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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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Various system aspects of a 300-kW nuclear reactor power system for spacecraft have been investigated. Special attention is given to the cases of a reusable OTV and a space-based radar. It is demonstrated that the stowed length of the power system is important to mission design, and that orbital storage for months to years may be needed for missions involving orbital assembly.
This publication presents a thermophysical property survey on materials that could potentially be used for future spacecraft thermal protection systems (TPS). This includes data that was reported in the 1960's as well as more current information reported through the 1980's. An attempt was made to cite the manufacturers as well as the data source in the bibliography. This volume represents an attempt to provide in a single source a complete set of thermophysical data on a large variety of materials used in spacecraft TPS analysis. The property data is divided into two categories: ablative and reusable. The ablative materials have been compiled into twelve categories that are descriptive of the material composition. An attempt was made to define the Arrhenius equation for each material although this data may not be available for some materials. In a similar manner, char data may not be available for some of the ablative materials. The reusable materials have been divided into three basic categories: thermal protection materials (such as insulators), adhesives, and structural materials.
The development of the Space Shuttle Orbiter thermal protection system heating environment is described from a design stand point that began in the early 1970s. The desire for a light weight, reusable heat shield required the development of new technology, relative to previous manned spacecraft, and a systems approach to the design of the vehicle, entry guidance, and thermal protection system. Several unanticipated issues had to be resolved in both the entry and ascent phases of flight, which are discussed at a high level. During the life of the Program, significant improvements in computing power and numerical methods have been applied to Space Shuttle aerodynamic and aerothermodynamic issues, with the Shuttle Program often being the motivation, and or sponsor of the analysis development.
Solar Electric Propulsion (SEP) when used for station keeping and final orbit insertion has been shown to increase a geostationary satellite's payload when launched by existing expendable launch vehicles. In the case of reusable launch vehicles or expendable launch vehicles where an upper stage is an expensive option, this methodology can be modified by using the existing on-board apogee chemical system to perform a perigee burn and then letting the electric propulsion system complete the transfer to geostationary orbit. The elimination of upper stages using on-board chemical and electric propulsion systems was thus examined for GEO spacecraft. Launch vehicle step-down from an Atlas IIAR to a Delta 7920 (no upper stage) was achieved using expanded on-board chemical tanks, 40 kW payload power for electric propulsion, and a 60 day elliptical to GEO SEP orbit insertion. Optimal combined chemical and electric trajectories were found using SEPSPOT. While Hall and ion thrusters provided launch vehicle step-down and even more payload for longer insertion times, NH3 arcjets had insufficient performance to allow launch vehicle step-down. Degradation levels were only 5% to 7% for launch step-down cases using advanced solar arrays. Results were parameterized to allow comparisons for future reusable launch vehicles. Results showed that for an 8 W/kg initial power/launch mass power density spacecraft, 50% to 100% more payload can be launched using this method.
Fuel container design for space vehicles, discussing impact protection, reusability and long term cryogenic storage
The goal of this project is to design a reusable safety device for a waist tether which will absorb the kinetic energy of an astronaut drifting away from the Space Station. The safety device must limit the tension of the tether line in order to prevent damage to the astronaut's space suit or to the structure of the spacecraft. The tether currently used on shuttle missions must be replaced after the safety feature has been developed. A reusable tether for the Space Station would eliminate the need for replacement tethers, conserving space and mass. This report presents background information, scope and limitations, methods of research and development, alternative designs, a final design solution and its evaluation, and recommendations for further work.
The packaging of metal wool for reusable thermal heat shields applied to aerodynamic and other surfaces for the space shuttle was analyzed and designed, and samples were fabricated and experimentally studied. Parametric trends were prepared for selected configurations. An all-metal thermally efficient, reliable, reusable and producible heat shield system was designed and structurally tested for use on spacecraft aerodynamic surfaces where temperatures do not exceed 810 K. Stainless steel sheet, primarily for structure and secondarily in the transverse plane for thermal expansion, was shown to accommodate thermal expansion in all directions when restrained at the edges and heated to 1360 K. Aerodynamic loads of 0.35 x 1000,000 newtons/sq meter, and higher, may be easily accepted by structures of this design. Seven all-metal thermal protection specimens, 12.7 cm square and 2.5 cm thick were fabricated and are being experimentally evaluated at simulated shuttle entry conditions in an arc jet facility.
Modem aircraft, UAVs, and robotic spacecraft pose substantial requirements on controllers in the light of ever increasing demands for reusability, affordability, and reliability. The individual systems (which are often nonlinear) must be controlled safely and reliably in environments where it is virtually impossible to analyze-ahead of time- all the important and possible scenarios and environmental factors. For example, system components (e.g., gyros, bearings of reaction wheels, valves) may deteriorate or break during autonomous UAV operation or long-lasting space missions, leading to a sudden, drastic change in vehicle performance. Manual repair or replacement is not an option in such cases. Instead, the system must be able to cope with equipment failure and deterioration. Controllability of the system must be retained as good as possible or re-established as fast as possible with a minimum of deactivation or shutdown of the system being controlled. In such situations the control engineer has to employ adaptive control systems that automatically sense and correct themselves whenever drastic disturbances and/or severe changes in the plant or environment occur.
This paper presents a software engineering perspective to designing and building fault protection monitor software for spacecraft.
Earlier studies show that future human explorations missions are composed of multi-vehicle assemblies with interconnected electric power systems. Some vehicles are often intended to serve as flexible multi-purpose or multi-mission platforms. This drives the need for power architectures that can be reconfigured to support this level of flexibility. Power system developmental costs can be reduced, program wide, by utilizing a common set of modular building blocks. Further, there are mission operational and logistics cost benefits of using a common set of modular spares. These benefits are the goals of the Advanced Exploration Systems (AES) Modular Power System (AMPS) project. A common set of modular blocks requires a substantial level of standardization in terms of the Electrical, Data System, and Mechanical interfaces. The AMPS project is developing a set of proposed interface standards that will provide useful guidance for modular hardware developers but not needlessly constrain technology options, or limit future growth in capability. In 2015 the AMPS project focused on standardizing the interfaces between the elements of spacecraft power distribution and energy storage. The development of the modular power standard starts with establishing mission assumptions and ground rules to define design application space. The standards are defined in terms of AMPS objectives including Commonality, Reliability-Availability, Flexibility-Configurability and Supportability-Reusability. The proposed standards are aimed at assembly and sub-assembly level building blocks. AMPS plans to adopt existing standards for spacecraft command and data, software, network interfaces, and electrical power interfaces where applicable. Other standards including structural encapsulation, heat transfer, and fluid transfer, are governed by launch and spacecraft environments and bound by practical limitations of weight and volume. Developing these mechanical interface standards is more difficult but an essential part of defining physical building blocks of modular power. This presentation describes the AMPS projects progress towards standardized modular power interfaces.
Aerogel composites formed by infiltrating organic and/or inorganic aerogels into fiber matrix materials enable us to exploit the low thermal conductivity and low density of aerogels while maintaining the strength, structure and other useful properties of a porous fiber matrix. New materials for extreme heating ranges are needed to insulate future spacecraft against the extreme heat of planetary atmospheric entry, but the insulation mass must be minimized in order to maximize the payload. A reusable system passively insulates to survive heating unchanged for relatively low heating. Ablators, which sacrifice mass to control heating, are used to protect vehicles against more extreme heating for a single use thermal protection system (TPS). Aerogel composites were fabricated and tested for spacecraft thermal protection. The high-temperaturey high heat flux tests described in this paper were performed in NASA Ames arc-jet facilities to simulate spacecraft atmospheric entry, and include heating conditions predicted for the forebody and backshell of the Mars Science Lander (MSL) entry probe. The aerogel composites tested showed excellent thermal performance in the arc-jet tests, functioning both as reusuable insulation under lower heat fluxes, and as ablative aerogels under the extreme heating predicted for the MSL forebody.
To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined that utilize electric propulsion supported by the nuclear reactor's power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the orbital transfer vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925-km altitude, 28.5-deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an orbital maneuvering vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO for transfer by OMV to the Shuttle. OTV propellant is resupplied, and the ion thrusters are replaced, by the OMV before each sortie to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg, with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo or by changing to mercury propellant.
Conceptual tank with replaceable liner holds liquid propellant and is suitable for reuse in spacecraft propulsion systems. Main components are: composite outer shell for containment of pressure, replaceable membrane liner, and foam-filled honeycomb sandwich bonded to inside of tank to insulate and stabilize wall of tank. Concept maximizes potential of filament-wound composite pressure shells with strength-to-density ratios five times those of comparable aluminum shells. Reduction in weight as much as 50 percent.
This paper describes the multi-mission Dshell++ simulation framework for high fidelity, physics-based simulation of spacecraft, robotic manipulation and mobility systems. Dshell++ is a C++/Python library which uses modern script driven object-oriented techniques to allow component reuse and a dynamic run-time interface for complex, high-fidelity simulation of spacecraft and robotic systems. The goal of the Dshell++ architecture is to manage the inherent complexity of physicsbased simulations while supporting component model reuse across missions. The framework provides several features that support a large degree of simulation configurability and usability.
The Delta launch vehicle has played a significant role in spaceand airborne studies since its first launch in 1960, and a high volume period of service is planned for the 1980s. The historical role played by Delta in launching satellites from 1960 to the present is discussed, and vehicle modifications made during this period are summarized. It is shown that out of 154 launches, 143 proved successful for an overall reliability of 93%. The forecasted launch schedule through 1985 is also presented. Various modifications are now under way to provide spacecraft interchangeability with the Shuttle: a Payload Assist Module (PAM) is proposed to provide an orderly transition from the Delta expendable vehicle to the Shuttle reusable vehicle; the new Delta 3920 Improved Second Stage is the result of a need for improved Delta performance to meet 3910 payload capabilities; the firing sequence of the solid rocket motors was altered from five at liftoff and four during ascent to a sequence of six and three, thereby increasing spacecraft weight in geosynchronous transfer orbit. Potential future improvements discussed include the Delta 4920, 9-ft-diam fairing, booster engine performance, PAM solid motor performance, a universal second stage, a hydrogen-oxygen second stage, and large strap-on solids.
The concept of reusing spacecraft and launch vehicles is currently exemplified by the Space Shuttle Orbiter and the two recoverable and reusable solid rocket boosters (SRBs) which are part of its launch system. It is shown that it is possible to recover launch systems weighing as much as one million pounds at reasonable impact velocities using a hybrid parachute-retrorocket recovery system. The new Kevlar materials, which have a much higher strength-to-weight ratio than nylon materials, make it possible to recover a one-million-pound booster for a reasonable weight penalty. The selected input parameters to the parachute weight equation result in an accurate estimation of the current Space Shuttle solid rocket booster parachute recovery system.
There are many future science instruments with high performance pointing (sub microradian) requirements. To build a separate spacecraft for each payload is prohibitively expensive, especially as not all instruments need to be in space for a long duration. Putting multiple payloads on a single basebody that supplies power, communications, and orbit maintenance is cheaper, easier to service, and allows for the spacecraft bus to be reused as new instruments become available to replace old instruments. Once several payloads are mounted together, the articulation of one may disturb another. The situation is even more extreme when the basebody serves multiple purposes, such as space station which has construction, satellite servicing, and man motion adding to the disturbance environment. The challenge then is to maintain high performance at low cost in a multiple payload environment. The goal is to supply many future science instruments with high performance pointing (sub microradian). The options are independent spacecraft for each payload (expensive); or multiple payloads on a single basebody (cheaper, easier to service, basebody reusable for several short duration payloads). The problems are one payload can disturb another, and other activities create large disturbances (construction, satellite servicing, and man motion).
The Long Duration Exposure Facility (LDEF) is a reusable hollow-cylindrical satellite sustaining a total of 57 different experiments. The 130 sq m of spacecraft surface area included anodized 6061-T6 Al alloy bay frames and clamps for holding experiment trays in the bay areas. Attention is presently given to the micrometeoroid impact crater features observed on two tray clamps recovered from the LDEF leading-edge locations. It is found that even very subtle surface modifications in structural alloy anodizing can influence micrometeoroid impact crater cracking, notable radial cracking due to the ejecta-rim of the impact craters.