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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Develop a Model Component

During my internship at NASA, I was a model developer for Ground Support Equipment (GSE). The purpose of a model developer is to develop and unit test model component libraries (fluid, electrical, gas, etc.). The models are designed to simulate software for GSE (Ground Special Power, Crew Access Arm, Cryo, Fire and Leak Detection System, Environmental Control System (ECS), etc. ~.) before they are implemented into hardware. These models support verifying local control and remote software for End-Item Software Under Test (SUT). The model simulates the physical behavior (function, state, limits and 110) of each end-item and it's dependencies as defined in the Subsystem Interface Table, Software Requirements & Design Specification (SRDS), Ground Integrated Schematic (GIS), and System Mechanical Schematic.(SMS). The software of each specific model component is simulated through MATLAB's Simulink program. The intensiv~ model development life cycle is a.s follows: Identify source documents; identify model scope; update schedule; preliminary design review; develop model requirements; update model.. scope; update schedule; detailed design review; create/modify library component; implement library components reference; implement subsystem components; develop a test script; run the test script; develop users guide; send model out for peer review; the model is sent out for verific~tionlvalidation; if there is empirical data, a validation data package is generated; if there is not empirical data, a verification package is generated; the test results are then reviewed; and finally, the user. requests accreditation, and a statement of accreditation is prepared. Once each component model is reviewed and approved, they are intertwined together into one integrated model. This integrated model is then tested itself, through a test script and autotest, so that it can be concluded that all models work conjointly, for a single purpose. The component I was assigned, specifically, was a fluid component, a discrete pressure switch. The switch takes a fluid pressure input, and if the pressure is greater than a designated cutoff pressure, the switch would stop fluid flow.

Ensey, Tyler S.↗
NASA NTRS2024

Investigating the Spatial and Temporal Limitations for Remote Sensing of Wildfire Smoke Using Satellite and Airborne Imagers During FIREX_AQ

Starting from point sources, wildfire smoke is important in the global aerosol system. The ability to characterize smoke near-source is key to modeling smoke dispersion and predicting air quality. With hemispheric views and 10-min refresh, imagers in Geostationary (GEO) orbit have advantages monitoring smoke over once-per-day sensors in low-earth orbit (LEO). However, both can be inadequate in capturing the characteristics of smoke plumes close to their sources due to too-coarse spatial resolution (both detector and product resolution), too-sparse temporal resolution (from LEO sensors), and too-conservative masking. In addition to satellite observations, the Fire Influence on Regional to Global Environments and Air Quality experiment offered sub-orbital enhanced-MODIS Airborne Simulator (eMAS) imagery at 50 m pixel resolution—including multiple eMAS flight tracks over individual fires in short time periods. It provided opportunity to explore smoke plume characterization at various spatial and temporal scales and quantify the limitations of space sensors for describing smoke magnitude near source as well as its temporal evolution. Here we applied modified aerosol algorithm to different imagers, relaxing its masking to estimate smoke's aerosol optical depth (AOD) as close as possible to its source. We found that GEO sensors with nominal 1 km spatial resolution can match the much finer resolution eMAS retrieved mean plume AOD, as long as the retrieval spatial resolution is finer than the width of the plumes. However, the plume's maximum AOD may be drastically underestimated by satellite products.

remote sensing↗
NASA NTRS1993

Mesoscale simulations of the November 25-26 and December 5-6 cirrus cases using the RAMS model

The Regional Atmospheric Modeling System (RAMS), developed at Colorado State University, was used during the First ISCCP (International Satellite Cloud Climatology Project) Regional Experiment (FIRE) 2 (13 Nov. through 6 Dec. 1991) to provide real time forecasts of cirrus clouds. Forecasts were run once a day, initializing with the 0000 UTC dataset provided by NOAA (Forecast Systems Laboratory (FSL) Mesoscale Analysis and Prediction System (MAPS)). In order to obtain better agreement with observations, a second set of simulations were done for the FIRE 2 cases that occurred on 25-26 Nov. and 5-6 Dec. In this set of simulations, a more complex radiation scheme was used, the Chen/Cotton radiation scheme, along with the nucleation of ice occurring at ice supersaturations as opposed to nucleation occurring at water supersaturations that was done in the actual forecast version. The runs using these more complex schemes took longer wall clock time (7-9 hours for the actual forecasts as compared to 12-14 hrs for the runs using the more complex schemes) however, the final results of the simulations were definitely improved upon. Comparisons between these two sets of simulations are given. Now underway are simulations of these cases using a closed analytical solution for the auto-conversion of ice from a pristine ice class (sizes less than about 50 microns in effective diameter) to a snow class (effective diameters on the order of several hundred microns). This solution is employed along with a new scheme for the nucleation of ice crystals due to Meyers et al and Demott et al. The scheme is derived assuming complete gamma distributions for both the pristine and snow classes. The time rate of change of the number concentration and mass mixing-ratio of each distribution is found by calculating either the flux of crystals that grow beyond a certain critical diameter by vapor deposition in an ice supersaturated regime or by calculating the flux of crystals that evaporate to sizes below that same critical effective diameter.

Harrington, J. L.↗
NASA NTRS1982

Development of a flash, bang, and smoke simulation of a shell burst

A large number of experiments (cue test firings) were performed in the definition of the cue concepts and packaging configurations. A total of 344 of these experiments were recorded with instrumentation photography to allow a quantitative analysis of the smoke cloud to be made as a function of time. These analyses were predominantly made using a short test site. Supplementary long range visibility tests were conducted to insure the required 3 kilometer visibility of the smoke signature.

Williamson, F. R.↗
NASA NTRS2016

Evaluation of Containment Boxes as a Fire Mitigation Method in Elevated Oxygen Conditions

NASA performed testing to evaluate the efficacy of fire containment boxes without forced ventilation. Configurational flammability testing was performed on a simulation avionics box replicating critical design features and filled with materials possessing representative flammability characteristics. This paper discusses the box's ability, under simulated end-use conditions, to inhibit the propagation of combustion to surrounding materials. Analysis was also performed to evaluate the potential for the fire containment box to serve as an overheat/ignition source to temperature sensitive equipment (such as items with lithium-ion batteries). Unrealistically severe combustion scenarios were used as a means to better understand the fire containment mechanism. These scenarios were achieved by utilizing materials/fuels not typically used in space vehicles due to flammability concerns. Oxygen depletion, during combustion within the fire containment boxes, drove self-extinguishment and proved an effective method of fire containment

Juarez, Alfredo↗
NASA NTRS1974

Method and apparatus for checking fire detectors

A fire detector checking method and device are disclosed for nondestructively verifying the operation of installed fire detectors of the type which operate on the principle of detecting the rate of temperature rise of the ambient air to sound an alarm and/or which sound an alarm when the temperature of the ambient air reaches a preset level. The fire alarm checker uses the principle of effecting a controlled simulated alarm condition to ascertain wheather or not the detector will respond. The checker comprises a hand-held instrument employing a controlled heat source, e.g., an electric lamp having a variable input, for heating at a controlled rate an enclosed mass of air in a first compartment, which air mass is then disposed about the fire detector to be checked. A second compartment of the device houses an electronic circuit to sense and adjust the temperature level and heating rate of the heat source.

Clawson, G. T.↗
NASA NTRS2026

Investigation of Bipropellant Plume-Induced Contamination Effects on Coverglass Materials

Contamination and degradation of external spacecraft materials by unburned and partially combusted species from bipropellant thruster plumes has long been observed as a key component of the induced space environment. Space shuttle flight experiments and returned flight hardware from the International Space Station (ISS) have both experienced microscopic impact features induced by high-velocity thruster plume droplets. Analytical results have shown that droplet impingement angle relative to a receiving surface plays a key role in the surface damage. Although impacts with normal impingement angles contribute more severely to surface degradation than highly oblique angles, surface effects at higher impingement angles should not be dismissed. Thruster plume-induced materials degradation is a complex phenomenon that depends on a variety of parameters, including but not limited to material type, system temperature and pressure, plume composition, and thruster firing specifications such as number of pulses, pulse duration, and sample distance from the thruster. For space applications, attaining the vacuum pressure and temperature conditions necessary for flight-like plume expansion and exposure conditions is not a trivial task. The German Aerospace Center (Deutsches Zentrum für Luft- und Raumfahrt, DLR) is a facility uniquely capable of simulating such conditions. Test coupons were exposed to bipropellant thruster firings under high vacuum at the DLR facility. Percent area coverage (PAC) and droplet size distributions were evaluated for the uncoated and coated solar array coverglass materials over a range of impingement angles (0̊ to 75̊). A post-test imaging workflow was developed that aimed to quantify changes in sample surface morphology obtained from scanning electron microscopy (SEM) images using the Image Processing and Analysis in Java (ImageJ) tool; an opensource image processing software. The goal was to create a framework through which to evaluate the effect of bipropellant-induced PAC and droplet size distribution on solar array coverglass optical transmission losses. Understanding this relationship is important because optical transmission losses are known to lead to current reduction in solar power generation systems. In addition to the development of surface characterization workflows, valuable lessons learned as they pertain to future investigations and experiments will be discussed. The authors hope that sharing these lessons will facilitate more utilization of DLR’s unique capabilities as well as open the conversation for how best to address experimental characterization of flight-like plume expansion and its impacts on materials surface degradation effects.

Gateway↗
NASA NTRS2010

Rocket Engine Altitude Simulation Technologies

John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

Woods, Jody L.↗
NASA NTRS1973

Aircraft Engine Sump Fire Mitigation

An investigation was performed of the conditions in which fires can result and be controlled within the bearing sump simulating that of a gas turbine engine; Esso 4040 Turbo Oil, Mobil Jet 2, and Monsanto MCS-2931 lubricants were used. Control variables include the oil inlet temperature, bearing temperature, oil inlet and scavenge rates, hot air inlet temperature and flow rate, and internal sump baffling. In addition to attempting spontaneous combustion, an electric spark and a rub (friction) mechanism were employed to ignite fires. Spontaneous combustion was not obtained; however, fires were readily ignited with the electric spark while using each of the three test lubricants. Fires were also ignited using the rub mechanism with the only test lubricant evaluated, Esso 4040. Major parameters controlling ignitions were: Sump configuration; Bearing and oil temperatures, hot air temperature and flow and bearing speed. Rubbing between stationary parts and rotating parts (eg. labyrinth seal and mating rub strip) is a very potent fire source suggesting that observed accidental fires in gas turbine sumps may well arise from this cause.

Rosenlieb, J. W.↗
NASA NTRS2001

Physical and Chemical Aspects of Fire Suppression in Extraterrestrial Environments

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.

Takahashi, F.↗
NASA NTRSDate not supplied

Update on AeroCom Biomass Burning Emission Injection Height experiment (BBEIH)

In BBEIH, we introduced the biomass burning injection heights based on MISR stereo-derived plume-height retrievals (Val Martin et al., 2010; 2018) in four participating CTMs. Specifically, we proposed 4 simulations for the year 2008: 1) BASE: using the biomass burning emissions from Global Fire Emissions Database version 4 with small fires (GFED4s) and the model default fire emission injection height; 2) BBIH: Same as BASE, but using the injection height derived from the seasonally and regionally varying MISR-retrieved plume heights; 3) NOBB: no fire emission; 4) Same as BASE, but using Fire Energetics and Emissions Research version 1.0 (FEER 1.0). We will address the following scientific questions: 1) To what extent are the model simulations sensitive to the assumed biomass burning injection height? 2) In which regions/seasons/surface-types are the aforementioned sensitivities most important? We will update the multi-model inter-comparison with a focus on the vertical aerosol distribution in near-source characteristics and downwind plume evolution." In order to test the sensitivity of model results to smoke injection height, we proposed the biomass burning emission injection height experiment (BBEIH) in the AeroCom 2019 workshop. The description of implementation methods and input data for BBEIH can be found in https://wiki.met.no/aerocom/phase3-experiments#biomass_burning_emission_injection_height_experiment_bbeih.

Xiaohua Pan↗
NASA NTRS1993

Space Shuttle Main Engine real time stability analysis

The Space Shuttle Main Engine (SSME) is a reusable, high performance, liquid rocket engine with variable thrust. The engine control system continuously monitors the engine parameters and issues propellant valve control signals in accordance with the thrust and mixture ratio commands. A real time engine simulation lab was installed at MSFC to verify flight software and to perform engine dynamic analysis. A real time engine model was developed on the AD100 computer system. This model provides sufficient fidelity on the dynamics of major engine components and yet simplified enough to be executed in real time. The hardware-in-the-loop type simulation and analysis becomes necessary as NASA is continuously improving the SSME technology, some with significant changes in the dynamics of the engine. The many issues of interfaces between new components and the engine can be better understood and be resolved prior to the firing of the engine. In this paper, the SSME real time simulation Lab at the MSFC, the SSME real time model, SSME engine and control system stability analysis, both in real time and non-real time is presented.

Kuo, F. Y.↗
NASA NTRS1982

Space Shuttle status and performance improvements

Following nearly a decade of studies and development, the Space Shuttle is beginning its flight test phase. This paper reviews the status of the program, and preparations of the space vehicle and facilities for flight. The launch preparations, including conduct of the Mission Verification Test, which comprises a Flight Readiness Firing prior to the actual launch, is discussed. This simulation was designed to test the readiness of the complete organization and all ground and flight equipment for commitment to the first manned orbital flight. The plan of the orbital flight program is outlined indicating how each succeeding flight is made more complex and severe to incrementally test the Space Shuttle. Completion of this four flight program readies the Space Shuttle for operations in late 1982. Payload performance to orbit is increased as the program progresses through the development flights and on into operations.

Day, L. E.↗
NASA NTRS1964

Dynamic Model Investigation of the Landing Characteristics of a Manned Spacecraft

Investigations were made to study the water-landing and certain grounds-surface landing characteristics of a Gemini spacecraft model. The water landing experiments were made by simulating paraglider and parachute letdowns with two 1/6- scale model configurations. Parameters included various combinations of attitude, horizontal speed, vertical speed, and landing skids extended and retracted. Investigations were made in calm water and in waves. The paraglider landings at horizontal speeds of 63 feet per second (19.8 m/sec) which resulted in a noseover or tumbling shortly after initial water contact. The maximum longitudinal acceleration of the model in calm water was about 14g units, and the maximum angular acceleration was 66 radians per second squared. In the parachute landings with the heat shield forward, the model skidded along the water surface on the heat shield. Parachute landings with the small end forward resulted in behavior similar to that of the paraglider landings. The ground-surface landings were made with a 1/3-scale model by simulating a parachute letdown with braking rockets, which were fired prior to touchdown to dissipate vertical velocity. In these landings, control of timing and aligning the rockets on the model was very critical, and violent behavior resulted when either rocket alignment or timing was in error. In the landings that were correctly controlled, the model either remained upright or slowly rolled over on its side.

EXPERIMENTS↗
NASA NTRS2016

A Hydrogen Containment Process for Nuclear Thermal Engine Ground testing

The objective of this study is to propose a new total hydrogen containment process to enable the testing required for NTP engine development. This H2 removal process comprises of two unit operations: an oxygen-rich burner and a shell-and-tube type of heat exchanger. This new process is demonstrated by simulation of the steady state operation of the engine firing at nominal conditions.

Wang, Ten-See↗
NASA NTRS2017

Coil-On-Plug Ignition for Oxygen/Methane Liquid Rocket Engines in Thermal-Vacuum Environments

A coil-on-plug ignition system has been developed and tested for Liquid Oxygen (LOX)/liquid methane (LCH4) rocket engines operating in thermal vacuum conditions. The igniters were developed and tested as part of the Integrated Cryogenic Propulsion Test Article (ICPTA), previously tested as part of the Project Morpheus test vehicle. The ICPTA uses an integrated, pressure-fed, cryogenic LOX/LCH4 propulsion system including a reaction control system (RCS) and a main engine. The ICPTA was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. A coil-on-plug ignition system has been developed to successfully demonstrate ignition reliability at these conditions while preventing corona discharge issues. The ICPTA uses spark plug ignition for both the main engine igniter and the RCS. The coil-on-plug configuration eliminates the conventional high-voltage spark plug cable by combining the coil and the spark plug into a single component. Prior to ICPTA testing at Plum Brook, component-level reaction control engine (RCE) and main engine igniter testing was conducted at NASA Johnson Space Center (JSC), which demonstrated successful hot-fire ignition using the coil-on-plug from sea-level ambient conditions down to 10(exp -2) torr. Integrated vehicle hot-fire testing at JSC demonstrated electrical and command/data system performance. Lastly, hot-fire testing at Plum Brook demonstrated successful ignitions at simulated altitude conditions at 30 torr and cold thermal-vacuum conditions at 6 torr. The test campaign successfully proved that coil-on-plug technology will enable integrated LOX/LCH4 propulsion systems in future spacecraft.

Melcher, John C.↗
NASA NTRS2006

Modeling the Optical Properties of Biomass Burning Aerosols: Young Smoke Aerosols From Savanna Fires and Comparisons to Observations from SAFARI 2000

Annually, farmers in southern Africa manage their land resources and prepare their fields for cultivation by burning crop residual debris, with a peak in the burning season occurring during August and September. The emissions from these fires in southern Africa are among the greatest from fires worldwide, and the gases and aerosol particles produced adversely affect air quality large distances from their source regions, and can even be tracked in satellite imagery as they cross the Atlantic and Pacific Ocean basins. During August and September 2000 an international group of researchers participating in the Southern African Regional Science Initiate field experiment (SAFARI 2000) made extensive ground-based, airborne, and satellite measurements of these gases and aerosols in order to quantify their amounts and effects on Earth's atmosphere. In this study we interpreted the measurements of smoke aerosol particles made during SAFARI 2000 in order to better represent these particles in a numerical model simulating their transport and fate. Typically, smoke aerosols emitted from fires are concentrated by mass in particles about 0.3 micrometers in diameter (1,000,000 micrometers = 1 meter, about 3 feet); for comparison, the thickness of a human hair is about 50 micrometers, almost 200 times as great. Because of the size of these particles, at the surface they can be easily inhaled into the lungs, and in high concentrations have deleterious health effects on humans. Additionally, these particles reflect and absorb sunlight, impacting both visibility and the balance of sunlight reaching -Earth's surface, and ultimately play a role in modulating Earth's climate. Because of these important effects, it is important that numerical models used to estimate Earth's climate response to changes in atmospheric composition accurately represent the quantity and evolution of smoke particles. In our model, called the Community Aerosol and Radiation Model for Atmospheres (CARMA) we used estimates of smoke emissions based on field studies and observations made with the NASA Terra and TRMM satellites. The meteorology used to calculate the transport was based on an assimilation of observed meteorological conditions provided by the National Center for Atmospheric Research.

Matichuk, R. I.↗
NASA NTRS1984

Viscous flow effects on hydrogen leaks from cracks in the Orbiter Challenger main engines

An analytical model was developed to provide additional insight and understanding of the factors that influence the simulation and prediction of leak rates from small cracks in pressurized containers. Specifically, the analysis was aimed at developing an analytical model capable of predicting the hydrogen leak rates from a crack in the combustion chamber coolant discharge manifold on main engine 1 of the Orbiter Challenger that was discovered during flight readiness firings 1 and 2. This model was based on viscous pipe flow analyses and calibrated for the crack geometry by using helium leak-rate data obtained from both low- and high-pressure tests used to simulate the flight readiness firing test conditions. In addition, this model includes the effects of crack width changes caused by different working stresses associated with the different test conditions. Because of the combination of the small crack dimensions and the wide range of pressures used for the test conditions, either laminar or turbulent viscous effects dominated the flows at all test conditions. This model was used to illustrate the sensitivity of the predicted leak rates to considerations of test conditions, viscous flow effects, and geometric features of the crack. In addition, the model was certified by comparing the hydrogen leak-rate prediction for the flight readiness firing test condition to the actual measured leak rate. The prediction was within 9 percent of the measured value.

Goodrich, W. D.↗