Risk-aware Mission Design Around Small Celestial Bodies

Robotic exploration of small celestial bodies is integral to current and future space programs. However, complex, uncertain dynamical environments around small bodies pose challenges for spacecraft to safely navigate around them. For successful small-body exploration, the mission design process needs to be aware of potential risks (e.g., collision with the body, loss of science opportunities), appropriately quantify these risks, and mitigate them to ensure achieving the expected performance (e.g., safety, cost) in the presence of uncertainty. To this end, this dissertation develops risk-aware mission design frameworks for robust small-body exploration under uncertainty, by merging techniques and insights in the fields of astrodynamics, stochastic control, optimal control, and optimization. In particular, this dissertation is focused on developing frameworks for robust spacecraft guidance and space trajectory optimization under uncertainty. The developed guidance framework combines chance-constrained optimal control, convex programming, and local approximation of orbital dynamics, to optimize a sequence of guidance policies that guarantee, with a user-defined confidence level, precise control of spacecraft orbits about a reference trajectory in the presence of uncertainty. The framework for trajectory optimization under uncertainty leverages dynamical properties of orbital mechanics to enable efficient semi-analytical uncertainty quantification within the optimization routine, formulated as a class of the indirect method by applying the calculus of variations. Furthermore, this dissertation also develops an orbit control framework that exploits one of the major disturbances in small-body environments, solar radiation pressure, to effectively utilize the natural force as a primary source of orbit controls, allowing greater flexibility in mission design for small-body exploration. These frameworks provide mathematical and computational tools for the design of robust exploration missions under uncertainty, paving the way for better access to and safer operations at small celestial bodies in our Solar System.</p

EQUULEUS Trajectory Design

This paper presents the trajectory design for EQUilibriUm Lunar-Earth point 6U Spacecraft (EQUULEUS), which aims to demonstrate orbit control capability of CubeSats in the cislunar space. The mission plans to observe the far side of the Moon from an Earth-Moon L2 (EML2) libration point orbit. The EQUULEUS trajectory design needs to react to uncertainties of mission design parameters such as the launch conditions, errors, and thrust levels. The main challenge is to quickly design science orbits at EML2 and low-energy transfers from the post-deployment trajectory to the science orbits within the CubeSat’s limited propulsion capabilities. To overcome this challenge, we develop a systematic trajectory design approach that 1) designs over 13,000 EML2 quasi-halo orbits in a full-ephemeris model with a statistical stationkeeping cost evaluation, and 2) identifies families of low-energy transfers to the science orbits using lunar flybys and solar perturbations. The approach is successfully applied for the trajectory design of EQUULEUS

EQUULEUS: Mission to Earth - Moon Lagrange Point by a 6U Deep Space CubeSat

A 6U Deep Space CubeSat EQUULEUS (EQUilibriUm Lunar-Earth point 6U Spacecraft) will be the world’s smallest spacecraft to explore the Earth―Moon Lagrange point. T he spacecraft is jointly developed by the University of Tokyo and JAXA and will belaunched by NASA’s SLS (Space Launch System) EM-1 (Exploration Mission-1) in 2018. The spacecraft will fly to a libration orbit around the Earth-Moon L2 point and demonstrate trajectory control techniques within the Sun-Earth-Moon region (e.g. low-energy transfers using weak stability regions) for the first time by a nano spacecraft. This mission will contribute to the realization of the future efficient cargo transfers to deep space ports located at the Lagrange points. This mission also carries several scientific observation instruments. The first one, named PHOENIX (Plasmaspheric Helium ion Observation by Enhanced New Imager in eXtreme ultraviolet), will conduct the imaging of the Earth\u27s plasmasphere by extreme UV wavelength. The observation will complement and enhance the geospace in-situ observation conducted by the ERG (JAXA\u27s small space science mission to be launched in 2016) and Van Allen probe (NASA) missions. As a result, we can improve our understanding of the radiation environment around the Earth, which is one of the critical issues for human cis-lunar exploration. The second scientific observation instrument, named CLOTH (Cis-Lunar Object Detector within Thermal Insulation), will detect and evaluate the meteoroid impact flux in the cis-lunar region by using dust detectors implemented in the spacecraft’s MLI (Multi-Layer Insulation). The goal of this mission is to understand the size and spatial distribution of solid objects in the cis-lunar space. The third scientific observation instrument, named DELPHINUS (DEtection camera for Lunar impact PHenomena IN 6U Spacecraft), will observe the impact flash at the far side of the moon from Earth—Moon L2 point (EML2) for the first time. This observation will characterize the flux of impacting meteors, and the results will contribute to the risk evaluation for future human activity and/or infrastructure on the lunar surface. EQUULEUS will use X-band and Ka-band frequencies for the deep space telecommunication. Japanese deep space antenna (64-meter antenna and 34-meter antenna) will be nominally used for the spacecraft operation, and the support from DSN (Deep Space Network) of JPL is also being planned. This paper describes mission outline, spacecraft system design, and some newly developed technologies