1. In-situ Propellant Production and ERV Propulsion System
Critical Design Review
Adam Butt
2/27/01

2. Overview In-situ Propellant Production ERV Propulsion system
Production Methods
ERV Propulsion system
Single Stage To Orbit (SSTO)
Tank sizing and positioning
Engine selection
Power and Reliability

3. In-situ Propellant Selection
Methanol has been chosen to the sole fuel to be produced in-situ
Reasons for the selection:
Does not require cryogenic storage
Higher density leads to smaller tanks
Greater yield per tonne of H2 than Methane
One tonne of H2 yields 5.3 tonnes of CH3OH and 5.7 tonnes of H2O, while only 2 tonnes of CH4 (and 4.5 tonnes of H2O)
Ease of using as rover propellant
The oxidizer to be produced in-situ will be O2

4. Method of Production (Upper portion of picture from JPL – Advanced Propulsion Concepts Website)
Zirconia cell process
CO2  CO + O2
Liquid Hydrogen feedstock added to the separated CO,
CO+3H2CH3OH+2H2O
Methanol stored, and water is electrolysized to cycle back hydrogen, and store oxygen
System Volume and mass
under 50m3 and 1 tonne H2 O2
H2 Storage
H2O
Electrolysis
H2O
Catalyst Bed CO
Methanol Storage
CH3OH

5. ERV Propulsion - SSTO Advantages of a Single Stage To Orbit mission:
Lack of staging decreases complexity and increases overall mission success rate
Allows for the use of one liquid rocket engine to provide the entire DV required to return the astronauts to Earth

6. Tank Sizing and Positioning
(8) Methanol Tanks
Tanks are sized according to the constrained volumes, and geometries suitable for pressurization
All propellant necessary for launch off of Mars, and departure from orbit are contained in the same tanks, and are used by the same engine
Propellant mass evenly distributed about vehicle
Thrust vectoring possible
(4) LOX Tanks
F-1A Saturn V Engine

7. Tank Sizing and Positioning
Assuming a CTV mass of 5 tonnes and an all up ERV weight around 50 tonnes,
Total mass of LOX is around 100tonnes
Total mass of methanol is around 65 tonnes

8. Engine Selection F-1A Saturn V Liquid Rocket Engine
Chosen for the large thrust requirements of around 9000kN
Existing technology that has been involved in many successful launches
Picture from

9. Power Requirements
Three conflicting power equations (or correlations) found:
Based on: Pe = [0.145(MT)1.238] kW, around 220kW would be necessary to produce all the propellant in 15 months
Based on NASA’s Sample Return Mission (which incorporates in-situ prop. Prod.), around 25% less power would be necessary, or 160kW
Based on an in-depth study from the University of Washington, in which they use more advanced processes to produce 200tonnes of propellant with 30kW, putting us around 22kW
No final decision made as of yet on which to use

10. Failure Probabilities
Engine failure probability very low, due to a proven and tried technology
Large amounts of simulated testing going on right now for in-situ production, therefore failure probability will decrease a mission time draws near
More in-depth analysis into newly designed system to come