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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Mars Ascent Vehicle Hybrid Propulsion Effort

A technology development program was undertaken to determine if the benefits of hybrid rocket propulsion could be realized for a Mars Ascent Vehicle (MAV) application. Specifically, the goal of this program was to increase the Technology Readiness Level (TRL) of a hybrid propulsion system such that it could be considered for a potential Mars Sample Return Campaign. Over the course of approximately five years, a new, wax-based fuel was developed, characterized and tested with Mixed Oxides of Nitrogen (from MON-3 to MON-25). Various ignition mechanisms were tested and solid hypergolic additives were evaluated for their potential to ignite the hybrid motor. Several hot-fire test campaigns were completed at both sub- and full-scale. The technology development program was born out of a JPL study documented by Shotwell and Karp that identified the Single Stage to Orbit (SSTO) hybrid propulsion system for a potential Mars Ascent Vehicle to be the lowest Gross Lift Off Mass option from the ten configurations evaluated. Benefits of the hybrid option included its predicted favorable low temperature storage and operation behavior, high performance and ability to restart (enabling the SSTO). The main disadvantage of the hybrid option was its low TRL, as it essentially started as a blank sheet design with a new fuel development. Near the end of the technology development program, a Preliminary Architecture Assessment (PAA) was completed to evaluate a single-stage hybrid propulsion system and a two-stage solid propulsion system for a potential MAV. The solid propulsion system was selected for further study in 2019, primarily because of its flight heritage. While the hybrid technology development program made many strides, there was still a list of challenges to overcome. This paper will discuss the substantial progress and remaining challenges from the technology development effort from 2015 through 2019. This includes fuel development and reformulation, sub- and full-scale testing at Space Propulsion Group, full scale testing at Parabilis, full scale testing at Whittinghill Aerospace, hypergolic additive testing at Purdue and Penn State and the evaluation of adding hypergolic additives to a fullscale grain. It also includes a discussion of the design that resulted from the Preliminary Architecture Assessment.

hybrid rocket↗
NASA NTRS1974

High-efficiency, low-temperature cesium diodes with lanthanum-hexaboride electrodes

Lanthanum-hexaboride electrodes in 1700 K cesium diodes may triple power outputs compared with those demonstrated for nuclear thermionic space applications. Still greater relative gains seem possible for emitters below 1700 K. Further improvements in cesium-diode performance should result from the lower collector temperatures allowed for earth and low-power-space duties. Decreased temperatures will lessen thermal-transport losses that attend thermionic-conversion mechanisms. Such advantages will add to those from collector-Carnot and electrode effects. If plasma ignition difficulties impede diode temperature reductions, recycling small fractions of the output power could provide ionization. So high-efficiency, low-temperature cesium diodes with lanthanum-hexaboride electrodes appear feasible.

Morris, J. F.↗
NASA NTRS1972

Characterization of the space shuttle reaction control system engine

A computer program was developed and written in FORTRAN 5 which predicts the transient and steady state performance and heat transfer characteristics of a pulsing GO2/GH2 rocket engine. This program predicts the dynamic flow and ignition characteristics which, when combined in a quasi-steady state manner with the combustion and mixing analysis program, will provide the thrust and specific impulse of the engine as a function of time. The program also predicts the transient and steady state heat transfer characteristics of the engine using various cooling concepts. The computer program, test case, and documentation are presented. The program is applicable to any system capable of utilizing the FORTRAN 4 or FORTRAN 5 language.

Wilson, M. S.↗
NASA NTRS1984

High-temperature gaseous oxygen/hydrogen thrusters for space station

Efforts to determine the critical technologies necessary for the development of a high-temperature GO2/GH2 thruster are reviewed, for space station applications. Two types of thrust chambers are evaluated: one operable at high temperature and the other incorporating regenerative cooling. The high temperature chamber made of rhenium requires minimum cooling of the chamber wall, however, an oxidation barrier should be incorporated to prevent the rhenium thruster from readily oxidizing. The use of rhenium brought about a lower cost, lower weight, and simplicity of fabrication. The igniter-injector, the chamber, the test facility, and the test program are discussed. It is found that an obtained flow split of 58/42 increased the thruster performance from 3893 N-s/Kg to 4030 N-s/Kg for the same chamber pressure and overall mixture ratio. An increase in performance is also observed when the core mixture ratio is lowered. The thruster was tested for 1.9 hrs with no well degradation.

Dowdy, M. W.↗
NASA NTRS2000

Model of ASTM Flammability Test in Microgravity: Iron Rods

There is extensive qualitative results from burning metallic materials in a NASA/ASTM flammability test system in normal gravity. However, this data was shown to be inconclusive for applications involving oxygen-enriched atmospheres under microgravity conditions by conducting tests using the 2.2-second Lewis Research Center (LeRC) Drop Tower. Data from neither type of test has been reduced to fundamental kinetic and dynamic systems parameters. This paper reports the initial model analysis for burning iron rods under microgravity conditions using data obtained at the LERC tower and modeling the burning system after ignition. Under the conditions of the test the burning mass regresses up the rod to be detached upon deceleration at the end of the drop. The model describes the burning system as a semi-batch, well-mixed reactor with product accumulation only. This model is consistent with the 2.0-second duration of the test. Transient temperature and pressure measurements are made on the chamber volume. The rod solid-liquid interface melting rate is obtained from film records. The model consists of a set of 17 non-linear, first-order differential equations which are solved using MATLAB. This analysis confirms that a first-order rate, in oxygen concentration, is consistent for the iron-oxygen kinetic reaction. An apparent activation energy of 246.8 kJ/mol is consistent for this model.

Steinberg, Theodore A↗
NASA NTRS1971

The combustion of pure and composite propellants - The expansion and application of laminar flame theory to heterogeneous solid propellants.

Development of a combustion theory based on the laminarized solutions to the energy and flow conservation equations, which is more realistic in recognizing the nature of the heating-rate problem and in obtaining a practical solution to estimating its magnitude. A new experimental approach is used for studying the combustion behavior of pure monopropellants and composite propellants which uses a laser beam to supply additional heat feedback to a burning surface. New experimental data are presented for the laser-induced combustion rate and ignition delay of pure ammonium perchlorate. The pure monopropellant theory is generalized to include such nonadiabatic effects, and the new experimental data are in good agreement with the nonadiabatic theory.-

Hertzberg, M.↗
NASA NTRS1977

Current status of droplet and liquid combustion

The present understanding of spray combustion in rocket engine, gas turbine, Diesel engine and industrial furnace applications is reviewed. In some cases, spray combustion can be modeled by ignoring the details of spray evaporation and treating the system as a gaseous diffusion flame; however, in many circumstances, this simplification is not adequate and turbulent two-phase flow must be considered. The behavior of individual droplets is a necessary component of two-phase models and recent work on transient droplet evaporation, ignition and combustion is considered, along with a discussion of important simplifying assumptions involved with modeling these processes. Methods of modeling spray evaporation and combustion processes are also discussed including: one-dimensional models for rocket engine and prevaporized combustion systems, lumped zone models (utilizing well-stirred reactor and plug flow regions) for gas turbine and furnace systems, locally homogeneous turbulent models, and two-phase models. The review highlights the need for improved injector characterization methods, more information of droplet transport characteristics in turbulent flow and continued development of more complete two-phase turbulent models

Faeth, G. M.↗
DOE OSTI2025

Key advancements toward eliminating the “drive deficit” in ICF hohlraum simulations

This paper, extended from a Letter [Chen et al., Phys. Rev. E 110, L013201 (2024)] published in Phys. Rev. E in 2024, presents detailed results from a set of experiments studying the x-ray drive produced in laser-heated ignition-scale hohlraums at the National Ignition Facility. The results reveal that deficiencies in non-local thermodynamic equilibrium atomics physics calculations of gold emissivity lead to the overestimation of the “M-band” radiation and are responsible for discrepancies in drive predictions vs measurements in these hohlraums. These experiments were performed on ViewFactor targets with the same gas fill and laser pulse shapes as currently used in ongoing ignition experiments. The common hohlraum model with a magnetohydrodynamics heat transport model and a 10%–20% reduction to the M-band absorption and emission opacity can bring the simulations into agreement with several independent measurements, including the absolute radiation flux, the radiation flux in the 2–4 keV spectral range, and spectroscopic measurements of both K-shell dopant and L-shell gold. This new model is applicable to ICF and high-energy density experiments that use gold hohlraums as drivers.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
DOE OSTI2026

Ultra Wide Bandgap Laser Addressable Photoconductors

Wide Band Gap (WBG) and Ultra Wide Band Gap (UWBG) photoconductors have the potential to meet several critical mission needs for LLNL and the broader USG due to their exceptional characteristics related to large critical electric field. In particular, high energy laser systems such as the National Ignition Facility would benefit from the development of new materials for active optical devices such as the Optically Addressable Light Valve (OALV) used in the Programmable Spatial Shaper System. Missions in Global Security would benefit from higher performance Photoconductive Semiconductor Switches for pulsed power and radiofrequency amplifiers. This project examined the suitability of both commercially available and custom synthesized semi insulating WBG and UWBG materials for several of these applications. We developed high laser damage threshold OALVs for optical systems based on several of these materials as well as new RF, pulsed power, and power electronics devices based on the same. These devices show superior power handling in terms of laser damage threshold, electric field, and power output. A number of the technologies developed have been transitioned to both internal and external customers.

42 ENGINEERING↗
DOE OSTI2026

Prioritizing Uncertainties in Hydrogen Contribution to Risk in Post-Crash Outcomes for Rail

This report presents analysis from Sandia National Laboratories predicting contributions to risk associated with the use of hydrogen technology for rail. Event sequence diagrams are used to describe possible accident scenarios and progressions. Initiating event frequencies and branch event probabilities for each scenario are quantified with uncertainty using distributions fit to Federal Railroad Administration and U.S. Department of Transportation Pipeline and Hazardous Materials Safety Administration data on applicable accidents from 2000 to 2020. Uncertainty is propagated through the event sequence diagram to estimate the frequency and conditional probability of accident end states. The analysis identifies four scenarios with significant contributions to risk from hydrogen that are predicted to occur relatively frequently, which may inform priorities for reducing uncertainty. These scenarios are 1) overpressure events resulting from collisions with hydrogen release due to mechanical damage and delayed ignition, 2) jet fire events resulting from collisions with hydrogen release due to mechanical damage and immediate ignition, 3) jet fires resulting from fire or explosion initiating events involving the hydrogen tank and correct operation of the thermally-activated pressure relief device (TPRD) subsequent to the thermal insult, and 4) pressure burst resulting from fire or explosion initiating events involving the hydrogen tank and failure of the TPRD. Delayed and immediate hydrogen ignition probabilities are identified as being highly uncertain and potential candidates for reducing conservatism in the predicted frequencies for these two scenarios.

08 HYDROGEN↗
NASA NTRS1990

Definition of experiments to investigate fire suppressants in microgravity

Defined and justified here are the conceptual design and operation of a critical set of experiments expected to yield information on suppressants and on suppressant delivery systems under realistic spacecraft-fire conditions (smoldering). Specific experiment parameters are provided on the solid fuel (carbon), oxidants (habitable spacecraft atmospheres), fuel/oxidant supply, mixing mode, and rate (quiescent and finite; ventilated and replenishable), ignition mode, event, and reignition tendency, fire-zone size, fire conditions, lifetime, and consequences (toxicity), suppressants (CO2, H2O, N2) and suppressant delivery systems, and diagnostics. Candidate suppressants were identified after an analysis of how reduced gravity alters combustion, and how these alterations may influence the modes, mechanisms, and capacities of terrestrial agents to suppress unwanted combustion, or fire. Preferred spacecraft suppression concepts included the local, near-quiescent application of a gas, vapor, or mist that has thermophysical fire-suppression activity and is chemically inert under terrestrial (normal gravity) combustion conditions. The scale, number, and duration (about 1 hour) of the proposed low-gravity experiments were estimated using data not only on the limitations imposed by spacecraft-carrier (Shuttle or Space Station Freedom) accommodations, but also data on the details and experience of standardized smolder-suppression experiments at normal gravity. Deliberately incorporated into the conceptual design was sufficient interchangeability for the prototype experimental package to fly either on Shuttle now or Freedom later. This flexibility is provided by the design concept of up to 25 modular fuel canisters within a containment vessel, which permits both integration into existing low-gravity in-space combustion experiments and simultaneous testing of separate experiments to conserve utilities and time.

Reuther, James J.↗
DOE OSTI2025

Acoustic Rocket Signatures Collected by Smartphones

Rockets generate complex acoustic signatures that can be detected over a thousand kilometers from their source. While many far-field acoustic rocket signatures have been collected and released to the public, very few signatures collected at distances less than 100 km are available. This work presents a curated and annotated dataset of acoustic signatures of 243 rocket launches collected by a network of smartphones stationed at distances between 10 and 70 km from the launch sites, resulting in 1089 individual recordings. Due to the frequency dependence of atmospheric attenuation and the relatively short propagation distances, higher-frequency features not preserved in most publicly available data are observed. The signals are time-aligned to allow for different segments of the signal (ignition, launch, trajectory, chronology) to be more easily examined and compared. Initial analysis of the features of these rocket launch stages is performed, observed features are compared to those found in the existing literature, and comparisons between signals from launches of different rocket types are made. The dataset is annotated and made available to the public to aid future analysis of the characteristics and source mechanisms of rocket acoustics as well as applications such as rocket detection and classification models.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1977

Current status of droplet and liquid combustion

The present understanding of spray combustion in rocket engine, gas turbine, Diesel engine and industrial furnace applications is reviewed. In some cases, spray combustion can be modeled by ignoring the details of spray evaporation and treating the system in the same manner as a gaseous diffusion flame; however, in many circumstances, this type of simplification is not adequate and the turbulent two-phase flow must be considered. The behavior of individual droplets is a necessary component of two-phase models and recent work on transient droplet evaporation, ignition and combustion is considered, along with a discussion of important simplifying assumptions involved with modeling these processes. Methods of modeling spray evaporation and combustion processes are also discussed including: one-dimensional models for rocket engine and prevaporized combustion systems, lumped zone models (utilizing well-stirred reactor and plug flow regions) for gas turbine and furnace systems, locally homogeneous turbulent models, and two-phase models. The review highlights the need for improved injector characterization methods, more information of droplet transport characteristics in turbulent flow and continued development of more complete two-phase turbulent models.

Faeth, G. M.↗
NASA NTRS2008

Liquid Oxygen/Liquid Methane Ascent Main Engine Technology Development

The National Aeronautics & Space Administration (NASA) has identified Liquid Oxygen (LO2)/Liquid Methane (LCH4) as a potential propellant combination for future space vehicles based upon the Exploration Systems Architecture Study (ESAS). The technology is estimated to have higher performance and lower overall systems mass compared to existing hypergolic propulsion systems. The current application considering this technology is the lunar ascent main engine (AME). AME is anticipated to be an expendable, pressure-fed engine to provide ascent from the moon at the completion of a 210 day lunar stay. The engine is expected to produce 5,500 lbf (24,465 N) thrust with variable inlet temperatures due to the cryogenic nature of the fuel and oxidizer. The primary technology risks include establishing reliable and robust ignition in vacuum conditions, maximizing specific impulse, developing rapid start capability for the descent abort, providing the capability for two starts and producing a total engine bum time over 500 seconds. This paper will highlight the efforts of the Marshall Space Flight Center (MSFC) in addressing risk reduction activities for this technology.

Robinson, Joel W.↗
NASA NTRS2009

Space Systems - Safety and Compatibility of Materials - Method to Determine the Flammability Thresholds of Materials

Spacecraft fire safety emphasizes fire prevention, which is achieved primarily through the use of fire-resistant materials. Materials selection for spacecraft is based on conventional flammability acceptance tests, along with prescribed quantity limitations and configuration control for items that are non-pass or questionable. ISO 14624-1 and -2 are the major methods used to evaluate flammability of polymeric materials intended for use in the habitable environments of spacecraft. The methods are upward flame-propagation tests initiated in static environments and using a well-defined igniter flame at the bottom of the sample. The tests are conducted in the most severe flaming combustion environment expected in the spacecraft. The pass/fail test logic of ISO 14624-1 and -2 does not allow a quantitative comparison with reduced gravity or microgravity test results; therefore their use is limited, and possibilities for in-depth theoretical analyses and realistic estimates of spacecraft fire extinguishment requirements are practically eliminated. To better understand the applicability of laboratory test data to actual spacecraft environments, a modified ISO 14624 protocol has been proposed that, as an alternative to qualifying materials as pass/fail in the worst-expected environments, measures the actual upward flammability limit for the material. A working group established by NASA to provide recommendations for exploration spacecraft internal atmospheres realized the importance of correlating laboratory data with real-life environments and recommended NASA to develop a flammability threshold test method. The working group indicated that for the Constellation Program, the flammability threshold information will allow NASA to identify materials with increased flammability risk from oxygen concentration and total pressure changes, minimize potential impacts, and allow for development of sound requirements for new spacecraft and extravehicular landers and habitats. Furthermore, recent research has shown that current normal gravity materials flammability tests do not correlate with flammability in ventilated, micro- or reduced-gravity conditions. Currently, the materials selection for spacecraft is based on the assumption of commonality between ground flammability test results and spacecraft environments, which does not appear to be valid. Materials flammability threshold data acquired in normal gravity can be correlated with data obtained in microgravity or reduced-gravity experiments, and consequently a more accurate assessment of the margin of safety of the material in the real environment can be made. In addition, the method allows the option of selecting better or best space system materials, as opposed to what would be considered just acceptable from a flammability point of view and realistic assessment of spacecraft fire extinguishment needs, which could result in significant weight savings. The knowledge afforded by this technique allows for limited extrapolations of flammability behavior to conditions not specifically tested and that could potentially result in significant cost and time savings. The intent of this Technical Specification is to bring to the attention of International Aerospace Community the importance of correlating laboratory test data with real-life space systems applications. The method presented is just one of the possibilities that are believed will lead to better understanding the applicability of laboratory aerospace materials flammability test data. International feedback on improving the proposed method, as well as suggestions for correlating other laboratory aerospace test data with real-life applications relevant to space systems are being sought.

Hirsch, David↗
DOE OSTI2024

Milestone report for MRT 8504: Commission EXAFS spectrometer with spectral range of 17.95-18.6 keV design to obtain EXAFS data for Pu L shell experiments

This milestone describes the development and fielding of EXAFS Phase II spectrometer designed to provide high quality (SNR>=5) EXAFS spectra for Plutonium measurements at the National Ignition facility (NIF). Benefits: the EXAFS data will be used by the High Energy Density (HED) Physics and more specifically the Focused Material Science program to measure the thermal state of compressed materials with good temperature sensitivity (~100-1000 Kelvin). Data with good signal to noise ratio (SNR>=5) and good spectral resolution allows the FMS program to fill a gap in diagnosing temperature to completely constrain material properties for weapon physics and Stockpile Stewardship applications.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS1982

Pulsed plasma thruster ignitor plug ignition characteristics

The operating characteristics of the semiconductor spark gap ignitor plug used to initiate the discharge in a pulsed plasma thruster are investigated. Current and voltage waveform measurements for the ignitor plug alone, and for the plug mounted in the thruster cathode indicate an average dynamic impedance of 0.2 ohms for a wide range of operating conditions, and a trigger circuit energy transfer efficiency to the plug on the order of 25%. Two modes of trigger ignition are found which are related to the rise time of the applied voltage pulse. Analysis of plasma characteristics reveals plug erosion to occur primarily in the semiconductor and anode regions, by mechanisms including plug metal embrittlement, plasma sputtering and vaporization caused by the discharge current pulse. Measurements also indicate a plume velocity on the order of 1580,000 cm/sec which increases with trigger circuit stored energy. The results establish levels of performance for future applications of semiconductor spark gap ignitor plugs in plasma devices.

Brady, M. E.↗
NASA NTRS2014

A Monte Carlo Approach to Modeling the Breakup of the Space Launch System EM-1 Core Stage with an Integrated Blast and Fragment Catalogue

The Liquid Propellant Fragment Overpressure Acceleration Model (L-FOAM) is a tool developed by Bangham Engineering Incorporated (BEi) that produces a representative debris cloud from an exploding liquid-propellant launch vehicle. Here it is applied to the Core Stage (CS) of the National Aeronautics and Space Administration (NASA) Space Launch System (SLS launch vehicle). A combination of Probability Density Functions (PDF) based on empirical data from rocket accidents and applicable tests, as well as SLS specific geometry are combined in a MATLAB script to create unique fragment catalogues each time L-FOAM is run-tailored for a Monte Carlo approach for risk analysis. By accelerating the debris catalogue with the BEi blast model for liquid hydrogen / liquid oxygen explosions, the result is a fully integrated code that models the destruction of the CS at a given point in its trajectory and generates hundreds of individual fragment catalogues with initial imparted velocities. The BEi blast model provides the blast size (radius) and strength (overpressure) as probabilities based on empirical data and anchored with analytical work. The coupling of the L-FOAM catalogue with the BEi blast model is validated with a simulation of the Project PYRO S-IV destruct test. When running a Monte Carlo simulation, L-FOAM can accelerate all catalogues with the same blast (mean blast, 2 σ blast, etc.), or vary the blast size and strength based on their respective probabilities. L-FOAM then propagates these fragments until impact with the earth. Results from L-FOAM include a description of each fragment (dimensions, weight, ballistic coefficient, type and initial location on the rocket), imparted velocity from the blast, and impact data depending on user desired application. LFOAM application is for both near-field (fragment impact to escaping crew capsule) and far-field (fragment ground impact footprint) safety considerations. The user is thus able to use statistics from a Monte Carlo set of L-FOAM catalogues to quantify risk for a multitude of potential CS destruct scenarios. Examples include the effect of warning time on the survivability of an escaping crew capsule or the maximum fragment velocities generated by the ignition of leaking propellants in internal cavities.

Richardson, Erin↗