Michael L. Belair, Charles J. Sarmiento, and Thomas M. Lavelle

Glenn Research Center, Cleveland, Ohio

Nuclear Thermal Rocket Simulation in NPSS

NASA/TM—2013-216553

July 2013
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Michael L. Belair, Charles J. Sarmiento, and Thomas M. Lavelle

Glenn Research Center, Cleveland, Ohio

Nuclear Thermal Rocket Simulation in NPSS

NASA/TM—2013-216553

July 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

49th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

San Jose, California, July 14–17, 2013
Acknowledgments

The authors would like to thank Stan Borowski, Jim Fittje, and Bruce Schnitlzer for their support and technical input on the 

project. This work was performed with funding from the Nuclear Cryogenic Propulsion Stage (NCPS) project established as part 

of the Advanced Exploration Systems Program.

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Level of Review: This material has been technically reviewed by technical management. 
Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories.

One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor

developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in

the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramicmetallic (CERMET) fuel element. Within each category, a small and a large thrust engine

are modeled. The small engine models utilize RL-10 turbomachinery performance maps

and have a thrust of approximately 33.4 kN (7,500 lbf). The large engine models utilize

scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2

kN (25,000 lbf). Power deposition profiles for each reactor were obtained from a detailed

Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as

thermodynamic state points, thrust, specific impulse, reactor power level, and maximum

fuel temperature are analyzed for each engine design.

Nomenclature

AR Area Ratio

CERMET Ceramic-Metallic fuel element

DRA Design Reference Architecture

g Acceleration of gravity at sea level

γ Ratio of specific heats

GRC Glenn Research Center

Isp Specific Impulse

kN kilownewtons force

lbf pounds force

MCNP Monte Carlo N-Particle transport code

MW Megawatts

MW Molecular Weight

NCPS Nuclear Cryogenic Propulsion Stage

NERVA Nuclear Engine for Rocket Vehicle Application

NESS Nuclear Engine System Simulation

NPSS Numerical Propulsion System Simulation code

NTP Nuclear Thermal Propulsion

NTR Nuclear Thermal Rocket

Pe Exhaust Pressure

Pi Inlet Pressure

R Universal Gas Constant

SNRE Small Nuclear Rocket Engine

T Temperature

NASA/TM—2013-216553 1

Nuclear Thermal Rocket Simulation in NPSS 

Michael L. Belair, Charles J. Sarmiento, and Thomas M. Lavelle 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 
I. Introduction

N

uclear thermal propulsion (NTP) is identified as the preferred propulsion technology for manned missions to Mars in NASA’s Design Reference Architecture (DRA) 5.0.1 NTP is the preferred propulsion

technology because it has heritage and higher performance than current propulsion systems available at the

thrust levels necessary for manned missions. The Nuclear Cryogenic Propulsion Stage (NCPS) project is

engaged in conceptual design of a small 33.4 kN (7,500 lbf) and a full-size/large 111.2 kN (25,000 lbf) class

nuclear rocket. The small engine could be used for a flight demonstration mission as well as possible robotic

probe missions. The large nuclear engine is intended to be used in a clustered configuration for human

exploration missions.2

In support of the NCPS project, the NASA Glenn Research Center (GRC) has been working to model

four nuclear thermal rockets in the Numerical Propulsion System Simulation (NPSS) framework. The main

intent of this work is to analyze the overall system performance and operation of a nuclear thermal rocket.

Existing turbomachinery design were simulated to mitigate the eventual cost of a turbomachinery development effort. The four engine models are broken into two categories based on the fuel element type. One fuel

element is based on the design from the Nuclear Engine for Rocket Vehicle Application (NERVA) project

and another is a ceramic-metallic (CERMET) fuel element design. The small engine models utilize a RL-10

type hydrogen turbopump and the larger engine models attempt to utilize a dual RL-60 type hydrogen

turbopump configuration.

The Rover program which included the Nuclear Engine for Rocket Vehicle Application (NERVA) project

demonstrated the concept of a nuclear rocket was feasible. The Rover/NERVA program was quite successful

and developed twenty rocket reactors that were built and ground tested. The reactor design that resulted

from the NERVA project is the basis for one of the reactor sets analyzed.3

A. Advantages of Nuclear Propulsion

The overall concept of a nuclear rocket is simple. In most conventional rocket systems, an extremely energetic

and hot gas is produced and that energy is converted into kinetic energy by expanding the gas through a

nozzle. In chemical propulsion systems, the energy is derived from the energy stored in chemical bonds.

In a nuclear rocket, the energy is supplied by heating the hydrogen in a nuclear reactor core. The most

common attribute evaluated to determine the performance of a rocket system is the specific impulse, Isp.

For the acceleration of gases in a nozzle, the specific impulse is a function of the absolute temperature and

the molecular weight as given in Equation 1.

Isp =

1

g

vuut

2γ

γ − 1

RT

MW 

1 −



Pe

Pi

 γ−1

γ

!

(1)

The advantage of the nuclear rocket is that you are no longer dependent upon chemical reactions to obtain

the high temperatures, so the molecular weight can be minimized and specific impulse can be maximized.

Currently the highest performance propellant combination commonly used for chemical systems is hydrogen

and oxygen. At a mixture ratio of 5.85:1, the combustion temperature is approximately 3420 K, the average

molecular weight is 13.8 g/mol, and the maximum theoretical specific impulse is near 480 seconds. In

a nuclear rocket, the propellant is typically H2 with a molecular weight of 2.016 g/mol and is heated to

temperatures near 2700 K. Based on the equation for specific impulse, the nuclear rocket has twice the

expected specific impulse, approximately 900 seconds, as hydrogen and oxygen chemical propulsion.

The other advantage of nuclear systems is that technology advances could potentially increase performance. In a chemical system, you are ultimately limited by the energy available from the chemical reaction;

however, in a nuclear rocket, the temperature of the propellant is only limited by the temperature limits of the reactor. Advanced reactor fuel design may be capable of increasing the propellant temperature

significantly.

B. Numerical Propulsion System Simulation (NPSS) Framework

The Numerical Propulsion System Simulation (NPSS) software is a propulsion system simulation tool developed by NASA GRC and its industry and Government partners.4 NPSS is used in the aerospace industry

to simulate aircraft engines and their components in a timely and cost-effective manner. NPSS has also

NASA/TM—2013-216553 2
been expanded to spacecraft applications due to its flexible nature. It allows developers to create their own

custom elements that can be used in system simulations. For the NCPS modeling project, custom reactor

and tie tube elements were developed to interface in the NPSS framework. The theory behind the reactor

and tie tube thermodynamics and fluid mechanics is summarized elsewhere.5, 6

II. Turbomachinery Modeling

A. RL-10 Hydrogen Pump and Turbine

The small class of engines utilizes the RL-10 hydrogen pump and turbine to power the engine. All RL-10

variants operate based on an expander cycle where the liquid hydrogen is heated in the regenerative nozzle

and chamber and used to drive the turbine. The same type of cycle can be run for a nuclear rocket where

the heat is obtained from the nozzle cooling section and the tie tubes. As of 2012, there are two variants of

the RL-10 in use.7 The two engines are summarized in Table 1.

Table 1. RL-10 Variants7

RL-10A4-2 RL-10B-2

Propellants LOX/LH2 LOX/LH2

Thrust (kN) 99.2 110.1

Chamber Pressure (MPa) 4.21 4.41

Mixture Ratio 5.5:1 5.85:1

Specific Impulse (seconds) 451 464

Area Ratio 84:1 285:1

Hydrogen Mass Flow Rate*(kg/sec) 3.45 3.54

* Calculated based on thrust, specific impulse and mixture ratio.

The decision to simulate RL-10 turbomachinery was made because the hydrogen flow rate for the small

nuclear rocket engine is very similar to the hydrogen flow rate of the RL-10. The expected flow rate needed

for the small engine class is approximately 3.81 kg/sec based on a thrust of 33.4 kN and a specific impulse

of 900 seconds. This flow rate represents an increase less than ten percent from the RL-10B-2 design.

B. RL-60 Hydrogen Pump and Turbine

The large class of nuclear engines, 111.2 kN thrust, utilizes the turbomachinery from the RL-60 design to

pump the liquid hydrogen through the reactor core. The RL-60 is an upgraded design from the RL-10. The

RL-60 was designed to have a thrust of 266.9 kN and a specific impulse of 465 seconds. The mixture ratio

of the RL-60 was intended to be operated at a range of 5.0 to 6.0:1. The hydrogen flow rate through the

engine is calculated to be 8.35-9.75 kg/sec. This flow rate is smaller than the expected flow rate of the large

nuclear engines. The large nuclear engines are expected to have a flow rate of approximately 12.7 kg/sec

based on a thrust of 111.2 kN and a specific impulse near 900 seconds.

Since there is approximately a 30-40 percent greater mass flow rate than the design level of the RL-60, it

was decided to use a dual turbopump configuration. In this configuration, two RL-60 turbopump assemblies

will operate in parallel at approximately 70 percent capacity in order to provide the necessary mass flow

rate. In the event of a failure of one turbopump assembly, the engine could be operated at a lower power

and thrust level if desirable.

III. Fuel Element Descriptions

A. NERVA

The NERVA derived engines utilize the (U,Zr)C-graphite hexagonal fuel element with structural support tie

tubes that was the baseline design for the Small Nuclear Rocket Engine (SNRE).8 The fuel element has a

flat to flat measurement of 1.905 cm and is 88.9 cm long for the small engine and 132.1 cm long for the

large engine. Each fuel element contains 19 coolant passages with 0.257 cm diameter holes for the hydrogen

NASA/TM—2013-216553 3
coolant. The outer surface of the fuel element and the interior of the coolant channels are coated with

a protective layer of ZrC. The outer surface is coated to a thickness of 50 micrometers, and the coolant

channels have a tapered coating varying from 50 micrometers at the inlet to 150 micrometers at the exit.

More information on the NERVA derived fuel elements can be found in the references written by Schnitlzer

and Fittje.9, 10 Figure 1 below shows the details for the fuel element and tie tube geometry.10

Figure 1. NERVA Derived Fuel Element and Tie Tube

B. ANL-200 CERMET

The CERMET engines utilize the fuel element known as ANL-200. The fuel element is composed of Gd2O3

stabilized UO2 particles in a tungsten matrix. The fuel element has a flat to flat measurement of 2.774 cm

and is 60.96 cm long for the small engine and 86.36 cm long for the large engine. Each fuel element contains

61 coolant passages with 0.170 cm diameter holes for the hydrogen coolant. The outer surface of the fuel

element and the interior of the coolant channels are coated with a 0.01778 cm protective layer of W-25 wt%

Re. Figure 2 below shows the general arrangement of the fuel element in a reactor core.11

Figure 2. ANL-200 Fuel Element

NASA/TM—2013-216553 4
IV. Small NERVA Derived Engine

A. Engine Configuration

Since thrust is proportional to reactor power, the small engine is modeled in order to have enough thermal

power to produce approximately 33.4 kN of thrust. All of the engines tested during the NERVA project were

designed for much higher thrust levels; Peewee was the smallest at 25 klbf.

3

In order to maintain criticality

in the engine at such low thrust levels, a slightly different engine configuration had to be devised where the

number of tie tubes was increased to a ratio near 1:1. The tie tubes improve criticality of the engine because

they contain a zirconium hydride moderator. Figure 3 shows a detailed view of the general pattern concept.9

The engine configuration has a total of 260 fuel elements and 251 tie tube elements. The reactor thermal

power output is set at approximately 162 MW. The reactor core is surrounded by a 14.7 cm thick radial

reflector and control drums. The engine core is estimated to weigh 1432 kg, containing approximately 27.5

kg of 235U. Figure 4 is an illustration of the entire core configuration.9

Figure 3. Dense (1:1) Fuel Element and Tie Tube Configuration

Figure 4. Small NERVA Core Configuration

A general description of the reactor and power distribution through the system is shown in Table 2.

NASA/TM—2013-216553 5
Table 2. Small NERVA Derived Engine Summary

Description Value Units

Number of Fuel Elements 260 NA

Number of Tie Tubes 251 NA

Avg. Power Deposited Fuel Element 0.586579 MW

Reactor Total Power 161.6 MW

Fuel Element Hydrogen Heating 140.7 MW

Tie Tube Hydrogen Heating 16.6 MW

Conduction from Fuel Elements 11.78 MW

Nuclear Deposited 4.84 MW

Control Drum Hydrogen Heating 3.45 MW

B. Engine Flow Path

One of the main design constraints of a nuclear rocket system is obtaining enough thermal energy to run the

turbomachinery. For the NERVA derived engines, this is typically not an issue because the pressure drop

through the engine is relatively low and there is plenty of available thermal energy from the tie tubes. In

fact, there is significantly more thermal energy available than necessary to run the turbomachinery. In order

to keep the turbine inlet temperature closer to the heritage RL-10 temperatures, an innovative flow path

had to be considered to lower this temperature as much as possible.

First, the liquid hydrogen leaves the tank and passes through the fuel inlet valve. The hydrogen is then

compressed through a two stage pump. Then the flow is split into two parallel paths. Part of the flow is

directed to the tie tubes which provide structural support for the reactor as well as criticality. The remaining

flow enters the nozzle and chamber regenerative cooling section. After the flow leaves the regenerative cooling

section, it is directed up along the reactor to cool the control drums and radial reflectors. Once the flow

exits the control drum/ radial reflector, it is recombined with a small fraction of the flow from the tie tubes.

This combination is sent to the turbine. Some of the flow is allowed to bypass the turbine in order to control

the flow rate of hydrogen through the pumps. Once the flow from the turbine, turbine bypass valve, and

remaining tie tube flow recombine, it is directed to the reactor inlet. The gas that enters the reactor is then

superheated to extremely high temperatures before it is exhausted through a nozzle. An illustration of the

flow path can be seen in the model results for this engine, shown in Figure 5.

C. Model Results

The model was run to obtain a chamber pressure of 4.14 MPa (600 psia) and a peak fuel temperature of 2860

K (5148 R). This specific state was chosen due to turbomachinery pump discharge pressure limitations and

fuel element temperature limitations. It should be noted that the pressure drop through the regenerative

cooling sections are estimated using resistance values that scale the pressure drop with the mass flow rate

of the engine.

NASA/TM—2013-216553 6
Figure 5. Small NERVA System Results

A comparison summary of the reactor is shown in Table 4. It should be noted that the component mass

section is only an estimate. The reactor mass is estimated based on the results from the MCNP analysis.9

The pressure vessel, nozzle, and other numbers are based on scaling factors developed during the NERVA

project. The scaling factors are used to estimate mass as a function of the thrust and the overall size of the

reactor core.

D. Discussion

There were several difficulties in the design of this engine. The biggest issue to resolve was the temperature

of the hydrogen entering the turbine. Since the RL-10 operates on an expander cycle, the temperatures

experienced by the turbomachinery are quite low. However, the total energy available from the small

NERVA engine design is much greater than necessary. This is due to the large fraction of tie tubes in the

engine that are necessary to maintain criticality at such a low thrust level. In order to compensate for this, a

small bleed from the tie tubes was used rather than the mixing the entire flow. However, this design feature

introduces further risk since it complicates the flow path of the engine and also results in a very high tie tube

outlet temperature. Further analysis is necessary to see if the tie tube elements can withstand the thermal

stress from this high temperature environment.

As shown in Table 4, the engine achieved a thrust to weight ratio of 1.96 and a specific impulse of approximately 900 seconds. The specific impulse is primarily driven by the maximum allowable fuel temperature.

In this analysis, the assumed maximum fuel temperature was 2860 K (5148 R). The thrust to weight of the

engine is primarily related to the mass of the reactor core since this is the largest contributor. A lower thrust

to weight is another complication from the low thrust level. A low thrust level requires more tie tubes and

larger radial reflectors. This significantly reduces the thrust to weight of this engine. More work will need

NASA/TM—2013-216553 7
to be performed to identify if the reactor mass can be reduced by returning to a 2:1 fuel element to tie tube

ratio from the current 1:1 ratio with some other approach to enhance criticality.

V. Large NERVA Derived Engine

A. Engine Configuration

The large engine is designed to produce enough thermal power to generate approximately 112.1 kN of thrust.

Since this engine is much larger than the 33.4 kN NERVA derived engine, fewer tie tubes are needed to provide

adequate criticality. For this engine, the fuel element to tie tube ratio was changed to approximately 2:1

which is the ratio studied for the small nuclear rocket engine (SNRE). Figure 6 shows a detailed view of

the general pattern concept.9In total, the engine has 564 fuel elements and 241 tie tube elements with a

total thermal power of 555 MW. The reactor core is surrounded by 14.7 cm thick radial reflector and control

drums. The core is estimated to have a total mass of 2645 kg, containing 36.8 kg of 235U.

Figure 6. SNRE (2:1) Fuel Element and Tie Tube Configuration

The fuel element to tie tube pattern shown in Figure 6 is repeated throughout the reactor. Near the edge

of the core, there are beryllium filler elements that are used to increase criticality and to make to the core

circular. The power distribution in this reactor is given in Table 3.

Table 3. Large NERVA Derived Engine Summary

Description Value Units

Number of Fuel Elements 564 NA

Number of Tie Tubes 241 NA

Avg. Power Deposited Fuel Element 0.9386027 MW

Reactor Total Power 555 MW

Fuel Element Hydrogen Heating 529.3 MW

Tie Tube Hydrogen Heating 39.1 MW

Conduction from Fuel Elements 27.1 MW

Nuclear Deposited 12.0 MW

Control Drum Hydrogen Heating 10.9 MW

B. Engine Flow Path

As with the small NERVA engine, there is plenty of thermal energy available to power the turbomachinery.

The main difference between the small engine and the large engine is that the engine is operated with dual

NASA/TM—2013-216553 8
turbopumps in parallel. Since the ratio of tie tubes to fuel elements is lower for this reactor, there is less

need to bypass some of the tie tube exit flow. The flow path is illustrated in in the results section in Figure 7.

C. Model Results

The model was run to obtain a chamber pressure of 6.89 MPa (1000 psia) and a peak fuel temperature of

2860 K (5148 R). This specific state was chosen due to turbomachinery pump discharge pressure limitations

and fuel element temperature limitations. It should be noted that the pressure drop through the regenerative

cooling sections are estimated using resistance values that scale the pressure drop with the mass flow rate

of the engine.

Figure 7. Large NERVA System Results

A comparison summary of the reactor is shown in Table 4. It should be noted that the component mass

section is only an estimate. The reactor mass is estimated based on the results from the MCNP analysis.9

The pressure vessel, nozzle, and other numbers are based on scaling factors developed during the NERVA

project. The scaling factors are used to estimate mass as a function of the thrust and the overall size of the

reactor core.

NASA/TM—2013-216553 9
D. Discussion

One major concern for the NERVA derived engines is that the turbine inlet temperatures may exceed the

nominal temperatures experienced in an expander cycle. As with the small NERVA derived engine, the large

engine transfers a significant amount of energy to the hydrogen in the tie tubes. As discussed before, this

is not considered as much of an issue for the large engine since the fuel element to tie tube ratio is greater

which lowers the turbine inlet temperature.

Another interesting result is that the specific impulse of the large engine is greater than the small engine

by approximately 12 seconds. Both the small and large NERVA derived engines were analyzed with an

assumed peak fuel temperature of 2860 K (5148 R). The higher specific impulse of the large engine is

attributed to the longer fuel elements which have a greater heat transfer area. This larger area results in

a lower thermal gradient between the fuel element and hydrogen coolant. The large engine also has much

greater performance than the small engine in terms of thrust to weight. This is mostly attributed to the

larger fuel element to tie tube ratio which resulted in a approximately half the number of reactor elements

per unit thrust.

VI. Small ANL-200 Engine

A. Engine Configuration

The small CERMET engine is based on the ANL-200 fuel element design. The engine core contains a total

of 121 fuel elements for a total reactor thermal power of 160 MW.9 Since this is a fast spectrum reactor,

no neutron moderation is required and the core is only composed of fuel elements. The reactor core is

surrounded by a 10 cm thick BeO radial reflector as well as control drums to improve criticality. In order to

reduce neutron leakage, there is a 15.24 cm axial reflector at the inlet of the reactor core. The reactor core

is estimated to have a mass of 1000 kg and contains 177.3 kg of 235U.

B. System Results

The small CERMET engine utilizes a flow path that is nearly identical to the flow path for the small NERVA

derived engine. The only difference is that the CERMET engine does not have a parallel flow for tie tubes

since there are no tie tubes in the system. The simulation set point was a chamber pressure of 3.79 MPa

(550 psia) and a maximum fuel temperature of 2860 K (5148 R). In this system, the chamber pressure is

limited due to the available energy to run the turbomachinery.

NASA/TM—2013-216553 10
Figure 8. Small ANL-200 Engine System Results

A comparison summary of the reactor is shown in Table 4. It should be noted that the component mass

section is only an estimate. The reactor mass is estimated based on the results from the MCNP analysis.9

The pressure vessel, nozzle, and other numbers are based on scaling factors developed during the NERVA

project. The scaling factors are used to estimate mass as a function of the thrust and the overall size of the

reactor core.

C. Discussion

The small ANL-200 CERMET engine achieves a thrust of 33.8 kN (7591 lbf) and a specific impulse of

885.4 seconds. The specific impulse is lower than the comparable small NERVA derived engine primarily

because the chamber temperature is lower for the CERMET engine for the same peak fuel temperature.

This is likely because the fuel element is thirty percent shorter than the NERVA fuel element which results

in a higher thermal gradient between the fuel element and the hydrogen gas. The major concern for the

CERMET engine is the available thermal energy to run the turbomachinery. In this analysis, the maximum

chamber pressure attainable is near the operation point, 3.79 MPa (550 psia). This occurs primarily because

the CERMET engine does not have regeneratively cooled tie tubes to increase propellant temperature. The

CERMET engine also has approximately one third of the energy deposition in the reflector and control drums

as compared to the NERVA derived engine. The amount of available energy is an issue when attempting to

scale this design to a higher thrust level.

The small CERMET engine achieves a higher thrust to weight ratio than the small NERVA engine. This

is primarily because the CERMET engine is more compact since it does not require tie tubes. The majority

of the mass in any system is the reactor core. Although CERMET fuel is more dense, the smaller volume core

is lighter and requires a lighter structure, radial reflector, and pressure vessel. All of these factors contribute

to making the reactor substantially lighter than the comparable NERVA engine.

NASA/TM—2013-216553 11
VII. Large ANL-200 Engine

A. Engine Configuration

The large CERMET engine is also based on the ANL-200 fuel element design. The reactor design analyzed

is regarded as the radial growth option since the additional power compared to the small engine is obtained

primarily by adding fuel elements. However, the fuel element length was also extended to 86.36 cm (34

inches). The engine core consists of 241 fuel elements for a total reactor power of 513 MW. The reactor core

is surrounded by a 10 cm thick radial reflector as well as control drums to improve criticality. In order to

reduce neutron leakage, there is a 15.24 cm axial reflector at the inlet of the reactor core. The reactor core

is estimated to have a mass of 2356 kg and contain 376.7 kg of 235U.

B. System Results

The overall engine system for the large ANL-200 reactor is based around a hot bleed cycle rather than the

typical expander cycle. This is due to a combination of factors. First, lack of sufficient thermal energy

for the turbomachinery in the CERMET systems results in a limit to the maximum thrust obtainable.

Secondly, the reactor design for the large ANL-200 engine is an exasperating condition. In order to maintain

a relatively compact engine core diameter, the fuel element had to be extended to 86.36 cm from 60.96

cm which correspondingly increased the power per fuel element. The combination of increased length and

increased power per fuel element greatly increased the pressure drop across the reactor and therefore required

a higher pump discharge pressure.

In order to obtain the necessary energy to drive the turbomachinery, a hot bleed cycle had to be used.

However, the hot bleed cycle reduces the average specific impulse of the system by as much as 65 seconds

as illustrated in Figure 9. The hot bleed cycle is implemented by mixing a portion of the flow exiting the

control drum/reflector cooling loop with a bleed from the outlet of the reactor core. The ratio of cool to

hot flow was proportioned in order to obtain a turbine inlet temperature of 833 K (1500 R). The turbine

exhaust is directed through a 2:1 area ratio nozzle in order to recuperate some of the losses from a hot bleed

cycle. The remaining flow from the control drum/reflector outlet is directly fed into the reactor core. The

results are shown in Figure 9 and Table 4.

NASA/TM—2013-216553 12
Figure 9. Large ANL-200 Hot Bleed Cycle Results

A comparison summary of the reactor is shown in Table 4. It should be noted that the component mass

section is only an estimate. The reactor mass is estimated based on the results from the MCNP analysis.9

The pressure vessel, nozzle, and other numbers are based on scaling factors developed during the NERVA

project. The scaling factors are used to estimate mass as a function of the thrust and the overall size of the

reactor core.

C. Discussion

The major issue with the large CERMET engine using the ANL-200 fuel element geometry is the pressure

drop across the reactor and the lack of available thermal energy for the turbomachinery. For this analysis, a

hot bleed cycle was analyzed. There are other potential options for providing the necessary energy, such as

using some outer elements as heater elements in parallel with the nozzle cooling.12 Since a hot bleed cycle

was examined, it is also assumed that the RL-60 turbomachinery will not be suitable. Therefore this design

may involve added development costs relative to other designs examined.

VIII. Discussion

Each engine design has its own strengths and weaknesses as can be seen in Table 4. For the small engine

design, the CERMET engine has better thrust to weight and a higher specific impulse. The higher thrust

to weight is due to the smaller, more compact engine design. Also, the small CERMET engine has a much

NASA/TM—2013-216553 13
simpler flow path design than the NERVA derived engine since it does not have tie tubes. However, the

small CERMET engine requires over six times the amount of highly enriched 235U. If the cost or availability

of enriched uranium is a major contributor to the engine, this may preclude the use of a CERMET engine.

When the desired thrust is increased to 112.1 kN (25 klbf), the benefits of CERMET seen in the small

engine start to decrease. This is primarily driven by the high pressure drop across the high thrust CERMET

core and the limited available energy for the turbopump operation. In order to get the engine cycle balance

to close, a hot bleed cycle was adopted which lowered the overall specific impulse of the engine and increased

the complexity of the design. Not only is the performance of the engine worse than the large NERVA engine,

it also requires approximately ten times the amount of enriched uranium. However, a lower pressure drop

and a reduction of enriched uranium might be achieved by redesigning the fuel element and reactor core.

Table 4. Engine Comparison

Description Small

NERVA

Small

CERMET

Large

NERVA

Large

CERMET

Units

Reactor

Active Fuel Length 88.9 60.96 132 86.36 cm

Radial Reflector Thickness 14.7 10 14.7 10 cm

Axial Reflector Thickness – 15.24 – 15.24 cm

Engine Diameter 87.7 55 98.5 68 cm

Element Pattern Type Dense – 2:1 –

Number of Fuel Elements 260 121 564 241

Number of Tie Tube Elements 251 – 241 –

235U mass 27.5 177.3 36.8 376.7 kg

Component Masses

Reactor 1435 1215 2645 2722 kg

Pressure Vessel 110 50 280 150 kg

Nozzle 115 115 240 270 kg

Turbomachinery-Piping 40 40 90 135 kg

Gimbal 30 30 50 50 kg

Engine Total 1730 1450 3305 3327 kg

Engine System

Thrust 33.2 33.8 113.3 111.2 kN

Maximum Fuel Temperature 2860 2860 2860 2860 K

Chamber Inlet Temperature 2756 2682 2794 2681 K

Chamber Pressure 4.14 3.79 6.89 6.89 MPa

Nozzle Expansion Ratio 300 300 300 300

Specific Impulse 899.6 885.4 911.5 824.2 sec

Thrust to Weight 1.96 2.38 3.49 3.41

IX. Conclusions

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. Each engine was designed with the intent of potentially using existing turbomachinery. However, this was not attainable for the large CERMET engine based on the ANL-200 fuel element

design. The other engine results compare closely to prior analysis performed by Fittje and Schnitzler using

the Nuclear Engine System Simulation (NESS) tool.12, 13 Potential issues for each engine design have been

identified and future work will be directed to reduce these risks.

NASA/TM—2013-216553 14
References

1Borowski, S., McCurdy, D., and Packard, T., “7-Launch NTR Space Transportation System for NASA’s Mars Design

Reference Architecture (DRA) 5.0,” AIAA-2009-5309, Aug. 2009.

2Houts, M., Borowski, S., George, J., Kim, T., Emrich, W., Hickman, R., Broadway, J., Gerrish, H., and Adams, R.,

“Affordable Development of a Nuclear Cryogenic Propulsion Stage,” AIAA-2012-4308, 2012.

3Finseth, J., “Overview of Rover Engine Test: Final Report,” NASA-CR-184270, Feb. 1991.

4Stauber, L., “Numerical Propulsion System Simulation (NPSS): An Award Winning Propulsion System Simulation Tool,”

NASA-TM-2002-211333, 2002.

5Clough, J., Integrated Propulsion and Power Modeling for Bimodal Nuclear Thermal Rockets , Ph.D. thesis, University

of Maryland, College Park, 2007.

6Clough, J., Starkey, R., Lewis, M., and Lavelle, T., “Reactor Modeling for Bimodal Nuclear Thermal Rockets,” AIAA2006-4556, 2006.

7Technology, A. S., “RL-10 Specifications,” http://www.spaceandtech.com/spacedata/engines/rl10 specs.shtml.

8Durham, F., “Nuclear Engine Definition Study Volumes I-III,” LA-5044-MS, 1972.

9Schnitzler, B., “Small Reactor Designs Suitable for Direct Nuclear Thermal Propulsion: Interim Report,” INL-EXT-12-

24776, Feb. 2012.

10Fittje, J., “Upgrades to the Ness (Nuclear Engine System Simulation) Code,” AIAA-2007-5620, 2007.

11Schnitzler, B. and Borowski, S., “Small Fast Spectrum Reactor Design Suitable for Direct Nuclear Thermal Propulsion,”

AIAA-2012-3958, 2012.

12Fittje, J. and Schnitzler, B., “Cycle Analysis of a 200 MW Class Cermet Based NTR Engine,” AIAA-2012-3960, 2012.

13Schnitzler, B., Borowski, S., and Fittje, J., “Lower Thrust Engine Options Based on the Small Nuclear Rocket Engine

Design,” AIAA-2011-5846, 2011.

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