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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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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1961

Estimates of Minimum Energy Requirements for Range-Controlled Return of a Nonlifting Satellite from a Circular Orbit

Existing expressions are used to obtain the minimum propellant fraction required for return from a circular orbit as a function of vacuum trajectory range. trajectory are matched to those of the atmospheric trajectory to obtain a complete return from orbit to earth. The results are restricted by the assumptions of (1) impulsive velocity change, (2) nearly circular transfer trajectory, ( 3) spherical earth, atmosphere, and gravitational field, (4) exponential atmospheric density variation with attitude, and (5) a nonrotating atmosphere. The solutions for the parameters of the vacuum Calculations are made t o determine the effects of longitudinal and lateral range on required propeUant fraction and reentry loading for a nonrotating earth and for several orbital altitudes. the single- and two-impulse method of return is made and the results indicate a "trade off" between propellant fraction required and landing- position accuracy. A comparison of An example of a return mission from a polar orbit is discussed where the initial deorbit point is the intersection of the North Pole horizon with the satellite orbit. Some effects of a rotating earth are also considered. It is found that, for each target-orbital-plane longitudinal difference, there exists a target latitude for which the required propellant fraction is a minimum.

Jackson, Charlie M., Jr.↗
NASA NTRS2012

Trajectory Browser Website

The Trajectory Browser is a web-based tool developed at the NASA Ames Research Center to be used for the preliminary assessment of trajectories to small-bodies and planets and for providing relevant launch date, time-of-flight and V requirements. The site hosts a database of transfer trajectories from Earth to asteroids and planets for various types of missions such as rendezvous, sample return or flybys. A search engine allows the user to find trajectories meeting desired constraints on the launch window, mission duration and delta V capability, while a trajectory viewer tool allows the visualization of the heliocentric trajectory and the detailed mission itinerary. The anticipated user base of this tool consists primarily of scientists and engineers designing interplanetary missions in the context of pre-phase A studies, particularly for performing accessibility surveys to large populations of small-bodies. The educational potential of the website is also recognized for academia and the public with regards to trajectory design, a field that has generally been poorly understood by the public. The website is currently hosted on NASA-internal URL http://trajbrowser.arc.nasa.gov/ with plans for a public release as soon as development is complete.

Foster, Cyrus↗
NASA NTRS2018

Low-Thrust Trajectory Design via Direct Transcription Leveraging Structure from the Low-Thrust Restricted Problem

A primary challenge of low-thrust mission design is the development of an initial guess for the state and control history of a trajectory. To address this challenge, one technique assembles dynamical structures, such as periodic orbits and their associated manifolds, into discontinuous chains that are corrected to locally optimal transfer solutions via direct transcription. In this investigation, dynamical structures that leverage low-thrust in a multi-body regime are incorporated into the orbit chain approach to expand the options available for construction of an initial guess and to guide the direct transcription algorithm toward different categories of locally optimal solutions. The properties and structures of the low-thrust model which facilitate orbit chain construction are demonstrated in representative transfer scenarios. Direct transcription is applied to converge upon locally optimal transfer trajectories in both simplified and ephemeris models. Results indicate that low-thrust dynamical structures offer a promising new catalog of options for use in the orbit chain approach to designing optimal low-thrust trajectories.

Grebow, Daniel↗
NASA NTRS1990

Return to the moon - The Lunar Observer mission

The Lunar Observer mission seeks to understand the moon in its global context. Motivation for this project comes from scientific interest and the need for robotic precursor missions in the proposed Space Exploration Initiative. The multiple roles to be performed by this mission mean that the trajectory design process promises to be quite complex and unique. The choice of launch vehicle, launch opportunity and daily launch window are affected by the selection of the translunar transfer trajectory, as well as the arrival conditions desired at lunar orbit insertion. The length of the mission and the tight lunar orbit control requirements determine the design of the circumlunar portions of the Lunar Observer mission.

Cook, Richard A.↗
NASA NTRS1990

Launching lunar missions from Space Station Freedom

The relative orbital motion of Space Station Freedom and the moon places practical constraints on the timing of launch/return transfer trajectories. This paper describes the timing characteristics as well as the Delta-V variations over a representative cycle of launch/return opportunities. On average, the minimum-Delta-V transfer opportunities occur at intervals of 9 days. However, there is a significant nonuniform variation in this timing interval, as well as the minimum stay time at the moon, over the short cycle (51 days) and the long cycle (18.6 years). The advantage of three-impulse transfers for extending the launch window is also described.

Friedlander, Alan↗
NASA NTRS1990

Ballistic trajectory options for manned Mars Missions

Mars Mission profile options and mission requirements data are presented for earth-Mars opposition and conjunction class round-trip flyby and stopover mission opportunities. The opposition-class flyby and sprint mission uses direct transfer trajectories to and on return from Mars. The opposition-class stopover mission employs the gravitational field of Venus to accelerate the space vehicle on either the outbound or inbound leg in order to reduce the propulsion requirement associated with the opposition-class mission. The conjunction-class mission minimizes propulsion requirements by optimizing the stopover time at Mars. Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low earth orbit to one mission profile option and mission opportunity to another.

Young, Archie C.↗
NASA NTRS1986

Mars mission concepts and opportunities

Trajectory and mission requirement data are presented for Earth Mars opposition and conjunction class roundtrip flyby and stopover mission opportunities available between 1997 and 2045. The opposition class flyby mission uses direct transfer trajectories to and on return from Mars. The opposition class stopover mission employs the gravitational field of Venus to accelerate the space vehicle on either the outbound or inbound leg in order to reduce the propulsion requirement associated with the opposition class mission. The conjunction class mission minimizes propulsion requirements by optimizing the stopover time at Mars.

Archie C Young↗
NASA NTRS1971

An analysis and comparison of several trajectory optimization methods

The sensitivities of the convergence characteristics of the methods to initially assumed parameters and trial solution, convergence times, computer logic, and storage requirements are discussed. Numerical comparison of the convergence characteristics is made by considering a minimum time, low thrust, Earth-Mars transfer trajectory. A modified quasi-linearization method reduces convergence time by approximately 70% when compared with the generalized Newton-Raphson method and allows the terminal boundary to be specified by a general function of the problem variables. A uniquely specified and easily determined, time dependent weighting matrix for the gradient techniques accelerates the shaping of the optimal control program and improves the convergence characteristics during the terminal iterations. Convergence envelopes, indicating how sensitive the convergence characteristics are to initially assumed parameters, are plotted for the perturbation and quasi-linearization methods. Several iteration schemes are proposed which increase the size of the convergence envelopes and decrease the sensitivity of the method to initially assumed parameters.

Lewallen, J. M.↗
NASA NTRS2017

Understanding the Sun-Earth Libration Point Orbit Formation Flying Challenges For WFIRST and Starshade

In order to fly an occulter in formation with a telescope at the Sun-Earth L2 (SEL2) Libration Point, one must have a detailed understanding of the dy-namics that govern the restricted three body system. For initial purposes, a linear approximation is satisfactory, but operations will require a high-fidelity modeling tool along with strategic targeting methods in order to be successful. This paper focuses on the challenging dynamics of the transfer trajectories to achieve the relative positioning of two spacecraft to fly in formation at SEL2, in our case, the Wide-Field Infrared Survey Telescope (WFIRST) and a proposed Starshade. By modeling the formation transfers using a high fidelity tool, an accurate V approximation can be made to as-sist with the development of the subsystem design required for a WFIRST and Starshade formation flight mission.

Sun-Earth-Moon↗
NASA NTRS1998

Lunar prospector mission design and trajectory support

The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

Lozier, David↗
NASA NTRS2023

Trajectory Design and Early Mission Operations for the Lunar IceCube Mission

The Lunar IceCube (LIC) mission, a Next Space Technologies for Exploration Partnerships (NSTEP) program selection, was launched as a rideshare onboard Artemis-I on November 16th 2022, and deployed into a high energy lunar flyby trajectory. The final destination of the mission was a polar elliptical lunar orbit from which it could conduct spectroscopy observations of the lunar surface; however, a near rectilinear halo orbit (NRHO) was planned to be used as a staging orbit that divided the lunar transfer and low-thrust spiral phases of the mission. This paper presents the process used to design the LIC transfer trajectory from the high-energy deployment state to a 9:2 synodic resonance NRHO. Additional analyses performed, to assess the critical deployment to lunar flyby phase of the trajectory and to generate recovery trajectories following a loss of contact with the spacecraft, are described as well. Lessons learned from working on the LIC mission are presented to inform the design of similar future CubeSat missions.

Robert E Pritchett↗
NASA NTRS1993

Trajectory design for the Deep Space Program Science Experiment (DSPSE) mission

In 1994, the Deep Space Program Science Experiment (DSPSE) spacecraft will become the first spacecraft to perform, in succession, both a lunar orbiting mission and a deep-space asteroid encounter mission. The primary mission objective is to perform a long-duration flight-test of various new-technology lightweight components, such as sensors, in a deep-space environment. The mission has two secondary science objectives: to provide high-resolution imaging of the entire lunar surface for mapping purposes and flyby imaging of the asteroid 1620 Geographos. The DSPSE mission is sponsored by the Strategic Defense Initiative Organization (SDIO). As prime contractor, the Naval Research Laboratory (NRL) is building the spacecraft and will conduct mission operations. The Goddard Space Flight Center's (GSFC) Flight Dynamics Division is supporting NRL in the areas of The Deep Space Network (DSN) will provide tracking support. The DSPSE mission will begin with a launch from the Western Test Range in late January 1994. Following a minimum 1.5-day stay in a low-Earth parking orbit, a solid kick motor burn will boost DSPSE into an 18-day, 2.5-revolution phasing orbit transfer trajectory to the Moon. Two burns to insert DSPSE into a lunar polar orbit suitable for the mapping mission will be followed by mapping orbit maintenance and adjustment operations over a period of 2 sidereal months. In May 1994, a lunar orbit departure maneuver, in conjunction with a lunar swingby 26 days later, will propel DSPSE onto a heliocentric transfer that will intercept Geographos on September 1, 1994. This paper presents the characteristics, deterministic delta-Vs, and design details of each trajectory phase of this unique mission, together with the requirements, constraints, and design considerations to which each phase is subject. Numerous trajectory plots and tables of significant trajectory events are included. Following a discussion of the results of a preliminary launch window analysis, a summary of the deterministic impulsive delta-V budget required to establish the baseline mission trajectory design is presented.

Carrington, D.↗
NASA NTRS1976

Mission design for a halo orbiter of the earth

The International Sun-Earth Explorer (ISEE) scientific satellite to be stationed in 1978 in the vicinity of the sun-earth interior libration point to continuously monitor the space between the sun and the earth, including the distant geomagnetic tail is described. Orbit selection considerations for the ISEE-C are discussed along with stationkeeping requirements and fuel-optimal trajectories. Due to the alignment of the interior libration point with the sun as viewed from the earth, it will be necessary to place the satellite into a 'halo orbit' around the libration point, in order to eliminate solar interference with down-link telemetry. Parametric data for transfer trajectories between an earth parking orbit (altitude about 185 km) and a libration-point orbit are presented. It is shown that the insertion magnitude required for placing a satellite into an acceptable halo orbit is rather modest.

Farquhar, R. W.↗
NASA NTRS2009

Direct Multiple Shooting Optimization with Variable Problem Parameters

Taking advantage of a novel approach to the design of the orbital transfer optimization problem and advanced non-linear programming algorithms, several optimal transfer trajectories are found for problems with and without known analytic solutions. This method treats the fixed known gravitational constants as optimization variables in order to reduce the need for an advanced initial guess. Complex periodic orbits are targeted with very simple guesses and the ability to find optimal transfers in spite of these bad guesses is successfully demonstrated. Impulsive transfers are considered for orbits in both the 2-body frame as well as the circular restricted three-body problem (CRTBP). The results with this new approach demonstrate the potential for increasing robustness for all types of orbit transfer problems.

Whitley, Ryan J.↗
NASA NTRS2014

An Integrated Approach to Modeling Solar Electric Propulsion Vehicles During Long Duration, Near-Earth Orbit Transfers

Recent NASA interest in utilizing solar electronic propulsion (SEP) technology to transfer payloads, e.g. from low-Earth orbit (LEO) to higher energy geostationary-Earth orbit (GEO) or to Earth escape, has necessitated the development of high fidelity SEP vehicle models and simulations. These models and simulations need to be capable of capturing vehicle dynamics and sub-system interactions experienced during the transfer trajectories which are typically accomplished with continuous-burn (potentially interrupted by solar eclipse), long duration "spiral out" maneuvers taking several months or more to complete. This paper presents details of an integrated simulation approach achieved by combining a high fidelity vehicle simulation code with a detailed solar array model. The combined simulation tool gives researchers the functionality to study the integrated effects of various vehicle sub-systems (e.g. vehicle guidance, navigation and control (GN&C), electric propulsion system (EP)) with time varying power production. Results from a simulation model of a vehicle with a 50 kW class SEP system using the integrated tool are presented and compared to the results from another simulation model employing a 50 kW end-of-life (EOL) fixed power level assumption. These models simulate a vehicle under three degree of freedom dynamics (i.e. translational dynamics only) and include the effects of a targeting guidance algorithm (providing a "near optimal" transfer) during a LEO to near Earth escape (C (sub 3) = −2.0 km (sup 2) / sec (sup −2) spiral trajectory. The presented results include the impact of the fully integrated, time-varying solar array model (e.g. cumulative array degradation from traversing the Van Allen belts, impact of solar eclipses on the vehicle and the related temperature responses in the solar arrays due to operating in the Earth's thermal environment, high fidelity array power module, etc.); these are used to assess the impact on vehicle performance (i.e. propellant consumption) and transit times.

vehicle↗
NASA NTRS2002

Benefits of Nuclear Electric Propulsion for Outer Planet Exploration

Nuclear electric propulsion (NEP) offers significant benefits to missions for outer planet exploration. Reaching outer planet destinations, especially beyond Jupiter, is a struggle against time and distance. For relatively near missions, such as a Europa lander, conventional chemical propulsion and NEP offer similar performance and capabilities. For challenging missions such as a Pluto orbiter, neither chemical nor solar electric propulsion are capable while NEP offers acceptable performance. Three missions are compared in this paper: Europa lander, Pluto orbiter, and Titan sample return, illustrating how performance of conventional and advanced propulsion systems vary with increasing difficulty. The paper presents parametric trajectory performance data for NEP. Preliminary mass/performance estimates are provided for a Europa lander and a Titan sample return system, to derive net payloads for NEP. The NEP system delivers payloads and ascent/descent spacecraft to orbit around the target body, and for sample return, delivers the sample carrier system from Titan orbit to an Earth transfer trajectory. A representative scientific payload 500 kg was assumed, typical for a robotic mission. The resulting NEP systems are 100-kWe class, with specific impulse from 6000 to 9000 seconds.

Kos, Larry↗
NASA NTRS1977

A comparison of solar sail and ion drive trajectories for a Halley's comet rendezvous mission

According to the propulsion concept of solar sail spacecraft the thrust force is produced by the specular reflection of sunlight from a large, essentially flat, reflecting surface. The magnitude of this force is approximately 9 newtons for a perfectly reflecting sail with an area of 1 square kilometer oriented normal to the sunline at a distance of one astronomical unit from the sun. There exists a restriction in the types of orbit transfer trajectories which can be considered with this propulsion system. In the case of the second propulsion system being considered for the Halley's comet rendezvous mission, thrust is produced by the acceleration of ionized mercury atoms by an electric field. Power to the ion thrusters is supplied by lightweight solar arrays which can provide up to 100 kW of electrical power at a distance of 1 AU from the sun. Because of differing thrust constraints for the two propulsion systems, trajectories for a Halley's comet rendezvous mission are significantly different for the Ion Drive and Solar Sail spacecraft. Details concerning the trajectory characteristics are shown with the aid of a number of graphs.

Sauer, C. G., Jr.↗
NASA NTRS2004

Sun-Earth L1 Region Halo-To-Halo Orbit and Halo-To-LisaJous Orbit Transfers

Practical techniques for designing transfer trajectories between Libration Point Orbits (LPOs) are presented. Motivation for development of these techniques was provided by a hardware contingency experienced by the Solar Heliospheric Observatory (SOHO), a joint mission of the European Space Agency (ESA) and the National Aeronautics and Space Administration (NASA) orbiting the L1 point of the Sun-Earth system. A potential solution to the problem involved a transfer from SOHO s periodic halo orbit to a new LPO of substantially different dimensions. Assuming the SOHO halo orbit as the departure orbit, several practical LPO transfer techniques were developed to obtain new Lissajous or periodic halo orbits that satisfy mission requirements and constraints. While not implemented for the SOHO mission, practical LPO transfer techniques were devised that are generally applicable to current and future LPO missions.

Roberts, Craig E.↗