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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Test System to Study the Ignition of Metals by Polymers in Oxygen

A new test system that uses Laser energy to ignite a polymer promoter has been developed at the NASA White Sands Test Facility. It will facilitate the study of the spread of fire from a burning polymer material to the metal surrounding it. The system can be used to answer questions regarding the effects of configuration on ignition and combustion. The data obtained from this test could also be used to develop mathematical models for analyzing the effects of configuration on ignition and combustion. The system features a 10,000-psi (69-MPa) test chamber with sight glass windows on either end and a 25-watt carbon dioxide Laser for an ignition source. The test system can be used with gaseous oxygen, nitrogen or any mixture of the two gases. To minimize the effect of preheating the metallic, the polymer is ignited with a minimal amount of Laser energy. Igniting the polymer in this fashion also simplifies the thermodynamic analysis of the ignition and propagation reactions. The system is very robust, versatile and straightforward to use. Depending on the test pressure and configuration, the test system operator can perform as many as 20 tests per day. Test results verify that ignition and combustion of the metallic sample is not only dependent on pressure, material type and temperature, but configuration of both the polymer promoter and metallic sample. Both 6061 aluminum and 316 stainless steel 0.25-inch (6.35-mm) diameter rods with a standard 0-ring groove were tested with Buna-N, Silicone, Teflon and Viton 0-rings. The system ignited all four types of 0-rings in oxygen at pressures ranging from ambient to 10,000 psi (69 MPa). However, neither the stainless steel nor the aluminum rods on which the O-rings were mounted ignited in any test conditions. Future testing may be done on the 0.25-inch (6.35-mm) rod and O-ring configuration to evaluate the lack of ignition in these tests. Future configurations may include a plug of polymer in the base of the sample and replicas of fire-damaged components. Furthermore, the test system may be used in the future to analyze the oxidation rate of Laser-heated metals in gaseous oxygen.

Shoffstall, Michael S.↗
NASA NTRS2001

Laser Induced Aluminum Surface Breakdown Model

Laser powered propulsion systems involve complex fluid dynamics, thermodynamics and radiative transfer processes. Based on an unstructured grid., pressure-based computational aerothermodynamics, platform, several sub-nio"'dels describing such underlying physics as laser ray tracing and focusing, thermal non-equilibrium, plasma radiation and air spark ignition have been developed. This proposed work shall extend the numerical platform and existing sub-models to include the aluminum wall surface Inverse Bremsstrahlung (113) effect from which surface ablation and free-electron generation can be initiated without relying on the air spark ignition sub-model. The following tasks will be performed to accomplish the research objectives.

Chen, Yen-Sen↗
NASA NTRS1989

Thermodynamic and fluid mechanic analysis of rapid pressurization in a dead-end tube

Three models have been applied to very rapid compression of oxygen in a dead-ended tube. Pressures as high as 41 MPa (6000 psi) leading to peak temperatures of 1400 K are predicted. These temperatures are well in excess of the autoignition temperature (750 K) of teflon, a frequently used material for lining hoses employed in oxygen service. These findings are in accord with experiments that have resulted in ignition and combustion of the teflon, leading to the combustion of the stainless steel braiding and catastrophic failure. The system analyzed was representative of a capped off-high-pressure oxygen line, which could be part of a larger system. Pressurization of the larger system would lead to compression in the dead-end line, and possible ignition of the teflon liner. The model consists of a large plenum containing oxygen at the desired pressure (500 to 6000 psi). The plenum is connected via a fast acting valve to a stainless steel tube 2 cm inside diameter. Opening times are on the order of 15 ms. Downstream of the valve is an orifice sized to increase filling times to around 100 ms. The total length from the valve to the dead-end is 150 cm. The distance from the valve to the orifice is 95 cm. The models describe the fluid mechanics and thermodynamics of the flow, and do not include any combustion phenomena. A purely thermodynamic model assumes filling to be complete upstream of the orifice before any gas passes through the orifice. This simplification is reasonable based on experiment and computer modeling. Results show that peak temperatures as high as 4800 K can result from recompression of the gas after expanding through the orifice. An approximate transient model without an orifice was developed assuming an isentropic compression process. An analytical solution was obtained. Results indicated that fill times can be considerably shorter than valve opening times. The third model was a finite difference, 1-D transient compressible flow model. Results from the code show the recompression effect but predict much lower peak temperatures than the thermodynamic model.

Leslie, Ian H.↗
NASA NTRS2006

Free-jet Testing of a REST Scramjet at Off-Design Conditions

Scramjet flowpaths employing elliptical combustors have the potential to improve structural efficiency and performance relative to those using planar geometries. NASA Langley has developed a scramjet flowpath integrated into a lifting body vehicle, while transitioning from a rectangular capture area to both an elliptical throat and combustor. This Rectangular-to-Elliptical Shape Transition (REST) scramjet, has a design point of Mach 7.1, and is intended to operate with fixed-geometry between Mach 4.5 and 8.0. This paper describes initial free-jet testing of the heat-sink REST scramjet engine model at conditions simulating Mach 5.3 flight. Combustion of gaseous hydrogen fuel at equivalence ratios between 0.5 and 1.5 generated robust performance after ignition with a silane-hydrogen pilot. Facility model interactions were experienced for fuel equivalence ratios above 1.1, yet despite this, the flowpath was not unstarted by fuel addition at the Mach 5.3 test condition. Combustion tests at reduced stagnation enthalpy indicated that the engine self-started following termination of the fuel injection. Engine data is presented for the largest fuel equivalence ratio tested without facility interaction. These results indicate that this class of three-dimensional scramjet engine operates successfully at off-design conditions.

Smart, Michael K.↗
NASA NTRS1994

Theoretical Basis for Estimated Test Times and Conditions for Drop Tower and Space-Based Droplet Burning Experiments With Methanol and N-Heptane

In order to develop an extensive envelope of test conditions for NASA's space-based Droplet Combustion Experiment (DCE) as well those droplet experiments which can be performed using a drop tower, the transient vaporization and combustion of methanol and n-heptane droplets were simulated using a recently developed fully time-dependent, spherically symmetric droplet combustion model. The transient vaporization of methanol and n-heptane was modeled to characterize the instantaneous gas phase composition surrounding the droplet prior to the introduction of an ignition source. The results for methanol/air showed that the entire gas phase surrounding a 2 mm methanol droplet deployed in zero-g .quickly falls outside the lean flammability limit. The gas phase surrounding an identically-sized n-heptane droplet, on the other hand, remains flammable. The combustion of methanol was then modeled considering a detailed gas phase chemical kinetic mechanism (168 steps, 26 species) and the effect of the dissolution of flame-generated water into the liquid droplet. These results were used to determine the critical ignition diameter required to achieve quasi-steady droplet combustion in a given oxidizing environment. For droplet diameters greater than the critical ignition diameter, the model predicted a finite diameter at which the flame would extinguish. These extinction diameters were found to vary significantly with initial droplet diameter. This phenomenon appears to be unique to the transient heat transfer, mass transfer and chemical kinetics of the system and thus has not been reported elsewhere to date. The extinction diameter was also shown to vary significantly with the liquid phase Lewis number since the amount of water present in the droplet at extinction is largely governed by the rate at which water is transported into the droplet via mass diffusion. Finally, the numerical results for n-heptane combustion were obtained using both 2 step and 96 step semi-emperical chemical kinetic mechanisms. Neither mechanism exhibited the variation of extinction diameter with initial diameter.

Marchese, Anthony J.↗
NASA NTRS1974

An analysis of wildfire prevention

A model of the production of wildfire ignitions and damages is developed and used to determine wildland activity-regulation decisions, which minimize total expected cost-plus-loss due to wildfires. In this context, the implications of various policy decisions are considered. The resulting decision rules take a form that makes it possible for existing wildfire management agencies to readily adopt them upon collection of the required data.

Heineke, J. M.↗
NASA NTRS1980

Combustion and operating characteristics of spark-ignition engines

The spark-ignition engine turbulent flame propagation process was investigated. Then, using a spark-ignition engine cycle simulation and combustion model, the impact of turbocharging and heat transfer variations or engine power, efficiency, and NO sub x emissions was examined.

Heywood, J. B.↗
NASA NTRS1985

A Fundamental Study of Smoldering with Emphasis on Experimental Design for Zero-G

The objective of this section of the microgravity project is to identify key sets of low-gravity experiments which would critically compliment a larger set of more easily performed normal-gravity experiments to explain the phenomena found in smoldering combustion. It is planned to follow through on the conceptual design of these experiments by participating in the future in the fabrication of the refined apparatus and in the data collection and interpretation. Low-gravity experiments are appropriate for smoldering combustion because of the complexity of smoldering which requires every means possible to discriminate among the many chemical and physical mechanisms active in most smoldering combustion scenarios. Efforts will be primarily analytical, attempting to identify appropriate approximations and dominant dimensionless groups based on existing data and state-of-the-art combustion modelling. Transient stability questions such as ignition, extinction and the choices among charring, tarring, or flaming modes will be included.

Pagni, P. J.↗
NASA NTRS1992

Arcing in space structures in low Earth orbit

This report describes results of an experimental and theoretical program to investigate arcing of structures containing dielectric and conducting materials when they are biased negatively with respect to a plasma. An argon ion source generated Ar(+) ions of directed energy 20 to 40 eV and density approximately 10(exp 7) cm(exp -3) that impinged upon samples containing a dielectric material on top of a negatively biased Kovar plate. Arcing events were studied for bias voltages between -300 and -1000V with respect to the ion beam. The samples were Dow Corning 93-500 adhesive on Kovar, fused silica cover slips bonded on Kovar, and silicon solar cells mounted on Kovar. Measurements of discharge current, Kovar plate voltage, and radiation from the arc versus time were carried out. Microsecond duration exposure images and optical spectra in the 0.24 to 0.40 micron band were also acquired during arcing events. Arcing events were found to be associated with exposed adhesive and means were found to eliminate arcing altogether. The charging of a silica cover plate and the fields around the plate were calculated using a particle-in-cell code. Models were developed to explain the ignition of the arc and the physical processes occurring during the discharge.

Upschulte, B. L.↗
NASA NTRS1994

Axisymmetric Time-Dependent Computations of Expansion Tube Flow

The goal of this work is to add insight about the flow within expansion tubes by using computational fluid dynamics. This is accomplished by comparing the results of axisymmetric numerical simulations with finite-rate chemistry to data from the HYPULSE expansion tube facility which was previously the NASA Langley expansion tube. The numerical simulations begin at the opening of the primary diaphragm and compute the flow throughout the whole facility and, thus, are able to follow and assess the effect of many of the flow features created during operation of the facility. One particular issue that will be investigated is the effect of boundary layer formation in the acceleration tube on the test gas volume and test gas conditions. Both laminar and turbulent boundary layers will be implemented. The effect of momentary shock reflection off the secondary diaphragm will also be investigated. There is concern that such a reflection will stagnate the test gas and create high levels of dissociated molecules. This is particularly important in propulsion experiments where a freestream composition different from flight conditions may influence ignition and burning data. Several different models of diaphragm rupture will be implemented in order to help understand the importance of this issue.

Wilson, Gregory J.↗
NASA NTRS2015

Minimum deltaV Burn Planning for the International Space Station Using a Hybrid Optimization Technique, Level 1

The International Space Station's (ISS) trajectory is coordinated and executed by the Trajectory Operations and Planning (TOPO) group at NASA's Johnson Space Center. TOPO group personnel routinely generate look-ahead trajectories for the ISS that incorporate translation burns needed to maintain its orbit over the next three to twelve months. The burns are modeled as in-plane, horizontal burns, and must meet operational trajectory constraints imposed by both NASA and the Russian Space Agency. In generating these trajectories, TOPO personnel must determine the number of burns to model, each burn's Time of Ignition (TIG), and magnitude (i.e. deltaV) that meet these constraints. The current process for targeting these burns is manually intensive, and does not take advantage of more modern techniques that can reduce the workload needed to find feasible burn solutions, i.e. solutions that simply meet the constraints, or provide optimal burn solutions that minimize the total DeltaV while simultaneously meeting the constraints. A two-level, hybrid optimization technique is proposed to find both feasible and globally optimal burn solutions for ISS trajectory planning. For optimal solutions, the technique breaks the optimization problem into two distinct sub-problems, one for choosing the optimal number of burns and each burn's optimal TIG, and the other for computing the minimum total deltaV burn solution that satisfies the trajectory constraints. Each of the two aforementioned levels uses a different optimization algorithm to solve one of the sub-problems, giving rise to a hybrid technique. Level 2, or the outer level, uses a genetic algorithm to select the number of burns and each burn's TIG. Level 1, or the inner level, uses the burn TIGs from Level 2 in a sequential quadratic programming (SQP) algorithm to compute a minimum total deltaV burn solution subject to the trajectory constraints. The total deltaV from Level 1 is then used as a fitness function by the genetic algorithm in Level 2 to select the number of burns and their TIGs for the next generation. In this manner, the two levels solve their respective sub-problems separately but collaboratively until a burn solution is found that globally minimizes the deltaV across the entire trajectory. Feasible solutions can also be found by simply using the SQP algorithm in Level 1 with a zero cost function. This paper discusses the formulation of the Level 1 sub-problem and the development of a prototype software tool to solve it. The Level 2 sub-problem will be discussed in a future work. Following the Level 1 formulation and solution, several look-ahead trajectory examples for the ISS are explored. In each case, the burn targeting results using the current process are compared against a feasible solution found using Level 1 in the proposed technique. Level 1 is then used to find a minimum deltaV solution given the fixed number of burns and burn TIGs. The optimal solution is compared with the previously found feasible solution to determine the deltaV (and therefore propellant) savings. The proposed technique seeks to both improve the current process for targeting ISS burns, and to add the capability to optimize ISS burns in a novel fashion. The optimal solutions found using this technique can potentially save hundreds of kilograms of propellant over the course of the ISS mission compared to feasible solutions alone. While the software tool being developed to implement this technique is specific to ISS, the concept is extensible to other long-duration, central-body orbiting missions that must perform orbit maintenance burns to meet operational trajectory constraints.

Brown, Aaron J.↗
NASA NTRS1977

Symposium on Combustion /International/, 16th, Massachusetts Institute of Technology, Cambridge, Mass., August 15-20, 1976, Proceedings

Aspects of combustion technology in power systems are considered, taking into account a combustion in large boilers, the control of over-all thermal efficiency of combustion heating systems, a comparison of mathematical models of the radiative behavior of a large-scale experimental furnace, a concentric multiannular swirl burner, and the effects of water introduction on diesel engine combustion and emissions. Attention is also given to combustion and related processes in energy production from coal, spray and droplet combustion, soot formation and growth, the kinetics of elementary reactions, flame structure and chemistry, propellant ignition and combustion, fire and explosion research, mathematical modeling, high output combustion systems, turbulent flames and combustion, and ignition, optical, and electrical properties.

Source record↗
NASA NTRS1979

Global enclosure fire modeling with applications

A brief review of LERC (Limiting Energy Release Criteria) is presented. Application of LERC to JSC/Boeing ignition source full-scale tests and to JSC/Dacfir math-model validation tests is outlined. It is concluded that: (1) a complete LERC application to the JSC/Boeing tests verifies the fuel load criterion as the limiting constraint; (2) the variability of magnitude and form of the results of repeated tests with and without small variations in parameters emphasizes the significance of the local flow, species concentration, and heat transfer distribution; and (3) weight loss measurements of recent JSC tests show consistent results with prior methods, fuel load constrained.

Stuart, J. W.↗
NASA NTRS1998

Numerical Simulations of a Lobed Fuel Injector

Numerical modeling of the nonreactive mixing processes associated with a lobed fuel injector in a coflowing air stream is presented. The lobed fuel injector is a device which generates strong streamwise vorticity, producing locally high strain rates which can enhance the molecular mixing of reactants while delaying ignition in a controlled manner. Vortex element modeling is used to simulate flow field evolution and fuel element mixing characteristics for this lobed fuel injector. Quantitative predictions for vorticity generation and qualitative results for streamwise rollup compare well qualitatively with recent experimental investigations of this flow field [Smith et al, Phys. Fluids 9, 667 (1997)]. Parametric studies of the effects of lobe amplitude-to-wavelength ratio, lobe angle, and lobe shape for given flow conditions suggest that geometrical features may be optimized to enhance mixing and control reaction processes.

Strickland, J. H.↗
NASA NTRS2015

Characterization of the Ignition Over-Pressure/Sound Suppression Water in the Space Launch System Mobile Launcher Using Volume of Fluid Modeling

The Space Launch System (SLS) Vehicle consists of a Core Stage with four RS-25 engines and two Solid Rocket Boosters (SRBs). This vehicle is launched from the Launchpad using a Mobile Launcher (ML) which supports the SLS vehicle until its liftoff from the ML under its own power. The combination of the four RS-25 engines and two SRBs generate a significant Ignition Over-Pressure (IOP) and Acoustic Sound environment. One of the mitigations of these environments is the Ignition Over-Pressure/Sound Suppression (IOP/SS) subsystem installed on the ML. This system consists of six water nozzles located parallel to and 24 inches downstream of each SRB nozzle exit plane as well as 16 water nozzles located parallel to and 53 inches downstream of the RS-25 nozzle exit plane. During launch of the SLS vehicle, water is ejected through each water nozzle to reduce the intensity of the transient pressure environment imposed upon the SLS vehicle. While required for the mitigation of the transient pressure environment on the SLS vehicle, the IOP/SS subsystem interacts (possibly adversely) with other systems located on the Launch Pad. One of the other systems that the IOP/SS water is anticipated to interact with is the Hydrogen Burn-Off Igniter System (HBOI). The HBOI system's purpose is to ignite the unburned hydrogen/air mixture that develops in and around the nozzle of the RS-25 engines during engine start. Due to the close proximity of the water system to the HBOI system, the presence of the IOP/SS may degrade the effectiveness of the HBOI system. Another system that the IOP/SS water may interact with adversely is the RS-25 engine nozzles and the SRB nozzles. The adverse interaction anticipated is the wetting, to a significant degree, of the RS-25 nozzles resulting in substantial weight of ice forming and water present to a significant degree upstream of the SRB nozzle exit plane inside the nozzle itself, posing significant additional blockage of the effluent that exits the nozzle upon motor start leading to detrimental effects. The purpose of the CFD simulations were to i) characterize the location of the IOP/SS water after it is ejected from the IOP/SS nozzles, ii) characterize the interaction of the IOP/SS system with the HBOI system and iii) characterize the interaction of the IOP/SS water with the RS-25 nozzles and the SRB nozzles.

West, Jeff↗
NASA NTRS2024

Fundamental Experimental Tests and Modeling of LOX/CH4 Engines at High Pressures

In recent years, the space travel industry has grown exponentially, resulting in the need for a low-cost, efficient rocket engine fuel. As such, there has been renewed interest in utilizing liquid natural gas (LNG) as it is less likely to soot than kerosene-based fuels and has widespread availability and low cost of use as compared to the traditional RP-X and even liquid methane (LCH4) fuels due to the reduced need for refinement. However, current literature does not fully cover varying blends of natural gas and the effects that impurities have on natural gas/methane fuels, with most studies being confined to low-pressure applications. Therefore, new experimental ignition delay time measurements at rocket engine relevant pressures are essential to ensure current chemical kinetic models capture the ignition behavior at these elevated conditions and to determine the optimal purity levels for a reliable, cost-efficient, aerospace-grade LCH4/LNG fuel. The current work explores various blends of natural gas/methane fuel with various impurities, including higher hydrocarbons as well as nitrogen-carriers. A shock tube study was carried out to study the ignition delay times and carbon monoxide time histories of these natural gas blends utilizing chemiluminescence and laser absorption spectroscopy at 20, 50 and 100 atm over temperatures ranging from 1400-1700K and at an equivalence ratio of 1. The experimental data was then compared to two different chemical kinetic mechanisms: the industry-standard GRI 3.0 and the in-house UCF 2022. The data is used to improve chemical kinetic mechanisms and modeling of LNG fuels for rocket engine applications.

Methane↗
NASA NTRSDate not supplied

Clouds in the Arctic Climate System: A Force for Change?

Clouds play a central role in the global climate system through the modulation of Earth’s energy flows and as a mediator of precipitation. In the Arctic, clouds are also a major player; however, the processes that govern cloud evolution in the Arctic differ from most of the other regions of the globe. Thus, clouds are a key “wildcard” within the Arctic climate system that could have a substantial influence on the Arctic climate system response to anthropogenic forcing. What is the role of clouds within the phenomenon known as Arctic Amplification? This is question does not have a clear answer. Clouds seem to be a center to the important processes driving Arctic Amplification (e.g., the atmospheric response to sea ice loss and airmass transformation), however feedback analysis studies indicate that net cloud feedback in the Arctic is small. This seminar discusses the role of clouds within the Arctic Amplification processes reviewing aspects of what we know about Arctic clouds and the uncertainties that limit our ability to model them. Results from recently published and ongoing work are presented that provide an observationally-based estimate of the cloud-sea ice feedback and evaluate cloud properties within models. The goal of this presentation is to ignite a discussion and new collaborations around the best approaches to resolving uncertainties related to the role of clouds with the Arctic system and how to better represent Arctic clouds in models.

Patrick C Taylor↗