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NASA NTRS1939

Knocking in the Otto-cycle Engine

Engine knock is, as is known, preceded by normal burning of the first part of the charge, and only the part burned last (residual charge), knocks. The aim of the present measurements was, first, to reexamine the combustion form in this residual charge, because of the absence of uniform and frequently contradictory results in the very extensive literature on the subject. On top of that, an attempt was to be made to gain a deeper insight into the mechanism accompanying the combustion process, by means of the electrical test equipment perfected in recent years.

Weinhart, H↗

NASA NTRS1993

CFD analyses of combustor and nozzle flowfields

The objectives of the research are to improve design capabilities for low thrust rocket engines through understanding of the detailed mixing and combustion processes. A Computational Fluid Dynamic (CFD) technique is employed to model the flowfields within the combustor, nozzle, and near plume field. The computational modeling of the rocket engine flowfields requires the application of the complete Navier-Stokes equations, coupled with species diffusion equations. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an ongoing experimental program at NASA LeRC. The numerical procedure is performed on both time-marching and time-accurate algorithms, using an LU approximate factorization in time, flux split upwinding differencing in space. The integrity of fuel film cooling along the wall, its effectiveness in the mixing with the core flow including unsteady large scale effects, the resultant impact on performance and the assessment of the near plume flow expansion to finite pressure altitude chamber are addressed.

Tsuei, Hsin-Hua↗

NASA NTRS1993

Small rocket research and technology

Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a ceramic composite of mixed hafnium carbide and tantalum carbide reinforced with graphite fibers.

Schneider, Steven↗

NASA NTRS1993

Inherent stability of central element coaxial liquid-liquid injectors

Most TRW liquid bi-propellant rocket engines built over the past thirty-plus years have employed a central element coaxial pintle injector and have operated with liquid/liquid propellant injection. This injector is a patented design exclusive to TRW and has unique features that make the rocket engine combustion characteristics different from those of other types of injector engine designs. Its many benefits include excellent combustion performance, efficient deep throttling, adaptability to low cost manufacturing, and high reliability. Approximately 200 pintle injector engines of various sizes and operating on a variety of propellants have been flown without a single inflight failure. An especially important feature of the pintle injector engine is its apparent inherent combustion stability. In over thirty years of development, testing, and production, TRW has never experienced combustion instability in any of its pintle injector engine designs. This has been true of engines operating over a range of thrust from 5 to 250,000 lbs. on earth-storable hypergolic propellants and a large number of smaller engines operating on a variety of propellants (21 combinations) in long duration-firing, pulsing (down to 2 msec), and deep throttling (as much as 19:1) modes. Operating chamber pressures have ranged from 10 to 3,500 psia. This record is particularly impressive given that typical TRW design practice does not consider combustion instability as an issue and no pintle engine has ever employed stability-enhancing features, such as baffles or acoustically resonant chambers. In spite of this, TRW engines have operated stably in regimes not possible with other types of injectors. Various physical explanations and combustion process models for this favorable stability characteristic have been postulated. However, a definitive study that unequivocally establishes the important stabilizing mechanisms still remains to be conducted.

Stoddard, Frank J.↗

NASA NTRS1994

Computational study of generic hypersonic vehicle flow fields

The geometric data of the generic hypersonic vehicle configuration included body definitions and preliminary grids for the forebody (nose cone excluded), midsection (propulsion system excluded), and afterbody sections. This data was to be augmented by the nose section geometry (blunt conical section mated with the noncircular cross section of the forebody initial plane) along with a grid and a detailed supersonic combustion ramjet (scramjet) geometry (inlet and combustor) which should be merged with the nozzle portion of the afterbody geometry. The solutions were to be obtained by using a Navier-Stokes (NS) code such as TUFF for the nose portion, a parabolized Navier-Stokes (PNS) solver such as the UPS and STUFF codes for the forebody, a NS solver with finite rate hydrogen-air chemistry capability such as TUFF and SPARK for the scramjet and a suitable solver (NS or PNS) for the afterbody and external nozzle flows. The numerical simulation of the hypersonic propulsion system for the generic hypersonic vehicle is the major focus of this entire work. Supersonic combustion ramjet is such a propulsion system, hence the main thrust of the present task has been to establish a solution procedure for the scramjet flow. The scramjet flow is compressible, turbulent, and reacting. The fuel used is hydrogen and the combustion process proceeds at a finite rate. As a result, the solution procedure must be capable of addressing such flows.

Narayan, Johnny R.↗

NASA NTRS1993

Laser-saturated fluorescence measurements in laminar sooting diffusion flames

The hydroxyl radical is known to be one of the most important intermediate species in the combustion processes. The hydroxyl radical has also been considered a dominant oxidizer of soot particles in flames. In this investigation the hydroxyl concentration profiles in sooting diffusion flames were measured by the laser-saturated fluorescence (LSF) method. The temperature distributions in the flames were measured by the two-line LSF technique and by thermocouple. In the sooting region the OH fluorescence was too weak to make accurate temperature measurements. The hydroxyl fluorescence profiles for all four flames presented herein show that the OH fluorescence intensities peaked near the flame front. The OH fluorescence intensity dropped sharply toward the dark region of the flame and continued declining to the sooting region. The OH fluorescence profiles also indicate that the OH fluorescence decreased with increasing height in the flames for all flames investigated. Varying the oxidizer composition resulted in a corresponding variation in the maximum OH concentration and the flame temperature. Furthermore, it appears that the maximum OH concentration for each flame increased with increasing flame temperature.

Wey, Changlie↗

NASA NTRS1994

Multi-dimensional combustor flowfield analyses in gas-gas rocket engine

The objectives of the present research are to improve design capabilities for low thrust rocket engines through understanding of the detailed mixing and combustions processes. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an on-going experimental program at NASA LeRC. Detailed computational modeling requires the application of the full three-dimensional Navier Stokes equations, coupled with species diffusion equations. The numerical procedure is performed on both time-marching and time-accurate algorithms and using an LU approximate factorization in time, flux split upwinding differencing in space. The emphasis in this paper is focused on using numerical analysis to understand detailed combustor flowfields, including the shear layer dynamics created between fuel film cooling and the core gas in the vicinity on the nearby combustor wall; the integrity and effectiveness of the coolant film; three-dimensional fuel jets injection/mixing/combustion characteristics; and their impacts on global engine performance.

Tsuei, Hsin-Hua↗

NASA NTRS1995

NASA Lewis Research Center's Preheated Combustor and Materials Test Facility

The Preheated Combustor and Materials Test Facility (PCMTF) in the Engine Research Building (ERB) at the NASA Lewis Research Center is one of two unique combustor facilities that provide a nonvitiated air supply to two test stands, where the air can be used for research combustor testing and high-temperature materials testing. Stand A is used as a research combustor stand, whereas stand B is used for cyclic and survivability tests of aerospace materials at high temperatures. Both stands can accommodate in-house and private industry research programs. The PCMTF is capable of providing up to 30 lb/s (pps) of nonvitiated, 450 psig combustion air at temperatures ranging from 850 to 1150 g F. A 5000 gal tank located outdoors adjacent to the test facility can provide jet fuel at a pressure of 900 psig and a flow rate of 11 gal/min (gpm). Gaseous hydrogen from a 70,000 cu ft (CF) tuber is also available as a fuel. Approximately 500 gpm of cooling water cools the research hardware and exhaust gases. Such cooling is necessary because the air stream reaches temperatures as high as 3000 deg F. The PCMTF provides industry and Government with a facility for studying the combustion process and for obtaining valuable test information on advanced materials. This report describes the facility's support systems and unique capabilities.

Nemets, Steve A.↗

NASA NTRS1995

Dynamic response of a pulsed Burke-Schumann diffusion flame

Turbulent flames are often envisioned as an ensemble of random vortices interacting with the combustion process. A better understanding of the vortex-flame interactions therefore would be useful in improving the modeling of turbulent diffusion flames. Substantial simplification may be made by investigating controlled interactions in a laminar flame, as opposed to random interactions in a turbulent flame. The general goals of the research project are to improve our understanding of (1) the influence of buoyancy on co-flow diffusion flames and (2) the effects of buoyancy on vortex-flame interactions in co-flow diffusion flames. As a first step toward objective (2), we conducted a joint experimental and numerical investigation of the vortex-flame interaction. Vortices were produced by mechanically pulsing the fuel flow at a low frequency, e.g., 10 Hz. Experiments were conducted using a nonflickering Burke-Schumann flame in both microgravity (mu-g) and normal gravity (1g) as a means of varying the buoyant force without modification of the pressure (i.e., density). The effects of buoyant convection may then be determined by a comparison of the mu-g and 1g results. The mu-g results may also reveal the important mechanisms which are masked or overwhelmed by buoyant convection in 1g. A numerical investigation was conducted using a validated, time-accurate numerical code to study the underlying physics during the flame interaction and to assist the interpretation of the experimental results.

Sheu, Jyh-Cherng↗

NASA NTRS1995

Structure Of Flame Balls At Low Lewis-number (SOFBALL)

The work has encompassed several topics related to the experimental and theoretical study of combustion limits in premixed flames at microgravity. These topics include (1) flame structure and stability at low Lewis number (which is the basis for the SOFBALL space flight experiment), (2) flame propagation and extinction in cylindrical tubes, and (3) experimental simulation of combustion processes using autocatalytic chemical reactions. Progress on each of these topics is outlined.

Ronney, Paul D.↗

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 NTRS1996

NASA flow fields analysis

The objectives of the present research are to improve design capabilities for low thrust rocket engines through understanding the detailed mixing and combustion processes in a representative combustor. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an on-going experimental program at NASA LERC. Detailed computational modeling involves the solution of both the two- and three-dimensional Navier Stokes equations, coupled with chemical reactions and the species diffusion equations. Computations of interest include both steady state and time-accurate flowfields and are obtained by means of LU approximate factorization in time and flux split upwinding differencing in space. The emphasis in the research is focused on using numerical analysis to understand detailed combustor flowfields, including the shear layer dynamics created between fuel film cooling and the core gas in the vicinity on the nearby combustor wall; the integrity and effectiveness of the coolant film; and three-dimensional fuel and oxidizer jet injection/mixing/combustion characteristics in the primary combustor along with their joint impacts on global engine performance.

Merkle, Charles L.↗

NASA NTRS1997

A Dual-Line Detection Rayleigh Scattering Diagnostic Technique for the Combustion of Hydrocarbon Fuels and Filtered UV Rayleigh Scattering for Gas Velocity Measurements

Non-intrusive techniques for the dynamic measurement of gas flow properties such as density, temperature and velocity, are needed in the research leading to the development of new generation high-speed aircraft. Accurate velocity, temperature and density data obtained in ground testing and in-flight measurements can help understand the flow physics leading to transition and turbulence in supersonic, high-altitude flight. Such non-intrusive measurement techniques can also be used to study combustion processes of hydrocarbon fuels in aircraft engines. Reliable, time and space resolved temperature measurements in various combustor configurations can lead to a better understanding of high temperature chemical reaction dynamics thus leading to improved modeling and better prediction of such flows. In view of this, a research program was initiated at Polytechnic University's Aerodynamics Laboratory with support from NASA Lewis Research Center through grants NAG3-1301 and NAG3-1690. The overall objective of this program has been to develop laser-based, non-contact, space- and time-resolved temperature and velocity measurement techniques. In the initial phase of the program a ND:YAG laser-based dual-line Rayleigh scattering technique was developed and tested for the accurate measurement of gas temperature in the presence of background laser glare. Effort was next directed towards the development of a filtered, spectrally-resolved Rayleigh/Mie scattering technique with the objective of developing an interferometric method for time-frozen velocity measurements in high-speed flows utilizing the uv line of an ND:YAG laser and an appropriate molecular absorption filter. This effort included both a search for an appropriate filter material for the 266 nm laser line and the development and testing of several image processing techniques for the fast processing of Fabry-Perot images for velocity and temperature information. Finally, work was also carried out for the development of a new laser-based strain-rate and vorticity technique for the time-resolved measurement of vorticity and strain-rates in turbulent flows.

Otugen, M. Volkan↗

NASA NTRS1993

Analysis of Radiation-natural Convection Interactions in 1-g and low-g Environments using the Discrete Exchange Factor Method

A new numerical method is presented for the analysis of combined natural convection and radiation heat transfer with applications in many engineering situations such as materials processing, combustion and fire research. Because of the recent interest in the low gravity environment of space, attention is devoted to both 1-g and low-g applications. The two-dimensional mathematical model is represented by a set of coupled nonlinear integro-partial differential equations. Radiative exchange is formulated using the Discrete Exchange Factor method (DEF). This method considers point to point exchange and provides accurate results over a wide range of radiation parameters. Numerical results show that radiation significantly influences the flow and heat transfer in both low-g and 1-g applications. In the low-g environment, convection is weak, and radiation can easily become the dominant heat transfer mode. It is also shown that volumetric heating by radiation gives rise to an intricate cell pattern in the top heated enclosure.

Kassemi, M.↗

NASA NTRS1998

Investigation of Thermal High Cycle and Low Cycle Fatigue Mechanisms of Thick Thermal Barrier Coatings

Ceramic thermal barrier coatings have attracted increased attention for diesel engine applications. The advantages of using the ceramic coatings include a potential increase in efficiency and power density and a decrease in maintenance cost. Zirconia-based ceramics are the most important coating materials for such applications because of their low thermal conductivity, relatively high thermal expansivity and excellent mechanical properties. However, durability of thick thermal barrier coatings (TBCS) under severe temperature cycling encountered in engine conditions, remains a major question. The thermal transients associated with the start/stop and no-load/full-load engine cycle, and with the in-cylinder combustion process, generate thermal low cycle fatigue (LCF) and thermal high cycle fatigue (HCF) in the coating system. Therefore, the failure mechanisms of thick TBCs are expected to be quite different from those of thin TBCs under these temperature transients. The coating failure is related not only to thermal expansion mismatch and oxidation of the bond coats and substrates, but also to the steep thermal stress gradients induced in the coating systems. Although it has been reported that stresses generated by thermal transients can initiate surface and interface cracks in a coating system, the mechanisms of the crack propagation and of coating failure under the complex LCF and HCF conditions are still not understood. In this paper, the thermal fatigue behavior of an yttria partially stabilized zirconia coating system under simulated LCF and HCF engine conditions is investigated. The effects of LCF and HCF on surface crack initiation and propagation are also discussed.

Zhu, Dong-Ming↗

NASA NTRS1961

Determination of Local Experimental Heat-Transfer Coefficients on Combustion Side of an Ammonia-Oxygen Rocket

Local experimental heat-transfer coefficients were measured in the chamber and throat of a 2400-pound-thrust ammonia-oxygen rocket engine with a nominal chamber pressure of 600 pounds per square inch absolute. Three injector configurations were used. The rocket engine was run over a range of oxidant-fuel ratio and chamber pressure. The injector that achieved the best performance also produced the highest rates of heat flux at design conditions. The heat-transfer data from the best-performing injector agreed well with the simplified equation developed by Bartz at the throat region. A large spread of data was observed for the chamber. This spread was attributed generally to the variations of combustion processes. The spread was least evident, however, with the best-performing injector.

Liebert, Curt H.↗

NASA NTRS1998

Challenges to Laser-Based Imaging Techniques in Gas Turbine Combustor Systems for Aerospace Applications

Increasingly severe constraints on emissions, noise and fuel efficiency must be met by the next generation of commercial aircraft powerplants. At NASA Lewis Research Center (LeRC) a cooperative research effort with industry is underway to design and test combustors that will meet these requirements. To accomplish these tasks, it is necessary to gain both a detailed understanding of the combustion processes and a precise knowledge of combustor and combustor sub-component performance at close to actual conditions. To that end, researchers at LeRC are engaged in a comprehensive diagnostic investigation of high pressure reacting flowfields that duplicate conditions expected within the actual engine combustors. Unique, optically accessible flame-tubes and sector rig combustors, designed especially for these tests. afford the opportunity to probe these flowfields with the most advanced, laser-based optical diagnostic techniques. However, these same techniques, tested and proven on comparatively simple bench-top gaseous flame burners, encounter numerous restrictions and challenges when applied in these facilities. These include high pressures and temperatures, large flow rates, liquid fuels, remote testing, and carbon or other material deposits on combustor windows. Results are shown that document the success and versatility of these nonintrusive optical diagnostics despite the challenges to their implementation in realistic systems.

Locke, Randy J.↗

NASA NTRS1999

Heterogeneous Combustion of Porous Graphite Particles in Normal and Microgravity

Combustion of solid fuel particles has many important applications, including power generation and space propulsion systems. The current models available for describing the combustion process of these particles, especially porous solid particles, include various simplifying approximations. One of the most limiting approximations is the lumping of the physical properties of the porous fuel with the heterogeneous chemical reaction rate constants. The primary objective of the present work is to develop a rigorous model that could decouple such physical and chemical effects from the global heterogeneous reaction rates. For the purpose of validating this model, experiments with porous graphite particles of varying sizes and porosity are being performed at NASA Lewis Research Center. The details of the experimental and theoretical model development effort are described.

Chelliah, Harsha↗