SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
This paper provides a Safety Checklist for use during the analysis of software requirements for spacecraft and other safety-critical, embedded systems.
The Microgravity Science Glovebox (MSG) is a double rack facility aboard the International Space Station (ISS) designed for gravity-dependent phenomena investigation handling. The MSG has been operating in the ISS US Laboratory Module since July 2002. The MSG facility provides an enclosed working area for investigation manipulation and observation, The MSG's unique design provides two levels of containment to protect the ISS crew from hazardous operations. Research investigations operating inside the MSG are provided a large 255 liter work volume, 1000 watts of dc power via a versatile supply interface (120, 28, +/-12, and 5 Vdc), 1000 watts of cooling capability, video and data recording and real time downlink, ground commanding capabilities, access to ISS Vacuum Exhaust and Vacuum Resource Systems, and gaseous nitrogen supply. With these capabilities, the MSG is an ideal platform for research required to advance the technology readiness levels (TRL) needed for the Crew Exploration Vehicle and the Exploration Initiative. Areas of research that will benefit from investigations in the MSG include thermal management, fluid physics, spacecraft fire safety, materials science, combustion, reaction control systems, in situ fabrication and repair, and advanced life support technologies. This paper will provide a detailed explanation of the MSG facility, a synopsis of the research that has already been accomplished in the MSG and an overview of investigations planning to operate in the MSG. In addition, this paper will address possible changes to the MSG utilization process that will be brought about by the transition to ISS as a National Laboratory.
Appropriate design of fire detection systems requires knowledge of both the expected fire signature and the background aerosol levels. Terrestrial fire detection systems have been developed based on extensive study of terrestrial fires. Unfortunately there is no corresponding data set for spacecraft fires and consequently the fire detectors in current spacecraft were developed based upon terrestrial designs. In low gravity, buoyant flow is negligible which causes particles to concentrate at the smoke source, increasing their residence time, and increasing the transport time to smoke detectors. Microgravity fires have significantly different structure than those in 1-g which can change the formation history of the smoke particles. Finally the materials used in spacecraft are different from typical terrestrial environments where smoke properties have been evaluated. It is critically important to detect a fire in its early phase before a flame is established, given the fixed volume of air on any spacecraft. Consequently, the primary target for spacecraft fire detection is pyrolysis products rather than soot. Experimental investigations have been performed at three different NASA facilities which characterize smoke aerosols from overheating common spacecraft materials. The earliest effort consists of aerosol measurements in low gravity, called the Smoke Aerosol Measurement Experiment (SAME), and subsequent ground-based testing of SAME smoke in 55-gallon drums with an aerosol reference instrument. Another set of experiments were performed at NASAs Johnson Space Center White Sands Test Facility (WSTF), with additional fuels and an alternate smoke production method. Measurements of these smoke products include mass and number concentration, and a thermal precipitator was designed for this investigation to capture particles for microscopic analysis. The final experiments presented are from NASAs Gases and Aerosols from Smoldering Polymers (GASP) Laboratory, with selected results focusing on realistic fuel preparations and heating profiles with regards to early detection of smoke. SAFFIRE is the upcoming large-scale fire experiment which will be executed in a Cygnus vehicle after it undocks from the ISS.
Hydrazine and its derivatives have dominated the class of hypergolic liquid propellants for bipropellant propulsion systems in rockets such as the Titan, MX Missile, and Ariane; it has also been widely utilized as a monopropellant in auxiliary power units and in thrusters for altitude and in-orbit control of satellites and spacecraft. With continued use of hydrazine in current and future spacecraft and payloads, it is necessary to understand the historical and current states of synthesis for the commodity and possible implications that may arise from changes in production processes for the United States stock.
Development of reliable power-by-wire actuation systems for both aeronautical and space applications has been sought recently to eliminate hydraulic systems from aircraft and spacecraft and thus improve safety, efficiency, reliability, and maintainability. The Electrically Powered Actuation Design (EPAD) program was a joint effort between the Air Force, Navy, and NASA to develop and fly a series of actuators validating power-by-wire actuation technology on a primary flight control surface of a tactical aircraft. To achieve this goal, each of the EPAD actuators was installed in place of the standard hydraulic actuator on the left aileron of the NASA F/A-18B Systems Research Aircraft (SRA) and flown throughout the SRA flight envelope. Numerous parameters were recorded, and overall actuator performance was compared with the performance of the standard hydraulic actuator on the opposite wing. This paper discusses the integration and testing of the EPAD electromechanical actuator (EMA) on the SRA. The architecture of the EMA system is discussed, as well as its integration with the F/A-18 Flight Control System. The flight test program is described, and actuator performance is shown to be very close to that of the standard hydraulic actuator it replaced. Lessons learned during this program are presented and discussed, as well as suggestions for future research.
Development of reliable power-by-wire actuation systems for both aeronautical and space applications has been sought recently to eliminate hydraulic systems from aircraft and spacecraft and thus improve safety, efficiency, reliability, and maintainability. The Electrically Powered Actuation Design (EPAD) program was a joint effort between the Air Force, Navy, and NASA to develop and fly a series of actuators validating power-by-wire actuation technology on a primary flight control surface of a tactical aircraft. To achieve this goal, each of the EPAD actuators was installed in place of the standard hydraulic actuator on the left aileron of the NASA F/A-18B Systems Research Aircraft (SRA) and flown throughout the SRA flight envelope. Numerous parameters were recorded, and overall actuator performance was compared with the performance of the standard hydraulic actuator on the opposite wing. This paper discusses the integration and testing of the EPAD electromechanical actuator (EMA) on the SRA. The architecture of the EMA system is discussed, as well as its integration with the F/A-18 Flight Control System. The flight test program is described, and actuator performance is shown to be very close to that of the standard hydraulic actuator it replaced. Lessons learned during this program are presented and discussed, as well as suggestions for future research.
The potential for collisions or close passes with other space objects presents a significant issue for many space tether applications, representing a potential risk both to the integrity of the tether system and t o the safety of other spacecraft. Potential collisions between tethers and other space objects may be possible to avoid if close encounters can be predicted with sufficient precision and advance notice. In order to provide a method for predicting the frequency with which a tether must be maneuvered to avoid collisions, and to provide a resource for accurate close-encounter prediction during tether flight experiments, we have developed a software tool that compares the trajectory of a tether object with that of all of the objects in the NORAD space catalogue. In this paper we describe the models and algorithms used in this tool, and discuss results of test cases conducted to predict the close-encounter frequency of a tether systems ranging from a short nanosatellite-based tether experiment to a hundred-kilometer long MXER tether system.
The objective of this project is the testing and evaluation of the effectiveness of a variety of fire suppressants and fire-response techniques that will be used in the next generation of spacecraft (Crew Exploration Vehicle, CEV) and planetary habitats. From the many lessons learned in the last 40 years of space travel, there is common agreement in the spacecraft fire safety community that a new fire suppression system will be needed for the various types of fire threats anticipated in new space vehicles and habitats. To date, there is no single fire extinguishing system that can address all possible fire situations in a spacecraft in an effective, reliable, clean, and safe way. The testing conducted under this investigation will not only validate the various numerical models that are currently being developed, but it will provide new design standards on fire suppression that can then be applied to the next generation of spacecraft extinguishment systems. The test program will provide validation of scaling methods by conducting small, medium, and large scale fires. A variety of suppression methods will be tested, such as water mist, carbon dioxide, and nitrogen with single and multiple injection points and direct or distributed agent deployment. These injection methods cover the current ISS fire suppression method of a portable hand-held fire extinguisher spraying through a port in a rack and also next generation spacecraft units that may have a multi-point suppression delivery system built into the design. Consideration will be given to the need of a crew to clean-up the agent and recharge the extinguishers in flight in a long-duration mission. The fire suppression methods mentioned above will be used to extinguish several fire scenarios that have been identified as the most relevant to spaceflight, such as overheated wires, cable bundles, and circuit boards, as well as burning cloth and paper. Further testing will be conducted in which obstructions and ventilation will be added to represent actual spacecraft conditions (e.g., a series of cards in a card rack).
Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using the Goddard Trajectory Determination System (GTDS) and a real-time extended Kalman filter estimation system to process Tracking Data and Relay Satellite (TDRS) System (TDRSS) measurements in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. GTDS is the operational orbit determination system used by the FDD, and the extended Kalman fliter was implemented in an analysis prototype system, the Real-Time Orbit Determination System/Enhanced (RTOD/E). The Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generates an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the Geodynamics (GEODYN) orbit determination system with laser ranging tracking data. The TOPEX/Poseidon trajectories were estimated for the October 22 - November 1, 1992, timeframe, for which the latest preliminary POD results were available. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch cases were assessed using overlap comparisons, while the sequential cases were assessed with covariances and the first measurement residuals. The batch least-squares and forward-filtered RTOD/E orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 10 meters (m) for the batch least squares and less than 18 m for the sequential estimation solutions. The differences among the POD, GTDS, and RTOD/E solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.
The Microgravity Science Glovebox (MSG) is a double rack facility aboard the International Space Station (ISS) designed for gravity-dependent phenomena investigation handling. The MSG has been operating in the ISS US Laboratory Module since July 2002. The MSG facility provides an enclosed working area for investigation manipulation and observation in the ISS. The MSG s unique design provides two levels of containment to protect the ISS crew from hazardous operations. Research investigations operating inside the MSG are provided a large 255 liter work volume, 1000 watts of dc power via a versatile supply interface (120,28, +/-12, and 5 Vdc), 1000 watts of cooling capability, video and data recording and real time downlink, ground commanding capabilities, access to ISS Vacuum Exhaust and Vacuum Resource Systems, and gaseous nitrogen supply. With these capabilities, the MSG is an ideal platform for research required to advance the technology readiness levels (TRL) needed for the Crew Exploration Vehicle and the Exploration Initiative. Areas of research that will benefit from investigations in the MSG include thermal management, fluid physics, spacecraft fire safety, materials science, combustion and reacting control systems, in situ fabrication and repair, and advanced life support technologies. This paper will provide a detailed explanation of the MSG facility, a synopsis of the research that has already been accomplished in the MSG, an overview of investigations planning to operate in the MSG, and possible augmentations that can be added to the MSG facility to further enhance the resources provided to investigations.
The Microgravity Science Glovebox (MSG) is a double rack facility aboard the International Space Station (ISS) designed for gravity-dependent phenomena investigation handling. The MSG has been operating in the ISS US Laboratory Module since July 2002. The MSG facility provides an enclosed working area for investigation manipulation and observation in the ISS. The MSG's unique design provides two levels of containment to protect the ISS crew from hazardous operations. Research investigations operating inside the MSG are provided a large 255 liter work volume, 1000 watts of dc power via a versatile supply interface (120,28, plus or minus 12, and 5 Vdc), 1000 watts of cooling capability, video and data recording and real time downlink, ground commanding capabilities, access to ISS Vacuum Exhaust' and Vacuum Resource 'Systems, and gaseous nitrogen supply. With these capabilities, the MSG is an ideal platform for research required to advance the technology readiness levels (TRL) needed for the Crew Exploration Vehicle and the Exploration Initiative. Areas of research that will benefit from investigations in the MSG include thermal management, fluid physics, spacecraft fire safety, materials science, combustion and reacting control systems, in situ fabrication and repair, and advanced life support technologies. This paper will provide a detailed explanation of the MSG facility, a synopsis of the research that has already been accomplished in the MSG, an overview of investigations planning to operate in the MSG, and possible augmentations that can be added to-the MSG facility to further enhance the resources provided to investigations.
Manned spacecraft recovery safety noting operational testing of postlanding systems, recovery support and retrieval procedures for Apollo
A program was initiated to study the feasibility of using advanced terrestrial power system protection techniques for spacecraft power systems. It was designed to enhance and automate spacecraft power distribution systems in the areas of safety, reliability and maintenance. The proposed power management/distribution system is described as well as security assessment and control, incipient and low current fault detection, and the proposed spaceborne protection system. It is noted that the intelligent remote power controller permits the implementation of digital relaying algorithms with both adaptive and programmable characteristics.
Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.
The Apollo Mission employs more than 210 pyrotechnic devices per mission.These devices are either automatic of commanded from the Apollo spacecraft systems. All devices require high reliability and safety and most are classified as either crew safety critical or mission critical. Pyrotechnic devices have a wide variety of applications including: launch escape tower separation, separation rocket ignition, parachute deployment and release and electrical circuit opening and closing. This viewgraph presentation identifies critical performance, design requirements and safety measures used to ensure quality, reliability and performance of Apollo pyrotechnic/explosive devices. The major components and functions of a typical Apollo pyrotechnic/explosive device are listed and described (initiators, cartridge assemblies, detonators, core charges). The presentation also identifies the major locations and uses for the devices on: the Command and Service Module, Lunar Module and all stages of the launch vehicle.
Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.
A heat transfer investigation was conducted for liquid and two-phase methane. The tests were conducted at the NASA Glenn Research Center Heated Tube Facility (HTF) using resistively heated tube sections to simulate conditions encountered in regeneratively cooled rocket engines. This testing is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. Nontoxic propellants, such as liquid oxygen/liquid methane (LO2/LCH4), offer potential benefits in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications. Regeneratively cooled thrust chambers are one solution for high performance, robust LO2/LCH4 engines, but cooling data on methane is limited. Several test runs were conducted using three different diameter Inconel 600 tubes, with nominal inner diameters of 0.0225-, 0.054-, and 0.075-in. The mass flow rate was varied from 0.005 to 0.07 lbm/sec. As the current focus of the PCAD project is on pressure fed engines for LO2/LCH4, the average test section outlet pressures were targeted to be 200 psia or 500 psia. The heat flux was incrementally increased for each test condition while the test section wall temperatures were monitored. A maximum average heat flux of 6.2 Btu/in.2 sec was achieved and, at times, the temperatures of the test sections reached in excess of 1800 R. The primary objective of the tests was to produce heat transfer correlations for methane in the liquid and two-phase regime. For two-phase flow testing, the critical heat flux values were determined where the fluid transitions from nucleate boiling to film boiling. A secondary goal of the testing was to measure system pressure drops in the two-phase regime.
The Ice, Cloud and land Elevation Satellite-2 was launched on 15 September 2018 with mission goals of collecting surface elevation data that can precisely measure ice sheet topography, cloud and aerosol heights, land topography, as well as vegetation height for ecosystems studies. The Constraint Analysis and Monitoring System (CAMS) was implemented as an ICESat-2 ground system element to perform Mission Planning and Spacecraft Safety Monitoring. Mission Planning requires CAMS to serve as an interface between the Project Science Office, the Instrument Support Facility and the Mission Operation Control Center. By ingesting inputs received from all three groups, the CAMS builds an optimized and deconflicted timeline of science and instrument activities. Based upon the timeline of activities, the CAMS utilizes a sophisticated set of algorithms to model and predict the location and pointing of the spacecraft instrument relative to the Earth and Sun. From the predicted position and pointing, the CAMS provides precise monitoring of instrument health, and performs space asset laser conjunction detection. Furthermore, the CAMS determines alternative plans to prevent detected health constraint violations or mitigate potential laser conjunctions. This paper provides an overview of the CAMS and presents the operational performance of planning science and instrument activities, the monitoring of instrument constraints for health and safety, and the space asset laser conjunction screening and mitigation process.