Feasibility of aerocapture in a rapid crewed mission to Mars

Rapid missions would be a desirable way for people to reach Mars, ensuring crew safety against the hazards of sustained exposure to reduced gravity and radiation, although posing significant challenges in terms of mass and deceleration requirements, for which suitable propulsion systems are yet to be developed. This work investigates the integration of aerocapture as a strategy to reduce propellant expenditure at Mars, aiming to reduce overall mission mass and duration. The interplanetary trajectory that leads to rapidly reaching Mars from Low Earth Orbit is defined alongside a spacecraft that can safely accommodate the crew, subsystems, multiple tanks, and a poten- tially massive propulsion system. Following a detailed assessment of the atmospheric entry environment, a 70◦ sphere-cone aeroshield is designed to ensure sufficient deceleration for the spacecraft. Key constraints to ensure crew safety and mission success are identified and shown to impose limits on the entry velocity, entry mass, and aeroshield size. Pareto fronts for the Initial Mass in Low Earth Orbit are traced for missions employing aerocapture versus those relying solely on propulsion systems, considering various mission durations and propulsion technologies. Results reveal the potential of aerocapture to enable shorter transfer durations across all propulsion systems, with mass benefits exceeding 80% for available options, albeit diminishing with increasing propulsion system efficiency. Aerocapture seems to offer a viable way to significantly diminish the mass, and therefore the price, of rapid crewed missions to Mars.