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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRSDate not supplied

Trajectory Design for a Spacecraft Capable of Deploying Probes to the Martian Surface en Route to Low Mars Orbit

The presented trajectory design and analysis was performed for the Aeolus spacecraft mission concept. The Aeolus spacecraft consists of an orbiter (i.e., “mothership”) with the goal of transferring from Earth to low Mars orbit via propulsive and/or atmospheric braking (i.e., aerobraking). During various phases of flight from hyperbolic approach of Mars through low Mars Orbit, the Aeolus orbiter will deploy multiple probes which are targeted to land on the Martian surface with the goal of achieving global surface coverage.

Mars hyperbolic approach↗
NASA NTRS2004

SIM trajectory design

SIM (Space Interferometry Mission) is a mission scheduled to launch in 2010 and will be the first spacecraft to use interferometry to measure the positions of stars to within 1 micro-arcsecond - a degree of precision never to before achieve. The flight hardware required to achieve this level of precision is very sensitive to its external environment, which places a number of challenging constraints on the trajectory design. This paper discusses the various trajectory options that were considered.

Titan Mission↗
NASA NTRS1992

Trajectory design and navigation analysis for Cargo Transfer Vehicle proximity operations

The paper documents an investigation of alternative (or off-nominal) final approach trajectories and terminal guidance algorithms for Cargo Transfer Vehicle (CTV) automated rendezvous and proximity operation in the vicinity of Space Station Freedom (SSF). Ten candidate trajectory designs, representing a survey of all approaches that reasonably satisfy the mission constraints, are compared qualitatively and quantitatively. It is shown that several fundamentally different types of approach strategies are feasible in the CTV application and enjoy more favorable passive collision avoidance and plume impingement properties than the nominal. They are, however, generally more costly in propellant. The extent to which an alternative approach is more costly is determined largely by the abort (or specifically the 'hold' or 'wave-off') requirements. These approach techniques are verified in a software simulation using a parametrically defined candidate navigation system, and proximity sensor accuracy requirements are derived.

Marsh, S. M.↗
NASA NTRS2001

Trajectory Design for the Microwave Anisotropy Probe (MAP)

The Microwave Anisotropy, Probe (MAP) is a Medium Class Explorers (MIDEX) Mission produced in partnership between Goddard Space Flight Center (GSFC) and Princeton University. The goal of the MAP mission is to produce an accurate fill-sky, map of the cosmic microwave background temperature fluctuations (anisotropy). The mission orbit is a Lissajous orbit about the L(sub 2) Sun-Earth Lagrange point. The trajectory design for MAP is complex, having many requirements that must be met including shadow avoidance, sun angle constraints, Lissqjous size and shape characteristics, and limited Delta-V budget. In order to find a trajectory that met the design requirements for the entire 4-year mission lifetime goal, GSFC Flight Dynamics engineers performed many analyses, the results of which are presented herein. The paper discusses the preliminary trade-offs to establish a baseline trajectory, analysis to establish the nominal daily trajectory, and the launch window determination to widen the opportunity from instantaneous to several minutes for each launch date.

Newman, Lauri Kraft↗
NASA NTRS2002

An Overview of Trajectory Design Operations for the Microwave Anisotropy Probe (MAP) Mission

The main science objective of the Microwave Anisotropy Probe (MAP) mission is to produce an accurate full-sky map of the cosmic microwave background temperature fluctuations - anisotropy. MAP will collect these measurements from a lissajous orbit about the Sun-Earth/Moon L2 Lagrange Point. The NASA Goddard Space Flight Center (GSFC) Flight Dynamics Analysis Branch provided mission analysis, maneuver planning and maneuver calibration for the MAP spacecraft. This paper will provide an overview of the MAP trajectory design, a summary of the maneuvers executed. Differences from the pre-launch nominal plan will also be discussed. During the MAP phasing loops, MAP performed three calibration maneuvers in order to characterize the performance of the primary sets of thrusters - +X, +Z, and -Z. The calibration maneuvers were designed to minimize their impact on the trajectory. Four maneuvers were performed to set up the gravity assist of the Moon - required to propel MAP out to its orbit about L2. These maneuvers were performed at the three phasing loop perigees and at 18 hours after the final perigee. It became necessary to alter some of the perigee maneuvers in order to shape the gravity assist. This shaping was done to help meet some mission goals. In particular, the gravity assist was changed slightly in order to remove lunar shadows in both the cruise out to L2 and in the first revolution about L2. This amounted to a change in the phasing loop AV of less than 1 m/s. After the gravity assist, two mid-course correction (MCC) maneuvers were performed in order to fine-tune the trajectory. MCC1 was used to clean up and errors which resulted from the gravity assist. MCC2 was performed in order to mitigate a large stationkeeping maneuver following a crucial instrument calibration period during the cruise phase. MAP executed it's first stationkeeping maneuver in January 16th and is ready for a second calibration period during late Winter / early Spring. Further information concerning subsequent stationkeeping maneuver will be added as they become available.

Cuevas, Osvaldo O.↗
NASA NTRS2009

Trading Robustness Requirements in Mars Entry Trajectory Design

One of the most important metrics characterizing an atmospheric entry trajectory in preliminary design is the size of its predicted landing ellipse. Often, requirements for this ellipse are set early in design and significantly influence both the expected scientific return from a particular mission and the cost of development. Requirements typically specify a certain probability level (6-level) for the prescribed ellipse, and frequently this latter requirement is taken at 36. However, searches for the justification of 36 as a robustness requirement suggest it is an empirical rule of thumb borrowed from non-aerospace fields. This paper presents an investigation into the sensitivity of trajectory performance to varying robustness (6-level) requirements. The treatment of robustness as a distinct objective is discussed, and an analysis framework is presented involving the manipulation of design variables to effect trades between performance and robustness objectives. The scenario for which this method is illustrated is the ballistic entry of an MSL-class Mars entry vehicle. Here, the design variable is entry flight path angle, and objectives are parachute deploy altitude performance and error ellipse robustness. Resulting plots show the sensitivities between these objectives and trends in the entry flight path angles required to design to these objectives. Relevance to the trajectory designer is discussed, as are potential steps for further development and use of this type of analysis.

Lafleur, Jarret M.↗
NASA NTRS2011

Applications of Multi-body Dynamical Environments: The ARTEMIS Transfer Trajectory Design

The application of forces in multi-body dynamical environments to permit the transfer of spacecraft from Earth orbit to Sun-Earth weak stability regions and then return to the Earth-Moon libration (L1 and L2) orbits has been successfully accomplished for the first time. This demonstrated that transfer is a positive step in the realization of a design process that can be used to transfer spacecraft with minimal Delta-V expenditures. Initialized using gravity assists to overcome fuel constraints; the ARTEMIS trajectory design has successfully placed two spacecrafts into Earth-Moon libration orbits by means of these applications.

Libration Orbits↗
NASA NTRS2010

Applications of Multi-Body Dynamical Environments: The ARTEMIS Transfer Trajectory Design

The application of forces in multi-body dynamical environments to pennit the transfer of spacecraft from Earth orbit to Sun-Earth weak stability regions and then return to the Earth-Moon libration (L1 and L2) orbits has been successfully accomplished for the first time. This demonstrated transfer is a positive step in the realization of a design process that can be used to transfer spacecraft with minimal Delta-V expenditures. Initialized using gravity assists to overcome fuel constraints; the ARTEMIS trajectory design has successfully placed two spacecraft into EarthMoon libration orbits by means of these applications.

Folta, David C.↗
NASA NTRS2016

Monte Carlo Analysis as a Trajectory Design Driver for the TESS Mission

The Transiting Exoplanet Survey Satellite (TESS) will be injected into a highly eccentric Earth orbit and fly 3.5 phasing loops followed by a lunar flyby to enter a mission orbit with lunar 2:1 resonance. Through the phasing loops and mission orbit, the trajectory is significantly affected by lunar and solar gravity. We have developed a trajectory design to achieve the mission orbit and meet mission constraints, including eclipse avoidance and a 30-year geostationary orbit avoidance requirement. A parallelized Monte Carlo simulation was performed to validate the trajectory after injecting common perturbations, including launch dispersions, orbit determination errors, and maneuver execution errors. The Monte Carlo analysis helped identify mission risks and is used in the trajectory selection process.

trajectory↗
NASA NTRS2006

Innovative Trajectory Designs to meet Exploration Challenges

This document is a viewgraph presentation of the conference paper. Missions incorporated into NASA's Vision for Space Exploration include many different destinations and regions; are challenging to plan; and need new and innovative trajectory design methods to enable them. By combining proven methods with chaos dynamics, exploration goals that require maximum payload mass or minimum duration can be achieved. The implementation of these innovative methods, such as weak stability boundaries, has altered NASA's approach to meet exploration challenges and is described to show how exploration goals may be met in the next decade. With knowledge that various perturbations play a significant role, the mission designer must rely on both traditional design strategies as well as these innovative methods. Over the past decades, improvements have been made that would at first glance seem dramatic. This paper provides a brief narrative on how a fundamental shift has occurred and how chaos dynamics improve the design of exploration missions with complex constraints.

Folta, David C.↗
NASA NTRS2011

Trajectory design considerations for small-body touch-and-go

"Touch-and-Go," or TAG, is an approach to small-body surface interrogation missions in which the spacecraft descends to the surface, remains in contact for a short time, and then ascends without coming to rest. Appropriate trajectory design solutions to support TAG missions vary widely based on the spacecraft dynamics, small-body environment, spacecraft and ground systems capabilities, and mission objectives. This paper discusses various factors that are considered during the process of developing a TAG mission trajectory and presents a few case study examples to demonstrate how TAG trajectories may vary from mission to mission.

Cangahuala, Alberto↗
NASA NTRS2015

Integrated Vehicle and Trajectory Design of Small Spacecraft with Electric Propulsion for Earth and Interplanetary Missions

This paper investigates the feasibility of Earth-transfer and interplanetary mission architectures for miniaturized spacecraft using emerging small solar electric propulsion technologies. Emerging small SEP thrusters offer significant advantages relative to existing technologies and will enable U-class systems to perform trajectory maneuvers with significant Delta V requirements. The approach in this paper is unique because it integrates trajectory design with vehicle sizing and accounts for the system and operational constraints of small U-class missions. The modeling framework includes integrated propulsion, orbit, energy, and external environment dynamics and systems-level power, energy, mass, and volume constraints. The trajectory simulation environment models orbit boosts in Earth orbit and flyby and capture trajectories to interplanetary destinations. A family of small spacecraft mission architectures are studied, including altitude and inclination transfers in Earth orbit and trajectories that escape Earth orbit and travel to interplanetary destinations such as Mercury, Venus, and Mars. Results are presented visually to show the trade-offs between competing performance objectives such as maximizing available mass and volume for payloads and minimizing transfer time. The results demonstrate the feasibility of using small spacecraft to perform significant Earth and interplanetary orbit transfers in less than one year with reasonable U-class mass, power, volume, and mission durations.

Spangelo, Sara↗
NASA NTRS2017

Trajectory Design Considerations for Exploration Mission 1

Exploration Mission 1 (EM-1) will be the first mission to send an uncrewed Orion vehicle to cislunar space in 2018, targeted to a Distant Retrograde Orbit (DRO). Analysis of EM-1 DRO mission opportunities in 2018 help characterize mission parameters that are of interest to other subsystems (e.g., power, thermal, communications, flight operations, etc). Subsystems request mission design trades which include: landing lighting, addition of an Orion main engine checkout burn, and use of auxiliary thruster only cases. This paper examines the evolving trade studies that incorporate subsystem feedback and demonstrate the feasibility of these constrained mission trajectory designs and contingencies.

Dawn, Timothy F.↗
NASA NTRS2011

Trajectory Design Considerations for Small Body Touch-and-Go

Outline: (1) Trajectory Description (2) Design Drivers: (2a) Dynamics (2b) Environment (2c) Spacecraft and Ground and System Capabilities (2d) Mission Objectives (3) Design Choices (4) Historical Precedents (5) Case Studies. What is Touch-and-Go (TAG)? (1) Descent to the surface (2) Brief contact (3) Ascends to a safe distance

comets↗
NASA NTRS2017

Trajectory Design for the Lunar Polar Hydrogen Mapper Mission

The presented trajectory was designed for the Lunar Polar Hydrogen Mapper (LunaH-Map) 6U CubeSat, which was awarded a ride on NASAs Space Launch System (SLS) with Exploration Mission 1 (EM-1) via NASAs 2015 SIMPLEX proposal call. After deployment from EM-1s upper stage (which is planned to enter heliocentric space via a lunar flyby), the LunaH-Map CubeSat will alter its trajectory via its low-thrust ion engine to target a lunar flyby that yields a Sun-Earth-Moon weak stability boundary transfer to set up a ballistic lunar capture. Finally, the orbit energy is lowered to reach the required quasi-frozen science orbit with periselene above the lunar south pole.

lunar polar orbit↗
NASA NTRS2019

Hybrid Transportation System Integrated Trajectory Design and Optimization for Mars Landing Site Accessibility

NASA’s Mars Study Capability Team continues the agency’s efforts to study and refine the nation’s plan to field a sustainable human Mars campaign. One of the primary open issues left unanswered during previous analysis cycles was the uncertainty of the impact to the vehicle performance requirement to deliver crew and cargo to the same landing site across multiple mission opportunities. The Mars Study Capability Team has recently developed an integrated trajectory optimization and system closure model to solve the complex interplanetary trajectory optimization using both low-thrust and high-thrust maneuvers. This paper demonstrates the capability of this new integrated trajectory design and optimization method as it applies to the landing site accessibility problem for the Hybrid transportation architecture. The results showed that the current vehicle is capable of reaching up to +20 degree latitude and down to -20 degree latitude in every mission opportunity from 2033 to 2052. However, reaching latitudes beyond +/- 20 required more propellant than the spacecraft is currently designed to carry for many of the mission opportunities.

Chai, Patrick R.↗