Jon Sims

Implementation of a Low-Thrust Trajectory Optimization Algorithm for Preliminary Design

A tool developed for the preliminary design of low-thrust trajectories is described. The trajectory is discretized into segments and a nonlinear programming method is used for optimization. The tool is easy to use, has robust convergence, and can handle many intermediate encounters. In addition, the tool has a wide variety of features, including several options for objective function and different low-thrust propulsion models (e.g., solar electric propulsion, nuclear electric propulsion, and solar sail). High-thrust, impulsive trajectories can also be optimized.

Mars Free Return Trajectories

Trajectories to Mars have been separated into various classes by previous a ~ t h o r s . l ~ These classes include Opposition, Sprint, Conjunction, Free Return, and Cycler. Opposition class missions are characterized by a high-energy trajectory and a relatively short Mars stay time (<3 months). The name stems from the fact that the Earth leaves opposition with Mars a t the Mars arrival. The total mission duration for this class ranges from 1 to 2.5 years. Also, the Opposition class trajectories have high arrival velocities a t both Mars and Earth. Since the periapsis of the transfers can be lower than the orbit of Venus, the addition of a Venus swingby can improve mission performance. A subset of the Opposition class is the Sprint class. This class has a mission duration of approximately 1 to 1.4 years with a 30 day stay time. Sprint missions are intended for piloted missions because of their short flight times, but they have higher AV requirements. The most traditional mission class is the Conjunction class. In this class the Earth is moving into conjunction with Mars at the time of Mars arrival. These missions are characterized by low-energy trajectories and have a relatively long stay time (0.8 to 1.5 years). They can be used during the early exploratory phase where many tasks need to be done on the planetary surface. For initial piloted missions, the Free Return class will most likely be the mission of choice since these trajectories do not require a deterministic maneuver to return the spacecraft to the Earth in the event of an emergency (e.g., Apollo 13). A subset of the Free Return class is the Cycler class,3 which includes

Aerogravity-assist trajectories to the outer planets

Abstract Aerogravity assist can significantly reduce the required launch energy and time of flight to the outer planets. An automated trajectory design program provides a thorough search of the near-future trajectory space for missions to Jupiter, Saturn, Uranus, Neptune and Pluto, using Venus and/or Mars for aerogravity assists. It is shown analytically that if high lift-to-drag vehicles (e.g. waveriders) are used, then the drag loss has minimal effect on the interplanetary trajectories.

Navigation Challenges of a Kinetic Energy Asteroid Deflection Spacecraft

*† ‡ § ** †† , A simple asteroid deflection experiment can be conducted by a small and inexpensive impactor spacecraft in cooperation with an independent asteroid rendezvous mission. The rendezvous mission would arrive at a near-Earth asteroid in advance of the impactor in order to characterize the asteroid while the impactor makes its way to the target. Then the rendezvous spacecraft would be in position to record the impact event, as well as to measure the deflection imparted as a result of the collision. This paper describes the mission concept and addresses the navigational challenges in delivering the impactor, as well as measuring the resultant deflection. I. Introduction simple and robust asteroid deflection experiment can be conducted by a relatively small and inexpensive impactor spacecraft, in close cooperation with an independent asteroid rendezvous mission. The rendezvous mission would arrive at a benign near-Earth asteroid in advance of the impactor launch in order to characterize the asteroid, acquire samples, etc., while the impactor makes its way to the target. Then the rendezvous spacecraft would be in position to record the impact event and document the post-impact scene, as well as to measure the deflection imparted as a result of the impactor’s collision with the asteroid. This concept requires that the rendezvous spacecraft already be in orbit around a candidate asteroid to take data before, during and after the impact to verify that a measurable change in the asteroid’s velocity took place and to study the dynamics of the impact itself. Such a mission would answer the critical question as to what momentum enhancement is provided by the ejected asteroidal material during the cratering process. This “blow back” momentum enhancement effect is thought to be significantly larger than the momentum provided by the impacting spacecraft itself. In addition to the obvious planetary defense aspects of such a mission, the scientific returns would be immense. Spacecraft information could be used to characterize the asteroid’s size, shape, mass, morphology, bulk density, porosity, rotation state, and composition, while an in situ study of the cratering event itself would allow insight into the asteroid’s interior structure.

Analysis of System Margins on Deep Space Missions Using Solar Electric Propulsion

NASAs Jet Propulsion Laboratory has conducted a study focused on the analysis of appropriate margins for deep space missions using solar electric propulsion (SEP). The purpose of this study is to understand the links between disparate system margins (power, mass, duty cycle, etc.) and their impact on overall mission performance and robustness. It is shown that the various sources of uncertainty and technical risk associated with electric propulsion mission design can be summarized into three relatively independent parameters 1) Electric Propulsion Power Margin, 2) Propellant Margin and 3) Duty Cycle Margin. The overall relationship between these parameters and other major sources of uncertainty is presented. A detailed trajectory analysis is conducted to examine the impact that various assumptions related to power, duty cycle, destination, and thruster performance including missed thrust periods have on overall performance. Recommendations are presented for system margins for deep space missions utilizing solar electric propulsion.

A Uranus-Neptune-Pluto opportunity

Abstract Recently, NASA has announced a change of philosophy toward smaller, faster spacecraft. A notable example of the new philosophy are plans for a Pluto fast flyby mission, involving a 150 kg spacecraft which can reach Pluto in about 7 years. In this paper, automated software is used to investigate multiple-encounter missions to the outer planets which can also be flown by the Pluto fast flyby spacecraft. A particularly important result of this analysis is the discovery of a three-planet opportunity with Uranus, Neptune and Pluto.

Trajectory options for low-cost missions to asteroids

Abstract We consider a wide variety of gravity-assist trajectories using Venus, Earth and Mars to obtain low launch energy trajectories to four large asteroids in the main belt. These trajectories are constructed by analytic and numeric search techniques. We optimize promising trajectories for minimum total Δ V and search for additional (nontargeted) asteroid flybys. Several optimized opportunities with multiple asteroid flybys are reported, followed by a discussion of general characteristics of the various trajectory types.

Electric Propulsion System Selection Process for Interplanetary Missions

† Engineer, Trajectory Design and Navigation, M/S 301-121, Member AIAA. ** Staff Engineer, Flight Systems Engineering, M/S 264-623, Member AIAA. ‡‡ Senior Engineer; Guidance, Navigation, & Control; M/S 301-150; Member AIAA. * Senior Engineer, Flight Systems Engineering, M/S T1722, Senior Member AIAA. ‡ Project Element Manager, Propulsion and Materials Engineering, M/S 125-109, Senior Member AIAA. †† Principle Member of Engineering Staff, Guidance Navigation and Control, M/S 301-121, Senior Member AIAA. § Senior Power Systems Engineer, Power Systems, M/S 303-310K.