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Results for “limited hydrogenation ability”

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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NASA NTRS1984

An upper limit for stratospheric hydrogen peroxide

It has been postulated that hydrogen peroxide is important in stratospheric chemistry as a reservoir and sink for odd hydrogen species, and for its ability to interconvert them. The present investigation is concerned with an altitude dependent upper limit curve for stratospheric hydrogen peroxide, taking into account an altitude range from 21.5 to 38.0 km for January 23, 1983. The data employed are from balloon flight No. 1316-P, launched from the National Scientific Balloon Facility (NSBF) in Palestine, Texas. The obtained upper limit curve lies substantially below the data reported by Waters et al. (1981), even though the results are from the same latitude and are both wintertime measurements.

Chance, K. V.↗
NASA NTRS2006

Engineering Analysis Studies for Preliminary Design of Lightweight Cryogenic Hydrogen Tanks in UAV Applications

A series of engineering analysis studies were conducted to investigate the potential application of nanoclay-enhanced graphite/epoxy composites and polymer cross-linked silica aerogels in cryogenic hydrogen storage tank designs. This assessment focused on the application of these materials in spherical tank designs for unmanned aeronautic vehicles with mission durations of 14 days. Two cryogenic hydrogen tank design concepts were considered: a vacuum-jacketed design and a sandwiched construction with an aerogel insulating core. Analyses included thermal and structural analyses of the tank designs as well as an analysis of hydrogen diffusion to specify the material permeability requirements. The analyses also provided material property targets for the continued development of cross-linked aerogels and nanoclay-enhanced graphite/epoxy composites for cryogenic storage tank applications. The results reveal that a sandwiched construction with an aerogel core is not a viable design solution for a 14-day mission. A vacuum-jacketed design approach was shown to be far superior to an aerogel. Aerogel insulation may be feasible for shorter duration missions. The results also reveal that the application of nanoclay-enhanced graphite/epoxy should be limited to the construction of outer tanks in a vacuum-jacketed design, since a graphite/epoxy inner tank does not provide a significant weight savings over aluminum and since the ability of nanoclay-enhanced graphite/epoxy to limit hydrogen permeation is still in question.

Sullivan, Roy M.↗
NASA NTRS2006

Design Considerations for the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today s best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. The goal of these simulations would be directed toward expanding the performance envelope of NTR engines over that which was demonstrated during the Rover and NERVA nuclear rocket programs of the 1970's. Currently, such a simulator is nearing completion at the Marshall Space Flight Center, and will shortly be used in the future to evaluate a wide variety of he1 element designs and the materials of which they are constructed. This present work addresses the initial experimental objectives of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Emrich, Bill↗
NASA NTRS2007

NTREES Testing and Operations Status

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today's best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. Currently, the construction of such a simulator has been completed at the Marshall Space Flight Center, and will be used in the future to evaluate a wide variety of fuel element designs and the materials of which they are fabricated. This present work addresses the operational status of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Emrich, Bill↗
NASA NTRS2010

Second Generation International Space Station (ISS) Total Organic Carbon Analyzer (TOCA) Verification Testing and On-Orbit Performance Results

The Total Organic Carbon Analyzer (TOGA) is designed to autonomously determine recovered water quality as a function of TOC. The current TOGA has been on the International Space Station since November 2008. Functional checkout and operations revealed complex operating considerations. Specifically, failure of the hydrogen catalyst resulted in the development of an innovative oxidation analysis method. This method reduces the activation time and limits the hydrogen produced during analysis, while retaining the ability to indicate TOC concentrations within 25% accuracy. Subsequent testing and comparison to archived samples returned from the Station and tested on the ground yield high confidence in this method, and in the quality of the recovered water.

Bentley, Nicole L.↗
NASA NTRS2014

Grain Destruction in a Supernova Remnant Shock Wave

Dust grains are sputtered away in the hot gas behind shock fronts in supernova remnants, gradually enriching the gas phase with refractory elements. We have measured emission in C IV (lambda)1550 from C atoms sputtered from dust in the gas behind a non-radiative shock wave in the northern Cygnus Loop. Overall, the intensity observed behind the shock agrees approximately with predictions from model calculations that match the Spitzer 24 micron and the X-ray intensity profiles. Thus these observations confirm the overall picture of dust destruction in SNR shocks and the sputtering rates used in models. However, there is a discrepancy in that the CIV intensity 10'' behind the shock is too high compared to the intensities at the shock and 25'' behind it. Variations in the density, hydrogen neutral fraction and the dust properties over parsec scales in the pre- shock medium limit our ability to test dust destruction models in detail.

dust↗
NASA NTRS2010

Robotic Lunar Landers for Science and Exploration

NASA Marshall Space Flight Center (MSFC) and The Johns Hopkins University Applied Physics Laboratory (APL) have been conducting mission studies and performing risk reduction activities for NASA s robotic lunar lander flight projects. This paper describes some of the lunar lander concepts derived from these studies conducted by the MSFC/APL Robotic Lunar Lander Development Project team. In addition, the results to date of the lunar lander development risk reduction efforts including high pressure propulsion system testing, structure and mechanism development and testing, long cycle time battery testing and combined GN&C and avionics testing will be addressed. The most visible elements of the risk reduction program are two autonomous lander flight test vehicles: a compressed air system with limited flight durations and a second version using hydrogen peroxide propellant to achieve significantly longer flight times and the ability to more fully exercise flight sensors and algorithms.

Chavers, D. G.↗
NASA NTRS2010

Robotic Lunar Landers for Science and Exploration

NASA Marshall Space Flight Center and The Johns Hopkins University Applied Physics Laboratory has been conducting mission studies and performing risk reduction activities for NASA s robotic lunar lander flight projects. In 2005, the Robotic Lunar Exploration Program Mission #2 (RLEP-2) was selected as a Exploration Systems Mission Directorate precursor robotic lunar lander mission to demonstrate precision landing and definitively determine if there was water ice at the lunar poles; however, this project was canceled. Since 2008, the team has been supporting NASA s Science Mission Directorate designing small lunar robotic landers for diverse science missions. The primary emphasis has been to establish anchor nodes of the International Lunar Network (ILN), a network of lunar science stations envisioned to be emplaced by multiple nations. This network would consist of multiple landers carrying instruments to address the geophysical characteristics and evolution of the moon. Additional mission studies have been conducted to support other objectives of the lunar science community and extensive risk reduction design and testing has been performed to advance the design of the lander system and reduce development risk for flight projects. This paper describes the current status of the robotic lunar mission studies that have been conducted by the MSFC/APL Robotic Lunar Lander Development team, including the ILN Anchor Nodes mission. In addition, the results to date of the lunar lander development risk reduction efforts including high pressure propulsion system testing, structure and mechanism development and testing, long cycle time battery testing and combined GN&C and avionics testing will be addressed. The most visible elements of the risk reduction program are two autonomous lander test articles: a compressed air system with limited flight durations and a second version using hydrogen peroxide propellant to achieve significantly longer flight times and the ability to more fully exercise flight sensors and algorithms. Robotic Lunar Lander design and development will have significant feed-forward to other missions to the Moon and, indeed, to other airless bodies such as Mercury, asteroids, and Europa, to which similar science and exploration objectives are applicable.

Cohen, B. A.↗
NASA NTRS2002

Ion Storage with the High Performance Antiproton Trap (HiPAT)

The matter antimatter reaction represents the densest form of energy storage/release known to modern physics: as such it offers one of the most compact sources of power for future deep space exploration. To take the first steps along this path, NASA-Marshall Space Flight Center is developing a storage system referred to as the High Performance Antiproton Trap (HiPAT) with a goal of maintaining 10(exp 12) particles for up to 18 days. Experiments have been performed with this hardware using normal matter (positive hydrogen ions) to assess the device's ability to hold charged particles. These ions are currently created using an electron gun method to ionize background gas; however, this technique is limited by the quantity that can be captured. To circumvent this issue, an ion source is currently being commissioned which will greatly increase the number of ions captured and more closely simulate actual operations expected at an antiproton production facility. Ions have been produced, stored for various time intervals, and then extracted against detectors to measure species, quantity and energy. Radio frequency stabilization has been tested as a method to prolong ion lifetime: results show an increase in the baseline 1/e lifetime of trapped particles from hours to days. Impurities in the residual background gas (typically carbon-containing species CH4, CO, CO2, etc.) present a continuing problem by reducing the trapped hydrogen population through the mechanism of ion charge exchange.

Martin, James↗
NASA NTRS1979

Laser rocket system analysis

The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

Jones, W. S.↗
NASA NTRS2012

Robotic Lunar Lander Development Status

NASA Marshall Space Flight Center and John Hopkins University Applied Physics Laboratory have developed several mission concepts to place scientific and exploration payloads ranging from 10 kg to more than 200 kg on the surface of the moon. The mission concepts all use a small versatile lander that is capable of precision landing. The results to date of the lunar lander development risk reduction activities including high pressure propulsion system testing, structure and mechanism development and testing, and long cycle time battery testing will be addressed. The most visible elements of the risk reduction program are two fully autonomous lander flight test vehicles. The first utilized a high pressure cold gas system (Cold Gas Test Article) with limited flight durations while the subsequent test vehicle, known as the Warm Gas Test Article, utilizes hydrogen peroxide propellant resulting in significantly longer flight times and the ability to more fully exercise flight sensors and algorithms. The development of the Warm Gas Test Article is a system demonstration and was designed with similarity to an actual lunar lander including energy absorbing landing legs, pulsing thrusters, and flight-like software implementation. A set of outdoor flight tests to demonstrate the initial objectives of the WGTA program was completed in Nov. 2011, and will be discussed.

Ballard, Benjamin↗
NASA NTRS2012

NASA's Robotic Lunar Lander Development Program

NASA Marshall Space Flight Center and the Johns Hopkins University Applied Physics Laboratory have developed several mission concepts to place scientific and exploration payloads ranging from 10 kg to more than 200 kg on the surface of the moon. The mission concepts all use a small versatile lander that is capable of precision landing. The results to date of the lunar lander development risk reduction activities including high pressure propulsion system testing, structure and mechanism development and testing, and long cycle time battery testing will be addressed. The most visible elements of the risk reduction program are two fully autonomous lander flight test vehicles. The first utilized a high pressure cold gas system (Cold Gas Test Article) with limited flight durations while the subsequent test vehicle, known as the Warm Gas Test Article, utilizes hydrogen peroxide propellant resulting in significantly longer flight times and the ability to more fully exercise flight sensors and algorithms. The development of the Warm Gas Test Article is a system demonstration and was designed with similarity to an actual lunar lander including energy absorbing landing legs, pulsing thrusters, and flight-like software implementation. A set of outdoor flight tests to demonstrate the initial objectives of the WGTA program was completed in Nov. 2011, and will be discussed.

Ballard, Benjamin W.↗
NASA NTRS1994

Dynamic Simulation of a Periodic 10 K Sorption Cryocooler

A transient thermal simulation model has been developed to simulate the dynamic performance of a multiple-stage 10 K sorption cryocooler for spacecraft sensor cooling applications that require periodic quick-cooldown (under 2 minutes) , negligible vibration, low power consumption, and long life (5 to 10 years). The model was specifically designed to represent the Brilliant Eyes Ten-Kelvin Sorption Cryocooler Experiment (BETSCE), but it can be adapted to represent other sorption cryocooler systems as well. The model simulates the heat transfer, mass transfer, and thermodynamic processes in the cryostat and the sorbent beds for the entire refrigeration cycle, and includes the transient effects of variable hydrogen supply pressures due to expansion and overflow of hydrogen during the cooldown operation. The paper describes model limitations and simplifying assumptions, with estimates of errors induced by them, and presents comparisons of performance predictions with ground experiments. An important benefit of the model is its ability to predict performance sensitivities to variations of key design and operational parameters. The insights thus obtained are expected to lead to higher efficiencies and lower weights for future designs.

cryocooler design performance sorption simulation ↗
NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high‐temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high‐melting‐point ceramics‐metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non‐nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small‐scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Gomez, C. F.↗
NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high‐temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high‐melting‐point ceramics‐metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non‐nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small‐scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Gomez, Carlos F.↗
NASA NTRS2006

The Role of Flight Experiments in the Development of Cryogenic Fluid Management Technologies

This paper reviews the history of cryogenic fluid management technology development and infusion into both the Saturn and Centaur vehicles. Ground testing and analysis proved inadequate to demonstrate full scale performance. As a consequence flight demonstration with a full scale vehicle was required by both the Saturn and Centaur programs to build confidence that problems were addressed. However; the flight vehicles were highly limited on flight instrumentation and the flight demonstration locked-in the design without challenging the function of design elements. Projects reviewed include: the Aerobee Sounding Rocket Cryogenic Fluid Management (CFM) tests which served as a valuable stepping stone to flight demonstration and built confidence in the ability to handle hydrogen in low gravity; the Saturn IVB Fluid Management Qualification flight test; the Atlas Centaur demonstration flights to develop two burn capability; and finally the Titan Centaur two post mission flight tests.

Chato, David J.↗
NASA NTRS1995

Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies: A Preliminary Assessment

The feasibility of conducting human missions to the Moon is examined assuming the use of three 'high leverage' technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) 'in-situ' resource utilization (ISRU)--specifically 'lunar-derived' liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the 'compact' dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of approximately 60 t (3 SSTO launches). Using approximately 8 t of LUNOX to 'reoxidize' the LERV for a 'direct return' flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine's ability to operate at any oxygen/ hydrogen mixture ratio from 0 to 7 with high specific impulse (approximately 940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV's 'propulsion' and 'propellant modules'. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes. Concluding remarks address the issue of lunar transportation system costs from the launch vehicle perspective.

Borowski, Stanley K.↗
NASA NTRS1972

Application of transport techniques to the analysis of NERVA shadow shields

A radiation shield internal to the NERVA nuclear rocket reactor required to limit the neutron and photon radiation levels at critical components located external to the reactor was evaluated. Two significantly different shield mockups were analyzed: BATH, a composite mixture of boron carbide, aluminum and titanium hydride, and a borated steel-liquid hydrogen system. Based on the comparisons between experimental and calculated neutron and photon radiation levels, the following conclusions were noted: (1) The ability of two-dimensional discrete ordinates code to predict the radiation levels internal to and at the surface of the shield mockups was clearly demonstrated. (2) Internal to the BATH shield mockups, the one-dimensional technique predicted the axial variation of neutron fluxes and photon dose rates; however, the magnitude of the neutron fluxes was about a factor of 1.8 lower than the two-dimensional analysis and the photon dose rate was a factor of 1.3 lower.

Capo, M. A.↗