Mars Ascent Vehicle

The Mars Ascent Vehicle (MAV), shown in Fig. 16, serves three purposes. First, the MAV engines provide the DV for the landing retro-burn. Second, the MAV serves as the storage tanks for the return propellant. Finally, the MAV serves as a single-stage booster that imparts the SRC and Earth Transfer Cruise Stage (ETCS) with the necessary DV for Earth transfer. There are two versions of the MAV: one for ISRU propellants (methane/oxygen), and one for terrestrial propellants (MMH/NTO). However, both the ISRU MAV and the TERP MAV share similar configuration and construction. The major components of the MAV are the tank assembly, the thermal shell, the thrust frame, and the rocket engines.

The tank assembly consists of four cylindrical tanks with spherical ends, and tank connectors. The tank configuration was chosen to minimize the height of the MAV. The tanks are made of filament wound graphite epoxy because of its high strength and stiffness, and are lined with a thin aluminum shell that prevents leakage through the composite. The tank connectors are composed of a primary aluminum frame, graphite fiber separators that reduce thermal conduction, and connection rings that attach the tank assembly to the rest of the MAV.

The ISRU propellants (125 K) and TERP propellants (280 K) must be stored at significantly different temperatures from Mars ambient (215 K). In order to reduce the need for active cooling and heating, the tank assembly is enclosed in a thermal shell. To eliminate conductive heat transfer to the tank assembly on Mars, the shell is evacuated by opening a valve to space during Mars transfer and closing it shortly before Mars entry. The shell is made of aluminum honeycomb sized for the external atmospheric pressure at Mars. To reduce radiative heat transfer to the tank assembly, the shell is lined with 1 cm of MLI and a layer of highly polished gold film. This system reduces heat transfer to the tanks to 4 W. Since the thermal shell would add a significant amount to the burnout mass of the MAV, shaped charges are used to split the thermal shell and separate it from the MAV shortly after launch.

The cross-shaped thrust frame joins the MAV to the MLV and is composed of four aluminum I-beams. The thrust frame is designed for cantilever loads resulting from launch and entry accelerations. Since the moment in each beam increases along the length of the beam, the width of the flange tapers along the beam. The engines and pressurant system also attach to the thrust frame.

Rocket Engines

The ISRU MAV is equipped with four rocket engines (Fig. 11) that serve as retro-rockets for Mars landing and to lift the spacecraft off Mars and inject it into the transfer orbit back to Earth. These engines run on the methane and oxygen stored in the MAV tanks during propellant production. Engine operating characteristics are summarized in Table 3.

The engines operate at an oxidizer to fuel mass ratio (O/F) of 3.5. This mixture ratio was chosen because it provides the highest theoretical Isp, as calculated with the CET89 version of the Gordon McBride code. A high Isp minimizes the propellant required to achieve the DV of 6110 m/s for Earth return.

The rocket engines used are partly based on a previous design developed at the University of Washington. Each of the four engines produces 1650 N of thrust, for a total thrust of 6600 N. This thrust is sufficient for Mars liftoff, providing an initial thrust to weight ratio of 3.27 based on an initial MAV mass of 550 kg. Vehicle acceleration at liftoff is 0.86 times Earth’s gravity (0.86 ge), while at burnout the acceleration is 6.3 g_e.

The nozzle expansion ratio, ε = 125, was chosen to minimize the required propellant mass. A 70% bell nozzle shape is used. After applying a correction factor of 0.98 for losses to non-axial components of kinetic energy in a 70% bell nozzle, the Isp calculated is 383 sec.2 To account for some of the real world losses that one might expect (frozen composition, entropy production, and atmospheric effects on the surface of Mars), a conservative Isp value of 368 sec was used for design purposes.

The CET89 computer code estimates the chamber temperature at 3450 K. At this high chamber temperature, the thrust chamber will need some type of active cooling. Regenerative cooling using methane cools the combustion chamber and the nozzle up to an expansion ratio of 25, beyond which the nozzle extension uses radiative cooling. The methane enters a manifold at the injector end of the engine and passes through cooling channels to the expansion ratio 25 plane, where another manifold reverses the flow. The methane then flows up separate channels into the injector. The flow of methane limits the inside wall temperature of the nozzle throat to 900 K, and all other locations are cooler than this.

To reduce complexity and increase reliability, the rocket engines are pressure-fed rather than turbopump-fed. The propellant tanks are pressurized to 3.1 MPa. The propellant pressure drops across the isolation valves and main valves, in piping, in the manifold, and across the injectors. In addition, the fuel pressure drops through the regenerative cooling jacket. The pressure drop for both methane and oxygen across the injectors is 400 kPa, 20% of chamber pressure, to ensure pressure stability. This injector pressure drop reduces the propellant pressure to chamber pressure, 2.0 MPa.

The TERP engine is a single Marquardt R-40B modified with an extendible nozzle. A mass breakdown of the ISRU MAV and the TERP MAV is shown in Table 4.

Sample Return Capsule

The design of the SRC (Fig. 18) is based around the sample containment canister. The majority of the SRC mass and volume is comprised of the reentry aeroshell required to slow the capsule as it enters Earth’s atmosphere.

Table 5 shows the mass breakdown for the components of the SRC. The SRC frontal heat shield is 6 cm thick at the stagnation point to withstand a max heating load of 350 W/cm.2 The shield consists of a 4.4 cm thick layer of ablative AVCO 5026-H/CG, 0.2 mm of RTV 560 bonding material sandwiching 4 mm of Nomex felt strain isolation pad, and 0.4 mm thick layers of graphite epoxy sandwiching a 15 mm Nomex honeycomb core. The backshell has a 0.25 mm layer of reaction cured glass (RCG) coating, a 2.0 cm layer of fibrous refractory composite insulation (FRCI-12), and a 0.4 mm layer of graphite epoxy sandwiching a 5 mm layer of graphite epoxy honeycomb. , Because the samples need to remain at a temperature below -10 C, the heat shield is approximately 20% thicker than necessary to help insulate the samples.

The SRC itself does not contain the necessary equipment to make the trip from Mars to Earth. That equipment is contained on the Earth Transfer Cruise Stage (ETCS) described below.

Cruise Stages

The cruise stages provide the communications, navigation, and RCS systems necessary for the six-month transfer obits from Earth to Mars and from Mars to Earth. Both are based on JPL’s Pathfinder cruise stage.

Earth Transfer Cruise Stage (ETCS)

The ETCS serves as a service module to the SRC during Earth transfer (Fig. 19). The ETCS provides 130 m/s of DV necessary for trajectory correction maneuvers and attitude control, and also provides for communication with Earth and navigation.

Acceptable SRC temperatures are maintained by shielding the SRC from the sun with the ETCS. Finally, the ETCS must provide the power necessary to run its own systems. The total mass of the ETCS is 22 kg.

The ETCS is 0.7 m in diameter with a 0.2-m-high side wall and is constructed of aluminum honeycomb. The required electronics for the ETCS are gathered into two electronics compartments in order to maintain them between 280 K and 298 K. The electronics compartments are insulated from the rest of the ETCS on five sides and exposed to space via a radiator that penetrates the side walls. The SRC is insulated from the ETCS by polystyrene foam and MLI. This insulation is sufficient to maintain the SRC temperature at no greater than 140 K.

Power for the ETCS is provided by 22% efficient GaAs solar cells covering the 0.38 m2 frontal area of the ETCS with a packing efficiency of 88%. This solar array provides the minimum 37 W for the ETCS systems in a 500 W/m2 solar flux. As the ETCS approaches Earth, the solar flux will exceed 500 W/m2, so excess power will be shunted.

Propulsion for the ETCS is provided by eight 0.9 N Olin MR-103c monopropellant hydrazine thrusters. The MR-103c is widely used in communications satellite applications. Hydrazine is stored in two titanium tanks and is supplied to the thrusters by a blow down system. Attitude control and navigation is assisted by a star tracker, sun sensors, and the Interferometric Fiber Optic Gyro and Inertial Measuring Unit (IFOG/IMU). The ETCS contains a computer, transponder, and an antenna in order to maintain contact with Earth.

Mars Transfer Cruise Stage (MTCS)

During launch from Earth, the MTCS (Fig. 20) serves as the mating adapter between the MLV and the upper stage of the Delta II. During transfer, the MTCS performs tasks similar to those performed by the ETCS. The MTCS must provide 130 m/s of course correction and attitude control DV, provide navigation and communication with Earth, and provide power for MTCS and MLV systems en route to Mars.

The MTCS is constructed out of a 2.8 m diameter, 1 cm thick graphite epoxy and Nomex honeycomb panel. The MLV aerobrake sits in the MLV support frame which joins the MLV to the upper stage of the Delta II. Straps that are connected by explosive bolts to the structural ring on the back shell hold the MLV down into the MLV support frame. RCS propulsion is provided by a monopropellant blow down system similar to that of the ETCS. However, the MTCS uses 4.4 N Olin MR-111 thrusters. Navigation and communications are facilitated by a transponder, antenna, sun sensors, and a star tracker located on the MTCS. The computer and IFOG/IMU on the ETCS are used by the MTCS to reduce mass. The power of 120 W required by the MTCS is provided by 1.2 m2 of solar panel on the front side of the MTCS.