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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Transfer Trajectory Options for Servicing Sun-Earth-Moon Libration Point Missions

Future missions to the Sun-Earth Libration L1 and L2 regions will require scheduled servicing to maintain hardware and replenish consumables. While there have been statements made by various NASA programs regarding servicing of vehicles at these locations or in Cis-lunar space, a practical transfer study has not been extensively investigated in an operational fashion to determine the impacts of navigation and maneuver errors. This investigation uses dynamical systems and operational models to design transfer trajectories between the Sun-Earth Libration region (QuasiHalo orbit) and the Earth-Moon vicinity (Distant Retrograde Orbit, QuasiHalo Orbit, Halo Orbit, and Near Rectilinear Halo Orbit). We address the total delta V cost of transfers between each pair of locations using a Monte Carlo analysis.

Earth↗
NASA NTRS1990

Examples of the nonlinear dynamics of ballistic capture and escape in the earth-moon system

An example of a trajectory is given which is initially captured in an elliptic resonant orbit about the earth and then ballistically escapes the earth-moon system. This is demonstrated by a numerical example in three-dimensions using a planetary ephemeris. Another example shows a mechanism of how an elliptic orbit about the earth can increase its energy by performing a complex nonlinear transition to an elliptic orbit of a larger semi-major axis. Capture is also considered. An application of ballistic capture at the moon via an unstable periodic orbit using the four-body sun-earth-moon-S/C interaction is described.

Belbruno, Edward A.↗
NASA NTRS2010

Optimal Low Energy Earth-Moon Transfers

The optimality of a low-energy Earth-Moon transfer is examined for the first time using primer vector theory. An optimal control problem is formed with the following free variables: the location, time, and magnitude of the transfer insertion burn, and the transfer time. A constraint is placed on the initial state of the spacecraft to bind it to a given initial orbit around a first body, and on the final state of the spacecraft to limit its Keplerian energy with respect to a second body. Optimal transfers in the system are shown to meet certain conditions placed on the primer vector and its time derivative. A two point boundary value problem containing these necessary conditions is created for use in targeting optimal transfers. The two point boundary value problem is then applied to the ballistic lunar capture problem, and an optimal trajectory is shown. Additionally, the ballistic lunar capture trajectory is examined to determine whether one or more additional impulses may improve on the cost of the transfer.

Griesemer, Paul Ricord↗
NASA NTRS2019

Transfer Trajectory Options for Servicing Sun-Earth-Moon Libration Point Missions

Future missions to the Sun-Earth Libration L1 and L2 regions will require scheduled servicing to maintain hardware and replenish consumables. While there have been statements made by various NASA programs regarding servicing of vehicles at these locations or in Cis-lunar space, a practical transfer study has not been extensively investigated in an operational fashion to determine the impacts of navigation and maneuver errors. This investigation uses dynamical systems and operational models to design transfer trajectories between the Sun-Earth Libration region (QuasiHalo orbit) and the Earth-Moon vicinity (Distant Retrograde Orbit, QuasiHalo Orbit, Halo Orbit, and Near Rectilinear Halo Orbit). We address the total ΔV cost of transfers and operational considerations between each pair of locations using a Monte Carlo analysis.

Folta, David C.↗
NASA NTRS2012

Simulating the Liaison Navigation Concept in a Geo + Earth-Moon Halo Constellation

Linked Autonomous Interplanetary Satellite Orbit Navigation, or LiAISON, is a novel satellite navigation technique where relative radiometric measurements between two or more spacecraft in a constellation are processed to obtain the absolute state of all spacecraft. The method leverages the asymmetry of the gravity field that the constellation exists in. This paper takes a step forward in developing a high fidelity navigation simulation for the LiAISON concept in an Earth-Moon constellation. In particular, we aim to process two-way Doppler measurements between a satellite in GEO orbit and another in a halo orbit about the Earth-Moon L1 point.

Linked Autonomous Interplanetary Satellite Orbit N↗
NASA NTRS1986

Orbital acrobatics in the Sun-Earth-Moon system

Unconventional trajectory techniques for space missions in the Sun-Earth-Moon system, including libration-point orbits, gravity-assist maneuvers, and Earth-return trajectories are reviewed. The ISEE-3/ICE flight experience is used to illustrate the utility of libration-point orbits called halo-orbits. Five lunar gravity-assist maneuvers used by the ISEE-3/ICE spacecraft are discussed. The final lunar swingby sent the spacecraft into a heliocentric trajectory that will eventually intercept Comet Giacobini-Zinner. As an example of the Earth-return trajectory concept, a proposed mission that includes flybys of three comets and two asteroids is described.

Farquhar, Robert W.↗
NASA NTRS1980

Tidal-friction theory of the earth-moon system

Serious errors contained in Jeffreys' (1952, 1959, 1970, 1976) discussion of tidal friction in the earth-moon system are identified and their consequences are discussed. A direct solution of the dynamical tidal equations for the couple from the earth acting upon the moon and the couple from the earth acting upon the sun, which were left unsolved by Jeffreys, is found to be incompatible with observations and the predictions of linear or quadratic friction theory, due to his failure to take into account the possible change of the moment of inertia of the earth with time in the derivation of the dynamical equations. Consideration of this factor leads to the conclusion that the earth must be contracting at a rate of 14.7 x 10 to the -11th/year, which can be accounted for only by the Ramsey theory, in which the terrestrial core is considered as a phase change rather than a change in chemical composition. Implications of this value for the rates of changes in day length and lunar distance are also indicated.

Lyttleton, R. A.↗
NASA NTRS2013

Earth-Moon Libration Point Orbit Stationkeeping: Theory, Modeling and Operations

Collinear Earth-Moon libration points have emerged as locations with immediate applications. These libration point orbits are inherently unstable and must be maintained regularly which constrains operations and maneuver locations. Stationkeeping is challenging due to relatively short time scales for divergence effects of large orbital eccentricity of the secondary body, and third-body perturbations. Using the Acceleration Reconnection and Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) mission orbit as a platform, the fundamental behavior of the trajectories is explored using Poincare maps in the circular restricted three-body problem. Operational stationkeeping results obtained using the Optimal Continuation Strategy are presented and compared to orbit stability information generated from mode analysis based in dynamical systems theory.

Folta, David C.↗
NASA NTRS1980

Utilization of multi-body trajectories in the Sun-Earth-Moon system

An overview of three uncommon trajectory concepts for space missions in the Sun-Earth-Moon System is presented. One concept uses a special class of libration-point orbits called 'halo orbits.' It is shown that members of this orbit family are advantageous for monitoring the solar wind input to the Earth's magnetosphere, and could also be used to establish a continuous communications link between the Earth and the far side of the Moon. The second concept employs pretzel-like trajectories to explore the Earth's geomagnetic tail. These trajectories are formed by using the Moon to carry out a prescribed sequence of gravity-assist maneuvers. Finally, there is the 'boomerang' trajectory technique for multiple-encounter missions to comets and asteroids. In this plan, Earth-swingby maneuvers are used to retarget the original spacecraft trajectory. The boomerang method could be used to produce a triple-encounter sequence which includes flybys of comets Halley and Tempel-2 as well as the asteroid Geographos.

Farquhar, R. W.↗
NASA NTRS1971

The use of the Earth-Moon system as a radio tracking facility

The use of the earth-moon system for radio tracking has several potentially useful aspects, from a scientific standpoint: extension of the baseline by two orders of magnitude, absence of lunar atmosphere, higher-velocity parallax with a longer period, and the moon's more modellable rotation. The facility would be a convenient base station for other research, not directly involving the tracking capability, but utilizing its orbiting data link features.

Mulholland, J. D.↗
NASA NTRS2000

Dynamics of the Final Stages of Terrestrial Planet Growth and the Formation of the Earth-Moon System

An overview of current theories of star and planet formation, with emphasis on terrestrial planet accretion and the formation of the Earth-Moon system is presented. These models predict that rocky planets should form around most single stars, although it is possible that in some cases such planets are lost to orbital decay within the protoplanetary disk. The frequency of formation of gas giant planets is more difficult to predict theoretically. Terrestrial planets are believed to grow via pairwise accretion until the spacing of planetary orbits becomes large enough that the configuration is stable for the age of the system. Giant impacts during the final stages of growth can produce large planetary satellites, such as Earth's Moon. Giant planets begin their growth like terrestrial planets, but they become massive enough that they are able to accumulate substantial amounts of gas before the protoplanetary disk dissipates.

Lissauer, Jack J.↗
NASA NTRS2015

A Free-Return Earth-Moon Cycler Orbit for an Interplanetary Cruise Ship

A periodic circumlunar orbit is presented that can be used by an interplanetary cruise ship for regular travel between Earth and the Moon. This Earth-Moon cycler orbit was revealed by introducing solar gravity and modest phasing maneuvers (average of 39 m/s per month) which yields close-Earth encounters every 7 or 10 days. Lunar encounters occur every 26 days and offer the chance for a smaller craft to depart the cycler and enter lunar orbit, or head for a Lagrange point (e.g., EM-L2 halo orbit), distant retrograde orbit (DRO), or interplanetary destination such as a near-Earth object (NEO) or Mars. Additionally, return-to-Earth abort options are available from many points along the cycling trajectory.

Interplanetary cruise ship↗
NASA NTRS1971

Experiment S029: Earth-Moon libration regions photography

Eleven photographs were taken of the L sub 4 earth-moon libration region during the Gemini 12 mission. The objective was to investigate the possible existence of clouds or particles of dust orbiting the earth in the L sub 4 region by the use of photographic techniques. No identifiable data were found in the first sequence of five photographs. Star field images could be recognized in two photographs of the second sequence of pictures. A light flare, of unknown source but possibly caused by light refraction from the window of the spacecraft, was recorded in both pictures. A reconstructed star field region (as photographed) is shown Isodensitraces of the two recognizable star fields were made; however, no conclusive results were obtained.

Morris, E. C.↗
NASA NTRS2012

Using Gravity Assists in the Earth-moon System as a Gateway to the Solar System

For spacecraft departing the Earth - Moon system, lunar flybys can significantly increase the hype rbolic escape energy (C3, in km 2 /sec 2 ) for a modest increase in flight time. Within ~2 months, lunar flybys can produce a C3 of ~2. Over 4 - 6 months, lunar flybys alone can increase the C3 to ~4.5, or they can provide for additional periapsis burns to increase the C3 from ~2 -3 to 10 or more, suitable for planetary missions. A lunar flyby departure can be followed by additional ∆ -V (such as that efficiently provided by a low thrust system, eg. Solar Electric Propulsion (SEP)) to raise the Earth - relative velocity (at a ratio of more than 2:1) before a subsequent Earth flyby, which redirects that velocity to a more di stant target, all within not much more than a year. This paper describes the applicability of lunar flybys for different flight times and propulsi on systems, and illustrates this with instances of past usage and future possibilities. Examples discussed i nclude ISEE - 3, Nozomi, STEREO, 2018 Mars studies (which showed an 8% payload increase), and missions to Near Earth Objects (NEOs). In addition, the options for the achieving the initial lunar flyby are systematically discussed, with a view towards their p ractical use with in a compact launch period. In particular, we show that launches to geosynchronous transfer orbit (GTO) as a secondary payload provide a feasible means of obtaining a lunar flyby for an acceptable cost, even for SEP systems that cannot ea sily deliver large ∆ - Vs at periapsis. Taken together, these results comprise a myriad of options for increasing the mission performance, by the efficient use of lunar flybys within an acceptable extension of the flight time.

lunar fly-bys↗