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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Initial Satellite Formation Flight Results from the Magnetospheric Multiscale Mission

This paper will describe the results that have been obtained to date concerning MMS formation flying. The MMS spacecraft spin at a rate of 3.1 RPM, with spin axis roughly aligned with Ecliptic North. Several booms are used to deploy instruments: two 5 m magnetometer booms in the spin plane, two rigid booms of length 12.5 m along the positive and negative spin axes, and four flexible wire booms of length 60 m in the spin plane. Minimizing flexible motion of the wire booms requires that reorientation of the spacecraft spin axis be kept to a minimum: this is limited to attitude maneuvers to counteract the effects of gravity-gradient and apparent solar motion. Orbital maneuvers must therefore be carried out in essentially the nominal science attitude. These burns make use of a set of monopropellant hydrazine thrusters: two (of thrust 4.5 N) along the spin axis in each direction, and eight (of thrust 18 N) in the spin plane; the latter are pulsed at the spin rate to produce a net delta-v. An on-board accelerometer-based controller is used to accurately generate a commanded delta-v. Navigation makes use of a weak-signal GPS-based system: this allows signals to be received even when MMS is flying above the GPS orbits, producing a highly accurate determination of the four MMS orbits. This data is downlinked to the MMS Mission Operations Center (MOC) and used by the MOC Flight Dynamics Operations Area (FDOA) for maneuver design. These commands are then uplinked to the spacecraft and executed autonomously using the controller, with the ground monitoring the burns in real time.

satellite↗
NASA NTRS2017

LEO to GEO (and Beyond) Transfers Using High Power Solar Electric Propulsion (HP-SEP)

Rideshare, or Multi-Payload launch configurations, are becoming more and more commonplace but access to space is only one part of the overall mission needs. The ability for payloads to achieve their target orbits or destinations can still be difficult and potentially not feasible with on-board propulsion limitations. The High Power Solar Electric Propulsion (HP-SEP) Orbital Maneuvering Vehicle (OMV) provides transfer capabilities for both large and small payload in excess of what is possible with chemical propulsion. Leveraging existing secondary payload adapter technology like the ESPA provides a platform to support Multi-Payload launch and missions. When coupled with HP-SEP, meaning greater than 30 kW system power, very large delta-V maneuvers can be accomplished. The HP-SEP OMV concept is designed to perform a Low Earth Orbit to Geosynchronous Orbit (LEO-GEO) transfer of up to six payloads each with 300kg mass. The OMV has enough capability to perform this 6 kms maneuver and have residual capacity to extend an additional transfer from GEO to Lunar orbit. This high deltaV capability is achieved using state of the art 12.5kW Hall Effect Thrusters (HET) coupled with high power roll up solar arrays. The HP-SEP OMV also provides a demonstration platform for other SEP technologies such as advanced Power Processing Units (PPU), Xenon Feed Systems (XFS), and other HET technologies. The HP-SEP OMV platform can be leveraged for other missions as well such as interplanetary science missions and applications for resilient space architectures.

Leo and Geo and HP-SEP↗
NASA NTRS1976

Nonintrusive Flowmeter

The results of analytical and experimental work performed in the design, fabrication, and test of a prototype nonintrusive gaging system for use in monitoring the consumption of earth-storable fuels and oxidants in either a one-g or a zero-g environment are explained. The design specifications were those applicable to the reaction control system and to the orbital maneuvering system (OMS) fuel and oxidant on the space shuttle while in orbit. The major requirement was for the measurement of flow pulses with sufficient accuracy to provide a continuous knowledge of the fuel and oxidant remaining in the OMS system to within 1% or better. An ultrasonic frequency chirp technique was used having a high inherent rejection for signals traversing stray paths, and for random noise generated by the flowing liquid. A detailed analysis of the frequency chirp approach for two modes of operation (period and phase changes), including an error analysis are reported.

Greene, M. W.↗
NASA NTRS1994

Access to space

The goal of this conceptual design was to devise a reusable, commercially viable, single-stage-to-orbit vehicle. The vehicle has the ability to deliver a 9100 kg (20,000 lb) payload to a low earth orbit of 433 km to 933 km (250 n.mi. - 450 n.mi.). The SSTO vehicle is 51 meters in length and has a gross takeoff mass of 680,400 kg (1,500,000 lb). The vehicle incorporates three RD-701 engines for the main propulsion system and two RL-10 engines for the orbital maneuvering system. The vehicle is designed for a three day stay on orbit with two crew members.

Source record↗
NASA NTRS1993

Orbit determination covariance analysis for the Deep Space Program Science Experiment mission

To define an appropriate orbit support procedure for the DSPSE mission, detailed permission orbit determination covariance analyses have been performed for the translunar and trans-Geographos mission phases. Preliminary analyses were also performed for the lunar mapping mission phase. These analyses are designed to assess the tracking patterns and the amount of tracking data needed to obtain orbit solutions of required accuracy for each mission phase and before and after each major orbit perturbation, such as orbit maneuvers and flybys of the Earth and Moon. In addition to operational orbit determination procedures, these analyses identify major error sources, estimate their contribution to orbital errors, and address possible strategies to reduce orbit determination error. For the lunar orbit phase, several lunar gravity error modeling approaches have been investigated. The covariance analysis results presented in this paper will serve as a guide for providing orbit determination support for the DSPSE mission.

Beckman, M.↗
NASA NTRS1980

Shuttle program standard maneuver sequences for orbiter/upper-stage separation SSUS-A, SSUS-D, and IUS

Descriptions of standard post-ejection maneuver sequences for the deployment of IUS, SSUS-A, and SSUS-D upper stages from the space shuttle orbiter are presented. The sequences were designed to satisfy requirements for limiting the damage inflicted on the orbiter by upper-stage exhaust particles, subject to a further requirement for minimizing the impingement of orbiter thruster plumes on the deployed payload. In all cases it was assumed that the orbital maneuvering system engines would be used to apply the orbiter's major separation velocity increment.

Wilson, S. W.↗
NASA NTRS2017

Optimizing Parking Orbits for Roundtrip Mars Missions

A roundtrip Mars mission presents many challenges to the design of a transportation system and requires a series of orbital maneuvers within Mars vicinity to capture, reorient, and then return the spacecraft back to Earth. The selection of a Mars parking orbit is crucial to the mission design; not only can the parking or-bit choice drastically impact the ΔV requirements of these maneuvers but also it must be properly aligned to target desired surface or orbital destinations. This paper presents a method that can optimize the Mars parking orbits given the arrival and departure conditions from heliocentric trajectories, and it can also en-force constraints on the parking orbits to satisfy other architecture design requirements such as co-planar subperiapsis descent to planned landing sites, due east or co-planar ascent back to the parking orbit, or low cost transfers to and from Phobos and Deimos.

Qu, Min↗
NASA NTRS1973

Representative Space Shuttle missions and their impact on shuttle design.

Four representative Space Shuttle missions for the 1980's are described: mission 1, a third stage deployment and retrieval mission; mission 2, a combination orbiting element revisit and sortie mission; mission 3, a rapid payload retrieval and return mission; and mission 4, a 30-day earth observation sortie mission. The first three missions are proposed baseline reference missions, and, as such, their mission requirements must be met by the baseline shuttle configuration developed by NASA contractors. Typically, these missions impact the design in terms of such basics as shuttle launch payload capability, payload bay size, performance of the orbital maneuvering system (OMS) and the reaction control system (RCS), the basic 7-day orbital system lifetime, and the thermal protection system for high cross-range entry. To illustrate that vehicle design may be subject to requirements from other mission types as well, mission 4 is included. Each description stresses mission requirements having unique impact on vehicle design.

Young, K. A.↗
NASA NTRS1990

Goddard Mission Analysis System (GMAS)

Written to simulate complex satellite orbits including attitude control and orbital maneuvers, GMAS also performs shadow and station coverage studies and graph-selected orbital parameters for two satellites. Written in FORTRAN and ASSEMBLER.

Mcgarry, F. E.↗
NASA NTRS2001

A Cryogenic Propellant Production Depot for Low Earth Orbit

The cost of access to space beyond low Earth orbit can be lowered if vehicles can refuel in orbit. The power requirements for a propellant depot that electrolyzes water and stores cryogenic oxygen and hydrogen can be met using technology developed for space solar power. A propellant depot is described that will be deployed in a 400 km circular equatorial orbit, receive tanks of water launched into a lower orbit from Earth by gun launch or reusable launch vehicle, convert the water to liquid hydrogen and oxygen, and store up to 500 metric tonnes of cryogenic propellants. Orbital maneuvering vehicles will transfer the Earth-launched propellant tanks from the lower orbit to the depot orbit. The propellant stored in the depot can support transportation from low Earth orbit to geostationary Earth orbit, the Moon, LaGrange points, Mars, etc. The propellant tanks on the depot are modified versions of those used in the Delta IV-Heavy launch vehicle. The tanks are configured in an in-line gravity-gradient configuration to minimize drag and settle the propellant. Temperatures can be maintained by body-mounted radiators; these will also provide some shielding against orbital debris. Power is supplied by a pair of solar arrays mounted perpendicular to the orbital plane, which rotate once per orbit to track the Sun. The majority of the power will be used to run the electrolysis system. Technology needed for an orbiting propellant depot can be tested and demonstrated in the near-term on the ground, on a Shuttle-deployed free-flyer, and on the International Space Station. Further along, an orbital depot can be deployed that stores liquid hydrogen and oxygen launched from Earth, to be followed by a full conversion and storage depot.

Potter, Seth D.↗
NASA NTRS2005

Auxiliary Propulsion Activities in Support of NASA's Exploration Initiative

The Space Launch Initiative (SLI) procurement mechanism NRA8-30 initiated the Auxiliary Propulsion System/Main Propulsion System (APS/MPS) Project in 2001 to address technology gaps and development risks for non-toxic and cryogenic propellants for auxiliary propulsion applications. These applications include reaction control and orbital maneuvering engines, and storage, pressure control, and transfer technologies associated with on-orbit maintenance of cryogens. The project has successfully evolved over several years in response to changing requirements for re-usable launch vehicle technologies, general launch technology improvements, and, most recently, exploration technologies. Lessons learned based on actual hardware performance have also played a part in the project evolution to focus now on those technologies deemed specifically relevant to the Exploration Initiative. Formal relevance reviews held in the spring of 2004 resulted in authority for continuation of the Auxiliary Propulsion Project through Fiscal Year 2005 (FY05), and provided for a direct reporting path to the Exploration Systems Mission Directorate. The tasks determined to be relevant under the project were: continuation of the development, fabrication, and delivery of three 870 lbf thrust prototype LOX/ethanol reaction control engines; the fabrication, assembly, engine integration and testing of the Auxiliary Propulsion Test Bed at White Sands Test Facility; and the completion of FY04 cryogenic fluid management component and subsystem development tasks (mass gauging, pressure control, and liquid acquisition elements). This paper presents an overview of those tasks, their scope, expectations, and results to-date as carried forward into the Exploration Initiative.

Best, Philip J.↗
NASA NTRS2000

Natural Hazards of the Space Environment

Spacecraft in Low Earth Orbit (LEO) are subject to numerous environmental hazards. Here I'll briefly discuss three environment factors that pose acute threats to the survival of spacecraft systems and crew: atmospheric drag, impacts by meteoroids and orbital debris, and ionizing radiation. Atmospheric drag continuously opposes the orbital motion of a satellite, causing the orbit to decay. This decay will lead to reentry if not countered by reboost maneuvers. Orbital debris is a by-product of man's activities in space, and consists of objects ranging in size from miniscule paint chips to spent rocket stages and dead satellites. Ionizing radiation experienced in LEO has several components: geomagnetically trapped protons and electrons (Van Allen belts); energetic solar particles; galactic cosmic rays; and albedo neutrons. These particles can have several types of prompt harmful effects on equipment and crew, from single-event upsets, latchup, and burnout of electronics, to lethal doses to crew.All three types of prompt threat show some dependence on the solar activity cycle. Atmospheric drag mitigation and large debris avoidance require propulsive maneuvers. M/OD and ionizing radiation require some form of shielding for crew and sensitive equipment. Limiting exposure time is a mitigation technique for ionizing radiation and meteor streams.

Evans, Steven W.↗
NASA NTRS1993

MERCATOR: Methods and Realization for Control of the Attitude and the Orbit of spacecraft

Since 1974, CNES has been involved in geostationary positioning. Among different entities participating in operations and their preparation, the Flight Dynamics Center (FDC) is in charge of performing the following tasks: orbit determination; attitude determination; computation, monitoring, and calibration of orbit maneuvers; computation, monitoring, and calibration of attitude maneuvers; and operational predictions. In order to fulfill this mission, the FDC receives telemetry from the satellite and localization measurements from ground stations (e.g., CNES, NASA, INTELSAT). These data are processed by space dynamics programs integrated in the MERCATOR system which is run on SUN workstations (UNIX O.S.). The main features of MERCATOR are redundancy, modularity, and flexibility: efficient, flexible, and user friendly man-machine interface; and four identical SUN stations redundantly linked in an Ethernet network. Each workstation can perform all the tasks from data acquisition to computation results dissemination through a video network. A team of four engineers can handle the space mechanics aspects of a complete geostationary positioning from the injection into a transfer orbit to the final maneuvers in the station-keeping window. MERCATOR has been or is to be used for operations related to more than ten geostationary positionings. Initially developed for geostationary satellites, MERCATOR's methodology was also used for satellite control centers and can be applied to a wide range of satellites and to future manned missions.

Tavernier, Gilles↗
NASA NTRS1988

Near term space transportation systems for earth orbit and planetary applications

With the resumption of Space Shuttle flights in mid-1988, many of the initial flights will include a mix of upper stage propulsion systems for geosynchronous orbit and planetary mission applications. This paper presents a system definition and the modifications required to the Inertial Upper Stage and the Transfer Orbit Stage for the near term Space Shuttle missions, namely the Tracking and Data Relay Satellite, Magellan, Galileo, Ulysses, Mars Observer, and the Advanced Communications Technology Satellite. The definition and capabilities of the Orbital Maneuvering Vehicle, currently being developed by NASA to perform a wide range of on-orbit missions and services in support of orbiting spacecraft, are also included.

Saucier, Sidney↗
NASA NTRS2010

Update on PISCES

An updated version of the Platform Independent Software Components for the Exploration of Space (PISCES) software library is available. A previous version was reported in Library for Developing Spacecraft-Mission-Planning Software (MSC-22983), NASA Tech Briefs, Vol. 25, No. 7 (July 2001), page 52. To recapitulate: This software provides for Web-based, collaborative development of computer programs for planning trajectories and trajectory- related aspects of spacecraft-mission design. The library was built using state-of-the-art object-oriented concepts and software-development methodologies. The components of PISCES include Java-language application programs arranged in a hierarchy of classes that facilitates the reuse of the components. As its full name suggests, the PISCES library affords platform-independence: The Java language makes it possible to use the classes and application programs with a Java virtual machine, which is available in most Web-browser programs. Another advantage is expandability: Object orientation facilitates expansion of the library through creation of a new class. Improvements in the library since the previous version include development of orbital-maneuver- planning and rendezvous-launch-window application programs, enhancement of capabilities for propagation of orbits, and development of a desktop user interface.

Pearson, Don↗
NASA NTRS1978

Environmental Impact Statement for the Space Shuttle Program

Test firings and launches will release air pollutants causing a temporary localized small degradation in air quality near the tests or launch site. Areas adjacent to the site will also be subjected to moderate sound levels of predominantly low frequencies for short durations. During the launch phase, hydrogen chloride will be introduced into the stratosphere causing a small decrease in ozone. Temporary perturbations to the ionosphere will occur during orbital maneuvers and entry will have no significant effect on communication or radio wave propagation. As the Orbiter descends, a low magnitude sonic beam will be produced along the groundtrack with maximum overpressures occurring near the landing site. The overpressures will be infrequent, will vary in location and are of sufficiently low energy to be considered a momentary annoyance, if noticed at all. Major alternatives considered are discontinuation or postponement of the program, use of alternate propellants and neutralization of the ground cloud.

Malkin, M. S.↗
NASA NTRS1984

Satellite Servicing Technology Development Missions

A new capability regarding the U.S. space efforts will be related to the servicing of satellites in orbit utilizing first-generation space station as the collection point or base for Shuttle-delivered payloads. Orbital maneuvering vehicles could move payloads or spacecraft assembled at the Shuttle/space station terminus to other earth orbit locations. It is assumed that such a capability will be initially available in the early 1990's. The benefits provided by satellite servicing in orbit are discussed, taking into account extended satellite lifetimes, lower acquisition cost, improved satellite performance, the possibility to change a satellite's mission, optimized science, and higher satellite reliability. The requirements for Satellite Servicing Technology Development Missions (TDMs) are considered. It is found that existing technology is insufficient, in various areas, to perform the servicing operations. A list is provided of critical technologies which must be developed.

Middleton, R.↗