Chit Hong Yam

Preliminary Design of Nuclear Electric Propulsion Missions to the Outer Planets

*† * ‡ Nuclear electric propulsion is expected to open new doors in deep space exploration. We study direct rendezvous missions to the outer planets which employ a constant-thrust, constant specific-impulse engine. We also consider how gravity assist can augment the capability of nuclear electric propulsion. We present numerical examples of gravity-assist missions to the outer planets, which use an engine similar to that of the Jupiter Icy Moons Orbiter. For example, in an Earth-Venus-Earth-Jupiter-Pluto mission, the spacecraft launches with a V∞ of 2.2 km/s and rendezvous with Pluto in 10.5 years, with a propellant mass fraction of 50%. We demonstrate the benefit of using intermediate gravity-assist bodies (e.g. Venus, Earth and Mars) to decrease both mission duration and propellant cost.

Notable Two-Synodic-Period Earth-Mars Cycler

Earth-Mars cycler trajectories (cyclers) could play an important role in a future human transportation system to Mars. A particular cycler that repeats every two synodic periods and has one intermediate Earth encounter is very promising. In a circular-coplanar model it requires no propulsive maneuvers, has 153-day transfer times between Earth and Mars, and has arrival V ∞ magnitudes of 4.7 km/s at Earth and 5.0 km/s at Mars. A method to find an analog cycler in a more realistic model (i.e., using an accurate ephemeris for the states of Earth and Mars) is described. Two cost metrics are considered: total cycler AV and total cycler AV plus total taxi ΔV. Numerical solutions are presented for both metrics. The total required AV is very small, though not zero. If the Earth-Mars and Mars-Earth transit times are constrained, then the characteristics of the optimal cycler trajectory change. Tradeoffs between maximum transit time and other mission characteristics are analyzed for all possible launch periods.

Low-Thrust Trajectories to Jupiter via Gravity Assists from Venus, Earth, and Mars

We design low-thrust gravity-assist trajectories to Jupiter via gravity assists from Venus, Earth, and Mars. Trades between time of flight, mass, and hardware specifications are examined for various flyby bodies such as Mars, Earth, Venus-Earth, and Venus-Venus. We find (locally) propellant-optimal trajectories for ranges of specific impulse and specific mass that represent present-day technologies (such as an ion thruster with a radioisotope power source) and future technologies (such as a nuclear-powered magnetoplasmadynamic thruster which may be available in 10 to 15 years). We consider the 35-year launch period from January 2010 to December 2044.

Saturn Impact Trajectories for Cassini End-of-Mission

Potential end-of-mission scenarios to be considered for the Cassini spacecraft must satisfy planetary quarantine requirements designed to prevent contamination of a pristine environment, which could include Titan and the other Saturnian moons. One assumed acceptable option for safe disposal of the spacecraft includes Saturn impact trajectories. Two classes of impact trajectories are investigated: short-period orbits characterized by periods of 6– 10days and long-period orbitswithperiods greater than850days.To impact Saturnwith short-period orbits, a series of successive Titan flybys is required to increase inclination and decrease periapsis to within Saturn’s atmosphere, while simultaneously avoiding the rings and mitigating V expenditures. To ensure that the spacecraft is not prematurely damaged by material in the rings, Tisserand graphs are employed to determine when the ring-plane crossing distance is within the F–G ring gap: the necessary geometry for the penultimate transfer. For long-period impact trajectories, solar gravity is exploited to significantly lower periapsis. Depending on the size andorientation of the long-period orbit, a maneuver (<50 m=s) at apoapsis must be added to ensure impact. For sufficiently large orbits with favorable characteristics, solar gravity alone drops the spacecraft’s periapsis into Saturn’s atmosphere. No maneuver is necessary after the final Titan flyby, providing an attractive “flyby-and-forget” option.

Reduced Parameterization for Optimization of Low-Thrust Gravity-Assist Trajectories: Case Studies

Low-thrust trajectories can be modeled as a series of impulsive ( ΔV) maneuvers connected by conic arcs. We study new ways of parameterizin g the ΔV vectors with a reduced number of variables and constraints. When optimizing low-thrust gravity-assist trajectories, the ΔV magnitudes can be parameterized with switch on/off times of the engine; the steering angles can be parameterized with coefficients of a Chebyshev series. We present numerical results for several missions, including: Earth-Jupiter rendezvous, Earth-MarsVesta flyby, Earth-Mercury rendezvous and a seven-synodic- period Earth-Mars roundtrip mission. In most of these cases, we found significant improvements in convergence speed (with acceptable accuracy) with the new formulations.