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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Ignition and combustion characteristics of metallized propellants

Experimental and analytical investigations focusing on secondary atomization and ignition characteristics of aluminum/liquid hydrocarbon slurry propellants were conducted. Experimental efforts included the application of a laser-based, two-color, forward-scatter technique to simultaneously measure free-flying slurry droplet diameters and velocities for droplet diameters in the range of 10-200 microns. A multi-diffusion flame burner was used to create a high-temperature environment into which a dilute stream of slurry droplets could be introduced. Narrowband measurements of radiant emission were used to determine if ignition of the aluminum in the slurry droplet had occurred. Models of slurry droplet shell formation were applied to aluminum/liquid hydrocarbon propellants and used to ascertain the effects of solids loading and ultimate particle size on the minimum droplet diameter that will permit secondary atomization. For a 60 weight-percent Al slurry, the limiting critical diameter was predicted to be 34.7 microns which is somewhat greater than the 20-25 micron limiting diameters determined in the experiments. A previously developed model of aluminum ignition in a slurry droplet was applied to the present experiments and found to predict ignition times in reasonable agreement with experimental measurements. A model was also developed that predicts the mechanical stress in the droplet shell and a parametric study was conducted. A one-dimensional model of a slurry-fueled rocket combustion chamber was developed. This model includes the processes of liquid hydrocarbon burnout, secondary atomization, aluminum ignition, and aluminum combustion. Also included is a model for radiant heat transfer from the hot aluminum oxide particles to the chamber walls. Exercising this model shows that only a modest amount of secondary atomization is required to reduce residence times for aluminum burnout, and thereby maintain relatively short chamber lengths. The model also predicts radiant heat transfer losses to the walls to be only approximately 3 percent of the fuel energy supplied. Additional work is required to determine the effects of secondary atomization on two-phase losses in the nozzle.

Turns, Stephen R.↗
NASA NTRS1988

Evaporation And Ignition Of Dense Fuel Sprays

Simple theoretical model makes useful predictions of trends. Pair of reports presents theoretical model of evaporation and ignition of sprayed liquid fuel. Developed as part of research in combustion of oil and liquid fuels derived from coal, tar sand, and shale in furnace. Work eventually contributes to increase efficiency of combustion and decrease pollution generated by burning of such fuels.

Bellan, Josette↗
NASA NTRS1993

Ignition and subsequent flame spread over a thin cellulosic material

Both ignition and flame spread on solid fuels are processes that not only are of considerable scientific interest but that also have important fire safety applications. Both types of processes, ignition and flame spread, are complicated by strong coupling between chemical reactions and transport processes, not only in the gas phase but also in the condensed phase. In most previous studies, ignition and flame spread were studied separately with the result that there has been little understanding of the transition from ignition to flame spread. In fire safety applications this transition is crucial to determine whether a fire will be limited to a localized, temporary burn or will transition into a growth mode with a potential to become a large fire. In order to understand this transition, the transient mechanisms of ignition and subsequent flame spread must be studied. However, there have been no definitive experimental or modeling studies, because of the complexity of the flow motion generated by buoyancy near the heated sample surface. One must solve the full Navier-Stokes equations over an extended region to represent accurately the highly unstable buoyant plume and entrainment of surrounding gas from far away. In order to avoid the complicated nature of the starting plume problem under normal gravity, previous detailed radiative ignition models were assumed to be one-dimensional or were applied at a stagnation point. Thus, these models cannot be extended to include the transition to flame spread. The mismatch between experimental and calculated geometries means that theories cannot be compared directly with experimental results in normal gravity. To overcome the above difficulty, theoretical results obtained without buoyancy can be directly compared with experimental data measured in a microgravity environment. Thus, the objective of this study is to develop a theoretical model for ignition and the transition to flame spread and to make predictions using the thermal and chemical characteristics of a cellulosic material which are measured in normal gravity.

Nakabe, Kazuyoshi↗
DOE OSTI2024

A comprehensive numerical investigation on spray models for Direct-Injection Spark-Ignition engines

Gasoline direct-injection spark-ignition (DISI) engines generate a large portion of their unburned hydrocarbon (UHC) and soot emissions during the cold-start phase. A predictive computational fluid dynamics (CFD) modeling framework can be used to understand the physical processes that characterize fuel spray evolution and fuel-film formation at cold start conditions, which can help to reduce engine-out particulate emissions. This study systematically evaluated spray submodels and developed a set of simulation best practices for physical-numerical submodels with the goal of enabling accurate simulations of liquid spray behavior in a DISI engine. Three comprehensive experimental datasets containing free-spray projected liquid volume (PLV), liquid volume fraction (LVF), and near-field X-ray radiography data were used to validate the simulation results and evaluate the spray submodels. Systematic analysis delved into injected parcel distribution, droplet collision, spray breakup, and evaporation via a detailed assessment of the relevant spray submodels. Moreover, the effects of turbulence models and the initial turbulent flow properties on the liquid spray evolution were examined. Based on extensive calibration efforts, a set of simulation best practices for the free spray was developed and validated against the PLV/LVF data. Simulation results indicated that the uniform distribution for parcel initialization, coupled with appropriate droplet collision submodels, provides an improved spray morphology compared to the cluster distribution. The findings also underscored the importance of calibrating the Kelvin-Helmholtz Rayleigh-Taylor (KH-RT) breakup model constants and droplet heat transfer coefficient scaling factor to achieve favorable agreement regarding measured liquid penetration and spray widths. In conclusion, this study marks a substantial stride towards accurately predicting fuel film evolution and soot formation within DISI engine performance.

ECN Spray G↗
NASA NTRS1992

Microgravity ignition experiment

The purpose of this project is to develop a flight-ready apparatus of the microgravity ignition experiment for the GASCAN 2 program. The microgravity ignition experiment is designed to study how a microgravity environment affects the time to ignition of a sample of alpha-cellulose paper. A microgravity environment will result in a decrease in the heat transferred from the sample due to a lack of convection currents, which would decrease time to ignition. A lack of convection current would also cause the oxygen supply at the sample not to be renewed, which could delay or even prevent ignition. When this experiment is conducted aboard GASCAN 2, the dominant result of the lack of ignition will be determined. The experiment consists of four canisters containing four thermocouples and a sensor to detect ignition of the paper sample. This year the interior of the canister was redesigned and a mathematical model of the heat transfer around the sample was developed. This heat transfer model predicts an ignition time of approximately 5.5 seconds if the decrease of heat loss from the sample is the dominant factor of the lack of convection currents.

Motevalli, Vahid↗
NASA NTRS1997

Demonstration of Oxygen and Carbon Monoxide Propellants for Mars Missions

Currently, proposed planetary exploration missions must be small, with low costs and a short development time. Relatively high-risk technologies are being accepted for such missions if they meet these guidelines. For a Mars sample-return mission, one of the higher risk technologies is the use of return propellants produced from indigenous materials such as the Martian atmosphere. This consists of 96 percent carbon dioxide, which can be processed into oxygen and carbon monoxide. This year, the NASA Lewis Research Center completed the experimental evaluation and subscale technology development of an oxygen/carbon monoxide propellant combination. Previous research included ignition characterization, combustion performance, and heat transfer characterization with gaseous propellants at room temperature. In this year s tests, we studied the ignition characteristics and combustion of oxygen and carbon monoxide at near liquid temperatures. The mixture ratio boundaries for oxygen and carbon monoxide were determined as a function of propellant temperature in a spark torch igniter. With both propellants at room temperature, the ignition range was between 0.50 and 1.44; and with both propellants chilled to near-liquid temperatures, it was between 2.4 and 3.1. Statistical analysis of the mean value of the ignition boundaries provided models that describe the combination of oxygen temperature, carbon monoxide temperature, and mixture ratio that resulted in ignition. This range is the larger boxed area shown in the figure. The smaller boxed area indicates the range at which there is a 90-percent confidence that ignition will occur. The relatively small range at only 90-percent confidence indicates that using the oxygen/carbon monoxide combination as its own ignition source may not be the best design for a remote engine operating on Mars. Tests also were performed in a simulated small rocket engine that used oxygen/hydrogen combustion gases as the ignition source for oxygen/carbon monoxide. In these experiments, the oxygen/carbon monoxide was successfully ignited in eight of eight tests at a mixture ratio of 0.52. In addition, the oxygen/carbon monoxide maintained steady combustion after the oxygen/hydrogen ignition source was removed, verifying that the oxygen/carbon monoxide rocket engine should continue to be included in mission plans as return propulsion from Mars.

Linne, Diane L.↗
NASA NTRS1995

Droplet Cluster Behavior in Dense and Dilute Regions of a Spray

A review is presented of modeling work relevant to evaporation and combustion in liquid rocket engine chamber. Highlights of models describing evaporation, ignition, and combustion of clusters of binary-fuel drops are discussed along with strategies for using these models to construct a model of a liquid rocket combustion chamber.

rockets combustion droplets↗
NASA NTRS1995

Ignition and subsequent transition to flame spread in a microgravity environment

The fire safety strategy in a spacecraft is (1) to detect any fire as early as possible, (2) to keep any fire as small as possible, and (3) to extinguish any fire as quickly as possible. This suggests that a material which undergoes a momentary, localized ignition might be tolerable but a material which permits a transition to flame spread would significantly increase the fire hazard. Therefore, it is important to understand how the transition from localized ignition to flame spread occurs and what parameters significantly affect the transition. The fundamental processes involved in ignition and flame spread have been extensively studied, but they have been studied separately. Some of the steady state flame models start from ignition to reach a steady state, but since the objective of such a calculation is to obtain the steady state flame spread rate, the calculation through the transition process is made without high accuracy to save computational time. We have studied the transition from a small localized ignition at the center of a thermally thin paper in a microgravity environment. The configuration for that study was axisymmetric, but more general versions of the numerical scheme have been developed by including the effects of a slow, external flow in both two and three dimensions. By exploiting the non-buoyant nature of the flow, it is possible to achieve resolution of fractions of millimeters for 3D flow domains on the order of 10 centimeters. Because the calculations are time dependent, we can study the evolution of multiple flame fronts originating from a localized ignition source. The interaction of these fronts determines whether or not they will eventually achieve steady state spread. Most flame spread studies in microgravity consider two-dimensional flame spread initiated by ignition at one end of a sample strip with or against a slow external flow. In this configuration there is only one flame front. A more realistic scenario involves separate, oppositely directed fronts in two dimensions, or a continuous, radially directed front in three dimensions. We present here some results of both the two and three dimensional codes.

Kashiwagi, Takashi↗
NASA NTRS2013

State Machine Modeling of the Space Launch System Solid Rocket Boosters

The Space Launch System is a Shuttle-derived heavy-lift vehicle currently in development to serve as NASA's premiere launch vehicle for space exploration. The Space Launch System is a multistage rocket with two Solid Rocket Boosters and multiple payloads, including the Multi-Purpose Crew Vehicle. Planned Space Launch System destinations include near-Earth asteroids, the Moon, Mars, and Lagrange points. The Space Launch System is a complex system with many subsystems, requiring considerable systems engineering and integration. To this end, state machine analysis offers a method to support engineering and operational e orts, identify and avert undesirable or potentially hazardous system states, and evaluate system requirements. Finite State Machines model a system as a finite number of states, with transitions between states controlled by state-based and event-based logic. State machines are a useful tool for understanding complex system behaviors and evaluating "what-if" scenarios. This work contributes to a state machine model of the Space Launch System developed at NASA Ames Research Center. The Space Launch System Solid Rocket Booster avionics and ignition subsystems are modeled using MATLAB/Stateflow software. This model is integrated into a larger model of Space Launch System avionics used for verification and validation of Space Launch System operating procedures and design requirements. This includes testing both nominal and o -nominal system states and command sequences.

model-based systems engineering↗
NASA NTRS1992

Burning Of Dense Clusters Of Fuel Drops

Report presents theoretical study of evaporation, ignition, and combustion of rich and relatively dense clusters of drops of liquid fuel. Focus on interactions between heterogenous liquid/gas mixture in cluster and flame surrounding it. Theoretical model of evaporation, ignition, and combustion presented.

Bellan, Josette↗
NASA NTRS1991

Experimental investigation of the flow field in the head-end star slot section of a solid rocket motor

A test program has been developed to investigate the complex flowfield in the head-end, star grain section of a solid rocket motor. Documentation of the flowfield pattern is to be obtained using both flow visualization techniques and quantitative methods, including static pressure measurements, heat transfer coefficient measurements, and hot-wire anemometry. A cold-flow, one-tenth-scale model based on the geometry of the Space Shuttle SRM head-end section and its single-port igniter, along with two models of a four-port igniter, have been constructed. The scale factor is derived from a scaling analysis which matches Reynolds number between the model and the full-scale flowfield in the star grain section. The NASA Marshall Space Flight Center 14 x 14 inch trisonic wind tunnel will be utilized for the cold flow tests. Since the work is still in progress, no results are included.

Ciucci, A.↗
DOE OSTI2024

Hierarchically Informed Engineering Models for Predictive Modeling of Turbulent Premixed Flame Propagation in Pre- chamber Turbulent Jet Ignition

The goal of the project is to improve the predictive accuracy and efficiency of turbulent combustion sub-models for pre-chamber turbulent jet ignition (TJI). This goal is achieved through the development of a hierarchically informed engineering model for turbulent combustion in TJI. The model development starts with the highest level of model description of turbulent combustion with direct numerical simulation (DNS) from which fundamental characteristics and scaling properties of turbulent premixed flame propagation under TJI relevant conditions are obtained.

42 ENGINEERING↗
NASA NTRS1995

Ignition and combustion of bulk metals at normal, elevated and reduced gravity

Knowledge of the oxidation, ignition, and combustion of bulk metals is important for fire safety in the production, management, and utilization of liquid and gaseous oxygen for ground based and space applications. This proposal outlines studies in continuation of research initiated earlier under NASA support to investigate the ignition and combustion characteristics of bulk metals under varying gravity conditions. Metal ignition and combustion have not been studied previously under these conditions and the results are important not only for improved fire safety but also to increase knowledge of basic ignition and combustion mechanisms. The studies completed to date have led to the development of a clean and reproducible ignition source and diagnostic techniques for combustion measurements and have provided normal, elevated, and reduced gravity combustion data on a variety of different pure metals. The research conducted under this grant will use the apparatus and techniques developed earlier to continue the elevated and low gravity experiments, and to develop the overall modeling of the ignition and combustion process. Metal specimens are to be ignited using a xenon short-arc lamp and measurements are to be made of the ignition energy, surface temperature history, burning rates, spectroscopy of surface and gas products, and surface morphology and chemistry. Elevated gravity will be provided by the University of Colorado Geotechnical Centrifuge and microgravity will be obtained in NASA's DC-9 Reduced Gravity aircraft.

Branch, Melvyn C.↗
NASA NTRS1981

A pocket model for aluminum agglomeration in composite propellants

This paper presents a model for the purpose of estimating the fraction of aluminum powder that will form agglomerates at the surface of deflagrating composite propellants. The basic idea is that the fraction agglomerated depends upon the amount of aluminum that melts within effective binder pocket volumes framed by oxidizer particles. The effective pocket depends upon the ability of ammonium perchlorate modals to encapsulate the aluminum and provide a local temperature sufficient to ignite the aluminum. Model results are discussed in the light of data showing effects of propellant formulation variables and pressure.

Cohen, N. S.↗
NASA NTRS1992

Results of an experimental investigation of the flow field in the head-end star slot section of a Solid Rocket Motor

Cold flow tests of a one-tenth scale model based on the geometry of the Space Shuttle Solid Rocket Motor, SRM, head-end section and its single-port igniter, along with two models of a four port igniter, have been conducted. The tests were done to establish quantitative as well as qualitative data on the behavior of the flow inside the star slot during the early part of the ignition transient. This paper presents the data obtained from these tests and discusses the implications of the data with respect to ignition transients of solid rocket motors which have head-end star grains.

Ciucci, A.↗
DOE OSTI2026

An analysis of alcohol compression ignition strategies on different engine platforms

Strategies that enable ignition of the low cetane alcohol fuels (methanol and ethanol) include intake heating (direct or via hot residual trapping), increased compression ratio (CR), multi-injection strategies, ignition enhancers, and active prechambers. The literature consistently shows alcohol compression ignition (CI) is more efficient than diesel at high loads irrespective of ignition strategy but shows contradictory results at low loads. To understand this, this work combines data from four different engine platforms ranging from a displaced volume of 0.4 to 2.1 L/cyl. using different alcohol CI ignition strategies. Using modeling to support experimental analysis, alcohol-specific efficiency penalties are highlighted to inform alcohol combustion system development. The alcohols’ lower energy density means more fuel must be injected near top dead center to achieve a given load, increasing the relative penalty of sensible/latent heating. Pilot injections help reduce this penalty. Intake heating incurs both a heat transfer penalty and a thermodynamic penalty due to a reduced ratio of specific heats during compression. A strategy that combines elevated CR, large pilot injections, and ignition enhancers can avoid this penalty. Alternatively, a prechamber strategy can achieve this without changing the CR, nor needing ignition enhancers in the fuel. At low loads, diesel burns in a partially premixed mode characterized by high burn rates and low heat losses, reducing alcohol-specific combustion benefits. Alcohol CI on the light-duty platform showed abnormally high heat transfer over diesel compared to other platforms, favoring a more premixed/partially premixed strategy despite a combustion efficiency tradeoff.

33 ADVANCED PROPULSION SYSTEMS↗