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Part of American Mars Expeditions
Credit: © Mark Wade
American manned Mars expedition. Study 1991. The solar electric propulsion (SEP) Mars transfer concept was the only non-nuclear advanced propulsion option in the STCAEM study.

Status: Study 1991. Specific impulse: 5,500 s.

It offered advantages of the lowest IMLEO of the four reference vehicles; a reusable, extremely high Isp (5,000 sec) system; a fully propulsive capture at Mars and Earth which avoided the need for high energy aerobraking; good mission flexibility (relative insensitivity to mission opportunity, capture orbit astrodynamics, or changes in payload mass); and low resupply mass (the argon propellant required amounted to roughly a third of total vehicle mass). Disadvantages included uncertainty of the cost of production of acres of solar arrays, and the need to deploy and control a relatively fragile vehicle, which was bigger than six football fields, in space.

STCAEM (Space Transfer Concepts and Analyses for Exploration Missions) was a major NASA funded study produced by Boeing in 1991. It provided an exhaustive trade analysis of mission profiles and trajectories for manned Mars missions using four different propulsion technologies (cryogenic chemical with aerobraking, nuclear thermal, nuclear electric, and solar electric). Within each study alternate mission profiles using split/sprint missions, flyby rendezvous, and additional aerobraking were examined. Only the baseline for the solar electric mission is presented here.

Nominal Mission Outline

Vehicle Systems

Primary vehicle systems were: power plant; main propulsion; vehicle bus; and crew systems.

The power plant consisted primarily of a field of solar arrays kept normal to the sun line at all times. The solar array area required to produce 10 MWe of power was 35,000 m2 and was maintained sufficiently rigid and in position by a deployable area truss (spaceframe) one bay deep. Details of deployment of the lightweight solar cell blankets across the structure were not worked out in the study.

The propulsion system included engine assembly, propellant storage subsystem, and plumbing components, split into two identical modules located at distal ends of the vehicle bus. Each engine assembly had five individual ion thrusters (the total of ten included two spares) in a 5 x 8 rectangular array. Each thruster was 1 m wide by 5 m long; beam neutralizers were located between the thrusters. The argon propellant was stored cryogenically in insulated, spherical tanks, mounted on the forward sides of the engine assemblies via structural and fluid quick-disconnects. Including tanks, the propellant storage system massed about 35% overall vehicle IMLEO. This relatively low propellant mass was a strong resupply advantage.

Thrust loads were extremely low for the electric propulsion (EP) system. Probable maximum loading was from impulses like attitude control system (ACS) firings, berthing operations, and construction and maintenance activity. The primary vehicle bus structure had two components: the area mass covered by the solar array field, and truss outriggers extending sufficiently far beyond the edge of the solar array that the ion engine plumes did not impinge on, and therefore erode, the power system. The crew systems were attached to the underbelly of the area truss (in the center for mass balance). Two communications satellites were also attached to the truss near the crew systems, to be deployed in Mars orbit for maintaining communication with Earth. Also mounted to the truss near the habitation system were thermal radiators for the power conditioning equipment.

The crew systems consisted of a long duration transit habitat and one or more MEVs (the reference design had one MEV). All habitable volumes were contiguous throughout each mission. Electric propulsion had the least sensitivity to increased payload mass, so an important option was provision for multiple MEVs. A multiple docking adapter would allow several MEVs to be used without altering the vehicle configuration (additional propellant tanks would be required).

SEP Mass Breakdown


Subtotal payload: 117.5 metric tons


Subtotal Propulsion: 28.0 metric tons

Solar array blanket

Subtotal Solar array blanket: 27.7 metric tons


Subtotal structure: 11.3 metric tons


Subtotal Utilities: 14.9 metric tons

Propellant system

Subtotal propellant system: 193.9 metric tons

Subtotal: 393.3 metric tons

  • 15% Growth (propulsion and other): 6.4 metric tons Total Initial Mass in Low Earth Orbit: 403.6 metric tons

    Mass to resupply after first use: 335.2 metric tons

    STCAEM SEP Mission Summary:

    • Summary: Major NASA funded study produced by Boeing in 1991; focus on in-space propulsion
    • Propulsion: Solar Electric
    • Braking at Mars: aerodynamic
    • Mission Type: opposition
    • Lunar swingby: yes
    • Split or All-Up: all up
    • ISRU: no ISRU
    • Launch Year: 2016
    • Crew: 4
    • Mars Surface payload-metric tons: 115
    • Outbound time-days: 220
    • Mars Stay Time-days: 20
    • Return Time-days: 310
    • Total Mission Time-days: 550
    • Total Payload Required in Low Earth Orbit-metric tons: 410
    • Total Propellant Required-metric tons: 180
    • Propellant Fraction: 0.43
    • Mass per crew-metric tons: 102
    • Launch Vehicle Payload to LEO-metric tons: 140
    • Number of Launches Required to Assemble Payload in Low Earth Orbit: 8
    • Launch Vehicle: Shuttle Z

    Family: Mars Expeditions. Country: USA. Spacecraft: STCAEM MEV. Propellants: Electric/Xenon. Agency: NASA, Boeing. Bibliography: 1985, 4421.
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    STCAEM ComparisonSTCAEM Comparison
    Credit: © Mark Wade

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