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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 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
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 NTRS1996

New TDRSS communications options for small satellites

The NASA uses a space network which includes the tracking and data relay satellite system (TDRSS) for the provision of reliable low data rate and high data rate relay services between user spacecraft in earth orbit and the ground. In relation to future small satellite designs, new communication options for the TDRSS support of small spacecraft missions are reported on. The technologies considered include new transponder technologies, evolution to Ka band frequencies and a TDRSS demand access service capability. Multiple access aspects are considered.

Zillig, David J.↗
NASA NTRS1991

High temperature thruster technology for spacecraft propulsion

A technology program intended to develop high-temperature oxidation-resistant thrusters for spacecraft applications is considered. The program will provide the requisite material characterizations and fabrication to incorporate iridium coated rhenium material into small rockets for spacecraft propulsion. This material increases the operating temperature of thrusters to 2200 C, a significant increase over the 1400 C of the silicide-coated niobium chambers currently used. Stationkeeping class 22 N engines fabricated from iridium-coated rhenium have demonstrated steady state specific impulses 20-25 seconds higher than niobium chambers. These improved performances are obtained by reducing or eliminating the fuel film cooling requirements in the combustion chamber while operating at the same overall mixture ratio as conventional engines.

Schneider, Steven J.↗
NASA NTRS1991

High temperature thruster technology for spacecraft propulsion

A technology program intended to develop high-temperature oxidation-resistant thrusters for spacecraft applications is considered. The program will provide the requisite material characterizations and fabrication to incorporate iridium coated rhenium material into small rockets for spacecraft propulsion. This material increases the operating temperature of thrusters to 2200 C, a significant increase over the 1400 C of the silicide-coated niobium chambers currently used. Stationkeeping class 22 N engines fabricated from iridium-coated rhenium have demonstrated steady state specific impulses 20-25 seconds higher than niobium chambers. These improved performances are obtained by reducing or eliminating the fuel film cooling requirements in the combustion chamber while operating at the same overall mixture ratio as conventional engines.

Schneider, Steven J.↗
NASA NTRS2004

Advanced mobility avionics : a reconfigurable mirco-avionics platform for the future needs of small planetary rovers and micospacecraft

Future small and micro-missions, such as Mars Scouts and Deep Space probes, require a new look at highly integrated, re-configurable, low power avionics. This paper will present our plans for developing a scalable, configurable, and highly integrated 32-bit embedded platform capable of implementing computationally intensive signal processing and control algorithms in space flight instruments and systems. This platform is designed to service the need of both small and large spacecraft and planetary rovers that will operate within moderate radiation environments. Some of the key characteristics of this platform are its small size, low power, high performance, and flexibility. This estimated 10 fold reduction in both size and power over state-of-the-art processing platforms will enable this new product to act as the core of a low-cost mobility system for a wide range of missions.

reconfigurable↗
NASA NTRS1989

The shapes of small satellites

Limb coordinates from spacecraft images have been used to measure the shapes of the small satellites of Mars, Jupiter, Saturn, and Uranus, with a view to the quantitative testing of those shapes' physical significance. A comparison of eight small satellites' ellipsoidal models with equilibrium ellipsoids indicates that the shapes of even moderately irregular satellites are poor predictors of mean densities; icy satellites smaller than about 150 km in mean radius are irregularly shaped, and exhibit no trend in roughness. Larger icy satellites are ellipsoidal, and rocky satellites with mean radius in the 6-110 km range are also irregularly shaped.

Thomas, Peter C.↗
NASA NTRS2013

30-kW SEP Spacecraft as Secondary Payloads for Low-Cost Deep Space Science Missions

The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.

Dawn spacecraft↗
NASA NTRS2014

Optical Properties of Nanosatellite Hardware

Over the last decade, a number of very small satellites have been launched into space. These have been called nanosatellites (generally of a weight between 1 and 10 kg) or picosatellites (weight <1 kg). This also includes CubeSats, which are based on 10-cm cube units. With the addition of the Japanese Experiment Module (JEM) Small Satellite Orbital Deployer (J-SSOD) to the International Space Station (ISS), CubeSats are easily cycled through the JEM airlock and deployed into space (fig. 1). The number of CubeSats launched since 2003 was approaching 100 at the time of publication, and the authors expect this trend in research to continue, particularly for high school and college flight experiments. Because these spacecraft are so small, there is usually no allowance for shielding or active heating or cooling of the avionics and other hardware. Parts that are usually ignored in the thermal analysis of larger spacecraft may contribute significantly to the heat load of a tiny satellite. In addition, many small satellites have commercial-off-the-shelf (COTS) components. To reduce costs, many providers of COTS components do not include the optical and physical parameters necessary for accurate thermal analysis. Marshall Space Flight Center participated in the development and analysis of the Space Missile Defense Command-Operational Nanosatellite Effect (SMDC-ONE) and the Edison Demonstration of Smallsat Networks (EDSN) nanosatellites. These optical property measurements are documented here in hopes that they may benefit future nanosatellite and picosatellite programs and aid thermal analysis to ensure project goals are met, with the understanding that material properties may vary by vendor, batch, manufacturing process, and preflight handling. Where possible, complementary data are provided from ground simulations of the space environment and flight experiments, such as the Materials on International Space Station Experiment (MISSE) series. NASA gives no recommendation, endorsement, or preference, either expressed or implied, concerning materials and vendors used. Solar absorptance was calculated from spectral reflectance measurements made from 250 to 2,800 nm with an AZ Technology Laboratory Portable Spectroreflectometer (LPSR) model 300. ASTM E-903 was the test method used under normal laboratory conditions, and ASTM E-490 was the solar spectral irradiance data used to calculate solar absorptance. Most of the samples were flat, but stray light was minimized as much as possible with either a blackbody or black cloth as sample background. The LPSR has repeatability of approximately +/-1%, where solar absorptance is given as range, that is, from actual measurements taken across the sample. Infrared emittance measurements were made with an AZ Technology TEMP 2000A infrared reflectometer. This instrument measures the total hemispheric reflectance averaged over 3-35 micrometer wavelengths. ASTM E-408 was the test method used under normal laboratory conditions. 3 Stray light was minimized as much as possible. The TEMP 2000A has repeatability of approximately +/-0.5%, where infrared emittance is given as a range, that is, from actual measurements taken across the sample.

Finckenor, M. M.↗
NASA NTRS1995

Autonomous control system reconfiguration for spacecraft with non-redundant actuators

The Small Satellite Technology Initiative (SSTI) 'CLARK' spacecraft is required to be single-failure tolerant, i.e., no failure of any single component or subsystem shall result in complete mission loss. Fault tolerance is usually achieved by implementing redundant subsystems. Fault tolerant systems are therefore heavier and cost more to build and launch than non-redundent, non fault-tolerant spacecraft. The SSTI CLARK satellite Attitude Determination and Control System (ADACS) achieves single-fault tolerance without redundancy. The attitude determination system system uses a Kalman Filter which is inherently robust to loss of any single attitude sensor. The attitude control system uses three orthogonal reaction wheels for attitude control and three magnetic dipoles for momentum control. The nominal six-actuator control system functions by projecting the attitude correction torque onto the reaction wheels while a slower momentum management outer loop removes the excess momentum in the direction normal to the local B field. The actuators are not redundant so the nominal control law cannot be implemented in the event of a loss of a single actuator (dipole or reaction wheel). The spacecraft dynamical state (attitude, angular rate, and momentum) is controllable from any five-element subset of the six actuators. With loss of an actuator the instantaneous control authority may not span R(3) but the controllability gramian integral(limits between t,0) Phi(t, tau)B(tau )B(prime)(tau) Phi(prime)(t, tau)d tau retains full rank. Upon detection of an actuator failure the control torque is decomposed onto the remaining active axes. The attitude control torque is effected and the over-orbit momentum is controlled. The resulting control system performance approaches that of the nominal system.

Grossman, Walter↗
NASA NTRS2023

Small-Scale Magnetic Flux Ropes in Stream Interaction Regions From Parker Solar Probe and Wind Spacecraft Observations

Using in situ measurements from the Parker Solar Probe and Wind spacecraft, we investigate the small-scale magnetic flux ropes (SFRs) and their properties inside stream interaction regions (SIRs). Within SIRs from ∼0.15 to 1 au, SFRs are found to exist in a wide range of solar wind speeds with more frequent occurrences after the stream interface, and the Alfvénicity of these structures decreases significantly with increasing heliocentric distances. Furthermore, we examine the variation of five corresponding SIRs from the same solar sources. The enhancements of suprathermal electrons within these SIRs persist at 1 au and are observed multiple times. An SFR appears to occur repeatedly with the recurring SIRs and is traversed by the Wind spacecraft at least twice. This set of SFRs has similarities in variations of the magnetic field components, plasma bulk properties, density ratio of solar wind alpha and proton particles, and unidirectional suprathermal electrons. We also show, through the detailed time-series plots and Grad–Shafranov reconstruction results, that they possess the same chirality and carry comparable amounts of magnetic flux. Lastly, we discuss the possibility for these recurring SFRs to be formed via interchange reconnection, maintain the connection with the Sun, and survive up to 1 au.

Solar wind↗
DOE OSTI2020

Hypervelocity impact flash expansion geometry under various spacecraft surface electrical conditions

A more congested space environment due to increasing space activities introduces new threats to spacecraft operators. Among these threats include hypervelocity impacts from microparticles (< 1 µg) in the form of meteoroids and space debris. These microparticles travel at speeds between 11.2 and 72.8 km/s with respect to the Earth and can impact spacecraft forming a small (~1 µm) and dense (~ 10 26 m -3 ) plasma. This plasma can generate a strong optical emission (impact flash) and a radio frequency (RF) emission. These emissions can lead to spacecraft electrical anomalies when the impacted spacecraft surface (target) carries high electrical potential due to various space weather effects. To understand the microparticle hypervelocity impact plasma and its associated threat to spacecraft electronics, we studied the hypervelocity impact flash expansion geometry under various spacecraft surface electrical conditions. Three high-speed broadband photomultiplier tubes (PMT) are used to capture the temporal evolution of the impact flash and the temporal evolution of the angular distribution of the impact flash expansion geometry in a 3 MV electrostatic dust accelerator during a ground-based experiment. Overall, results indicate that the impact surface electrical condition has a statistically significant influence on the impact flash geometry, and a strong similarity between the impact plasma and flash expansion geometry. Thus, this paper provides new experimental support for the impact flash origin from and connection with the impact plasma, and serves as the first impact flash expansion geometry characterization under various impact surface electrical conditions in a collisionless environment (< 10 -5 Torr).

42 ENGINEERING↗
NASA NTRS1992

Planetary missions using microspacecraft technology

Some of the planetary missions using microspacecraft of mass on the order of 10 kg which are presently being considered at JPL are discussed. The primary benefits offered by this type of spacecraft is the small and relatively inexpensive launch vehicles and short program development schedules. A mission presently considered by the Asteroid Investigation with Microspacecraft (AIM) study is described with special attention given to the spacecraft configuration; the AIM mission will launch three spacecraft from a single Pegasus vehicle, achieving separate flybys for three near-earth asteroids. Attention is also given to the salient features of the Mars Rover Sample Return mission, the Venus Atmospheric Sounder, and the Comet Nucleus Mission, for which the application of microspacecraft technology is considered.

Short, Kendra L.↗
NASA NTRSDate not supplied

Commanding Constellations (Pipeline Architecture)

Providing ground command software for constellations of spacecraft is a challenging problem. Reliable command delivery requires a feedback loop; for a constellation there will likely be an independent feedback loop for each constellation member. Each command must be sent via the proper Ground Station, which may change from one contact to the next (and may be different for different members). Dynamic configuration of the ground command software is usually required (e.g. directives to configure each member's feedback loop and assign the appropriate Ground Station). For testing purposes, there must be a way to insert command data at any level in the protocol stack. The Pipeline architecture described in this paper can support all these capabilities with a sequence of software modules (the pipeline), and a single self-identifying message format (for all types of command data and configuration directives). The Pipeline architecture is quite simple, yet it can solve some complex problems. The resulting solutions are conceptually simple, and therefore, reliable. They are also modular, and therefore, easy to distribute and extend. We first used the Pipeline architecture to design a CCSDS (Consultative Committee for Space Data Systems) Ground Telecommand system (to command one spacecraft at a time with a fixed Ground Station interface). This pipeline was later extended to include gateways to any of several Ground Stations. The resulting pipeline was then extended to handle a small constellation of spacecraft. The use of the Pipeline architecture allowed us to easily handle the increasing complexity. This paper will describe the Pipeline architecture, show how it was used to solve each of the above commanding situations, and how it can easily be extended to handle larger constellations.

Tim Ray↗
NASA NTRS1991

Disinfectants for spacecraft applications - An overview

The review of disinfectants for use on manned missions emphasizes the need for contamination control to prevent the detrimental effects of bacteria growth on crew health. Microbial control is possible by means of biocides, but the selected product has to meet stringent toxicity requirements for the small environments in spacecraft. The testing and evaluation is described of four biocide candidates: hydrogen peroxide, quaternary ammonium compounds, iodine, and glutaraldehyde. The effectiveness of the disinfectants are analyzed in terms of the ability to treat typical microbial counts from Skylab missions in a closed environment. It is shown that many biocide candidates are not compatible with the ECLSS, water-recovery management, and air-revitalization subsystems of the Space Station Freedom. The use of hydrogen peroxide is proposed with a secondary stronger agent for microbial spills from biological experiments.

Koenig, David W.↗
NASA NTRS2018

Method and Apparatus for Spacecraft Gyroscope Scale Factor Calibration

A method and apparatus for estimating gyro scale factor during normal spacecraft operations, using small attitude motions that are compliant with mission pointing accuracy and stability requirements and a signal processing method that specifically detects the intentionally induced motions. This process increases operational availability by avoiding the need to take the spacecraft offline for large calibration maneuvers.

Reth, Alan D↗
NASA NTRS2003

Catalyzed Combustion In Micro-Propulsion Devices: Project Status

In recent years, there has been a tendency toward shrinking the size of spacecraft. New classes of spacecraft called micro-spacecraft have been defined by their mass, power, and size ranges. Spacecraft in the range of 20 to 100 kg represent the class most likely to be utilized by most small sat users in the near future. There are also efforts to develop 10 to 20 kg class spacecraft for use in satellite constellations. More ambitious efforts will be to develop spacecraft less than 10 kg, in which MEMS fabrication technology is required. These new micro-spacecraft will require new micro-propulsion technology. Although micro-propulsion includes electric propulsion approaches, the focus of this proposed program is micro-chemical propulsion which requires the development of microcombustors. As combustors are scaled down, the surface to volume ratio increases. The heat release rate in the combustor scales with volume, while heat loss rate scales with surface area. Consequently, heat loss eventually dominates over heat release when the combustor size becomes smaller, thereby leading to flame quenching. The limitations imposed on chamber length and diameter has an immediate impact on the degree of miniaturization of a micro-combustor. Before micro-combustors can be realized, such a difficulty must be overcome. One viable combustion alternative is to take advantage of surface catalysis. Micro-chemical propulsion for small spacecraft can be used for primary thrust, orbit insertion, trajectory-control, and attitude control. Grouping micro-propulsion devices in arrays will allow their use for larger thrust applications. By using an array composed of hundreds or thousands of micro-thruster units, a particular configuration can be arranged to be best suited for a specific application. Moreover, different thruster sizes would provide for a range of thrust levels (from N s to mN s) within the same array. Several thrusters could be fired simultaneously for thrust levels higher than the basic units, or in a rapid sequence in order to provide gradual but steady low-g acceleration. These arrays of micro-propulsion systems would offer unprecedented flexibility and redundancy for satellite propulsion and reaction control for launch vehicles. A high-pressure bi-propellant micro-rocket engine is already being developed using MEMS technology. High pressure turbopumps and valves are to be incorporated onto the rocket chip . High pressure combustion of methane and O2 in a micro-combustor has been demonstrated without catalysis, but ignition was established with a spark. This combustor has rectangular dimensions of 1.5 mm by 8 mm (hydraulic diameter 3.9 mm) and a length of 4.5 mm and was operated at 1250 kPa with plans to operate it at 12.7 MPa. These high operating pressures enable the combustion process in these devices, but these pressures are not practical for pressure fed satellite propulsion systems. Note that the use of these propellants requires an ignition system and that the use of a spark would impose a size limitation to this micro-propulsion device because the spark unit cannot be shrunk proportionately with the thruster. Results presented in this paper consist of an experimental evaluation of the minimum catalyst temperature for initiating/supporting combustion in sub-millimeter diameter tubes. The tubes are resistively heated and reactive premixed gases are passed through the tubes. Tube temperature and inlet pressure are monitored for an indication of exothermic reactions and composition changes in the gases.

Sung, C. J.↗