Propulsion Systems for Manned Mars Missions

Propulsion Systems for Manned Mars Missions

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With the advent of human space travel becoming more and more viable with a continued effort and drive from the private sector in the recent 10 years the dream of space travel for regular people is slowly but surely becoming realized. Private space companies such as Virgin Galactic, Space X, Ad Astra and many more around the world are racing to the ‘final frontier’ trying to produce a viable and commercially successful, profitable business. With many opportunities both financial and ideological to be had, the push for space exploration is reaching levels that were present during the US – USSR space race of the 1950’s, 60’s and 70’s. The interest this time not coming from a military drive for dominance, but rather from the newer, younger generation of CEO’s, aspiring scientists and engineers sold on the romanticized ideals of space exploration. But the question of why remains an existential one for many earth bound mammals. Many who support the idea of space travel as not only a means of greater understanding of our universe but of a greater quest to understand our place in it. A destination frequently purported by supporters of the scientific philosophy and one of space exploration is, in terms of average distance, is our second closest neighbor, Mars. With the closest orbital distance to earth at approximately INSERT FIGURE it would take approximately LOTS M8 days for a return trip for travel time alone with conventional chemical propulsion, assuming the mission starts in low earth orbit (LEO) then into an orbit around mars then return to LEO. Nuclear propulsion systems are becoming an increasingly popular and practical technology for reducing that travel time by up to 200 days, with some estimates claiming with a large enough power supply transit times of VALUE weeks one way can be achieved CITATION. This leaves the question of what are these propulsion systems and where do they get their power requirements.
This report will focus on the most popular ideas and projects being tested today which include some already well-established technologies as well some that have only just entered the laboratory.
There will also be some focus on the power plants that will provide the energy necessary to cross the interspatial void between the planets.

Nuclear Thermal Rocket (NTR)

The NTR presents a great advantage over traditional chemical propulsion systems. NTR uses fission reactions to generate thermal power, which in turn heats a liquid propellant to produce thrust.

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The two main types of NTR that will be discussed are the; solid and and gas core NTR designs, with the main focus on solid core being the only ground tested NTR.

Comparison of propulsion types for mars mission. [1]

Open cycle Gas-Core Nuclear Rockets (GCNR) is the most efficient NTR concept with potential to Isp range of 3000-5000 seconds. [2]
The open cycle GCNR design has never been fully tested due to the complexity of the fluid dynamics involved. The core is essentially comprised of a sphere of nuclear plasma suspended by liquid hydrogen (LH2) in a reaction chamber with direct contact between the fission source and the propellant. The design has the advantage of being able to theoretically increase the exhaust temperature of the propellant to 10’s of thousands of Kelvins (turning the propellant in to a plasma), hence dramatically increasing the Isp, but this then raises an issue of nozzle cooling. Most designs state that the use of a magnetic nozzle would be the solution to this problem. The complexity of GCNR arises due to the issue of containment of the fissile material that generates the thermal power. The design is unlikely to be developed due to the intense radioactivity of the exhaust gas.

Schematic of the Open-Cycle GCNR design. [2]

Another GCNC concept is the ‘nuclear light bulb’ design, which does not allow contact of the radioactive plasma core with the propellant. The fissile material (typically an isotope of uranium) is encased in a transparent fused silica containment vessel, which transfers energy into a flow of LH2, which is then expanded through a de Laval nozzle to produce thrust. The Isp of the GCNR is typically around 2000 seconds with potential to reach much higher Isp with containment, maximum working temperature and nozzle limitations (cooling etc.) being the limiting factors. This concept has yet to be developed experimentally. [3]

Schematic of the Nuclear Light bulb design. [3]

Solid core NTR designs are the most common and researched of NTR. They employ a traditional solid fusion reactor design and the process of operation is well understood and has a high controllability and simplicity when compared with its contemporaries.

The ROVER/NERVA program (1955 -1972) successfully demonstrated 20 compact solid-core reactor designs for space application. These tests provided proof of concept and proof of design that NTR technology is a feasible option for interplanetary travel. [4]


Solid core NTR’s can attain specific impulses (Isp) upwards of 900s, more than twice the Isp of chemical rockets and achieving high levels of thrust (10’s-100’s of Klbf). In the Mars Design Reference Architecture (DRA) 5.0 study conducted by NASA in 2009 concluded that solid core NTP was the optimal choice to carry crews and cargo in Earth to Mars transit. The DRA 5.0 based their calculations on the 25 Klbf engine design from the Rover-NERVA project, a configuration of 3 “Pewee” engines (diagram can be found in appendix), the highest performing engine from the Rover Program, to be used in parallel. The Pewee-1 engine design was successfully test at a peak thermal power output operating at ~507MWt for 40 minutes. [6]

The design itself is relatively simple and has the ability to; self-start, stop, restart, has demonstrated sustained engine operation and can operate at a wide range of thrust levels.
The fission of Uranium-235 within the composite fuel element generates thermal energy, which heats a liquid hydrogen propellant. The LH2 first moves from the storage tank through a duct wound about the exit nozzle for cooling (additional propellant heating) where it is then passed along the outside of the reactor to be heated further then through the cooling ducts in the super heated reactor core and subsequently expelled through a supersonic nozzle. Before being fed through the core some LH2 is syphoned off and channeled through turbines to drive turbo-pumps, which initially compresses the liquid fuel (shown in figure below). Hydrogen that has passed though the turbines is then routed back in to coolant channels surrounding the core.
Schematic of solid-core NTR design. [4]

Control of the engine output is through the matching of the LH2 flow to core cooling requirements. Reactor core output is controlled through the use of external control rods surrounding the outside of the fuel element vessel. These fuel rods are made from a combination of materials, typically half of the control absorbs neutrons (B4C) and the other half reflects them (BeO). The control rods are subsequently rotated to achieve a desired thermal output. Control face with neutron absorbing material facing the core for lower temperature output (less neutrons are reflected back into the core) and control face with neutron reflective material orientated towards the core for higher thermal output. [7]

Engine Performance
Taking parameters from the Pewee 1 tests shown in table 1 calculations can be made for; specific impulse (Isp), throat and exit diameters exhaust exit velocity, thrust produced and the mass of fuel needed to reach the mission delta v.

Engine Performance characteristics outlined in Pewee 1 Tests. [6]
Tcore ~2556 K
Pcore ~6895 KPa
ε (exit to throat area ratio) ~300:1
LH2 flow rate ~18.6 Kg/s
Thrust to weight ratio ~3.5
Critical mass U-235 in core 36.4 kg
Nozzle chamber temperature ~1837 K
Nozzle chamber Pressure 4275 kPa
Mission delta V budget* 3.83 km/s
Total mass of Spacecraft* 139 t
*Taken from DRA 5.0 study. [7]

Assuming isentropic processes exhaust velocities (ue) can easily be calculated through the following equation. [15][21]
u_e=√(2γ(R_0/M)/(γ-1) T_02 [1-(P_e/P_02 )^((γ-1)/γ) ] )
Where R0 is the universal gas constant, T02 is the nozzle chamber temperature, p02 is the nozzle chamber pressure, γ is the ratio of specific heats, Pe is the nozzle exit pressure and M is the molecular weight of the propellant.
Given an exit to throat area ratio (A/A*) of 300 a nozzle chamber to exit pressure ratio (Pe/P02) can be evaluated through isentropic relations.

P_e/P_02 =5.358×〖10〗^(-5)

Using liquid hydrogen (LH2) as propellant yields (table of LH2 properties in appendix).

u_e=7056.12 m/s

Force of thrust (F_T) can then be calculated through the basic thrust equation.

F_T=m ̇u_e+(P_e-P_a ) A_e
F_T=m ̇c^* C_F
c^*=√(((R_0/M) T_02)/γ ((γ+1)/2)^((γ+1)/(γ-1)) )
C_F=A_e/A^* (P_e-P_a)/P_02 +√((2γ^2)/(γ-1) (2/(γ-1))^((γ+1)/(γ-1)) [1-(P_e/P_02 )^((γ-1)/γ) ] )
Subbing in values gives
F_T=132445 N

m ̇ Is the mass flow rate of propellant, P_a is the ambient pressure (assumed to be zero), Ae is the exit area of the nozzle, c* is the characteristic velocity and CF is the nozzle coefficient. [15]

The specific impulse can now be evaluated.

I_sp=(c^* C_F)/g_0 =725.862 s

g_0 Is the acceleration due to gravity at the Earth’s surface (~9.81 m/s2). The mass of LH2 propellant needed to achieve the mission delta v budget is found using a derivation of Tsiolkovsky’s rocket equation. [19]

m_propellant=m_initial (1-e^(-Δv/(I_sp g_0 )) )

Inserting values from table ## and result obtained from equation # gives.

m_propellant=57.825 tonnes of LH_2

Hence the burn time would be.

t=m_propellant/m ̇ =m_initial/m ̇ (1-e^(-Δv/(I_sp g_0 )) )

t=3108.88 s=51.815 minutes

This burn time is reasonable as the pewee design was tested at full power for over 40 minutes. [6]

The mass of propellant used is for a one-way acceleration of the craft. It states in the Rover test files that the Isp of the Pewee engine was not optimised as it was primarily used as a test bed for reactor core elements. However specific impulses of this magnitude still hold a huge advantage over bi-propellant chemical propulsion when considering inter planetary travel. [6]
If the Isp were improved to the purported values of 900s and above this would reduce the weight of propellant needed by approx. 10 tonnes, which in turn reduce the burn time and overall mass of the craft, hence the mass of propellant required.
Improvements in specific impulse can be achieved though increased core temperatures, which a fission reaction can produce easily, but the limitation lay with the material properties of the core.

The engine examined in the DRA 5.0 utilizes U-235 fuel in a graphite matrix fuel element (UC2). U-235 has an energy density of 1.68e+13 J/kg, while Liquid hydrogen (1.20e+8 J/kg) is used for propellant. [8]

The thermal core in the Pewee design is comprised of 402 hexagonal fuel elements with internal cooling channels for the LH2 to flow through. A typical core design developed in the Rover-NERVA programs is shown in figure ##.


The internal flow channels for the heat transfer to the LH2 is moderated by a coating of niobium carbide or zirconium carbide to prevent corrosion of the uranium-carbon matrix by the super heated LH2. Clusters of six fuel elements were place around a central tie rod to place the fuel elements under axial compression to minimize damage caused by flow vibrations. [10]

A fraction of the LH2 flow is syphoned in to coolant channels through the supporting structure to reduce excess heat damage.
Several fuel element designs were tested by both the US and Russian NTP Programs with several inquiries in to different fuel element compositions and a particular focus on CERMET (ceramic-metal composite) fuel elements.
Figure ## below demonstrates the effect that core operating temperature has on Isp and the difference in operating temperatures of different fuel matrixes.


Soviet efforts of developing fuel elements produced unique geometries, resembling a drill piece, to maximise heat transfer. The ternary carbide fuel in the ‘twisted ribbon’ configuration provided the highest performance characteristics reportedly operating at maximum core temperatures of 3200 K.[7][11]

‘Twisted ribbon’ design. [12]

There is a multitude of support for nuclear thermal rocket designs and application and with available technology seems to be the best option compared to the competing propulsion systems such as chemical combustion or electric propulsion for long term space missions.
The prospect of sending humans to mars is of great interest to many in the scientific community and members of the general public and there are no international space treaties that explicitly prevent the use of nuclear reactor technology in space. One mission parameter designated in NASA’s Prometheus project that could be taken into consideration was that nuclear reactors could not be in operation until the time for the spacecraft’s orbit to decay is longer than the decay time of the fissile material. Generally speaking, at the point where the craft will remain in orbit indefinitely. [13]
Mission parameters and safety considerations will help to alleviate public and administerial concerns. NASA literature dictates that several design factors should be considered when employing the use of fission reactor systems in space.

“The ability [of the reactor] to operate reliably without continual actions from ground control, the ability to keep the reactor in a subcritical state prior to startup and under various accident scenarios, the ability to remove operational and decay heat during specified normal and off-normal operating conditions, and the ability to reliably perform all necessary control and safety functions.”[14]

At a distance in excess of 10ft (3.048m) is proportional to 1/r^2 , where r is the distance from the reactor core. The data in the table states at 0 degree orientation from reactor core (where crew module would be located), at a distance of 6ft, 1.8*10^7 rads/hour is released. [17] Assuming that the radiation from 6-10 ft. is also proportional to distance as stated above,
Rads per hour=k 1/r^2 .
Where for this purpose k will be called the coefficient of radiation environment and will have the units, (rads.ft2)/hour. It is a relative measure of the strength of the radiation source. K can be calculated from the tabulated data.
k=1.8×〖10〗^7×6^2=64.8×〖10〗^7 rads×(ft^2)/(hour)
Maximum allowable occupational radiation absorption standards give 5rems/year or 0.0005704 rems/hour. [16]
Assuming rads ≈ rems (approximate maximum possible dose).
Without shielding the crew would have to be situated in front of reactor 1.066 ×〖10〗^6 ft (approx. 325 Km). The shielding required reducing radiation to the maximum occupational exposure level at ~7ft from reactor core would be approx. 40cm of lead shielding. [18]

With further advancement of nuclear thermal technology, the reactor for the NTR can be augmented for ‘Bi-modal’ operation, producing both the propulsive power for the craft but also can be run to produce electrical power, generated through a Brayton cycle configuration. This electrical power generated is estimated at .1-1 MWe, which can be used to run the onboard systems such as communication, refrigeration and life support systems as well as the possibility of an electric thruster to increase the delta v capabilities of the craft as well as performing orbital maneuvers. [7]


Thrust augmentation through injecting liquid Oxygen (LOX) into the divergent, supersonic section of the nozzle can deliver additional propulsive force by adding chemical energy and mass to the exhaust gas. The addition thrust achieved would result in shorted burn times thus reducing the LH2 mass required. Due to the higher density of LOX smaller containment vessels can be used further reducing the overall mass of the system. The table ## shows the variation of Isp versus LOX/H2 mixture in the exhaust. [20]



LH2 Properties [8]
Ratio of specific heats, γ ~1.4
Gas constant, R 4.12 kJ/kg
Specific heat at constant pressure, cp 14.32 kJ/kg

Cross section of Pewee engine design [6]


Nuclear Engine for Rocket Vehicle Application (NERVA)


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[2]. Ragsdale, R. (1991). Open cycle gas core nuclear rockets. [PDF] NASA Technical Reports Server.

[3]. Latham, T. (1991). NUCLEAR LIGHT BULB. [PDF] NASA Technical Reports Server.


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[6]. Finseth, J. (2014). Rover nuclear rocket engine program: Overview of rover engine tests. [PDF] NASA Techical Reports Server.

[7]. Borowski, S., McCurdy, D. and Packard, T. (2012). Nuclear Thermal Propulsion (NTP): A Proven Growth Technology for Human NEO / Mars Exploration Missions. [PDF] NASA Technical Report Server.

[8]. University of Washington: School of Oceanography, (2005). Energy in Natural Processes and Human Consumption. [online] Available at: [Accessed 6 Apr. 2014].


[10]. Bhattacharyya, S. K. An Assessment of Fuels for Nuclear Thermal Propulsion. Rep. no. ANL/TD/TM01-22. Argonne, IL: Argonne National Laboratory, 2001.

[11]. Benensky, K. (2013). Summary of Historical Solid Core Nuclear Thermal Propulsion Fuels. [PDF] Penn State Scholar Sphere.

[12]. Borowski, S., Corban, R., McGuire, M. and Beke, E. (1993). Nuclear Thennal RocketNehicle Design Options for Future NASA Missions to the Moon and Mars. In: Space Programs and Technologies Conference and Exhibit.

[13]. Joseph J. MacAvoy, Nuclear Space and the Earth Environment: The Benefits, Dangers, and Legality of Nuclear Power and Propulsion in Outer Space, 29 Wm. & Mary Envtl. L. & Pol'y Rev. 191 (2004),

[14]. NASA Space Science, Space Fission Reactor Power Systems: Their Use and Safety (Feb. 2003), at 2, available at

[15]. Stengel, R. (2014). Launch Vehicle Design: Propulsion. 1st ed. [PDF] Princeton University. Available at: [Accessed 8 May. 2014].

[16]. ANS / Public Information / Resources / Radiation Dose Chart. 2014. ANS / Public Information / Resources / Radiation Dose Chart. [ONLINE] Available at: [Accessed 24 May 2014].

[17]. NERVA. Performance/Design and Qualification Requirements. (1970). 1st ed. [ebook] Aerojet Nuclear Systems Company, p.74. Available at: [Accessed 17 Apr. 2014]

[18]. A Compass DeRose Guide by Steven J. DeRose. 2014. A Compass DeRose Guide by Steven J. DeRose. [ONLINE] Available at: [Accessed 25 May 2014].

[19]. Braeunig, R. (2012). Basics of Space Flight: Rocket Propulsion. [online] Available at: [Accessed 3 May. 2014].

[20]. Borowski, S. and Dudzinski, L. (2003). 2001: A Space Odyssey” Revisited—The Feasibility of 24 Hour Commuter Flights to the Moon Using NTR Propulsion with LUNOX Afterburners. Cleveland, OH, United States: NASA Glenn Research Center.

[21]. Hill, P. and Peterson, C., Mechanics and Thermodynamics of Propulsion, 2ndEdition, Addison Wesley, Reading, Massachusetts, 1992.

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