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Contents include the foloving: 1. Spacecraft Fire Prevention, Detection, and Suppression. 2. Advanced Life Support. Air/water revitalization, waste management. 3. In Situ Resource Utilization (ISRU). Fuel/consumables from regolith/atmosphere. 4. Extra vehicular Activity. Air revitalization, power systems (MEMS scale combustors). 5. In-situ Fabrication and Repair.Of these we have the lead responsibility in Fire Safety.
In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or regression rates as a function of the head-end oxidizer mass flux, G=rho(e)U(e), and the chamber pressure. Furthermore, computation of the full slab burner configuration has also been obtained for various stages of the burn. Comparisons with available experimental data from small scale tests conducted by General Dynamics-Thiokol-Rocketdyne suggest reasonable agreement in the predicted regression rates. Future work will include: (1) a model for soot generation in the flame for more quantitative radiative transfer modelling, (2) a parametric study of combustion efficiency, and (3) transient calculations to help determine the possible mechanisms responsible for combustion instability in hybrid rocket motors.
Phase 1 of the Lean Premixed-Prevaporized Comubstor Design Study is a nine month analytical study effort with no experimental or testing activities included. The program has the objective to design and analyze advanced combustor concepts with features for fuel premixing and prevaporization upstream of the combustion zone for use in future subsonic aircrafts with features for fuel premixing and prevaporization upstream of the combustion zone for use in future subsonic aircraft engines. All of the designs also embody some form of variable geometry for combustor flow modulation. The primary criterion for these designs is low oxides of nitrogen emissions at stratospheric cruise conditions. Four combustor concepts are being designed for the NASA/GE Energy Efficient Engine (EEE) envelope and cycle. Current status of the program is that the four concepts sized for the EEE were designed and are currently undergoing analysis and evaluation.
Program cost and planning data based on 1980 technology and shown in 1979 dollars for a 20K lb Thrust Staged Combustion Cycle Engine are presented. These data were compared with those for the Advanced Expander Cycle Engine at 10K lb and 20K lb thrust levels.
For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.
Long term durability testing of candidate ceramic materials for advanced gas turbine applications was performed. Specimens of the materials were exposed to combustion atmospheres under temperature cyclic conditions for times up to 3500 hours and at peak temperatures of 1200 and 1370 C. The specimen strengths after exposure were measured at room temperature and at 1200 C. In addition, a number of physical measurements were taken; however, no microstructural examination of the specimens was performed. Materials examined include: sintered alpha silicon carbide, sintered silicon nitride, and reaction bonded silicon nitride. Sample preparation techniques were developed to provide high quality thin-section specimens which can be examined by transmitted light or thinned using an ion mill for subsequent transmission electron microscopic examination. Specimen preparation techniques and the results of optical and electron microscopic examination of the reaction interfaces of the selected specimens are presented.
An overview of the Space Station Freedom is given. Its modules are discussed and illustrated along with its microgravity research facilities. These facilities include the advanced protein crystal growth facility, the containerless processing facility, a furnace facility, a combustion facility, and a fluid physics/dynamics facility. The topic of living in space is also addressed.
For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.
For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.
This report is a compilation of the FY 1992 Principal Investigator program task descriptions funded by the Microgravity Science and Applications Division (MSAD), NASA Headquarters, Washington, DC. The document also provides a bibliography of FY 1992 publications and presentations cited by MSAD Principal Investigators, and an index of the Principal Investigators and their affiliations. The purpose of the document is to provide an overview and progress report for the funded tasks for scientists and researchers in industry, university, and government communities. The tasks are grouped into three categories appropriate to the type of research being done-space flight, ground based, and advanced technology development-and by science discipline. The science disciplines are: biotechnology, combustion science,, electronic materials, fluid physics, fundamental physics, glass and ceramics, metals and alloys, and protein crystal growth.
Key elements of the microgravity research program as conducted by the Microgravity Science and Applications Division (MSAD) within the Office of Space Science and Applications (OSSA) during fiscal year (FY) 1992 are described. This NASA funded program supported investigators from the university, industry, and government research communities. The program's goals, the approach taken to achieve those goals, and the resources that were available are summarized. It provides a 'snapshot' of the Program's status at the end of FY 1992 and reviews highlights and progress in the ground and flight-based research during the year. It also describes four major space missions that flew during FY 1992, the advanced technology development (ATD) activities, and the plans to use the research potential of Space Station Freedom and other advanced carriers. The MSAD program structure encompassed five research areas: (1) Biotechnology, (2) Combustion Science, (3) Fluid Physics, (4) Materials Science, and (5) Benchmark Physics.
A fire, whether in a spacecraft or in occupied spaces on extraterrestrial bases, can lead to mission termination or loss of life. While the fire-safety record of US space missions has been excellent, the advent of longer duration missions to Mars, the moon, or aboard the International Space Station (ISS) increases the likelihood of fire events, with more limited mission termination options. The fire safety program of NASA's manned space flight program is based largely upon the principles of controlling the flammability of on-board materials and greatly eliminating sources of ignition. As a result, very little research has been conducted on fire suppression in the microgravity or reduced-gravity environment. The objectives of this study are: to obtain fundamental knowledge of physical and chemical processes of fire suppression, using gravity and oxygen concentration as independent variables to simulate various extraterrestrial environments, including spacecraft and surface bases in Mars and moon missions; to provide rigorous testing of analytical models, which include comprehensive descriptions of combustion and suppression chemistry; and to provide basic research results useful for technological advances in fire safety, including the development of new fire-extinguishing agents and approaches, in the microgravity environment associated with ISS and in the partial-gravity Martian and lunar environments.
The mission of the International Space Station is to provide a working laboratory in orbit for research in engineering, life sciences, and microgravity. Among the microgravity disciplines that are preparing to utilize this international resource are materials processing, combustion, fluid dynamics, biotechnology, and fundamental physics. The Station promises to enable significant advances in each of these areas by making available a research facility in which gravitational and other accelerations, and their corresponding buoyancy and diffusion effects on various physical processes, are orders of magnitude lower than they are on Earth. In order to fulfill this promise, it is not enough for the Space Station to simply replicate a typical terrestrial scientific laboratory in orbit. Although an orbiting laboratory is free of most of the effects of gravitational acceleration by virtue of its free fall condition, it also produces structural vibration or jitter that can interfere with the processes under study. To ensure the quality of the acceleration environment and enable a successful mission, the Space Station Program has limited potential disturbances in two ways: first, by isolating the most sensitive payloads from the vehicle structure, and second, by quieting major disturbances at their sources. The first area, payload isolation, is implemented inside the pressurized modules at the rack level. Sub-rack level isolators have also been developed. This paper addresses the second area, disturbance source limits, for one of the major sources of mechanical noise on the Space Station: the Solar Alpha Rotary Joints. Due to the potential for large disturbances to the microgravity environment, an initial analytical prediction of rotary joint vibration output was made. Key components were identified and tested to validate the analytical predictions. Based on the component test results, the final vibration output of the joints was verified by a test on each fully assembled flight unit. This paper describes the Space Station microgravity requirements, the rotary joint hardware, and the disturbance producing aspects of joint operation. The test setup, instrumentation, test conditions, and results for the component level and system level measurements are described. An overall forcing function that describes the maximum torque imparted to the Station is created based on the test results, and these disturbances are shown to meet the applicable Space Station microgravity requirements.
Topics include: Single-Photon-Sensitive HgCdTe Avalanche Photodiode Detector; Surface-Enhanced Raman Scattering Using Silica Whispering-Gallery Mode Resonators; 3D Hail Size Distribution Interpolation/Extrapolation Algorithm; Color-Changing Sensors for Detecting the Presence of Hypergolic Fuels; Artificial Intelligence Software for Assessing Postural Stability; Transformers: Shape-Changing Space Systems Built with Robotic Textiles; Fibrillar Adhesive for Climbing Robots; Using Pre-Melted Phase Change Material to Keep Payloads in Space Warm for Hours without Power; Development of a Centrifugal Technique for the Microbial Bioburden Analysis of Freon (CFC-11); Microwave Sinterator Freeform Additive Construction System (MS-FACS); DSP/FPGA Design for a High-Speed Programmable S-Band Space Transceiver; On-Chip Power-Combining for High-Power Schottky Diode-Based Frequency Multipliers; FPGA Vision Data Architecture; Memory Circuit Fault Simulator; Ultra-Compact Transputer-Based Controller for High-Level, Multi-Axis Coordination; Regolith Advanced Surface Systems Operations Robot Excavator; Magnetically Actuated Seal; Hybrid Electrostatic/Flextensional Mirror for Lightweight, Large-Aperture, and Cryogenic Space Telescopes; System for Contributing and Discovering Derived Mission and Science Data; Remote Viewer for Maritime Robotics Software; Stackfile Database; Reachability Maps for In Situ Operations; JPL Space Telecommunications Radio System Operating Environment; RFI-SIM: RFI Simulation Package; ION Configuration Editor; Dtest Testing Software; IMPaCT - Integration of Missions, Programs, and Core Technologies; Integrated Systems Health Management (ISHM) Toolkit; Wind-Driven Wireless Networked System of Mobile Sensors for Mars Exploration; In Situ Solid Particle Generator; Analysis of the Effects of Streamwise Lift Distribution on Sonic Boom Signature; Rad-Tolerant, Thermally Stable, High-Speed Fiber-Optic Network for Harsh Environments; Towed Subsurface Optical Communications Buoy; High-Collection-Efficiency Fluorescence Detection Cell; Ultra-Compact, Superconducting Spectrometer-on-a-Chip at Submillimeter Wavelengths; UV Resonant Raman Spectrometer with Multi-Line Laser Excitation; Medicine Delivery Device with Integrated Sterilization and Detection; Ionospheric Simulation System for Satellite Observations and Global Assimilative Model Experiments - ISOGAME; Airborne Tomographic Swath Ice Sounding Processing System; flexplan: Mission Planning System for the Lunar Reconnaissance Orbiter; Estimating Torque Imparted on Spacecraft Using Telemetry; PowderSim: Lagrangian Discrete and Mesh-Free Continuum Simulation Code for Cohesive Soils; Multiple-Frame Detection of Subpixel Targets in Thermal Image Sequences; Metric Learning to Enhance Hyperspectral Image Segmentation; Basic Operational Robotics Instructional System; Sheet Membrane Spacesuit Water Membrane Evaporator; Advanced Materials and Manufacturing for Low-Cost, High-Performance Liquid Rocket Combustion Chambers; Motor Qualification for Long-Duration Mars Missions.
The National Aerospace Plane is a horizontal takeoff and landing, single stage-to-orbit vehicle using hydrogen fuel. The first flights are planned for the mid 1990's. The success of this important national program requires advances in virtually every discipline associated with both airbreathing and space flight. The high heating value, cooling capacity, and combustion properties make hydrogen the fuel of choice, but low density results in a large vehicle. Both fuel cooling capacity and density are increased with the use of slush hydrogen and result in significant reductions in vehicle size. A national program to advance this technology and to find engineering solutions to the many design issues is now under way. The program uses the expertise of the cryogenics production and services industry, the instrumentation industry, universities and governments. The program will be discussed to highlight the major issues and display progress to date.
The overall objective of this project was to evaluate the advantages of transformational polybenzimidazole (PBI) polymer hollow-fiber membrane (HFM)-based, carbon dioxide (CO 2 ) capture and purification technology at bench-scale using an actual coal-derived syngas stream from a coal gasification facility. The project was carried out over two budget periods. The technical objectives in Budget Period 1 (BP1) included preparing HFs and modules and upgrading the available skid for field testing. The technical objectives for BP2 were to field-test the skid unit with actual coal-derived syngas from an oxygen-blown gasifier to obtain performance data, update the Techno-Economic Analysis (TEA) that would assist with future process scale-up, and provide information on the design of a small pilot-scale test unit. The goal was to advance the PBI-HFM CO 2 capture and gas separation system for pre-combustion applications beyond second-generation economic performance predictions and make progress toward meeting overall fossil energy performance goals of CO 2 capture with 95% CO 2 purity at a cost of electricity (COE) 30% less than baseline capture approaches. The research program was designed with progressive technical tasks leading to both dynamic and steady-state testing of the PBI-HFM skid with actual coal-derived syngas. The work plan was to: (1) fabricate sufficient Generation-2 (GEN-2) fibers for module fabrication; (2) upgrade the fiber skid to accommodate large fiber modules for bench-scale field testing; (3) conduct dynamic and steady-state testing with coal-derived syngas from an oxygen-blown gasifier and obtain system performance data; (4) perform a TEA and environmental, health, and safety (EH&S) assessment; (5) update the State-Point Data Table, Technology Gap Analysis (TGA), and Technology Maturation Plan (TMP); (6) uninstall and return the test skid to the Recipient’s facilities; and (7) submit a Final Report that describes the results and analysis of the project research effort.
Radiation-cooled bipropellant thrust chambers are being considered for ascent/ descent engines and reaction control systems on various NASA missions and spacecraft, such as the Mars Sample Return and Orion Multi-Purpose Crew Vehicle (MPCV). Currently, iridium (Ir)-lined rhenium (Re) combustion chambers are the state of the art for in-space engines. NASA's Advanced Materials Bipropellant Rocket (AMBR) engine, a 150-lbf Ir-Re chamber produced by Plasma Processes and Aerojet Rocketdyne, recently set a hydrazine specific impulse record of 333.5 seconds. To withstand the high loads during terrestrial launch, Re chambers with improved mechanical properties are needed. Recent electrochemical forming (EL-Form"TM") results have shown considerable promise for improving Re's mechanical properties by producing a multilayered deposit composed of a tailored microstructure (i.e., Engineered Re). The Engineered Re processing techniques were optimized, and detailed characterization and mechanical properties tests were performed. The most promising techniques were selected and used to produce an Engineered Re AMBR-sized combustion chamber for testing at Aerojet Rocketdyne.
A combustor program was conducted to evolve and to identify the technology needed for, and to establish the credibility of, using combustors with catalytic reactors in modern high-pressure-ratio aircraft turbine engines. Two selected catalytic combustor concepts were designed, fabricated, and evaluated. The combustors were sized for use in the NASA/General Electric Energy Efficient Engine (E3). One of the combustor designs was a basic parallel-staged double-annular combustor. The second design was also a parallel-staged combustor but employed reverse flow cannular catalytic reactors. Subcomponent tests of fuel injection systems and of catalytic reactors for use in the combustion system were also conducted. Very low-level pollutant emissions and excellent combustor performance were achieved. However, it was obvious from these tests that extensive development of fuel/air preparation systems and considerable advancement in the steady-state operating temperature capability of catalytic reactor materials will be required prior to the consideration of catalytic combustion systems for use in high-pressure-ratio aircraft turbine engines.