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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Thick thermal barrier coatings for diesel components

An engineered thick thermal barrier coating consisting of multiple layers of zirconia and CoCrAlY with a zirconia top layer and having a system thermal conductance less than 410 w/m(exp 2)K exceeded the 100 hour engine durability goals set forth in this program. The thermal barrier coatings were intact at the test conclusion. Back to back single cylinder research engine tests were conducted with watercooled, metal hardware and oil-cooled, thermal barrier coating insulated hardware to determine apparent heat release and fuel economy. Apparent heat release data revealed that the insulated engine had a shorter ignition delay and a longer combustion duration than the metal engine. The insulated engine fuel economy was approximately two percent worse on average for this series of tests. There was no attempt to optimize engine efficiency of the insulated engine by modifying the engine timing, coating, or other techniques.

Yonushonis, T. M.↗
NASA NTRS1989

Experimental and analytical investigations of wave enhanced supersonic combustors

Supersonic combustion ramjet (scramjet) engines rely on rapid mixing and combustion of fuel for good performance. However, both of these processes are relatively slow compared to the residence times in supersonic combusters. Methods of improving the mixing process include subjecting the fuel streams to shock waves and the generation of vortices by spiriling, bifurcating struts. The combustion process can also be enhanced by shock or detonation waves. An oblique shock wave can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture and thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. While the existence of standing oblique detonation waves has been proven in computer simulations, the experimental validation is still in progress. Currently, there is an experimental and analytical program at NASA-Ames Research Center to study detonations and other means of wave enhanced supersonic mixing and combustion.

Adelman, Henry G.↗
NASA NTRS1995

Oxidation Mechanisms of Toluene and Benzene

An expanded and improved version of a previously published benzene oxidation mechanism is presented and shown to model published experimental data fairly successfully. This benzene submodel is coupled to a modified version of a toluene oxidation submodel from the recent literature. This complete mechanism is shown to successfully model published experimental toluene oxidation data for a highly mixed flow reactor and for higher temperature ignition delay times in a shock tube. A comprehensive sensitivity analysis showing the most important reactions is presented for both the benzene and toluene reacting systems. The NASA Lewis toluene mechanism's modeling capability is found to be equivalent to that of the previously published mechanism which contains a somewhat different benzene submodel.

Bittker, David A.↗
NASA NTRS1996

Flamelet Model Application for Non-Premixed Turbulent Combustion

The current Final Report contains results of the study which was performed in Scientific Research Center 'ECOLEN' (Moscow, Russia). The study concerns the development and verification of non-expensive approach for modeling of supersonic turbulent diffusion flames based on flamelet consideration of the chemistry/turbulence interaction (FL approach). Research work included: development of the approach and CFD tests of the flamelet model for supersonic jet flames; development of the simplified procedure for solution of the flamelet equations based on partial equilibrium chemistry assumption; study of the flame ignition/extinction predictions provided by flamelet model. The performed investigation demonstrated that FL approach allowed to describe satisfactory main features of supersonic H 2/air jet flames. Model demonstrated also high capabilities for reduction of the computational expenses in CFD modeling of the supersonic flames taking into account detailed oxidation chemistry. However, some disadvantages and restrictions of the existing version of approach were found in this study. They were: (1) inaccuracy in predictions of the passive scalar statistics by our turbulence model for one of the considered test cases; and (2) applicability of the available version of the flamelet model to flames without large ignition delay distance only. Based on the results of the performed investigation, we formulated and submitted to the National Aeronautics and Space Administration our Project Proposal for the next step research directed toward further improvement of the FL approach.

Secundov, A.↗
NASA NTRS1996

NOx Emissions from a Lobed Fuel Injector/Burner

The present experimental study examines the performance of a novel fuel injector/burner configuration with respect to reduction in nitrogen oxide NOx emissions. The lobed injector/burner is a device in which very rapid initial mixing of reactants can occur through strong streamwise vorticity generation, producing high fluid mechanical strain rates which can delay ignition and thus prevent the formation of stoichiometric diffusion flames. Further downstream of the rapid mixing region. this flowfield produces a reduced effective strain rate, thus allowing ignition to occur in a premixed mode, where it is possible for combustion to take place under locally lean conditions. potentially reducing NOx emissions from the burner. The present experiments compare NO/NO2/NOx emissions from a lobed fuel injector configuration with emissions from a straight fuel injector to determine the net effect of streamwise vorticity generation. Preliminary results show that the lobed injector geometry can produce lean premixed flame structures. while for comparable flow conditions, a straight fuel injector geometry produces much longer. sooting diffusion flames or slightly rich pre-mixed flames. NO measurements show that emissions from a lobed fuel injector/burner can be made significantly lower than from a straight fuel injector under comparable flow conditions.

Mitchell, M. G.↗
NASA NTRS1999

Catalyst Development for Hydrogen Peroxide Rocket Engines

The development of various catalysts of hydrogen peroxide was conducted for the applications of liquid rocket engines. The catalyst development includes silver screen technology, solid catalyst technology, and homogeneous catalyst technology. The silver screen technology development was performed with 85% (by weight) hydrogen peroxide. The results of this investigation were used as the basis for the catalyst design of a pressure-fed liquid-fueled upper stage engine. Both silver-plated nickel 200 screens and pure silver screens were used as the active metal catalyst during the investigation, The data indicate that a high decomposition efficiency (greater than 90%) of 85% hydrogen peroxide can be achieved at a bed loading of 0.5 lbm/sq in/sec with both pure silver and silver plated screens. Samarium oxide coating, however, was found to retard the decomposition process and the catalyst bed was flooded at lower bed loading. A throughput of 200 lbm of hydrogen peroxide (1000 second run time) was tested to evaluate the catalyst aging issue and performance degradation was observed starting at approximately 400 seconds. Catalyst beds of 3.5 inch in diameter was fabricated using the same configuration for a 1,000-lbf rocket engine. High decomposition efficiency was obtained with a low pressure drop across the bed. Solid catalyst using precious metal was also developed for the decomposition of hydrogen peroxide from 85% to 98% by weight. Preliminary results show that the catalyst has a strong reactivity even after 15 minutes of peroxide decomposition. The development effort also includes the homogeneous catalyst technology. Various non-toxic catalysts were evaluated with 98% peroxide and hydrocarbon fuels. The results of open cup drop tests indicate an ignition delay around 11 ms.

Morlan, P. 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 NTRS1999

Durability, Performance, and Emission of Diesel Engines Using Carbon Fiber Piston and Liner

This report summarizes the research conducted by NC State University in investigating the durability, performance and emission of a carbon fiber piston and liner in our single cylinder research Diesel engine. Both the piston and liner were supplied to NC State University by NASA LaRC and manufactured by C-CAT under a separate contract to NASA LaRC. The carbon-carbon material used to manufacture the piston and liner has significantly lower thermal conductivity, coefficient of thermal expansion, and superior strength characteristics at elevated temperatures when compared to conventional piston materials such as aluminum. The results of the carbon-carbon fiber piston testing were compared to a baseline configuration, which used a conventional aluminum piston in a steel liner. The parameters measured were the brake specific fuel consumption, ignition delay, frictional horsepower, volumetric efficiency, and durability characteristics of the two pistons. Testing was performed using a naturally aspirated Labeco Direct Injection single cylinder diesel engine. Two test cases were performed over a range of loads and speeds. The fixed test condition between the aluminum and carbon-carbon piston configurations was the brake mean effective pressure. The measured data was the fuel consumption rate, volumetric efficiency, load, speed, cylinder pressure, needle lift, and exhaust gas temperature. The cylinder pressure, and fuel consumption, exhaust gas temperature, and needle lift were recorded using a National Instruments DAQ board and a PC. All test cases used Diesel no. 2 for fuel.

Afify, E. M.↗
NASA NTRS2001

Mars Flyer Rocket Propulsion Risk Assessment Kaiser Marquardt Testing

This report describes the investigation of a 10-N, bipropellant thruster, operating at -40 C, with monomethylhydrazine (MMH) and 25% nitric oxide in nitrogen tetroxide (MON-25). The thruster testing was conducted as part of a risk reduction activity for the Mars Flyer, a proposed mission to fly a miniature airplane in the Martian atmosphere. Testing was conducted using an existing thruster, designed for MMH and MON-3 propellants. The nitric oxide content of MON-3 was increased to 25%, to lower its freezing point to -55 C. The thruster was conditioned, along with the propellants, to temperature prior to hot firing. Thruster operating parameters included oxidizer-to-fuel mixture ratios of 1.6 to 2.7 and inlet pressure ranging from 689 to 2070 kPa. The test matrix consisted of many 10-second firings and several 60-, 300-, 600-, and 1200-second firings, as well as pulse testing. The thruster successfully accumulated nearly 10,000 seconds of operation without failure, at temperatures ranging from -40 C to 22 C. At nominal inlet pressures, the ignition delay was comparable to MMH/MON-3 operation. The optimal performance for the 8.9-N thruster was determined to be at a mixture ratio of 1.93 with an average specific impulse of 298 sec.

Marquardt, Kaiser↗
NASA NTRS1995

Review of the PDWA Concept for Combustion Enhancement in a Supersonic Air-Breathing Combustor Environment

This paper reviews the design of the Pulsed Detonation Wave Augmentor (PDWA) concept and the preliminary computational fluid dynamics studies that supported it. The PDWA relies on the rapid generation of detonation waves in a small tube, which are then injected into the supersonic stream of the main combustor. The blast waves thus generated are used to stimulate the mixing and combustion inside the main combustor. The mixing enhancement relies on various forms of the baroclinic interaction, where misaligned pressure and density gradients combine to produce vortical flow. By using unsteady shock waves, the concept also uses the Richtmyer-Meshkov effect to further increase the rate of mixing. By carefully designing the respective configurations of the combustor and the detonation tubes, one can also increase the penetration of the fuel into the supersonic air stream. The unsteady shocks produce lower stagnation pressure losses than steady shocks. Combustion enhancement can also be obtained through the transient shock-heating of the fuel-air interface, and the lowering of the ignition delay in these regions. The numerical simulations identify these processes, and show which configurations give the best results. Engineering considerations are also presented, and discuss the feasibility of the concept. Of primary importance are the enhancements in performance, the design simplicity, the minimization of the power, cost, and weight, and the methods to achieve very rapid cycling.

Canbier, Jean-Luc↗
NASA NTRS2003

An Earth-Based Equivalent Low Stretch Apparatus to Assess Material Flammability for Microgravity and Extraterrestrial Fire-Safety Applications

The objective of this project is to modify the standard oxygen consumption (cone) calorimeter (described in ASTM E 1354 and NASA STD 6001 Test 2) to provide a reproducible bench-scale test environment that simulates the buoyant or ventilation flow that would be generated by or around a burning surface in a spacecraft or extraterrestrial gravity level. This apparatus will allow us to conduct normal gravity experiments that accurately and quantitatively evaluate a material's flammability characteristics in the real-use environment of spacecraft or extra-terrestrial gravitational acceleration. The Equivalent Low Stretch Apparatus (ELSA) uses an inverted cone geometry with the sample burning in a ceiling fire configuration that provides a reproducible bench-scale test environment that simulates the buoyant or ventilation flow that would be generated by a flame in a spacecraft or extraterrestrial gravity level. Prototype unit testing results are presented in this paper. Ignition delay times and regression rates for PMMA are presented over a range of radiant heat flux levels and equivalent stretch rates which demonstrate the ability of ELSA to simulate key features of microgravity and extraterrestrial fire behavior.

Olson, S. L.↗
NASA NTRS1997

An Investigation to Improve Quality Evaluations of Primers and Propellant for 20mm Munitions

To reduce the frequency of electrically initiated, 20mm munition hangfires (delayed ignitions), a joint Army/NASA investigation was conducted to recommend quality evaluation improvements for acceptance of both primers and gun propellant. This effort focused only on evaluating ignition and combustion performance as potential causes of hangfires: poor electrical initiation of the primer, low output performance of the primer, low ignition sensitivity of the gun propellant, and the effects of cold temperature. The goal was to determine the "best" of the Army and NASA test methods to assess the functional performance of primers and gun propellants. The approach was to evaluate the performance of both high-quality and deliberately defective primers to challenge the sensitivity of test methods. In addition, the ignition sensitivity of different manufacturing batches of gun propellants was evaluated. The results of the investigation revealed that improvements can be made in functional evaluations that can assist in identifying and reducing ignition and performance variations. The "best" functional evaluation of primers and propellant is achieved through a combination of both Army and NASA test methods. Incorporating the recommendations offered in this report may provide for considerable savings in reducing the number of cartridge firings, while significantly lowering the rejection rate of primer, propellant and cartridge lots. The most probable causes for ignition and combustion-related hangfires were the lack of calcium silicide in the primer mix, a low output performance of primers, and finally, poor ignition sensitivity of gun propellant. Cold temperatures further reduce propellant ignition sensitivity, as well as reducing burn rate and chamber pressures.

Bement, L. J.↗
NASA NTRS2006

Fuel-Air Mixing and Combustion in Scramjets

At flight speeds, the residence time for atmospheric air ingested into a scramjet inlet and exiting from the engine nozzle is on the order of a millisecond. Therefore, fuel injected into the air must efficiently mix within tens of microseconds and react to release its energy in the combustor. The overall combustion process should be mixing controlled to provide a stable operating environment; in reality, however, combustion in the upstream portion of the combustor, particularly at higher Mach numbers, is kinetically controlled where ignition delay times are on the same order as the fluid scale. Both mixing and combustion time scales must be considered in a detailed study of mixing and reaction in a scramjet to understand the flow processes and to ultimately achieve a successful design. Although the geometric configuration of a scramjet is relatively simple compared to a turbomachinery design, the flow physics associated with the simultaneous injection of fuel from multiple injector configurations, and the mixing and combustion of that fuel downstream of the injectors is still quite complex. For this reason, many researchers have considered the more tractable problem of a spatially developing, primarily supersonic, chemically reacting mixing layer or jet that relaxes only the complexities introduced by engine geometry. All of the difficulties introduced by the fluid mechanics, combustion chemistry, and interactions between these phenomena can be retained in the reacting mixing layer, making it an ideal problem for the detailed study of supersonic reacting flow in a scramjet. With a good understanding of the physics of the scramjet internal flowfield, the designer can then return to the actual scramjet geometry with this knowledge and apply engineering design tools that more properly account for the complex physics. This approach will guide the discussion in the remainder of this section.

Drummond, J. Philip↗
NASA NTRS2008

Gaseous Surrogate Hydrocarbons for a Hifire Scramjet that Mimic Opposed Jet Extinction Limits for Cracked JP Fuels

This paper describes, first, the top-down methodology used to define simple gaseous surrogate hydrocarbon (HC) fuel mixtures for a hypersonic scramjet combustion subtask of the HiFIRE program. It then presents new and updated Opposed Jet Burner (OJB) extinction-limit Flame Strength (FS) data obtained from laminar non-premixed HC vs. air counterflow diffusion flames at 1-atm, which follow from earlier investigations. FS represents a strain-induced extinction limit based on cross-section-average air jet velocity, U(sub air), that sustains combustion of a counter jet of gaseous fuel just before extinction. FS uniquely characterizes a kinetically limited fuel combustion rate. More generally, Applied Stress Rates (ASRs) at extinction (U(sub air) normalized by nozzle or tube diameter, D(sub n or t) can directly be compared with extinction limits determined numerically using either a 1-D or (preferably) a 2-D Navier Stokes simulation with detailed transport and finite rate chemistry. The FS results help to characterize and define three candidate surrogate HC fuel mixtures that exhibit a common FS 70% greater than for vaporized JP-7 fuel. These include a binary fuel mixture of 64% ethylene + 36% methane, which is our primary recommendation. It is intended to mimic the critical flameholding limit of a thermally- or catalytically-cracked JP-7 like fuel in HiFIRE scramjet combustion tests. Our supporting experimental results include: (1) An idealized kinetically-limited ASR reactivity scale, which represents maximum strength non-premixed flames for several gaseous and vaporized liquid HCs; (2) FS characterizations of Colket and Spadaccini s suggested ternary surrogate, of 60% ethylene + 30% methane + 10% n-heptane, which matches the ignition delay of a typical cracked JP fuel; (3) Data showing how our recommended binary surrogate, of 64% ethylene + 36% methane, has an identical FS; (4) Data that characterize an alternate surrogate of 44% ethylene + 56% ethane with identical FS and nearly equal molecular weights; this could be useful when systematically varying the fuel composition. However, the mixture liquefies at much lower pressure, which limits on-board storage of gaseous fuel; (5) Dynamic Flame Weakening results that show how oscillations in OJB input flow (and composition) can weaken (extinguish) surrogate flames up to 200 Hz, but the weakening is 2.5x smaller compared to pure methane; and finally, (6) FS limits at 1-atm that compare with three published 1-D numerical OJB extinction results using four chemical kinetic models. The methane kinetics generally agree closely at 1-atm, whereas, the various ethylene models predict extinction limits that average ~ 45% high, which represents a significant problem for numerical simulation of surrogate-based flameholding in a scramjet cavity. Finally, we continue advocating the FS approach as more direct and fundamental for assessing idealized scramjet flameholding potentials than measurements of "unstrained" premixed laminar burning velocity or blowout in a Perfectly Stirred Reactor.

Pellett, Gerald L.↗
NASA NTRS2012

Large Scale Experiments on Spacecraft Fire Safety

Full scale fire testing complemented by computer modelling has provided significant know how about the risk, prevention and suppression of fire in terrestrial systems (cars, ships, planes, buildings, mines, and tunnels). In comparison, no such testing has been carried out for manned spacecraft due to the complexity, cost and risk associated with operating a long duration fire safety experiment of a relevant size in microgravity. Therefore, there is currently a gap in knowledge of fire behaviour in spacecraft. The entire body of low-gravity fire research has either been conducted in short duration ground-based microgravity facilities or has been limited to very small fuel samples. Still, the work conducted to date has shown that fire behaviour in low-gravity is very different from that in normal-gravity, with differences observed for flammability limits, ignition delay, flame spread behaviour, flame colour and flame structure. As a result, the prediction of the behaviour of fires in reduced gravity is at present not validated. To address this gap in knowledge, a collaborative international project, Spacecraft Fire Safety, has been established with its cornerstone being the development of an experiment (Fire Safety 1) to be conducted on an ISS resupply vehicle, such as the Automated Transfer Vehicle (ATV) or Orbital Cygnus after it leaves the ISS and before it enters the atmosphere. A computer modelling effort will complement the experimental effort. Although the experiment will need to meet rigorous safety requirements to ensure the carrier vehicle does not sustain damage, the absence of a crew removes the need for strict containment of combustion products. This will facilitate the possibility of examining fire behaviour on a scale that is relevant to spacecraft fire safety and will provide unique data for fire model validation. This unprecedented opportunity will expand the understanding of the fundamentals of fire behaviour in spacecraft. The experiment is being developed by an international topical team that is collaboratively defining the experiment requirements and performing supporting analysis, experimentation and technology development. This paper presents the objectives, status and concept of this project.

Urban, David L.↗
NASA NTRS2012

Large Scale Experiments on Spacecraft Fire Safety

Full scale fire testing complemented by computer modelling has provided significant knowhow about the risk, prevention and suppression of fire in terrestrial systems (cars, ships, planes, buildings, mines, and tunnels). In comparison, no such testing has been carried out for manned spacecraft due to the complexity, cost and risk associated with operating a long duration fire safety experiment of a relevant size in microgravity. Therefore, there is currently a gap in knowledge of fire behaviour in spacecraft. The entire body of low-gravity fire research has either been conducted in short duration ground-based microgravity facilities or has been limited to very small fuel samples. Still, the work conducted to date has shown that fire behaviour in low-gravity is very different from that in normal gravity, with differences observed for flammability limits, ignition delay, flame spread behaviour, flame colour and flame structure. As a result, the prediction of the behaviour of fires in reduced gravity is at present not validated. To address this gap in knowledge, a collaborative international project, Spacecraft Fire Safety, has been established with its cornerstone being the development of an experiment (Fire Safety 1) to be conducted on an ISS resupply vehicle, such as the Automated Transfer Vehicle (ATV) or Orbital Cygnus after it leaves the ISS and before it enters the atmosphere. A computer modelling effort will complement the experimental effort. Although the experiment will need to meet rigorous safety requirements to ensure the carrier vehicle does not sustain damage, the absence of a crew removes the need for strict containment of combustion products. This will facilitate the possibility of examining fire behaviour on a scale that is relevant to spacecraft fire safety and will provide unique data for fire model validation. This unprecedented opportunity will expand the understanding of the fundamentals of fire behaviour in spacecraft. The experiment is being developed by an international topical team that is collaboratively defining the experiment requirements and performing supporting analysis, experimentation and technology development. This paper presents the objectives, status and concept of this project.

Urban, David↗
NASA NTRS2014

Cool-flame Extinction During N-Alkane Droplet Combustion in Microgravity

Recent droplet combustion experiments onboard the International Space Station (ISS) have revealed that large n-alkane droplets can continue to burn quasi-steadily following radiative extinction in a low-temperature regime, characterized by negative-temperaturecoefficient (NTC) chemistry. In this study we report experimental observations of n-heptane, n-octane, and n-decane droplets of varying initial sizes burning in oxygen/nitrogen/carbon dioxide and oxygen/helium/nitrogen environments at 1.0, 0.7, and 0.5 atmospheric pressures. The oxygen concentration in these tests varied in the range of 14% to 25% by volume. Large n-alkane droplets exhibited quasi-steady low-temperature burning and extinction following radiative extinction of the visible flame while smaller droplets burned to completion or disruptively extinguished. A vapor-cloud formed in most cases slightly prior to or following the "cool flame" extinction. Results for droplet burning rates in both the hot-flame and cool-flame regimes as well as droplet extinction diameters at the end of each stage are presented. Time histories of radiant emission from the droplet captured using broadband radiometers are also presented. Remarkably the "cool flame" extinction diameters for all the three n-alkanes follow a trend reminiscent of the ignition delay times observed in previous studies. The similarities and differences among the n-alkanes during "cool flame" combustion are discussed using simplified theoretical models of the phenomenon

microgravity combustion↗
NASA NTRS2017

Fuel and Combustor Concerns for Future Commercial Combustors

Civil aircraft combustor designs will move from rich-burn to lean-burn due to the latter's advantage in low NOx and nvPM emissions. However, the operating range of lean-burn is narrower, requiring premium mixing performance from the fuel injectors. As the OPR increases, the corresponding combustor inlet temperature increase can benefit greatly with fuel composition improvements. Hydro-treatment can improve coking resistance, allowing finer fuel injection orifices to speed up mixing. Selective cetane number control across the fuel carbon-number distribution may allow delayed ignition at high power while maintaining low-power ignition characteristics.

Combustor↗