Although the optimal choice for a power plant for the Mars sample return mission would be a Radioisotope Thermoelectric Generator (RTG) or similar device, current policies regarding nuclear devices prohibit their use on this mission. To meet the power requirements for all phases of the mission, solar photovoltaic cells and rechargeable batteries are employed. Mars is relatively close to the Sun; the solar flux varies from 500 W/m2 at aphelion to 730 W/m2 at perihelion. The range of the solar flux available at Mars is sufficient to power the spacecraft without resorting to inordinately large solar arrays.

To determine the actual size of the solar array that is required for each component (cruise stages, landing vehicle, and rover) in space and on the surface of Mars, a detailed analysis of the solar flux has been performed. The analysis takes into account the distance of Mars from the sun, its axial tilt with respect to the sun, the length of day and position of the spacecraft on Mars. It can be seen from Fig. 9 that both the northern and southern hemispheres are basically equivalent for the given surface stay, with the equator being the best possible location for the Ares Acquire mission.

The energy available varies greatly over the stay on the surface of the planet. The rates at which the PPPs produce propellant depend directly on the available energy. At times of high energy availability, propellant production is correspondingly high, and then decreases as the solar flux decreases. The propellant production (for all three ISRU scenarios) averaged throughout the entire stay on the surface of Mars is sufficient to produce all the propellant needed.

To keep the solar array as small as possible, the most efficient solar cells available, 22% efficient gallium arsenide (GaAs) cells, are employed. Due to limitations of both mass and volume the largest solar array that can be carried by the Mars landing vehicle (MLV) is 24 m2, with a mass of 66 kg. The SHMP and TERP Mars landing vehicles neither require nor can carry the 24 m2 solar array. They instead carry a 10 m2, and 2 m2 solar array, respectively. Nickel/Zinc batteries with a specific energy of 57 Wh/kg are used for storage of the nighttime power requirement in all options. To decrease the mass of batteries carried by the MLV, aspects of propellant production are limited to daylight periods (~12 hrs). The battery mass per mission is shown in Table 2. The minimum solar array area required for surface operations in all cases is smaller than the actual solar array being taken to Mars. The safety margin present in each design is meant to compensate for dust storms (local, regional, or global), which could prevent the successful production of the necessary fuel for the return to Earth of the samples.

Two cruise stages, one for each transfer orbit, contain the solar array and battery system that provide power while the spacecraft is in space. The power available to the cruise stages is simply a function of the spacecraft's distance to the sun and the orientation of the solar array. To allow for maximum energy collection, the solar array in each case is aimed directly at the sun. The cruise stages operate in an environment where the minimum solar flux available is never less than 500 W/m2.

In all, the spacecraft carries a total of three solar arrays and battery storage systems, as well as the power system for the rover. It is evident that using solar power for such an extended time on the surface of Mars poses several risks, including winds and dust storms. However, the solar array carried by the Ares Acquire mission in all four scenarios will allow for successful mission completion.

The design of the vehicle is divided into the Mars Landing Vehicle (MLV), Mars Ascent Vehicle (MAV), Sample Return Capsule (SRC), and transfer cruise stages. The SRC, MAV and cruise stages are common to all scenarios.

Mars Landing Vehicles

The MLVs for all options are similar in several respects and share some common components. They all require a sample acquisition system and a solar array. All MLVs are constructed of aluminum I-beams. For the ISRU options, the MLV must also have a propellant production plant, and in the case of the SHMP mission, seed hydrogen tanks. All components of the MLV are left behind on the surface of Mars when the MAV ascends. Mass summaries are provided at the end of the vehicle design section.

Seed Hydrogen-Methane Produced (SHMP)

The SHMP option requires the transport to Mars of the seed hydrogen for the Sabatier reactor. The transport of the liquid hydrogen presents not only the inescapable problems of the cryogenic systems, but also size limitations, as the tanks required for the 65 kg of H2 are voluminous, and the spacecraft must fit within the requisite Delta II launch vehicle. The SHMP MLV is illustrated in Fig. 10. Attached to the base of the MLV are four titanium hydrogen tanks (heavily insulated with SOFI and MLI) that keep the 2.5 MPa hydrogen at 20 K. The hydrogen is fed to the Sabatier reactor where it is processed into methane and oxygen propellant. The SHMP MLV requires a different planform shape from the other three MLVs due to the voluminous hydrogen tankage. While the other MLVs have a triangular shape with a three-legged landing gear, the SHMP MLV requires a square planform with a four-legged landing gear so that the four hydrogen tanks can be carried on the MLV structure and still fit within the launch vehicle. The power for the SHMP MLV is provided by 10 m2 of solar array deployed in four folding wing-like arrays.

Water-Sabatier

The WATS MLV (Fig. 11) is slightly different from the SHMP MLV. The square frame is replaced by a triangular planform frame and the voluminous hydrogen tanks are replaced by the WAVAR unit which connects to the Sabatier reactor as shown in Fig. 6. The Sabatier reactor is identical to the SHMP Sabatier reactor shown in Figs. 1 and 2. The solar array for the WATS option is also much larger than that of the SHMP scenario, due to the increased power requirements of the WAVAR unit. The solar array (Fig. 12) is deployed in a petal-like fashion, and is stayed by cables.

Methane Transported-Oxygen Produced (MTOP)

The MTOP scenario has an MLV which is configured identically to the WATS MLV (Fig. 11). The MTOP option is unique among the ISRU missions in that it is importing methane fuel from Earth. The methane remains in the MAV tank until needed for ascent, thereby eliminating the external tankage needed for the SHMP MLV. The power requirement for MTOP is comparable to that of WATS. Therefore, the same solar array design is used.

Terrestrial Propellant (TERP)

The TERP MLV (Fig. 13), while structurally similar to the others, does not have a propellant production plant. This greatly reduces the size of the solar array required. The framework of the TERP MLV, however, must be structurally heavier to account for the increased stresses experienced due to the much larger Earth launch mass and Mars landing mass. The propellants for the TERP mission are MMH and NTO, as stated earlier.

MLV Aerobrakes

The MLV aerobrake is designed according to the amount of heat flux expected during atmospheric entry. A maximum heat flux of 63 W/cm2 is expected at the stagnation point, at a Mars entry velocity of 7 km/sec and flight path angle of 11° for the ISRU options. The peak surface temperature at the stagnation point is ~1500 K in each case.

The aerobrake for the ISRU MLVs (Fig. 14) is derived from the Mars Pathfinder aerobrake design. It is a 70° half cone with a 2.8 m base diameter and a 0.7 m nose radius. The primary thermal protection material is FRCI-12, with carbon-carbon composite at the stagnation region. The structural substrate is 3 cm thick aluminum honeycomb sandwiched between graphite-epoxy sheets.

In order to maintain the heat load and the deceleration levels of the more massive TERP vehicle comparable to those of the ISRU vehicles, the TERP aerobrake must have an area approximately twice that of the other options, i.e., the TERP aerobrake diameter needs to be 3.95 m. However, this figure exceeds the internal diameter of even the largest faring of an Atlas II AS launch vehicle. Accordingly, the TERP aerobrake is sized at 3.65 m diameter so that it will fit that launch vehicle. This necessitates a somewhat steeper atmospheric entry and results in a slightly higher heating rate and deceleration.

Due to the height of the MLV, a backshield is necessary to protect it from flow impingement, as shown in Fig. 15. The backshield is constructed of a 1 cm thick Nomex honeycomb, sandwiched by 1 mm thick high temperature graphite epoxy face sheets. The framework of the backshield consists of aluminum 2024 T-beams, in the form of 3 struts and 6 stringers. The TERP MLV’s backshell is larger and requires additional struts and stringers.