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Results for “Low Thrust Trajectory”

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 145 records · Page 8
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NASA NTRS1977

Optimal low-thrust takeoff from an orbit about an oblate planet

Future space missions to the outer planets may depend upon the use of low-thrust propulsion systems. As these planets are decidedly oblate, the question of the effect of that oblateness on a low-thrust trajectory is of some interest. In this paper the problem of optimal energy increase is attacked under the assumption that the coefficients for the second zonal harmonic, and the nondimensional thrust acceleration are the same order of magnitude. By means of a two-variable asymptotic expansion technique, a near optimal control program is generated and the first-order uniformly valid approximation for the corresponding trajectory is obtained. Tangential thrust is shown to be a good near-optimal thrust program even in the presence of oblateness effects. The optimal control program is found to be oscillatory and quite similar to the optimal control for energy increase in an inverse square gravitational field.

Jacobson, R. A.↗
NASA NTRS1992

Comet nucleus and asteroid sample return missions

During the 1991-92 academic year, the Pennsylvania State University has developed three sample return missions: one to the nucleus of comet Wild 2, one to the asteroid Eros, and one to three asteroids located in the Main Belt. The primary objective of the comet nucleus sample return mission is to rendezvous with a short period comet and acquire a 10 kg sample for return to Earth. Upon rendezvous with the comet, a tethered coring and sampler drill will contact the surface and extract a two-meter core sample from the target site. Before the spacecraft returns to Earth, a monitoring penetrator containing scientific instruments will be deployed for gathering long-term data about the comet. A single asteroid sample return mission to the asteroid 433 Eros (chosen for proximity and launch opportunities) will extract a sample from the asteroid surface for return to Earth. To limit overall mission cost, most of the mission design uses current technologies, except the sampler drill design. The multiple asteroid sample return mission could best be characterized through its use of future technology including an optical communications system, a nuclear power reactor, and a low-thrust propulsion system. A low-thrust trajectory optimization code (QuickTop 2) obtained from the NASA LeRC helped in planning the size of major subsystem components, as well as the trajectory between targets.

Melton, Robert G.↗
NASA NTRS2004

Chemical and Solar Electric Propulsion Systems Analyses for Mars Sample Return Missions

Conceptual in-space transfer stages, including those utilizing solar electric propulsion, chemical propulsion, and chemical propulsion with aerobraking or aerocapture assist at Mars, were evaluated. Roundtrip Mars sample return mission vehicles were analyzed to determine how specific system technology selections influence payload delivery capability. Results show how specific engine, thruster, propellant, capture mode, trip time and launch vehicle technology choices would contribute to increasing payload or decreasing the size of the required launch vehicles. Heliocentric low-thrust trajectory analyses for Solar Electric Transfer were generated with the SEPTOP code.

Donahue, Benjamin B.↗
NASA NTRS2006

Performance of Solar Electric Powered Deep Space Missions Using Hall Thruster Propulsion

Power limited, low-thrust trajectories were assessed for missions to Jupiter, Saturn, and Neptune utilizing a single Venus Gravity Assist (VGA) and a primary propulsion system based on either a 3-kW high voltage Hall thruster, of the type being developed by the NASA In-Space Propulsion Technology Program, or an 8-kW variant of this thruster. These Hall thrusters operate with specific impulses below 3,000 seconds. A trade study was conducted to examine mission parameters that include: net delivered mass (NDM), beginning-of-life (BOL) solar array power, heliocentric transfer time, required launch vehicle, number of operating thrusters, and throttle profile. The top performing spacecraft configuration was defined to be the one that delivered the highest mass for a range of transfer times. In order to evaluate the potential future benefit of using next generation Hall thrusters as the primary propulsion system, comparisons were made with the advanced state-of-the-art (ASOA), 7-kW, 4,100 second NASA's Evolutionary Xenon Thruster (NEXT) for the same mission scenarios. For the BOL array powers considered in this study (less than 30 kW), the results show that the performance of the Hall thrusters, relative to NEXT, is largely dependant on the performance capability of the launch vehicle, and that at least a 10 percent performance gain, equating to at least an additional 200 kg dry mass at each target planet, is achieved over the higher specific impulse NEXT when launched on an Atlas 551.

Witzberger, Kevin E.↗
NASA NTRS1998

Negative C3 Launch Options for Solar System Exploration

Low thrust trajectory analyses were used to examine the feasibility of using solar electronic propulsion for Earth escape from a negative C3 launch for deep space missions in order to significantly increase the net delivered mass capability of inexpensive launch vehicles.

negative C3 Solar System Exploration Earth mass co↗
NASA NTRS2006

A Quick Method for Estimating Vehicle Characteristics Appropriate for Continuous Thrust Round Trip Missions Within the Solar System

A simple method of estimating vehicle parameters appropriate for interplanetary travel can provide a useful tool for evaluating the suitability of particular propulsion systems to various space missions. Although detailed mission analyses for interplanetary travel can be quite complex, it is possible to derive hirly simple correlations which will provide reasonable trip time estimates to the planets. In the present work, it is assumed that a constant thrust propulsion system propels a spacecraft on a round trip mission having equidistant outbound and inbound legs in which the spacecraft accelerates during the first portion of each leg of the journey and decelerates during the last portion of each leg of the journey. Comparisons are made with numerical calculations from low thrust trajectory codes to estimate the range of applicability of the simplified correlations.

Emrich, Bill↗
NASA NTRS2008

Chasing a Comet with a Solar Sail

Solar sail propulsion systems enable a wide range of missions that require constant thrust or high delta-V over long mission times. One particularly challenging mission type is a comet rendezvous mission. This paper presents optimal low-thrust trajectory designs for a range of sailcraft performance metrics and mission transit times that enables a comet rendezvous mission. These optimal trajectory results provide a trade space which can be parameterized in terms of mission duration and sailcraft performance parameters such that a design space for a small satellite comet chaser mission is identified. These results show that a feasible space exists for a small satellite to perform a comet chaser mission in a reasonable mission time.

Stough, Robert W.↗
NASA NTRS2008

Genetic Algorithm Optimizes Q-LAW Control Parameters

A document discusses a multi-objective, genetic algorithm designed to optimize Lyapunov feedback control law (Q-law) parameters in order to efficiently find Pareto-optimal solutions for low-thrust trajectories for electronic propulsion systems. These would be propellant-optimal solutions for a given flight time, or flight time optimal solutions for a given propellant requirement. The approximate solutions are used as good initial solutions for high-fidelity optimization tools. When the good initial solutions are used, the high-fidelity optimization tools quickly converge to a locally optimal solution near the initial solution. Q-law control parameters are represented as real-valued genes in the genetic algorithm. The performances of the Q-law control parameters are evaluated in the multi-objective space (flight time vs. propellant mass) and sorted by the non-dominated sorting method that assigns a better fitness value to the solutions that are dominated by a fewer number of other solutions. With the ranking result, the genetic algorithm encourages the solutions with higher fitness values to participate in the reproduction process, improving the solutions in the evolution process. The population of solutions converges to the Pareto front that is permitted within the Q-law control parameter space.

Lee, Seungwon↗
NASA NTRS2012

Mars Sample Return Using Solar Sail Propulsion

Many Mars Sample Return (MSR) architecture studies have been conducted over the years. A key element of them is the Earth Return Stage (ERS) whose objective is to obtain the sample from the Mars Ascent Vehicle (MAV) and return it safely to the surface of the Earth. ERS designs predominantly use chemical propulsion [1], incurring a significant launch mass penalty due to the low specific impulse of such systems coupled with the launch mass sensitivity to returned mass. It is proposed to use solar sail propulsion for the ERS, providing a high (effective) specific impulse propulsion system in the final stage of the multi-stage system. By doing so to the launch mass of the orbiter mission can be significantly reduced and hence potentially decreasing mission cost. Further, solar sailing offers a unique set of non-Keplerian low thrust trajectories that may enable modifications to the current approach to designing the Earth Entry Vehicle by potentially reducing the Earth arrival velocity. This modification will further decrease the mass of the orbiter system. Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like surface made of a lightweight, reflective material. The continuous photonic pressure provides propellantless thrust to conduct orbital maneuvering and plane changes more efficiently than conventional chemical propulsion. Because the Sun supplies the necessary propulsive energy, solar sails require no onboard propellant, thus reducing system mass. This technology is currently at TRL 7/8 as demonstrated by the 2010 flight of the Japanese Aerospace Exploration Agency, JAXA, IKAROS mission. [2]

Johnson, Les↗
NASA NTRS2011

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Near Earth Object (NEO)↗
NASA NTRS2016

Hybrid Differential Dynamic Programming with Stochastic Search

Differential dynamic programming (DDP) has been demonstrated as a viable approach to low-thrust trajectory optimization, namely with the recent success of NASA's Dawn mission. The Dawn trajectory was designed with the DDP-based Static/Dynamic Optimal Control algorithm used in the Mystic software.1 Another recently developed method, Hybrid Differential Dynamic Programming (HDDP),2, 3 is a variant of the standard DDP formulation that leverages both first-order and second-order state transition matrices in addition to nonlinear programming (NLP) techniques. Areas of improvement over standard DDP include constraint handling, convergence properties, continuous dynamics, and multi-phase capability. DDP is a gradient based method and will converge to a solution nearby an initial guess. In this study, monotonic basin hopping (MBH) is employed as a stochastic search method to overcome this limitation, by augmenting the HDDP algorithm for a wider search of the solution space.

Aziz, Jonathan↗
NASA NTRS2016

Hybrid Differential Dynamic Programming with Stochastic Search

Differential dynamic programming (DDP) has been demonstrated as a viable approach to low-thrust trajectory optimization, namely with the recent success of NASAs Dawn mission. The Dawn trajectory was designed with the DDP-based Static Dynamic Optimal Control algorithm used in the Mystic software. Another recently developed method, Hybrid Differential Dynamic Programming (HDDP) is a variant of the standard DDP formulation that leverages both first-order and second-order state transition matrices in addition to nonlinear programming (NLP) techniques. Areas of improvement over standard DDP include constraint handling, convergence properties, continuous dynamics, and multi-phase capability. DDP is a gradient based method and will converge to a solution nearby an initial guess. In this study, monotonic basin hopping (MBH) is employed as a stochastic search method to overcome this limitation, by augmenting the HDDP algorithm for a wider search of the solution space.

Aziz, Jonathan↗
NASA NTRS2018

Analysis of Near Rectilinear Halo Orbit Insertion with a 40-kW Solar Electric Propulsion System

This paper examines low thrust trajectories for delivery of a 40-kW solar electric propulsion spacecraft and potential additional payload to a desired NRHO. One option considered is a trans-lunar injection launch as a co-manifested payload on the Space Launch System. For this option, a reference trajectory is designed and a scan of launch dates is completed to understand the propellant mass sensitivity. A 15-day period cyclical variation in required propellant is observed that is attributed to solar gravity effects. A second option considered is to launch on a smaller commercial launch vehicle to a less energetic elliptical orbit and use SEP to spiral out to NRHO. For this option, analysis is completed to understand the trades between delivered mass to NRHO, total propellant required, time of flight, and solar array degradation. Results show that, while launching to lower altitudes can deliver greater payload mass to NRHO, significant solar array degradation can be observed.

Mccarty, Steven L.↗
NASA NTRSDate not supplied

Integrated Framework to Enable Design Space Exploration of In-Space Transportation Architectures

Even with the recent shift in focus of NASA’s human spaceflight program towards the Moon, the long-term goal continues to be crewed missions to Mars. Regardless of the accepted architecture, in-space transportation systems are a critical portion to achieving this goal. An emphasis is placed on presenting decision makers with many options early on to make an informed down-selection. Although current processes are able to address any aspects of the challenging underlying multi-disciplinary analysis and optimization problem of these advanced concepts, they are unable to comprehensively perform design space exploration inexpensively,potentially leaving attractive alternatives on the table. A design framework is proposed that is capable of enabling the integrated mission analysis, vehicle sizing and synthesis problem for in-space transportation architectures. This framework is enabled through the use of the vehicle synthesis tool, Dynamic Rocket Equation Tool, various subsystem sizing models, and low-thrust trajectory optimization codes. As a case study to demonstrate the framework, two design points of a hybrid propulsion stage concept, studied by Georgia Tech’s 2018 Revolutionary Aerospace Systems Concepts-Academic Linkage team, is explored.

Akshay Prasad↗
NASA NTRSDate not supplied

Integrated Framework to Enable Design Space Exploration of In-Space Transportation Architectures

Even with the recent shift in focus of NASA’s human spaceflight program towards the Moon, the long-term goal continues to be crewed missions to Mars. Regardless of the accepted architecture, in-space transportation systems are a critical portion to achieving this goal. An emphasis is placed on presenting decision makers with many options early on to make an informed down-selection. Although current processes are able to address any aspects of the challenging underlying multi-disciplinary analysis and optimization problem of these advanced concepts, they are unable to comprehensively perform design space exploration inexpensively, potentially leaving attractive alternatives on the table. A design framework is proposed that is capable of enabling the integrated mission analysis, vehicle sizing and synthesis problem for in-space transportation architectures. This framework is enabled through the use of the vehicle synthesis tool, Dynamic Rocket Equation Tool, various subsystem sizing models, and low-thrust trajectory optimization codes. As a case study to demonstrate the framework, two design points of a hybrid propulsion stage concept, studied by Georgia Tech’s 2018 Revolutionary Aerospace Systems Concepts-Academic Linkage team, is explored.

Akshay Prasad↗
NASA NTRSDate not supplied

Analytical Partial Derivatives of the Q-Law Guidance Algorithm

One method of producing low-thrust trajectories is the closed-loop Q-Law guidance algorithm. This work wraps Q-Law inside a nonlinear programming problem to search for locally optimal gain combinations and enforce nonlinear constraints on the initial state. Gradient-based trajectory optimization has been shown to benefit greatly when analytical partial derivatives are supplied to the optimizer. This paper presents the Q-Law optimization problem setup and derives the Q-Law thrust vector partial derivatives for the State Transition Matrix required by standard nonlinear optimization software.

Jackson L Shannon↗
NASA NTRS2017

Characteristics of a high-power ion beam deflection system necessary to deflect the hypothetical asteroid 2017 PD

The July 2027 impact date for the hypothetical asteroid 2017 PDC, that is the subject of an emergency response exercise, leaves just over ten years to implement a deflection approach. The analyses herein allocates four years to the design, fabrication, assembly, test and launch of a notional high-power Ion Beam Deflection (IBD) vehicle to meet a launch readiness date no later than May 2021. Using this launch date along with estimates for the vehicle mass and performance characteristics of the electric propulsion system, low-thrust trajectory analyses indicate a 2.56-year flight time to rendezvous with 2017 PDC. This would leave 3.6 years to execute the actual deflection phase. A 160-kW IBD vehicle could deflect 2017 PDC by at least one Earth radius within this time provided the asteroid’s actual diameter was less than about 140 m and its density was 2 g/cm3 or less. Larger diameters and/or higher densities would require a higher power IBD vehicle, multiple IBD vehicles, or a longer deflection phase. Ion beam deflection is largely independent of the characteristics of the threat object, but its effectiveness, like all deflection approaches, is sensitive to both the asteroid mass and the time available for deflection. The characteristics of IBD, i.e., large standoff distance between the spacecraft and the asteroid surface, as well as ion beam divergence angles of a few degrees, facilitates the possible simultaneous use of multiple IBD vehicles to improve the performance and robustness of the asteroid deflection. The 65-cm diameter, 20-kW NEXIS ion thruster developed in support of the Jupiter Icy Moons Orbiter mission concept is particularly well suited for application to a high-power IBD system. It has demonstrated an ion beam divergence angle of approximately 2 degrees and operation at 20 kW for over 2,000 hours. Completion of the development and flight qualification of this technology would significantly reduce the risk and time necessary to respond to the discovery of potentially hazardous asteroid in the size range of 50- to 100-m diameter.

Reeves, David↗