Ares Acquire launches during the March 2001 launch opportunity on top of a Delta II 7925 launch vehicle into low Earth orbit (LEO), as shown in Fig. 10. After achieving LEO, the spin-stabilized PAM-D third stage of the Delta II 7925 provides the energy required to inject the Ares Acquire spacecraft on a six-month Type I conjunction-class trajectory to Mars.

Following separation from the PAM-D, spacecraft roll thrusters are fired to reduce spacecraft spin to the cruise rate of 2 RPM. Then the navigation system initiates a Sun/Earth acquisition sequence using sensors located on the MTCS. A communications link is established with a visible Deep Space Network complex and stored launch telemetry and real-time spacecraft performance data are downlinked for ground analysis. The cruise command sequence is then uplinked to initiate the cruise phase.

No science is conducted during the cruise phase. However, all surface operation equipment is subjected to health checks twice during cruise phase: once just after launch and once prior to entry.

Mars Entry and Landing

The spacecraft assumes the Mars entry attitude 24 hours before initial atmospheric interface (Fig. 21). An attitude update is performed and ground verification of proper attitude is obtained one hour before atmospheric entry. The MTCS jettisons 30 minutes prior to atmospheric entry for the descent and landing sequence. The aeroshell, consisting of a heat shield and backshell encapsulating the lander, enters the atmosphere at 11° with a velocity of 7 km/s and employs aerobraking to reduce the vehicle speed to 455 m/s prior to parachute deployment. The parachute is deployed at an altitude of 7 km, as processed by the navigation system, to decrease the descent velocity to approximately 70 m/s.

The aerobrake separates with the deployment of the parachute. The rockets on the Mars Ascent Vehicle are then fired at an altitude of 400 m, after a release from the backshell/parachute, for a soft landing (Fig. 22).

Surface Operations

Once the craft has arrived on the Martian surface, the solar arrays are deployed and communications with Earth are reestablished with the lander high-gain antenna. The landing telemetry and lander status are downlinked and command sequences for surface operations updated and initiated.

Surface operations begin on the first sol, immediately after all surface operation command sequences are uplinked from ground control. An atmospheric sample is taken before the propellant production is started to avoid contamination of the sample with vented gases. The device for remote sample acquisition (rover) is deployed to pick up samples of the Martian surface outside the exhaust plume contamination area. Sample acquisition is controlled from Earth. In addition, a core drilling device attached to a Remote Manipulator Arm (RMA) takes a sample of the Martian crust at the landing site. The basic Pathfinder meteorological package is deployed and processes data throughout the surface stay.

The sample acquisition portion of the surface operations is completed within seven sols, and then the rover begins an extended autonomous exploration mission. Once all sample cylinders are stowed in the sample containment canister and the containment canister is stowed and locked in the SRC, the propellant production begins. The propellant production plant must produce enough propellant during the rest of the surface stay on Mars to give the MAV a DV of 6110 m/s (6300 m/s for the TERP MAV) for the return trip to Earth. Propellant production is completed in 480 sols.

Return to Earth

After an approximately 570 sol surface stay, the MAV launches from the MLV using the indigenously manufactured propellant and begins the return to Earth. The SRC and Earth Transfer Cruise Stage (ETCS) combination separates from the rest of the MAV after the MAV reaches burnout. The ETCS provides any trajectory correction maneuvers needed for the transfer to Earth.

The spacecraft assumes entry attitude 24 hours before Earth atmospheric entry. The ETCS performs an attitude update with ground verification and releases the SRC 30 minutes prior to atmospheric entry. The SRC enters the Earth atmosphere at an entry angle of 6° with a velocity of 11.4 km/s. Aerobraking reduces the SRC speed to approximately 50 m/s and a parachute is deployed at 12 km. The parachute reduces the speed of the SRC to 11 m/s and the SRC is recovered via an aerosnatch maneuver at 7 km.

The other objective of this cost analysis is to determine the relative costs of the four missions presented in this study. As shown in Table 1, the three ISRU missions are, by a good margin, more expensive to develop than the TERP mission. Among the three ISRU missions, WATS is the most expensive, followed by MTOP and finally SHMP.

However, these results should be interpreted carefully for several reasons. First, the errors involved in predicting specific costs are almost certainly greater than relative differences between the mission costs themselves. Second, imposing a hard selection criterion, such as a $200 million cost cap, on numbers as soft as these may create a false sense of decisiveness. Third, once the cost of the launcher is added in, the picture becomes very different. Because the TERP mission has such a high Earth launch mass, an Atlas IIAS launch vehicle is required. This adds $70 million to the bottom line cost and effectively negates any gains made by the scaled down R&D budget.

Finally, ISRU technology will eventually have to be used if a truly ambitious effort is ever made to explore Mars. Because of this, an ISRU technology demonstrator mission such as this one can be considered to be a significant investment in later, larger missions. This brings up the possibility that some of the research required by this mission could be paid for by other, more ambitious projects. Though this may seem more like accounting than engineering, it is reasonable that later missions should help fund the research that would make their existence possible.

The Seed Hydrogen-Methane Produced (SHMP) option presents problems with transporting liquid hydrogen from Earth to Mars and storing it on-planet until propellant production is complete, yet has the advantage of a more mature propellant production technology as compared to Methane Transported-Oxygen Produced (MTOP) or Water-Sabatier (WATS) options. The MTOP scenario has the advantages of avoiding the thermal difficulties associated with the long-term storage of the hydrogen feedstock. However, the zirconia electrolyzers present thermal problems of their own as the electrolyzers’ operating temperature is >1000K. WATS, while incorporating a proven technology (Sabatier reactors) with a new concept (the Water Vapor Adsorption Reactor), is advantageous in that it is a completely ISRU mission with no propellant or feedstock imported from Earth. While the Terrestrial Propellant (TERP) mission is designed to fulfill the mission requirements (except for the ISRU and launch vehicle requirements), it is much larger than the other missions, and is therefore useful only for baseline comparison.

As Fig. 24 shows, among the ISRU options, the WATS mission shows a slight advantage in mass comparison. The SHMP option is advantageous as far as power considerations, with both MTOP and WATS having roughly equivalent power needs (Fig. 25).

A cost estimate for each of these missions has been performed. The requirements as set by NASA for a total mission cost (not including launch vehicle) of $200 million appears to be a reasonable goal for this type of mission. Two of the ISRU missions (SHMP and MTOP) came in under the limit, while WATS was slightly over due to higher development costs. A graphical representation of the cost comparison is shown in Fig. 26.

The WATS scenario, while costing more, presents a completely ISRU option that develops technology which will be important for future manned and unmanned missions, while fulfilling the mission requirements of a Mars sample return mission. It is recommended that future research be concentrated on the WATS scenario and that it be developed to its fullest potential.

Despite the challenges of developing ISRU technology, its inherent advantages are obvious. In the future, ISRU technology will lead mankind to the next phase of space exploration: manned exploration of the solar system. This exposes the greatest weakness of the Terrestrial Propellant option and the strength behind ISRU technology: progress toward the future.

We would like to thank the following people for their support and guidance: Robert Zubrin at Lockheed Martin; Diane Linne at NASA Lewis; Dana Andrews, Eric Wetzel, and Tim Vinopal at Boeing; Kumar Ramohalli at University of Arizona; David Kaplan and John Connolly at NASA JSC; Kim Aaron, Robert Frisbee, and Jim Clawson at JPL; Larry Berg and Chuck Solomon at McDonnell Douglas Delta Project; Rodney Brooks and Helen Greiner at IS Robotics; Jim Tillman at University of Washington Dept. of Atmospheric Sciences; and Eckart Schmidt and Joe Cassady at Olin Aerospace.

We would also like to thank the students of previous AA 420/421 Space Systems Design classes at the University of Washington, whose work provided much inspiration. If we have seen farther, it is because we have stood on their shoulders.