific Impulse (s) 368 375 1.9% 381 3.5%
Nozzle Flow Rate (lbm/s) 1,122 1,230 9.6% 1,314 17.1%
Chamber Pressure (psia) 3,008 3,300 9.7% 3,532 17.4%
Chamber Temperature (◦
R) 6,556 6,572 0.2% 6,582 0.4%
LPOP Speed (rpm) 5,315 5,640 6.1% 5,898 11.0%
LPOP Discharge Pressure (psia) 425 447 5.2% 466 9.6%
HPOP Speed (rpm) 23,479 25,164 7.2% 26,461 12.7%
LPFP Speed (rpm) 15,344 16,336 6.5% 17,045 11.1%
LPFP Discharge Pressure (psia) 273 299 9.5% 316 15.8%
HPFP Speed (rpm) 35,218 37,660 6.9% 39,619 12.5%
HPFP Discharge Pressure (psia) 6,459 7,325 13.4% 8,048 24.5%
Increasing the chamber pressure succeeded in increasing the engine thrust, as expected, but the other
effects of the higher pressures are important to note. Much of this thrust increase came from higher mass
flow through the engine. This will cause the vehicle’s tanks to empty more quickly, which may or may not be
acceptable depending on vehicle design but must be considered. The higher pressures also boosted specific
impulse slightly, which linearly increases the amount of velocity change available for a given propellant
mass, so even a small change in specific impulse can provide significant benefit to the vehicle design as a
whole. Examining the engine powerhead, it is observed that the higher chamber pressures also drive higher
pump speeds and discharge pressures. This may be problematic for the implementation of a higher chamber
pressure variant, as the pumps are already operating at high speed and may not have much rotordynamic
margin to be run even faster. Higher pump discharge pressures may lead to housings or plumbing carrying
unsafely high stress, especially since the pump discharge pressures (for the high-pressure pumps) increase by
more than the chamber pressure to account for higher line losses and required power to drive the turbine.
These higher pressures and speeds may be acceptable for an expendable engine, depending on the specific
details of the existing design and margins.
The other operational change examined was changing the oxidizer-to-fuel ratio. Generally engines are
designed to support a range of O/F values, so relatively small changes in O/F likely do not require a change
in the engine design. Two analyses were performed with different O/F ratios, one running rich (low O/F)
and one lean (high O/F). For both cases the chamber pressure was maintained at the 10% boosted level
described in Table 5. As can be seen in Table 6, the impact on most high-level performance parameters was
small, even relative to the small change in O/F which was analyzed. Utilizing a low O/F (running rich)
had several positive performance impacts, though. Both the nozzle flow rate and the chamber temperature
decreased, and specific impulse was improved. Thrust was very slightly reduced, however. A high O/F had
the opposite effect. Depending on the goal of the design changes, either approach might be justifiable. If there
are concerns about high chamber temperatures presenting a dangerous operating condition for the engine,
decreasing O/F could counteract the temperature increase due to higher chamber pressures, without giving
up much of the thrust gained by doing so. O/F trades cannot be viewed in isolation, however. Decreasing
the O/F requires more fuel for the same mass of oxygen, increasing required tank mass and volume, which
can be significant even for small changes. For the Space Shuttle, going from an O/F of 6.032 to 6.000, as
analyzed here, would require an additional 1,263 lbm of liquid hydrogen, corresponding to 285 more cubic
feet of tank volume.24
Any change to the oxidizer-to-fuel ratio must incorporate analysis of the full vehicle
9 of 12

for this reason.
Table 6. SLS Oxidizer-to-Fuel Ratio Performance Impacts
Performance Parameter High Pressure
High O/F Low O/F
Oxidizer-to-Fuel Ratio 6.032 6.080 0.796% 6.000 -0.531%
Gross Thrust (lbf) 461,300 461,500 0.043% 461,200 -0.022%
Specific Impulse (s) 375.1 374.6 -0.133% 375.6 0.133%
Nozzle Flow Rate (lbm/s) 1,230 1,232 0.163% 1,228 -0.163%
Chamber Temperature (◦
R) 6,572 6,594 0.335% 6,558 -0.213%
The final trade study performed for this project was a design change to the engine’s pumps. Pump
analysis often makes use of the dimensionless parameter ψ, the head coefficient, to describe the pressure rise
across the pump and compare pumps of dissimilar sizes or rotational speeds. This parameter is a function of
the design of the pump and is given by Equation 3, where the rotational speed N is in radians per second.27
ψ =
∆p
ρ (Nr)
2 (3)
For two pumps with the same rotational speed and radius acting on the same fluid, the pump with the
higher head coefficient will produce a higher discharge pressure. Conversely, a pump with a lower head
coefficient and the same radius and working fluid will have to rotate more rapidly to produce the same
pressure rise. Generally, a higher rotational speed increases the stress within the pump body and shaft, and
may cause the pump to operate with less rotordynamic margin, but can reduce the weight of the pump by
reducing the radius required for a given pressure rise. This is why rocket engine turbopumps tend to operate
at much higher speeds than industrial pumps.28
For this trade study, a scalar multiplication factor was applied to the head coefficient maps used by the
ROCETS model. In the interest of mass reduction, factors less than 1 were considered. Since pressure rise
is a function of radius for a centrifugal pump,28
this is roughly equivalent to scaling down the physical size
of the pump while keeping the overall design the same, or even trimming down the radius of an existing
pump, though there are other factors to consider in that case. This would reduce mass by lowering the mass
of the pump itself, as well as the pump housing, which is generally a fairly large component of engine mass.
The changes in this study were compared to the SLS baseline model, and were operated at the baseline
O/F and chamber pressure. As shown in Table 7, the effect on pump speeds due to the head coefficient
reduction is as expected. The pump discharge pressures were largely unchanged, though the distribution
of the pressure rise did tend to shift toward the low-pressure side. Most likely, a 65% reduction in head
coefficient is unreasonably large, especially due to the significantly increased rotational speeds it requires,
but the trend provides a useful direction in which to proceed.
Table 7. SLS Pump Head Coefficient Performance Impacts
Performance Parameter SLS Baseline
85% ψ 65% ψ
LPOP Speed (rpm) 5,315 5,649 6.3% 6,272 18.0%
LPOP Discharge Pressure (psia) 425 438 3.1% 444 4.5%
HPOP Speed (rpm) 23,479 24,865 5.9% 27,421 16.8%
LPFP Speed (rpm) 15,344 16,621 8.3% 19,199 25.1%
LPFP Discharge Pressure (psia) 273 283 3.7% 298 9.2%
HPFP Speed (rpm) 35,218 37,638 6.9% 42,757 21.4%
HPFP Discharge Pressure (psia) 6,459 6,449 -0.2% 6,472 0.2%
10 of 12

IV. Conclusions
This investigation has produced a proof of concept for parametric design studies of rocket engines. With
advanced technologies, novel architectures, and an increased emphasis on integration, the need for such
capability early in the design process has increased dramatically. The modeling components provided by
ROCETS allow the development of high-fidelity engine models which can both be used early to inform design
decisions as well as during detailed design to refine and optimize the engine. The trade studies performed
in this project are only a preliminary look at the options available, but already some trends can be seen
and possible opportunities identified. For example, in order to get the maximum engine performance, it
is desirable to have the highest possible chamber pressure. Mass can be reduced by trimming pumps for
the same chamber pressure, increasing the pump rotational speed. In both cases, hardware limitations like
rotordynamic margin and the stress in pressure vessel walls must be taken into account to ensure the safety
of the engine.
When investigating different design options as analyzed here, the ultimate goals of the change must also
be considered. It is unlikely that one solution will improve everything of interest – there will almost certainly
be tradeoffs. For example, for the application of the RS-25 to the SLS, affordability has been deemed of
higher importance than performance.12
Therefore, a push to increase engine thrust might not be the best
solution. Instead, if a lower thrust is acceptable, the chamber pressure could be decreased relative to the
SSME baseline in concert with trimming the pumps. This reduces the thrust and efficiency of the engine,
but the lower pressures and pump diameter could reduce the mass and cost of the engine through smaller
and thinner pump housings. If the vehicle design allows, the specific impulse lost from the lower chamber
pressure could be partially recovered through a shift in oxidizer-to-fuel ratio to run slightly more rich (lower
O/F).
While the capability and utility of parametric rocket engine analysis has been developed in this investiga-
tion, there is much future work that could be done. In particular, improving the interface with the analysis
tools would greatly assist the designer in performing trade studies more rapidly. Examination of a larger
set of design parameters as well as the varying levels of fidelity which ROCETS can provide would identify
which parameters have the greatest impact on metrics of interest and allow the selection of the appropriate
level of fidelity for different applications, omitting or defaulting less impactful parameters. The ultimate
goal with this tool is to integrate it into a unified sizing and synthesis environment for launch and space
vehicles, allowing the consideration of the system-level effects of subsystem changes even at a preliminary
design stage.
References
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13Stanfield, J., “Space Launch System,” Tech. rep., NASA George C. Marshall Space Flight Center, June 2012.
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Engines, Cambridge University Press, 2nd ed., 2003.
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