Applications of Ultra-Low Ballistic Coefficient Entry Vehicles to Existing and Future Space Missions

The ballistic coefficient (β = m cDA ) of entry vehicles has traditionally not been a design parameter: the mass was fixed by the launch vehicle payload, the drag coefficient was determined by the aerodynamic configuration, and the reference area (typically cross-section area) was fixed by the launch vehicle diameter. If the ballistic coefficient is instead considered as a potential design parameter, several benefits result from driving to lower and lower values. As the ballistic coefficient decreases, the peak stagnation point heating rate and temperature decrease. As β reaches the range of 150-300 Pa, the heat shield can be deployed using a mechanical framework (much like an umbrella) supporting existing ceramic fabric as the heat shield. Offsetting the center of gravity from the vehicle centerline allows lift/drag ratios in the range of 0.15-0.25, which mitigates entry decelerations and provides active targeting capability for a designated landing site. Due to the lower entry temperatures, ionization of the surrounding air stream is reduced or eliminated, allowing communications and GPS-based navigation throughout the entry trajectory. As the spacecraft enters the dense lower atmosphere, the low areal loading results in terminal velocities in the range of 15-20 m/sec, requiring only terminal decelerators (rockets or airbags) to mitigate the landing impact. This paper reviews the concept and development history of ultra-low ballistic coefficient (ULβ) entry vehicles, as introduction to its focus on the applications of this class of vehicles to existing (ISS) and future (exploration) missions. Results are presented for the use of ULβ vehicles in future missions, including intact cargo down-mass from ISS and aerocapture and direct hypervelocity entry, descent, and landing (EDL) for lunar and Mars missions. Of particular interest is the potential for a ULβ-type vehicle as an alternate approach to crew transfer vehicles currently being developed under the NASA COTS program. Such a system offers several unique advantages over more conventional capsule or winged designs. The large wake area and relatively low airstream enthalpies allow the use of simple cylindrical shapes for the crew cabin, rather than the low volumetric efficiencies of conical configurations. Since the ParaShield itself supplies the capabilities for entry, descent, and landing, the EDL components of the spacecraft could be easily detached for lunar missions, allowing the ULβ crew cabin to be used for lunar landing and exploration missions while the deployable entry shield is parked in lunar orbit to reduce landed mass. Similar advantages accrue to a human Mars mission, including the use of ULβ for aerobraking into Mars orbit and direct hypervelocity EDL upon return to Earth from both the Moon and Mars.