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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 127 records · Page 7
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NASA NTRS1994

Tracking strategies for laser ranging to multiple satellite targets

By the middle of the decade, several new Laser Geodynamic Satellites will be launched to join the current constellation comprised of the laser geodynamic satellite (LAGEOS) (US), Starlette (France), Ajisai (Japan), and Etalon I and II (USSR). The satellites to be launched, LAGEOS II and III (US & Italy), and Stella (France), will be injected into orbits that differ from the existing constellation so that geodetic and gravimetric quantities are sampled to enhance their resolution and accuracy. An examination of various possible tracking strategies adopted by the network of laser tracking stations has revealed that the recovery of precise geodetic parameters can be obtained over shorter intervals than is currently obtainable with the present constellation of satellites. This is particularly important in the planning of mobile laser tracking operations, given a network of permanently operating tracking sites. Through simulations, it is shown that laser tracking of certain satellite passes, pre-selected to provide optimal sky-coverage, provides the means to acquire a sufficient amount of data to allow the recovery of 1 cm station positions.

Robbins, J. W.↗
NASA NTRS1993

TOPEX/POSEIDON Orbit Acquisition Manuever Sequence

TOPEX/POSEIDON....was launched...on August 10, 1992...The Project wanted to achieve operational orbit as soon as possible in order to begin oceanographic data acquisition through altimeters. A sequence of six manuevers was implemented to acquire the operational orbit following injection...

TOPEX TOPEX/POSEIDON manuever sequence design↗
NASA NTRS2013

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

A crewed mission to Mars poses a significant challenge in dealing with the physiological issues that arise with the crew being exposed to a near zero-gravity environment as well as significant solar and galactic radiation for such a long duration. While long surface stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological effects on the crew. However, for a 1-year round trip mission, the outbound and inbound hyperbolic velocity at Earth and Mars can be very large resulting in a significant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power levels (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower specific mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for efficient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo-Newtons of thrust at reasonably high specific impulse (Isp) of 900 seconds for impulsive transplanetary injection and orbital insertion maneuvers. When in power generation/EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Burke, Laura M.↗
NASA NTRS2013

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

A crewed mission to Mars poses a signi cant challenge in dealing with the physiolog- ical issues that arise with the crew being exposed to a near zero-gravity environment as well as signi cant solar and galactic radiation for such a long duration. While long sur- face stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological e ects on the crew. However, for a 1-year round trip mission, the outbound and inbound hy- perbolic velocity at Earth and Mars can be very large resulting in a signi cant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power lev- els (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower speci c mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for ecient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo- Newtons of thrust at reasonably high speci c impulse (Isp) of 900 seconds for impulsive trans-planetary injection and orbital insertion maneuvers. When in power generation / EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Burke, Laura A.↗
NASA NTRS1975

Utility of Space Transportation System to Space Communication Community

A potentially cost effective technique was investigated of launching operational satellites into synchronous orbit using the space transportation system (STS). This technique uses an unguided spinning solid rocket motor as the means for boosting a satellite from a low altitude shuttle parking orbit into a synchronous transfer orbit. The spacecraft is then injected into a geosynchronous orbit by an apogee kick motor fired at transfer orbit apogee. The approach is essentially that used on all Delta and Atlas-Centaur launches of synchronous satellites with the shuttle orbiter performing the function of the first two stages of the Delta three stage launch vehicle and the perigee kick motor performing the function of the Delta third state. It is concluded that the STS can be useful to the space communication community as well as to other geostationary satellite system users if the recommended actions are implemented.

Bronstein, L. M.↗
NASA NTRSDate not supplied

Artemis I: Test Flight Buys Down Risk for Humanity's Return to the Moon

The Artemis program successfully demonstrated its foundational deep space human transportation system during the Artemis I test flight in late 2022. This bold mission put the Artemis program on a course to accomplish increasingly complex missions to return humans to the Moon and to prepare for Mars and other destinations. The Artemis I uncrewed mission tested the Space Launch System (SLS) rocket and the Orion spacecraft using the ground processing, launch and recovery services provided by the Exploration Ground Systems (EGS) program at the Kennedy Space Center (KSC). The Artemis team accomplished key mission priorities that deliberately bought down risk for upcoming crewed missions as part of a build-up flight test strategy. The Artemis I mission priorities were 1) demonstrate Orion heatshield at lunar re-entry conditions; 2) operate systems in flight environment; and 3) retrieve the spacecraft. The successful performance and demonstration of these primary mission objectives from lift-off thru orbital insertion, translunar injection, outbound transit towards the Moon, entry into lunar orbit, coast in the lunar Distant Retrograde Orbit (DRO), exit from lunar orbit, transit back to Earth, atmospheric re-entry, descent, splashdown and recovery successfully showed that NASA has a foundational deep space transportation capability in place for human class missions. In addition to the primary mission objectives, a set of scientific payloads, approved Flight Test Objectives (FTOs) and Development Flight Test Objectives (DFTOs) were incorporated into the nominal mission plan and were designated as secondary, or priority 4, and were considered “bonus objectives” in terms of crewed flight risk buydown. During the mission, the team found itself in a position where it was comfortable achieving even more content than originally planned. As a result, DFTOs were added during the mission to further buy down risk to later crewed test flights. This paper presents an overview of the Artemis I test flight, including mission priorities and key test objectives, the launch campaign and as-flown mission profile, test objectives accomplished over the course of the 25.5- day mission, and implications to the crewed Artemis II and later test flight missions.

Artemis↗
NASA NTRS1986

Gas ionization induced by a high speed plasma injection in space

Gas ionization induced by a fast plasma injection has been observed with the Space Experiments with Particle Accelerators (SEPAC) Experiment on Spacelab-1. When an impulsive high-density plasma was injected from the orbiter, waves near the lower-hybrid frequency were enhanced, and the surrounding gas drifting with the orbiter was ionized for several tens of milliseconds after the plasma injection. The long-duration gas ionization was observed only when the plasma flux incoming to the orbiter cargo bay and the orbital velocity perpendicular to the magnetic field were relatively large. This effect has been explained by the concept of critical velocity ionization (CVI) for the gas drifting with the orbiter, although the gas velocity perpendicular to the magnetic field was less than the Alfven critical velocity.

Sasaki, S.↗
NASA NTRS1994

SEPAC co-investigator support

Mission preparation and support, data reduction and analysis, and modelling of beam phenomenon activities are described with respect to the Space Experiments with Particle Accelerators (SEPAC) project. The SEPAC Electron Beam Accelerator was used in observing the life time and drift velocity of ionization streaks in the upper atmosphere by ground based radars. A second active source, the plasma contactor (PC) electrically neutralized the shuttle orbiter during beam injections. Several observations of plasma density enhancements were correlated with orbiter crossings of the magnetic meridian of the Jicamarca, Peru radar located near the magnetic equator. One of several successful experiments was the artificial aurora experiment in which downward directed electron beam pulses were observed by the onboard AEPI camera. Theoretical and modelling efforts focused on phenomena related to electron beam and plasma cloud injection from spacecraft.

Source record↗
NASA NTRS2015

Lunar Cube Transfer Trajectory Options

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can be considered which have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfers and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

Libration Orbits↗
NASA NTRS2015

Lunar Cube Transfer Trajectory Options

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can b e considered which have a wide range of transfer duration, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO) geostationary transfer orbits (GTO) and higher energy direct lunar transfer and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

Libration Orbits↗
NASA NTRS1974

A feasibility study of unmanned rendezvous and docking in Mars orbit. Volume 1: Summary

The technical feasibility of achieving automatic rendezvous and docking in Mars orbit as a part of a surface sample return mission was investigated based on using as much existing Viking '75 Orbiter and Lander hardware as possible. Both 1981 and 1983/84 mission opportunities were considered. The principle result of the study was the definition of a three stage 289 kg Mars Ascent Vehicle (MAV) capable of accepting a 1 kg sample, injecting itself into a 2200 km circular orbit, and rendezvousing with an orbiting spacecraft carrying an earth return vehicle. Conclusions are that with state of the art systems plus limited application of new developments in areas where feasibility has already been demonstrated, e.g., solid rocket motor sterilization, it is possible to land a small ascent vehicle capable of automatically ascending and rendezvousing with a modified Viking '75 orbiter spacecraft. The mission can be flown in 1981 or 1983/84, but a dual launch or a larger launch vehicle than the Viking Titan 3 Centaur, or the use of space storable propellants for Mars orbit injection, would be required in the 1983/84 opportunity.

Source record↗
NASA NTRS1985

Argon ion pollution of the magnetosphere

Construction of a Solar Power Satellite (SPS) would require the injection of large quantities of propellant to transport material from Low Earth Orbit (LEO) to the construction site at Geostationary Earth Orbit (GEO). This injection, in the form of approx 10 to the 32nd power, 2 KeV argon ions (and associated electrons) per SPS, is comparable to the content of the plasmasphere (approx 10 to the 31st power ions). In addition to the mass deposited, this represents a considerable injection of energy. The injection is examined in terms of a simple model for the expansion of the beam plasma. General features of the subsequent magnetospheric convection of the argon are also examined.

Lopez, R. E.↗
NASA NTRS1992

Lunar lander conceptual design

This paper is a first look at the problems of building a lunar lander to support a small lunar surface base. A series of trade studies was performed to define the lander. The initial trades concerned choosing number of stages, payload mass, parking orbit altitude, and propellant type. Other important trades and issues included plane change capability, propellant loading and maintenance location, and reusability considerations. Given a rough baseline, the systems were then reviewed. A conceptual design was then produced. The process was carried through only one iteration. Many more iterations are needed. A transportation system using reusable, aerobraked orbital transfer vehicles (OTV's) is assumed. These OTV's are assumed to be based and maintained at a low Earth orbit (LEO) space station, optimized for transportation functions. Single- and two-stage OTV stacks are considered. The OTV's make the translunar injection (TLI), lunar orbit insertion (LOI), and trans-Earth injection (TEI) burns, as well as midcourse and perigee raise maneuvers.

Stecklein, J. M.↗
NASA NTRS1985

An enhancement of plasma density by neutral gas injection observed in SEPAC Spacelab-1 experiment

An enhancement of plasma density observed during a neutral gas injection in Space Experiments with Particle Accelerators by the Space Shuttle/Spacelab-1 is presented. When a plume of nitrogen gas was injected from the orbiter into space, a large amount of plasma was detected by an onboard plasma probe. The observed density often increased beyond the background plasma density and was strongly dependent on the attitude of the orbiter with respect to the velocity vector. This effect has been explained by a collisional interaction between the injected gas molecules and the ionospheric ions relatively drifting at the orbital speed.

Sasaki, S.↗
NASA NTRS1975

Dynamic analysis of the long-term zonal earth energy budget experiment (LZEEBE) spacecraft

The motions of the flexible LZEEBE spacecraft after injection into a circular orbit were analyzed. The spacecraft consists of three spherical balloons connected to a central hub by three long flexible booms which lie in a plane such that the angle between any pair of booms is 120 degrees. The major torques acting on the spacecraft are gravity-gradient torques and torques due to solar radiation pressure acting on the balloons which have different reflective properties. A development of the equations of motion is presented. Computer simulations indicate that the spacecraft will have random motion, provided the injection into orbit does not create conditions favorable to gravity-gradient capture.

Meirovitch, L.↗
NASA NTRS2005

Proof-of-Concept Trajectory Designs for a Multi-Spacecraft, Low-Thrust Heliocentric Solar Weather Buoy Mission

A new solar weather mission has been proposed, involving a dozen or more small spacecraft spaced at regular, constant intervals in a mutual heliocentric circular orbit between the orbits of Earth and Venus. These solar weather buoys (SWBs) would carry instrumentation to detect and measure the material in solar flares, solar energetic particle events, and coronal mass ejections as they flowed past the buoys, serving both as science probes and as a radiation early warning system for the Earth and interplanetary travelers to Mars. The baseline concept involves placing a mothercraft carrying the SWBs into a staging orbit at the Sun-Earth L1 libration point. The mothercraft departs the L1 orbit at the proper time to execute a trailing-edge lunar flyby near New Moon, injecting it into a heliocentric orbit with its perihelion interior to Earth s orbit. An alternative approach would involve the use of a Double Lunar Swingby (DLS) orbit, rather than the L1 orbit, for staging prior to this flyby. After injection into heliocentric orbit, the mothercraft releases the SWBs-all equipped with low-thrust pulsed plasma thrusters (PPTs)-whereupon each SWB executes a multi-day low-thrust finite bum around perihelion, lowering aphelion such that each achieves an elliptical phasing orbit of different orbital period from its companions. The resulting differences in angular rates of motion cause the spacecraft to separate. While the lead SWB achieves the mission orbit following an insertion burn at its second perihelion passage, the remaining SWBs must complete several revolutions in their respective phasing orbits to establish them in the mission orbit with the desired longitudinal spacing. The complete configuration for a 14 SWB scenario using a single mothercraft is achieved in about 8 years, and the spacing remains stable for at least a further 6 years. Flight operations can be simplified, and mission risk reduced, by employing two mothercraft instead of one. In this scenario: the second mothercraft stays in a libration-point or DLS staging orbit until the first mothercraft has achieved nearly 180 separation from the Earth. The timing of the second mothercraft's subsequent lunar flyby is planned such that this spacecraft will be located 180 from the first mothercraft upon completion of its heliocentric circularization maneuvers. Both groups of satellites then only have to spread out over 180 to obtain full 360 coverage around the Sun.

Muller, Ronald↗
NASA NTRS1991

German telecommunications satellite (Deutscher fernmelde satellit) (DFS-1 and -2)

The German Telecommunications Satellite (DFS) Program is to provide telecommunications service for high data rate transmission of text and video data to the Federal Republic of Germany within the 11-14 GHz and 20-30 GHz bands. The space segment of this program is composed of three satellites, DFS-1, DFS-2, and DFS-3, which will be located at 23.5 degrees E longitude of the geostationary orbit. The DFS will be launched from the Center Spatial Guyanis in French Giana on an Ariane launch vehicle. The mission follows the typical injection sequence: parking orbit, transfer orbit, and earth orbit. Attitude maneuvers will be performed to orient the spacecraft prior to Apogee Kick Motor (AKM) firing. After AKM firing, drift phase orbital and attitude maneuvers will be performed to place the spacecraft in its final geostationary position. The Deep Space Network (DSN) will support the transfer and drift orbit mission phases. Information is presented in tabular form for the following areas: DSN support, compatibility testing, frequency assignments, telemetry, command, and tracking support responsibilities.

Hiendlmeier, G.↗
NASA NTRSDate not supplied

Cislunar Trajectory Design and Maneuver Autonomy for NASA's Moon to Mars Architecture

NASA’s Moon to Mars architecture is an ambitious roadmap of manned cislunar and deep space exploration. The extensive amount of orbital assets required will place a significant burden on ground-based resources, such as communication networks and operations facilities. Spacecraft autonomy is essential for maintaining a vast number of complex missions beyond Earth orbit. To achieve full autonomy, spacecraft must be able to employ methods of robust maneuver design without an explicit dependence on commands sent from the ground. This level of autonomy is needed not only for stationkeeping, but also for outbound transfers. To address the need of spacecraft maneuver design autonomy, this work investigates the use of neural networks (NNs) in a supervised learning environment. A supervised learning approach for NNs allows for a curated training data set, consisting exclusively of perturbations applied to a desired mission concept of operations (ConOps). The proposed approach allows humans on the ground to design a specific mission ConOps before flight, then employ NNs to fly the mission robustly and autonomously. This investigation numerically tests maneuver autonomy in four highly sensitive regions of flight: orbit raising, translunar injection burns, powered lunar flybys, and invariant manifold insertion burns. These straining cases are contextualized by testing them in a demonstration mission, targeting an Earth-Moon L3 orbit. The study first establishes feasibility by automating impulsive burn maneuvers. However, some guidance algorithms will need more intensive commands, such as inertial pointing and angular rates. To validate this method, NN maneuver autonomy is applied to a finite burn model of the demonstration mission. The use of sequential, mission specific maneuvers provide an appropriate testbed to demonstrate the robustness of a NN trained on feasible perturbed states. Moreover, these scenarios provide preliminary proof-of-concept for fully autonomous missions that execute maneuvers without dependence upon explicit command uplinks. As a result, the technological advancement proposed in this work may significantly ease the strain on ground-based mission operations. This would enable complex and autonomous mission execution in cislunar and deep space regimes, filling a technology gap required to support future manned missions.

NASA↗