Mars Global Surveyor aerodynamics for maneuvers in Martian atmosphere

The aerodynamic and heating characteristics of the Mars Global Surveyor spacecraft in its planned aerobraldng maneuver were investigated. Analytical and Direct Simulation Monte Carlo codes were employed to obtain results for aerobraking flight through free molecular and transition flow regimes. The spacecraft was found to be stable in various configurations. Solar panel heating was found to be higher on the edges than in the center. Attitude control system thruster plume freestream interaction leads to thrust reversal and diminished control system effectiveness. Drag and heat transfer coefficients were found to be highly dependent on atmospheric density. Nomenclature A«f = reference area Lref = reference length CD = drag coefficient, drag / 0.5pVArcf CL = lift coefficient, lift / O.SpV'A,^ CM = moment coefficient, moment /0.5pVArefLref X= mean free path Kn = Knudsen number, A, / Lref p = freestream mass density, kg/km Introduction In order to learn more about the Martian atmosphere, surface, and gravity field, the Mars Surveyor Program will launch a pair of spacecraft to Mars every twenty-six months. Mars Global Surveyor (MGS), one of the first pair of spacecraft built in the program, was successfully launched on November 7, 1996. MGS is expected to arrive at Mars on September 11, 1997, and will be captured in an elliptical orbit using conventional chemical propulsion. Aerobraking in the upper atmosphere will then be used to reduce orbital energy and establish a nearly circular mapping orbit. During aerobraking, the spacecraft will pass through different atmospheric densities and be subjected to various aerodynamic and heating loads. To ensure that the MGS spacecraft can maintain aerodynamic stability and encounter heating rates not exceeding design values, an analysis of the spacecraft aerothermodynamics and the effectiveness of the Attitude Control System (ACS) thrusters was performed. This paper will present the results of this analysis and the methods used to obtain them. The first two sections will describe the spacecraft, the aerobraking maneuver, and the configurations associated with it. MGS will pass through altitudes that include the free molecule and transition flow regimes. The Graduate Student, Member A1AA. Aero-Space Technologist, Aerosol Research Branch, Atmospheric Sciences Division. Member AIAA. Professor of Engineering and Applied Science. Associate Fellow AIAA. Copyright <8) 1997 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental Purposes. All other rights are reserved by the copyright owner. 1 American Institute of Aeronautics and Astronautics Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. third section will discuss the numerical methods used. An analytical code was used for the free molecule calculations whereas Direct Simulation Monte Carlo (DSMC) code was employed for the transition regime. A highly detailed geometry was used for the computational model. The next section will present the results for the aerobraking configuration, which was designed to provide sufficient drag and high static stability. The fifth section will discuss the "safing" configuration and the aerodynamics associated with it. The sixth section contains results for an analysis of thruster plumefreestream interaction. Thrust reversal may occur for certain aerobraking altitudes. The last section investigates the dependence of drag coefficient and heat transfer coefficient on atmospheric density. Mars Global Surveyor Spacecraft The Mars Global Surveyor Spacecraft was built by Lockheed Martin Astronautics and consists of four subassemblies: the equipment module, propulsion module, solar arrays, and communication antennas. The spacecraft is shown in Figure 1 in its mapping configuration. The equipment module houses the avionics and science instruments and measures 1.22 x 1.22 x 0.76 meters. All of the instruments except the magnetometer are bolted to the nadir equipment deck on the +Z side. The propulsion module on the -Z side contains the main engine, attitude control thrusters, and propellant tanks. It consists of a rectangular shaped box 1.06 m on a side, not including the thermal blankets. There are three thrusters per corner; two aft facing, and one for roll control. Each thruster burns hydrazine in a mono-propellant, pulse mode. The two solar arrays, which will provide power for the spacecraft, measure 3.53 meters in length and 1.85 meters in width. Each array consists of an inner and outer panel which are comprised of gallium arsenide and silicon cells, respectively. The structure which supports the solar array and connects it to the equipment and propulsion modules is herein referred to as the "yoke. " Metal "drag flaps" 0.81 meters in length are mounted on the ends of both arrays. These serve to decrease the spacecraft's ballistic coefficient during aerobraking. The primary communication antenna is the 1.5 meter high gain antenna (HGA). During aerobraking, the HGA will remain body fixed to the +X side of the spacecraft and its protective cover attached. Aerobraking Maneuver and Configurations The aerobraking maneuver will progressively lower MGS from its highly elliptical capture orbit (56,600 km altitude apoapsis, 353 km altitude periapsis) to a nearly circular one (450 km altitude apoapsis, 350 km periapsis ) for mapping of the planet. This maneuver will be accomplished in three phases. The first phase, set to last for seven or eight orbits, will gradually lower the periapsis to 112 km using four propulsive impulses with the attitude control thrusters. This gradual "walk-in" is necessitated by the large uncertainty in the atmospheric density model and will allow time to study the atmosphere and its orbit-to-orbit variation. The second or main phase follows will last about three months. MGS will make 300-400 successive passes through the atmosphere, bringing the apoapsis down to 2000 km. The periapsis altitude will be maintained within a specified corridor to produce enough drag within time constraints yet remain high enough to avoid excessive heating and maximum dynamic pressure limits. The last phase follows for about three weeks. During this time, the apoapsis altitude will be lowered to its final value of 450 km, and periapsis gradually raised to 143 km. Aerobraking will end with a termination burn to raise the periapsis to its final altitude of 350 km. The aerodynamic forces during aerobraking are expected to account for a AV decrease of 1300 m/s. The originally planned aerobraking configuration for MGS is shown in Figure 2 with a representative computational model. The panels are swept 30° relative to the Y axis with the solar cell side opposite that exposed to the freestream. After launch and trajectory insertion, it was discovered that the -Y solar array had not fully deployed and snapped into place. The array, from the inner panel on out, remains about 20.5° from full extension to the plane of the yoke. It is suspected that a shaft in the deployment damper mechanism broke shortly after launch, thereby causing the damper arm to wedge into the hinge joint between the inner panel and yoke. Concerned that the hinge would not be able to withstand the aerodynamic torques associated with the aerobraking maneuver, MGS engineers revised the aerobraking configuration; the new configuration is shown in Figure 3. American Institute of Aeronautics and Astronautics Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. The panel assembly is rotated 180° about the Y axis, and the yoke is now oriented 51° from the -Y axis instead of 30°. This retains the 30° sweep while ensuring that any deflection of the array would be into the secured position (with the wedged damper arm doubling as a mechanical "stop"). The solar cells are now exposed to the freestream; aerodynamic and heating issues related to this configuration are addressed in a latter section. In the event of contingencies, the MGS spacecraft can adopt a "safing" configuration. This configuration will cause the spacecraft to roll over until the solar arrays are aligned in the direction of the sun. This allows the navigation team to be fairly certain of the MGS attitude once it conies out of a drag pass. To obtain this configuration from the aerobraking configuration, the panel assemblies are rotated 25° about the Y axis toward the HGA. This configuration can be seen in Figure 4. There is a revised safing configuration associated with the revised aerobraking configuration; the characteristics of this new safing configuration are currently being evaluated and are not included in this paper.