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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1941

The efficiency of combustion turbines with constant-pressure combustion

Of the two fundamental cycles employed in combustion turbines, namely, the explosion (or constant-volume) cycle and the constant-pressure cycle, the latter is considered more in detail and its efficiency is derived with the aid of the cycle diagrams for the several cases with adiabatic and isothermal compression and expansion strokes and with and without utilization of the exhaust heat. Account is also taken of the separate efficiencies of the turbine and compressor and of the pressure losses and heat transfer in the piping. The results show that without the utilization of the exhaust heat the efficiencies for the two cases of adiabatic and isothermal compression is offset by the increase in the heat supplied. It may be seen from the curves that it is necessary to attain separate efficiencies of at least 80 percent in order for useful results to be obtained. There is further shown the considerable effect on the efficiency of pressure losses in piping or heat exchangers.

Piening, Werner↗
NASA NTRS1990

Critical evaluation of Jet-A spray combustion using propane chemical kinetics in gas turbine combustion simulated by KIVA-2

Jet-A spray combustion has been evaluated in gas turbine combustion with the use of propane chemical kinetics as the first approximation for the chemical reactions. Here, the numerical solutions are obtained by using the KIVA-2 computer code. The KIVA-2 code is the most developed of the available multidimensional combustion computer programs for application of the in-cylinder combustion dynamics of internal combustion engines. The released version of KIVA-2 assumes that 12 chemical species are present; the code uses an Arrhenius kinetic-controlled combustion model governed by a four-step global chemical reaction and six equilibrium reactions. Researchers efforts involve the addition of Jet-A thermophysical properties and the implementation of detailed reaction mechanisms for propane oxidation. Three different detailed reaction mechanism models are considered. The first model consists of 131 reactions and 45 species. This is considered as the full mechanism which is developed through the study of chemical kinetics of propane combustion in an enclosed chamber. The full mechanism is evaluated by comparing calculated ignition delay times with available shock tube data. However, these detailed reactions occupy too much computer memory and CPU time for the computation. Therefore, it only serves as a benchmark case by which to evaluate other simplified models. Two possible simplified models were tested in the existing computer code KIVA-2 for the same conditions as used with the full mechanism. One model is obtained through a sensitivity analysis using LSENS, the general kinetics and sensitivity analysis program code of D. A. Bittker and K. Radhakrishnan. This model consists of 45 chemical reactions and 27 species. The other model is based on the work published by C. K. Westbrook and F. L. Dryer.

Nguyen, H. L.↗
NASA NTRS2013

Characterization of Swirl-Venturi Lean Direct Injection Designs for Aviation Gas-Turbine Combustion

Injector geometry, physical mixing, chemical processes, and engine cycle conditions together govern performance, operability and emission characteristics of aviation gas-turbine combustion systems. The present investigation explores swirl-venturi lean direct injection combustor fundamentals, characterizing the influence of key geometric injector parameters on reacting flow physics and emission production trends. In this computational study, a design space exploration was performed using a parameterized swirl-venturi lean direct injector model. From the parametric geometry, 20 three-element lean direct injection combustor sectors were produced and simulated using steady-state, Reynolds-averaged Navier-Stokes reacting computations. Species concentrations were solved directly using a reduced 18-step reaction mechanism for Jet-A. Turbulence closure was obtained using a nonlinear κ-ε model. Results demonstrate sensitivities of the geometric perturbations on axially averaged flow field responses. Output variables include axial velocity, turbulent kinetic energy, static temperature, fuel patternation and minor species mass fractions. Significant trends have been reduced to surrogate model approximations, intended to guide future injector design trade studies and advance aviation gas-turbine combustion research.

Combustion Chamber↗
NASA NTRS1990

Critical evaluation of Jet-A spray combustion using propane chemical kinetics in gas turbine combustion simulated by KIVA-II

Numerical solutions of the Jet-A spray combustion were obtained by means of the KIVA-II computer code after Jet-A properties were added to the 12 chemical species the program had initially contained. Three different reaction mechanism models are considered. The first model consists of 131 reactions and 45 species; it is evaluated by comparing calculated ignition delay times with available shock tube data, and it is used in the evaluation of the other two simplified models. The simplified mechanisms consider 45 reactions and 27 species and 5 reactions and 12 species, respectively. In the prediction of pollutants NOx and CO, the full mechanism of 131 reactions is considered to be more reliable. The numerical results indicate that the variation of the maximum flame temperature is within 20 percent as compared with that of the full mechanism of 131 reactions. The chemical compositions of major components such as C3H8, H2O, O2, CO2, and N2 are of the same order of magnitude. However, the concentrations of pollutants are quite different.

Nguyen, H. L.↗
NASA NTRS1977

Analytical investigation of thermal barrier coatings for advanced power generation combustion turbines

An analytical evaluation was conducted to determine quantitatively the improvement potential in cycle efficiency and cost of electricity made possible by the introduction of thermal barrier coatings to power generation combustion turbine systems. The thermal barrier system, a metallic bond coat and yttria stabilized zirconia outer layer applied by plasma spray techniques, acts as a heat insulator to provide substantial metal temperature reductions below that of the exposed thermal barrier surface. The study results show the thermal barrier to be a potentially attractive means for improving performance and reducing cost of electricity for the simple, recuperated, and combined cycles evaluated.

Amos, D. J.↗
NASA NTRS1987

Vapor deposition and condensate flow on combustion turbine blades - Theoretical model to predict/understand some corrosion rate consequences of molten alkali sulfate deposition in the field or laboratory

An analysis is undertaken of aerodynamically- and centrifugally-driven liquid condensate layers on nonisothermal combustion turbines' stator vanes and rotor blades. Attention is given to the quantitative consequences of one possible mechanism for the initiation of 'hot corrosion' in the underlying blade material through a 'fluxing' of the protective oxide coating by the molten salt of the Newtonian condensate film. Illustrative calculations are presented for the condensate streamline pattern and the distributions of the steady-state condensate layer thickness, together with the corresponding oxide dissolution rate, for a test turbine blade.

Rosner, Daniel E.↗
NASA NTRS1991

Multi-dimensional modelling of gas turbine combustion using a flame sheet model in KIVA II

A flame sheet model for heat release is incorporated into a multi-dimensional fluid mechanical simulation for gas turbine application. The model assumes that the chemical reaction takes place in thin sheets compared to the length scale of mixing, which is valid for the primary combustion zone in a gas turbine combustor. In this paper, the details of the model are described and computational results are discussed.

Cheng, W. K.↗
NASA NTRS1980

Effect of fuel molecular structure on soot formation in gas turbine combustion

The effect of fuel variations at the same hydrogen content on the formation of soot in a gas turbine combustor was studied. Six fuels were burned to a combustor over a matrix of about 50 test conditions with test conditions ranging over 500-1800 kPa (5-18 atm) pressure and 500-1000 K burner inlet temperature; fuel-air ratios were varied from 0.008-0.024. Flame radiation measurements were made through a sapphire window toward the end of the primary zone. The hydrogen content of the six test fuels ranged from 12.80 to 12.88%. Five fuels emphasized hydrocarbon types: (mono, di, and tricyclic), naphthenes (decalin) and partially hydrogenated aromatics (tetralin); the sixth fuel emphasized final boiling point.

Naegeli, D. W.↗
NASA NTRS1994

A study of low emissions gas turbine combustions

Analytical studies have been conducted to determine the best methods of reducing NO(x) emissions from proposed civilian supersonic transports. Modifications to the gas turbine engine combustors and the use of additives were both explored. It was found that combustors which operated very fuel rich or lean appear to be able to meet future emissions standards. Ammonia additives were also effective in removing NO(x), but residual ammonia remained a problem. Studies of a novel combustor which reduces emissions and improves performance were initiated. In a related topic, a study was begun on the feasibility of using supersonic aircraft to obtain atmospheric samples. The effects of shock heating and compression on sample integrity were modeled. Certain chemical species, including NO2, HNO3, and ClONO2 were found to undergo changes to their composition after they passed through shock waves at Mach 2. The use of detonation waves to enhance mixing and combustion in supersonic airflows was also investigated. This research is important to the use of airbreathing propulsion to obtain orbital speeds and access to space. Both steady and pulsed detonation waves were shown to improve engine performance.

Adelman, Henry G.↗
NASA NTRS1979

Fundamentals of Gas Turbine combustion

Combustion problems and research recommendations are discussed in the areas of atomization and vaporization, combustion chemistry, combustion dynamics, and combustion modelling. The recommendations considered of highest priority in these areas are presented.

Gerstein, M.↗
NASA NTRS1993

Integrated Intelligent training and job aiding for combustion turbine engines

This paper describes an ongoing program to augment such an expert system gas turbine startup advisor, known as the EPRI SA VANT System, by including an intelligent training package. It will give a brief background on the SA VANT development and an overview of its evolution into a full-blown Gas Turbine Information System (GTIS) for rapid access of on-line documentation, diagnostics, and training. In particular, the paper will address: (1) the conversion of the knowledge base used by the SA VANT startup advisor so that it can be used for both training and job aiding; and (2) the hypertext-oriented user manuals being incorporated into the system for rapidly accessing on-line documentation at the job site.

Mckeithan, Clifford M., Jr.↗
NASA NTRS2006

Advanced Seal Development for Large Industrial Gas Turbines

Efforts are in progress to develop advanced sealing for large utility industrial gas turbine engines (combustion turbines). Such seals have been under developed for some time for aero gas turbines. It is desired to transition this technology to combustion turbines. Brush seals, film riding face and circumferential seals, and other dynamic and static sealing approaches are being incorporated into gas turbines for aero applications by several engine manufacturers. These seals replace labyrinth or other seals with significantly reduced leakage rates. For utility industrial gas turbines, leakage reduction with advanced sealing can be even greater with the enormous size of the components. Challenges to transitioning technology include: extremely long operating times between overhauls; infrequent but large radial and axial excursions; difficulty in coating larger components; and maintenance, installation, and durability requirements. Advanced sealing is part of the Advanced Turbine Systems (ATS) engine development being done under a cooperative agreement between Westinghouse and the US Department of Energy, Office of Fossil Energy. Seal development focuses on various types of seals in the 501ATS engine both at dynamic and static locations. Each development includes rig testing of candidate designs and subsequent engine validation testing of prototype seals. This presentation gives an update of the ongoing ATS sealing efforts with special emphasis on brush seals.

Chupp, Raymond E.↗
NASA NTRS2003

A Nonlinear Model for Fuel Atomization in Spray Combustion

Most gas turbine combustion codes rely on ad-hoc statistical assumptions regarding the outcome of fuel atomization processes. The modeling effort proposed in this project is aimed at developing a realistic model to produce accurate predictions of fuel atomization parameters. The model involves application of the nonlinear stability theory to analyze the instability and subsequent disintegration of the liquid fuel sheet that is produced by fuel injection nozzles in gas turbine combustors. The fuel sheet is atomized into a multiplicity of small drops of large surface area to volume ratio to enhance the evaporation rate and combustion performance. The proposed model will effect predictions of fuel sheet atomization parameters such as drop size, velocity, and orientation as well as sheet penetration depth, breakup time and thickness. These parameters are essential for combustion simulation codes to perform a controlled and optimized design of gas turbine fuel injectors. Optimizing fuel injection processes is crucial to improving combustion efficiency and hence reducing fuel consumption and pollutants emissions.

Liu, Nan-Suey↗
NASA NTRS1983

Comparative evaluation of gas-turbine engine combustion chamber starting and stalling characteristics for mechanical and air-injection

The effectiveness of propellant atomization with and without air injection in the combustion chamber nozzle of a gas turbine engine is studied. Test show that the startup and burning performance of these combustion chambers can be improved by using an injection during the mechanical propellant atomization process. It is shown that the operational range of combustion chambers can be extended to poorer propellant mixtures by combined air injection mechanical atomization of the propellant.

Dyatlov, I. N.↗
NASA NTRS2011

Apparatus and method for gas turbine active combustion control system

An Active Combustion Control System and method provides for monitoring combustor pressure and modulating fuel to a gas turbine combustor to prevent combustion dynamics and/or flame extinguishments. The system includes an actuator, wherein the actuator periodically injects pulsed fuel into the combustor. The apparatus also includes a sensor connected to the combustion chamber down stream from an inlet, where the sensor generates a signal detecting the pressure oscillations in the combustor. The apparatus controls the actuator in response to the sensor. The apparatus prompts the actuator to periodically inject pulsed fuel into the combustor at a predetermined sympathetic frequency and magnitude, thereby controlling the amplitude of the pressure oscillations in the combustor by modulating the natural oscillations.

Umeh, Chukwueloka↗
NASA NTRSDate not supplied

Pressure-Gain Combustion for Gas Turbines

Pressure Gain Combustion (PGC) is under investigation as a means to improve the thermal efficiency of gas turbines. PGC is a fundamentally unsteady combustion process which, through some means of confinement, raises the total pressure of the working fluid relative to the initial process state. When implemented as a gas turbine combustor it yields a total pressure gain across the device instead of the typical loss seen in a conventional combustor. This pressure gain can significantly improve gas turbine performance. This presentation is a modified and updated version of a 2018 tutorial of the same name. It is tailored to a graduate student audience that is familiar with the PGC concept, but has focused entirely on a specific type of implementation. It will provide a brief review of the concept of PGC and why it's important. The majority of the presentation will describe the various approaches to implementation that are under investigation by the PGC community. Emphasis will be on strategies other than the one that the students are investigating.

Combustion↗
NASA NTRS2018

Pressure-Gain Combustion for Gas Turbines

Pressure Gain Combustion (PGC) is under investigation as a means to improve the thermal efficiency of gas turbines. PGC is a fundamentally unsteady combustion process which, through some means of confinement, raises the total pressure of the working fluid relative to the initial process state. When implemented in a combustor it yields a total pressure gain across the device instead of the typical loss seen in a conventional combustor. This pressure gain can significantly improve gas turbine performance. The tutorial will provide an introduction to the concept of PGC. It will begin with a process description in fundamental thermodynamic terms, and will quantify potential benefits. The majority of the presentation will describe the various approaches to implementation that are under investigation by the PGC.

Paxson, Dan↗