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Results for “low-thrust trajectory design”

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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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Space shuttle rendezous, radiation and reentry analysis code

A preliminary space shuttle mission design and analysis tool is reported emphasizing versatility, flexibility, and user interaction through the use of a relatively small computer (IBM-7044). The Space Shuttle Rendezvous, Radiation and Reentry Analysis Code is used to perform mission and space radiation environmental analyses for four typical space shuttle missions. Included also is a version of the proposed Apollo/Soyuz rendezvous and docking test mission. Tangential steering circle to circle low-thrust tug orbit raising and the effects of the trapped radiation environment on trajectory shaping due to solar electric power losses are also features of this mission analysis code. The computational results include a parametric study on single impulse versus double impulse deorbiting for relatively low space shuttle orbits as well as some definitive data on the magnetically trapped protons and electrons encountered on a particular mission.

Mcglathery, D. M.↗
NASA NTRS1977

Navigation capability for an ion drive rendezvous with Halley's Comet

An analysis has been conducted in connection with plans for a study of Halley's Comet during its 1986 apparition. The use of low-thrust vehicles, utilizing an ion drive system, is being considered for a comet rendezvous mission. A preliminary trajectory for the Halley rendezvous mission calls for launch on June 20, 1982, followed by rendezvous on December 21, 1985. The navigation analysis described focuses on the terminal approach to Halley, the 60-day period preceding rendezvous. Navigation analysis assumptions are examined, taking into account navigation error sources, radio tracking, onboard optical data, earth-based comet observations, and orbit determination and guidance strategies. The preliminary mission design considers a rendezvous at approximately 56,000 km from the comet nucleus (6,000 km outside the dust envelope). Navigation performance is measured in terms of comet-relative position and velocity errors at encounter. Variations to the baseline navigation study provide illustrations concerning the close link between delivery accuracy and stochastic thrust errors.

Thornton, C. L.↗
NASA NTRS2016

Low-Thrust Transfers from Distant Retrograde Orbits to L2 Halo Orbits in the Earth-Moon System

This paper presents a study of transfers between distant retrograde orbits (DROs) and L2 halo orbits in the Earth-Moon system that could be flown by a spacecraft with solar electric propulsion (SEP). Two collocation-based optimal control methods are used to optimize these highly-nonlinear transfers: Legendre pseudospectral and Hermite-Simpson. Transfers between DROs and halo orbits using low-thrust propulsion have not been studied previously. This paper offers a study of several families of trajectories, parameterized by the number of orbital revolutions in a synodic frame. Even with a poor initial guess, a method is described to reliably generate families of solutions. The circular restricted 3-body problem (CRTBP) is used throughout the paper so that the results are autonomous and simpler to understand.

trajectory design↗
NASA NTRS2002

Solar Electric Propulsion Technologies Being Designed for Orbit Transfer Vehicle Applications

There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1. Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space telescopes, and planetary mission staging. In these scenarios, an SEP stage, or "tug," is used to transport payloads to L1--such as the habitat module, lunar excursion and return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics, lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity trajectory model to refine the analysis.

Sarver-Verhey, Timothy R.↗
NASA NTRS2016

Low-Thrust Transfers from Distant Retrograde Orbits to L2 Halo Orbits in the Earth-Moon System

Enable future missions Any mission to a DRO or halo orbit could benefit from the capability to transfer between these orbits Chemical propulsion could be used for these transfers, but at high propellant cost Fill gaps in knowledge A variety of transfers using SEP or solar sails have been studied for the Earth-Moon system Most results in literature study a single transfer This is a step toward understanding the wide array of types of transfers available in an N-body force model.

optimization↗
NASA NTRS1973

Low-thrust mission risk analysis, with application to a 1980 rendezvous with the comet Encke

A computerized failure process simulation procedure is used to evaluate the risk in a solar electric space mission. The procedure uses currently available thrust-subsystem reliability data and performs approximate simulations of the thrust sybsystem burn operation, the system failure processes, and the retargeting operations. The method is applied to assess the risks in carrying out a 1980 rendezvous mission to the comet Encke. Analysis of the results and evaluation of the effects of various risk factors on the mission show that system component failure rates are the limiting factors in attaining a high mission relability. It is also shown that a well-designed trajectory and system operation mode can be used effectively to partially compensate for unreliable thruster performance.

Yen, C. L.↗
NASA NTRS1973

Low-thrust mission risk analysis.

A computerized multi-stage failure process simulation procedure is used to evaluate the risk in a solar electric space mission. The procedure uses currently available thrust-subsystem reliability data and performs approximate simulations of the thrust subsystem burn operation, the system failure processes, and the retargetting operations. The application of the method is used to assess the risks in carrying out a 1980 rendezvous mission to Comet Encke. Analysis of the results and evaluation of the effects of various risk factors on the mission show that system component failure rates is the limiting factor in attaining a high mission reliability. But it is also shown that a well-designed trajectory and system operation mode can be used effectively to partially compensate for unreliable thruster performance.

Yen, C. L.↗
NASA NTRS2004

Refinement and Use of the OTIS Program for RLV and Low-Thrust Mission Design

NASA Glenn's Systems Analysis Branch (SAB) is involved in mission analysis and associated software tool development for Earth-to-Orbit and In-Space (low thrust) missions. One of the key software analysis programs used by the SAB is the Optimal Trajectories by Implicit Simulation (OTIS) program. This program is developed and managed by NASA Glenn s SAB and is used on a national level by government and contractor affiliates. The SAB is also involved in the mission design activities for the Next Generation Launch Technology (NGLT) program as well as In-Space mission design involving new In-Space propulsion technologies. It is in these two main areas that support, in-depth studies, and development were performed. Below is a brief summary of the study and development activities.

Source record↗
NASA NTRS2002

A review of some exact solutions to the planar equations of motion of a thrusting spacecraft

With the complexities in computing optimal low thrust trajectories, easily-computed, good sub-optimal trajectories provide both a practical alternative for mission designers and a starting point for optimisation. The present paper collects in one place for easy reference and comparison several exact solutions that have been obtained in the literature over the last few decades: the logarithmic spiral, Pinkham's variant thereof, Forbes spiral, the exponential sinusoid, the case of constant radial thrust, Markopoulos's Keplerian thrust arcs, Lawden's spiral, and the analogous Bishop and Azimov spiral.

low-thrust trajectories electric propulsion solar ↗
NASA NTRS2003

An investigation of a Jupiter Galilean Moon Orbiter trajectory

NASA's proposed mission to send a single electric propulsion spacecraft to orbit Callisto, Ganymeds, and Europa within a decade would require a very complex trajectory. Strong multi-body effects combined with low-thrust control of capture and escape will make the trajectory challenging. This paper describes an optimal trajectory that begins in low Earth orbit and ends in low Io orbit.

Galilean Moon Orbiter↗
NASA NTRS2006

Comparison of Performance Predictions for New Low-Thrust Trajectory Tools

Several low thrust trajectory optimization tools have been developed over the last 3% years by the Low Thrust Trajectory Tools development team. This toolset includes both low-medium fidelity and high fidelity tools which allow the analyst to quickly research a wide mission trade space and perform advanced mission design. These tools were tested using a set of reference trajectories that exercised each tool s unique capabilities. This paper compares the performance predictions of the various tools against several of the reference trajectories. The intent is to verify agreement between the high fidelity tools and to quantify the performance prediction differences between tools of different fidelity levels.

Polsgrove, Tara↗
NASA NTRS2007

Efficient Optimization of Low-Thrust Spacecraft Trajectories

A paper describes a computationally efficient method of optimizing trajectories of spacecraft driven by propulsion systems that generate low thrusts and, hence, must be operated for long times. A common goal in trajectory-optimization problems is to find minimum-time, minimum-fuel, or Pareto-optimal trajectories (here, Pareto-optimality signifies that no other solutions are superior with respect to both flight time and fuel consumption). The present method utilizes genetic and simulated-annealing algorithms to search for globally Pareto-optimal solutions. These algorithms are implemented in parallel form to reduce computation time. These algorithms are coupled with either of two traditional trajectory- design approaches called "direct" and "indirect." In the direct approach, thrust control is discretized in either arc time or arc length, and the resulting discrete thrust vectors are optimized. The indirect approach involves the primer-vector theory (introduced in 1963), in which the thrust control problem is transformed into a co-state control problem and the initial values of the co-state vector are optimized. In application to two example orbit-transfer problems, this method was found to generate solutions comparable to those of other state-of-the-art trajectory-optimization methods while requiring much less computation time.

Lee, Seungwon↗
NASA NTRS2015

An Automatic Medium to High Fidelity Low-Thrust Global Trajectory Toolchain; EMTG-GMAT

Solving the global optimization, low-thrust, multiple-flyby interplanetary trajectory problem with high-fidelity dynamical models requires an unreasonable amount of computational resources. A better approach, and one that is demonstrated in this paper, is a multi-step process whereby the solution of the aforementioned problem is solved at a lower-fidelity and this solution is used as an initial guess for a higher-fidelity solver. The framework presented in this work uses two tools developed by NASA Goddard Space Flight Center: the Evolutionary Mission Trajectory Generator (EMTG) and the General Mission Analysis Tool (GMAT). EMTG is a medium to medium-high fidelity low-thrust interplanetary global optimization solver, which now has the capability to automatically generate GMAT script files for seeding a high-fidelity solution using GMAT's local optimization capabilities. A discussion of the dynamical models as well as thruster and power modeling for both EMTG and GMAT are given in this paper. Current capabilities are demonstrated with examples that highlight the toolchains ability to efficiently solve the difficult low-thrust global optimization problem with little human intervention.

low thrust↗
NASA NTRS2013

Thrust Direction Optimization: Satisfying Dawn's Attitude Agility Constraints

The science objective of NASA's Dawn Discovery mission is to explore the two largest members of the main asteroid belt, the giant asteroid Vesta and the dwarf planet Ceres. Dawn successfully completed its orbital mission at Vesta. The Dawn spacecraft has complex, difficult to quantify, and in some cases severe limitations on its attitude agility. The low-thrust transfers between science orbits at Vesta required very complex time varying thrust directions due to the strong and complex gravity and various science objectives. Traditional thrust design objectives (like minimum (Delta)V or minimum transfer time) often result in thrust direction time evolutions that can not be accommodated with the attitude control system available on Dawn. This paper presents several new optimal control objectives, collectively called thrust direction optimization that were developed and necessary to successfully navigate Dawn through all orbital transfers at Vesta.

trajectory design↗
NASA NTRS2017

Low-Thrust Trajectory Optimization with Simplified SQP Algorithm

The problem of low-thrust trajectory optimization in highly perturbed dynamics is a stressing case for many optimization tools. Highly nonlinear dynamics and continuous thrust are each, separately, non-trivial problems in the field of optimal control, and when combined, the problem is even more difficult. This paper de-scribes a fast, robust method to design a trajectory in the CRTBP (circular restricted three body problem), beginning with no or very little knowledge of the system. The approach is inspired by the SQP (sequential quadratic programming) algorithm, in which a general nonlinear programming problem is solved via a sequence of quadratic problems. A few key simplifications make the algorithm presented fast and robust to initial guess: a quadratic cost function, neglecting the line search step when the solution is known to be far away, judicious use of end-point constraints, and mesh refinement on multiple shooting with fixed-step integration.In comparison to the traditional approach of plugging the problem into a “black-box” NLP solver, the methods shown converge even when given no knowledge of the solution at all. It was found that the only piece of information that the user needs to provide is a rough guess for the time of flight, as the transfer time guess will dictate which set of local solutions the algorithm could converge on. This robustness to initial guess is a compelling feature, as three-body orbit transfers are challenging to design with intuition alone. Of course, if a high-quality initial guess is available, the methods shown are still valid.We have shown that endpoints can be efficiently constrained to lie on 3-body repeating orbits, and that time of flight can be optimized as well. When optimizing the endpoints, we must make a trade between converging quickly on sub-optimal endpoints or converging more slowly on end-points that are arbitrarily close to optimal. It is easy for the mission design engineer to adjust this trade based on the problem at hand.The biggest limitation to the algorithm at this point is that multi-revolution transfers (greater than 2 revolutions) do not work nearly as well. This restriction comes in because the relationship between node 1 and node N becomes increasingly nonlinear as the angular distance grows. Trans-fers with more than about 1.5 complete revolutions generally require the line search to improve convergence. Future work includes: Comparison of this algorithm with other established tools; improvements to how multiple-revolution transfers are handled; parallelization of the Jacobian computation; in-creased efficiency for the line search; and optimization of many more trajectories between a variety of 3-body orbits.

Parrish, Nathan L.↗
NASA NTRS2017

Low Thrust Mission Trajectories to Near Earth Asteroids

The discovery of 2016 HO3 and its classification as a quasi-satellite has sparked a stronger interest towards Near Earth Asteroids (NEAs). This work presents low-thrust low-power mission designs to various NEAs using an EELV Secondary Payload Adapter (ESPA). A global trajectory optimizer (EMTG) was used to generate mission solutions to a select 13 NEAs using a 200 watt BHT-200 thruster as a proof of concept. The missions presented here demonstrate that a low-cost electric propulsion ESPA mission to NEAs is a feasible concept for many asteroids.

Asteroid↗
NASA NTRS2016

Spiral Lamber's Problem with Generalized Logarithmic Spirals

Lambert’s problem subject to a continuous acceleration is solved using the family of generalized logarithmic spirals. Thanks to the existence of two first integrals related to the energy and angular momentum surprising analogies with the Keplerian case are found. A minimum-energy spiral transfer exists. Increasing the value of the constant of the generalized energy yields pairs of conjugate spiral trajectories. The properties of such spirals are strongly connected with the properties of conjugate Keplerian orbits. When the generalized constant of the energy reaches a critical value the two solutions degenerate into a pair of parabolic spirals, one of which connects the two vectors through infinity. From that point the spiral transfers become hyperbolic. Generalized logarithmic spirals admit closed-form solutions to all the required magnitudes including the time of flight, providing a deep insight into the dynamics of the problem. In addition, the maximum acceleration along the transfer is found analytically so the solutions that violate the design constraints on the maximum thrust acceleration can be rejected without any further computations. When the time of flight is fixed there is still a degree of freedom in the solution, related to a control parameter. Resonant transfers appear naturally thanks to the symmetry properties of the generalized logarithmic spirals. The problem of designing a low-thrust transfer between two bodies can be reduced to solving the corresponding spiral Lambert’s problem. In order to show the versatility of the method it is applied to the design of an asteroid tour and to explore launch opportunities to Mars.\

Roa, Javier↗