Design Of Lunar-Gravity-Assisted Escape Trajectories

AbstractLunar gravity assist is a means to boost the energy and C3 of an escape trajectory. Trajectories with two lunar gravity assists are considered and analyzed. Two approaches are applied and tested for the design of missions aimed at Near-Earth asteroids. In the first method, indirect optimization of the heliocentric leg is combined to an approximate analytical treatment of the geocentric phase for short escape trajectories. In the second method, the results of pre-computed maps of escape C3 are employed for the design of longer Sun-perturbed escape sequences combined with direct optimization of the heliocentric leg. Features are compared and suggestions about a combined use of the approaches are presented. The techniques are efficiently applied to the design of a mission to a near-Earth asteroid

Magnetour: Surfing Planetary Systems on Electromagnetic and Multi-Body Gravity Fields

In this NIAC Phase One study, we propose a new mission concept, named Magnetour, to facilitate the exploration of outer planet systems and address both power and propulsion challenges. Our approach would enable a single spacecraft to orbit and travel between multiple moons of an outer planet, with no propellant required. Our approach would enable a single spacecraft to orbit and travel between multiple moons of an outer planet, with no propellant nor onboard power source required. To achieve this free-lunch _Grand Tour', we exploit the unexplored combination of magnetic and multi-body gravitational fields of planetary systems, with a unique focus on using a bare tether for power and propulsion. The main objective of the study is to develop this conceptually novel mission architecture, explore its design space, and investigate its feasibility and applicability to enhance the exploration of planetary systems within a 10-year timeframe. Propellantless propulsion technology offers enormous potential to transform the way NASA conducts outer planet missions. We hope to demonstrate that our free-lunch tour concept can replace heavy, costly, traditional chemical-based missions and can open up a new variety of trajectories around outer planets. Leveraging the powerful magnetic and multi-body gravity fields of planetary systems to travel freely among planetary moons would allow for long-term missions and provide unique scientific capabilities and flagship-class science for a fraction of the mass and cost of traditional concepts. New mission design techniques are needed to fully exploit the potential of this new concept.This final report contains the results and findings of the Phase One study, and is organized as follows. First, an overview of the Magnetour mission concept is presented. Then, the research methodology adopted for this Phase One study is described, followed by a brief outline of the main findings and their correspondence with the original Phase One task plan. Next, an overview of the environment of outer planets is provided, including magnetosphere, radiation belt and planetary moons. Then performance of electrodynamic tethers is assessed, as well as other electromagnetic systems. A method to exploit multi-body dynamics is given next. These analyses allow us to carry out a Jovian mission design to gain insight in the benefits of Magnetour. In addition, a spacecraft configuration is presented that fully incorporates the tether in the design. Finally technology roadmap considerations are discussed

Overview of the Mission Design Reference Trajectory for NASA's Asteroid Redirect Robotic Mission

The National Aeronautics and Space Administration's (NASA's) recently cancelled Asteroid Redirect Mission was proposed to rendezvous with and characterize a 100 m plus class near-Earth asteroid and provide the capability to capture and retrieve a boulder off of the surface of the asteroid and bring the asteroidal material back to cislunar space. Leveraging the best of NASA's science, technology, and human exploration efforts, this mission was originally conceived to support observation campaigns, advanced solar electric propulsion, and NASA's Space Launch System heavy-lift rocket and Orion crew vehicle. The asteroid characterization and capture portion of ARM was referred to as the Asteroid Redirect Robotic Mission (ARRM) and was focused on the robotic capture and then redirection of an asteroidal boulder mass from the reference target, asteroid 2008 EV5, into an orbit near the Moon, referred to as a Near Rectilinear Halo Orbit where astronauts would visit and study it. The purpose of this paper is to document the final reference trajectory of ARRM and the challenges and unique methods employed in the trajectory design of the mission

Potential Cislunar and Interplanetary Proving Ground Excursion Trajectory Concepts

NASA has been investigating potential translunar excursion concepts to take place in the 2020s that would be used to test and demonstrate long duration life support and other systems needed for eventual Mars missions in the 2030s. These potential trajectory concepts could be conducted in the proving ground, a region of cislunar and near-Earth interplanetary space where international space agencies could cooperate to develop the technologies needed for interplanetary spaceflight. Enabled by high power Solar Electric Propulsion (SEP) technologies, the excursion trajectory concepts studied are grouped into three classes of increasing distance from the Earth and increasing technical difficulty: the first class of excursion trajectory concepts would represent a 90-120 day round trip trajectory with abort to Earth options throughout the entire length, the second class would be a 180-210 day round trip trajectory with periods in which aborts would not be available, and the third would be a 300-400 day round trip trajectory without aborts for most of the length of the trip. This paper provides a top-level summary of the trajectory and mission design of representative example missions of these three classes of excursion trajectory concepts

Near Earth Asteroid Scout - Mission Update

After its deployment from NASA’s Space Launch System (SLS), the Near-Earth Asteroid (NEA) Scout mission will travel to and image an asteroid during a close flyby using an 86m2 solar sail as its primary propulsion. Solar sails are large, mirror-like structures made of a lightweight material that reflects sunlight to propel the spacecraft. The continuous solar photon pressure provides thrust with no need for the heavy, expendable propellants used by conventional chemical and electric propulsion systems. Developed by NASA’s Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL), the NEA Scout is based on the industry-standard CubeSat form factor. The spacecraft measures 11 cm x 24 cm x 36 cm and weighs less than 14 kilograms. Following deployment from the Space Launch System (SLS), the solar sail will deploy, and the spacecraft will begin its 2.0 – 2.5-year journey. About one month before the asteroid flyby, NEA Scout will search for the target and start its Approach Phase using a combination of radio tracking and optical navigation and perform a relatively slow flyby (10-20 m/s) of the target. A summary of the mission, sailcraft, mission design, and its first several months of deep space operation will be described

Simulating Mars on Earth: Georgia Tech's Crew 47 at the Mars Desert Research Station

Presentation for the Mars Society @ Georgia Tech

Comet Hitchhiker: NIAC Phase 1 Final Report

Summary of Activities-Developed the Comet Hitchhiker concept, which is to hitch rides on small bodies (asteroids and comets) using a tethered spacecraft. (Section 2)-Identified five scientifically important missions that would be enabled or significantly benefited by the Comet Hitchhiker concept.The five mission concepts are: KBO rendezvous, Centaur rendezvous, Trojan rendezvous, Damocloid rendezvous, and Main asteroid belt tour to rendezvous with multiple (10) objects. (Section 3)-Derived the Space Hitchhike Equation, or "the rocket equation for hitchhiker", which relates the specific strength of tether, mass ratio, and V. (Section 4.1)-Performed in-depth feasibility analysis of the critical components of the concept through: Finite-element simulations of tether and spacecraft dynamics, as shown in Figure 1 (Section 4.4); Supercomputer simulations of the hypervelocity impact of harpoon on a small body, as shown in Figure 2. (Section 6)-Performed public outreach activities including the collaboration with a concept artist of the Museum of Science Fiction, exposure to media, and public presentations. (Section 8

A methodology for robust optimization of low-thrust trajectories in multi-body environments

Future ambitious solar system exploration missions are likely to require ever larger propulsion capabilities and involve innovative interplanetary trajectories in order to accommodate the increasingly complex mission scenarios. Two recent advances in trajectory design can be exploited to meet those new requirements: the use of low-thrust propulsion which enables larger cumulative momentum exchange relative to chemical propulsion; and the consideration of low-energy transfers relying on full multi-body dynamics. Yet the resulting optimal control problems are hypersensitive, time-consuming and extremely difficult to tackle with current optimization tools. Therefore, the goal of the thesis is to develop a methodology that facilitates and simplifies the solution finding process of low-thrust optimization problems in multi-body environments. Emphasis is placed on robust techniques to produce good solutions for a wide range of cases despite the strong nonlinearities of the problems. The complete trajectory is broken down into different component phases, which facilitates the modeling of the effects of multiple bodies and makes the process less sensitive to the initial guess. A unified optimization framework is created to solve the resulting multi-phase optimal control problems. Interfaces to state-of-the-art solvers SNOPT and IPOPT are included. In addition, a new, robust Hybrid Differential Dynamic Programming (HDDP) algorithm is developed. HDDP is based on differential dynamic programming, a proven robust second-order technique that relies on Bellman's Principle of Optimality and successive minimization of quadratic approximations. HDDP also incorporates nonlinear mathematical programming techniques to increase efficiency, and decouples the optimization from the dynamics using first- and second-order state transition matrices. Crucial to this optimization procedure is the generation of the sensitivities with respect to the variables of the system. In the context of trajectory optimization, these derivatives are often tedious and cumbersome to estimate analytically, especially when complex multi-body dynamics are considered. To produce a solution with minimal effort, an new approach is derived that computes automatically first- and high-order derivatives via multicomplex numbers. Another important aspect of the methodology is the representation of low-thrust trajectories by different dynamical models with varying degrees of fidelity. Emphasis is given on analytical expressions to speed up the optimization process. In particular, one novelty of the framework is the derivation and implementation of analytic expressions for motion subjected to Newtonian gravitation plus an additional constant inertial force. Example applications include low-thrust asteroid tour design, multiple flyby trajectories, and planetary inter-moon transfers. In the latter case, we generate good initial guesses using dynamical systems theory to exploit the chaotic nature of these multi-body systems. The developed optimization framework is then used to generate low-energy, inter-moon trajectories with multiple resonant gravity assists.Ph.D.Committee Chair: Russell, Ryan; Committee Member: Braun, Robert; Committee Member: Clarke, John-Paul; Committee Member: Dargent, Thierry; Committee Member: Sims, Jon; Committee Member: Tsiotras, Panagioti

DESIGN OF LUNAR-GRAVITY-ASSISTED ESCAPE MANEUVERS

Lunar gravity assist is a means to boost the energy and C3 of an escape maneuver. Two approaches are applied and tested for the design of trajectories aimed at Near-Earth asteroids. Maneuvers with two lunar gravity assists are considered and analyzed. Indirect optimization of the heliocentric leg is combined to an approximate analytical treatment of the geocentric phase for short escape maneuvers. The results of pre-computed maps of escape C3 are used for the design of longer sun-perturbed escape sequences. Features are compared and suggestions about a combined use of the approaches are presented. The techniques are efficiently applied to the design of a mission to a near-Earth asteroid