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Sample records for catalytic combustors

  1. Clean catalytic combustor program

    NASA Technical Reports Server (NTRS)

    Ekstedt, E. E.; Lyon, T. F.; Sabla, P. E.; Dodds, W. J.

    1983-01-01

    A combustor program was conducted to evolve and to identify the technology needed for, and to establish the credibility of, using combustors with catalytic reactors in modern high-pressure-ratio aircraft turbine engines. Two selected catalytic combustor concepts were designed, fabricated, and evaluated. The combustors were sized for use in the NASA/General Electric Energy Efficient Engine (E3). One of the combustor designs was a basic parallel-staged double-annular combustor. The second design was also a parallel-staged combustor but employed reverse flow cannular catalytic reactors. Subcomponent tests of fuel injection systems and of catalytic reactors for use in the combustion system were also conducted. Very low-level pollutant emissions and excellent combustor performance were achieved. However, it was obvious from these tests that extensive development of fuel/air preparation systems and considerable advancement in the steady-state operating temperature capability of catalytic reactor materials will be required prior to the consideration of catalytic combustion systems for use in high-pressure-ratio aircraft turbine engines.

  2. Transient catalytic combustor model

    NASA Technical Reports Server (NTRS)

    Tien, J. S.

    1981-01-01

    A quasi-steady gas phase and thermally thin substrate model is used to analyze the transient behavior of catalytic monolith combustors in fuel lean operation. The combustor response delay is due to the substrate thermal inertia. Fast response is favored by thin substrate, short catalytic bed length, high combustor inlet and final temperatures, and small gas channel diameters. The calculated gas and substrate temperature time history at different axial positions provides an understanding of how the catalytic combustor responds to an upstream condition change. The computed results also suggest that the gas residence times in the catalytic bed in the after bed space are correlatable with the nondimensional combustor response time. The model also performs steady state combustion calculations; and the computed steady state emission characteristics show agreement with available experimental data in the range of parameters covered. A catalytic combustor design for automotive gas turbine engine which has reasonably fast response ( 1 second) and can satisfy the emission goals in an acceptable total combustor length is possible.

  3. Steam reformer with catalytic combustor

    DOEpatents

    Voecks, Gerald E.

    1990-03-20

    A steam reformer is disclosed having an annular steam reforming catalyst bed formed by concentric cylinders and having a catalytic combustor located at the center of the innermost cylinder. Fuel is fed into the interior of the catalytic combustor and air is directed at the top of the combustor, creating a catalytic reaction which provides sufficient heat so as to maintain the catalytic reaction in the steam reforming catalyst bed. Alternatively, air is fed into the interior of the catalytic combustor and a fuel mixture is directed at the top. The catalytic combustor provides enhanced radiant and convective heat transfer to the reformer catalyst bed.

  4. Steam reformer with catalytic combustor

    NASA Technical Reports Server (NTRS)

    Voecks, Gerald E. (Inventor)

    1990-01-01

    A steam reformer is disclosed having an annular steam reforming catalyst bed formed by concentric cylinders and having a catalytic combustor located at the center of the innermost cylinder. Fuel is fed into the interior of the catalytic combustor and air is directed at the top of the combustor, creating a catalytic reaction which provides sufficient heat so as to maintain the catalytic reaction in the steam reforming catalyst bed. Alternatively, air is fed into the interior of the catalytic combustor and a fuel mixture is directed at the top. The catalytic combustor provides enhanced radiant and convective heat transfer to the reformer catalyst bed.

  5. Two stage catalytic combustor

    NASA Technical Reports Server (NTRS)

    Alvin, Mary Anne (Inventor); Bachovchin, Dennis (Inventor); Smeltzer, Eugene E. (Inventor); Lippert, Thomas E. (Inventor); Bruck, Gerald J. (Inventor)

    2010-01-01

    A catalytic combustor (14) includes a first catalytic stage (30), a second catalytic stage (40), and an oxidation completion stage (49). The first catalytic stage receives an oxidizer (e.g., 20) and a fuel (26) and discharges a partially oxidized fuel/oxidizer mixture (36). The second catalytic stage receives the partially oxidized fuel/oxidizer mixture and further oxidizes the mixture. The second catalytic stage may include a passageway (47) for conducting a bypass portion (46) of the mixture past a catalyst (e.g., 41) disposed therein. The second catalytic stage may have an outlet temperature elevated sufficiently to complete oxidation of the mixture without using a separate ignition source. The oxidation completion stage is disposed downstream of the second catalytic stage and may recombine the bypass portion with a catalyst exposed portion (48) of the mixture and complete oxidation of the mixture. The second catalytic stage may also include a reticulated foam support (50), a honeycomb support, a tube support or a plate support.

  6. Modeling a Transient Catalytic Combustor

    NASA Technical Reports Server (NTRS)

    Tien, J. S.

    1985-01-01

    Transient model of monolith catalytic combustor presented in report done under NASA/DOE contract. Model assumes quasi-steady gas phase and thermally "thin" solid. In gas-phase treatment, several quasi-global chemical reactions assumed capable of describing CO and unburnt hydrocarbon emissions in fuel-lean operations. In steady-state computation presented, influence of selected operating and design parameters on minimum combustor length studied. When fast transient responses required, both steady and unsteady studies made to achieve meaningful compromise in design.

  7. Diesel engine catalytic combustor system. [aircraft engines

    NASA Technical Reports Server (NTRS)

    Ream, L. W. (Inventor)

    1984-01-01

    A low compression turbocharged diesel engine is provided in which the turbocharger can be operated independently of the engine to power auxiliary equipment. Fuel and air are burned in a catalytic combustor to drive the turbine wheel of turbine section which is initially caused to rotate by starter motor. By opening a flapper value, compressed air from the blower section is directed to catalytic combustor when it is heated and expanded, serving to drive the turbine wheel and also to heat the catalytic element. To start, engine valve is closed, combustion is terminated in catalytic combustor, and the valve is then opened to utilize air from the blower for the air driven motor. When the engine starts, the constituents in its exhaust gas react in the catalytic element and the heat generated provides additional energy for the turbine section.

  8. FEASIBILITY OF BURNING COAL IN CATALYTIC COMBUSTORS

    EPA Science Inventory

    The report gives results of a study, showing that pulverized coal can be burned in a catalytic combustor. Pulverized coal combustion in catalytic beds is markedly different from gaseous fuel combustion. Gas combustion gives uniform bed temperatures and reaction rates over the ent...

  9. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

    W. R. Laster; E. Anoshkina

    2008-01-31

    Under the sponsorship of the U. S. Department of Energy's National Energy Technology Laboratory, Siemens Westinghouse has conducted a three-year program to develop an ultra low NOx, fuel flexible catalytic combustor for gas turbine application in IGCC. The program is defined in three phases: Phase 1 - Implementation Plan, Phase 2 - Validation Testing and Phase 3 - Field Testing. Both Phase 1 and Phase 2 of the program have been completed. In IGCC power plants, the gas turbine must be capable of operating on syngas as a primary fuel and an available back-up fuel such as natural gas. In this program the Rich Catalytic Lean (RCLTM) technology is being developed as an ultra low NOx combustor. In this concept, ultra low NOx is achieved by stabilizing a lean premix combustion process by using a catalytic reactor to oxidize a portion of the fuel, increasing the temperature of fuel/air mixture prior to the main combustion zone. In Phase 1, the feasibility of the catalytic concept for syngas application has been evaluated and the key technology issues identified. In Phase II the technology necessary for the application of the catalytic concept to IGCC fuels was developed through detailed design and subscale testing. Phase III (currently not funded) will consist of full-scale combustor basket testing on natural gas and syngas.

  10. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

    Laster, W. R.; Anoshkina, E.

    2008-01-31

    Under the sponsorship of the U. S. Department of Energy’s National Energy Technology Laboratory, Siemens Westinghouse has conducted a three-year program to develop an ultra low NOx, fuel flexible catalytic combustor for gas turbine application in IGCC. The program is defined in three phases: Phase 1- Implementation Plan, Phase 2- Validation Testing and Phase 3 – Field Testing. Both Phase 1 and Phase 2 of the program have been completed. In IGCC power plants, the gas turbine must be capable of operating on syngas as a primary fuel and an available back-up fuel such as natural gas. In this program the Rich Catalytic Lean (RCLTM) technology is being developed as an ultra low NOx combustor. In this concept, ultra low NOx is achieved by stabilizing a lean premix combustion process by using a catalytic reactor to oxidize a portion of the fuel, increasing the temperature of fuel/air mixture prior to the main combustion zone. In Phase 1, the feasibility of the catalytic concept for syngas application has been evaluated and the key technology issues identified. In Phase II the technology necessary for the application of the catalytic concept to IGCC fuels was developed through detailed design and subscale testing. Phase III (currently not funded) will consist of full-scale combustor basket testing on natural gas and syngas.

  11. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

    W. R. Laster; E. Anoshkina; P. Szedlacsek

    2006-03-31

    Under the sponsorship of the U.S. Department of Energy's National Energy Technology Laboratory, Siemens Westinghouse is conducting a three-year program to develop an ultra low NOx, fuel flexible catalytic combustor for gas turbine application in IGCC. The program is defined in three phases: Phase 1-Implementation Plan, Phase 2-Validation Testing and Phase 3-Field Testing. The Phase 1 program has been completed. Phase II was initiated in October 2004. In IGCC power plants, the gas turbine must be capable of operating on syngas as a primary fuel and an available back-up fuel such as natural gas. In this program the Rich Catalytic Lean (RCL{trademark}) technology is being developed as an ultra low NOx combustor. In this concept, ultra low NOx is achieved by stabilizing a lean premix combustion process by using a catalytic reactor to react part of the fuel, increasing the fuel/air mixture temperature. In Phase 1, the feasibility of the catalytic concept for syngas application has been evaluated and the key technology issues identified. In Phase II the catalytic concept will be demonstrated through subscale testing. Phase III will consist of full-scale combustor basket testing on natural gas and syngas.

  12. Advanced low emissions catalytic combustor program at General Electric

    NASA Technical Reports Server (NTRS)

    Dodds, W. J.

    1979-01-01

    The Advanced Low Emissions Catalytic Combustors Program (ALECC) is being undertaken to evaluate the feasibility of employing catalytic combustion technology in aircraft gas turbine engines as a means to control emission of oxides of nitrogen during subsonic stratospheric cruise operation. The ALECC Program is being conducted in three phases. The first phase, which was completed in November, 1978, consisted of a design study to identify catalytic combustor designs having the greatest potential to meet the emissions and performance goals specified. The primary emissions goal of this program was to obtain cruise NO emissions of less than 1g/kg (compared with levels of 15 to 20 g/x obtained with current designs)/ However, good overall performance and feasibility for engine development were heavily weighted in the evaluation of combustor designs.

  13. Advanced catalytic combustors for low pollutant emissions, phase 1

    NASA Technical Reports Server (NTRS)

    Dodds, W. J.

    1979-01-01

    The feasibility of employing the known attractive and distinguishing features of catalytic combustion technology to reduce nitric oxide emissions from gas turbine engines during subsonic, stratospheric cruise operation was investigated. Six conceptual combustor designs employing catalytic combustion were defined and evaluated for their potential to meet specific emissions and performance goals. Based on these evaluations, two parallel-staged, fixed-geometry designs were identified as the most promising concepts. Additional design studies were conducted to produce detailed preliminary designs of these two combustors. Results indicate that cruise nitric oxide emissions can be reduced by an order of magnitude relative to current technology levels by the use of catalytic combustion. Also, these combustors have the potential for operating over the EPA landing-takeoff cycle and at cruise with a low pressure drop, high combustion efficiency and with a very low overall level of emission pollutants. The use of catalytic combustion, however, requires advanced technology generation in order to obtain the time-temperature catalytic reactor performance and durability required for practical aircraft engine combustors.

  14. Piloted rich-catalytic lean-burn hybrid combustor

    DOEpatents

    Newburry, Donald Maurice

    2002-01-01

    A catalytic combustor assembly which includes, an air source, a fuel delivery means, a catalytic reactor assembly, a mixing chamber, and a means for igniting a fuel/air mixture. The catalytic reactor assembly is in fluid communication with the air source and fuel delivery means and has a fuel/air plenum which is coated with a catalytic material. The fuel/air plenum has cooling air conduits passing therethrough which have an upstream end. The upstream end of the cooling conduits is in fluid communication with the air source but not the fuel delivery means.

  15. Steam Reformer With Fibrous Catalytic Combustor

    NASA Technical Reports Server (NTRS)

    Voecks, Gerald E.

    1987-01-01

    Proposed steam-reforming reactor derives heat from internal combustion on fibrous catalyst. Supplies of fuel and air to combustor controlled to meet demand for heat for steam-reforming reaction. Enables use of less expensive reactor-tube material by limiting temperature to value safe for material yet not so low as to reduce reactor efficiency.

  16. Catalytic combustor for integrated gasification combined cycle power plant

    DOEpatents

    Bachovchin, Dennis M.; Lippert, Thomas E.

    2008-12-16

    A gasification power plant 10 includes a compressor 32 producing a compressed air flow 36, an air separation unit 22 producing a nitrogen flow 44, a gasifier 14 producing a primary fuel flow 28 and a secondary fuel source 60 providing a secondary fuel flow 62 The plant also includes a catalytic combustor 12 combining the nitrogen flow and a combustor portion 38 of the compressed air flow to form a diluted air flow 39 and combining at least one of the primary fuel flow and secondary fuel flow and a mixer portion 78 of the diluted air flow to produce a combustible mixture 80. A catalytic element 64 of the combustor 12 separately receives the combustible mixture and a backside cooling portion 84 of the diluted air flow and allows the mixture and the heated flow to produce a hot combustion gas 46 provided to a turbine 48. When fueled with the secondary fuel flow, nitrogen is not combined with the combustor portion.

  17. Industrial Gas Turbine Engine Catalytic Pilot Combustor-Prototype Testing

    SciTech Connect

    Etemad, Shahrokh; Baird, Benjamin; Alavandi, Sandeep; Pfefferle, William

    2010-04-01

    PCI has developed and demonstrated its Rich Catalytic Lean-burn (RCL®) technology for industrial and utility gas turbines to meet DOE's goals of low single digit emissions. The technology offers stable combustion with extended turndown allowing ultra-low emissions without the cost of exhaust after-treatment and further increasing overall efficiency (avoidance of after-treatment losses). The objective of the work was to develop and demonstrate emission benefits of the catalytic technology to meet strict emissions regulations. Two different applications of the RCL® concept were demonstrated: RCL® catalytic pilot and Full RCL®. The RCL® catalytic pilot was designed to replace the existing pilot (a typical source of high NOx production) in the existing Dry Low NOx (DLN) injector, providing benefit of catalytic combustion while minimizing engine modification. This report discusses the development and single injector and engine testing of a set of T70 injectors equipped with RCL® pilots for natural gas applications. The overall (catalytic pilot plus main injector) program NOx target of less than 5 ppm (corrected to 15% oxygen) was achieved in the T70 engine for the complete set of conditions with engine CO emissions less than 10 ppm. Combustor acoustics were low (at or below 0.1 psi RMS) during testing. The RCL® catalytic pilot supported engine startup and shutdown process without major modification of existing engine controls. During high pressure testing, the catalytic pilot showed no incidence of flashback or autoignition while operating over a wide range of flame temperatures. In applications where lower NOx production is required (i.e. less than 3 ppm), in parallel, a Full RCL® combustor was developed that replaces the existing DLN injector providing potential for maximum emissions reduction. This concept was tested at industrial gas turbine conditions in a Solar Turbines, Incorporated high-pressure (17 atm.) combustion rig and in a modified Solar Turbines

  18. Fuel Flexible, Low Emission Catalytic Combustor for Opportunity Fuel Applications

    SciTech Connect

    Eteman, Shahrokh

    2013-06-30

    Limited fuel resources, increasing energy demand and stringent emission regulations are drivers to evaluate process off-gases or process waste streams as fuels for power generation. Often these process waste streams have low energy content and/or highly reactive components. Operability of low energy content fuels in gas turbines leads to issues such as unstable and incomplete combustion. On the other hand, fuels containing higher-order hydrocarbons lead to flashback and auto-ignition issues. Due to above reasons, these fuels cannot be used directly without modifications or efficiency penalties in gas turbine engines. To enable the use of these wide variety of fuels in gas turbine engines a rich catalytic lean burn (RCL®) combustion system was developed and tested in a subscale high pressure (10 atm.) rig. The RCL® injector provided stability and extended turndown to low Btu fuels due to catalytic pre-reaction. Previous work has shown promise with fuels such as blast furnace gas (BFG) with LHV of 85 Btu/ft3 successfully combusted. This program extends on this work by further modifying the combustor to achieve greater catalytic stability enhancement. Fuels containing low energy content such as weak natural gas with a Lower Heating Value (LHV) of 6.5 MJ/m3 (180 Btu/ft3 to natural gas fuels containing higher hydrocarbon (e.g ethane) with LHV of 37.6 MJ/m3 (1010 Btu/ft3) were demonstrated with improved combustion stability; an extended turndown (defined as the difference between catalytic and non-catalytic lean blow out) of greater than 250oF was achieved with CO and NOx emissions lower than 5 ppm corrected to 15% O2. In addition, for highly reactive fuels the catalytic region preferentially pre-reacted the higher order hydrocarbons with no events of flashback or auto-ignition allowing a stable and safe operation with low NOx and CO emissions.

  19. Emissions and performance of catalysts for gas turbine catalytic combustors

    NASA Technical Reports Server (NTRS)

    Anderson, D. N.

    1977-01-01

    Three noble-metal monolithic catalysts were tested in a 12-centimeter diameter combustion test rig to obtain emissions and performance data at conditions simulating the operation of a catalytic combustor for an automotive gas turbine engine. Tests with one of the catalysts at 800 K inlet mixture temperature 300,000 pa (3 atm) pressure, and a reference velocity (catalyst bed inlet velocity) of 10 m/sec demonstrated greater than 99 percent combustion efficiency for reaction temperatures higher than 1300 K. With a reference velocity of 25 m/sec the reaction temperature required to achieve the same combustion efficiency increased to 1380 K. The exit temperature pattern factors for all three catalysts were below 0.1 when adiabatic reaction temperatures were higher than 1400 K. The highest pressure drop was 4.5 percent at 25 m/sec reference velocity. Nitrogen oxides emissions were less than 0.1 g NO2/kg fuel for all test conditions.

  20. Advanced Low-Emissions Catalytic-Combustor Program, phase 1. [aircraft gas turbine engines

    NASA Technical Reports Server (NTRS)

    Sturgess, G. J.

    1981-01-01

    Six catalytic combustor concepts were defined, analyzed, and evaluated. Major design considerations included low emissions, performance, safety, durability, installations, operations and development. On the basis of these considerations the two most promising concepts were selected. Refined analysis and preliminary design work was conducted on these two concepts. The selected concepts were required to fit within the combustor chamber dimensions of the reference engine. This is achieved by using a dump diffuser discharging into a plenum chamber between the compressor discharge and the turbine inlet, with the combustors overlaying the prediffuser and the rear of the compressor. To enhance maintainability, the outer combustor case for each concept is designed to translate forward for accessibility to the catalytic reactor, liners and high pressure turbine area. The catalytic reactor is self-contained with air-cooled canning on a resilient mounting. Both selected concepts employed integrated engine-starting approaches to raise the catalytic reactor up to operating conditions. Advanced liner schemes are used to minimize required cooling air. The two selected concepts respectively employ fuel-rich initial thermal reaction followed by rapid quench and subsequent fuel-lean catalytic reaction of carbon monoxide, and, fuel-lean thermal reaction of some fuel in a continuously operating pilot combustor with fuel-lean catalytic reaction of remaining fuel in a radially-staged main combustor.

  1. Fuel-Flexible, Low-Emissions Catalytic Combustor for Opportunity Fuels

    SciTech Connect

    2009-11-01

    Precision Combustion, Inc. will develop a unique, fuel-flexible Rich Catalytic Lean-Burn (RCL®) injector with catalytic combustor capable of enabling ultralow-emission, lean premixed combustion of a wide range of gaseous opportunity fuels. This will broaden the range of opportunity fuels that can be utilized to include low- and ultralow-Btu gases, such as digester and blast furnace gases, and fuels containing reactive species, such as refinery, wellhead, and industrial byproduct gases.

  2. Catalytic honeycomb combustor - Steady-state model and comparison with experiment

    NASA Technical Reports Server (NTRS)

    Tien, J. S.

    1980-01-01

    A steady-state lean combustion model for monolithic catalytic combustors is given. The model, consisting of several semi-global chemical reaction steps in the gas-phase and on the surface, is capable of analyzing CO and THC emissions. In the model computation presented, the influence of operating and design parameters on the minimum combustor length is studied. Special attention is given to the effect of after-bed gas-phase reaction space. Comparison with experimental data indicates good agreement in the range of parameters covered.

  3. Gas phase oxidation downstream of a catalytic combustor

    NASA Technical Reports Server (NTRS)

    Tien, J. S.; Anderson, D. N.

    1979-01-01

    Effect of the length available for gas-phase reactions downstream of the catalytic reactor on the emission of CO and unburned hydrocarbons was investigated. A premixed, prevaporized propane/air feed to a 12/cm/diameter catalytic/reactor test section was used. The catalytic reactor was made of four 2.5 cm long monolithic catalyst elements. Four water cooled gas sampling probes were located at positions between 0 and 22 cm downstream of the catalytic reactor. Measurements of unburned hydrocarbon, CO, and CO2 were made. Tests were performed with an inlet air temperature of 800 K, a reference velocity of 10 m/s, pressures of 3 and 600,000 Pa, and fuel air equivalence ratios of 0.14 to 0.24. For very lean mixtures, hydrocarbon emissions were high and CO continued to be formed downstream of the catalytic reactor. At the highest equivalence ratios tested, hydrocarbon levels were much lower and CO was oxidized to CO2 in the gas phase downstream. To achieve acceptable emissions, a downstream region several times longer than the catalytic reactor could be required.

  4. Low and medium heating value coal gas catalytic combustor characterization

    NASA Technical Reports Server (NTRS)

    Schwab, J. A.

    1982-01-01

    Catalytic combustion with both low and medium heating value coal gases obtained from an operating gasifier was demonstrated. A practical operating range for efficient operation was determined, and also to identify potential problem areas were identified for consideration during stationary gas turbine engine design. The test rig consists of fuel injectors, a fuel-air premixing section, a catalytic reactor with thermocouple instrumentation and a single point, water cooled sample probe. The test rig included inlet and outlet transition pieces and was designed for installation into an existing test loop.

  5. Transient Catalytic Combustor Model With Detailed Gas and Surface Chemistry

    NASA Technical Reports Server (NTRS)

    Struk, Peter M.; Dietrich, Daniel L.; Mellish, Benjamin P.; Miller, Fletcher J.; Tien, James S.

    2005-01-01

    In this work, we numerically investigate the transient combustion of a premixed gas mixture in a narrow, perfectly-insulated, catalytic channel which can represent an interior channel of a catalytic monolith. The model assumes a quasi-steady gas-phase and a transient, thermally thin solid phase. The gas phase is one-dimensional, but it does account for heat and mass transfer in a direction perpendicular to the flow via appropriate heat and mass transfer coefficients. The model neglects axial conduction in both the gas and in the solid. The model includes both detailed gas-phase reactions and catalytic surface reactions. The reactants modeled so far include lean mixtures of dry CO and CO/H2 mixtures, with pure oxygen as the oxidizer. The results include transient computations of light-off and system response to inlet condition variations. In some cases, the model predicts two different steady-state solutions depending on whether the channel is initially hot or cold. Additionally, the model suggests that the catalytic ignition of CO/O2 mixtures is extremely sensitive to small variations of inlet equivalence ratios and parts per million levels of H2.

  6. CHARACTERIZATION OF CATALYTIC COMBUSTOR TURBULENCE AND ITS INFLUENCE ON VANE AND ENDWALL HEAT TRANSFER AND ENDWALL FILM COOLING

    SciTech Connect

    Forrest E. Ames

    2002-10-01

    Endwall heat transfer distributions taken in a large-scale low speed linear cascade facility are documented for mock catalytic and dry low NOx (DLN) combustion systems. Inlet turbulence levels range from about 1.0 percent for the mock Catalytic combustor condition to 14 percent for the mock dry low NOx combustor system. Stanton number contours are presented at both turbulence conditions for Reynolds numbers based on true chord length and exit conditions ranging from 500,000 to 2,000,000. Catalytic combustor endwall heat transfer shows the influence of the complex three-dimensional flow field, while the effects of individual vortex systems are less evident for the mock dry low NOx cases. Turbulence scales have been documented for both cases. Inlet boundary layers are relatively thin for the mock catalytic combustor case while inlet flow approximates a channel flow with high turbulence for the mock DLN combustor case. Inlet boundary layer parameters are presented across the inlet passage for the three Reynolds numbers and both the mock catalytic and DLN combustor inlet cases. Both midspan and 95 percent span pressure contours are included. This research provides a well-documented database taken across a range of Reynolds numbers and turbulence conditions for assessment of endwall heat transfer predictive capabilities.

  7. The effect of catalyst length and downstream reactor distance on catalytic combustor performance

    NASA Technical Reports Server (NTRS)

    Anderson, D.

    1980-01-01

    A study was made to determine the effects on catalytic combustor performance which resulted from independently varying the length of a catalytic reactor and the length available for gas-phase reactions downstream of the catalyst. Monolithic combustion catalysts from three manufacturers were tested in a combustion test rig with no. 2 diesel fuel. Catalytic reactor lengths of 2.5 and 5.4 cm, and downstream gas-phase reaction distances of 7.3, 12.4, 17.5, and 22.5 cm were evaluated. Measurements of carbon monoxide, unburned hydrocarbons, nitrogen oxides, and pressure drop were made. The catalytic-reactor pressure drop was less than 1 percent of the upstream total pressure for all test configurations and test conditions. Nitrogen oxides and unburned hydrocarbons emissions were less than 0.25 g NO2/kg fuel and 0.6 g HC/kg fuel, respectively. The minimum operating temperature (defined as the adiabatic combustion temperature required to obtain carbon monoxide emissions below a reference level of 13.6 g CO/kg fuel) ranged from 1230 K to 1500 K for the various conditions and configurations tested. The minimum operating temperature decreased with increasing total (catalytic-reactor-plus-downstream-gas-phase-reactor-zone) residence time but was independent of the relative times spent in each region when the catalytic-reactor residence time was greater than or equal to 1.4 ms.

  8. Experimental evaluation of two premixing-prevaporizing fuel injection concepts for a gas turbine catalytic combustor

    NASA Technical Reports Server (NTRS)

    Tacina, R.

    1976-01-01

    A premixing-prevaporizing fuel system to be used with a catalytic combustor was evaluated for possible application in an automotive gas turbine. Spatial fuel distribution and degree of vaporization were measured using jet A fuel. Two types of air blast injectors were tested, a splash groove injector and a multiple jet cross stream injector. Air swirlers with vane angles of 15 deg and 30 deg were used to improve the spatial fuel distribution in a 12 cm diameter tubular rig. Distribution and vaporization measurements were made 35.5 cm downstream of the injector. The spatial fuel distribution was nearly uniform with the multiple jet contrastream injector and the splash-groove injector with a 30 deg air swirler. The vaporization was nearly 100 percent at an inlet air temperature of 600 K, and at 800 K inlet air temperature fuel oxidation reactions were observed. The total pressure loss was less than 0.5 percent of the total pressure for the multiple jet cross stream injector and the splash groove injector (without air swirler) and less than 1 percent for the splash groove with a 30 deg air swirler.

  9. Emissions and performance of catalysts for gas turbine catalytic combustors. [automobile engines

    NASA Technical Reports Server (NTRS)

    Anderson, D. N.

    1977-01-01

    Three noble-metal monolithic catalysts were tested in a 12-cm-dia. combustion test rig to obtain emissions and performance data at conditions simulating the operation of a catalytic combustor for an automotive gas turbine engine. Tests with one of the catalysts at 800 K inlet mixture temperature, 3 x 10 to the 5th Pa pressure, and a reference velocity (catalyst bed inlet velocity) of 10 m/sec demonstrated greater than 99 percent combustion efficiency for reaction temperatures higher than 1300 K. With a reference velocity of 25 m/sec the reaction temperature required to achieve the same combustion-efficiency increased to 1380 K. The exit temperature pattern factors for all three catalysts were below 0.1 when adiabatic reaction temperatures were higher than 1400 K. The highest pressure drop was 4.5 percent at 25 m/sec reference velocity. Nitrogen oxides emissions were less than 0.1 g NO2/kg fuel for all test conditions.

  10. Experimental evaluation of premixing-prevaporizing fuel injection concepts for a gas turbine catalytic combustor

    NASA Technical Reports Server (NTRS)

    Tacina, R. R.

    1977-01-01

    Experiments were performed to evolve and evaluate a premixing-prevaporizing fuel system to be used with a catalytic combustor for possible application in an automotive gas turbine. Spatial fuel distribution and degree of vaporization were measured using Jet A fuel. Three types of air blast injectors, an air assist nozzle and a simplex pressure atomizer were tested. Air swirlers with vane angles up to 30 deg were used to improve the spatial fuel distribution. The work was done in a 12-cm (4.75-in.) diameter tubular rig. Test conditions were: a pressure of 0.3 and 0.5 MPa (3 and 5 atm), inlet air temperatures up to 800 K (980 F), velocity of 20 m/sec (66 ft/sec) and fuel-air ratios of 0.01 and 0.025. Uniform spatial fuel distributions that were within plus or minus 10 percent of the mean were obtained. Complete vaporization of the fuel was achieved with air blast configurations at inlet air temperatures of 550 K (530 F) and higher. The total pressure loss was less than 0.5 percent for configurations without air swirlers and less than 1 percent for configurations with a 30 deg vane angle air swirler.

  11. A passively-fed methanol steam reformer heated with two-stage bi-fueled catalytic combustor

    NASA Astrophysics Data System (ADS)

    Lo, Kai-Fan; Wong, Shwin-Chung

    2012-09-01

    This paper presents further progress on our simple novel passively-fed methanol steam reformer. The present study focuses on the development of a catalytic combustor workable with both hydrogen and methanol fuels. The aim is to reutilize the exhaust hydrogen from a fuel cell under stable operation but burn methanol during the start-up. On a copper plate, the catalytic combustor in a u-turn channel is integrally machined under a two-turn serpentine-channel reformer. To resolve the highly different fuel reactivities, a suitably diluted catalyst formula demonstrates uniform temperature distributions burning with either liquid methanol or an H2/CO2 mixture simulating the exhaust gas from a fuel cell. In a two-stage process, it first takes 25 min to reach 270 °C by burning methanol. After the fuel is switched to the H2/CO2 mixture, another 20 min is needed to attain an optimal steady state which yields a high methanol conversion of 95% and acceptably low CO fraction of 1.04% at a reaction temperature of 278 °C. The H2 and CO2 concentrations are 75.1% and 23.6%.

  12. Experimental study of an integral catalytic combustor: Heat exchanger for Stirling engines

    NASA Technical Reports Server (NTRS)

    Bulzan, D. L.

    1982-01-01

    The feasibility of using catalytic combustion with heat removal for the Stirling engine to reduce exhaust emissions and also improve heat transfer to the working fluid was studied using spaced parallel plates. An internally air-cooled heat exchanger was placed between two noble metal catalytic plates. A preheated fuel-air mixture passed between the plates and reacted on the surface of the catalyzed plates. Heat was removed from the catalytic surface by radiation and convection to the aircooled heat exchangers to control temperature and minimize thermal nitrogen oxide emissions. Test conditions were inlet combustion air temperatures from 850 to 900 K, inlet velocities of about 10 m/s, equivalence ratios from 0.5 to 0.9, and pressures from 1.3x10 to the 5th power to 2.0x10 to the 5th power Pa. Propane fuel was used for all testing. Combustion efficiencies greater than 99.5 percent were measured. Nitrogen oxide emissions ranged from 1.7 to 3.3 g NO2/kg fuel. The results demonstrate the feasibility of the concept and indicate that further investigation of the concept is warranted.

  13. Evaluation of Durable Metallic Supports for Catalytic Combustors, CRADA Final Report ORNL 00-0570

    SciTech Connect

    Pint, B. A.; Wright, I. G.; Lara-Curzio, E.; McCarty, J.; Barnes, J.

    2003-10-01

    In 2000, a Cooperative Research and Development Agreement (CRADA) was undertaken between the Oak Ridge National Laboratory (ORNL) and Catalytica Energy Systems Incorporated (CESI) to determine the properties of current metallic catalyst supports and examine new candidate alloys for this application. A team was established at ORNL to examine oxidation-limited lifetime of these thin-walled metallic components using standard lifetime models and to measure the mechanical properties of the foils (40-200:m in thickness) which can differ substantially from bulk properties. Oxidation experiments were conducted on foil specimens at 700/-1100/C in laboratory air and in air with 10 vol.% water vapor to better simulate the combustor environment. At the higher test temperatures, time to oxidation-induced (i.e. breakaway oxidation) failure was determined in 1h cycles in order to verify predictions from a standard reservoir-type oxidation lifetime model. Selected specimens were run for >10,000h in 100 or 500h cycles at lower test temperatures in order to determine the oxidation kinetics for the model. The creep properties of selected foils were measured for 4,000-8,000h at operation-relevant stresses and temperatures. None of the new candidate alloys significantly out-performed currently used alloys in laboratory testing, particularly in oxidation lifetime testing. Therefore, engine testing was not performed on any of the new candidate alloys. Both the oxidation- and creep-resistance of FeCrAl alloys was greater than expected and the results of the CRADA allowed CESI to extend life or increase operating temperatures for these lower cost substrate alloys in the next generation of catalyst modules.

  14. Experimental investigation into NO sub x control of a gas-turbine combustor and augmentor tube incorporating a catalytic-reduction system. Master's thesis

    SciTech Connect

    Behrens, C.K.

    1990-03-01

    An initial experimental investigation was conducted to examine the feasibility of NOx emission control using catalytic reduction techniques in the jet engine test cell environment., A modified T-63 gas turbine combustor and an augmentor tube, 21 feet in length and containing a perlite catalyst, were used as a gas generator and catalytic reduction system. Four data runs were made. Three runs were completed without the catalyst installed. Temperature and velocity profile measurements were obtained in order to calculate augmentation ratios for different engine fuel to air ratios. NOx, CO and unburned hydrocarbon concentrations in the exhaust were measured to provide a baseline for further tests. A fourth data run was made with the perlite catalyst installed in the augmentor tube. A 64 percent NOx reduction was observed, however the large pressure drop across the catalytic bed deemed the current configuration impractical. Recommendations for alternative configurations are presented. The results of the investigation have proven that further study is warranted.

  15. Development of a high-temperature durable catalyst for use in catalytic combustors for advanced automotive gas turbine engines

    NASA Technical Reports Server (NTRS)

    Tong, H.; Snow, G. C.; Chu, E. K.; Chang, R. L. S.; Angwin, M. J.; Pessagno, S. L.

    1981-01-01

    Durable catalytic reactors for advanced gas turbine engines were developed. Objectives were: to evaluate furnace aging as a cost effective catalytic reactor screening test, measure reactor degradation as a function of furnace aging, demonstrate 1,000 hours of combustion durability, and define a catalytic reactor system with a high probability of successful integration into an automotive gas turbine engine. Fourteen different catalytic reactor concepts were evaluated, leading to the selection of one for a durability combustion test with diesel fuel for combustion conditions. Eight additional catalytic reactors were evaluated and one of these was successfully combustion tested on propane fuel. This durability reactor used graded cell honeycombs and a combination of noble metal and metal oxide catalysts. The reactor was catalytically active and structurally sound at the end of the durability test.

  16. Development of a high-temperature durable catalyst for use in catalytic combustors for advanced automotive gas turbine engines

    SciTech Connect

    Tong, H; Snow, G C; Chu, E K :; Chang, R L.S.; Angwin, M J; Pessagno, S L

    1981-09-01

    An experimental program was performed to develop durable catalytic reactors for advanced gas turbine engines. This program was performed as part of DOE's Gas Turbine Highway Vehicle Systems Project. Objectives of this program were to evaluate furnace aging as a cost-effective catalytic reactor screening test, measure reactor degradation as a function of furnace aging, demonstrate 1000 h of combustion durability, and define a catalytic reactor system with a high probability of successfful integration into an automotive gas turbine engine. In the first phase of this program, 14 different catalytic reactor concepts were evaluated, leading to the selection of one for a durability combustion test with diesel fuel at 1700 K combustion coditions. The durability reactor, a proprietary UOP noble metal catalyst, failed structurally after about 136 h and the catalyst was essentially inactive after about 226 h. In Phase II, eight additional catalytic reactors were evalated and one of these was sucessfully combustion-tested for 1000 h at 1700 K on propane fuel. This durability reactor used graded-cell honeycombs and a combination of noble metal and metal oxide catalysts. The reactor was catalytically active and structurally sound at the end of the durability test.

  17. Fuel cell system with combustor-heated reformer

    DOEpatents

    Pettit, William Henry

    2000-01-01

    A fuel cell system including a fuel reformer heated by a catalytic combustor fired by anode effluent and/or fuel from a liquid fuel supply providing fuel for the fuel cell. The combustor includes a vaporizer section heated by the combustor exhaust gases for vaporizing the fuel before feeding it into the combustor. Cathode effluent is used as the principle oxidant for the combustor.

  18. Segmented combustor

    NASA Technical Reports Server (NTRS)

    Halila, Ely E. (Inventor)

    1994-01-01

    A combustor liner segment includes a panel having four sidewalls forming a rectangular outer perimeter. A plurality of integral supporting lugs are disposed substantially perpendicularly to the panel and extend from respective ones of the four sidewalls. A plurality of integral bosses are disposed substantially perpendicularly to the panel and extend from respective ones of the four sidewalls, with the bosses being shorter than the lugs. In one embodiment, the lugs extend through supporting holes in an annular frame for mounting the liner segments thereto, with the bosses abutting the frame for maintaining a predetermined spacing therefrom.

  19. CATALYTIC COMBUSTION COMPONENT AND SYSTEM PROTOTYPE DEVELOPMENT

    EPA Science Inventory

    The report gives results of a project to develop the components required for catalytic combustion system operation and evaluation. The systems investigated (firetube boiler, watertube boiler, and gas turbine), when integrated with the catalytic combustor, have potential for both ...

  20. Combustor and combustor screech mitigation methods

    DOEpatents

    Kim, Kwanwoo; Johnson, Thomas Edward; Uhm, Jong Ho; Kraemer, Gilbert Otto

    2014-05-27

    The present application provides for a combustor for use with a gas turbine engine. The combustor may include a cap member and a number of fuel nozzles extending through the cap member. One or more of the fuel nozzles may be provided in a non-flush position with respect to the cap member.

  1. Advanced Low NO Sub X Combustors for Supersonic High-Altitude Aircraft Gas Turbines

    NASA Technical Reports Server (NTRS)

    Roberts, P. B.; White, D. J.; Shekleton, J. R.

    1975-01-01

    A test rig program was conducted with the objective of evaluating and minimizing the exhaust emissions, in particular NO sub x, of three advanced aircraft combustor concepts at a simulated, high altitude cruise condition. The three combustor designs, all members of the lean reaction, premixed family, are the Jet Induced Circulation (JIC) combustor, the Vortex Air Blast (VAB) combustor, and a catalytic combustor. They were rig tested in the form of reverse flow can combustors in the 0.127 m. (5.0 in.) size range. Various configuration modifications were applied to each of the initial JIC and VAB combustor model designs in an effort to reduce the emissions levels. The VAB combustor demonstrated a NO sub x level of 1.1 gm NO2/kg fuel with essentially 100% combustion efficiency at the simulated cruise combustor condition of 50.7 N/sq cm (5 atm), 833 K (1500 R) inlet pressure and temperature respectively and 1778 K (3200 R) outlet temperature on Jet-A1 fuel. Early tests on the catalytic combustor were unsuccessful due to a catalyst deposition problem and were discontinued in favor of the JIC and VAB tests. In addition emissions data were obtained on the JIC and VAB combustors at low combustor inlet pressure and temperatures that indicate the potential performance at engine off-design conditions.

  2. Combustor concepts for aircraft gas turbine low-power emissions reduction

    NASA Technical Reports Server (NTRS)

    Mularz, E. J.; Gleason, C. C.; Dodds, W. J.

    1978-01-01

    Several combustor concepts were designed and tested to demonstrate significant reductions in aircraft engine idle pollutant emissions. Each concept used a different approach for pollutant reductions: the hot wall combustor employs a thermal barrier coating and impingement cooled liners; the recuperative cooling combustor preheats the air before entering the combustion chamber; and the catalytic converter combustor is composed of a conventional primary zone followed by a catalytic bed for pollutant cleanup. The designs are discussed in detail and test results are presented for a range of aircraft engine idle conditions. The results indicate that ultralow levels of unburned hydrocarbons and carbon monoxide emissions can be achieved.

  3. Combustor concepts for aircraft gas turbine low-power emissions reduction

    NASA Technical Reports Server (NTRS)

    Mularz, E. J.; Gleason, C. C.; Dodds, W. J.

    1978-01-01

    Three combustor concepts have been designed and tested to demonstrate significant reductions in aircraft engine idle pollutant emissions. Each concept used a different approach for pollutant reductions: the Hot Wall Combustor employs a thermal barrier coating and impingement cooled liners, the Recuperative Cooling Combustor preheats the air before entering the combustion chamber, and the Catalytic Converter Combustor is composed of a conventional primary zone followed by a catalytic bed for pollutant cleanup. The designs are discussed in detail and test results are presented for a range of aircraft engine idle conditions. The results indicate that ultra-low levels of unburned hydrocarbons and carbon monoxide emissions can be achieved with this technology.

  4. The effect of incomplete fuel-air mixing on the lean blowout limit, lean stability limit and NO(x) emissions in lean premixed gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Shih, W.-P.; Lee, J. G.; Santavicca, D. A.

    1994-01-01

    Gas turbine engines for both land-based and aircraft propulsion applications are facing regulations on NOx emissions which cannot be met with current combustor technology. A number of alternative combustor strategies are being investigated which have the potential capability of achieving ultra-low NOx emissions, including lean premixed combustors, direct injection combustors, rich burn-quick quench-lean burn combustors and catalytic combustors. The research reported in this paper addresses the effect of incomplete fuel-air mixing on the lean limit performance and the NOx emissions characteristics of lean premixed combustors.

  5. Gas turbine combustor transition

    DOEpatents

    Coslow, Billy Joe; Whidden, Graydon Lane

    1999-01-01

    A method of converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit.

  6. Gas turbine combustor transition

    DOEpatents

    Coslow, B.J.; Whidden, G.L.

    1999-05-25

    A method is described for converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit. 7 figs.

  7. Modular combustor dome

    NASA Technical Reports Server (NTRS)

    Glynn, Christopher Charles (Inventor); Halila, Ely Eskenazi (Inventor); Bibler, John David (Inventor); Morris, David Byron (Inventor)

    2001-01-01

    A combustor dome module includes a mixer tube having a hollow heat shield sealingly joined around the outlet end thereof. The modules may then be assembled in an array for defining the combustor dome, with each module being individually removable therefrom.

  8. Gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Burd, Steven W. (Inventor); Cheung, Albert K. (Inventor); Dempsey, Dae K. (Inventor); Hoke, James B. (Inventor); Kramer, Stephen K. (Inventor); Ols, John T. (Inventor); Smith, Reid Dyer Curtis (Inventor); Sowa, William A. (Inventor)

    2011-01-01

    A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber having a length, L. The liner assembly includes a radially inner liner, a radially outer liner that circumscribes the inner liner, and a bulkhead, having a height, H1, which extends between the respective forward ends of the inner liner and the outer liner. The combustor has an exit height, H3, at the respective aft ends of the inner liner and the outer liner interior. The annular combustor has a ratio H1/H3 having a value less than or equal to 1.7. The annular combustor may also have a ration L/H3 having a value less than or equal to 6.0.

  9. Combustor liner cooling system

    DOEpatents

    Lacy, Benjamin Paul; Berkman, Mert Enis

    2013-08-06

    A combustor liner is disclosed. The combustor liner includes an upstream portion, a downstream end portion extending from the upstream portion along a generally longitudinal axis, and a cover layer associated with an inner surface of the downstream end portion. The downstream end portion includes the inner surface and an outer surface, the inner surface defining a plurality of microchannels. The downstream end portion further defines a plurality of passages extending between the inner surface and the outer surface. The plurality of microchannels are fluidly connected to the plurality of passages, and are configured to flow a cooling medium therethrough, cooling the combustor liner.

  10. Dual-Mode Combustor

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J (Inventor); Dippold, Vance F (Inventor)

    2013-01-01

    A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

  11. Direct heating surface combustor

    NASA Technical Reports Server (NTRS)

    Beremand, D. G.; Shire, L. I.; Mroz, T. S. (Inventor)

    1978-01-01

    The combustor utilizes a non-adiabatic flame to provide low-emission combustion for gas turbines. A fuel-air mixture is directed through a porous wall, the other side of which serves as a combustion surface. A radiant heat sink disposed adjacent to and spaced from the combustion surface controls the combustor flame temperature in order to prevent the formation of oxides of nitrogen. A secondary air flow cools the heat sink. Additionally, up to 100% of secondary air flow is mixed with the combustion products at the direct heating surface combustor to dilute such products thereby reducing exit temperature. However, if less than 100% secondary air is mixed to the combustor, the remainder may be added to the combustion products further downstream.

  12. Combustor diffuser interaction program

    NASA Technical Reports Server (NTRS)

    Srinivasan, Ram; Thorp, Daniel

    1986-01-01

    Advances in gas turbine engine performance are achieved by using compressor systems with high stage loading and low part count, which result in high exit Mach numbers. The diffuser and combustor systems in such engines should be optimized to reduce system pressure loss and to maximize the engine thrust-to-weight ratio and minimize length. The state-of-the-art combustor-diffuser systems do not meet these requirements. Detailed understanding of the combustor-diffuser flow field interaction is required for designing advanced gas turbine engines. An experimental study of the combustor-diffuser interaction (CDI) is being conducted to obtain data for the evaluation and improvement of analytical models applicable to a wide variety of diffuser designs. The CDI program consists of four technical phases: Literature Search; Baseline Configuration; Parametric Configurations; and Performance Configurations. Phase 2 of the program is in progress.

  13. Combustor liner durability analysis

    NASA Technical Reports Server (NTRS)

    Moreno, V.

    1981-01-01

    An 18 month combustor liner durability analysis program was conducted to evaluate the use of advanced three dimensional transient heat transfer and nonlinear stress-strain analyses for modeling the cyclic thermomechanical response of a simulated combustor liner specimen. Cyclic life prediction technology for creep/fatigue interaction is evaluated for a variety of state-of-the-art tools for crack initiation and propagation. The sensitivity of the initiation models to a change in the operating conditions is also assessed.

  14. Correlations of catalytic combustor performance parameters

    NASA Technical Reports Server (NTRS)

    Bulzan, D. L.

    1978-01-01

    Correlations for combustion efficiency percentage drop and the minimum required adiabatic reaction temperature necessary to meet emissions goals of 13.6 g CO/kg fuel and 1.64 g HC/kg fuel are presented. Combustion efficiency was found to be a function of the cell density, cell circumference, reactor length, reference velocity, and adiabatic reaction temperature. The percentage pressure drop at an adiabatic reaction temperature of 1450 K was found to be proportional to the reference velocity to the 1.5 power and to the reactor length. It is inversely proportional to the pressure, cell hydraulic diameter, and fractional open area. The minimum required adiabatic reaction temperature was found to increase with reference velocity and decrease with cell circumference, cell density and reactor length. A catalyst factor was introduced into the correlations to account for differences between catalysts. Combustion efficiency, the percentage pressure drop, and the minimum required adiabatic reaction temperature were found to be a function of the catalyst factor. The data was from a 12 cm-diameter test rig with noble metal reactors using propane fuel at an inlet temperature of 800 K.

  15. Vortex combustor for low NOX emissions when burning lean premixed high hydrogen content fuel

    DOEpatents

    Steele, Robert C; Edmonds, Ryan G; Williams, Joseph T; Baldwin, Stephen P

    2012-11-20

    A trapped vortex combustor. The trapped vortex combustor is configured for receiving a lean premixed gaseous fuel and oxidant stream, where the fuel includes hydrogen gas. The trapped vortex combustor is configured to receive the lean premixed fuel and oxidant stream at a velocity which significantly exceeds combustion flame speed in a selected lean premixed fuel and oxidant mixture. The combustor is configured to operate at relatively high bulk fluid velocities while maintaining stable combustion, and low NOx emissions. The combustor is useful in gas turbines in a process of burning synfuels, as it offers the opportunity to avoid use of diluent gas to reduce combustion temperatures. The combustor also offers the possibility of avoiding the use of selected catalytic reaction units for removal of oxides of nitrogen from combustion gases exiting a gas turbine.

  16. Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel

    DOEpatents

    Steele, Robert C.; Edmonds, Ryan G.; Williams, Joseph T.; Baldwin, Stephen P.

    2009-10-20

    A trapped vortex combustor. The trapped vortex combustor is configured for receiving a lean premixed gaseous fuel and oxidant stream, where the fuel includes hydrogen gas. The trapped vortex combustor is configured to receive the lean premixed fuel and oxidant stream at a velocity which significantly exceeds combustion flame speed in a selected lean premixed fuel and oxidant mixture. The combustor is configured to operate at relatively high bulk fluid velocities while maintaining stable combustion, and low NOx emissions. The combustor is useful in gas turbines in a process of burning synfuels, as it offers the opportunity to avoid use of diluent gas to reduce combustion temperatures. The combustor also offers the possibility of avoiding the use of selected catalytic reaction units for removal of oxides of nitrogen from combustion gases exiting a gas turbine.

  17. Gas turbine topping combustor

    DOEpatents

    Beer, Janos; Dowdy, Thomas E.; Bachovchin, Dennis M.

    1997-01-01

    A combustor for burning a mixture of fuel and air in a rich combustion zone, in which the fuel bound nitrogen in converted to molecular nitrogen. The fuel rich combustion is followed by lean combustion. The products of combustion from the lean combustion are rapidly quenched so as to convert the fuel bound nitrogen to molecular nitrogen without forming NOx. The combustor has an air radial swirler that directs the air radially inward while swirling it in the circumferential direction and a radial fuel swirler that directs the fuel radially outward while swirling it in the same circumferential direction, thereby promoting vigorous mixing of the fuel and air. The air inlet has a variable flow area that is responsive to variations in the heating value of the fuel, which may be a coal-derived fuel gas. A diverging passage in the combustor in front of a bluff body causes the fuel/air mixture to recirculate with the rich combustion zone.

  18. Combustor and method for purging a combustor

    DOEpatents

    Berry, Jonathan Dwight; Hughes, Michael John

    2015-06-09

    A combustor includes an end cap. The end cap includes a first surface and a second surface downstream from the first surface, a shroud that circumferentially surrounds at least a portion of the first and second surfaces, a plate that extends radially within the shroud, a plurality of tubes that extend through the plate and the first and second surfaces, and a first purge port that extends through one or more of the plurality of tubes, wherein the purge port is axially aligned with the plate.

  19. Low NO/x/ and fuel flexible gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Lew, H. G.; Decorso, S. M.; Vermes, G.; Carl, D.; Havener, W. J.; Schwab, J.; Notardonato, J.

    1981-01-01

    The feasibility of various low NO(x) emission gas turbine combustor configurations was evaluated. The configurations selected for fabrication and testing at full pressure and temperature involved rich-lean staged combustion utilizing diffusion flames, rich-lean prevaporized/premix flames, and staged catalytic combustion. The test rig consisted of a rich burner module, a quench module, and a lean combustion module. Test results are obtained for the combustor while burning petroleum distillate fuel, a coal derived liquid, and a petroleum residual fuel. The results indicate that rich-lean diffusion flames with low fuel-bound nitrogen conversion are achievable with very high combustion efficiencies.

  20. Combustor burner vanelets

    DOEpatents

    Lacy, Benjamin; Varatharajan, Balachandar; Kraemer, Gilbert Otto; Yilmaz, Ertan; Zuo, Baifang

    2012-02-14

    The present application provides a burner for use with a combustor of a gas turbine engine. The burner may include a center hub, a shroud, a pair of fuel vanes extending from the center hub to the shroud, and a vanelet extending from the center hub and/or the shroud and positioned between the pair of fuel vanes.

  1. Fluidized bed combustor modeling

    NASA Technical Reports Server (NTRS)

    Horio, M.; Rengarajan, P.; Krishnan, R.; Wen, C. Y.

    1977-01-01

    A general mathematical model for the prediction of performance of a fluidized bed coal combustor (FBC) is developed. The basic elements of the model consist of: (1) hydrodynamics of gas and solids in the combustor; (2) description of gas and solids contacting pattern; (3) kinetics of combustion; and (4) absorption of SO2 by limestone in the bed. The model is capable of calculating the combustion efficiency, axial bed temperature profile, carbon hold-up in the bed, oxygen and SO2 concentrations in the bubble and emulsion phases, sulfur retention efficiency and particulate carry over by elutriation. The effects of bed geometry, excess air, location of heat transfer coils in the bed, calcium to sulfur ratio in the feeds, etc. are examined. The calculated results are compared with experimental data. Agreement between the calculated results and the observed data are satisfactory in most cases. Recommendations to enhance the accuracy of prediction of the model are suggested.

  2. Ceramic combustor mounting

    DOEpatents

    Hoffman, Melvin G.; Janneck, Frank W.

    1982-01-01

    A combustor for a gas turbine engine includes a metal engine block including a wall portion defining a housing for a combustor having ceramic liner components. A ceramic outlet duct is supported by a compliant seal on the metal block and a reaction chamber liner is stacked thereon and partly closed at one end by a ceramic bypass swirl plate which is spring loaded by a plurality of circumferentially spaced, spring loaded guide rods and wherein each of the guide rods has one end thereof directed exteriorly of a metal cover plate on the engine block to react against externally located biasing springs cooled by ambient air and wherein the rod spring support arrangement maintains the stacked ceramic components together so that a normal force is maintained on the seal between the outlet duct and the engine block under all operating conditions. The support arrangement also is operative to accommodate a substantial difference in thermal expansion between the ceramic liner components of the combustor and the metal material of the engine block.

  3. Gas turbine topping combustor

    DOEpatents

    Beer, J.; Dowdy, T.E.; Bachovchin, D.M.

    1997-06-10

    A combustor is described for burning a mixture of fuel and air in a rich combustion zone, in which the fuel bound nitrogen in converted to molecular nitrogen. The fuel rich combustion is followed by lean combustion. The products of combustion from the lean combustion are rapidly quenched so as to convert the fuel bound nitrogen to molecular nitrogen without forming NOx. The combustor has an air radial swirler that directs the air radially inward while swirling it in the circumferential direction and a radial fuel swirler that directs the fuel radially outward while swirling it in the same circumferential direction, thereby promoting vigorous mixing of the fuel and air. The air inlet has a variable flow area that is responsive to variations in the heating value of the fuel, which may be a coal-derived fuel gas. A diverging passage in the combustor in front of a bluff body causes the fuel/air mixture to recirculate with the rich combustion zone. 14 figs.

  4. Combustor technology for future aircraft

    NASA Technical Reports Server (NTRS)

    Tacina, Robert R.

    1990-01-01

    The continuing improvement of aircraft gas turbine engine operating efficiencies involves increases in overall engine pressure ratio increases that will result in combustor inlet pressure and temperature increases, greater combustion temperature rises, and higher combustor exit temperatures. These conditions entail the development of fuel injectors generating uniform circumferential and radial temperature patterns, as well as combustor liner configurations and materials capable of withstanding increased thermal radiation even as the amount of cooling air is reduced. Low NO(x)-emitting combustor concepts are required which will employ staged combustion. The development status of component technologies answering these requirements are presently evaluated.

  5. HYPULSE combustor analysis

    NASA Technical Reports Server (NTRS)

    Rizkalla, O. F.

    1993-01-01

    The analysis of selected data from tests of unit fuel injectors in a generic scramjet combustor model is presented. The tests were conducted in the NASA HYPULSE expansion tube at conditions typical of flight at Mach 13.5 and 17. The analysis used a three-stream tube method, with finite-rate chemistry, in which the fuel, test gas, and mixing/combustive streams were treated independently but with the same static pressure. Performance of three candidate fuel injectors is examined based on deduced mixing and combustion efficiencies.

  6. Staged cascade fluidized bed combustor

    DOEpatents

    Cannon, Joseph N.; De Lucia, David E.; Jackson, William M.; Porter, James H.

    1984-01-01

    A fluid bed combustor comprising a plurality of fluidized bed stages interconnected by downcomers providing controlled solids transfer from stage to stage. Each stage is formed from a number of heat transfer tubes carried by a multiapertured web which passes fluidizing air to upper stages. The combustor cross section is tapered inwardly from the middle towards the top and bottom ends. Sorbent materials, as well as non-volatile solid fuels, are added to the top stages of the combustor, and volatile solid fuels are added at an intermediate stage.

  7. Segmented annular combustor

    DOEpatents

    Reider, Samuel B.

    1979-01-01

    An industrial gas turbine engine includes an inclined annular combustor made up of a plurality of support segments each including inner and outer walls of trapezoidally configured planar configuration extents and including side flanges thereon interconnected by means of air cooled connector bolt assemblies to form a continuous annular combustion chamber therebetween and wherein an air fuel mixing chamber is formed at one end of the support segments including means for directing and mixing fuel within a plenum and a perforated header plate for directing streams of air and fuel mixture into the combustion chamber; each of the outer and inner walls of each of the support segments having a ribbed lattice with tracks slidably supporting porous laminated replaceable panels and including pores therein for distributing combustion air into the combustion chamber while cooling the inner surface of each of the panels by transpiration cooling thereof.

  8. Low NOx heavy fuel combustor concept program. Phase 1: Combustion technology generation

    NASA Technical Reports Server (NTRS)

    Lew, H. G.; Carl, D. R.; Vermes, G.; Dezubay, E. A.; Schwab, J. A.; Prothroe, D.

    1981-01-01

    The viability of low emission nitrogen oxide (NOx) gas turbine combustors for industrial and utility application. Thirteen different concepts were evolved and most were tested. Acceptable performance was demonstrated for four of the combustors using ERBS fuel and ultralow NOx emissions were obtained for lean catalytic combustion. Residual oil and coal derived liquids containing fuel bound nitrogen (FBN) were also used at test fuels, and it was shown that staged rich/lean combustion was effective in minimizing the conversion of FBN to NOx. The rich/lean concept was tested with both modular and integral combustors. While the ceramic lined modular configuration produced the best results, the advantages of the all metal integral burners make them candidates for future development. An example of scaling the laboratory sized combustor to a 100 MW size engine is included in the report as are recommendations for future work.

  9. Experimental clean combustor program, phase 1

    NASA Technical Reports Server (NTRS)

    Bahr, D. W.; Gleason, C. C.

    1975-01-01

    Full annular versions of advanced combustor designs, sized to fit within the CF6-50 engine, were defined, manufactured, and tested at high pressure conditions. Configurations were screened, and significant reductions in CO, HC, and NOx emissions levels were achieved with two of these advanced combustor design concepts. Emissions and performance data at a typical AST cruise condition were also obtained along with combustor noise data as a part of an addendum to the basic program. The two promising combustor design approaches evolved in these efforts were the Double Annular Combustor and the Radial/Axial Combustor. With versions of these two basic combustor designs, CO and HC emissions levels at or near the target levels were obtained. Although the low target NOx emissions level was not obtained with these two advanced combustor designs, significant reductions were relative to the NOx levels of current technology combustors. Smoke emission levels below the target value were obtained.

  10. Experimental clean combustor program, phase 2

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Peduzzi, A.; Vitti, G. E.

    1976-01-01

    Combustor pollution reduction technology for commercial CTOL engines was generated and this technology was demonstrated in a full-scale JT9D engine in 1976. Component rig refinement of the two best combustor concepts were tested. These concepts are the vorbix combustor, and a hybrid combustor which combines the pilot zone of the staged premix combustor and the main zone of the swirl-can combustor. Both concepts significantly reduced all pollutant emissions relative to the JT9D-7 engine combustor. However, neither concept met all program goals. The hybrid combustor met pollution goals for unburned hydrocarbons and carbon monoxide but did not achieve the oxides of nitrogen goal. This combustor had significant performance deficiencies. The Vorbix combustor met goals for unburned hydrocarbons and oxides of nitrogen but did not achieve the carbon monoxide goal. Performance of the vorbix combustor approached the engine requirements. On the basis of these results, the vorbix combustor was selected for the engine demonstration program. A control study was conducted to establish fuel control requirements imposed by the low-emission combustor concepts and to identify conceptual control system designs. Concurrent efforts were also completed on two addendums: an alternate fuels addendum and a combustion noise addendum.

  11. Combustor flame flashback

    NASA Technical Reports Server (NTRS)

    Proctor, M. P.; Tien, J. S.

    1985-01-01

    A stainless steel, two-dimensional (rectangular), center-dump, premixed-prevaporized combustor with quartz window sidewalls for visual access was designed, built, and used to study flashback. A parametric study revealed that the flashback equivalence ratio decreased slightly as the inlet air temperature increased. It also indicated that the average premixer velocity and premixer wall temperature were not governing parameters of flashback. The steady-state velocity balance concept as the flashback mechanism was not supported. From visual observation several stages of burning were identified. High speed photography verified upstream flame propagation with the leading edge of the flame front near the premixer wall. Combustion instabilities (spontaneous pressure oscillations) were discovered during combustion at the dump plane and during flashback. The pressure oscillation frequency ranged from 40 to 80 Hz. The peak-to-peak amplitude (up to 1.4 psi) increased as the fuel/air equivalence ratio was increased attaining a maximum value just before flashback. The amplitude suddenly decreased when the flame stabilized in the premixer. The pressure oscillations were large enough to cause a local flow reversal. A simple test using ceramic fiber tufts indicated flow reversals existed at the premixer exit during flickering. It is suspected that flashback occurs through the premixer wall boundary layer flow reversal caused by combustion instability. A theoretical analysis of periodic flow in the premixing channel has been made. The theory supports the flow reversal mechanism.

  12. Combustor and method for distributing fuel in the combustor

    DOEpatents

    Uhm, Jong Ho; Ziminsky, Willy Steve; Johnson, Thomas Edward; York, William David

    2016-04-26

    A combustor includes a tube bundle that extends radially across at least a portion of the combustor. The tube bundle includes an upstream surface axially separated from a downstream surface. A plurality of tubes extends from the upstream surface through the downstream surface, and each tube provides fluid communication through the tube bundle. A baffle extends axially inside the tube bundle between adjacent tubes. A method for distributing fuel in a combustor includes flowing a fuel into a fuel plenum defined at least in part by an upstream surface, a downstream surface, a shroud, and a plurality of tubes that extend from the upstream surface to the downstream surface. The method further includes impinging the fuel against a baffle that extends axially inside the fuel plenum between adjacent tubes.

  13. Radial midframe baffle for can-annular combustor arrangement having tangentially oriented combustor cans

    SciTech Connect

    Rodriguez, Jose L.

    2015-09-15

    A can-annular gas turbine engine combustion arrangement (10), including: a combustor can (12) comprising a combustor inlet (38) and a combustor outlet circumferentially and axially offset from the combustor inlet; an outer casing (24) defining a plenum (22) in which the combustor can is disposed; and baffles (70) configured to divide the plenum into radial sectors (72) and configured to inhibit circumferential motion of compressed air (16) within the plenum.

  14. Pulse combustor with controllable oscillations

    DOEpatents

    Richards, George A.; Welter, Michael J.; Morris, Gary J.

    1992-01-01

    A pulse combustor having thermally induced pulse combustion in a continuously flowing system is described. The pulse combustor is fitted with at lease one elongated ceramic body which significantly increases the heat transfer area in the combustion chamber of the combustor. The ceramic body or bodies possess sufficient mass and heat capacity to ignite the fuel-air charge once the ceramic body or bodies are heated by conventional spark plug initiated combustion so as to provide repetitive ignition and combustion of sequentially introduced fuel-air charges without the assistance of the spark plug and the rapid quenching of the flame after each ignition in a controlled manner so as to provide a selective control over the oscillation frequency and amplitude. Additional control over the heat transfer in the combustion chamber is provided by employing heat exchange mechanisms for selectively heating or cooling the elongated ceramic body or bodies and/or the walls of the combustion chamber.

  15. Pulse combustor with controllable oscillations

    SciTech Connect

    Richards, G.A.; Morris, G.J.; Welter, M.J.

    1991-12-31

    A pulse combustor having thermally induced pulse combustion in a continuously flowing system is described. The pulse combustor is fitted with at lease one elongated ceramic body which significantly increases the heat transfer area in the combustion chamber of the combustor. The ceramic body or bodies possess sufficient mass and heat capacity to ignite the fuel-air charge once the ceramic body or bodies are heated by conventional spark plug initiated combustion so as to provide repetitive ignition and combustion of sequentially introduced fuel-air charges without the assistance of the spark plug and the rapid quenching of the flame after each ignition in a controlled manner so as to provide a selective control over the oscillation frequency and amplitude. Additional control over the heat transfer in the combustion chamber is provided by employing heat exchange mechanisms for selectively heating or cooling the elongated ceramic body or bodies and/or the walls of the combustion chamber.

  16. Methanol tailgas combustor control method

    DOEpatents

    Hart-Predmore, David J.; Pettit, William H.

    2002-01-01

    A method for controlling the power and temperature and fuel source of a combustor in a fuel cell apparatus to supply heat to a fuel processor where the combustor has dual fuel inlet streams including a first fuel stream, and a second fuel stream of anode effluent from the fuel cell and reformate from the fuel processor. In all operating modes, an enthalpy balance is determined by regulating the amount of the first and/or second fuel streams and the quantity of the first air flow stream to support fuel processor power requirements.

  17. Combustor with fuel preparation chambers

    NASA Technical Reports Server (NTRS)

    Zelina, Joseph (Inventor); Myers, Geoffrey D. (Inventor); Srinivasan, Ram (Inventor); Reynolds, Robert S. (Inventor)

    2001-01-01

    An annular combustor having fuel preparation chambers mounted in the dome of the combustor. The fuel preparation chamber comprises an annular wall extending axially from an inlet to an exit that defines a mixing chamber. Mounted to the inlet are an air swirler and a fuel atomizer. The air swirler provides swirled air to the mixing chamber while the atomizer provides a fuel spray. On the downstream side of the exit, the fuel preparation chamber has an inwardly extending conical wall that compresses the swirling mixture of fuel and air exiting the mixing chamber.

  18. Energy Efficient Engine: Combustor component performance program

    NASA Technical Reports Server (NTRS)

    Dubiel, D. J.

    1986-01-01

    The results of the Combustor Component Performance analysis as developed under the Energy Efficient Engine (EEE) program are presented. This study was conducted to demonstrate the aerothermal and environmental goals established for the EEE program and to identify areas where refinements might be made to meet future combustor requirements. In this study, a full annular combustor test rig was used to establish emission levels and combustor performance for comparison with those indicated by the supporting technology program. In addition, a combustor sector test rig was employed to examine differences in emissions and liner temperatures obtained during the full annular performance and supporting technology tests.

  19. HSCT Sector Combustor Evaluations for Demonstration Engine

    NASA Technical Reports Server (NTRS)

    Greenfield, Stuart; Heberling, Paul; Kastl, John; Matulaitis, John; Huff, Cynthia

    2004-01-01

    In LET Task 10, critical development issues of the HSCT lean-burn low emissions combustor were addressed with a range of engineering tools. Laser diagnostics and CFD analysis were applied to develop a clearer understanding of the fuel-air premixing process and premixed combustion. Subcomponent tests evaluated the emissions and operability performance of the fuel-air premixers. Sector combustor tests evaluated the performance of the integrated combustor system. A 3-cup sector was designed and procured for laser diagnostics studies at NASA Glenn. The results of these efforts supported the earlier selection of the Cyclone Swirler as the pilot stage premixer and the IMFH (Integrated Mixer Flame Holder) tube as the main stage premixer of the LPP combustor. In the combustor system preliminary design subtask, initial efforts to transform the sector combustor design into a practical subscale engine combustor met with significant challenges. Concerns about the durability of a stepped combustor dome and the need for a removable fuel injection system resulted in the invention and refinement of the MRA (Multistage Radial Axial) combustor system in 1994. The MRA combustor was selected for the HSR Phase II LPP subscale combustor testing in the CPC Program.

  20. Modelling of furnaces and combustors

    SciTech Connect

    Kahil, E.E.

    1985-01-01

    This book presents an account of the art of modelling for heat transfer and fluid flows in furnaces and combustors. After describing the different types of furnace flows, the author deals with the conservation equations. The different turbulence modelling assumptions, the more complicated problem of turbulent combustion modelling, and various types of turbulent flames are also described and reviewed, with appropriate models being assigned.

  1. Premixed Prevaporized Combustor Technology Forum

    NASA Technical Reports Server (NTRS)

    1979-01-01

    The Forum was held to present the results of recent and current work intended to provide basic information required for demonstration of lean, premixed prevaporized combustors for aircraft gas turbine engine application. Papers are presented which deal with the following major topics: (1) engine interfaces; (2) fuel-air preparation; (3) autoignition; (4) lean combustion; and (5) concept design studies.

  2. Low NOx heavy fuel combustor concept program, phase 1

    NASA Technical Reports Server (NTRS)

    Cutrone, M. B.

    1981-01-01

    Combustion tests were completed with seven concepts, including three rich/lean concepts, three lean/lean concepts, and one catalytic combustor concept. Testing was conducted with ERBS petroleum distillate, petroleum residual, and SRC-II coal-derived liquid fuels over a range of operating conditions for the 12:1 pressure ratio General Electric MS7001E heavy-duty turbine. Blends of ERBS and SRC-II fuels were used to vary fuel properties over a wide range. In addition, pyridine was added to the ERBS and residual fuels to vary nitrogen level while holding other fuel properties constant. Test results indicate that low levels of NOx and fuel-bound nitrogen conversion can be achieved with the rich/lean combustor concepts for fuels with nitrogen contents up to 1.0% by weight. Multinozzle rich/lean Concept 2 demonstrated dry low Nox emissions within 10-15% of the EPA New Source Performance Standards goals for SRC-II fuel, with yields of approximately 15%, while meeting program goals for combustion efficiency, pressure drop, and exhaust gas temperature profile. Similar, if not superior, potential was demonstrated by Concept 3, which is a promising rich/lean combustor design.

  3. Experimental clean combustor program, phase 2

    NASA Technical Reports Server (NTRS)

    Gleason, C. C.; Rogers, D. W.; Bahr, D. W.

    1976-01-01

    The primary objectives of this three-phase program are to develop technology for the design of advanced combustors with significantly lower pollutant emission levels than those of current combustors, and to demonstrate these pollutant emission reductions in CF6-50C engine tests. The purpose of the Phase 2 Program was to further develop the two most promising concepts identified in the Phase 1 Program, the double annular combustor and the radial/axial staged combustor, and to design a combustor and breadboard fuel splitter control for CF6-50 engine demonstration testing in the Phase 3 Program. Noise measurement and alternate fuels addendums to the basic program were conducted to obtain additional experimental data. Twenty-one full annular and fifty-two sector combustor configurations were evaluated. Both combustor types demonstrated the capability for significantly reducing pollutant emission levels. The most promising results were obtained with the double annular combustor. Rig test results corrected to CF-50C engine conditions produced EPA emission parameters for CO, HC, and NOX of 3.4, 0.4, and 4.5 respectively. These levels represent CO, HC, and NOX reductions of 69, 90, and 42 percent respectively from current combustor emission levels. The combustor also met smoke emission level requirements and development engine performance and installation requirements.

  4. Variable residence time vortex combustor

    DOEpatents

    Melconian, Jerry O.

    1987-01-01

    A variable residence time vortex combustor including a primary combustion chamber for containing a combustion vortex, and a plurality of louvres peripherally disposed about the primary combustion chamber and longitudinally distributed along its primary axis. The louvres are inclined to impel air about the primary combustion chamber to cool its interior surfaces and to impel air inwardly to assist in driving the combustion vortex in a first rotational direction and to feed combustion in the primary combustion chamber. The vortex combustor also includes a second combustion chamber having a secondary zone and a narrowed waist region in the primary combustion chamber interconnecting the output of the primary combustion chamber with the secondary zone for passing only lower density particles and trapping higher density particles in the combustion vortex in the primary combustion chamber for substantial combustion.

  5. Combustor modelling for scramjet engines

    NASA Technical Reports Server (NTRS)

    Drummond, J. P.; Rogers, R. C.; Evans, J. S.

    1979-01-01

    A system of computer programs is being developed to analyse and predict the complex flow fields found in hydrogen-fueled scramjet combustors. Each program is designed to solve the governing equation system for the type of flow present in a particular combustor region. A two-dimensional parabolic program has been found to be valuable in the development and experimental evaluation of turbulence and chemistry models for supersonic flow, and in the development of a program to model supersonic flow downstream of the fuel injection struts by means of solutions to the three-dimensional parabolic Navier-Stokes equations and species equations. A partially elliptic code has been derived to account for local subsonic flow regions, and fully elliptic programs have been developed by the consideration of streamwise diffusion effects for the recirculating flow fields near transverse fuel injectors. The programs are currently being applied to problems of scramjet engine development.

  6. Combustor with multistage internal vortices

    DOEpatents

    Shang, Jer Y.; Harrington, Richard E.

    1989-01-01

    A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard area to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard.

  7. Combustor with multistage internal vortices

    DOEpatents

    Shang, Jer Yu; Harrington, R.E.

    1987-05-01

    A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard areas to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard. 2 figs.

  8. Vertical combustor for particulate refuse

    NASA Astrophysics Data System (ADS)

    Chung, P. M.; Carlson, L.

    1981-03-01

    A one-dimensional model is constructed of a vertical combustor for refuse particle combustion in order to analyze it for waste energy recovery. The three components of the model, fuel particles, inert solid particles and the gaseous mixture are described by momentum, energy, and mass conservation equations, resulting in three different flow velocities and temperatures for the medium. The gaseous component is further divided into six chemical species that evolve in combustion at temperatures below about 1367 K. A detailed description is given of the fuel particle combustion through heating, devolatilization, and combustion of the volatile gas in the boundary layer, return of the flame sheet to the fuel surface, and char combustion. The solutions show the combustor to be viable for U.S. refuse which consists of combustibles that can be volatilized up to 85 to 95% below 1366 K. Char combustion, however, is found to be too slow to be attempted in the combustor, where the fuel residence time is of the order of 2 s.

  9. Alternate-Fueled Combustor-Sector Performance: Part A: Combustor Performance Part B: Combustor Emissions

    NASA Technical Reports Server (NTRS)

    Shouse, D. T.; Neuroth, C.; Henricks, R. C.; Lynch, A.; Frayne, C.; Stutrud, J. S.; Corporan, E.; Hankins, T.

    2010-01-01

    Alternate aviation fuels for military or commercial use are required to satisfy MIL-DTL-83133F(2008) or ASTM D 7566 (2010) standards, respectively, and are classified as drop-in fuel replacements. To satisfy legacy issues, blends to 50% alternate fuel with petroleum fuels are certified individually on the basis of feedstock. Adherence to alternate fuels and fuel blends requires smart fueling systems or advanced fuel-flexible systems, including combustors and engines without significant sacrifice in performance or emissions requirements. This paper provides preliminary performance (Part A) and emissions and particulates (Part B) combustor sector data for synthetic-parafinic-kerosene- (SPK-) type fuel and blends with JP-8+100 relative to JP-8+100 as baseline fueling.

  10. Component Development to Accelerate Commercial Implementation of Ultra-Low Emissions Catalytic Combustion

    SciTech Connect

    McCarty, Jon; Berry, Brian; Lundberg, Kare; Anson, Orris

    2003-03-31

    This final report describes a 2000-2003 program for the development of components and processes to enhance the commercialization of ultra-low emissions catalytic combustion in industrial gas turbines. The range of project tasks includes: development of more durable, lower-cost catalysts and catalytic combustor components; development and design of a catalytic pre-burner and a catalytic pilot burner for gas turbines, and on-site fuel conversion processing for utilization of liquid fuel.

  11. Low NO(x) Combustor Development

    NASA Technical Reports Server (NTRS)

    Kastl, J. A.; Herberling, P. V.; Matulaitis, J. M.

    2005-01-01

    The goal of these efforts was the development of an ultra-low emissions, lean-burn combustor for the High Speed Civil Transport. The HSCT Mach 2.4 FLADE C1 Cycle was selected as the baseline engine cycle. A preliminary compilation of performance requirements for the HSCT combustor system was developed. The emissions goals of the program, baseline engine cycle, and standard combustor performance requirements were considered in developing the compilation of performance requirements. Seven combustor system designs were developed. The development of these system designs was facilitated by the use of spreadsheet-type models which predicted performance of the combustor systems over the entire flight envelope of the HSCT. A chemical kinetic model was developed for an LPP combustor and employed to study NO(x) formation kinetics, and CO burnout. These predictions helped to define the combustor residence time. Five fuel-air mixer concepts were analyzed for use in the combustor system designs. One of the seven system designs, one using the Swirl-Jet and Cyclone Swirler fuel-air mixers, was selected for a preliminary mechanical design study.

  12. Combustor with non-circular head end

    SciTech Connect

    Kim, Won -Wook; McMahan, Kevin Weston

    2015-09-29

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a head end with a non-circular configuration, a number of fuel nozzles positioned about the head end, and a transition piece extending downstream of the head end.

  13. Analytical fuel property effects--small combustors

    NASA Technical Reports Server (NTRS)

    Sutton, R. D.; Troth, D. L.; Miles, G. A.

    1984-01-01

    The consequences of using broad-property fuels in both conventional and advanced state-of-the-art small gas turbine combustors are assessed. Eight combustor concepts were selected for initial screening, of these, four final combustor concepts were chosen for further detailed analysis. These included the dual orifice injector baseline combustor (a current production 250-C30 engine combustor) two baseline airblast injected modifications, short and piloted prechamber combustors, and an advanced airblast injected, variable geometry air staged combustor. Final predictions employed the use of the STAC-I computer code. This quasi 2-D model includes real fuel properties, effects of injector type on atomization, detailed droplet dynamics, and multistep chemical kinetics. In general, fuel property effects on various combustor concepts can be classified as chemical or physical in nature. Predictions indicate that fuel chemistry has a significant effect on flame radiation, liner wall temperature, and smoke emission. Fuel physical properties that govern atomization quality and evaporation rates are predicted to affect ignition and lean-blowout limits, combustion efficiency, unburned hydrocarbon, and carbon monoxide emissions.

  14. HSCT Sector Combustor Hardware Modifications for Improved Combustor Design

    NASA Technical Reports Server (NTRS)

    Greenfield, Stuart C.; Heberling, Paul V.; Moertle, George E.

    2005-01-01

    An alternative to the stepped-dome design for the lean premixed prevaporized (LPP) combustor has been developed. The new design uses the same premixer types as the stepped-dome design: integrated mixer flameholder (IMFH) tubes and a cyclone swirler pilot. The IMFH fuel system has been taken to a new level of development. Although the IMFH fuel system design developed in this Task is not intended to be engine-like hardware, it does have certain characteristics of engine hardware, including separate fuel circuits for each of the fuel stages. The four main stage fuel circuits are integrated into a single system which can be withdrawn from the combustor as a unit. Additionally, two new types of liner cooling have been designed. The resulting lean blowout data was found to correlate well with the Lefebvre parameter. As expected, CO and unburned hydrocarbons emissions were shown to have an approximately linear relationship, even though some scatter was present in the data, and the CO versus flame temperature data showed the typical cupped shape. Finally, the NOx emissions data was shown to agree well with a previously developed correlation based on emissions data from Configuration 3 tests performed at GEAE. The design variations of the cyclone swirler pilot that were investigated in this study did not significantly change the NOx emissions from the baseline design (GEAE Configuration 3) at supersonic cruise conditions.

  15. TRW advanced slagging coal combustor utility demonstration

    SciTech Connect

    Not Available

    1990-01-01

    The TRW Advanced Entrained Coal Combustor Demonstration Project consists of retrofitting Orange and Rockland (O R) Utility Corporation's Lovett Plant Unit No. 3 with four (4) slagging combustors which will allow the gas/oil unit to fire 2.5% sulfur coal. The slagging combustor process will provide NO{sub x} and SO{sub x} emissions that meet NSPS and New York State Environmental Standards. The TRW-Utility Demonstration Unit (UDU) is responsible for the implementation of program policies and overall direction of the project. The following projects will be carried out: process and design development of clean coal technology CCT-1 the development and operation of the entrained coal combustor will enable the boiler to burn low and medium sulfur coal while meeting all the Federal/State emission requirements; demonstrate sulfur dioxide emissions control by pulverized limestone injection into the entrained coal combustor system.

  16. Experimental clean combustor program, phase 3

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Fiorentino, A.; Greene, W.

    1977-01-01

    A two-stage vortex burning and mixing combustor and associated fuel system components were successfully tested at steady state and transient operating conditions. The combustor exceeded the program goals for all three emissions species, with oxides of nitrogen 10 percent below the goal, carbon monoxide 26 percent below the goal, and total unburned hydrocarbons 75 percent below the goal. Relative to the JT9D-7 combustor, the oxides of nitrogen were reduced by 58 percent, carbon monoxide emissions were reduced by 69 percent, and total unburned hydrocarbons were reduced by 9 percent. The combustor efficiency and exit temperature profiles were comparable to those of production combustor. Acceleration and starting characteristics were deficient relative to the production engine.

  17. Combustor for fine particulate coal

    DOEpatents

    Carlson, L.W.

    1988-01-26

    A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover. 4 figs.

  18. Combustor for fine particulate coal

    DOEpatents

    Carlson, Larry W.

    1988-01-01

    A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover.

  19. Combustor for fine particulate coal

    DOEpatents

    Carlson, L.W.

    1988-11-08

    A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover. 4 figs.

  20. The role of surface generated radicals in catalytic combustion

    NASA Technical Reports Server (NTRS)

    Santavicca, D. A.; Stein, Y.; Royce, B. S. H.

    1985-01-01

    Experiments were conducted to better understand the role of catalytic surface reactions in determining the ignition characteristics of practical catalytic combustors. Hydrocarbon concentrations, carbon monoxide and carbon dioxide concentrations, hydroxyl radical concentrations, and gas temperature were measured at the exit of a platinum coated, stacked plate, catalytic combustor during the ignition of lean propane-air mixtures. The substrate temperature profile was also measured during the ignition transient. Ignition was initiated by suddenly turning on the fuel and the time to reach steady state was of the order of 10 minutes. The gas phase reaction, showed no pronounced effect due to the catalytic surface reactions, except the absence of a hydroxyl radical overshoot. It is found that the transient ignition measurements are valuable in understanding the steady state performance characteristics.

  1. Solid Fuel Ramjet Combustor Design

    NASA Astrophysics Data System (ADS)

    Krishnan, S.; George, Philmon

    1998-03-01

    Combustion aspects of solid fuel ramjet (SFRJ) are reviewed. On the point of view of the ability of an SFRJ to operate satisfactorily at all off-design conditions the areas of concern to propulsion system designer are (1) selection of a fuel type, (2) flame holding requirements that limit maximum fuel loading, (3) understanding the fuel regression rate behaviour as a function of flight speed and altitude, (4) diffusion-controlled combustion process and its efficiency enhancement, and (5) inlet/combustor matching. Considering these areas, the following aspects are reviewed from the information available in open literature: (1) different experimental set-up conditions adopted in combustor research, (2) various suitable fuel types, (3) flammability limits, (4) fuel regression rate behaviour, (5) methods of achieving high efficiency in metallized fuel, and (6) various modelling efforts. Detailed discussion is presented on two different types of regression rate mechanism in SFRJ: one that is controlled by the heat transfer processes downstream of the reattachment region and the other by that in the region itself. With a view to demonstrate the use of the information collected through this review, a preliminary design procedure is presented for an SFRJ-assisted gun launched projectile of pseudo-vacuum trajectory.

  2. Wedge edge ceramic combustor tile

    DOEpatents

    Shaffer, James E.; Holsapple, Allan C.

    1997-01-01

    A multipiece combustor has a portion thereof being made of a plurality of ceramic segments. Each of the plurality of ceramic segments have an outer surface and an inner surface. Each of the plurality of ceramic segments have a generally cylindrical configuration and including a plurality of joints. The joints define joint portions, a first portion defining a surface being skewed to the outer surface and the inner surface. The joint portions have a second portion defining a surface being skewed to the outer surface and the inner surface. The joint portions further include a shoulder formed intermediate the first portion and the second portion. The joints provide a sealing interlocking joint between corresponding ones of the plurality of ceramic segments. Thus, the multipiece combustor having the plurality of ceramic segment with the plurality of joints reduces the physical size of the individual components and the degradation of the surface of the ceramic components in a tensile stress zone is generally eliminated reducing the possibility of catastrophic failures.

  3. Wedge edge ceramic combustor tile

    DOEpatents

    Shaffer, J.E.; Holsapple, A.C.

    1997-06-10

    A multipiece combustor has a portion thereof being made of a plurality of ceramic segments. Each of the plurality of ceramic segments have an outer surface and an inner surface. Each of the plurality of ceramic segments have a generally cylindrical configuration and including a plurality of joints. The joints define joint portions, a first portion defining a surface being skewed to the outer surface and the inner surface. The joint portions have a second portion defining a surface being skewed to the outer surface and the inner surface. The joint portions further include a shoulder formed intermediate the first portion and the second portion. The joints provide a sealing interlocking joint between corresponding ones of the plurality of ceramic segments. Thus, the multipiece combustor having the plurality of ceramic segment with the plurality of joints reduces the physical size of the individual components and the degradation of the surface of the ceramic components in a tensile stress zone is generally eliminated reducing the possibility of catastrophic failures. 7 figs.

  4. Low NOx Advanced Vortex Combustor

    SciTech Connect

    Edmonds, R.G.; Williams, J.T.; Steele, R.C.; Straub, D.L.; Casleton, K.H.; Bining, Avtar

    2008-05-01

    A lean-premixed advanced vortex combustor (AVC) has been developed and tested. The natural gas fueled AVC was tested at the U.S. Department of Energy’s National Energy Technology Laboratory in Morgantown, WV. All testing was performed at elevated pressures and inlet temperatures and at lean fuel-air ratios representative of industrial gas turbines. The improved AVC design exhibited simultaneous NOx /CO/unburned hydrocarbon (UHC) emissions of 4/4/0 ppmv (all emissions corrected to 15% O2 dry). The design also achieved less than 3 ppmv NOx with combustion efficiencies in excess of 99.5%. The design demonstrated marked acoustic dynamic stability over a wide range of operating conditions, which potentially makes this approach significantly more attractive than other lean-premixed combustion approaches. In addition, the measured 1.75% pressure drop is significantly lower than conventional gas turbine combustors, which could translate into an overall gas turbine cycle efficiency improvement. The relatively high velocities and low pressure drop achievable with this technology make the AVC approach an attractive alternative for syngas fuel applications.

  5. Low NOx Heavy Fuel Combustor Concept Program

    NASA Technical Reports Server (NTRS)

    Novick, A. S.; Troth, D. L.

    1981-01-01

    The development of the technology required to operate an industrial gas turbine combustion system on minimally processed, heavy petroleum or residual fuels having high levels of fuel-bound nitrogen (FBN) while producing acceptable levels of exhaust emissions is discussed. Three combustor concepts were designed and fabricated. Three fuels were supplied for the combustor test demonstrations: a typical middle distillate fuel, a heavy residual fuel, and a synthetic coal-derived fuel. The primary concept was an air staged, variable-geometry combustor designed to produce low emissions from fuels having high levels of FBN. This combustor used a long residence time, fuel-rich primary combustion zone followed by a quick-quench air mixer to rapidly dilute the fuel rich products for the fuel-lean final burnout of the fuel. This combustor, called the rich quench lean (RQL) combustor, was extensively tested using each fuel over the entire power range of the model 570 K engine. Also, a series of parameteric tests was conducted to determine the combustor's sensitivity to rich-zone equivalence ratio, lean-zone equivalence ratio, rich-zone residence time, and overall system pressure drop. Minimum nitrogen oxide emissions were measured at 50 to 55 ppmv at maximum continuous power for all three fuels. Smoke was less than a 10 SAE smoke number.

  6. Chaos in an imperfectly premixed model combustor.

    PubMed

    Kabiraj, Lipika; Saurabh, Aditya; Karimi, Nader; Sailor, Anna; Mastorakos, Epaminondas; Dowling, Ann P; Paschereit, Christian O

    2015-02-01

    This article reports nonlinear bifurcations observed in a laboratory scale, turbulent combustor operating under imperfectly premixed mode with global equivalence ratio as the control parameter. The results indicate that the dynamics of thermoacoustic instability correspond to quasi-periodic bifurcation to low-dimensional, deterministic chaos, a route that is common to a variety of dissipative nonlinear systems. The results support the recent identification of bifurcation scenarios in a laminar premixed flame combustor (Kabiraj et al., Chaos: Interdiscip. J. Nonlinear Sci. 22, 023129 (2012)) and extend the observation to a practically relevant combustor configuration. PMID:25725637

  7. Combustor oscillating pressure stabilization and method

    DOEpatents

    Gemmen, Randall S.; Richards, George A.; Yip, Mui-Tong Joseph; Robey, Edward H.; Cully, Scott R.; Addis, Richard E.

    1998-01-01

    High dynamic pressure oscillations in hydrocarbon-fueled combustors typically occur when the transport time of the fuel to the flame front is at some fraction of the acoustic period. These oscillations are reduced to acceptably lower levels by restructuring or repositioning the flame front in the combustor to increase the transport time. A pilot flame front located upstream of the oscillating flame and pulsed at a selected frequency and duration effectively restructures and repositions the oscillating flame in the combustor to alter the oscillation-causing transport time.

  8. Chaos in an imperfectly premixed model combustor

    SciTech Connect

    Kabiraj, Lipika Saurabh, Aditya; Paschereit, Christian O.; Karimi, Nader; Sailor, Anna; Mastorakos, Epaminondas; Dowling, Ann P.

    2015-02-15

    This article reports nonlinear bifurcations observed in a laboratory scale, turbulent combustor operating under imperfectly premixed mode with global equivalence ratio as the control parameter. The results indicate that the dynamics of thermoacoustic instability correspond to quasi-periodic bifurcation to low-dimensional, deterministic chaos, a route that is common to a variety of dissipative nonlinear systems. The results support the recent identification of bifurcation scenarios in a laminar premixed flame combustor (Kabiraj et al., Chaos: Interdiscip. J. Nonlinear Sci. 22, 023129 (2012)) and extend the observation to a practically relevant combustor configuration.

  9. Combustor oscillating pressure stabilization and method

    DOEpatents

    Gemmen, R.S.; Richards, G.A.; Yip, M.T.J.; Robey, E.H.; Cully, S.R.; Addis, R.E.

    1998-08-11

    High dynamic pressure oscillations in hydrocarbon-fueled combustors typically occur when the transport time of the fuel to the flame front is at some fraction of the acoustic period. These oscillations are reduced to acceptably lower levels by restructuring or repositioning the flame front in the combustor to increase the transport time. A pilot flame front located upstream of the oscillating flame and pulsed at a selected frequency and duration effectively restructures and repositions the oscillating flame in the combustor to alter the oscillation-causing transport time. 7 figs.

  10. Gas turbine combustor stabilization by heat recirculation

    NASA Technical Reports Server (NTRS)

    Ganji, A.; Short, J.; Branch, M. C.; Oppenheim, A. K.

    1975-01-01

    The feasibility of heat recirculation for stabilization of lean mixtures and emission reduction has been studied in detail for a typical aircraft gas turbine combustor. Thermodynamic calculations have indicated temperature and heat recirculation rates for operation of the combustor over a range of combustion zone equivalence ratios and for varying modes of desired engine operation. Calculations indicate the feasibility of stabilizing the combustion zone at equivalence ratios as low as 0.2 with achievable heat recirculation rates. Detailed chemical kinetic calculations suggest that combustor heat release is maintained with reaction completion substantially before the NO forming reactions, even though CO is rapidly oxidized in this same region.

  11. Combustor design and analysis using the Rocket Combustor Interactive Design (ROCCID) methodology

    NASA Technical Reports Server (NTRS)

    Klem, Mark D.; Pieper, Jerry L.; Walker, Richard E.

    1990-01-01

    The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure.

  12. Combustor design and analysis using the ROCket Combustor Interactive Design (ROCCID) Methodology

    NASA Technical Reports Server (NTRS)

    Klem, Mark D.; Pieper, Jerry L.; Walker, Richard E.

    1990-01-01

    The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure.

  13. Thermally-Choked Combustor Technology

    NASA Technical Reports Server (NTRS)

    Knuth, William H.; Gloyer, P.; Goodman, J.; Litchford, R. J.

    1993-01-01

    A program is underway to demonstrate the practical feasibility of thermally-choked combustor technology with particular emphasis on rocket propulsion applications. Rather than induce subsonic to supersonic flow transition in a geometric throat, the goal is to create a thermal throat by adding combustion heat in a diverging nozzle. Such a device would have certain advantages over conventional flow accelerators assuming that the pressure loss due to heat addition does not severely curtail propulsive efficiency. As an aid to evaluation, a generalized one-dimensional compressible flow analysis tool was constructed. Simplified calculations indicate that the process is fluid dynamically and thermodynamically feasible. Experimental work is also being carried out in an attempt to develop, assuming an array of practical issues are surmountable, a practical bench-scale demonstrator using high flame speed H2/O2 combustibles.

  14. Alternate-Fueled Combustor-Sector Performance. Parts A and B; (A) Combustor Performance; (B) Combustor Emissions

    NASA Technical Reports Server (NTRS)

    Shouse, D. T.; Hendricks, R. C.; Lynch, A.; Frayne, C. W.; Stutrud, J. S.; Corporan, E.; Hankins, T.

    2012-01-01

    Alternate aviation fuels for military or commercial use are required to satisfy MIL-DTL-83133F(2008) or ASTM D 7566 (2010) standards, respectively, and are classified as "drop-in" fuel replacements. To satisfy legacy issues, blends to 50% alternate fuel with petroleum fuels are certified individually on the basis of processing and assumed to be feedstock agnostic. Adherence to alternate fuels and fuel blends requires "smart fueling systems" or advanced fuel-flexible systems, including combustors and engines, without significant sacrifice in performance or emissions requirements. This paper provides preliminary performance (Part A) and emissions and particulates (Part B) combustor sector data. The data are for nominal inlet conditions at 225 psia and 800 F (1.551 MPa and 700 K), for synthetic-paraffinic-kerosene- (SPK-) type (Fisher-Tropsch (FT)) fuel and blends with JP-8+100 relative to JP-8+100 as baseline fueling. Assessments are made of the change in combustor efficiency, wall temperatures, emissions, and luminosity with SPK of 0%, 50%, and 100% fueling composition at 3% combustor pressure drop. The performance results (Part A) indicate no quantifiable differences in combustor efficiency, a general trend to lower liner and higher core flow temperatures with increased FT fuel blends. In general, emissions data (Part B) show little differences, but with percent increase in FT-SPK-type fueling, particulate emissions and wall temperatures are less than with baseline JP-8. High-speed photography illustrates both luminosity and combustor dynamic flame characteristics.

  15. Introducing the VRT gas turbine combustor

    NASA Astrophysics Data System (ADS)

    Melconian, Jerry O.; Mostafa, Abdu A.; Nguyen, Hung Lee

    1990-07-01

    An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.

  16. Experimental clean combustor program, phase 2

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Peduzzi, A.; Vitti, G. E.

    1976-01-01

    The alternate fuels investigation objective was to experimentally determine the impacts, if any, on exhaust emissions, performance, and durability characteristics of the hybrid and vorbix low pollution combustor concepts when operated on test fuels which simulate composition and property changes which might result from future broadened aviation turbine fuel specifications or use of synthetically derived crude feedstocks. Results of the program indicate a significant increase in CO and small NOX increase in emissions at idle for both combustor concepts, and an increase in THC for the vorbix concept. Minimal impact was observed on gaseous emissions at high power. The vorbix concept exhibited significant increase in exhaust smoke with increasing fuel aromatic content. Altitude stability was not affected for the vorbix combustor, but was substantially reduced for the hybrid concept. Severe carbon deposition was observed in both combustors following limited endurance testing with No. 2 home heat fuel. Liner temperature levels were insensitive to variations in aromatic content over the range of conditions investigated.

  17. TRW Advanced Slagging Coal Combustor Utility Demonstration

    SciTech Connect

    Not Available

    1989-01-01

    The TRW Advanced Slagging Coal Combustor Demonstration Project consists of retrofitting Orange and Rockland (O R) Utility Corporation's Lovett Plant Unit No. 3 with four (4) slagging combustors which will allow the gas/ou desip unit to fire 2.5 sulfur coal. The slogging combustor process will provide NO[sub x] and SO[sub x] emissions that meet NSPS and New York State Envirommental Standards. TRW-CBU scope of work includes the engineering, design and supply of the slogging combustors, coal and limestone feed systems and a control system for these components. During this report period, the design activities for all systems progressed to permit the release of specifications and requests for proposals. Award of contracts for long-delivery items and major equipment are being placed to meet the revised program schedule.

  18. Development of an Advanced Annular Combustor

    NASA Technical Reports Server (NTRS)

    Rusnak, J. P.; Shadowen, J. H.

    1969-01-01

    The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

  19. Introducing the VRT gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Melconian, Jerry O.; Mostafa, Abdu A.; Nguyen, Hung Lee

    1990-01-01

    An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.

  20. Scramjet including integrated inlet and combustor

    SciTech Connect

    Kutschenreuter, P.H. Jr.; Blanton, J.C.

    1992-02-04

    This patent describes a scramjet engine. It comprises: a first surface including an aft facing step; a cowl including: a leading edge and a trailing edge; an upper surface and a lower surface extending between the leading edge and the trailing edge; the cowl upper surface being spaced from and generally parallel to the first surface to define an integrated inlet-combustor therebetween having an inlet for receiving and channeling into the inlet-combustor supersonic inlet airflow; means for injecting fuel into the inlet-combustor at the step for mixing with the supersonic inlet airflow for generating supersonic combustion gases; and further including a spaced pari of sidewalls extending between the first surface to the cowl upper surface and wherein the integrated inlet-combustor is generally rectangular and defined by the sidewall pair, the first surface and the cowl upper surface.

  1. Combustor assembly in a gas turbine engine

    DOEpatents

    Wiebe, David J; Fox, Timothy A

    2013-02-19

    A combustor assembly in a gas turbine engine. The combustor assembly includes a combustor device coupled to a main engine casing, a first fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner disposed radially inwardly from the flow sleeve. The first fuel injection system provides fuel that is ignited with the pressurized air creating first working gases. The intermediate duct is disposed between the liner and the transition duct and defines a path for the first working gases to flow from the liner to the transition duct. An intermediate duct inlet portion is associated with a liner outlet and allows movement between the intermediate duct and the liner. An intermediate duct outlet portion is associated with a transition duct inlet section and allows movement between the intermediate duct and the transition duct.

  2. Experimental clean combustor program; noise measurement addendum, Phase 2

    NASA Technical Reports Server (NTRS)

    Emmerling, J. J.; Bekofske, K. L.

    1976-01-01

    Combustor noise measurements were performed using wave guide probes. Test results from two full scale annular combustor configurations in a combustor test rig are presented. A CF6-50 combustor represented a current design, and a double annular combustor represented the advanced clean combustor configuration. The overall acoustic power levels were found to correlate with the steady state heat release rate and inlet temperature. A theoretical analysis for the attenuation of combustor noise propagating through a turbine was extended from a subsonic relative flow condition to include the case of supersonic flow at the discharge side. The predicted attenuation from this analysis was compared to both engine data and extrapolated component combustor data. The attenuation of combustor noise through the CF6-50 turbine was found to be greater than 14 dB by both the analysis and the data.

  3. Scramjet Combustor Characteristics at Hypervelocity Condition over Mach 10 Flight

    NASA Astrophysics Data System (ADS)

    Takahashi, M.; Komuro, T.; Sato, K.; Kodera, M.; Tanno, H.; Itoh, K.

    2009-01-01

    To investigate possibility of reduction of a scramjet combustor size without thrust performance loss, a two-dimensional constant-area combustor of a previous engine model was replaced with the one with 23% lower-height. With the application of the lower-height combustor, the pressure in the combustor becomes 50% higher and the combustor length for the optimal performance becomes 43% shorter than the original combustor. The combustion tests of the modified engine model were conducted using a large free-piston driven shock tunnel at flow conditions corresponding to the flight Mach number from 9 to 14. CFD was also applied to the engine internal flows. The results showed that the mixing and combustion heat release progress faster to the distance and the combustor performance similar to that of the previous engine was obtained with the modified engine. The reduction of the combustor size without the thrust performance loss is successfully achieved by applying the lower-height combustor.

  4. Small Gas Turbine Combustor Primary Zone Study

    NASA Technical Reports Server (NTRS)

    Sullivan, R. E.; Young, E. R.; Miles, G. A.; Williams, J. R.

    1983-01-01

    A development process is described which consists of design, fabrication, and preliminary test evaluations of three approaches to internal aerodynamic primary zone flow patterns: (1) conventional double vortex swirl stabilization; (2) reverse flow swirl stabilization; and (3) large single vortex flow system. Each concept incorporates special design features aimed at extending the performance capability of the small engine combustor. Since inherent geometry of these combustors result in small combustion zone height and high surface area to volume ratio, design features focus on internal aerodynamics, fuel placement, and advanced cooling. The combustors are evaluated on a full scale annular combustor rig. A correlation of the primary zone performance with the overall performance is accomplished using three intrusion type gas sampling probes located at the exit of the primary zone section. Empirical and numerical methods are used for designing and predicting the performance of the three combustor concepts and their subsequent modifications. The calibration of analytical procedures with actual test results permits an updating of the analytical design techniques applicable to small reverse flow annular combustors.

  5. Rolling contact mounting arrangement for a ceramic combustor

    DOEpatents

    Boyd, Gary L.; Shaffer, James E.

    1995-01-01

    A combustor assembly having a preestablished rate of thermal expansion is mounted within a gas turbine engine housing having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the combustor assembly. The combustor assembly is constructed of a inlet end portion, a outlet end portion and a plurality of combustor ring segments positioned between the end portions. A mounting assembly is positioned between the combustor assembly and the gas turbine engine housing to allow for the difference in the rate of thermal expansion while maintaining axially compressive force on the combustor assembly to maintain contact between the separate components.

  6. Rolling contact mounting arrangement for a ceramic combustor

    DOEpatents

    Boyd, G.L.; Shaffer, J.E.

    1995-10-17

    A combustor assembly having a preestablished rate of thermal expansion is mounted within a gas turbine engine housing having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the combustor assembly. The combustor assembly is constructed of a inlet end portion, a outlet end portion and a plurality of combustor ring segments positioned between the end portions. A mounting assembly is positioned between the combustor assembly and the gas turbine engine housing to allow for the difference in the rate of thermal expansion while maintaining axially compressive force on the combustor assembly to maintain contact between the separate components. 3 figs.

  7. Experimental Clean Combustor Program (ECCP), phase 3. [commercial aircraft turbofan engine tests with double annular combustor

    NASA Technical Reports Server (NTRS)

    Gleason, C. C.; Bahr, D. W.

    1979-01-01

    A double annular advanced technology combustor with low pollutant emission levels was evaluated in a series of CF6-50 engine tests. Engine lightoff was readily obtained and no difficulties were encountered with combustor staging. Engine acceleration and deceleration were smooth, responsive and essentially the same as those obtainable with the CF6-50 combustor. The emission reductions obtained in carbon monoxide, hydrocarbons, and nitrogen oxide levels were 55, 95, and 30 percent, respectively, at an idle power setting of 3.3 percent of takeoff power on an EPA parameter basis. Acceptable smoke levels were also obtained. The exit temperature distribution of the combustor was found to be its major performance deficiency. In all other important combustion system performance aspects, the combustor was found to be generally satisfactory.

  8. A Comparison of Combustor-Noise Models

    NASA Technical Reports Server (NTRS)

    Hultgren, Lennart S.

    2012-01-01

    The present status of combustor-noise prediction in the NASA Aircraft Noise Prediction Program (ANOPP)1 for current-generation (N) turbofan engines is summarized. Several semi-empirical models for turbofan combustor noise are discussed, including best methods for near-term updates to ANOPP. An alternate turbine-transmission factor2 will appear as a user selectable option in the combustor-noise module GECOR in the next release. The three-spectrum model proposed by Stone et al.3 for GE turbofan-engine combustor noise is discussed and compared with ANOPP predictions for several relevant cases. Based on the results presented herein and in their report,3 it is recommended that the application of this fully empirical combustor-noise prediction method be limited to situations involving only General-Electric turbofan engines. Long-term needs and challenges for the N+1 through N+3 time frame are discussed. Because the impact of other propulsion-noise sources continues to be reduced due to turbofan design trends, advances in noise-mitigation techniques, and expected aircraft configuration changes, the relative importance of core noise is expected to greatly increase in the future. The noise-source structure in the combustor, including the indirect one, and the effects of the propagation path through the engine and exhaust nozzle need to be better understood. In particular, the acoustic consequences of the expected trends toward smaller, highly efficient gas-generator cores and low-emission fuel-flexible combustors need to be fully investigated since future designs are quite likely to fall outside of the parameter space of existing (semi-empirical) prediction tools.

  9. Combustor bulkhead heat shield assembly

    SciTech Connect

    Zeisser, M.H.

    1990-06-19

    This paper describes a gas turbine engine having an annular combustion chamber defined by an annular, inner liner, a concentric outer liner, and an upstream annular combustor head, wherein the head includes a radially extending bulkhead having circumferentially distributed openings for each receiving an individual fuel nozzle therethrough. It comprises: a segmented heat shield assembly, disposed between the combustion chamber interior and the bulkhead, including generally planar, sector shaped heat shields, each shield abutting circumferentially with two next adjacent shields and extending radially from proximate the inner liner to proximate the outer liner, the plurality of shields collectively defining an annular protective barrier, and wherein each sector shaped shield further includes an opening, corresponding to one of the bulkhead nozzle openings for likewise receiving the corresponding nozzle therethrough, the shield opening further including an annular lip extending toward the bulkhead and being received within the bulkhead opening, raised ridges on the shield backside, the ridges contacting the facing bulkhead surface and defining a flow path for a flow of cooling air issuing from a sized supply opening disposed in the bulkhead, the flow path running ultimately from adjacent the annular lip to the edges of each shield segment, wherein the raised edges extend fully along the lateral, circumferentially spaced edges of each shield segment and about the adjacent shield segments wherein the raised ridges further extend circumferentially between the annular lip and the abutting edge ridges.

  10. Low NO/sub x/ Heavy Fuel Combustor Concept Program. Phase I. Final report

    SciTech Connect

    Cutrone, M B

    1981-10-01

    Six combustor concepts were designed, fabricated, and underwent a series of combustion tests with the objective of evaluating and developing a combustor capable of meeting US New Source Performance Standards (NSPS), dry, for high-nitrogen liquid fuels. Three rich/lean and three lean/lean two-stage combustors were tested with ERBS distillate, petroleum residual, and SRC-II coal derived liquid (CDL) fuels with fuel-bound nitrogen contents of 0.0054, 0.23, and 0.87 weight percent, respectively. A lean/lean concept was demonstrated with ultralow NO/sub x/ emissions, dry, of 5 gm NO/sub x/kg fuel on ERBS, and NO/sub x/ emissions meeting the NSPS NO/sub x/ standard on residual fuel. This combustor concept met operational goals for pressure drop, smoke, exhaust pattern factor, and combustion efficiency. A rich/lean concept was identified and developed which demonstrated NO/sub x/ emissions approaching the NSPS standards, dry, for all liquid fuels including the 0.87 weight percent nitrogen SRC-II coal-derived liquid. Exhaust pattern factor and pressure drop met or approached goals. Smoke emissions were higher than the program goal. However, a significant improvement was made with only a minor modification of the fuel injector/air swirler system, and further development should result in meeting smoke goals for all fuels. Liner metal temperatures were higher than allowable for commercial application. Conceptual designs for further development of these two rich/lean and lean/lean concepts have been completed which address smoke and metal temperature concerns, and are available for the next phase of this NASA-sponsored, DOE-funded program. Tests of a rich/lean concept, and a catalytic combustor concept using low- and intermediate-Btu simulated coal-derived gases will be completed during the ongoing Phase IA extension of this program.

  11. Experimental clean combustor program, alternate fuels addendum, phase 2

    NASA Technical Reports Server (NTRS)

    Gleason, C. C.; Bahr, D. W.

    1976-01-01

    The characteristics of current and advanced low-emissions combustors when operated with special test fuels simulating broader range combustion properties of petroleum or coal derived fuels were studied. Five fuels were evaluated; conventional JP-5, conventional No. 2 Diesel, two different blends of Jet A and commercial aromatic mixtures - zylene bottoms and haphthalene charge stock, and a fuel derived from shale oil crude which was refined to Jet A specifications. Three CF6-50 engine size combustor types were evaluated; the standard production combustor, a radial/axial staged combustor, and a double annular combustor. Performance and pollutant emissons characteristics at idle and simulated takeoff conditions were evaluated in a full annular combustor rig. Altitude relight characteristics were evaluated in a 60 degree sector combustor rig. Carboning and flashback characteristics at simulated takeoff conditions were evaluated in a 12 degree sector combustor rig. For the five fuels tested, effects were moderate, but well defined.

  12. Exhaust gas measurements in a propane fueled swirl stabilized combustor

    NASA Technical Reports Server (NTRS)

    Aanad, M. S.

    1982-01-01

    Exhaust gas temperature, velocity, and composition are measured and combustor efficiencies are calculated in a lean premixed swirl stabilized laboratory combustor. The radial profiles of the data between the co- and the counter swirl cases show significant differences. Co-swirl cases show evidence of poor turbulent mixing across the combustor in comparison to the counter-swirl cases. NO sub x levels are low in the combustor but substantial amounts of CO are present. Combustion efficiencies are low and surprisingly constant with varying outer swirl in contradiction to previous results under a slightly different inner swirl condition. This difference in the efficiency trends is expected to be a result of the high sensitivity of the combustor to changes in the inner swirl. Combustor operation is found to be the same for propane and methane fuels. A mechanism is proposed to explain the combustor operation and a few important characteristics determining combustor efficiency are identified.

  13. Dish stirling solar receiver combustor test program

    NASA Technical Reports Server (NTRS)

    Bankston, C. P.; Back, L. H.

    1981-01-01

    The operational and energy transfer characteristics of the Dish Stirling Solar Receiver (DSSR) combustor/heat exchanger system was evaluated. The DSSR is designed to operate with fossil fuel augmentation utilizing a swirl combustor and cross flow heat exchanger consisting of a single row of 4 closely spaced tubes that are curved into a conical shape. The performance of the combustor/heat exchanger system without a Stirling engine was studied over a range of operating conditions and output levels using water as the working fluid. Results show that the combustor may be started under cold conditions, controlled safety, and operated at a constant air/fuel ratio (10 percent excess air) over the required range of firing rates. Furthermore, nondimensional heat transfer coefficients based on total heat transfer are plotted versus Reynolds number and compared with literature data taken for single rows of closely spaced tubes perpendicular to cross flow. The data show enhanced heat transfer for the present geometry and test conditions. Analysis of the results shows that the present system meets specified thermal requirements, thus verifying the feasibility of the DSSR combustor design for final prototype fabrication.

  14. Alternate-Fueled Combustor-Sector Performance

    NASA Technical Reports Server (NTRS)

    Thomas, Anna E.; Saxena, Nikita T.; Shouse, Dale T.; Neuroth, Craig; Hendricks, Robert C.; Lynch, Amy; Frayne, Charles W.; Stutrud, Jeffrey S.; Corporan, Edwin; Hankins, Terry

    2013-01-01

    In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MIL-DTL-83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as "drop-in" fuel replacements. This report provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fischer-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 degF (533 K), 125 psia (0.86 MPa) at 625 degF (603 K), 175 psia (1.21 MPa) at 725 degF (658 K), and 225 psia (1.55 MPa) at 790 degF (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3, 4, and 5 percent combustor pressure drop (DP) for fuel:air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade and vane lives.

  15. Alternate-Fueled Combustor-Sector Performance

    NASA Technical Reports Server (NTRS)

    Thomas, Anna E.; Saxena, Nikita T.; Shouse, Dale T.; Neuroth, Craig; Hendricks, Robert C.; Lynch, Amy; Frayne, Charles W.; Stutrud, Jeffrey S.; Corporan, Edwin; Hankins, Terry

    2012-01-01

    In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MILDTL- 83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as drop-in fuel replacements. This paper provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fisher-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 F (533 K), 125 psia (0.86 MPa) at 625 F (603 K), 175 psia (1.21 MPa) at 725 F (658 K), and 225 psia (1.55 MPa) at 790 F (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3%, 4%, and 5% combustor pressure drop (% delta P) for fuel: air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade/vane life.

  16. Combustor development for automotive gas turbines

    SciTech Connect

    Ross, P.T.; Anderson, D.N.; Williams, J.R.

    1983-09-01

    This paper describes the development of a combustion system for the AGT 100 automotive gas turbine engine. The AGT 100 is a 100 hp engine being developed by Detroit Diesel Allison Division of General Motors Corporation. To achieve optimum fuel economy, the AGT 100 engine operates on a regenerative cycle. A maximum turbine inlet temperature of 1288/sup 0/C (2350/sup 0/F) is reached, and air is supplied to the inlet of the combustor at temperatures as high as 1024/sup 0/C (1875/sup 0/F). To meet the low-emission and high-durability requirements at these conditions, a premix/prevaporization ceramic combustor employing variable geometry to control the temperature in the burning zone has been developed. A test section capable of handling 1024/sup 0/C (1875/sup 0/F) inlet air was designed and fabricated to evaluate this combustor. Testing of both metal (transpiration cooled) and ceramic combustors was conducted. Emissions were measured and found to be a function of burner inlet temperature. At 999/sup 0/C (1830/sup 0/F) burner inlet temperature, NO /SUB x/ emissions were two orders of magnitude below the program goals. At the same temperature but at a different variable-geometry position, the CO was 30 times below the program goal. Considerable testing was conducted to evaluate the behavior of the ceramic materials used in the combustor. No failures occurred during steady-state operation; however, some cracks developed in the dome during extended transient operation.

  17. Experimental clean combustor program noise measurement addendum, phase 1

    NASA Technical Reports Server (NTRS)

    Emmerling, J. J.

    1975-01-01

    The test results of combustor noise measurements taken with waveguide probes are presented. Waveguide probes were shown to be a viable measurement technique for determining high sound pressure level broadband noise. A total of six full-scale annular combustors were tested and included the three advanced combustor designs: swirl-can, radial/axial, and double annular.

  18. LDV measurements in an annular combustor model

    NASA Astrophysics Data System (ADS)

    Barron, Dean A.

    1986-08-01

    The design and setup of a Laser Doppler Velocimeter (LDV) system used to take velocity measurements in an annular combustor model are covered. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. The LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

  19. Combustor development for automotive gas turbines

    NASA Technical Reports Server (NTRS)

    Ross, P. T.; Williams, J. R.; Anderson, D. N.

    1982-01-01

    The development of a combustion system for the AGT 100 automotive gas turbine engine is described. A maximum turbine inlet temperature of 1288 C is reached during the regenerative cycle, and air up to 1024 C is supplied to the combustor inlet. A premix/prevaporization ceramic combustor employing variable geometry to control burning zone temperature was developed and tested. Tests on both metal and ceramic combustors showed that emissions were a function of burner inlet temperature (BIT). At 999 C BIT, NO(x) emissions were two orders of magnitude below program goals, and at the same temperature but at a different variable geometry position, the CO was 30 times below program goal. Tests to evaluate the durability of the ceramic materials showed no failures during steady-state operation; however, some cracks developed in the dome during extended transient operation.

  20. Flow establishment in a generic scramjet combustor

    NASA Technical Reports Server (NTRS)

    Jacobs, P. A.; Rogers, R. C.; Weidner, E. H.; Bittner, R. D.

    1990-01-01

    The establishment of a quasi-steady flow in a generic scramjet combustor was studied for the case of a time varying inflow to the combustor. Such transient flow is characteristic of the reflected shock tunnel and expansion tube test facilities. Several numerical simulations of hypervelocity flow through a straight duct combustor with either a side wall step fuel injector or a centrally located strut injector are presented. Comparisons were made between impulsively started but otherwise constant flow conditions (typical of the expansion tube or tailored operations of the reflected shock tunnel) and the relaxing flow produced by the 'undertailored' operations of the reflected shock tunnel. Generally the inviscid flow features, such as the shock pattern and pressure distribution, were unaffected by the time varying inlet conditions and approached steady state in approx. the times indicated by experimental correlations. However, viscous features, such as heat transfer and skin friction, were altered by the relaxing inlet flow conditions.

  1. LDV Measurements in an Annular Combustor Model

    NASA Technical Reports Server (NTRS)

    Barron, Dean A.

    1996-01-01

    This thesis covers the design and setup of a laser doppler velocimeter (LDV) system used to take velocity measurements in an annular combustor model. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

  2. Optical Detection Of Flameout In A Combustor

    NASA Technical Reports Server (NTRS)

    Borg, Stephen E.; West, James W.; Harper, Samuel E.; Alderfer, David W.; Lawrence, Robert M.

    1994-01-01

    Fuel supply shut down in time to prevent explosion. Optical flameout detector designed to signal control system of facility to cut off supply of fuel into combustion chamber if flame goes out. Combustor which optical flameout detector designed burns methane in air to provide hot gases for 8-ft high-temperature test chamber. Acoustical flameout detector for same combustor described in "Acoustical Detection of Flameout in Combustor" (LAR-14900). Fiber optic probes mounted to fuel-spray bar upstream of flame. No focusing optics used, and probes aimed across flow of gases at spot on combustion chamber wall downstream from spray bar. Arrangement enables flameout detection system to respond quickly to potential loss of flame since it detects movement of flame front away from spray bar face. Overall response time of detection system under 10 milliseconds.

  3. Combustor assembly in a gas turbine engine

    DOEpatents

    Wiebe, David J; Fox, Timothy A

    2015-04-28

    A combustor assembly in a gas turbine engine includes a combustor device, a fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner surrounded by the flow sleeve. The fuel injection system provides fuel to be mixed with the pressurized air and ignited in the liner to create combustion products. The intermediate duct is disposed between the liner and the transition duct so as to define a path for the combustion products to flow from the liner to the transition duct. The intermediate duct is associated with the liner such that movement may occur therebetween, and the intermediate duct is associated with the transition duct such that movement may occur therebetween. The flow sleeve includes structure that defines an axial stop for limiting axial movement of the intermediate duct.

  4. Preliminary calibration of a generic scramjet combustor

    NASA Technical Reports Server (NTRS)

    Jacobs, P. A.; Morgan, R. G.; Rogers, R. C.; Wendt, M.; Brescianini, C.; Paull, A.; Kelly, G.

    1991-01-01

    The results of a preliminary investigation of the combustion of hydrogen fuel at hypersonic flow conditions are provided. The tests were performed in a generic, constant-area combustor model with test gas supplied by a free-piston-driven reflected-shock tunnel. Static pressure measurements along the combustor wall indicated that burning did occur for combustor inlet conditions of P(static) approximately equal to 19kPa, T(static) approximately equal to 1080 K, and U approximately equal to 3630 m/s with a fuel equivalence ratio approximately equal to 0.9. These inlet conditions were obtained by operating the tunnel with stagnation enthalpy approximately equal to 8.1 MJ/kg, stagnation pressure approximately equal to 52 MPa, and a contoured nozzle with a nominal exit Mach number of 5.5.

  5. Ground idle performance improvement of a double-annular combustor by using simulated variable combustor geometry

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1975-01-01

    A test program was undertaken to determine if variable combustor geometry could be used to reduce exhaust emissions of a low-pressure-ratio jet engine operating at ground idle conditions. Three techniques for varying combustor geometry were simulated. Other techniques evaluated were radial fuel staging and the use of preheated fuel. When simulated variable combustor geometry was employed with radial fuel staging, combustion efficiency at a fuel-air ratio of 0.01 was increased from 77 to 95 percent, and exhaust emissions of unburned hydrocarbons and carbon monoxide were significantly reduced.

  6. Experiment of rocket-ram annular combustor

    NASA Astrophysics Data System (ADS)

    Yatsuyanagi, N.; Sakamoto, H.; Sato, K.; Ono, F.; Sasaki, M.; Takahashi, M.

    In this experiment, the double-nozzle type of rocket-ram annular combustor with a total thrust of 5kN was designed and tested with varying ratios of thrust produced by rocket and ram. Thrust and pressure distribution along the common expansion nozzle, i.e., the ram combustor nozzle, were measured to investigate the effect of interaction of the two expansion gases on thrust. Enhancement of specific impulse was verified by the experiments. That is, the specific impulse gains in rocket-ram parallel operation, the ratio of rocket thrust to ram thrust being 50 to 50, were found to be 190 percent of gains in pure rocket operation.

  7. Combustion characteristics of a double swirling combustor

    NASA Astrophysics Data System (ADS)

    Jiang, Zhi; Gao, Ge; Ning, Huang

    1991-01-01

    The combustion characteristics at atmospheric pressure for a coaxial double swirler combustor have been investigated. The results show that its combustion performances can be significantly improved by increasing properly swirl number for the inner swirler, as outer and inner flows are counterrotating. The annular recirculation zone is a good source for flame stabilization, which is a important factor for extending the weak extinction limit. The temperature distribution at the exit section of the combustor can be controlled by varying fuel-air ratio of outer and inner swirling flows.

  8. Predicting and Preventing Incipient Flameout in Combustors

    NASA Technical Reports Server (NTRS)

    Puster, Richard Lee

    2003-01-01

    A method of predicting and preventing incipient flameout in a combustor has been proposed. The method should be applicable to a variety of liquid- and gas-fueled combustors in furnaces and turbine engines. Until now, there have been methods of detecting flameouts after they have occurred, but there has been no way of predicting incipient flameouts and, hence, no way of acting in time to prevent them. Prevention of flameout could not only prevent damage to equipment but, in the case of aircraft turbine engines, could also save lives.

  9. Low NO.sub.x combustor

    DOEpatents

    Taylor, Jack R.

    1987-01-01

    A combustor having an annular first stage, a generally cylindrically-shaped second stage, and an annular conduit communicably connecting the first and second stages. The conduit has a relatively small annular height and a large number of quench holes in the walls thereof such that quench air injected into the conduit through the quench holes will mix rapidly with, or quench, the combustion gases flowing through the conduit. The rapid quenching reduces the amount of NO.sub.x produced in the combustor.

  10. Micro-combustor for gas turbine engine

    DOEpatents

    Martin, Scott M.

    2010-11-30

    An improved gas turbine combustor (20) including a basket (26) and a multiplicity of micro openings (29) arrayed across an inlet wall (27) for passage of a fuel/air mixture for ignition within the combustor. The openings preferably have a diameter on the order of the quenching diameter; i.e. the port diameter for which the flame is self-extinguishing, which is a function of the fuel mixture, temperature and pressure. The basket may have a curved rectangular shape that approximates the shape of the curved rectangular shape of the intake manifolds of the turbine.

  11. Laser diagnostics on a hypersonic combustor

    SciTech Connect

    Taylor, D.J.; Oldenborg, R.C.; Tiee, J.J.; Northam, G.B.; Antcliff, R.R.; Cutler, A.D.; Jarrett, O.; Smith, M.W. NASA, Langley Research Center, Hampton, VA )

    1991-01-01

    NASA-Langley has implemented a laser-based multipoint/multiparameter diagnostics system at its hypersonic direct-connect combustor, in order to measure both temperature and majority species densities in two dimensions, using spatially-scanned CARS; in addition, line-imaged measurements of radical densities are simultaneously generated by LIF at any of several planes downstream of the fuel injector. Initial experimental trials have demonstrated successful detection of one-dimensional images of OH density, as well as CARS N2-temperature measurements, in the turbulent reaction zone of the hypersonic combustor.

  12. System and method for controlling a combustor assembly

    DOEpatents

    York, William David; Ziminsky, Willy Steve; Johnson, Thomas Edward; Stevenson, Christian Xavier

    2013-03-05

    A system and method for controlling a combustor assembly are disclosed. The system includes a combustor assembly. The combustor assembly includes a combustor and a fuel nozzle assembly. The combustor includes a casing. The fuel nozzle assembly is positioned at least partially within the casing and includes a fuel nozzle. The fuel nozzle assembly further defines a head end. The system further includes a viewing device configured for capturing an image of at least a portion of the head end, and a processor communicatively coupled to the viewing device, the processor configured to compare the image to a standard image for the head end.

  13. Simulated Altitude Performance of Combustors for the 24C Jet Engine. 2; 24C-4 Combustor

    NASA Technical Reports Server (NTRS)

    Bernardo, Everett; Schroeter, Thomas T.; Miller, Robert C.

    1947-01-01

    The performance of a 24C-4 combustor was investigated with three different combustor baskets and five modifications of these baskets at conditions simulating static (zero-ram) operation of the 24C jet engine over ranges of altitude and engine speed to determine and improve the altitude operational limits of the 24C combustor. Information was also obtained regarding combustion characteristics, the fuel-flow characteristics of the fuel manifolds, and the combustor total-pressure drop. NACA modifications, which consisted of blocking rows of holes on the baskets, increased the minimum point on the altitude-operational-limit curve, which occurs at low engine speeds, for a narrow-upstream-end basket by 8000 feet (from 23, 000 to 31,000 ft_ and for a wide-upstream-end basket by 21,000 feet (from 12, 000 to 34,000 ft). These improvements were approximately maintained over the entire range of engine speeds investigated.

  14. Combustor for a low-emissions gas turbine engine

    DOEpatents

    Glezer, Boris; Greenwood, Stuart A.; Dutta, Partha; Moon, Hee-Koo

    2000-01-01

    Many government entities regulated emission from gas turbine engines including CO. CO production is generally reduced when CO reacts with excess oxygen at elevated temperatures to form CO2. Many manufactures use film cooling of a combustor liner adjacent to a combustion zone to increase durability of the combustion liner. Film cooling quenches reactions of CO with excess oxygen to form CO2. Cooling the combustor liner on a cold side (backside) away from the combustion zone reduces quenching. Furthermore, placing a plurality of concavities on the cold side enhances the cooling of the combustor liner. Concavities result in very little pressure reduction such that air used to cool the combustor liner may also be used in the combustion zone. An expandable combustor housing maintains a predetermined distance between the combustor housing and combustor liner.

  15. Active Control of High-Frequency Combustor Instability Demonstrated

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Chang, Clarence T.

    2003-01-01

    To reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities-high-pressure oscillations much like sound waves that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the combustor and turbine safe operating life. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Propulsion and Power Program, the NASA Glenn Research Center in partnership with Pratt & Whitney, United Technologies Research Center, and Georgia Institute of Technology is developing technologies for the active control of combustion instabilities.

  16. Stably operating pulse combustor and method

    DOEpatents

    Zinn, Ben T.; Reiner, David

    1990-01-01

    A pulse combustor apparatus adapted to burn either a liquid fuel or a pulverized solid fuel within a preselected volume of the combustion chamber. The combustion process is substantially restricted to an optimum combustion zone in order to attain effective pulse combustion operation.

  17. Stably operating pulse combustor and method

    DOEpatents

    Zinn, B.T.; Reiner, D.

    1990-05-29

    A pulse combustor apparatus is described which is adapted to burn either a liquid fuel or a pulverized solid fuel within a preselected volume of the combustion chamber. The combustion process is substantially restricted to an optimum combustion zone in order to attain effective pulse combustion operation. 4 figs.

  18. Flashback Arrestor for LPP, Low NOx Combustors

    NASA Technical Reports Server (NTRS)

    Kraemer, Gil; Lee, Chi-Ming

    1998-01-01

    Lean premixed, prevaporized (LPP) high temperature combustor designs as explored for the Advanced Subsonic Transport (AST) and High Speed Civil Transport (HSCT) combustors can achieve low NO(x), emission levels. An enabling device is needed to arrest flashback and inhibit preignition at high power conditions and during transients (surge and rapid spool down). A novel flashback arrestor design has demonstrated the ability to arrest flashback and inhibit preignition in a 4.6 cm diameter tubular reactor at full power inlet temperatures (725 C) using Jet-A fuel at 0.4 less than or equal To phi less than or equal to 3.5. Several low pressure loss (0.2 to 0.4% at 30 m/s) flashback arrestor designs were developed which arrested flashback at all of the test conditions. Flame holding was also inhibited off the flash arrestor face or within the downstream tube even velocities (less than or equal to 3 to 6 m/s), thus protecting the flashback arrestor and combustor components. Upstream flow conditions influence the specific configuration based on using either a 45% or 76% upstream geometric blockage. Stationary, lean premixed dry low NO(x) gas turbine combustors would also benefit from this low pressure drop flashback arrestor design which can be easily integrated into new and existing designs.

  19. Thermal Imaging Control of Furnaces and Combustors

    SciTech Connect

    David M. Rue; Serguei Zelepouga; Ishwar K. Puri

    2003-02-28

    The object if this project is to demonstrate and bring to commercial readiness a near-infrared thermal imaging control system for high temperature furnaces and combustors. The thermal imaging control system, including hardware, signal processing, and control software, is designed to be rugged, self-calibrating, easy to install, and relatively transparent to the furnace operator.

  20. Low NOx heavy fuel combustor concept program

    NASA Technical Reports Server (NTRS)

    White, D. J.; Lecren, R. T.; Batakis, A. P.

    1981-01-01

    A total of twelve low NOx combustor configurations, embodying three different combustion concepts, were designed and fabricated as modular units. These configurations were evaluated experimentally for exhaust emission levels and for mechanical integrity. Emissions data were obtained in depth on two of the configurations.

  1. Catalytic combustion with incompletely vaporized residual fuel

    NASA Technical Reports Server (NTRS)

    Rosfjord, T. J.

    1981-01-01

    Catalytic combustion of fuel lean mixtures of incompletely vaporized residual fuel and air was investigated. The 7.6 cm diameter, graded cell reactor was constructed from zirconia spinel substrate and catalyzed with a noble metal catalyst. Streams of luminous particles exited the rector as a result of fuel deposition and carbonization on the substrate. Similar results were obtained with blends of No. 6 and No. 2 oil. Blends of shale residual oil and No. 2 oil resulted in stable operation. In shale oil blends the combustor performance degraded with a reduced degree of fuel vaporization. In tests performed with No. 2 oil a similar effect was observed.

  2. Advanced Low NOx Combustors for Aircraft Gas Turbines

    NASA Technical Reports Server (NTRS)

    Roberts, P. B.; White, D. J.; Shekleton, J. R.; Butze, H. F.

    1976-01-01

    A test rig program was conducted with the objective of evaluating and minimizing the exhaust emissions, in particular NOx, of two advanced aircraft combustor concepts at a simulated high-altitude cruise condition. The two pre-mixed, lean-reaction designs are known as the Jet Induced Circulation (JIC) combustor and the Vortex Air Blast (VAB) combustor and were rig tested in the form of reverse flow can combustors in the 0.13 ni (5.0 in. ) size range. Various configuration modifications were applied to the JIC and VAB combustor designs in an effort to reduce the emissions levels. The VAB combustor demonstrated a NOx level of 1.11 gm NO2/kg fuel with essentially 100 percent combustion efficiency at the simulated cruise combustor condition of 507 kPa (5 atm), 833 K (1500 R), inlet pressure and temperature respectively, and 1778 K (3200 R) outlet temperature on Jet-Al fuel. These configuration screening tests were carried out on essentially reaction zones only, in order to simplify the construction and modification of the combustors and to uncouple any possible effects on the emissions produced by the dilution flow. Tests were also conducted however at typical engine idle conditions on both combustors equipped with dilution ports in order to better define the problem areas involved in the operation of such concepts over a complete engine operational envelope. Versions of variable-geometry, JIC and VAB annular combustors are proposed.

  3. Durability testing at 5 atmospheres of advanced catalysts and catalyst supports for gas turbine engine combustors

    NASA Technical Reports Server (NTRS)

    Olson, B. A.; Lee, H. C.; Osgerby, I. T.; Heck, R. M.; Hess, H.

    1980-01-01

    The durability of CATCOM catalysts and catalyst supports was experimentally demonstrated in a combustion environment under simulated gas turbine engine combustor operating conditions. A test of 1000 hours duration was completed with one catalyst using no. 2 diesel fuel and operating at catalytically-supported thermal combustion conditions. The performance of the catalyst was determined by monitoring emissions throughout the test, and by examining the physical condition of the catalyst core at the conclusion of the test. Tests were performed periodically to determine changes in catalytic activity of the catalyst core. Detailed parametric studies were also run at the beginning and end of the durability test, using no. 2 fuel oil. Initial and final emissions for the 1000 hours test respectively were: unburned hydrocarbons (C3 vppm):0, 146, carbon monoxide (vppm):30, 2420; nitrogen oxides (vppm):5.7, 5.6.

  4. STUDY OF THE EFFECTIVENESS OF A CATALYTIC COMBUSTION DEVICE ON A WOOD BURNING APPLIANCE

    EPA Science Inventory

    The report gives results of operating a wood stove, incorporating a catalytic combustor, while burning air-dried oak at low burning rates. Gas composition was measured continuously both at the entrance of the catalyst (after the gases had left the burning wood) and at the exit of...

  5. Systems Characterization of Combustor Instabilities With Controls Design Emphasis

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2004-01-01

    This effort performed test data analysis in order to characterize the general behavior of combustor instabilities with emphasis on controls design. The analysis is performed on data obtained from two configurations of a laboratory combustor rig and from a developmental aero-engine combustor. The study has characterized several dynamic behaviors associated with combustor instabilities. These are: frequency and phase randomness, amplitude modulations, net random phase walks, random noise, exponential growth and intra-harmonic couplings. Finally, the very cause of combustor instabilities was explored and it could be attributed to a more general source-load type impedance interaction that includes the thermo-acoustic coupling. Performing these characterizations on different combustors allows for more accurate identification of the cause of these phenomena and their effect on instability.

  6. Energy efficient engine combustor test hardware detailed design report

    NASA Technical Reports Server (NTRS)

    Zeisser, M. H.; Greene, W.; Dubiel, D. J.

    1982-01-01

    The combustor for the Energy Efficient Engine is an annular, two-zone component. As designed, it either meets or exceeds all program goals for performance, safety, durability, and emissions, with the exception of oxides of nitrogen. When compared to the configuration investigated under the NASA-sponsored Experimental Clean Combustor Program, which was used as a basis for design, the Energy Efficient Engine combustor component has several technology advancements. The prediffuser section is designed with short, strutless, curved-walls to provide a uniform inlet airflow profile. Emissions control is achieved by a two-zone combustor that utilizes two types of fuel injectors to improve fuel atomization for more complete combustion. The combustor liners are a segmented configuration to meet the durability requirements at the high combustor operating pressures and temperatures. Liner cooling is accomplished with a counter-parallel FINWALL technique, which provides more effective heat transfer with less coolant.

  7. Pollution measurements of a swirl-can combustor

    NASA Technical Reports Server (NTRS)

    Niedzwiecki, R. W.; Jones, R. E.

    1972-01-01

    Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for an experimental, annular, swirl can combustor. The combustor was 42 inches in diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 600, 900 and 1050 F, inlet pressures of 5 to 6 atmospheres, reference velocities of 69 to 120 feet per second and fuel-air ratios of 0.014 to 0.0695. Tests were also conducted at a simulated engine idle condition. Results demonstrated that swirl can combustors produce oxides of nitrogen levels substantially lower than conventional combustor designs. These reductions are attributed to reduced dwell times resulting from short combustor length, quick mixing of combustion gases with diluent air, and to uniform fuel distributions resulting from the swirl can approach. Radial staging of fuel at idle conditions resulted in increases in combustion efficiencies and corresponding reductions in pollutant levels.

  8. Predictive models for circulating fluidized bed combustors

    SciTech Connect

    Gidaspow, D.

    1989-11-01

    The overall objective of this investigation is to develop experimentally verified models for circulating fluidized bed (CFB) combustors. The purpose of these models is to help American industry, such as Combustion Engineering, design and scale-up CFB combustors that are capable of burning US Eastern high sulfur coals with low SO{sub x} and NO{sub x} emissions. In this report, presented as a technical paper, solids distributions and velocities were computed for a PYROFLOW circulating fluidized bed system. To illustrate the capability of the computer code an example of coal-pyrite separation is included, which was done earlier for a State of Illinois project. 24 refs., 20 figs., 2 tabs.

  9. Ring connection for porous combustor wall panels

    NASA Technical Reports Server (NTRS)

    Verdouw, Albert J. (Inventor)

    1980-01-01

    A gas turbine engine combustor assembly of unique configuration has an outer wall made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween by a reinforcing and heat dissipation ring and a unique weld configuration to prevent thermal erosion of the ends of the porous metal panels at the butt joints; the combustor further including a unique inner wall made up of a plurality of like axially extending multi-layered porous metal panels joined at butt joints by a reinforcing and heat dissipation ring on the inner surface of the inner wall panels and an improved butt weld joint that prevents thermal erosion of the ends of the porous metal inner wall panels.

  10. Advanced composite combustor structural concepts program

    NASA Technical Reports Server (NTRS)

    Sattar, M. A.; Lohmann, R. P.

    1984-01-01

    An analytical study was conducted to assess the feasibility of and benefits derived from the use of high temperature composite materials in aircraft turbine engine combustor liners. The study included a survey and screening of the properties of three candidate composite materials including tungsten reinforced superalloys, carbon-carbon and silicon carbide (SiC) fibers reinforcing a ceramic matrix of lithium aluminosilicate (LAS). The SiC-LAS material was selected as offering the greatest near term potential primarily on the basis of high temperature capability. A limited experimental investigation was conducted to quantify some of the more critical mechanical properties of the SiC-LAS composite having a multidirection 0/45/-45/90 deg fiber orientation favored for the combustor linear application. Rigorous cyclic thermal tests demonstrated that SiC-LAS was extremely resistant to the thermal fatigue mechanisms that usually limit the life of metallic combustor liners. A thermal design study led to the definition of a composite liner concept that incorporated film cooled SiC-LAS shingles mounted on a Hastelloy X shell. With coolant fluxes consistent with the most advanced metallic liner technology, the calculated hot surface temperatures of the shingles were within the apparent near term capability of the material. Structural analyses indicated that the stresses in the composite panels were low, primarily because of the low coefficient of expansion of the material and it was concluded that the dominant failure mode of the liner would be an as yet unidentified deterioration of the composite from prolonged exposure to high temperature. An economic study, based on a medium thrust size commercial aircraft engine, indicated that the SiC-LAS combustor liner would weigh 22.8N (11.27 lb) less and cost less to manufacture than advanced metallic liner concepts intended for use in the late 1980's.

  11. Fuel property effects in stirred combustors

    NASA Technical Reports Server (NTRS)

    1980-01-01

    Soot formation in strongly backmixed combustion was investigated using the jet-stirred combustor (JSC). This device provided a combustion volume in which temperature and combustion were uniform. It simulated the recirculating characteristics of the gas turbine primary zone; it was in this zone where mixture conditions were sufficiently rich to produce soot. Results indicate that the JSC allows study of soot formation in an aerodynamic situation revelant to gas turbines.

  12. Pulsed atmospheric fluidized bed combustor apparatus

    DOEpatents

    Mansour, Momtaz N.

    1993-10-26

    A pulsed atmospheric fluidized bed reactor system is disclosed and claimed along with a process for utilization of same for the combustion of, e.g. high sulfur content coal. The system affords a economical, ecologically acceptable alternative to oil and gas fired combustors. The apparatus may also be employed for endothermic reaction, combustion of waste products, e.g., organic and medical waste, drying materials, heating air, calcining and the like.

  13. A Comparison of Combustor-Noise Models

    NASA Technical Reports Server (NTRS)

    Hultgren, Lennart, S.

    2012-01-01

    The current status of combustor-noise prediction in the NASA Aircraft Noise Prediction Program (ANOPP) for current-generation (N) turbofan engines is summarized. Best methods for near-term updates are reviewed. Long-term needs and challenges for the N+1 through N+3 timeframe are discussed. This work was carried out under the NASA Fundamental Aeronautics Program, Subsonic Fixed Wing Project, Quiet Aircraft Subproject.

  14. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Modular excess air 50 4 Refuse-derived fuel stoker 150 24 Bubbling fluidized bed combustor 100 4 Circulating fluidized bed combustor 100 4 Pulverized coal/refuse-derived fuel mixed fuel-fired combustor 150...

  15. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... Modular excess air 50 4 Refuse-derived fuel stoker 150 24 Bubbling fluidized bed combustor 100 4 Circulating fluidized bed combustor 100 4 Pulverized coal/refuse-derived fuel mixed fuel-fired combustor 150...

  16. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... Modular excess air 50 4 Refuse-derived fuel stoker 150 24 Bubbling fluidized bed combustor 100 4 Circulating fluidized bed combustor 100 4 Pulverized coal/refuse-derived fuel mixed fuel-fired combustor 150...

  17. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... Modular excess air 50 4 Refuse-derived fuel stoker 150 24 Bubbling fluidized bed combustor 100 4 Circulating fluidized bed combustor 100 4 Pulverized coal/refuse-derived fuel mixed fuel-fired combustor 150...

  18. Low NOx Fuel Flexible Combustor Integration Project Overview

    NASA Technical Reports Server (NTRS)

    Walton, Joanne C.; Chang, Clarence T.; Lee, Chi-Ming; Kramer, Stephen

    2015-01-01

    The Integrated Technology Demonstration (ITD) 40A Low NOx Fuel Flexible Combustor Integration development is being conducted as part of the NASA Environmentally Responsible Aviation (ERA) Project. Phase 2 of this effort began in 2012 and will end in 2015. This document describes the ERA goals, how the fuel flexible combustor integration development fulfills the ERA combustor goals, and outlines the work to be conducted during project execution.

  19. High-temperature durability considerations for HSCT combustor

    NASA Technical Reports Server (NTRS)

    Jacobson, Nathan S.

    1992-01-01

    The novel combustor designs for the High Speed Civil Transport will require high temperature materials with long term environmental stability. Higher liner temperatures than in conventional combustors and the need for reduced weight necessitates the use of advanced ceramic matrix composites. The combustor environment is defined at the current state of design, the major degradation routes are discussed for each candidate ceramic material, and where possible, the maximum use temperatures are defined for these candidate ceramics.

  20. Advanced liner-cooling techniques for gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Riddlebaugh, S. M.

    1985-01-01

    Component research for advanced small gas turbine engines is currently underway at the NASA Lewis Research Center. As part of this program, a basic reverse-flow combustor geometry was being maintained while different advanced liner wall cooling techniques were investigated. Performance and liner cooling effectiveness of the experimental combustor configuration featuring counter-flow film-cooled panels is presented and compared with two previously reported combustors featuring: splash film-cooled liner walls; and transpiration cooled liner walls (Lamilloy).

  1. Rapid-quench axially staged combustor

    DOEpatents

    Feitelberg, Alan S.; Schmidt, Mark Christopher; Goebel, Steven George

    1999-01-01

    A combustor cooperating with a compressor in driving a gas turbine includes a cylindrical outer combustor casing. A combustion liner, having an upstream rich section, a quench section and a downstream lean section, is disposed within the outer combustor casing defining a combustion chamber having at least a core quench region and an outer quench region. A first plurality of quench holes are disposed within the liner at the quench section having a first diameter to provide cooling jet penetration to the core region of the quench section of the combustion chamber. A second plurality of quench holes are disposed within the liner at the quench section having a second diameter to provide cooling jet penetration to the outer region of the quench section of the combustion chamber. In an alternative embodiment, the combustion chamber quench section further includes at least one middle region and at least a third plurality of quench holes disposed within the liner at the quench section having a third diameter to provide cooling jet penetration to at least one middle region of the quench section of the combustion chamber.

  2. Low Emissions RQL Flametube Combustor Test Results

    NASA Technical Reports Server (NTRS)

    Chang, Clarence T.; Holdeman, James D.

    2001-01-01

    The overall objective of this test program was to demonstrate and evaluate the capability of the Rich-burn/Quick-mix/Lean-burn (RQL) combustor concept for HSR applications. This test program was in support of the Pratt & Whitney and GE Aircraft Engines HSR low-NOx Combustor Program. Collaborative programs with Parker Hannifin Corporation and Textron Fuel Systems resulted in the development and testing of the high-flow low-NOx rich-burn zone fuel-to-air ratio research fuel nozzles used in this test program. Based on the results obtained in this test program, several conclusions can be made: (1) The RQL tests gave low NOx and CO emissions results at conditions corresponding to HSR cruise. (2) The Textron fuel nozzle design with optimal multiple partitioning of fuel and air circuits shows potential of providing an acceptable uniform local fuel-rich region in the rich burner. (3) For the parameters studied in this test series, the tests have shown T3 is the dominant factor in the NOx formation for RQL combustors. As T3 increases from 600 to 1100 F, EI(NOx) increases approximately three fold. (4) Factors which appear to have secondary influence on NOx formation are P4, T4, infinity(sub rb), V(sub ref,ov). (5) Low smoke numbers were measured for infinity(sub rb) of 2.0 at P4 of 120 psia.

  3. Low pollution combustor designs for CTOL engines - Results of the Experimental Clean Combustor Program

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Peduzzi, A.; Niedzwiecki, R. W.

    1976-01-01

    The NASA/Pratt & Whitney Aircraft Experimental Clean Combustor Program is a multi-year, major contract effort. Primary program objectives are the generation of combustor technology for development of advanced commercial CTOL engines with lower exhaust emissions than current aircraft and demonstration of this technology in a full-scale JT9D engine in 1976. This paper describes the pollution and performance goals, Phase I and II test results, and the Phase III combustor hardware, pollution sampling techniques, and test plans. Best results were obtained with the Vorbix concept which employs multiple burning zones and improved fuel preparation and distribution. Substantial reductions were achieved in all pollutant categories, meeting the 1979 EPA standards for NOx, THC, and smoke when extrapolated to JT9D cycle conditions. The Vorbix concept additionally demonstrated the capability for acceptable altitude relight and did not appear to have unsolvable durability or exit temperature distribution problems.

  4. Numerical analysis of the flows in annular slinger combustors

    NASA Astrophysics Data System (ADS)

    Huebner, S.; Exley, T.

    1990-07-01

    Improved gas-turbine combustor design techniques are developed through the application of CFD flow predictions. The conservation equations of mass, momentum, and energy are solved using the finite-volume approach of Spalding. The geometry is a three-dimensional region of cyclic symmetry for a selected annular slinger combustor of reasonable performance. The flow is assumed nonreacting, isothermal, and turbulent. Mixing of the dilution jet stream with the bulk combustor flow is simulated by assuming different inlet temperatures for the two mass sources and noting the temperature profile at the combustor exit plane. A flow visualization experiment is performed on cold flow conditions and reasonably corroborates the CFD predictions.

  5. Apparatus and method for cooling a combustor cap

    DOEpatents

    Zuo, Baifang; Washam, Roy Marshall; Wu, Chunyang

    2014-04-29

    A combustor includes an end cap having a perforated downstream plate and a combustion chamber downstream of the downstream plate. A plenum is in fluid communication with the downstream plate and supplies a cooling medium to the combustion chamber through the perforations in the downstream plate. A method for cooling a combustor includes flowing a cooling medium into a combustor end cap and impinging the cooling medium on a downstream plate in the combustor end cap. The method further includes flowing the cooling medium into a combustion chamber through perforations in the downstream plate.

  6. Controlled pilot oxidizer for a gas turbine combustor

    DOEpatents

    Laster, Walter R.; Bandaru, Ramarao V.

    2010-07-13

    A combustor (22) for a gas turbine (10) includes a main burner oxidizer flow path (34) delivering a first portion (32) of an oxidizer flow (e.g., 16) to a main burner (28) of the combustor and a pilot oxidizer flow path (38) delivering a second portion (36) of the oxidizer flow to a pilot (30) of the combustor. The combustor also includes a flow controller (42) disposed in the pilot oxidizer flow path for controlling an amount of the second portion delivered to the pilot.

  7. Predicting Noise From Aircraft Turbine-Engine Combustors

    NASA Technical Reports Server (NTRS)

    Gliebe, P.; Mani, R.; Salamah, S.; Coffin, R.; Mehta, Jayesh

    2005-01-01

    COMBUSTOR and CNOISE are computer codes that predict far-field noise that originates in the combustors of modern aircraft turbine engines -- especially modern, low-gaseous-emission engines, the combustors of which sometimes generate several decibels more noise than do the combustors of older turbine engines. COMBUSTOR implements an empirical model of combustor noise derived from correlations between engine-noise data and operational and geometric parameters, and was developed from databases of measurements of acoustic emissions of engines. CNOISE implements an analytical and computational model of the propagation of combustor temperature fluctuations (hot spots) through downstream turbine stages. Such hot spots are known to give rise to far-field noise. CNOISE is expected to be helpful in determining why low-emission combustors are sometimes noisier than older ones, to provide guidance for refining the empirical correlation model embodied in the COMBUSTOR code, and to provide insight on how to vary downstream turbinestage geometry to reduce the contribution of hot spots to far-field noise.

  8. Combustor nozzle for a fuel-flexible combustion system

    DOEpatents

    Haynes, Joel Meier; Mosbacher, David Matthew; Janssen, Jonathan Sebastian; Iyer, Venkatraman Ananthakrishnan

    2011-03-22

    A combustor nozzle is provided. The combustor nozzle includes a first fuel system configured to introduce a syngas fuel into a combustion chamber to enable lean premixed combustion within the combustion chamber and a second fuel system configured to introduce the syngas fuel, or a hydrocarbon fuel, or diluents, or combinations thereof into the combustion chamber to enable diffusion combustion within the combustion chamber.

  9. Noise addendum experimental clean combustor program, phase 1

    NASA Technical Reports Server (NTRS)

    Sofrin, T. G.; Ross, D. A.

    1975-01-01

    The development of advanced CTOL aircraft engines with reduced exhaust emissions is discussed. Combustor noise information provided during the basic emissions program and used to advantage in securing reduced levels of combustion noise is included. Results are presented of internal pressure transducer measurements made during the scheduled emissions test program on ten configurations involving variations of three basic combustor designs.

  10. Combustor technology for future small gas turbine aircraft

    NASA Technical Reports Server (NTRS)

    Lyons, Valerie J.; Niedzwiecki, Richard W.

    1994-01-01

    To enhance fuel efficiency, future advanced small gas turbine engines will utilize engine cycles calling for overall engine pressure ratios, leading to higher combustor inlet pressures and temperatures. Further, the temperature rise through the combustor and the corresponding exit temperature are also expected to increase. This report describes future combustor technology needs for small gas turbine engines. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is anticipated in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors. Due to combustor size considerations, staged combustion is more easily accommodated in large engines. The inclusion of staged combustion in small engines will pose greater combustor design challenges.

  11. Preliminary Investigation of Combustion of Diborane in a Turbojet Combustor

    NASA Technical Reports Server (NTRS)

    Kaufman, Warner B; Gibbs, James B; Branstetter, J Robert

    1957-01-01

    Boron and its hydrides offer increased flight range relative to conventional fuels for turbojet engines. Preliminary evaluation has been made of the combustion characteristics and deposition problems resulting from burning diborone in a single, modified J33 combustor. A combustor relatively free of deposits for the limited test conditions has been developed. Three possible methods of alleviating deposits on the turbine blades are reported.

  12. Serial cooling of a combustor for a gas turbine engine

    DOEpatents

    Abreu, Mario E.; Kielczyk, Janusz J.

    2001-01-01

    A combustor for a gas turbine engine uses compressed air to cool a combustor liner and uses at least a portion of the same compressed air for combustion air. A flow diverting mechanism regulates compressed air flow entering a combustion air plenum feeding combustion air to a plurality of fuel nozzles. The flow diverting mechanism adjusts combustion air according to engine loading.

  13. Critical Propulsion Components. Volume 2; Combustor

    NASA Technical Reports Server (NTRS)

    2005-01-01

    Several studies have concluded that a supersonic aircraft, if environmentally acceptable and economically viable, could successfully compete in the 21st century marketplace. However, before industry can commit to what is estimated as a 15 to 20 billion dollar investment, several barrier issues must be resolved. In an effort to address these barrier issues, NASA and Industry teamed to form the High-Speed Research (HSR) program. As part of this program, the Critical Propulsion Components (CPC) element was created and assigned the task of developing those propulsion component technologies necessary to: (1) reduce cruise emissions by a factor of 10 and (2) meet the ever-increasing airport noise restrictions with an economically viable propulsion system. The CPC-identified critical components were ultra-low emission combustors, low-noise/high-performance exhaust nozzles, low-noise fans, and stable/high-performance inlets. Propulsion cycle studies (coordinated with NASA Langley Research Center sponsored airplane studies) were conducted throughout this CPC program to help evaluate candidate components and select the best concepts for the more complex and larger scale research efforts. The propulsion cycle and components ultimately selected were a mixed-flow turbofan (MFTF) engine employing a lean, premixed, prevaporized (LPP) combustor coupled to a two-dimensional mixed compression inlet and a two-dimensional mixer/ejector nozzle. Due to the large amount of material presented in this report, it was prepared in four volumes; Volume 1: Summary, Introduction, and Team. Propulsion System Studies, Volume 2: Combustor, Volume 3: Exhaust Nozzle, and Volume 4: Inlet and Fan/Inlet Acoustic Team.

  14. Pulse Combustor Design, A DOE Assessment

    SciTech Connect

    National Energy Technology Laboratory

    2003-07-31

    The goal of the U.S. Department of Energy's (DOE) Clean Coal Technology (CCT) program is to furnish the energy marketplace with a number of advanced, more efficient, and environmentally responsible coal utilization technologies through demonstration projects. These projects seek to establish the commercial feasibility of the most promising advanced coal technologies that have developed beyond the proof-of-concept stage. This document serves as a DOE post-project assessment (PPA) of a project selected in CCT Round IV, the Pulse Combustor Design Qualification Test, as described in a Report to Congress (U.S. Department of Energy 1992). Pulse combustion is a method intended to increase the heat-transfer rate in a fired heater. The desire to demonstrate the use of pulse combustion as a source of heat for the gasification of coal, thus avoiding the need for an oxygen plant, prompted ThermoChem, Inc. (TCI), to submit a proposal for this project. In October 1992, TCI entered into a cooperative agreement with DOE to conduct this project. In 1998, the project was restructured and scaled down, and in September 1998, a new cooperative agreement was signed. The site of the revised project was TCI's facilities in Baltimore, Maryland. The original purpose of this CCT project was to demonstrate a unit that would employ ten identical 253-resonance tube combustors in a coal gasification unit. The objective of the scaled-down project was to test a single 253-resonance-tube combustor in a fluidized sand bed, with gasification being studied in a process development unit (PDU). DOE provided 50 percent of the total project funding of $8.6 million. The design for the demonstration unit was completed in February 1999, and construction was completed in November 2000. Operations were conducted in March 2001.

  15. COMPUTATIONAL FLUID DYNAMICS MODELING ANALYSIS OF COMBUSTORS

    SciTech Connect

    Mathur, M.P.; Freeman, Mark; Gera, Dinesh

    2001-11-06

    In the current fiscal year FY01, several CFD simulations were conducted to investigate the effects of moisture in biomass/coal, particle injection locations, and flow parameters on carbon burnout and NO{sub x} inside a 150 MW GEEZER industrial boiler. Various simulations were designed to predict the suitability of biomass cofiring in coal combustors, and to explore the possibility of using biomass as a reburning fuel to reduce NO{sub x}. Some additional CFD simulations were also conducted on CERF combustor to examine the combustion characteristics of pulverized coal in enriched O{sub 2}/CO{sub 2} environments. Most of the CFD models available in the literature treat particles to be point masses with uniform temperature inside the particles. This isothermal condition may not be suitable for larger biomass particles. To this end, a stand alone program was developed from the first principles to account for heat conduction from the surface of the particle to its center. It is envisaged that the recently developed non-isothermal stand alone module will be integrated with the Fluent solver during next fiscal year to accurately predict the carbon burnout from larger biomass particles. Anisotropy in heat transfer in radial and axial will be explored using different conductivities in radial and axial directions. The above models will be validated/tested on various fullscale industrial boilers. The current NO{sub x} modules will be modified to account for local CH, CH{sub 2}, and CH{sub 3} radicals chemistry, currently it is based on global chemistry. It may also be worth exploring the effect of enriched O{sub 2}/CO{sub 2} environment on carbon burnout and NO{sub x} concentration. The research objective of this study is to develop a 3-Dimensional Combustor Model for Biomass Co-firing and reburning applications using the Fluent Computational Fluid Dynamics Code.

  16. Coal desulfurization in a rotary kiln combustor

    SciTech Connect

    Cobb, J.T. Jr.

    1992-09-11

    The purpose of this project was to demonstrate the combustion of coal and coal wastes in a rotary kiln reactor with limestone addition for sulfur control. The rationale for the project was the perception that rotary systems could bring several advantages to combustion of these fuels, and may thus offer an alternative to fluid-bed boilers. Towards this end, an existing wood pyrolysis kiln (the Humphrey Charcoal kiln) was to be suitably refurbished and retrofitted with a specially designed version of a patented air distributor provided by Universal Energy, Inc. (UEI). As the project progressed beyond the initial stages, a number of issues were raised regarding the feasibility and the possible advantages of burning coals in a rotary kiln combustor and, in particular, the suitability of the Humphrey Charcoal kiln as a combustor. Instead, an opportunity arose to conduct combustion tests in the PEDCO Rotary Cascading-Bed Boiler (RCBB) commercial demonstration unit at the North American Rayon CO. (NARCO) in Elizabethton, TN. The tests focused on anthracite culm and had two objectives: (a) determine the feasibility of burning anthracite culms in a rotary kiln boiler and (b) obtain input for any further work involving the Humphrey Charcoal kiln combustor. A number of tests were conducted at the PEDCO unit. The last one was conducted on anthracite culm procured directly from the feed bin of a commercial circulating fluid-bed boiler. The results were disappointing; it was difficult to maintain sustained combustion even when large quantities of supplemental fuel were used. Combustion efficiency was poor, around 60 percent. The results suggest that the rotary kiln boiler, as designed, is ill-suited with respect to low-grade, hard to burn solid fuels, such as anthracite culm. Indeed, data from combustion of bituminous coal in the PEDCO unit suggest that with respect to coal in general, the rotary kiln boiler appears inferior to the circulating fluid bed boiler.

  17. Active Control of Combustor Instability Shown to Help Lower Emissions

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Chang, Clarence T.

    2002-01-01

    In a quest to reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities, or high-pressure oscillations much like sound waves, that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Aerospace Propulsion and Power Base Research and Technology Program, the NASA Glenn Research Center, in partnership with Pratt & Whitney and United Technologies Research Center, is developing technologies for the active control of combustion instabilities. With active combustion control, the fuel is pulsed to put pressure oscillations into the system. This cancels out the pressure oscillations being produced by the instabilities. Thus, the engine can have lower pollutant emissions and long life.The use of active combustion instability control to reduce thermo-acoustic-driven combustor pressure oscillations was demonstrated on a single-nozzle combustor rig at United Technologies. This rig has many of the complexities of a real engine combustor (i.e., an actual fuel nozzle and swirler, dilution cooling, etc.). Control was demonstrated through modeling, developing, and testing a fuel-delivery system able to the 280-Hz instability frequency. The preceding figure shows the capability of this system to provide high-frequency fuel modulations. Because of the high-shear contrarotating airflow in the fuel injector, there was some concern that the fuel pulses would be attenuated to the point where they would

  18. Ultra Low NOx Catalytic Combustion for IGCC Power Plants

    SciTech Connect

    Shahrokh Etemad; Benjamin Baird; Sandeep Alavandi; William Pfefferle

    2008-03-31

    In order to meet DOE's goals of developing low-emissions coal-based power systems, PCI has further developed and adapted it's Rich-Catalytic Lean-burn (RCL{reg_sign}) catalytic reactor to a combustion system operating on syngas as a fuel. The technology offers ultra-low emissions without the cost of exhaust after-treatment, with high efficiency (avoidance of after-treatment losses and reduced diluent requirements), and with catalytically stabilized combustion which extends the lower Btu limit for syngas operation. Tests were performed in PCI's sub-scale high-pressure (10 atm) test rig, using a two-stage (catalytic then gas-phase) combustion process for syngas fuel. In this process, the first stage consists of a fuel-rich mixture reacting on a catalyst with final and excess combustion air used to cool the catalyst. The second stage is a gas-phase combustor, where the air used for cooling the catalyst mixes with the catalytic reactor effluent to provide for final gas-phase burnout and dilution to fuel-lean combustion products. During testing, operating with a simulated Tampa Electric's Polk Power Station syngas, the NOx emissions program goal of less than 0.03 lbs/MMBtu (6 ppm at 15% O{sub 2}) was met. NOx emissions were generally near 0.01 lbs/MMBtu (2 ppm at 15% O{sub 2}) (PCI's target) over a range on engine firing temperatures. In addition, low emissions were shown for alternative fuels including high hydrogen content refinery fuel gas and low BTU content Blast Furnace Gas (BFG). For the refinery fuel gas increased resistance to combustor flashback was achieved through preferential consumption of hydrogen in the catalytic bed. In the case of BFG, stable combustion for fuels as low as 88 BTU/ft{sup 3} was established and maintained without the need for using co-firing. This was achieved based on the upstream catalytic reaction delivering a hotter (and thus more reactive) product to the flame zone. The PCI catalytic reactor was also shown to be active in ammonia

  19. Numerical investigation of recirculation in the UTSI MHD combustor

    SciTech Connect

    Schulz, R.J.; Lee, J.J.; Giel, T.V. Jr.

    1983-09-01

    Numerical studies were carried out to investigate the gross structure of flow in cylindrical combustors. The combustor configurations studied are variations of a working design used at the University of Tennessee Space Institute to burn pulverized coal at temperatures in excess of 3000K for generation of a plasma feeding a magnetohydrodynamic channel. The numerical studies were conducted for an isothermal fluid; the main objective of the calculations was to study the effect of the oxidant injection pattern on the gross structure of recirculating flows within the combustor. The calculations illustrate the basic features of the flow in combustors of this type and suggest implications for the injection of coal and oxidizer in this type of combustor.

  20. Experimental evaluation of combustor concepts for burning broad property fuels

    NASA Technical Reports Server (NTRS)

    Kasper, J. M.; Ekstedt, E. E.; Dodds, W. J.; Shayeson, M. W.

    1980-01-01

    A baseline CF6-50 combustor and three advanced combustor designs were evaluated to determine the effects of combustor design on operational characteristics using broad property fuels. Three fuels were used in each test: Jet A, a broad property 13% hydrogen fuel, and a 12% hydrogen fuel blend. Testing was performed in a sector rig at true cruise and simulated takeoff conditions for the CF6-50 engine cycle. The advanced combustors (all double annular, lean dome designs) generally exhibited lower metal temperatures, exhaust emissions, and carbon buildup than the baseline CF6-50 combustor. The sensitivities of emissions and metal temperatures to fuel hydrogen content were also generally lower for the advanced designs. The most promising advanced design used premixing tubes in the main stage. This design was chosen for additional testing in which fuel/air ratio, reference velocity, and fuel flow split were varied.

  1. Parameters controlling nitric oxide emissions from gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Heywood, J. B.; Mikus, T.

    1973-01-01

    Nitric oxide forms in the primary zone of gas turbine combustors where the burnt gas composition is close to stoichiometric and gas temperatures are highest. It was found that combustor air inlet conditions, mean primary zone fuel-air ratio, residence time, and the uniformity of the primary zone are the most important variables affecting nitric oxide emissions. Relatively simple models of the flow in a gas turbine combustor, coupled with a rate equation for nitric oxide formation via the Zeldovich mechanism are shown to correlate the variation in measured NOx emissions. Data from a number of different combustor concepts are analyzed and shown to be in reasonable agreement with predictions. The NOx formation model is used to assess the extent to which an advanced combustor concept, the NASA swirl can, has produced a lean well-mixed primary zone generally believed to be the best low NOx emissions burner type.

  2. Thermal and emission characteristics of a CAN combustor

    NASA Astrophysics Data System (ADS)

    Shah, Rupesh D.; Banerjee, Jyotirmay

    2016-03-01

    Experimental investigations are carried out to establish the thermal and emission characteristics of a CAN combustor. Temperature and emission levels at the combustor exit are measured for different swirler vane angles and air fuel ratios (AFR). Swirler vane angle is varied from 15° to 60° in steps of 15°. AFR is varied in the range of 41-51. Experimental analysis is carried out using methane as fuel. Measured temperature variation at combustor outlet indicates that the hot product of combustor flows near the liner wall. Gradient of temperature near the wall decreases as the swirler vane angle (and corresponding swirl number) is increased. The peak temperature reduces at higher value of AFR. Emission level of carbon monoxide decreases with increase in AFR and swirler vane orientation. A higher level of NOX emission is observed for AFR of 45. This is due to change in shape and strength of the recirculation region in the primary zone of the combustor.

  3. Combustor air flow control method for fuel cell apparatus

    DOEpatents

    Clingerman, Bruce J.; Mowery, Kenneth D.; Ripley, Eugene V.

    2001-01-01

    A method for controlling the heat output of a combustor in a fuel cell apparatus to a fuel processor where the combustor has dual air inlet streams including atmospheric air and fuel cell cathode effluent containing oxygen depleted air. In all operating modes, an enthalpy balance is provided by regulating the quantity of the air flow stream to the combustor to support fuel cell processor heat requirements. A control provides a quick fast forward change in an air valve orifice cross section in response to a calculated predetermined air flow, the molar constituents of the air stream to the combustor, the pressure drop across the air valve, and a look up table of the orifice cross sectional area and valve steps. A feedback loop fine tunes any error between the measured air flow to the combustor and the predetermined air flow.

  4. Energy efficient engine sector combustor rig test program

    NASA Technical Reports Server (NTRS)

    Dubiel, D. J.; Greene, W.; Sundt, C. V.; Tanrikut, S.; Zeisser, M. H.

    1981-01-01

    Under the NASA-sponsored Energy Efficient Engine program, Pratt & Whitney Aircraft has successfully completed a comprehensive combustor rig test using a 90-degree sector of an advanced two-stage combustor with a segmented liner. Initial testing utilized a combustor with a conventional louvered liner and demonstrated that the Energy Efficient Engine two-stage combustor configuration is a viable system for controlling exhaust emissions, with the capability to meet all aerothermal performance goals. Goals for both carbon monoxide and unburned hydrocarbons were surpassed and the goal for oxides of nitrogen was closely approached. In another series of tests, an advanced segmented liner configuration with a unique counter-parallel FINWALL cooling system was evaluated at engine sea level takeoff pressure and temperature levels. These tests verified the structural integrity of this liner design. Overall, the results from the program have provided a high level of confidence to proceed with the scheduled Combustor Component Rig Test Program.

  5. Parametric test results of a swirl-can combustor

    NASA Technical Reports Server (NTRS)

    Niedzwiecki, R. W.; Jones, R. E.

    1973-01-01

    Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for three models of an experimental, annular swirl can combustor. The combustor was 1.067 meters in outer diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 589, 756 and 839 K, inlet pressures of 3 to 6.4 atmospheres, reference velocities of 21 to 38 meters per second and combustor equivalence ratios, based on total combustor flows of 0.206 to 1.028. Maximum oxides of nitrogen emission index values occurred at an equivalence ratio of 0.7 with lower values measured for both higher and lower equivalence ratios. Oxides of nitrogen concentrations, to the 0.7 level with 756 K inlet air, were correlated for the three models by a combined parameter consisting of measured flow and geometric parameters. Effects of the individual parameters comprising the correlation are also presented.

  6. Radial inlet guide vanes for a combustor

    SciTech Connect

    Zuo, Baifang; Simons, Derrick; York, William; Ziminsky, Willy S

    2013-02-12

    A combustor may include an interior flow path therethrough, a number of fuel nozzles in communication with the interior flow path, and an inlet guide vane system positioned about the interior flow path to create a swirled flow therein. The inlet guide vane system may include a number of windows positioned circumferentially around the fuel nozzles. The inlet guide vane system may also include a number of inlet guide vanes positioned circumferentially around the fuel nozzles and adjacent to the windows to create a swirled flow within the interior flow path.

  7. Lean stability augmentation for premixing, prevaporizing combustors

    NASA Technical Reports Server (NTRS)

    Mcvey, J. B.; Kennedy, J. B.

    1979-01-01

    An experimental program was conducted to investigate techniques for improving the lean combustion limits of premixing, prevaporizing combustors applicable to gas turbine engine main burners. Augmented flameholders employing recessed perforated plates, catalyzed tube bundles, and configurations in which pilot fuel was injected into the wakes of V-gutters or perforated plates were designed and tested. Stable operation of the piloted designs was achieved at equivalence ratios as low as 0.25; NOx emissions of less than 1.0 g/kg at simulated turbine engine cruise conditions were obtained. A piloted perforated plate employing four percent pilot fuel flow produced the best performance while meeting severe NOx constraints.

  8. Testing and Characterization of CMC Combustor Liners

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Verrilli, Michael J.

    2003-01-01

    Multiple combustor liner applications, both segmented and fully annular designs, have been configured for exposure in NASA's High Pressure Burner Rig (HPBR). The segmented liners were attached to the rig structure with SiC/SiC fasteners and exposed to simulated gas turbine conditions for nearly 200 hours. Test conditions included pressures of 6 atm., gas velocity of 42 m/s, and gas temperatures near 1450 C. The temperatures of both the cooled and combustion flow sides of the liners were measured using optical and contact measurement techniques. Minor weight loss was observed, but the liners remained structural sound, although damage was noted in some fasteners.

  9. Active Suppression of Instabilities in Engine Combustors

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2004-01-01

    A method of feedback control has been proposed as a means of suppressing thermo-acoustic instabilities in a liquid- fueled combustor of a type used in an aircraft engine. The basic principle of the method is one of (1) sensing combustor pressure oscillations associated with instabilities and (2) modulating the rate of flow of fuel to the combustor with a control phase that is chosen adaptively so that the pressure oscillations caused by the modulation oppose the sensed pressure oscillations. The need for this method arises because of the planned introduction of advanced, lean-burning aircraft gas turbine engines, which promise to operate with higher efficiencies and to emit smaller quantities of nitrogen oxides, relative to those of present aircraft engines. Unfortunately, the advanced engines are more susceptible to thermoacoustic instabilities. These instabilities are hard to control because they include large dead-time phase shifts, wide-band noise characterized by amplitudes that are large relative to those of the instabilities, exponential growth of the instabilities, random net phase walks, and amplitude fluctuations. In this method (see figure), the output of a combustion-pressure sensor would be wide-band-pass filtered and then further processed to generate a control signal that would be applied to a fast-actuation valve to modulate the flow of fuel. Initially, the controller would rapidly take large phase steps in order to home in, within a fraction of a second, to a favorable phase region within which the instability would be reduced. Then the controller would restrict itself to operate within this phase region and would further restrict itself to operate within a region of stability, as long as the power in the instability signal was decreasing. In the phase-shifting scheme of this method, the phase of the control vector would be made to continuously bounce back and forth from one boundary of an effective stability region to the other. Computationally

  10. Rapid Deployment of Rich Catalytic Combustion

    SciTech Connect

    Richard S. Tuthill

    2004-06-10

    The overall objective of this research under the Turbines Program is the deployment of fuel flexible rich catalytic combustion technology into high-pressure ratio industrial gas turbines. The resulting combustion systems will provide fuel flexibility for gas turbines to burn coal derived synthesis gas or natural gas and achieve NO{sub x} emissions of 2 ppmvd or less (at 15 percent O{sub 2}), cost effectively. This advance will signify a major step towards environmentally friendly electric power generation and coal-based energy independence for the United States. Under Phase 1 of the Program, Pratt & Whitney (P&W) performed a system integration study of rich catalytic combustion in a small high-pressure ratio industrial gas turbine with a silo combustion system that is easily scalable to a larger multi-chamber gas turbine system. An implementation plan for this technology also was studied. The principal achievement of the Phase 1 effort was the sizing of the catalytic module in a manner which allowed a single reactor (rather than multiple reactors) to be used by the combustion system, a conclusion regarding the amount of air that should be allocated to the reaction zone to achieve low emissions, definition of a combustion staging strategy to achieve low emissions, and mechanical integration of a Ceramic Matrix Composite (CMC) combustor liner with the catalytic module.

  11. Experiment of rocket-ram annular combustor

    NASA Astrophysics Data System (ADS)

    Yatsuyanagi, Nobuyuki; Sakamoto, Hiroshi; Sato, Kazuo; Sasaki, Masaki; Ono, Fumiei

    Superiority in specific impulse of the double-nozzle type of rocket-ram combined engine over the ducted type of combined engine was shown by performance calculations. Then, a double-nozzle type of rocket-ram annular combustor with a total thrust of 5 kN was designed and experimentally tested with varying ratios of thrust produced by rocket and ram. With the combustor having different diverging half-angles, namely 10 deg 18 arcmin, and 6 deg 40 arcmin, thrust and pressure distribution along the common expansion nozzle were measured to investigate the effect of interaction of the two expansion gases on thrust. Enhancement of specific impulse was experimentally verified. That is, the specific impulse gained in rocket-ram parallel operations, the ratio of rocket thrust to ram thrust being 50 to 50, were found to be 190 percent of those in pure rocket operations. However, in the downstream region of the common nozzle, the flow might separate due to the generation of shock waves in either type of nozzle configuration.

  12. Combustor Computations for CO2-Neutral Aviation

    NASA Technical Reports Server (NTRS)

    Hendricks, Robert C.; Brankovic, Andreja; Ryder, Robert C.; Huber, Marcia

    2011-01-01

    Knowing the pure component C(sub p)(sup 0) or mixture C(sub p) (sup 0) as computed by a flexible code such as NIST-STRAPP or McBride-Gordon, one can, within reasonable accuracy, determine the thermophysical properties necessary to predict the combustion characteristics when there are no tabulated or computed data for those fluid mixtures 3or limited results for lower temperatures. (Note: C(sub p) (sup 0) is molar heat capacity at constant pressure.) The method can be used in the determination of synthetic and biological fuels and blends using the NIST code to compute the C(sub p) (sup 0) of the mixture. In this work, the values of the heat capacity were set at zero pressure, which provided the basis for integration to determine the required combustor properties from the injector to the combustor exit plane. The McBride-Gordon code was used to determine the heat capacity at zero pressure over a wide range of temperatures (room to 6,000 K). The selected fluids were Jet-A, 224TMP (octane), and C12. It was found that each heat capacity loci were form-similar. It was then determined that the results [near 400 to 3,000 K] could be represented to within acceptable engineering accuracy with the simplified equation C(sub p) (sup 0) = A/T + B, where A and B are fluid-dependent constants and T is temperature (K).

  13. Error Reduction Program. [combustor performance evaluation codes

    NASA Technical Reports Server (NTRS)

    Syed, S. A.; Chiappetta, L. M.; Gosman, A. D.

    1985-01-01

    The details of a study to select, incorporate and evaluate the best available finite difference scheme to reduce numerical error in combustor performance evaluation codes are described. The combustor performance computer programs chosen were the two dimensional and three dimensional versions of Pratt & Whitney's TEACH code. The criteria used to select schemes required that the difference equations mirror the properties of the governing differential equation, be more accurate than the current hybrid difference scheme, be stable and economical, be compatible with TEACH codes, use only modest amounts of additional storage, and be relatively simple. The methods of assessment used in the selection process consisted of examination of the difference equation, evaluation of the properties of the coefficient matrix, Taylor series analysis, and performance on model problems. Five schemes from the literature and three schemes developed during the course of the study were evaluated. This effort resulted in the incorporation of a scheme in 3D-TEACH which is usuallly more accurate than the hybrid differencing method and never less accurate.

  14. Coal desulfurization in a rotary kiln combustor

    SciTech Connect

    Cobb, J.T. Jr.

    1990-08-15

    BCR National Laboratory (BCRNL) has initiated a project aimed at evaluating the technical and economic feasibility of using a rotary kiln, suitably modified, to burn Pennsylvania anthracite wastes, co-fired with high-sulfur bituminous coal. Limestone will be injected into the kiln for sulfur control, to determine whether high sulfur capture levels can be achieved with high sorbent utilization. The principal objectives of this work are: (1) to prove the feasibility of burning anthracite refuse, with co-firing of high-sulfur bituminous coal and with limestone injection for sulfur emissions control, in a rotary kiln fitted with a Universal Energy International (UEI) air injector system; (2) to determine the emissions levels of SO{sub x} and NO{sub x} and specifically to identify the Ca/S ratios that are required to meet New Source Performance Standards; (3) to evaluate the technical and economic merits of a commercial rotary kiln combustor in comparison to fluidized bed combustors; and, (4) to ascertain the need for further work, including additional combustion tests, prior to commercial application, and to recommend accordingly a detailed program towards this end.

  15. Azimuthally forced flames in an annular combustor

    NASA Astrophysics Data System (ADS)

    Worth, Nicholas; Dawson, James; Mastorakos, Epaminondas

    2015-11-01

    Thermoacoustic instabilities are more likely to occur in lean burn combustion systems, making their adoption both difficult and costly. At present, our knowledge of such phenomena is insufficient to produce an inherently stable combustor by design, and therefore an improved understanding of these instabilities has become the focus of a significant research effort. Recent experimental and numerical studies have demonstrated that the symmetry of annular chambers permit a range of self-excited azimuthal modes to be generated in annular geometry, which can make the study of isolated modes difficult. While acoustic forcing is common in single flame experiments, no equivalent for forced azimuthal modes in an annular chamber have been demonstrated. The present investigation focuses on the novel application of acoustic forcing to a laboratory scale annular combustor, in order to generate azimuthal standing wave modes at a prescribed frequency and amplitude. The results focus on the ability of the method to isolate the mode of oscillation using experimental pressure and high speed OH* measurements. The successful excitation of azimuthal modes demonstrated represents an important step towards improving our fundamental understanding of this phenomena in practically relevant geometry.

  16. Multi-fuel pre combustor unit

    SciTech Connect

    Paul, M.A.; Paul, A.

    1993-07-06

    A pre combustor unit is described for installation in a thermal engine having means for generating compressed air for mixing with a fuel for combustion in a main combustion chamber in the engine, the pre combustor unit comprising: a housing having means for installing the precombustion unit in the engine, the housing having an internal precombustion chamber with a discharge passage that communicates with the main combustion chamber of the engine when the unit is installed in the engine; a displaceable valve head in the housing, where the housing includes a valve seat at the discharge passage and the valve head is seatable on the valve seat to block discharge passage from communicating with the precombustion chamber; actuating means connected to the valve head for selectively displacing the valve head into seating engagement with the valve seat, wherein the discharge passage is closed, and for retracting the valve head from a seating engagement, wherein the discharge passage is open; a compressed air passage communicating with the precombustion chamber and having means for selectively communicating with the means of the thermal engine for generating compressed air, wherein compressed air from the engine is delivered to the precombustion chamber; and a fuel supply means communicating with the compressed air passage for supplying fuel to the compressed air passage wherein a fuel-air mixture is delivered to the precombustion chamber.

  17. Low NO.sub.x multistage combustor

    DOEpatents

    Becker, Frederick E.; Breault, Ronald W.; Litka, Anthony F.; McClaine, Andrew W.; Shukla, Kailash

    2000-01-01

    A high efficiency, Vortex Inertial Staged Air (VIStA) combustor provides ultra-low NO.sub.X production of about 20 ppmvd or less with CO emissions of less than 50 ppmvd, both at 3% O.sub.2. Prompt NO.sub.X production is reduced by partially reforming the fuel in a first combustion stage to CO and H.sub.2. This is achieved in the first stage by operating with a fuel rich mixture, and by recirculating partially oxidized combustion products, with control over stoichiometry, recirculation rate and residence time. Thermal NO.sub.X production is reduced in the first stage by reducing the occurrence of high temperature combustion gas regions. This is achieved by providing the first stage burner with a thoroughly pre-mixed fuel/oxidant composition, and by recirculating part of the combustion products to further mix the gases and provide a more uniform temperature in the first stage. In a second stage combustor thermal NO.sub.X production is controlled by inducing a large flow of flue gas recirculation in the second stage combustion zone to minimize the ultimate temperature of the flame. One or both of the first and second stage burners can be cooled to further reduce the combustion temperature and to improve the recirculation efficiency. Both of these factors tend to reduce production of NO.sub.X.

  18. Analytical fuel property effects: Small combustors

    NASA Technical Reports Server (NTRS)

    Cohen, J. D.

    1984-01-01

    The study performed in Phase 1 of this program applies only to a T700/CT7 engine family type combustor functioning in the engine as defined and does not necessarily apply to other cycles or combustors of differing stoichiometry. The study was not extended to any of the fuel delivery accessories such as pumps or control systems, nor was there any investigation of potential systems problems which might arise as a consequence of abnormal properties such as density which might affect delivery schedules or aromatics content which might affect fuel system seals. The T700/CT7 engine is a front drive turboshaft or turboprop engine in the 1500-1800 shp (1120-1340 kW) class as currently configured with highpower core flows of about 10 lb/sec (4.5 kg/sec). It employs a straight-through annular combustion system less than 5 in. (12.5 cm) in length utilizing a machined ring film cooled construction and twelve low-pressure air blast fuel injectors. Commercial and Naval versions employ two 0.5 Joule capacitive discharge surface gap ignitors.

  19. Ultra-Low NOx Advanced Vortex Combustor

    SciTech Connect

    Edmonds, R.G.; Steele, R.C.; Williams, J.T.; Straub, D.L.; Casleton, K.H.; Bining, Avtar

    2006-05-01

    An ultra lean-premixed Advanced Vortex Combustor (AVC) has been developed and tested. The natural gas fueled AVC was tested at the U.S. Department of Energy’s National Energy Technology Laboratory (USDOE NETL) test facility in Morgantown (WV). All testing was performed at elevated pressures and inlet temperatures and at lean fuel-air ratios representative of industrial gas turbines. The improved AVC design exhibited simultaneous NOx/CO/UHC emissions of 4/4/0 ppmv (all emissions are at 15% O2 dry). The design also achieved less than 3 ppmv NOx with combustion efficiencies in excess of 99.5%. The design demonstrated tremendous acoustic dynamic stability over a wide range of operating conditions which potentially makes this approach significantly more attractive than other lean premixed combustion approaches. In addition, a pressure drop of 1.75% was measured which is significantly lower than conventional gas turbine combustors. Potentially, this lower pressure drop characteristic of the AVC concept translates into overall gas turbine cycle efficiency improvements of up to one full percentage point. The relatively high velocities and low pressure drops achievable with this technology make the AVC approach an attractive alternative for syngas fuel applications.

  20. Combustor technology for future small gas turbine aircraft

    NASA Technical Reports Server (NTRS)

    Lyons, Valerie J.; Niedzwiecki, Richard W.

    1993-01-01

    Future engine cycles proposed for advanced small gas turbine engines will increase the severity of the operating conditions of the combustor. These cycles call for increased overall engine pressure ratios which increase combustor inlet pressure and temperature. Further, the temperature rise through the combustor and the corresponding exit temperature also increase. Future combustor technology needs for small gas turbine engines is described. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is expected in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors.

  1. Catalytic reactor

    DOEpatents

    Aaron, Timothy Mark; Shah, Minish Mahendra; Jibb, Richard John

    2009-03-10

    A catalytic reactor is provided with one or more reaction zones each formed of set(s) of reaction tubes containing a catalyst to promote chemical reaction within a feed stream. The reaction tubes are of helical configuration and are arranged in a substantially coaxial relationship to form a coil-like structure. Heat exchangers and steam generators can be formed by similar tube arrangements. In such manner, the reaction zone(s) and hence, the reactor is compact and the pressure drop through components is minimized. The resultant compact form has improved heat transfer characteristics and is far easier to thermally insulate than prior art compact reactor designs. Various chemical reactions are contemplated within such coil-like structures such that as steam methane reforming followed by water-gas shift. The coil-like structures can be housed within annular chambers of a cylindrical housing that also provide flow paths for various heat exchange fluids to heat and cool components.

  2. Multiple jet mixing flowfields in an isothermal model combustor

    NASA Astrophysics Data System (ADS)

    Ghosh, A.; Schulz, R. J.; Giel, T. V., Jr.

    1986-01-01

    The purpose of the present experimental investigation of confined, multiple turbulent jet-mixing with recirculation, in an axisymmetric duct that simulated a combustor, was the examination of flow fields that employ injector plates for the mixing of fuels and oxidizers. Quantitative descriptions of the velocity and turbulence fields were obtained with a vectorized, two-component laser Doppler velocimeter. The results obtained indicate that the annular slit injector jet generates a two-dimensional combustor flow that is in accord with theoretical studies, although rings of discrete injector jets create very complex, fully three-dimensional combustor flow fields.

  3. YF 102 in-duct combustor noise measurement, volume 1

    NASA Technical Reports Server (NTRS)

    Wilson, C. A.

    1977-01-01

    The combustion chamber from a YF 102 gas turbine engine was instrumented with semi-infinite acoustic wave guide probes and installed in a test rig to complement the combustor noise test. These combustor rig tests are described and the recorded data are listed. Internal dynamic pressure level measurements were made at the same locations and at the same operating conditions of the NASA YF 102 test. In addition, the combustor was operated at various off-designed points where one parameter at a time was varied. Background noise recordings were made to determine the magnitude of facility or test rig noise present.

  4. Multifuel evaluation of rich/quench/lean combustor

    NASA Technical Reports Server (NTRS)

    Notardonato, J. J.; Novick, A. S.; Troth, D. L.

    1982-01-01

    The fuel flexible combustor technology was developed for application to the Model 570-K industrial gas turbine engine. The technology, to achieve emission goals, emphasizes dry NOx reduction methods. Due to the high levels of fuel-bound nitrogen (FBN), control of NOx can be effected through a staged combustor with a rich initial combustion zone. A rich/quench/lean variable geometry combustor utilizes the technology presented to achieve low NOx from alternate fuels containing FBN. The results focus on emissions and durability for multifuel operation.

  5. Small gas-turbine combustor study: Fuel injector evaluation

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Riddlebaugh, S. M.

    1981-01-01

    As part of a continuing effort at the Lewis Research Center to improve performance, emissions, and reliability of turbine machinery, an investigation of fuel injection technique and effect of fuel type on small gas turbine combustors was undertaken. Performance and pollutant emission levels are documented over a range of simulated flight conditions for a reverse flow combustor configuration using simplex pressure-atomizing, spill-flow return, and splash cone airblast injectors. A parametric evaluation of the effect of increased combustor loading with each of the fuel injector types was obtained. Jet A and an experimental referee broad specification fuel were used to determine the effect of fuel type.

  6. Exhaust emissions of a double annular combustor: Parametric study

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1974-01-01

    A full scale double-annular ram-induction combustor designed for Mach 3.0 cruise operation was tested. Emissions of oxides of nitrogen, carbon monoxide, unburned hydrocarbons, and smoke were measured over a range of combustor operating variables including reference velocity, inlet air temperature and pressure, and exit average temperature. ASTM Jet-A fuel was used for these tests. An equation is provided relating oxides of nitrogen emissions as a function of the combustor, operating variables. A small effect of radial fuel staging on reducing exhaust emissions (which were originally quite low) is demonstrated.

  7. Design study of a gas turbine combustor with heat recirculation

    NASA Technical Reports Server (NTRS)

    Ganji, A.; Branch, M. C.; Oppenheim, A. K.

    1976-01-01

    A means of avoiding stoichiometric combustion, reducing emissions, and yet providing stable burning for lean mixtures is based on the use of heat recirculation rather than flow recirculation. This paper is concerned with the calculations of the design parameters of a gas turbine combustor with heat exchanger to produce the desired preheat temperature. The combustor inlet temperature, maximum temperature, equivalence ratio and recirculated heat are determined by thermodynamic analysis. The heat transfer analysis then provides the dimensions of the system to produce the predetermined boundary conditions. It is indicated that practical combustor design may be feasible for reactant mixtures as low as equivalence ratio 0.2.

  8. Achieving improved cycle efficiency via pressure gain combustors

    SciTech Connect

    Gemmen, R.S.; Janus, M.C.; Richards, G.A.; Norton, T.S.; Rogers, W.A.

    1995-04-01

    As part of the Department of Energy`s Advanced Gas Turbine Systems Program, an investigation is being performed to evaluate ``pressure gain`` combustion systems for gas turbine applications. This paper presents experimental pressure gain and pollutant emission data from such combustion systems. Numerical predictions for certain combustor geometries are also presented. It is reported that for suitable aerovalved pulse combustor geometries studied experimentally, an overall combustor pressure gain of nearly 1 percent can be achieved. It is also shown that for one combustion system operating under typical gas turbine conditions, NO{sub x} and CO emmissions, are about 30 ppmv and 8 ppmv, respectively.

  9. Systems and methods for detection of blowout precursors in combustors

    DOEpatents

    Lieuwen, Tim C.; Nair, Suraj

    2006-08-15

    The present invention comprises systems and methods for detecting flame blowout precursors in combustors. The blowout precursor detection system comprises a combustor, a pressure measuring device, and blowout precursor detection unit. A combustion controller may also be used to control combustor parameters. The methods of the present invention comprise receiving pressure data measured by an acoustic pressure measuring device, performing one or a combination of spectral analysis, statistical analysis, and wavelet analysis on received pressure data, and determining the existence of a blowout precursor based on such analyses. The spectral analysis, statistical analysis, and wavelet analysis further comprise their respective sub-methods to determine the existence of blowout precursors.

  10. Induction time effects in pulse combustors

    SciTech Connect

    Bell, J B; Marcus, D L; Pember, R B

    1999-04-09

    Combustion systems that take advantage of a periodic combustion process have many advantages over conventional systems. Their rate of heat transfer is greatly enhanced and their pollutant emissions are lower. They draw in their own supply of fuel and air and they are self-venting. They have few moving parts. The most common type of pulse combustor is based on a Helmholtz resonator - a burning cycle drives a resonant pressure wave, which in turn enhances the rate of combustion, resulting in a self-sustaining, large-scale oscillation. Although the basic physical mechanisms controlling such a process were explained by Rayleigh over a century ago, a full understanding of the operation of a pulse combustor still does not exist. The dominant processes in such a system--combustion, turbulent fluid dynamics, acoustics--are highly coupled and interact nonlinearly, which has reduced the design process to a costly and inefficient trial-and-error procedure. Several recent numerical and experimental studies, however, have been focused towards a better understanding of the basic underlying physics. Barr et al. [l] have elucidated the relative roles of the time scales governing the energy release, the turbulent mixing, and the acoustics. Keller et al. [5] have demonstrated the importance of the phase relation between the resonant pressure field in the tailpipe and the periodic energy release. Marcus et al. [6] have developed the capability for a fully three-dimensional simulation of the reacting flow in a pulse combustor. This paper is an application of that methodology to a detailed investigation of the frequency response of the model to changes in the chemical kinetics. The methodology consists of a fully conservative second-order Godunov algorithm for the inviscid, reacting gas dynamics equations coupled to an adaptive mesh refinement procedure[2]. The axisymmetric and three-dimensional simulations allow us to explore in detail the interaction between the transient fluid

  11. Advanced Low Emissions Subsonic Combustor Study

    NASA Technical Reports Server (NTRS)

    Smith, Reid

    1998-01-01

    Recent advances in commercial and military aircraft gas turbines have yielded significant improvements in fuel efficiency and thrust-to-weight ratio, due in large part to increased combustor operating pressures and temperatures. However, the higher operating conditions have increased the emission of oxides of nitrogen (NOx), which is a pollutant with adverse impact on the atmosphere and environment. Since commercial and military aircraft are the only important direct source of NOx emissions at high altitudes, there is a growing consensus that considerably more stringent limits on NOx emissions will be required in the future for all aircraft. In fact, the regulatory communities have recently agreed to reduce NOx limits by 20 percent from current requirements effective in 1996. Further reductions at low altitude, together with introduction of limits on NOx at altitude, are virtual certainties. In addition, the U.S. Government recently conducted hearings on the introduction of federal fees on the local emission of pollutants from all sources, including aircraft. While no action was taken regarding aircraft in this instance, the threat of future action clearly remains. In these times of intense and growing international competition, the U.S. le-ad in aerospace can only be maintained through a clear technological dominance that leads to a product line of maximum value to the global airline customer. Development of a very low NOx combustor will be essential to meet the future needs of both the commercial and military transport markets, if additional economic burdens and/or operational restrictions are to be avoided. In this report, Pratt & Whitney (P&W) presents the study results with the following specific objectives: Development of low-emissions combustor technologies for advances engines that will enter into service circa 2005, while producing a goal of 70 percent lower NOx emissions, compared to 1996 regulatory levels. Identification of solution approaches to

  12. Analysis of the effect on combustor noise measurements of acoustic waves reflected by the turbine and combustor inlet

    NASA Technical Reports Server (NTRS)

    Huff, R. G.

    1984-01-01

    Spectral analyses of static pressure fluctuations measured in turbine engine combustors at low engine speed show good agreement with theory. At idle speed the high pressure turbine is unchoked. Above idle speed the turbine chokes and a significant change in the shape of the measured combustor pressure spectrum is observed. A simplified theoretical model of the acoustic pressure generated in the combustor due to the turbulence-flame front interaction did not account for acoustic waves reflected from the turbine. By retaining this simplified combustion noise source model and adding a partial reflecting plane at the turbine and combustor inlet, a simple theoretical model was developed that reproduces the undulations in the combustor fluctuating pressure spectra. Plots of the theoretical combustor fluctuating pressure spectra are compared to the measured pressure spectra obtained from the CF6-50 turbofan engine over a range of engine operating speeds. The simplified combustion noise theory when modified by a simple turbine reflecting plane adequately accounts for the changes in measured combustor pressure spectra. It is further concluded that the shape of the pressure spectra downstream of the turbine, neglecting noise generated by the turbine itself, will be the combustion noise spectra unchanged except for the level reduction due to the energy blocked by the turbine.

  13. Analysis of the effect on combustor noise measurements of acoustic waves reflected by the turbine and combustor inlet

    NASA Astrophysics Data System (ADS)

    Huff, R. G.

    Spectral analyses of static pressure fluctuations measured in turbine engine combustors at low engine speed show good agreement with theory. At idle speed the high pressure turbine is unchoked. Above idle speed the turbine chokes and a significant change in the shape of the measured combustor pressure spectrum is observed. A simplified theoretical model of the acoustic pressure generated in the combustor due to the turbulence-flame front interaction did not account for acoustic waves reflected from the turbine. By retaining this simplified combustion noise source model and adding a partial reflecting plane at the turbine and combustor inlet, a simple theoretical model was developed that reproduces the undulations in the combustor fluctuating pressure spectra. Plots of the theoretical combustor fluctuating pressure spectra are compared to the measured pressure spectra obtained from the CF6-50 turbofan engine over a range of engine operating speeds. The simplified combustion noise theory when modified by a simple turbine reflecting plane adequately accounts for the changes in measured combustor pressure spectra. It is further concluded that the shape of the pressure spectra downstream of the turbine, neglecting noise generated by the turbine itself, will be the combustion noise spectra unchanged except for the level reduction due to the energy blocked by the turbine.

  14. Analysis of the effect on combustor noise measurements of acoustic waves reflected by the turbine and combustor inlet

    NASA Astrophysics Data System (ADS)

    Huff, R. G.

    1984-10-01

    Spectral analyses of static pressure fluctuations measured in turbine engine combustors at low engine speed show good agreement with theory. At idle speed the high pressure turbine is unchoked. Above idle speed the turbine chokes and a significant change in the shape of the measured combustor pressure spectrum is observed. A simplified theoretical model of the acoustic pressure generated in the combustor due to the turbulence-flame front interaction did not account for acoustic waves reflected from the turbine. By retaining this simplified combustion noise source model and adding a partial reflecting plane at the turbine and combustor inlet, a simple theoretical model was developed that reproduces the undulations in the combustor fluctuating pressure spectra. Plots of the theoretical combustor fluctuating pressure spectra are compared to the measured pressure spectra obtained from the CF6-50 turbofan engine over a range of engine operating speeds. The simplified combustion noise theory when modified by a simple turbine reflecting plane adequately accounts for the changes in measured combustor pressure spectra. It is further concluded that the shape of the pressure spectra downstream of the turbine, neglecting noise generated by the turbine itself, will be the combustion noise spectra unchanged except for the level reduction due to the energy blocked by the turbine.

  15. Liquid rocket combustor computer code development

    NASA Technical Reports Server (NTRS)

    Liang, P. Y.

    1985-01-01

    The Advanced Rocket Injector/Combustor Code (ARICC) that has been developed to model the complete chemical/fluid/thermal processes occurring inside rocket combustion chambers are highlighted. The code, derived from the CONCHAS-SPRAY code originally developed at Los Alamos National Laboratory incorporates powerful features such as the ability to model complex injector combustion chamber geometries, Lagrangian tracking of droplets, full chemical equilibrium and kinetic reactions for multiple species, a fractional volume of fluid (VOF) description of liquid jet injection in addition to the gaseous phase fluid dynamics, and turbulent mass, energy, and momentum transport. Atomization and droplet dynamic models from earlier generation codes are transplated into the present code. Currently, ARICC is specialized for liquid oxygen/hydrogen propellants, although other fuel/oxidizer pairs can be easily substituted.

  16. Low NOx heavy fuel combustor concept program

    NASA Technical Reports Server (NTRS)

    White, D. J.; Kubasco, A. J.

    1982-01-01

    Three simulated coal gas fuels based on hydrogen and carbon monoxide were tested during an experimental evaluation with a rich lean can combustor: these were a simulated Winkler gas, Lurgi gas and Blue Water gas. All three were simulated by mixing together the necessary pure component species, to levels typical of fuel gases produced from coal. The Lurgi gas was also evaluated with ammonia addition. Fuel burning in a rich lean mode was emphasized. Only the Blue Water gas, however, could be operated in such fashion. This showed that the expected NOx signature form could be obtained, although the absolute values of NOx were above the 75 ppm goals for most operating conditions. Lean combustion produced very low NOx well below 75 ppm with the Winkler and Lurgi gases. In addition, these low levels were not significantly impacted by changes in operating conditions.

  17. Numerical Analysis of the SCHOLAR Supersonic Combustor

    NASA Technical Reports Server (NTRS)

    Rodriguez, Carlos G.; Cutler, Andrew D.

    2003-01-01

    The SCHOLAR scramjet experiment is the subject of an ongoing numerical investigation. The facility nozzle and combustor were solved separate and sequentially, with the exit conditions of the former used as inlet conditions for the latter. A baseline configuration for the numerical model was compared with the available experimental data. It was found that ignition-delay was underpredicted and fuel-plume penetration overpredicted, while the pressure rise was close to experimental values. In addition, grid-convergence by means of grid-sequencing could not be established. The effects of the different turbulence parameters were quantified. It was found that it was not possible to simultaneously predict the three main parameters of this flow: pressure-rise, ignition-delay, and fuel-plume penetration.

  18. Mercury emissions from municipal solid waste combustors

    SciTech Connect

    Not Available

    1993-05-01

    This report examines emissions of mercury (Hg) from municipal solid waste (MSW) combustion in the United States (US). It is projected that total annual nationwide MSW combustor emissions of mercury could decrease from about 97 tonnes (1989 baseline uncontrolled emissions) to less than about 4 tonnes in the year 2000. This represents approximately a 95 percent reduction in the amount of mercury emitted from combusted MSW compared to the 1989 mercury emissions baseline. The likelihood that routinely achievable mercury emissions removal efficiencies of about 80 percent or more can be assured; it is estimated that MSW combustors in the US could prove to be a comparatively minor source of mercury emissions after about 1995. This forecast assumes that diligent measures to control mercury emissions, such as via use of supplemental control technologies (e.g., carbon adsorption), are generally employed at that time. However, no present consensus was found that such emissions control measures can be implemented industry-wide in the US within this time frame. Although the availability of technology is apparently not a limiting factor, practical implementation of necessary control technology may be limited by administrative constraints and other considerations (e.g., planning, budgeting, regulatory compliance requirements, etc.). These projections assume that: (a) about 80 percent mercury emissions reduction control efficiency is achieved with air pollution control equipment likely to be employed by that time; (b) most cylinder-shaped mercury-zinc (CSMZ) batteries used in hospital applications can be prevented from being disposed into the MSW stream or are replaced with alternative batteries that do not contain mercury; and (c) either the amount of mercury used in fluorescent lamps is decreased to an industry-wide average of about 27 milligrams of mercury per lamp or extensive diversion from the MSW stream of fluorescent lamps that contain mercury is accomplished.

  19. Low NO sub x heavy fuel combustor concept program

    NASA Technical Reports Server (NTRS)

    Russell, P.; Beal, G.; Hinton, B.

    1981-01-01

    A gas turbine technology program to improve and optimize the staged rich lean low NOx combustor concept is described. Subscale combustor tests to develop the design information for optimization of the fuel preparation, rich burn, quick air quench, and lean burn steps of the combustion process were run. The program provides information for the design of high pressure full scale gas turbine combustors capable of providing environmentally clean combustion of minimally of minimally processed and synthetic fuels. It is concluded that liquid fuel atomization and mixing, rich zone stoichiometry, rich zone liner cooling, rich zone residence time, and quench zone stoichiometry are important considerations in the design and scale up of the rich lean combustor.

  20. Method for operating a combustor in a fuel cell system

    DOEpatents

    Chalfant, Robert W.; Clingerman, Bruce J.

    2002-01-01

    A method of operating a combustor to heat a fuel processor in a fuel cell system, in which the fuel processor generates a hydrogen-rich stream a portion of which is consumed in a fuel cell stack and a portion of which is discharged from the fuel cell stack and supplied to the combustor, and wherein first and second streams are supplied to the combustor, the first stream being a hydrocarbon fuel stream and the second stream consisting of said hydrogen-rich stream, the method comprising the steps of monitoring the temperature of the fuel processor; regulating the quantity of the first stream to the combustor according to the temperature of the fuel processor; and comparing said quantity of said first stream to a predetermined value or range of predetermined values.

  1. SLUDGE COMBUSTOR USING SWIRL AND ACTIVE COMBUSTION CONTROL

    EPA Science Inventory

    A research program directed at developing technology for compact shipboard incinerators for sludges is described. The concept utilizes previously developed Vortex Containment Combustor (VCC) as a primary unit with an active combustion control afterburner (AB). The overall power s...

  2. Investigation of a low NOx full-scale annular combustor

    NASA Technical Reports Server (NTRS)

    1982-01-01

    An atmospheric test program was conducted to evaluate a low NOx annular combustor concept suitable for a supersonic, high-altitude aircraft application. The lean premixed combustor, known as the vortex air blast (VAB) concept, was tested as a 22.0-cm diameter model in the early development phases to arrive at basic design and performance criteria. Final demonstration testing was carried out on a full scale combustor of 0.66-m diameter. Variable geometry dilution ports were incorporated to allow operation of the combustor across the range of conditions between idle (T(in) = 422 K, T(out) = 917 K) and cruise (T(in) = 833 K, T(out) - 1778 K). Test results show that the design could meet the program NOx goal of 1.0 g NO2/kg fuel at a one-atmospheric simulated cruise condition.

  3. Mathematical modelling of coal fired fluidized bed combustors

    SciTech Connect

    Selcuk, N.; Siddall, R.G.; Sivrioglu, U.

    1980-12-01

    A system model of continuous fluidized bed combustors burning coal of wide size distribution has been derived, and applied to the investigation of the effect of excess air and recycle on bed concentration and temperature profiles and combustion efficiency of a pilot scale coal fired fluidized combustor. To demonstrate the effect of recycling, the behaviour of the fluidized combustor has been predicted for two extreme cases of recycle: complete and no recycle of elutriated char particles, the former was chosen to determine the behaviour of the model in the absence of elutriation, and the latter corresponds to the actual operating conditions of the fluidized combustor. Expected trends for concentration and temperature profiles and combustion efficiency are predicted correctly for both cases. The predictive ability and the flexibility of the model for incorporation of refinements such as a correlation for bubble growth and a detailed combustion mechanism, makes the model a promising one for the evaluation of performance of the fluid bed industrial boilers.

  4. A variable geometry combustor for broadened properties fuels

    NASA Technical Reports Server (NTRS)

    Dodds, W. J.; Fear, J. S.

    1987-01-01

    A program was conducted to design and develop a variable geometry combustor, sized for the cycle and envelope of a large commercial turbofan engine. The combustor uses a variable area swirl cup to control stoichiometry in the primary combustion zone. Potential advantages of this design include improved capability to burn non-standard fuels, short system length, and increased operating temperature range for advanced high performance engine cycles. After considerable development, key program emissons and performance goals were met with the variable geometry combustor. Primary development efforts were to evolve improved variable swirl cup configurations. In particular, air leakage through the variable area swirl cup had a strong effect on low power emissions and performance, while smoke level at high power was affected by features for improved mixing of the fuel and swirler air flow. Additional design and development is still needed to evolve a practical variable geometry combustor.

  5. Composite Matrix Cooling Scheme for Small Gas Turbine Combustors

    NASA Technical Reports Server (NTRS)

    Paskin, Marc D.; Ross, Phillip T.; Mongia, Hukam C.; Acosta, Waldo A.

    1990-01-01

    The design, manufacture, and testing of a compliant metal/ceramic (CMC) wall cooling concept-implementing combustor for small gas turbine engines has been undertaken by a joint U.S. Army/NASA technology development program. CMC in principle promises greater wall cooling effectiveness than conventional designs and materials, thereby facilitating a substantial reduction in combustor cooling air requirements and furnishing greater airflow for the control of burner outlet temperature patterns as well as improving thermodynamic efficiency and reducing pollutant emissions and smoke levels. Rig test results have confirmed the projected benefits of the CMC concept at combustor outlet temperatures of the order of 2460 F, at which approximately 80 percent less cooling air than conventionally required was being employed by the CMC combustor.

  6. Exhaust gas emissions of a vortex breakdown stabilized combustor

    NASA Technical Reports Server (NTRS)

    Yetter, R. A.; Gouldin, F. C.

    1976-01-01

    Exhaust gas emission data are described for a swirl stabilized continuous combustor. The combustor consists of confined concentric jets with premixed fuel and air in the inner jet and air in the outer jet. Swirl may be induced in both inner and outer jets with the sense of rotation in the same or opposite directions (co-swirl and counter-swirl). The combustor limits NO emissions by lean operation without sacrificing CO and unburned hydrocarbon emission performance, when commercial-grade methane and air fired at one atmosphere without preheat are used. Relative swirl direction and magnitude are found to have significant effects on exhaust gas concentrations, exit temperatures, and combustor efficiencies. Counter-swirl gives a large recirculation zone, a short luminous combustion zone, and large slip velocities in the interjet shear layer. For maximum counter-swirl conditions, the efficiency is low.

  7. Pollution technology program, can-annular combustor engines

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Fiorentino, A. J.; Greene, W.

    1976-01-01

    A Pollution Reduction Technology Program to develop and demonstrate the combustor technology necessary to reduce exhaust emissions for aircraft engines using can-annular combustors is described. The program consisted of design, fabrication, experimental rig testing and assessment of results and was conducted in three program elements. The combustor configurations of each program element represented increasing potential for meeting the 1979 Environmental Protection Agency (EPA) emission standards, while also representing increasing complexity and difficulty of development and adaptation to an operational engine. Experimental test rig results indicate that significant reductions were made to the emission levels of the baseline JT8D-17 combustor by concepts in all three program elements. One of the Element I single-stage combustors reduced carbon monoxide to a level near, and total unburned hydrocarbons (THC) and smoke to levels below the 1979 EPA standards with little or no improvement in oxides of nitrogen. The Element II two-stage advanced Vorbix (vortex burning and mixing) concept met the standard for THC and achieved significant reductions in CO and NOx relative to the baseline. Although the Element III prevaporized-premixed concept reduced high power NOx below the Element II results, there was no improvement to the integrated EPA parameter relative to the Vorbix combustor.

  8. The Numerical Investigation of a Dual-Mode Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Riggins, David

    1998-01-01

    A numerical investigation of a multiple-jet array dual-mode scramjet combustor has been performed utilizing a three-dimensional Navier-Stokes code with finite-rate chemistry. Results indicate substantial upstream interaction in the form of an oblique shock/expansion train upstream of the combustor, culminating in completely subsonic flow in the vicinity of fuel injectors. The flow returns to supersonic velocities in the downstream (diverging) portion of the combustor. Mixing and combustion are rapid in this flow and predicted combustion efficiency closely matches experimental data. However, comparisons of wall pressure between the simulation and the experiment show i) substantial underprediction of the upstream interaction distance and ii) moderate overprediction of peak pressure in the vicinity of the entrance of the combustor. This can be at least partially explained by examination of available experimental data; this data shows a very significant movement of the entering vitiated airflow to the sides of the combustor (around the injector array and the upstream interaction front as a whole). This important effect is currently being examined by an extension of the modeling to include the entire half-duct of the same combustor geometry.

  9. CFD Analysis of Emissions for a Candidate N+3 Combustor

    NASA Technical Reports Server (NTRS)

    Ajmani, Kumud

    2015-01-01

    An effort was undertaken to analyze the performance of a model Lean-Direct Injection (LDI) combustor designed to meet emissions and performance goals for NASA's N+3 program. Computational predictions of Emissions Index (EINOx) and combustor exit temperature were obtained for operation at typical power conditions expected of a small-core, high pressure-ratio (greater than 50), high T3 inlet temperature (greater than 950K) N+3 combustor. Reacting-flow computations were performed with the National Combustion Code (NCC) for a model N+3 LDI combustor, which consisted of a nine-element LDI flame-tube derived from a previous generation (N+2) thirteen-element LDI design. A consistent approach to mesh-optimization, spray-modeling and kinetics-modeling was used, in order to leverage the lessons learned from previous N+2 flame-tube analysis with the NCC. The NCC predictions for the current, non-optimized N+3 combustor operating indicated a 74% increase in NOx emissions as compared to that of the emissions-optimized, parent N+2 LDI combustor.

  10. CFD Analysis of Emissions for a Candidate N+3 Combustor

    NASA Technical Reports Server (NTRS)

    Ajmani, Kumud

    2015-01-01

    An effort was undertaken to analyze the performance of a model Lean-Direct Injection (LDI) combustor designed to meet emissions and performance goals for NASA's N+3 program. Computational predictions of Emissions Index (EINOx) and combustor exit temperature were obtained for operation at typical power conditions expected of a small-core, high pressure-ratio (greater than 50), high T3 inlet temperature (greater than 950K) N+3 combustor. Reacting-flow computations were performed with the National Combustion Code (NCC) for a model N+3 LDI combustor, which consisted of a nine-element LDI flame-tube derived from a previous generation (N+2) thirteen-element LDI design. A consistent approach to mesh-optimization, spraymodeling and kinetics-modeling was used, in order to leverage the lessons learned from previous N+2 flame-tube analysis with the NCC. The NCC predictions for the current, non-optimized N+3 combustor operating indicated a 74% increase in NOx emissions as compared to that of the emissions-optimized, parent N+2 LDI combustor.

  11. Hypersonic Combustor Model Inlet CFD Simulations and Experimental Comparisons

    NASA Technical Reports Server (NTRS)

    Venkatapathy, E.; TokarcikPolsky, S.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    Numerous two-and three-dimensional computational simulations were performed for the inlet associated with the combustor model for the hypersonic propulsion experiment in the NASA Ames 16-Inch Shock Tunnel. The inlet was designed to produce a combustor-inlet flow that is nearly two-dimensional and of sufficient mass flow rate for large scale combustor testing. The three-dimensional simulations demonstrated that the inlet design met all the design objectives and that the inlet produced a very nearly two-dimensional combustor inflow profile. Numerous two-dimensional simulations were performed with various levels of approximations such as in the choice of chemical and physical models, as well as numerical approximations. Parametric studies were conducted to better understand and to characterize the inlet flow. Results from the two-and three-dimensional simulations were used to predict the mass flux entering the combustor and a mass flux correlation as a function of facility stagnation pressure was developed. Surface heat flux and pressure measurements were compared with the computed results and good agreement was found. The computational simulations helped determine the inlet low characteristics in the high enthalpy environment, the important parameters that affect the combustor-inlet flow, and the sensitivity of the inlet flow to various modeling assumptions.

  12. Particle-jet interactions in an MHD second stage combustor

    SciTech Connect

    Lottes, S.A.; Chang, S.L.

    1992-07-01

    An Argonne two-phase combustion flow computer code is used to simulate reacting flows to aid in the development of an advanced combustor for magnetohydrodynamic power generation. The combustion code is a general hydrodynamics computer code for two-phase, two- dimensional, steady state, turbulent, and reacting flows, based on mass, momentum, and energy conservation laws for multiple gas species and particles. The combustion code includes turbulence, integral combustion, and particle evaporation submodels. A recently developed integral combustion submodel makes calculations more efficient and more stable while still preserving the major physical effects of the complex combustion processes. The combustor under investigation is a magnetohydrodynamic second stage combustor in which opposed jets of oxidizer are injected into a confined cross-stream of hot coal gas flow following a first stage swirl combustor. The simulation is intended to enhance the understanding the of seed particle evaporation in the combustor and evaluate the effects of combustor operating conditions on seed particle evaporation and vapor dispersion, which directly affect overall magnetohydrodynamic power generation. Computation results show that oxidizer jet angle and particle size may greatly affect particle evaporation and vapor dispersion. At a jet angle about 130 degrees, particle evaporation rate is the highest because of the highest average gas temperature. As particle size increases beyond 10 microns in diameter, the effects of particle size on wall deposition rate, evaporation delay, and downstream seed vapor dispersion become more pronounced. 16 refs., 10 figs.

  13. Low NOx, Lean Direct Wall Injection Combustor Concept Developed

    NASA Technical Reports Server (NTRS)

    Tacina, Robert R.; Wey, Changlie; Choi, Kyung J.

    2003-01-01

    The low-emissions combustor development at the NASA Glenn Research Center is directed toward advanced high-pressure aircraft gas turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low-power conditions. Low-NOx combustors can be classified into rich burn and lean burn concepts. Lean burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) combustors. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibly of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone and, thus, does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, the key is good atomization and mixing of the fuel quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP.

  14. Simulations of Small-Scale Liquid Film Combustors

    NASA Astrophysics Data System (ADS)

    Popov, Pavel; Sirignano, William

    2015-11-01

    Recent technological advances have generated need for small-scale combustor designs. The reduction of scale, however, leads to a higher area to volume ratio and thus greater relative heat loss. Liquid film combustors are one proposed design which aims to overcome this obstacle. In them, the fuel is injected as a liquid film on the combustor wall, and heat transfer is reduced due to evaporative cooling of the liquid film leading to reduced temperature gradients at the combustor walls. In this work, we present simulation results for a cylindrical small scale liquid film combustor, in which the reactants are liquid heptane and gaseous air. A computational procedure has been developed to simulate this two-phase combustion problem, using detailed chemical mechanisms. A cubic equation of state is applied for the simulation of the gaseous phase at high pressures. The present study examines the structure of the triple flame inside this combustor design, which has been analyzed in previous experimental work. Comparison between simulation and experimental work is made, with particular emphasis on the influence of the chemical mechanism, high-pressure equation of state, and the effect of swirl amplitudes in the liquid and gas phases on the structure of the flame. Supported by AFOSR grant FA9550-12-1-0156, AFOSR scientific manager: Dr. Mitat Birkan.

  15. Simulated Altitude Performance of Combustors for the Westinghouse 24C Jet Engine I-24C-2 Combustor

    NASA Technical Reports Server (NTRS)

    Manganiello, Eugene J.; Bernardo, Everett; Schroeter, Thomas T.

    1948-01-01

    A Westinghouse 24C-2 combustor was investigated at conditions simulating operation of the 24C Jet engine at zero ram over ranges of altitude and engine speed. The investigation was conducted to determine the altitude operational limits, that is, the maximum altitude for various engine speeds at which an average combustor-outlet gas temperature sufficient for operation of the jet engine could be obtained. Information was also obtained regarding the character of the flames, the combustion efficiency, the combustor-outlet gas temperature and velocity distributions, the extent of afterburning, the flow characteristics of the fuel manifolds, the combustor inlet-to-outlet total-pressure drop, and the durability of the combustor basket. The results of the investigation indicated that the altitude operational limits for zero ram decreased from 12,000 feet at an engine speed of 4000 rpm to a minimum of 9000 feet at 6000 rpm, and thence increased to 49,000 feet at 12,000 rpm.. At altitudes below the operational limits, flames were essentially steady, but, at altitudes above the operational limits, flames were often cycling and either blew out or caused violent explosions and vibrations. At conditions on the altitude operational limits the type of combustion varied from steady to cycling with increasing fuel-air ratio and the reverse occurred with decreasing fuel-air ratio. In the region of operation investigated, the combustion efficiency ranged from 75 to 95 percent at altitudes below the operational limits and dropped to 55 percent or less at some altitudes above the operational limits. The deviations in the local combustor-outlet gas temperatures were within +20 to -30 percent of the mean combustor temperature rise for inlet-air temperatures at the low end of the range investigated, but became more uneven (up to +/-100 percent) with increasing inlet-air temperatures. The distribution of the combustor-outlet gas velocity followed a similar trend. Practically no

  16. Raney nickel catalytic device

    DOEpatents

    O'Hare, Stephen A.

    1978-01-01

    A catalytic device for use in a conventional coal gasification process which includes a tubular substrate having secured to its inside surface by expansion a catalytic material. The catalytic device is made by inserting a tubular catalytic element, such as a tubular element of a nickel-aluminum alloy, into a tubular substrate and heat-treating the resulting composite to cause the tubular catalytic element to irreversibly expand against the inside surface of the substrate.

  17. Evaluation of catalytic combustion of actual coal-derived gas

    NASA Technical Reports Server (NTRS)

    Blanton, J. C.; Shisler, R. A.

    1982-01-01

    The combustion characteristics of a Pt-Pl catalytic reactor burning coal-derived, low-Btu gas were investigated. A large matrix of test conditions was explored involving variations in fuel/air inlet temperature and velocity, reactor pressure, and combustor exit temperature. Other data recorded included fuel gas composition, reactor temperatures, and exhaust emissions. Operating experience with the reactor was satisfactory. Combustion efficiencies were quite high (over 95 percent) over most of the operating range. Emissions of NOx were quite high (up to 500 ppm V and greater), owing to the high ammonia content of the fuel gas.

  18. Flow conditioner for fuel injector for combustor and method for low-NO.sub.x combustor

    DOEpatents

    Dutta, Partha; Smith, Kenneth O.; Ritz, Frank J.

    2013-09-10

    An injector for a gas turbine combustor including a catalyst coated surface forming a passage for feed gas flow and a channel for oxidant gas flow establishing an axial gas flow through a flow conditioner disposed at least partially within an inner wall of the injector. The flow conditioner includes a length with an interior passage opening into upstream and downstream ends for passage of the axial gas flow. An interior diameter of the interior passage smoothly reduces and then increases from upstream to downstream ends.

  19. Investigation into the effects of vermiculite on NOx reduction and additives on sooting and exhaust infrared signature from a gas-turbine combustor. Master's thesis

    SciTech Connect

    Engel, K.R.

    1990-09-01

    An experimental investigation was conducted to determine the feasibility of using catalytic reduction of NOX emissions from a typical jet engine combustor in the test cell environment. A modified T-63 combustor in combination with an instrumented 21 foot augmentation tube containing a vermiculite catalyst was used. Several methods for containing the vermiculite were attempted. Both vermiculite and vermiculite which had been coated with thiourea were used. Up to 19% reduction in NOX concentrations was obtained using the vermiculite coated with thiourea, however the pressure loss across the catalyst bed was measured to be 36 in. H2O. The techniques used proved ineffective and unacceptable for gas turbine engine test cell applications. Tests were conducted using both Wynn's 15/590 and Catane TM (ferrocene) fuel supplements in order to determine their effectiveness for soot reduction and whether or not the exhaust plume could be changed.

  20. Core Noise: Overview of Upcoming LDI Combustor Test

    NASA Technical Reports Server (NTRS)

    Hultgren, Lennart S.

    2012-01-01

    This presentation is a technical summary of and outlook for NASA-internal and NASA-sponsored external research on core (combustor and turbine) noise funded by the Fundamental Aeronautics Program Fixed Wing Project. The presentation covers: the emerging importance of core noise due to turbofan design trends and its relevance to the NASA N+3 noise-reduction goal; the core noise components and the rationale for the current emphasis on combustor noise; and the current and planned research activities in the combustor-noise area. Two NASA-sponsored research programs, with particular emphasis on indirect combustor noise, "Acoustic Database for Core Noise Sources", Honeywell Aerospace (NNC11TA40T) and "Measurement and Modeling of Entropic Noise Sources in a Single-Stage Low-Pressure Turbine", U. Illinois/U. Notre Dame (NNX11AI74A) are briefly described. Recent progress in the development of CMC-based acoustic liners for broadband noise reduction suitable for turbofan-core application is outlined. Combustor-design trends and the potential impacts on combustor acoustics are discussed. A NASA GRC developed nine-point lean-direct-injection (LDI) fuel injector is briefly described. The modification of an upcoming thermo-acoustic instability evaluation of the GRC injector in a combustor rig to also provide acoustic information relevant to community noise is presented. The NASA Fundamental Aeronautics Program has the principal objective of overcoming today's national challenges in air transportation. The reduction of aircraft noise is critical to enabling the anticipated large increase in future air traffic. The Quiet Performance Research Theme of the Fixed Wing Project aims to develop concepts and technologies to dramatically reduce the perceived community noise attributable to aircraft with minimal impact on weight and performance.

  1. CFD simulation of hydrodynamic characteristics on pulse combustor

    NASA Astrophysics Data System (ADS)

    Rahmatika, Annie Mufyda; Salihat, Efaning; Tikasari, Rachma; Widiyastuti, W.; Winardi, Sugeng

    2016-02-01

    The purpose of this research is to study the simulation of the combustion characteristics and performances in pulse combustor using different excess air composition and different pulse combustor geometry using CFD (Computational Fluid Dynamics) software Ansys FLUENT 15.0. The distribution of temperature, pressure, and fluid velocity using 2D axisymmetric with k-ɛ turbulence models. Two kind geometries of pulse combustors were selected and compared their performance. The first combustor, called geometry A has expanded tail-pipe with diameter 10 mm expanded to 20 mm with length 86 mm. The second combustor, called geometry B has cylinder tailpipe which 10 mm in diameter and 200 mm in length. Air and propane were selected as oxidizer and fuel, respectively, at temperature 27°C and pressure 1 atm with varied excess air of 0%, 23%, 200%, and 500%. The simulation result shows that the average temperature of outflow gas combustion decreased with increasing the excess air. On the other hand, the pressure amplitude increased with increasing the excess air. Amplitude of presure for excess air of 0%, 23%, 200% and 500% were 14,976.03 Pa; 26,100.19 Pa; 41,529.02 Pa; and 85,019.01 Pa, respectively. The geometry of pulse combustor affected the performance of gas combustion produced. Geometry A showed that the energy produced in the combustion cycle amounts to 538,639 to 958,639 J/kg. On the other hand, geometry B showed that the generated energy was in the range 864,502 to 1,280,814 J/kg. Combustor with geometry B provided more effective combustion performance rather than B caused by its larger heat transfer area sectional area.

  2. Low NO.sub.x combustor

    NASA Technical Reports Server (NTRS)

    Halila, Ely E. (Inventor)

    1994-01-01

    A combustor includes a dome assembly having radially outer and inner liners joined thereto and defining therebetween a combustion zone. The dome assembly includes at least one annular dome having a pair of axially extending first flanges between which are disposed a plurality of circumferentially spaced apart carburetors for discharging a fuel/air mixture into the combustion zone for generating combustion gases. An annular heat shield includes a pair of axially extending legs integrally joined to a radially extending face in a generally U-shaped configuration, with the face including a plurality of circumferentially spaced apart ports disposed concentrically with perspective ones of the carburetors for allowing the fuel/air mixture to be discharged therefrom through the heat shield. At least one of the heat shield legs includes a plurality of circumferentially spaced apart mounting holes disposed adjacent to a respective one of the dome flanges, and a plurality of mounting pins are fixedly joined to the dome flange and extend radially through respective ones of the mounting holes without interference therewith for allowing unrestrained thermal movement between the heat shield and the dome while supporting the heat shield against axial pressure loads thereon. In a preferred embodiment, the dome assembly includes three domes having respective ones of the heat shield, and respective baffles are spaced from the heat shields for providing impingement cooling thereof.

  3. Fluidized bed combustor and tube construction therefor

    DOEpatents

    De Feo, Angelo; Hosek, William

    1981-01-01

    A fluidized bed combustor comprises a reactor or a housing which has a windbox distributor plate adjacent the lower end thereof which contains a multiplicity of hole and air discharge nozzles for discharging air and coal into a fluidized bed which is maintained above the distributor plate and below a take-off connection or flue to a cyclone separator in which some of the products of combustion are treated to remove the dust which is returned into the fluidized bed. A windbox is spaced below the fluidized bed and it has a plurality of tubes passing therethrough with the passage of combustion air and fluidizing air which passes through an air space so that fluidizing air is discharged into the reaction chamber fluidized bed at the bottom thereof to maintain the bed in a fluidized condition. A fluid, such as air, is passed through the tubes which extend through the windbox and provide a preheating of the combustion air and into an annular space between telescoped inner and outer tubes which comprise heat exchanger tubes or cooling tubes which extend upwardly through the distributor plate into the fluidized bed. The heat exchanger tubes are advantageously arranged so that they may be exposed in groups within the reactor in a cluster which is arranged within holding rings.

  4. Tube construction for fluidized bed combustor

    DOEpatents

    De Feo, Angelo; Hosek, William

    1984-01-01

    A fluidized bed combustor comprises a reactor or a housing which has a windbox distributor plate adjacent the lower end thereof which contains a multiplicity of hole and air discharge nozzles for discharging air and coal into a fluidized bed which is maintained above the distributor plate and below a take-off connection or flue to a cyclone separator in which some of the products of combustion are treated to remove the dust which is returned into the fluidized bed. A windbox is spaced below the fluidized bed and it has a plurality of tubes passing therethrough with the passage of combustion air and fluidizing air which passes through an air space so that fluidizing air is discharged into the reaction chamber fluidized bed at the bottom thereof to maintain the bed in a fluidized condition. A fluid, such as air, is passed through the tubes which extend through the windbox and provide a preheating of the combustion air and into an annular space between telescoped inner and outer tubes which comprise heat exchanger tubes or cooling tubes which extend upwardly through the distributor plate into the fluidized bed. The heat exchanger tubes are advantageously arranged so that they may be exposed in groups within the reactor in a cluster which is arranged within holding rings.

  5. Alternate-Fueled Combustor-Sector Emissions

    NASA Technical Reports Server (NTRS)

    Saxena, Nikita T.; Thomas, Anna E.; Shouse, Dale T.; Neuroth, Craig; Hendricks, Robert C.; Lynch, Amy; Frayne, Charles W.; Stutrud, Jeffrey S.; Corporan, Edwin; Hankins, Terry

    2013-01-01

    In order to meet rapidly growing demand for fuel, as well as address environmental concerns, the aviation industry has been testing alternate fuels for performance and technical usability in commercial and military aircraft. In order to make alternate fuels (and blends) a viable option for aviation, the fuel must be able to perform at a similar or higher level than traditional petroleum fuel. They also attempt to curb harmful emissions, and therefore a truly effective alternate fuel would emit at or under the level of currently used fuel. This report analyzes data from gaseous and particulate emissions of an aircraft combustor sector. The data were evaluated at various inlet conditions, including variation in pressure and temperature, fuel-to-air ratios, and percent composition of alternate fuel. Traditional JP-8+100 data were taken as a baseline, and blends of JP-8+100 with synthetic-paraffinic-kerosene (SPK) fuel (Fischer-Tropsch (FT)) were used for comparison. Gaseous and particulate emissions, as well as flame luminosity, were assessed for differences between FT composition of 0, 50, and 100 percent. The data show that SPK fuel (an FT-derived fuel) had slightly lower harmful gaseous emissions, and smoke number information corroborated the hypothesis that SPK-FT fuels are cleaner burning fuels.

  6. Analysis of Regen Cooling in Rocket Combustors

    NASA Technical Reports Server (NTRS)

    Harper, Brent (Technical Monitor); Merkle, C. L.; Li, D.; Sankaran, V.

    2004-01-01

    The use of detailed CFD modeling for the description of cooling in rocket chambers is discussed. The overall analysis includes a complete three-dimensional analysis of the flow in the regenerative cooling passages, conjugate heat transfer in the combustor walls, and the effects of film cooling on the inside chamber. The results in the present paper omit the effects of film cooling and include only regen cooling and the companion conjugate heat transfer. The hot combustion gases are replaced by a constant temperature wall boundary condition. Load balancing for parallel cluster computations is ensured by using single-block unstructured grids for both fluids and solids, and by using a 'multiple physical zones' to account for differences in the number of equations. Validation of the method is achieved by comparing simple two-dimensional solutions with analytical results. Representative results for cooling passages are presents showing the effects of heat conduction in the copper walls with tube aspect ratios of 1.5:l.

  7. Heat transfer in circulating fluidized bed combustor

    SciTech Connect

    Bucak, O.; Dogan, O.M.; Uysal, B.Z.

    1999-07-01

    The importance of fluidized bed combustion in utilizing the energy of especially low quality coals is widely accepted. Among various fluidized bed combustion technologies, circulating fluidized beds are preferred as a result of the efforts to get higher combustion efficiencies. The aim of the present research was to investigate the applicability of this technology to Turkish lignites. To achieve this object a 6.5 m tall pilot circulating fluidized bed combustor with 155 mm diameter and all the auxiliary equipment were designed, constructed and tested using Seyitomer lignite of 0.9--2.38 mm in size. Heat transfer from the bed to the water cooling jackets was examined to recover the combustion energy. The inside heat transfer coefficient was determined to be around 121 W/m{sup 2} K for the suspension density of 20--55 kg/m{sup 3}. The agreement of the experimental findings with theoretical estimations was also checked. Furthermore, the thermal efficiency of the system for the heat recovered was found to be 63%.

  8. External combustor for gas turbine engine

    DOEpatents

    Santanam, Chandran B.; Thomas, William H.; DeJulio, Emil R.

    1991-01-01

    An external combustor for a gas turbine engine has a cyclonic combustion chamber into which combustible gas with entrained solids is introduced through an inlet port in a primary spiral swirl. A metal draft sleeve for conducting a hot gas discharge stream from the cyclonic combustion chamber is mounted on a circular end wall of the latter adjacent the combustible gas inlet. The draft sleeve is mounted concentrically in a cylindrical passage and cooperates with the passage in defining an annulus around the draft sleeve which is open to the cyclonic combustion chamber and which is connected to a source of secondary air. Secondary air issues from the annulus into the cyclonic combustion chamber at a velocity of three to five times the velocity of the combustible gas at the inlet port. The secondary air defines a hollow cylindrical extension of the draft sleeve and persists in the cyclonic combustion chamber a distance of about three to five times the diameter of the draft sleeve. The hollow cylindrical extension shields the drive sleeve from the inlet port to prevent discharge of combustible gas through the draft sleeve.

  9. Investigation on the flame dynamics of meso-combustors

    NASA Astrophysics Data System (ADS)

    Ahmed, Mahbub

    Miniature heat engines burning hydrogen and hydrocarbon fuels have significantly higher energy densities compared to conventional lithium batteries and thus will play an essential role in the portable production of power for future electronics, remote sensors, and micro aerial vehicles. Additionally, miniature heat engines will tremendously benefit next generation of environmental technologies such as steam reforming, ammonia decomposition and fuel cells. Successful miniaturization of heat engine components demand a more complete and broader understanding of micro-fluid dynamics and micro-combustion phenomena associated with the combustor design. This dissertation is aimed at investigating the details of the micro-mixing dynamics and the combustion behavior of the meso-combustor and to create fundamental understanding of physics based design methodology. The primary goals of the project are (i) to develop an understanding of fuel-air mixing inside a meso-combustor, (ii) to develop an understanding of the flame stability (flame quenching and velocity blowout) criteria of a meso-combustor, (iii) to understand the thermal behavior of the meso-combustor, and (iv) to correlate these with combustor operating conditions such as the Reynolds number, equivalent ratio, and thermal power etc. The present study shows that adequate mixing of fuel and air is achievable in millimeter scale combustors. Both computed results and experimental measurements of iso-thermal (non-burning) flows at different mixing configurations indicate that the laminar burning velocity remains higher than the local flow velocities in most of the combustor locations to support stable flame propagations. Stable flames of hydrogen are achieved for all mixing and flow configurations. The combustion of methane with air as oxidizer in the combustors is unreliable. However, highly stable combustion of methane at various mixing and flow conditions is achieved when pure oxygen is used as an oxidizer. The

  10. Ceramic composite liner material for gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Ercegovic, D. B.; Walker, C. L.; Norgren, C. T.

    1984-01-01

    Advanced commercial and military gas turbine engines may operate at combustor outlet temperatures in excess of 1920 K (3000 F). At these temperatures combustors liners experience extreme convective and radiative heat fluxes. The ability of a plasma sprayed ceramic coating to reduce liner metal temperature has been recognized. However, the brittleness of the ceramic layer and the difference in thermal expansion with the metal substrate has caused cracking, spalling and some separation of the ceramic coating. Research directed at turbine tip seals (or shrouds) has shown the advantage of applying the ceramic to a compliant metal pad. This paper discusses recent studies of applying ceramics to combustor liners in which yttria stabilized zirconia plasma sprayed on compliant metal substrates which were exposed to near stoichiometric combustion, presents performance and durability results, and describes a conceptual design for an advanced, small gas turbine combustor. Test specimens were convectively cooled or convective-transpiration cooled and were evaluated in a 10 cm square flame tube combustor at inlet air temperatures of 533 K (500 F) and at a pressure of 0.5 MPa (75 psia). The ceramics were exposed to flame temperatures in excess of 2000 K (3320 F). Results appear very promising with all 30 specimens surviving a screening test and one of two specimens surviving a cyclic durability test.

  11. Testing of felt-ceramic materials for combustor applications

    NASA Technical Reports Server (NTRS)

    Venkat, R. S.; Roffe, G.

    1983-01-01

    The feasibility of using composite felt ceramic materials as combustor liners was experimentally studied. The material consists of a porous felt pad sandwiched between a layer of ceramic and one of solid metal. Flat, rectangular test panels, which encompassed several design variations of the basic composite material, were tested, two at a time, in a premixed gas turbine combustor as sections of the combustor wall. Tests were conducted at combustor inlet conditions of 0.5 MPa and 533 K with a reference velocity of 25 m/s. The panels were subjected to a hot gas temperature of 2170 K with 1% of the total airflow used to film cool the ceramic surface of the test panel. In general, thin ceramic layers yield low ceramic stress levels with high felt ceramic interface temperatures. On the other hand, thick ceramic layers result in low felt ceramic interface temperatures but high ceramic stress levels. Extensive thermal cycling appears to cause material degradation, but for a limited number of cycles, the survivability of felt ceramic materials, even under extremely severe combustor operating conditions, was conclusively demonstrated.

  12. Flame structures in the pressurized methane-air combustor

    SciTech Connect

    Yamamoto, Tsuyoshi; Miyazaki, Tomonaga, Furuhata, Tomohiko; Arai, Norio

    1998-07-01

    This study has been carried out in order to investigate the applicability of a pressurized and fuel-rich burner at a first stage combustor for a newly proposed chemical gas turbine system. The flammability limits, exhaust gas composition and the NO{sub x} emission characteristics under the pressurized conditions of 1.1--4.1 MPa have been investigated in a model combustor. This paper focuses on the influence of pressure and F/A equivalence ratio on flame structures of pressurized combustion with methane and air to obtain detailed data for designing of fuel-rich combustor for gas turbine application. The flame under fuel-rich condition and pressure of 1 MPa showed underventilated structure like other atmospheric fuel-rich flames while the flame under pressure over 1.5 MPa had shapes as fuel-lean flame. The flame becomes longer as the pressure was increased under the fuel-lean conditions, which under fuel-rich condition the influence of pressure on flame length was smaller in comparison with the flame under fuel-lean conditions. These results give an opportunity for developing smaller combustor under fuel-rich and pressurized condition compared to fuel-lean one. Numerical simulation has been done for defining the temperature profile in the model combustor using the k-{var{underscore}epsilon} turbulence model and three-step reaction model. The comparison between theoretical results and experimental data showed fair agreements.

  13. Experimental studies on methane-fuel laboratory scale ram combustor

    SciTech Connect

    Kinoshita, Y.; Kitajima, J.; Seki, Y.; Tatara, A.

    1995-07-01

    The laboratory scale ram combustor test program has been investigating fundamental combustion characteristics of a ram combustor, which operates from Mach 2.5 to 5 for the super/hypersonic transport propulsion system. In the previous study, combustion efficiency had been found poor, less than 70 percent, due to a low inlet air temperature and a high velocity at Mach 3 condition. To improve the low combustion efficiency, a fuel zoning combustion concept was investigated by using a subscale combustor model first. Combustion efficiency more than 90 percent was achieved and the concept was found very effective. Then a laboratory scale ram combustor was fabricated and combustion tests were carried out mainly at the simulated condition of Mach 5. A vitiation technique wa used to simulate a high temperature of 1,263 K. The test results indicate that ignition, flame stability, and combustion efficiency were not significant, but the NO{sub x} emissions are a critical problem for the ram combustor at Mach 5 condition.

  14. Computational Simulation of Acoustic Modes in Rocket Combustors

    NASA Technical Reports Server (NTRS)

    Harper, Brent (Technical Monitor); Merkle, C. L.; Sankaran, V.; Ellis, M.

    2004-01-01

    A combination of computational fluid dynamic analysis and analytical solutions is being used to characterize the dominant modes in liquid rocket engines in conjunction with laboratory experiments. The analytical solutions are based on simplified geometries and flow conditions and are used for careful validation of the numerical formulation. The validated computational model is then extended to realistic geometries and flow conditions to test the effects of various parameters on chamber modes, to guide and interpret companion laboratory experiments in simplified combustors, and to scale the measurements to engine operating conditions. In turn, the experiments are used to validate and improve the model. The present paper gives an overview of the numerical and analytical techniques along with comparisons illustrating the accuracy of the computations as a function of grid resolution. A representative parametric study of the effect of combustor mean flow Mach number and combustor aspect ratio on the chamber modes is then presented for both transverse and longitudinal modes. The results show that higher mean flow Mach numbers drive the modes to lower frequencies. Estimates of transverse wave mechanics in a high aspect ratio combustor are then contrasted with longitudinal modes in a long and narrow combustor to provide understanding of potential experimental simulations.

  15. 40 CFR 60.54a - Standard for municipal waste combustor acid gases.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 7 2013-07-01 2013-07-01 false Standard for municipal waste combustor... for Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) (c) On...

  16. 40 CFR 60.56a - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 7 2013-07-01 2013-07-01 false Standards for municipal waste combustor... Performance for Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.56a Standards for municipal waste combustor operating practices....

  17. 40 CFR 60.54a - Standard for municipal waste combustor acid gases.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standard for municipal waste combustor... for Municipal Waste Combustors for Which Construction Is Commenced After December 20, 1989 and On or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) (c) On...

  18. 40 CFR 60.54a - Standard for municipal waste combustor acid gases.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 7 2012-07-01 2012-07-01 false Standard for municipal waste combustor... for Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) (c) On...

  19. 40 CFR 60.54a - Standard for municipal waste combustor acid gases.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 6 2011-07-01 2011-07-01 false Standard for municipal waste combustor... for Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) (c) On...

  20. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste combustor... Standards of Performance for Large Municipal Waste Combustors for Which Construction is Commenced After... Standards for municipal waste combustor operator training and certification. (a) No later than the date...

  1. 40 CFR 60.54a - Standard for municipal waste combustor acid gases.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standard for municipal waste combustor... for Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) (c) On...

  2. 40 CFR 60.56a - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... Performance for Municipal Waste Combustors for Which Construction Is Commenced After December 20, 1989 and On or Before September 20, 1994 § 60.56a Standards for municipal waste combustor operating practices....

  3. 40 CFR 60.36b - Emission guidelines for municipal waste combustor fugitive ash emissions.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... combustor fugitive ash emissions. 60.36b Section 60.36b Protection of Environment ENVIRONMENTAL PROTECTION... September 20, 1994 § 60.36b Emission guidelines for municipal waste combustor fugitive ash emissions. For approval, a State plan shall include requirements for municipal waste combustor fugitive ash emissions...

  4. 40 CFR Table 3 to Subpart Fff of... - Municipal Waste Combustor Operating Requirements

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Operating Requirements Municipal waste combustor technology Carbon monoxide emissions level (parts per... combustor 200 24 a Measured at the combustor outlet in conjunction with a measurement of oxygen concentration, corrected to 7 percent oxygen, dry basis. Calculated as an arithmetic average. b Averaging...

  5. Combustor with two stage primary fuel assembly

    DOEpatents

    Sharifi, Mehran; Zolyomi, Wendel; Whidden, Graydon Lane

    2000-01-01

    A combustor for a gas turbine having first and second passages for pre-mixing primary fuel and air supplied to a primary combustion zone. The flow of fuel to the first and second pre-mixing passages is separately regulated using a single annular fuel distribution ring having first and second row of fuel discharge ports. The interior portion of the fuel distribution ring is divided by a baffle into first and second fuel distribution manifolds and is located upstream of the inlets to the two pre-mixing passages. The annular fuel distribution ring is supplied with fuel by an annular fuel supply manifold, the interior portion of which is divided by a baffle into first and second fuel supply manifolds. A first flow of fuel is regulated by a first control valve and directed to the first fuel supply manifold, from which the fuel is distributed to first fuel supply tubes that direct it to the first fuel distribution manifold. From the first fuel distribution manifold, the first flow of fuel is distributed to the first row of fuel discharge ports, which direct it into the first pre-mixing passage. A second flow of fuel is regulated by a second control valve and directed to the second fuel supply manifold, from which the fuel is distributed to second fuel supply tubes that direct it to the second fuel distribution manifold. From the second fuel distribution manifold, the second flow of fuel is distributed to the second row of fuel discharge ports, which direct it into the second pre-mixing passage.

  6. Development of pressurized coal partial combustor

    SciTech Connect

    Yoshida, K.; Ino, T.; Yamamoto, T.; Kimura, N.

    1995-12-31

    The integrated gasification combined cycle (IGCC), an environment-friendly power generation system of high thermal efficiency, is being developed via various approaches around the world. The oxygen-blown entrained flow gasification process is a relatively simple method of producing medium calorie coal gas suitable for application to gas turbines. Various systems for this process have been developed to a demonstration level in Europe and America. Japan has actively been developing the air-blown process. However, taking stable molten slag discharge into consideration, coal must be supplied at two stages to raise the combustor temperature in ash molten part. Only two reports have been presented regarding two-stage coal supply. One is the report on an experiment with the Hycol gasifier, in which air feed ratio is varied, with coal feed fixed. The other is report on a simulation study with various gasifier coal feed ratios, conducted at Central Research Institute of Electric Power Industry. It seems that the appropriate feed ratio has not yet been established. Through this activity, a unique furnace construction has been established, and these influences of stoichiometric air ratio, of oxygen enrichment, of char recycling and of coal types on performance have been clarified. The purpose of the present study is to apply this developed CPC techniques to a Pressurized CPC (PCPC), thereby improving the IGCC technology. For the present study, we conducted systematic experiments on the air-blown process with a two stage dry feed system, using a 7 t/d-coal bench scale PCPC test facility, operated at the pressure of 0.4 MPa, and clarified the influence of coal feed ratio on coal gasification performance. This report describes the above-mentioned bench scale test procedures and results, and also some informations about a plan of a 25 t/d-coal pilot test system.

  7. Computational fluid dynamic analysis of hybrid rocket combustor flowfields

    NASA Technical Reports Server (NTRS)

    Venkateswaran, S.; Merkle, C. L.

    1995-01-01

    Computational fluid dynamic analyses of the Navier-Stokes equations coupled with solid-phase pyrolysis, gas-phase combustion, turbulence and radiation are performed to study hybrid rocket combustor flowfields. The computational study is closely co-ordinated with a companion experimental program using a planar slab burner configuration with HTPB as fuel and gaseous oxygen. Computational predictions agree reasonably well with measurement data of fuel regression rates and surface temperatures. Additionally, most of the parametric trends predicted by the model are in general agreement with experimental trends. The computational model is applied to extend the results from the lab-scale to a full-scale axisymmetric configuration. The numerical predictions indicate that the full-scale configuration burns at a slower rate than the lab-scale combustor under identical specific flow rate conditions. The results demonstrate that detailed CFD analyses can play a useful role in the design of hybrid combustors.

  8. Gas turbine engine combustor can with trapped vortex cavity

    DOEpatents

    Burrus, David Louis; Joshi, Narendra Digamber; Haynes, Joel Meier; Feitelberg, Alan S.

    2005-10-04

    A gas turbine engine combustor can downstream of a pre-mixer has a pre-mixer flowpath therein and circumferentially spaced apart swirling vanes disposed across the pre-mixer flowpath. A primary fuel injector is positioned for injecting fuel into the pre-mixer flowpath. A combustion chamber surrounded by an annular combustor liner disposed in supply flow communication with the pre-mixer. An annular trapped dual vortex cavity located at an upstream end of the combustor liner is defined between an annular aft wall, an annular forward wall, and a circular radially outer wall formed therebetween. A cavity opening at a radially inner end of the cavity is spaced apart from the radially outer wall. Air injection first holes are disposed through the forward wall and air injection second holes are disposed through the aft wall. Fuel injection holes are disposed through at least one of the forward and aft walls.

  9. Low NO(x) heavy fuel combustor program

    NASA Technical Reports Server (NTRS)

    Lister, E.; Niedzwiecki, R. W.; Nichols, L.

    1979-01-01

    The 'low nitrogen oxides heavy fuel combustor' program is described. Main program objectives are to generate and demonstrate the technology required to develop durable gas turbine combustors for utility and industrial applications, which are capable of sustained, environmentally acceptable operation with minimally processed petroleum residual fuels. The program will focus on 'dry' reductions of oxides of nitrogen, improved combustor durability, and satisfactory combustion of minimally processed petroleum residual fuels. Other technology advancements sought include: fuel flexibility for operation with petroleum distillates, blends of petroleum distillates and residual fuels, and synfuels (fuel oils derived from coal or shale); acceptable exhaust emissions of carbon monoxide, unburned hydrocarbons, sulfur oxides and smoke; and retrofit capability to existing engines.

  10. National Combustion Code: A Multidisciplinary Combustor Design System

    NASA Technical Reports Server (NTRS)

    Stubbs, Robert M.; Liu, Nan-Suey

    1997-01-01

    The Internal Fluid Mechanics Division conducts both basic research and technology, and system technology research for aerospace propulsion systems components. The research within the division, which is both computational and experimental, is aimed at improving fundamental understanding of flow physics in inlets, ducts, nozzles, turbomachinery, and combustors. This article and the following three articles highlight some of the work accomplished in 1996. A multidisciplinary combustor design system is critical for optimizing the combustor design process. Such a system should include sophisticated computer-aided design (CAD) tools for geometry creation, advanced mesh generators for creating solid model representations, a common framework for fluid flow and structural analyses, modern postprocessing tools, and parallel processing. The goal of the present effort is to develop some of the enabling technologies and to demonstrate their overall performance in an integrated system called the National Combustion Code.

  11. Numerical Simulation of Dual-Mode Scramjet Combustors

    NASA Technical Reports Server (NTRS)

    Rodriguez, C. G.; Riggins, D. W.; Bittner, R. D.

    2000-01-01

    Results of a numerical investigation of a three-dimensional dual-mode scramjet isolator-combustor flow-field are presented. Specifically, the effect of wall cooling on upstream interaction and flow-structure is examined for a case assuming jet-to-jet symmetry within the combustor. Comparisons are made with available experimental wall pressures. The full half-duct for the isolator-combustor is then modeled in order to study the influence of side-walls. Large scale three-dimensionality is observed in the flow with massive separation forward on the side-walls of the duct. A brief review of convergence-acceleration techniques useful in dual-mode simulations is presented, followed by recommendations regarding the development of a reliable and unambiguous experimental data base for guiding CFD code assessments in this area.

  12. Analytical fuel property effects: Small combustors, phase 2

    NASA Technical Reports Server (NTRS)

    Hill, T. G.; Monty, J. D.; Morton, H. L.

    1985-01-01

    The effects of non-standard aviation fuels on a typical small gas turbine combustor were studied and the effectiveness of design changes intended to counter the effects of these fuels was evaluated. The T700/CT7 turboprop engine family was chosen as being representative of the class of aircraft power plants desired for this study. Fuel properties, as specified by NASA, are characterized by low hydrogen content and high aromatics levels. No. 2 diesel fuel was also evaluated in this program. Results demonstrated the anticipated higher than normal smoke output and flame radiation intensity with resulting increased metal temperatures on the baseline T700 combustor. Three new designs were evaluated using the non standard fuels. The three designs incorporated enhanced cooling features and smoke reduction features. All three designs, when burning the broad specification fuels, exhibited metal temperatures at or below the baseline combustor temperatures on JP-5. Smoke levels were acceptable but higher than predicted.

  13. Energy efficient engine diffuser/combustor model technology

    NASA Technical Reports Server (NTRS)

    Gardner, W.

    1980-01-01

    A full scale, full annular diffuser/combustor model test rig was tested to investigate how configurational changes affect pressure loss and flow separation characteristics. The rig was characterized by five major modules: inlet; prediffuser; strut; simulated combustor; and full combustor. The prediffuser featured a short, curved wall dump design. Performance goals included: (1) a separation-free prediffuser flow field; (2) total pressure loss limited to 3.0 percent in the prediffuser and shrouds; and (3) an overall section pressure loss of 5.5 percent P sub T3 at the design airflow distribution. The results indicated that the prediffuser configurations operate well within the program goals for pressure loss and demonstrate separation free operation over a wide range of inlet conditions.

  14. Laser velocimetry measurements in a gas turbine research combustor

    NASA Technical Reports Server (NTRS)

    Driscoll, J. F.; Pelaccio, D. G.

    1979-01-01

    The effects of turbulence on the production of pollutant species in a gas-turbine research combustor are studied using laser diffraction velocimetry (LDV) techniques. Measurements that were made in the primary combustion zone include mean velocity, rms velocity fluctuations, velocity probability distributions, and autocorrelation functions. A unique combustor design provides relatively uniform flow conditions and independent control of drop size, equivalence ratio, inlet temperature, and combustor pressure. Parameters which characterize the nature of the spray combustion (i.e., whether single droplet or group combustion occurs), were determined from the LDV data. Turbulent diffusivity (eddy viscosity) reaches a value of 2930 sq cm/sec, corresponding to a convective integral length scale of 1.8 cm. The group combustion number, based on turbulent diffusivity, is measured to be 6.2

  15. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2007-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  16. Low NO/x/ heavy fuel combustor program

    NASA Technical Reports Server (NTRS)

    Lister, E.; Niedzwiecki, R. W.; Nichols, L.

    1980-01-01

    The paper deals with the 'Low NO/x/ Heavy Fuel Combustor Program'. Main program objectives are to generate and demonstrate the technology required to develop durable gas turbine combustors for utility and industrial applications, which are capable of sustained, environmentally acceptable operation with minimally processed petroleum residual fuels. The program will focus on 'dry' reductions of oxides of nitrogen (NO/x/), improved combustor durability and satisfactory combustion of minimally processed petroleum residual fuels. Other technology advancements sought include: fuel flexibility for operation with petroleum distillates, blends of petroleum distillates and residual fuels, and synfuels (fuel oils derived from coal or shale); acceptable exhaust emissions of carbon monoxide, unburned hydrocarbons, sulfur oxides and smoke; and retrofit capability to existing engines.

  17. Emissions from biomass combustion in a fluidized bed combustor and gas cleanup system

    SciTech Connect

    Burton, B.; Lighty, J.S.; Inkley, D.; Eddings, E.; Overacker, D.; Davis, K.; Lee, C.; Sarofim, A.

    1999-07-01

    The University of Utah Department of Chemical and Fuels Engineering and Reaction Engineering International have designed and tested a fluidized bed for resource recovery in a Mars or lunar space station for feed streams consisting of inedible plant biomass and solid human waste. In conjunction with the combustor, the system has an extensive flue gas clean-up system to meet Spacecraft Maximum Allowable Concentrations (SMACs). This paper discusses the selection of a rich low-temperature combustion mode that minimizes the ash fusion problems with the high potassium feed and which generates sufficient unburned carbon monoxide to enable the reduction of NO. The components of the gas clean-up stream include: particle removal; HCl removal; NO{sub x} reduction; hydrocarbon and carbon monoxide destruction; sulfur capture; and a final gas polishing unit. Major developmental efforts were required to develop systems for trouble-free waste feeding and NO{sub x} reduction. The combustor is operated at temperatures below 700 C since the ash component of the hydroponically grown inedible biomass has a very low melting point. Low temperature operation results in high levels of CO and unburned hydrocarbons, which can be used as reducing agents for NO{sub x} in the downstream catalytic unit. This is more desirable than using ammonia, which is hazardous, and an expendable reagent that must be stored in sufficient quantity for the duration of a mission. The paper will discuss the results of an innovative catalyst system to reduce NO{sub x}, hydrocarbons, and CO. One important feature of this totally regenerative system is the potential reuse of potassium and sulfur captured in the ash for the hydroponic plant solution.

  18. Combustor flow computations in general coordinates with a multigrid method

    NASA Astrophysics Data System (ADS)

    Shyy, Wei; Braaten, Mark E.

    The computational approach presented for single-phase combusting turbulent flowfields balances the requirements of complex physical and chemical flow interactions with those of resolving the three-dimensional geometrical constraints of the combustor contours, film cooling slots, and circular dilution holes. Attention is given to the three-dimensional grid-generation algorithm, the two-dimensional adaptive grid method applied to recirculating turbulent reacting flows, and theory/data assessments for three-dimensional combusting flows in an annular gas turbine combustor.

  19. Preliminary studies of combustor sensitivity to alternative fuels

    NASA Technical Reports Server (NTRS)

    Humenik, F. M.

    1980-01-01

    Combustion problems associated with using alternative fuels ground power and aeropropulsion applications were studied. Rectangular sections designed to simulate large annular combustor test conditions were examined. The effects of using alternative fuels with reduced hydrogen content, increased aromatic content, and a broad variation in fuel property characteristics were also studied. Data of special interest were collected which include: flame radiation characteristics in the various combustor zones; the correponding increase in liner temperature from increased radiant heat flux; the effect of fuel bound nitrogen on oxides of nitrogen (NO sub x) emissions; and the overall total effect of fuel variations on exhaust emissions.

  20. Combustion-acoustic stability analysis for premixed gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Darling, Douglas; Radhakrishnan, Krishnan; Oyediran, Ayo; Cowan, Lizabeth

    1995-01-01

    Lean, prevaporized, premixed combustors are susceptible to combustion-acoustic instabilities. A model was developed to predict eigenvalues of axial modes for combustion-acoustic interactions in a premixed combustor. This work extends previous work by including variable area and detailed chemical kinetics mechanisms, using the code LSENS. Thus the acoustic equations could be integrated through the flame zone. Linear perturbations were made of the continuity, momentum, energy, chemical species, and state equations. The qualitative accuracy of our approach was checked by examining its predictions for various unsteady heat release rate models. Perturbations in fuel flow rate are currently being added to the model.

  1. Variable volume combustor with center hub fuel staging

    DOEpatents

    Ostebee, Heath Michael; McConnaughhay, Johnie Franklin; Stewart, Jason Thurman; Keener, Christopher Paul

    2016-08-23

    The present application and the resultant patent provide a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a center hub for providing the flow of fuel therethrough. The center hub may include a first supply circuit for a first micro-mixer fuel nozzle and a second supply circuit for a second micro-mixer fuel nozzle.

  2. Method for operating a combustor in a fuel cell system

    DOEpatents

    Clingerman, Bruce J.; Mowery, Kenneth D.

    2002-01-01

    In one aspect, the invention provides a method of operating a combustor to heat a fuel processor to a desired temperature in a fuel cell system, wherein the fuel processor generates hydrogen (H.sub.2) from a hydrocarbon for reaction within a fuel cell to generate electricity. More particularly, the invention provides a method and select system design features which cooperate to provide a start up mode of operation and a smooth transition from start-up of the combustor and fuel processor to a running mode.

  3. Large-eddy simulations of flows in a ramjet combustor

    NASA Astrophysics Data System (ADS)

    Jou, Wen-Huei; Menon, Suresh

    The oscillatory cold flow in a ramjet combustor configuration is presently addressed by a numerical simulation method which gives attention to the interaction between the flowfield's vorticity and acoustic components, when the reduced frequency of the flow, based on the speed of sound, is of the order of unity. The numerical model has indicated that the combustor's interior must be isolated from the external region region by a choked nozzle. The numerical simulations thus obtained are able to exclude the effects of artificially imposed outflow-boundary conditions. The unsteady flow fields near the shear layer separation point in the nozzle region are investigated.

  4. Adaptive Instability Suppression Controls in a Liquid-fueled Combustor

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; DeLaat, John C.

    2002-01-01

    An adaptive control algorithm has been developed for the suppression of combustion thermo-acoustic instabilities. This technique involves modulating the fuel flow in the combustor with a control phase that continuously slides within the stable phase region, in a back and forth motion. The control method is referred to as Adaptive Sliding Phasor Averaged Control (ASPAC). The control method is evaluated against a simplified simulation of the combustion instability. Plans are to validate the control approach against a more physics-based model and an actual experimental combustor rig.

  5. Numerical Simulations of Static Tested Ramjet Dump Combustor

    NASA Astrophysics Data System (ADS)

    Javed, Afroz; Chakraborty, Debasis

    2016-06-01

    The flow field of a Liquid Fuel Ram Jet engine side dump combustor with kerosene fuel is numerically simulated using commercial CFD code CFX-11. Reynolds Averaged 3-D Navier-Stokes equations are solved alongwith SST turbulence model. Single step infinitely fast reaction is assumed for kerosene combustion. The combustion efficiency is evaluated in terms of the unburnt kerosene vapour leaving the combustor. The comparison of measured pressures with computed values show that the computation underpredicts (~5 %) pressures for non reacting cases but overpredicts (9-7 %) for reacting cases.

  6. Stagnation point reverse flow combustor for a combustion system

    NASA Technical Reports Server (NTRS)

    Zinn, Ben T. (Inventor); Neumeier, Yedidia (Inventor); Seitzman, Jerry M. (Inventor); Jagoda, Jechiel (Inventor); Hashmonay, Ben-Ami (Inventor)

    2007-01-01

    A combustor assembly includes a combustor vessel having a wall, a proximate end defining an opening and a closed distal end opposite said proximate end. A manifold is carried by the proximate end. The manifold defines a combustion products exit. The combustion products exit being axially aligned with a portion of the closed distal end. A plurality of combustible reactant ports is carried by the manifold for directing combustible reactants into the combustion vessel from the region of the proximate end towards the closed distal end.

  7. Nonlinear structural and life analyses of a combustor liner

    NASA Technical Reports Server (NTRS)

    Moreno, V.; Meyers, G. J.; Kaufman, A.; Halford, G. R.

    1982-01-01

    Three dimensional, nonlinear finite element structural analyses were performed for a simulated combustor liner specimen to assess the capability of nonlinear analyses using classical inelastic material models to represent the thermoplastic creep response of the one half scale component. Results indicate continued cyclic hardening and ratcheting while experimental data suggested a stable stress strain response after only a few loading cycles. The computed stress strain history at the critical location was put into two life prediction methods, strainrange partitioning and a Pratt and Whitney combustor life prediction method to evaluate their ability to predict cyclic crack initiation. It is found that the life prediction analyses over predicted the observed cyclic crack initiation life.

  8. Switchable catalytic DNA catenanes.

    PubMed

    Hu, Lianzhe; Lu, Chun-Hua; Willner, Itamar

    2015-03-11

    Two-ring interlocked DNA catenanes are synthesized and characterized. The supramolecular catenanes show switchable cyclic catalytic properties. In one system, the catenane structure is switched between a hemin/G-quadruplex catalytic structure and a catalytically inactive state. In the second catenane structure the catenane is switched between a catalytically active Mg(2+)-dependent DNAzyme-containing catenane and an inactive catenane state. In the third system, the interlocked catenane structure is switched between two distinct catalytic structures that include the Mg(2+)- and the Zn(2+)-dependent DNAzymes. PMID:25642796

  9. Experimental evaluation of fuel preparation systems for an automotive gas turbine catalytic combustor

    NASA Technical Reports Server (NTRS)

    Tacina, R. R.

    1977-01-01

    Spatial fuel distributions, degree of vaporization, pressure drop and air velocity profiles were measured. Three airblast injectors and an air-assist nozzle were tested. Air swirlers were used to improve the spatial fuel-air distribution. The work was done in a 12 cm tubular duct. Test conditions were: a pressure of 0.3 and 0.5 MPa, inlet air temperatures up to 800 K, air velocities of 10 20 m/s and fuel-air ratios up to 0.020. The fuel was Jet A. The best results were obtained with an airblast configuration that used multiple cones to provide high velocity air for atomization and also straightened the inlet airflow. With this configuration, uniform spatial fuel-air distributions were obtained with mixing lengths greater than 17.8 cm. In this length, vaporization of the fuel was 98.5 percent complete at an inlet air temperature of 700 K. The total pressure loss was 1.0 percent with a reference velocity of 20 m/s and 0.25 percent at 10m/s. The air velocity was uniform across the duct and no autoignition reactions were observed.

  10. Turbine combustor with fuel nozzles having inner and outer fuel circuits

    DOEpatents

    Uhm, Jong Ho; Johnson, Thomas Edward; Kim, Kwanwoo

    2013-12-24

    A combustor cap assembly for a turbine engine includes a combustor cap and a plurality of fuel nozzles mounted on the combustor cap. One or more of the fuel nozzles would include two separate fuel circuits which are individually controllable. The combustor cap assembly would be controlled so that individual fuel circuits of the fuel nozzles are operated or deliberately shut off to provide for physical separation between the flow of fuel delivered by adjacent fuel nozzles and/or so that adjacent fuel nozzles operate at different pressure differentials. Operating a combustor cap assembly in this fashion helps to reduce or eliminate the generation of undesirable and potentially harmful noise.

  11. Small gas turbine combustor experimental study: Compliant metal/ceramic liner and performance evaluation

    NASA Technical Reports Server (NTRS)

    Acosta, W. A.; Norgren, C. T.

    1986-01-01

    Combustor research relating to the development of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently underway at NASA Lewis. As part of this combustor research, a basic reverse-flow combustor has been used to investigate advanced liner wall cooling techniques. Liner temperature, performance, and exhaust emissions of the experimental combustor utilizing compliant metal/ceramic liners were determined and compared with three previously reported combustors that featured: (1) splash film-cooled liner walls; (2) transpiration cooled liner walls; and (3) counter-flow film cooled panels.

  12. NASA/Pratt and Whitney experimental clean combustor program: Engine test results

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Fiorentino, A. J.; Greene, W.

    1977-01-01

    A two-stage vorbix (vortex burning and mixing) combustor and associated fuel system components were successfully tested in an experimental JT9D engine at steady-state and transient operating conditions, using ASTM Jet-A fuel. Full-scale JT9D experimental engine tests were conducted in a phase three aircraft experimental clean combustor program. The low-pollution combustor, fuel system, and fuel control concepts were derived from phase one and phase two programs in which several combustor concepts were evaluated, refined, and optimized in a component test rig. Significant pollution reductions were achieved with the combustor which meets the performance, operating, and installation requirements of the engine.

  13. Computations of soot and and NO sub x emissions from gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Srivatsa, S. K.

    1982-01-01

    An analytical program was conducted to compute the soot and NOx emissions from a combustor and the radiation heat transfer to the combustor walls. The program involved the formulation of an emission and radiation model and the incorporation of this model into the Garrett 3-D Combustor Perfomance Computer Program. Computations were performed for the idle, cruise, and take-off conditions of a JT8D can combustor. The predicted soot and NOx emissions and the radiation heat transfer to the combustor walls agree reasonably well with the limited experimental data available.

  14. Hypersonic research engine project. Phase 2: Some combustor test results of NASA aerothermodynamic integration model

    NASA Technical Reports Server (NTRS)

    Sun, Y. H.; Gaede, A. E.; Sainio, W. C.

    1975-01-01

    Combustor test results of the NASA Aerothermodynamic Integration Model are presented of a ramjet engine developed for operation between Mach 3 and 8. Ground-based and flight experiments which provide the data required to advance the technology of hypersonic air-breathing propulsion systems as well as to evaluate facility and testing techniques are described. The engine was tested with synthetic air at Mach 5, 6, and 7. The hydrogen fuel was heated to 1500 R prior to injection to simulate a regeneratively cooled system. Combustor efficiencies up to 95 percent at Mach 6 were achieved. Combustor process in terms of effectiveness, pressure integral factor, total pressure recovery and Crocco's pressure-area relationship are presented and discussed. Interactions between inlet-combustor, combustor stages, combustor-nozzle, and the effects of altitude, combustor step, and struts are observed and analyzed.

  15. Non-reacting flow visualization of supersonic combustor based on cavity and cavity-strut flameholder

    NASA Astrophysics Data System (ADS)

    Zhao, Yanhui; Liang, Jianhan; Zhao, Yuxin

    2016-04-01

    Nano-particle planer laser scattering and particle image velocimetry technology are employed to observe the flow field of scramjet combustors based on cavity and cavity-strut flameholder. Density field and velocity distribution inside combustors are obtained. Mainstream fluid enters into cavity nearby side wall in experimental observation because side wall shock waves interact with bottom wall boundary layer. Cavity fluid is entrained into mainstream in the middle of combustor meanwhile. Flow past cavity displays obvious three dimensional characteristics in both combustors. But cavity-strut combustor displays asymmetrical flow field because of strut configuration. Mass exchange between mainstream and cavity fluid is evaluated by statistic mass flow rate into cavity. Mass flow rate near side wall is raised to 6.62 times of the value in the middle of cavity combustor while it is 5.1 times in cavity-strut combustor. Further study is needed to injection strategies and realistic flow characteristics on condition of combustion.

  16. MUNICIPAL SOLID WASTE (MSW) COMBUSTOR ASH DEMONSTRATION PROGRAM - "THE BOATHOUSE"

    EPA Science Inventory

    The report presents the results of a research program designed to examine the engineering and environmental acceptability of using municipal solid waste (MSW) combustor ash as an aggregate substitute in the manufacture of construction quality cement blocks. 50 tons of MSW combust...

  17. DEVELOPMENT OF A VORTEX CONTAINMENT COMBUSTOR FOR COAL COMBUSTION SYTEMS

    EPA Science Inventory

    The report describes the development of a vortex containment combustor (VCC) for coal combustion systems, designed to solve major problems facing the conversion of oil- and gas-fired boilers to coal (e.g., derating, inorganic impurities in coal, and excessive formation of NOx and...

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

    NASA Technical Reports Server (NTRS)

    Nemets, Steve A.; Ehlers, Robert C.; Parrott, Edith

    1995-01-01

    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.

  19. Acoustic modal analysis of a full-scale annular combustor

    NASA Technical Reports Server (NTRS)

    Karchmer, A. M.

    1982-01-01

    An acoustic modal decomposition of the measured pressure field in a full scale annular combustor installed in a ducted test rig is described. The modal analysis, utilizing a least squares optimization routine, is facilitated by the assumption of randomly occurring pressure disturbances which generate equal amplitude clockwise and counter-clockwise pressure waves, and the assumption of statistical independence between modes. These assumptions are fully justified by the measured cross spectral phases between the various measurement points. The resultant modal decomposition indicates that higher order modes compose the dominant portion of the combustor pressure spectrum in the range of frequencies of interest in core noise studies. A second major finding is that, over the frequency range of interest, each individual mode which is present exists in virtual isolation over significant portions of the spectrum. Finally, a comparison between the present results and a limited amount of data obtained in an operating turbofan engine with the same combustor is made. The comparison is sufficiently favorable to warrant the conclusion that the structure of the combustor pressure field is preserved between the component facility and the engine.

  20. EFFECT OF SOOT AND COPPER COMBUSTOR DEPOSITS ON DIOXIN EMISSIONS

    EPA Science Inventory

    An experimental study was conducted to investigate the effects of residual soot and copper combustor deposits on the formation of polychlorinated dibenzo-p-dioxins (PCDDs) and polychlorinated dibenzofurans (PCDFs) during the combustion of a chlorinated waste. In a bench-scale set...

  1. DEVELOPMENT OF GOOD COMBUSTION PRACTICE FOR MUNICIPAL WASTE COMBUSTORS

    EPA Science Inventory

    The paper summarizes the rationale for EPA's good combustion practice (GCP) strategy. OTE: The EPA is developing new air pollution rules for all new and existing municipal waste combustors (MWCs), rules requiring all MWCs to use GCP. The goals of GCP are to maximize furnace destr...

  2. Using the NASA GRC Sectored-One-Dimensional Combustor Simulation

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.; Mehta, Vishal R.

    2014-01-01

    The document is a user manual for the NASA GRC Sectored-One-Dimensional (S-1-D) Combustor Simulation. It consists of three sections. The first is a very brief outline of the mathematical and numerical background of the code along with a description of the non-dimensional variables on which it operates. The second section describes how to run the code and includes an explanation of the input file. The input file contains the parameters necessary to establish an operating point as well as the associated boundary conditions (i.e. how it is fed and terminated) of a geometrically configured combustor. It also describes the code output. The third section describes the configuration process and utilizes a specific example combustor to do so. Configuration consists of geometrically describing the combustor (section lengths, axial locations, and cross sectional areas) and locating the fuel injection point and flame region. Configuration requires modifying the source code and recompiling. As such, an executable utility is included with the code which will guide the requisite modifications and insure that they are done correctly.

  3. Performance characteristics of a slagging gasifier for MHD combustor systems

    NASA Technical Reports Server (NTRS)

    Smith, K. O.

    1979-01-01

    The performance of a two stage, coal combustor concept for magnetohydrodynamic (MHD) systems was investigated analytically. The two stage MHD combustor is comprised of an entrained flow, slagging gasifier as the first stage, and a gas phase reactor as the second stage. The first stage was modeled by assuming instantaneous coal devolatilization, and volatiles combustion and char gasification by CO2 and H2O in plug flow. The second stage combustor was modeled assuming adiabatic instantaneous gas phase reactions. Of primary interest was the dependence of char gasification efficiency on first stage particle residence time. The influence of first stage stoichiometry, heat loss, coal moisture, coal size distribution, and degree of coal devolatilization on gasifier performance and second stage exhaust temperature was determined. Performance predictions indicate that particle residence times on the order of 500 msec would be required to achieve gasification efficiencies in the range of 90 to 95 percent. The use of a finer coal size distribution significantly reduces the required gasifier residence time for acceptable levels of fuel use efficiency. Residence time requirements are also decreased by increased levels of coal devolatilization. Combustor design efforts should maximize devolatilization by minimizing mixing times associated with coal injection.

  4. Assessment, development and application of combustor aerothermal models

    NASA Technical Reports Server (NTRS)

    Holdeman, J. D.; Mongia, H. C.; Mularz, E. J.

    1988-01-01

    The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamics codes have led to their application to complex 3-D flows such as those in the gas turbine combustor. A number of U.S. Government and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aerothermal Modeling and supporting programs are described.

  5. LEVEL 2 CHEMICAL ANALYSIS OF FLUIDIZED-BED COMBUSTOR SAMPLES

    EPA Science Inventory

    The report gives results of a Level 1 data evaluation and prioritization and the Level 2 environmental assessment (EA) chemical data acquired on a set of fluidized-bed combustor (FBC) particulate samples. The Level 2 analysis followed the approach described in 'Approach to Level ...

  6. Acoustic modal analysis of a full-scale annular combustor

    NASA Technical Reports Server (NTRS)

    Karchmer, A. M.

    1983-01-01

    An acoustic modal decomposition of the measured pressure field in a full scale annular combustor installed in a ducted test rig is described. The modal analysis, utilizing a least squares optimization routine, is facilitated by the assumption of randomly occurring pressure disturbances which generate equal amplitude clockwise and counter-clockwise pressure waves, and the assumption of statistical independence between modes. These assumptions are fully justified by the measured cross spectral phases between the various measurement points. The resultant modal decomposition indicates that higher order modes compose the dominant portion of the combustor pressure spectrum in the range of frequencies of interest in core noise studies. A second major finding is that, over the frequency range of interest, each individual mode which is present exists in virtual isolation over significant portions of the spectrum. Finally, a comparison between the present results and a limited amount of data obtained in an operating turbofan engine with the same combustor is made. The comparison is sufficiently favorable to warrant the conclusion that the structure of the combustor pressure field is preserved between the component facility and the engine. Previously announced in STAR as N83-21896

  7. CFD analysis of jet mixing in low NOx flametube combustors

    NASA Technical Reports Server (NTRS)

    Talpallikar, M. V.; Smith, C. E.; Lai, M. C.; Holdeman, J. D.

    1991-01-01

    The Rich-burn/Quick-mix/Lean-burn (RQL) combustor was identified as a potential gas turbine combustor concept to reduce NO(x) emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NO(x) levels, cylindrical flametube versions of RQL combustors are being tested at NASA Lewis Research Center. A critical technology needed for the RQL combustor is a method of quickly mixing by-pass combustion air with rich-burn gases. Jet mixing in a cylindrical quick-mix section was numerically analyzed. The quick-mix configuration was five inches in diameter and employed twelve radial-inflow slots. The numerical analyses were performed with an advanced, validated 3-D Computational Fluid Dynamics (CFD) code named REFLEQS. Parametric variation of jet-to-mainstream momentum flux ratio (J) and slot aspect ratio was investigated. Both non-reacting and reacting analyses were performed. Results showed mixing and NO(x) emissions to be highly sensitive to J and slot aspect ratio. Lowest NO(x) emissions occurred when the dilution jet penetrated to approximately mid-radius. The viability of using 3-D CFD analyses for optimizing jet mixing was demonstrated.

  8. PERFORMANCE OF EMISSIONS CONTROL SYSTEMS ON MUNICIPAL WASTE COMBUSTORS

    EPA Science Inventory

    The paper reports results of several EPA-supported field evaluations of data on gaseous pollutant emissions from modern municipal waste combustors/incinerators and emissions control by flue gas cleaning systems. The results are presented in terms of acid gas (HCl and SO2), trace ...

  9. DEVELOPMENT OF A VORTEX CONTAINMENT COMBUSTOR FOR COAL COMBUSTION SYSTEMS

    EPA Science Inventory

    The report describes the development of a vortex containment combustor (VCC) for coal combustion systems, designed to solve major problems facing the conversion of oil- and gas-fired boilers to coal (e.g., derating, inorganic impurities in coal, and excessive formation of NOx and...

  10. Coanda injection system for axially staged low emission combustors

    DOEpatents

    Evulet, Andrei Tristan; Varatharajan, Balachandar; Kraemer, Gilbert Otto; ElKady, Ahmed Mostafa; Lacy, Benjamin Paul

    2012-05-15

    The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole. The tertiary coanda nozzles are configured to inject a third fluid comprising air, at least one other diluent, fuel, or combinations thereof to one or more remaining combustion zones among the plurality of combustion zones.

  11. COMBUSTION CONTROL OF ORGANIC EMISSIONS FROM MUNICIPAL WASTE COMBUSTORS

    EPA Science Inventory

    More than two decades ago, researchers identified benzo(a)pyrene and other organic species in the emissions from incineration of solid waste. Chlorinated dibenzo-p-dioxins and-furans (CDD/CDF) were first detected in municipal waste combustor (MWC) emissions in 1977. Since then, C...

  12. Rectangular capture area to circular combustor scramjet engine

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.

    1978-01-01

    A new concept for a scramjet engine design was presented. The inlet transformed a rectangular shaped capture stream into a cross section which was almost circular in shape at the inlet throat or combustor entrance. The inlet inner surface was designed by the method of streamline tracing. The high pressure and temperature regions of the combustor were almost circular in shape and thus the benefits of hoop stresses in relation to structural weight could be utilized to reduce combustor and engine weights. The engine had a center body consisting of a 20 deg included angle cone, followed by a constant diameter cylinder. Fuel injection struts were arranged in a radial array and were swept 54 deg from the center body to the inlet inner surface and had values of length to maximum average thickness between 5.6and 6.6 which were felt to be structurally reasonable. Combustor wetted areas were shown to be less than those of the present fully rectangular engine concept.

  13. MUNICIPAL SOLID WASTE COMBUSTOR ASH DEMONSTRATION PROGRAM - "THE BOATHOUSE"

    EPA Science Inventory

    The report presents the results of a research program designed to examine the engineering and environmental acceptability of using municipal solid waste (MSW) combustor ash as an aggregate substitute in the manufacture of construction quality cement blocks. 50 tons of MSW combust...

  14. Fluidized bed combustor and coal gun-tube assembly therefor

    DOEpatents

    Hosek, William S.; Garruto, Edward J.

    1984-01-01

    A coal supply gun assembly for a fluidized bed combustor which includes heat exchange elements extending above the bed's distributor plate assembly and in which the gun's nozzles are disposed relative to the heat exchange elements to only discharge granular coal material between adjacent heat exchange elements and in a path which is substantially equidistant from adjacent heat exchange elements.

  15. Transient/structural analysis of a combustor under explosive loads

    NASA Technical Reports Server (NTRS)

    Gregory, Peyton B.; Holland, Anne D.

    1992-01-01

    The 8-Foot High Temperature Tunnel (HTT) at NASA Langley Research Center is a combustion-driven blow-down wind tunnel. A major potential failure mode that was considered during the combustor redesign was the possibility of a deflagration and/or detonation in the combustor. If a main burner flame-out were to occur, then unburned fuel gases could accumulate and, if reignited, an explosion could occur. An analysis has been performed to determine the safe operating limits of the combustor under transient explosive loads. The failure criteria was defined and the failure mechanisms were determined for both peak pressures and differential pressure loadings. An overview of the gas dynamics analysis was given. A finite element model was constructed to evaluate 13 transient load cases. The sensitivity of the structure to the frequency content of the transient loading was assessed. In addition, two closed form dynamic analyses were conducted to verify the finite element analysis. It was determined that the differential pressure load or thrust load was the critical load mechanism and that the nozzle is the weak link in the combustor system.

  16. Hydrogen Fuel Capability Added to Combustor Flametube Rig

    NASA Technical Reports Server (NTRS)

    Frankenfield, Bruce J.

    2003-01-01

    Facility capabilities have been expanded at Test Cell 23, Research Combustor Lab (RCL23) at the NASA Glenn Research Center, with a new gaseous hydrogen fuel system. The purpose of this facility is to test a variety of fuel nozzle and flameholder hardware configurations for use in aircraft combustors. Previously, this facility only had jet fuel available to perform these various combustor flametube tests. The new hydrogen fuel system will support the testing and development of aircraft combustors with zero carbon dioxide (CO2) emissions. Research information generated from this test rig includes combustor emissions and performance data via gas sampling probes and emissions measuring equipment. The new gaseous hydrogen system is being supplied from a 70 000-standard-ft3 tube trailer at flow rates up to 0.05 lb/s (maximum). The hydrogen supply pressure is regulated, and the flow is controlled with a -in. remotely operated globe valve. Both a calibrated subsonic venturi and a coriolis mass flowmeter are used to measure flow. Safety concerns required the placement of all hydrogen connections within purge boxes, each of which contains a small nitrogen flow that is vented past a hydrogen detector. If any hydrogen leaks occur, the hydrogen detectors alert the operators and automatically safe the facility. Facility upgrades and modifications were also performed on other fluids systems, including the nitrogen gas, cooling water, and air systems. RCL23 can provide nonvitiated heated air to the research combustor, up to 350 psig at 1200 F and 3.0 lb/s. Significant modernization of the facility control systems and the data acquisition systems was completed. A flexible control architecture was installed that allows quick changes of research configurations. The labor-intensive hardware interface has been removed and changed to a software-based system. In addition, the operation of this facility has been greatly enhanced with new software programming and graphic operator interface

  17. Low Emissions RQL Flametube Combustor Component Test Results

    NASA Technical Reports Server (NTRS)

    Holdeman, James D.; Chang, Clarence T.

    2001-01-01

    This report describes and summarizes elements of the High Speed Research (HSR) Low Emissions Rich burn/Quick mix/Lean burn (RQL) flame tube combustor test program. This test program was performed at NASA Glenn Research Center circa 1992. The overall objective of this test program was to demonstrate and evaluate the capability of the RQL combustor concept for High Speed Civil Transport (HSCT) applications with the goal of achieving NOx emission index levels of 5 g/kg-fuel at representative HSCT supersonic cruise conditions. The specific objectives of the tests reported herein were to investigate component performance of the RQL combustor concept for use in the evolution of ultra-low NOx combustor design tools. Test results indicated that the RQL combustor emissions and performance at simulated supersonic cruise conditions were predominantly sensitive to the quick mixer subcomponent performance and not sensitive to fuel injector performance. Test results also indicated the mixing section configuration employing a single row of circular holes was the lowest NOx mixer tested probably due to the initial fast mixing characteristics of this mixing section. However, other quick mix orifice configurations such as the slanted slot mixer produced substantially lower levels of carbon monoxide emissions most likely due to the enhanced circumferential dispersion of the air addition. Test results also suggested that an optimum momentum-flux ratio exists for a given quick mix configuration. This would cause undesirable jet under- or over-penetration for test conditions with momentum-flux ratios below or above the optimum value. Tests conducted to assess the effect of quick mix flow area indicated that reduction in the quick mix flow area produced lower NOx emissions at reduced residence time, but this had no effect on NOx emissions measured at similar residence time for the configurations tested.

  18. ULTRA LOW NOx CATALYTIC COMBUSTION FOR IGCC POWER PLANTS

    SciTech Connect

    Lance L. Smith

    2004-03-01

    Tests were performed in PCI's sub-scale high-pressure (10 atm) test rig, using PCI's two-stage (catalytic / gas-phase) combustion process for syngas fuel. In this process, the first stage is a Rich-Catalytic Lean-burn (RCL{trademark}) catalytic reactor, wherein a fuel-rich mixture contacts the catalyst and reacts while final and excess combustion air cool the catalyst. The second stage is a gas-phase combustor, wherein the catalyst cooling air mixes with the catalytic reactor effluent to provide for final gas-phase burnout and dilution to fuel-lean combustion products. During the reporting period, PCI successfully achieved NOx = 0.011 lbs/MMBtu at 10 atm pressure (corresponding to 2.0 ppm NOx corrected to 15% O{sub 2} dry) with near-zero CO emissions, surpassing the project goal of < 0.03 lbs/MMBtu NOx. These emissions levels were achieved at scaled (10 atm, sub-scale) baseload conditions corresponding to Tampa Electric's Polk Power Station operation on 100% syngas (no co-firing of natural gas).

  19. Low emissions combustor development for an industrial gas turbine to utilize LCV fuel gas

    SciTech Connect

    Kelsall, G.J.; Smith, M.A. . Coal Research Establishment); Cannon, M.F. . Aero and Technology Products)

    1994-07-01

    Advanced coal-based power generation systems such as the British Coal Topping Cycle offer the potential for high-efficiency electricity generation with minimum environmental impact. An important component of the Topping cycle program is the gas turbine, for which development of a combustion system to burn low calorific value coal derived fuel gas, at a turbine inlet temperature of 1,260 C (2,300 F), with minimum pollutant emissions, is a key R and D issue. A phased combustor development program is underway burning low calorific value fuel gas (3.6--4.1 MJ/m[sup 3]) with low emissions, particularly NO[sub x] derived from fuel-bound nitrogen. The first phase of the combustor development program has now been completed using a generic tubo-annular, prototype combustor design. Tests were carried out at combustor loading and Mach numbers considerably greater than the initial design values. Combustor performance at these conditions was encouraging. The second phase of the program is currently in progress. This will assess, initially, an improved variant of the prototype combustor operating at conditions selected to represent a particular medium sized industrial gas turbine. This combustor will also be capable of operating using natural gas as an auxiliary fuel, to suite the start-up procedure for the Topping Cycle. The paper presents the Phase 1 test program results for the prototype combustor. Design of the modified combustor for Phase 2 of the development program is discussed, together with preliminary combustor performance results.

  20. Adaptive Controls Method Demonstrated for the Active Suppression of Instabilities in Engine Combustors

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2004-01-01

    An adaptive feedback control method was demonstrated that suppresses thermoacoustic instabilities in a liquid-fueled combustor of a type used in aircraft engines. Extensive research has been done to develop lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle to reduce the environmental impact of aerospace propulsion systems. However, these lean-burning combustors are susceptible to thermoacoustic instabilities (high-frequency pressure waves), which can fatigue combustor components and even the downstream turbine blades. This can significantly decrease the safe operating lives of the combustor and turbine. Thus, suppressing the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors under NASA's Propulsion and Power Program. This control methodology has been developed and tested in a partnership of the NASA Glenn Research Center, Pratt & Whitney, United Technologies Research Center, and the Georgia Institute of Technology. Initial combustor rig testing of the controls algorithm was completed during 2002. Subsequently, the test results were analyzed and improvements to the method were incorporated in 2003, which culminated in the final status of this controls algorithm. This control methodology is based on adaptive phase shifting. The combustor pressure oscillations are sensed and phase shifted, and a high-frequency fuel valve is actuated to put pressure oscillations into the combustor to cancel pressure oscillations produced by the instability.

  1. Catalytic distillation structure

    DOEpatents

    Smith, Jr., Lawrence A.

    1984-01-01

    Catalytic distillation structure for use in reaction distillation columns, a providing reaction sites and distillation structure and consisting of a catalyst component and a resilient component intimately associated therewith. The resilient component has at least about 70 volume % open space and being present with the catalyst component in an amount such that the catalytic distillation structure consist of at least 10 volume % open space.

  2. Catalytic cracking process

    DOEpatents

    Lokhandwala, Kaaeid A.; Baker, Richard W.

    2001-01-01

    Processes and apparatus for providing improved catalytic cracking, specifically improved recovery of olefins, LPG or hydrogen from catalytic crackers. The improvement is achieved by passing part of the wet gas stream across membranes selective in favor of light hydrocarbons over hydrogen.

  3. Computational Analysis of Dynamic SPK(S8)-JP8 Fueled Combustor-Sector Performance

    NASA Technical Reports Server (NTRS)

    Ryder, R.; Hendricks, Roberts C.; Huber, M. L.; Shouse, D. T.

    2010-01-01

    Civil and military flight tests using blends of synthetic and biomass fueling with jet fuel up to 50:50 are currently considered as "drop-in" fuels. They are fully compatible with aircraft performance, emissions and fueling systems, yet the design and operations of such fueling systems and combustors must be capable of running fuels from a range of feedstock sources. This paper provides Smart Combustor or Fuel Flexible Combustor designers with computational tools, preliminary performance, emissions and particulates combustor sector data. The baseline fuel is kerosene-JP-8+100 (military) or Jet A (civil). Results for synthetic paraffinic kerosene (SPK) fuel blends show little change with respect to baseline performance, yet do show lower emissions. The evolution of a validated combustor design procedure is fundamental to the development of dynamic fueling of combustor systems for gas turbine engines that comply with multiple feedstock sources satisfying both new and legacy systems.

  4. Effects of operating pressure on flame oscillation and emission characteristics in a partially premixed swirl combustor

    SciTech Connect

    Kim, Jong-Ryul; Choi, Gyung-Min; Kim, Duck-Jool

    2011-01-15

    The influence of varying combustor pressure on flame oscillation and emission characteristics in the partially premixed turbulent flame were investigated. In order to investigate combustion characteristics in the partially premixed turbulent flame, the combustor pressure was controlled in the range of -30 to 30 kPa for each equivalence ratio ({phi} = 0.8-1.2). The r.m.s. of the pressure fluctuations increased with decreasing combustor pressure for the lean condition. The combustor pressure had a sizeable influence on combustion oscillation, whose dominant frequency varied with the combustor pressure. Combustion instabilities could be controlled by increasing the turbulent intensity of the unburned mixture under the lean condition. An unstable flame was caused by incomplete combustion; hence, EICO greatly increased. Furthermore, EINO{sub x} simply reduced with decreasing combustor pressure at a rate of 0.035 g/10 kPa. The possibility of combustion control on the combusting mode and exhaust gas emission was demonstrated. (author)

  5. An Adaptive Instability Suppression Controls Method for Aircraft Gas Turbine Engine Combustors

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; DeLaat, John C.; Chang, Clarence T.

    2008-01-01

    An adaptive controls method for instability suppression in gas turbine engine combustors has been developed and successfully tested with a realistic aircraft engine combustor rig. This testing was part of a program that demonstrated, for the first time, successful active combustor instability control in an aircraft gas turbine engine-like environment. The controls method is called Adaptive Sliding Phasor Averaged Control. Testing of the control method has been conducted in an experimental rig with different configurations designed to simulate combustors with instabilities of about 530 and 315 Hz. Results demonstrate the effectiveness of this method in suppressing combustor instabilities. In addition, a dramatic improvement in suppression of the instability was achieved by focusing control on the second harmonic of the instability. This is believed to be due to a phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling. These results may have implications for future research in combustor instability control.

  6. Innovative Adaptive Control Method Demonstrated for Active Suppression of Instabilities in Engine Combustors

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2005-01-01

    This year, an improved adaptive-feedback control method was demonstrated that suppresses thermoacoustic instabilities in a liquid-fueled combustor of a type used in aircraft engines. Extensive research has been done to develop lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle to reduce the environmental impact of aerospace propulsion systems. However, these lean-burning combustors are susceptible to thermoacoustic instabilities (high-frequency pressure waves), which can fatigue combustor components and even downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppressing the thermoacoustic combustor instabilities is an enabling technology for meeting the low-emission goals of the NASA Ultra-Efficient Engine Technology (UEET) Project.

  7. Energy efficient engine pin fin and ceramic composite segmented liner combustor sector rig test report

    NASA Technical Reports Server (NTRS)

    Dubiel, D. J.; Lohmann, R. P.; Tanrikut, S.; Morris, P. M.

    1986-01-01

    Under the NASA-sponsored Energy Efficient Engine program, Pratt and Whitney has successfully completed a comprehensive test program using a 90-degree sector combustor rig that featured an advanced two-stage combustor with a succession of advanced segmented liners. Building on the successful characteristics of the first generation counter-parallel Finwall cooled segmented liner, design features of an improved performance metallic segmented liner were substantiated through representative high pressure and temperature testing in a combustor atmosphere. This second generation liner was substantially lighter and lower in cost than the predecessor configuration. The final test in this series provided an evaluation of ceramic composite liner segments in a representative combustor environment. It was demonstrated that the unique properties of ceramic composites, low density, high fracture toughness, and thermal fatigue resistance can be advantageously exploited in high temperature components. Overall, this Combustor Section Rig Test program has provided a firm basis for the design of advanced combustor liners.

  8. Primary zone dynamics in a gas turbine combustor

    NASA Astrophysics Data System (ADS)

    Sullivan, J. P.; Barron, D.; Seal, M.; Morgan, D.; Murthy, S. N. B.

    1989-01-01

    Fluid mechanical investigations simulating the flow in the primary zone of a gas turbine combustor are presented using three generic test rigs: (1) rotating pipe yielding a swirling jet of air; (2) primary zone model with a single swirler and various primary jet configurations, operated with air; and (3) two rectangular models of a (stretched-out) annular combustor with five swirlers in the backwall and with various primary jet configurations, one operated with air and the other with water. Concentration measurements are obtained using laser sheet imaging techniques and velocity measurements using a laser Doppler velocimeter. The results show recirculation zones, intense mixing, instabilities of the interacting jets and the presence of large random vortical motions. The flowfields are shown to exhibit bimodal behavior, have asymmetries despite symmetrical geometry and inlet conditions and display strong jet/swirler and swirler/swirler interactions.

  9. Measurements in an annular combustor-diffuser system

    NASA Astrophysics Data System (ADS)

    Srinivasan, R.; Freeman, W. G.; Mozumdar, S.; Grahmann, J. W.

    1990-07-01

    Results of three-component Laser Doppler Velocimeter (LDV) measurements in an annular combustor-diffuser system are presented in this paper. The LDV measurements were made at several locations in the prediffuser and dump regions of the combustor-diffuser test rig for three different inlet velocity profiles. The prediffuser average inlet Mach number was maintained at 0.305 during these tests. The LDV data are compared with predictions obtained from a boundary-fitted, two-dimensional elliptic analytical model. The agreement between the LDV data and the predicted results is very good for mean velocities. However, measured turbulence intensities are higher than the predicted values in the region adjacent to the prediffuser walls.

  10. Experimental results from a reverse flow annual combustor

    NASA Astrophysics Data System (ADS)

    Joubert, F. M.; Hattingh, H. V.

    Computer-predicted temperature distributions in the wall liners of a combustion chamber were compared to the experimentally obtained values from combustion tests carried out in a small, full-scale reverse-flow annular combustor at sea level take-off conditionns. The largest discrepancies between the measured and predicted linear temperatures occured in the primary zone, with most of the predictions falling above the measured values, and with neither of the two computer programs satisfying the accuracy of 4 percent (of the experimental values) needed for making estimates on the life of a combustor. On the other hand, the correlation between the measured and predicted liner pressure drop was satisfactory. The validity and usefulnes of simple computer models as aids in the design of gas turbine combustion chambers are discussed.

  11. Fuel injection assembly for gas turbine engine combustor

    NASA Technical Reports Server (NTRS)

    Candy, Anthony J. (Inventor); Glynn, Christopher C. (Inventor); Barrett, John E. (Inventor)

    2002-01-01

    A fuel injection assembly for a gas turbine engine combustor, including at least one fuel stem, a plurality of concentrically disposed tubes positioned within each fuel stem, wherein a cooling supply flow passage, a cooling return flow passage, and a tip fuel flow passage are defined thereby, and at least one fuel tip assembly connected to each fuel stem so as to be in flow communication with the flow passages, wherein an active cooling circuit for each fuel stem and fuel tip assembly is maintained by providing all active fuel through the cooling supply flow passage and the cooling return flow passage during each stage of combustor operation. The fuel flowing through the active cooling circuit is then collected so that a predetermined portion thereof is provided to the tip fuel flow passage for injection by the fuel tip assembly.

  12. A mathematical model for jet engine combustor pollutant emissions

    NASA Technical Reports Server (NTRS)

    Boccio, J. L.; Weilerstein, G.; Edelman, R. B.

    1973-01-01

    Mathematical modeling for the description of the origin and disposition of combustion-generated pollutants in gas turbines is presented. A unified model in modular form is proposed which includes kinetics, recirculation, turbulent mixing, multiphase flow effects, swirl and secondary air injection. Subelements of the overall model were applied to data relevant to laboratory reactors and practical combustor configurations. Comparisons between the theory and available data show excellent agreement for basic CO/H2/Air chemical systems. For hydrocarbons the trends are predicted well including higher-than-equilibrium NO levels within the fuel rich regime. Although the need for improved accuracy in fuel rich combustion is indicated, comparisons with actual jet engine data in terms of the effect of combustor-inlet temperature is excellent. In addition, excellent agreement with data is obtained regarding reduced NO emissions with water droplet and steam injection.

  13. Modeling scramjet combustor flowfields with a grid adaptation scheme

    NASA Technical Reports Server (NTRS)

    Ramakrishnan, R.; Singh, D. J.

    1994-01-01

    The accurate description of flow features associated with the normal injection of fuel into supersonic primary flows is essential in the design of efficient engines for hypervelocity aerospace vehicles. The flow features in such injections are complex with multiple interactions between shocks and between shocks boundary layers. Numerical studies of perpendicular sonic N2 injection and mixing in a Mach 3.8 scramjet combustor environment are discussed. A dynamic grid adaptation procedure based on the equilibration of spring-mass system is employed to enhanced the description of the complicated flow features. Numerical results are compared with experimental measurements and indicate that the adaptation procedure enhances the capability of the modeling procedure to describe the flow features associated with scramjet combustor components.

  14. Numerical Study of Low Emission Gas Turbine Combustor Concepts

    NASA Technical Reports Server (NTRS)

    Yang, Song-Lin

    2002-01-01

    To further reduce pollutant emissions, such as CO, NO(x), UHCs, etc., in the next few decades, innovative concepts of gas turbine combustors must be developed. Several concepts, such as the LIPP (Lean- Premixed- Prevaporized), RQL (Rich-Burn Quick-Quench Lean-Burn), and LDI (Lean-Direct-Injection), have been under study for many years. To fully realize the potential of these concepts, several improvements, such as inlet geometry, air swirler, aerothermochemistry control, fuel preparation, fuel injection and injector design, etc., must be made, which can be studied through the experimental method and/or the numerical technique. The purpose of this proposal is to use the CFD technique to study, and hence, to guide the design process for low emission gas turbine combustors. A total of 13 technical papers have been (or will be) published.

  15. Effect of fuel vapor concentrations on combustor emissions and performance

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Ingebo, R. D.

    1973-01-01

    Effects of fuel vaporization on the exhaust emission levels of oxides of nitrogen, carbon monoxide, total hydrocarbons, and smoke number were obtained in an experimental turbojet combustor segment. Two different fuel injectors were used in which liquid ASTM A-1 jet fuel and vapor propane fuel were independently controlled to simulate varying degrees of vaporization. Tests were conducted over a range of inlet-air temperatures from 478 to 700 K, pressures from 4 to 20 atm, and combustor reference velocities from 15.3 to 27.4 m/sec. Converting from liquid to complete vapor fuel resulted in oxides of nitrogen reductions of as much as 22 percent and smoke number reductions up to 51 percent. Supplement data are also presented on flame emissivity, flame temperature, and primary-zone liner wall temperatures.

  16. Reduction of nitric oxide emissions from a combustor

    NASA Technical Reports Server (NTRS)

    Craig, R. A.; Pritchard, H. O. (Inventor)

    1980-01-01

    A turbojet combustor and method for controlling nitric oxide emissions by employing successive combustion zones is described. After combustion of an initial portion of the fuel in a primary combustion zone, the combustion products of the primary zone are combined with the remaining portion of fuel and additional plenum air and burned in a secondary combustion zone under conditions that result in low nitric oxide emissions. Low nitric oxide emissions are achieved by a novel turbojet combustor arrangement which provides flame stability by allowing stable combustion to be accompanied by low nitric oxide emissions resulting from controlled fuel-lean combustion (ignited by the emission products from the primary zone) in a secondary combustion zone at a lower combustion temperature resulting in low emission of nitric oxide.

  17. Conceptual model of turbulent flameholding for scramjet combustors

    NASA Technical Reports Server (NTRS)

    Huber, P. W.

    1980-01-01

    New concepts and approaches to scramjet combustor design are presented. Blowoff was from failure of the recirculation-zone (RZ) flame to reach the dividing streamline (DS) at the rear stagnation zone. Increased turbulent exchange across the DS helped flameholding due to forward movement of the flame anchor point inside the RZ. Modeling of the blowoff phenomenon was based on a mass conservation concept involving the traverse of a flame element across the RZ and a flow element along the DS. The scale required to achieve flameholding, predicted by the model, showed a strong adverse effect of low pressure and low fuel equivalence ratio, moderate effect of flight Mach number, and little effect of temperature recovery factor. Possible effects of finite rate chemistry on flameholding and flamespreading in scramjets are discussed and recommendations for approaches to engine combustor design as well as for needed research to reduce uncertainties in the concepts are made.

  18. Analytical fuel property effects, small combustors, phase 1

    NASA Technical Reports Server (NTRS)

    Cohen, J. D.

    1983-01-01

    The effects of nonstandard aviation fuels on a typical small gas turbine combustor was analyzed. The T700/CT7 engine family was chosen as being representative of the class of aircraft power plants desired. Fuel properties, as specified by NASA, are characterized by low hydrogen content and high aromatics levels. Higher than normal smoke output and flame radiation intensity for the current T700 combustor which serves as a baseline were anticipated. It is, therefore, predicted that out of specification smoke visibility and higher than normal shell temperatures will exist when using NASA ERBS fuels with a consequence of severe reduction in cyclic life. Three new designs are proposed to compensate for the deficiencies expected with the existing design. They have emerged as the best of the eight originally proposed redesigns or combinations thereof. After the five choices that were originally made by NASA on the basis of competing performance factors, General Electric narrowed the field to the three proposed.

  19. Reliable and Affordable Control Systems Active Combustor Pattern Factor Control

    NASA Technical Reports Server (NTRS)

    McCarty, Bob; Tomondi, Chris; McGinley, Ray

    2004-01-01

    Active, closed-loop control of combustor pattern factor is a cooperative effort between Honeywell (formerly AlliedSignal) Engines and Systems and the NASA Glenn Research Center to reduce emissions and turbine-stator vane temperature variations, thereby enhancing engine performance and life, and reducing direct operating costs. Total fuel flow supplied to the engine is established by the speed/power control, but the distribution to individual atomizers will be controlled by the Active Combustor Pattern Factor Control (ACPFC). This system consist of three major components: multiple, thin-film sensors located on the turbine-stator vanes; fuel-flow modulators for individual atomizers; and control logic and algorithms within the electronic control.

  20. Fuel-Flexible Gas Turbine Combustor Flametube Facility

    NASA Technical Reports Server (NTRS)

    Little, James E.; Nemets, Stephen A.; Tornabene, Robert T.; Smith, Timothy D.; Frankenfield, Bruce J.; Manning, Stephen D.; Thompson, William K.

    2004-01-01

    Facility modifications have been completed to an existing combustor flametube facility to enable testing with gaseous hydrogen propellants at the NASA Glenn Research Center. The purpose of the facility is to test a variety of fuel nozzle and flameholder hardware configurations for use in aircraft combustors. Facility capabilities have been expanded to include testing with gaseous hydrogen, along with the existing hydrocarbon-based jet fuel. Modifications have also been made to the facility air supply to provide heated air up to 350 psig, 1100 F, and 3.0 lbm/s. The facility can accommodate a wide variety of flametube and fuel nozzle configurations. Emissions and performance data are obtained via a variety of gas sample probe configurations and emissions measurement equipment.

  1. Lean, premixed, prevaporized fuel combustor conceptual design study

    NASA Technical Reports Server (NTRS)

    Fiorentino, A. J.; Greene, W.; Kim, J.

    1979-01-01

    Four combustor concepts, designed for the energy efficient engine, utilize variable geometry or other flow modulation techniques to control the equivalence ratio of the initial burning zone. Lean conditions are maintained at high power to control oxides of nitrogen while near stoichometric conditions are maintained at low power for low CO and THC emissions. Each concept was analyzed and ranked for its potential in meeting the goals of the program. Although the primary goal of the program is a low level of nitric oxide emissions at stratospheric cruise conditions, both the ground level EPA emission standards and combustor performance and operational requirements typical of advanced subsonic aircraft engines are retained as goals as well. Based on the analytical projections made, two of the concepts offer the potential of achieving the emission goals; however, the projected operational characteristics and reliability of any concept to perform satisfactorily over an entire aircraft flight envelope would require extensive experimental substantiation before engine adaptation can be considered.

  2. A review of NASA combustor and turbine heat transfer research

    NASA Technical Reports Server (NTRS)

    Rudey, R. A.; Graham, R. W.

    1984-01-01

    The thermal design of the combustor and turbine of a gas turbine engine poses a number of difficult heat transfer problems. The importance of improved prediction techniques becomes more critical in anticipation of future generations of gas turbine engines which will operate at higher cycle pressure and temperatures. Research which addresses many of the complex heat transfer processes holds promise for yielding significant improvements in prediction of metal temperatures. Such research involves several kinds of program including: (1) basic experiments which delineate the fundamental flow and heat transfer phenomena that occur in the hot sections of the gas turbine but at low enthalpy conditions; (2) analytical modeling of these flow and heat transfer phenomena which results from the physical insights gained in experimental research; and (3) verification of advanced prediction techniques in facilities which operate near the real engine thermodynamic conditions. In this paper, key elements of the NASA program which involves turbine and combustor heat transfer research will be described and discussed.

  3. A study of the pulsation driving mechanism in pulsating combustors

    NASA Astrophysics Data System (ADS)

    Goldman, Y.; Timnat, Y. M.

    Experiments performed in a facility consisting of a Schmidt-type pulsating combustor, in which high-speed photographs were taken and pressure, temperature and gas composition measured, showed that the air supply conditions at the inlet and the volume of the combustor strongly influence the oscillation frequency. From the measurements, the existence of two separate regions, one containing cold air and the other containing fuel-rich gas, was found, and a pressure-volume diagram was drawn, showing the effect of chemical energy release and heat supply during the compression stroke and differentiating it from the expansion. A model of the interaction between the cyclic combustion process and the acoustic oscillations of the gas volume within the chamber and the tail-pipe is presented. The conditions for chemical energy release that result in high-pressure amplitude are described.

  4. Self-regulating fuel staging port for turbine combustor

    DOEpatents

    Van Nieuwenhuizen, William F.; Fox, Timothy A.; Williams, Steven

    2014-07-08

    A port (60) for axially staging fuel and air into a combustion gas flow path 28 of a turbine combustor (10A). A port enclosure (63) forms an air path through a combustor wall (30). Fuel injectors (64) in the enclosure provide convergent fuel streams (72) that oppose each other, thus converting velocity pressure to static pressure. This forms a flow stagnation zone (74) that acts as a valve on airflow (40, 41) through the port, in which the air outflow (41) is inversely proportion to the fuel flow (25). The fuel flow rate is controlled (65) in proportion to engine load. At high loads, more fuel and less air flow through the port, making more air available to the premixing assemblies (36).

  5. Design and evaluation of combustors for reducing aircraft engine pollution

    NASA Technical Reports Server (NTRS)

    Jones, R. E.; Grobman, J.

    1973-01-01

    Efforts in reducing exhaust emissions from turbine engines are reported. Various techniques employed and the results of testing are briefly described and referenced for detail. The experimental approaches taken to reduce oxides of nitrogen emissions include the use of: (1) multizone combustors incorporating reduced dwell times, (2) fuel-air premixing, (3) air atomization, (4) fuel prevaporization, and (5) gaseous fuel. Since emissions of unburned hydrocarbons and carbon monoxide are caused by poor combustion efficiency at engine idle, the studies of fuel staging in multizone combustors and air assist fuel nozzles have indicated that large reductions in these emissions can be achieved. Also, the effect of inlet-air humidity on oxides of nitrogen was studied as well as the very effective technique of direct water injection. The emission characteristics of natural gas and propane fuels were measured and compared with those of ASTM-Al kerosene fuel.

  6. System for supporting bundled tube segments within a combustor

    DOEpatents

    Melton, Patrick Benedict

    2016-03-01

    A system for supporting bundled tube segments within a combustor includes an annular sleeve that extends circumferentially and axially within the combustor, a support lug that extends radially inward from the annular sleeve and an annular support frame that is disposed within the annular sleeve. The annular support frame includes an inner ring portion, an outer ring portion and a plurality of spokes that extend radially between the inner and outer ring portions. The inner ring portion, the outer ring portion and the plurality of spokes define an annular array of openings for receiving a respective bundled tube segment. The inner ring portion is connected to each bundled tube segment and the outer ring portion is coupled to the support lug.

  7. System and method for reducing combustion dynamics in a combustor

    SciTech Connect

    Uhm, Jong Ho; Johnson, Thomas Edward; Zuo, Baifang; York, William David

    2015-09-01

    A system for reducing combustion dynamics in a combustor includes an end cap having an upstream surface axially separated from a downstream surface, and tube bundles extend from the upstream surface through the downstream surface. A divider inside a tube bundle defines a diluent passage that extends axially through the downstream surface, and a diluent supply in fluid communication with the divider provides diluent flow to the diluent passage. A method for reducing combustion dynamics in a combustor includes flowing a fuel through tube bundles, flowing a diluent through a diluent passage inside a tube bundle, wherein the diluent passage extends axially through at least a portion of the end cap into a combustion chamber, and forming a diluent barrier in the combustion chamber between the tube bundle and at least one other adjacent tube bundle.

  8. System and method for reducing combustion dynamics in a combustor

    SciTech Connect

    Uhm, Jong Ho; Johnson, Thomas Edward; Zuo, Baifang; York, William David

    2013-08-20

    A system for reducing combustion dynamics in a combustor includes an end cap having an upstream surface axially separated from a downstream surface, and tube bundles extend through the end cap. A diluent supply in fluid communication with the end cap provides diluent flow to the end cap. Diluent distributors circumferentially arranged inside at least one tube bundle extend downstream from the downstream surface and provide fluid communication for the diluent flow through the end cap. A method for reducing combustion dynamics in a combustor includes flowing fuel through tube bundles that extend axially through an end cap, flowing a diluent through diluent distributors into a combustion chamber, wherein the diluent distributors are circumferentially arranged inside at least one tube bundle and each diluent distributor extends downstream from the end cap, and forming a diluent barrier in the combustion chamber between at least one pair of adjacent tube bundles.

  9. Calculation of two-phase flow in gas turbine combustors

    SciTech Connect

    Tolpadi, A.K.

    1995-10-01

    A method is presented for computing steady two-phase turbulent combusting flow in a gas turbine combustor. The gas phase equations are solved in an Eulerian frame of reference. The two-phase calculations are performed by using a liquid droplet spray combustion a model and treating the motion of the evaporating fuel droplets in a Lagrangian frame of reference. The numerical algorithm employs nonorthogonal curvilinear coordinates, a multigrid iterative solution procedure, the standard k-{epsilon} turbulence model, and a combustion model comprising an assumed shape probability density function and the conserved scalar formulation. The trajectory computation of the fuel provides the source terms for all the gas phase equations. This two-phase model was applied to a real piece of combustion hardware in the form of a modern GE/SNECMA single annular CFM56 turbofan engine combustor. For the purposes of comparison, calculations were also performed by treating the fuel as a single gaseous phase. The effect on the solution of two extreme situations of the fuel as a gas and initially as a liquid was examined. The distribution of the velocity field and the conserved scalar within the combustor, as well as the distribution of the temperature field in the reaction zone and in the exhaust, were all predicted with the combustor operating both at high-power and low-power (ground idle) conditions. The calculated exit gas temperature was compared with test rig measurements. Under both low and high-power conditions, the temperature appeared to show an improved agreement with the measured data when the calculations were performed with the spray model as compared to a single-phase calculation.

  10. Lean, Premixed-Prevaporized (LPP) combustor conceptual design study

    NASA Technical Reports Server (NTRS)

    Dickman, R. A.; Dodds, W. J.; Ekstedt, E. E.

    1979-01-01

    Four combustion systems were designed and sized for the energy efficient engine. A fifth combustor was designed for the cycle and envelope of the twin-spool, high bypass ratio, high pressure ratio turbofan engine. Emission levels, combustion performance, life, and reliability assessments were made for these five combustion systems. Results of these design studies indicate that cruise NOx emission can be reduced by the use of lean, premixed-prevaporaized combustion and airflow modulation.

  11. Advanced composite combustor structural concepts program. Final Report

    SciTech Connect

    Sattar, M.A.; Lohmann, R.P.

    1984-12-01

    An analytical study was conducted to assess the feasibility of and benefits derived from the use of high temperature composite materials in aircraft turbine engine combustor liners. The study included a survey and screening of the properties of three candidate composite materials including tungsten reinforced superalloys, carbon-carbon and silicon carbide (SiC) fibers reinforcing a ceramic matrix of lithium aluminosilicate (LAS). The SiC-LAS material was selected as offering the greatest near term potential primarily on the basis of high temperature capability. A limited experimental investigation was conducted to quantify some of the more critical mechanical properties of the SiC-LAS composite having a multidirection 0/45/-45/90 deg fiber orientation favored for the combustor linear application. Rigorous cyclic thermal tests demonstrated that SiC-LAS was extremely resistant to the thermal fatigue mechanisms that usually limit the life of metallic combustor liners. A thermal design study led to the definition of a composite liner concept that incorporated film cooled SiC-LAS shingles mounted on a Hastelloy X shell. With coolant fluxes consistent with the most advanced metallic liner technology, the calculated hot surface temperatures of the shingles were within the apparent near term capability of the material. Structural analyses indicated that the stresses in the composite panels were low, primarily because of the low coefficient of expansion of the material and it was concluded that the dominant failure mode of the liner would be an as yet unidentified deterioration of the composite from prolonged exposure to high temperature. An economic study, based on a medium thrust size commercial aircraft engine, indicated that the SiC-LAS combustor liner would weigh 22.8N (11.27 lb) less and cost less to manufacture than advanced metallic liner concepts intended for use in the late 1980's.

  12. Pulsed atmospheric fluidized bed combustor apparatus and process

    DOEpatents

    Mansour, Momtaz N.

    1992-01-01

    A pulsed atmospheric fluidized bed reactor system is disclosed and claimed along with a process for utilization of same for the combustion of, e.g. high sulfur content coal. The system affords a economical, ecologically acceptable alternative to oil and gas fired combustors. The apparatus may also be employed for endothermic reaction, combustion of waste products, e.g. organic and medical waste, drying, calcining and the like.

  13. Development of a retrofit coal combustor for industrial applications

    SciTech Connect

    Not Available

    1993-01-01

    During this quarter the tandem pulse combustors were assembled and several definition-start-up tests were conducted on both single units and the tandem unit. The start-up tests indicated that several configuration modifications were required before the evaluation tests were initiated. The modifications were completed and both base-line performance for the single unit and the initial tests of the tandem unit were completed.

  14. Numerical modeling of combustion dynamics in a lean premixed combustor

    SciTech Connect

    Cannon, S.M.; Smith, C.E.

    1998-07-01

    The objective of this study was to evaluate the ability of a time-accurate, 2-D axi-symmetric CFD model to accurately predict combustion dynamics in a premixed pipe combustor driven by mixture feed variation. Independently measured data, including the magnitude and frequency of combustor pressure, were used to evaluate the model. The Smagorinsky, RGN k-{var{underscore}epsilon}, and molecular viscosity models were used to describe the subgrid turbulence, and a one-step, finite-rate reaction to equilibrium products model was used to describe the subgrid chemistry. Swirl source terms were included within the premix passage's computational domain and allowed the model to retain known boundary conditions at the choked flow inlet and the constant pressure exit. To ensure pressure waves were accurately captured, 1-D numerical analyses were first performed to assess the effects of boundary conditions, temporal and spatial differencing, time step, and grid size. It was found that the selected numerical details produced little numerical dissipation of the pressure waves. Then, 2-D axisymmetric analyses were performed in which the inlet temperature was varied. It was found that increases in the inlet temperature (keeping a constant mass flow rate) had a large effect on the unsteady combustor behavior since reaction and advection rates were increased. The correct trend of decreasing rms pressures with increasing inlet temperature was predicted. This agreement in rms pressure behavior supports the ability of the CFD model to accurately capture unsteady heat release and its coupling with resonant acoustic waves in multi-dimensional combustor systems. The effect of subgrid turbulence model was small for the unstable cases studied here.

  15. Examination of mode shapes in an unstable model combustor

    NASA Astrophysics Data System (ADS)

    Sisco, J. C.; Yu, Y. C.; Sankaran, V.; Anderson, W. E.

    2011-01-01

    The coupling between the fluid dynamics, heat addition, and the acoustics of a combustor system determine whether it is prone toward combustion instability. This paper presents results from a benchmark study of the eigenmodes in an unstable experimental combustor. The axisymmetric combustor configuration is representative of a number of practical systems and comprises an injector tube, geometric expansion into a combustion chamber, and a short converging nozzle. Instability limit cycle amplitudes ranged from 5% to nearly 50% of the mean 2.2 MPa pressure. Multiple harmonics were measured for the highly unstable cases. The model combustor was designed to provide a fairly comprehensive set of tested effects: sonic vs subsonic inlets; oxidizer tube lengths that were either quarter-wave, half-wave, or off-resonant acoustic equivalents to the combustion chamber; a significant injector mean flow with Ma˜0.4; and a varied combustion chamber length. The measured mode shape data were analyzed and reduced to provide comparison with results from a linearized one-dimensional Euler model, which included the effects of real boundary conditions, entropy generation, area change, and heat and mass addition, but did not include a model for unsteady heat addition. For low-amplitude instabilities, the measured resonance frequencies agreed with those calculated by the model for the injector tube-combustion chamber system. Resonance frequencies for the high-amplitude oscillation cases corresponded to the first longitudinal frequency of the combustion chamber and its integer multiples. Good quantitative agreement was obtained between computed and measured phase difference profiles, and mode envelopes agreed qualitatively. These results provide a basis for subsequent combustion response studies on the effects of unsteady heat addition.

  16. YF 102 in-duct combustor noise measurements with a turbine nozzle, volume 1

    NASA Astrophysics Data System (ADS)

    Wilson, C. A.; Oconnell, J. M.

    1981-09-01

    The internal noise generated by an Avco Lycoming YF-102 engine combustor installed in a test rig was recorded. Two configurations were tested one with and one without the first stage turbine nozzle installed. Acoustic probes and accessories were used. Internal dynamic pressure level measurements were made at ten locations within the combustor. The combustor rig, the test procedures, and data acquisition and reduction systems are described. Tables and plots of narrow band and one third octave band pressure level spectra are included.

  17. Numerical Prediction of Non-Reacting and Reacting Flow in a Model Gas Turbine Combustor

    NASA Technical Reports Server (NTRS)

    Davoudzadeh, Farhad; Liu, Nan-Suey

    2005-01-01

    The three-dimensional, viscous, turbulent, reacting and non-reacting flow characteristics of a model gas turbine combustor operating on air/methane are simulated via an unstructured and massively parallel Reynolds-Averaged Navier-Stokes (RANS) code. This serves to demonstrate the capabilities of the code for design and analysis of real combustor engines. The effects of some design features of combustors are examined. In addition, the computed results are validated against experimental data.

  18. YF 102 in-duct combustor noise measurements with a turbine nozzle, volume 1

    NASA Technical Reports Server (NTRS)

    Wilson, C. A.; Oconnell, J. M.

    1981-01-01

    The internal noise generated by an Avco Lycoming YF-102 engine combustor installed in a test rig was recorded. Two configurations were tested one with and one without the first stage turbine nozzle installed. Acoustic probes and accessories were used. Internal dynamic pressure level measurements were made at ten locations within the combustor. The combustor rig, the test procedures, and data acquisition and reduction systems are described. Tables and plots of narrow band and one third octave band pressure level spectra are included.

  19. Performance and Pollution Measurements of Two-Row Swirl-Can Combustor Having 72 Modules

    NASA Technical Reports Server (NTRS)

    Biaglow, James A.; Trout, Arthur M.

    1975-01-01

    A test program was conducted to evaluate the performance and gaseous-pollutant levels of an experimental full-annulus 72-module swirl-can combustor. A comparison of data with those for a 120-module swirl-can combustor showed no significant difference in performance or levels of gaseous pollutants. Oxides of nitrogen were correlated for the 72- and 120-swirl-can combustors by using a previously developed parameter.

  20. Experimental investigation of the low NOx vortex airblast annular combustor

    NASA Technical Reports Server (NTRS)

    Johnson, S. M.; Biaglow, J. A.; Smith, J. M.

    1984-01-01

    A low oxides of nitrogen vortex airblast annular combustor was evaluated which has attained the goal of 1 gm NO2/kg fuel or less during operation. The experimental combustor test conditions were a nominal inlet-air temperature of 703 K, inlet total pressures between 0.52 to 0.83 MPa, and a constant inlet Mach number of 0.26. Exit temperature pattern factors for all test points were between 0.16 and 0.20 and exit swirl flow angles were 47 degrees at isothermal conditions and 23 degrees during combustion. Oxides of nitrogen did not exceed 1.05 gm NO2/kg fuel at the highest inlet pressure and exhaust temperature tested. Previous correlations have related NOx proportionally to the combustor inlet pressure raised to some exponent. In this experiment, a band of exponents between 0.5 and 1.0 resulted for fuel-air ratios from 0.023 to 0.027 and inlet pressures from 0.52 to 0.83 MPa. Previously announced in STAR as N84-22567

  1. Combustion of oil palm solid wastes in fluidized bed combustor

    SciTech Connect

    Shamsuddin, A.H.; Sopian, K.

    1995-12-31

    The palm oil industry of Malaysia is the largest in the world producing about 55% of the world production. The industry has approximately 270 mills throughout the country with processing sizes ranging from 10 tonnes/hour to 120 tonnes/hour. All mills produce solid wastes, about 50% of the fresh fruit bunches in terms of weight. The solid wastes produced are in the form of empty fruit bunches, fibers and shells. These wastes have high energy value, ranging from 14 to 18 MJ/kg. The industry is currently self-sufficient in terms of energy. Fibers and shell wastes are being used as boiler fuel to raise steam for electrical power production and process steam. However, the combustion technology currently being employed is obsolete with low efficiency and polluting. A fluidized bed combustor pilot plant is designed and constructed at Combustion Research Laboratory, Universiti Kebangsaan Malaysia. The combustor is made up of 600 mm {times} 900 mm rectangular bed filled with sand up to 400 mm height, static. A bank of heat transfer tubes is imbedded in the bed, designed to absorb 50% of heat released by the fuel in the bed. The remaining heat is transferred in tubes placed on the wall of the freeboard area. Experimental studies were carried out in the pilot plant using palm oil solid wastes. The combustion temperatures were maintained in the range 800--900 C. The performance of the combustor was evaluated in terms of combustion and boiler efficiencies and flue gas emissions monitored.

  2. Characteristics of a trapped-vortex (TV) combustor

    NASA Technical Reports Server (NTRS)

    Hsu, K.-Y.; Gross, L. P.; Trump, D. D.; Roquemore, W. M.

    1994-01-01

    The characteristics of a Trapped-Vortex (TV) combustor are presented. A vortex is trapped in the cavity established between two disks mounted in tandem. Fuel and air are injected directly into the cavity in such a way as to increase the vortex strength. Some air from the annular flow is also entrained into the recirculation zone of the vortex. Lean blow-out limits of the combustor are determined for a wide range of annular air flow rates. These data indicate that the lean blow-out limits are considerably lower for the TV combustor than for flames stabilized using swirl or bluff-bodies. The pressure loss through the annular duct is also low, being less than 2% for the flow conditions in this study. The instantaneous shape of the recirculation zone of the trapped vortex is measured using a two-color PIV technique. Temperature profiles obtained with CARS indicate a well mixed recirculation zone and demonstrate the impact of primary air injection on the local equivalence ratio.

  3. NASA Lewis Research Center's combustor test facilities and capabilities

    NASA Technical Reports Server (NTRS)

    Bianco, Jean

    1995-01-01

    NASA Lewis Research Center (LeRC) presently accommodates a total of six combustor test facilities with unique capabilities. The facilities are used to evaluate combustor and afterburner concepts for future engine applications, and also to test the survivability and performance of innovative high temperature materials, new instrumentation, and engine components in a realistic jet engine environment. The facilities provide a variety of test section interfaces and lengths to allow for flametube, sector and component testing. The facilities can accommodate a wide range of operating conditions due to differing capabilities in the following areas: inlet air pressure, temperature, and flow; fuel flow rate, pressure, and fuel storage capacity; maximum combustion zone temperature; cooling water flow rate and pressure; types of exhaust - atmospheric or altitude; air heater supply pressure; and types of air heaters - vitiated or nonvitiated. All of the facilities have provisions for standard gas (emissions) analysis, and a few of the facilities are equipped with specialized gas analysis equipment, smoke and particle size measurement devices, and a variety of laser systems. This report will present some of the unique features of each of the high temperature/high pressure combustor test facilities at NASA LeRC.

  4. Cars Thermometry in a Supersonic Combustor for CFD Code Validation

    NASA Technical Reports Server (NTRS)

    Cutler, A. D.; Danehy, P. M.; Springer, R. R.; DeLoach, R.; Capriotti, D. P.

    2002-01-01

    An experiment has been conducted to acquire data for the validation of computational fluid dynamics (CFD) codes used in the design of supersonic combustors. The primary measurement technique is coherent anti-Stokes Raman spectroscopy (CARS), although surface pressures and temperatures have also been acquired. Modern- design- of-experiment techniques have been used to maximize the quality of the data set (for the given level of effort) and minimize systematic errors. The combustor consists of a diverging duct with single downstream- angled wall injector. Nominal entrance Mach number is 2 and enthalpy nominally corresponds to Mach 7 flight. Temperature maps are obtained at several planes in the flow for two cases: in one case the combustor is piloted by injecting fuel upstream of the main injector, the second is not. Boundary conditions and uncertainties are adequately characterized. Accurate CFD calculation of the flow will ultimately require accurate modeling of the chemical kinetics and turbulence-chemistry interactions as well as accurate modeling of the turbulent mixing

  5. Compliant Metal Enhanced Convection Cooled Reverse-Flow Annular Combustor

    NASA Technical Reports Server (NTRS)

    Paskin, Marc D.; Acosta, Waldo A.

    1994-01-01

    A joint Army/NASA program was conducted to design, fabricate, and test an advanced, reverse-flow, small gas turbine combustor using a compliant metal enhanced (CME) convection wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CME cooling technique and tben demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F (1922 K) burner outlet temperature (BOT). The CME concept offers significant improvements in wall cooling effectiveness resulting in a large reduction in cooling air requirements. Therefore, more air is available for control of burner outlet temperature pattern in addition to the benefit of improved efficiency, reduced emissions, and smoke levels. Rig test results demonstrated the benefits and viability of the CME concept meeting or exceeding the aerothermal performance and liner wall temperature characteristics of similar lower temperature-rise combustors, achieving 0.15 pattern factor at 3000 F (1922 K) BOT, while utilizing approximately 80 percent less cooling air than conventional, film-cooled combustion systems.

  6. Sectoral combustor for burning low-BTU fuel gas

    DOEpatents

    Vogt, Robert L.

    1980-01-01

    A high-temperature combustor for burning low-BTU coal gas in a gas turbine is disclosed. The combustor includes several separately removable combustion chambers each having an annular sectoral cross section and a double-walled construction permitting separation of stresses due to pressure forces and stresses due to thermal effects. Arrangements are described for air-cooling each combustion chamber using countercurrent convective cooling flow between an outer shell wall and an inner liner wall and using film cooling flow through liner panel grooves and along the inner liner wall surface, and for admitting all coolant flow to the gas path within the inner liner wall. Also described are systems for supplying coal gas, combustion air, and dilution air to the combustion zone, and a liquid fuel nozzle for use during low-load operation. The disclosed combustor is fully air-cooled, requires no transition section to interface with a turbine nozzle, and is operable at firing temperatures of up to 3000.degree. F. or within approximately 300.degree. F. of the adiabatic stoichiometric limit of the coal gas used as fuel.

  7. Transient Numerical Modeling of Catalytic Channels

    NASA Technical Reports Server (NTRS)

    Struk, Peter M.; Dietrich, Daniel L.; Miller, Fletcher J.; T'ien, James S.

    2007-01-01

    first case. Finally, the results show that different initial surface-species distribution leads to different steady-states under certain conditions. These results demonstrate the utility of a lumped two-phase model of a transient catalytic combustor with detailed chemistry.

  8. Reverse-flow combustor for small gas turbines with pressure-atomizing fuel injectors

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Mularz, E. J.; Riddlebaugh, S. M.

    1978-01-01

    A reverse flow combustor suitable for a small gas turbine (2 to 3 kg/s mass flow) was used to evalute the effect of pressure atomizing fuel injectors on combustor performance. In these tests an experimental combustor was designed to operate with 18 simplex pressure atomizing fuel injectors at sea level takeoff conditions. To improve performance at low power conditions, fuel was redistributed so that only every other injector was operational. Combustor performance, emissions, and liner temperature were compared over a range of pressure and inlet air temperatures corresponding to simulated idle, cruise, and takeoff conditions typical of a 16 to 1 pressure ratio turbine engine.

  9. Emissions reduction by varying the swirler airflow split in advanced gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Micklow, Gerald J.; Roychoudhury, Subir; Nguyen, H. L.; Cline, Michael C.

    1992-01-01

    A rich burn/quick mix/lean burn (RQL) combustor concept for reducing pollutant emissions is currently under investigation at the NASA Lewis Research Center. The current study investigates the effect of varying the mass flow rate split between the swirler passages for an equivalance ratio of 2.0 on fuel distribution, temperature distribution, and emissions for the fuel nozzle/rich burn section of an RQL combustor. It is seen that optimizing these parameters can substantially improve combustor performance and reduce combustor emissions. The optimal mass flow rate split for reducing NO(x) emissions based on the numerical study was the same as found by experiment.

  10. High-temperature combustor liner tests in structural component response test facility

    NASA Technical Reports Server (NTRS)

    Moorhead, Paul E.

    1988-01-01

    Jet engine combustor liners were tested in the structural component response facility at NASA Lewis. In this facility combustor liners were thermally cycled to simulate a flight envelope of takeoff, cruise, and return to idle. Temperatures were measured with both thermocouples and an infrared thermal imaging system. A conventional stacked-ring louvered combustor liner developed a crack at 1603 cycles. This test was discontinued after 1728 cycles because of distortion of the liner. A segmented or float wall combustor liner tested at the same heat flux showed no significant change after 1600 cycles. Changes are being made in the facility to allow higher temperatures.

  11. Effect of design features on performance of a double-annular ram-induction combustor

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1975-01-01

    An extensive test program was undertaken to determine the effect of many design features such as the size and number of air scoops, and the type of diffuser airflow distribution to use to optimize performance of a double-annular ram-induction combustor of 94 cm outer diameter. Six combustor configurations were tested. It was found that a snouted double annular combustor built with 256 ram-induction air scoops with a combustor open area giving a total pressure loss of 5.0 percent at a diffuser inlet Mach number of 0.25 gave the best overall performance of the configurations tested.

  12. Enhanced heat transfer rocket combustor technology component hot-fire test results

    NASA Technical Reports Server (NTRS)

    Brown, William S.

    1990-01-01

    The evaluation of a method for enhancing combustor hot-gas wall heat extraction by using hot-fire tests of a rocket engine combustor calorimeter with hot-gas wall ribs is presented. The capability for enhanced heat extraction is required to increase available turbine drive energy for high chamber pressure operation, and therefore higher overall expander cycle engine performance. Determination of the rib effectiveness for incorporation into the design of a high-performance combustor for an advanced expander cycle combustor intended for use in an orbital transfer vehicle or advanced space engine, was the objective of these tests.

  13. Method and apparatus for controlling combustor temperature during transient load changes

    DOEpatents

    Clingerman, Bruce J.; Chalfant, Robert W.

    2002-01-01

    A method and apparatus for controlling the temperature of a combustor in a fuel cell apparatus includes a fast acting air bypass valve connected in parallel with an air inlet to the combustor. A predetermined excess quantity of air is supplied from an air source to a series connected fuel cell and combustor. The predetermined excess quantity of air is provided in a sufficient amount to control the temperature of the combustor during start-up of the fuel processor when the load on the fuel cell is zero and to accommodate any temperature transients during operation of the fuel cell.

  14. Radiant heat transfer from flames in a single tubular turbojet combustor / Leonard Topper

    NASA Technical Reports Server (NTRS)

    Topper, Leonard

    1952-01-01

    An experimental investigation of thermal radiation from the flame of a single tubular turbojet-engine combustor to the combustor liner is presented. The effects of combustor inlet-air pressure, air mass flow, and fuel-air ratio on the radiant intensity and the temperature and emissivity of the flame are reported. The total radiation of the "luminous" flames (containing incandescent soot particles) was much greater (4 to 21 times) than the "nonluminous" molecular radiation. The intensity of radiation from the flame increased rapidly with an increase in combustor inlet-air pressure; it was affected to a lesser degree by variations in fuel-air ratio and air mass flow.

  15. Assessment of pulverized-coal-fired combustor performance. Models for coal-combustor performance: analytical tool verification

    SciTech Connect

    Richter, W.

    1981-02-01

    The development of mathematical models that describe the complex heat transfer processes which occur in industrial combustion chambers is discussed. These combustor models are grouped as either pure heat transfer models or as coupled fluid flow, combustion, and heat transfer models. Two models of the first type and one of the second type are described together with some basic assumptions and sample problems which illustrate their major features and capabilities. (LCL)

  16. Optical Diagnosis of Gas Turbine Combustors Being Conducted

    NASA Technical Reports Server (NTRS)

    Hicks, Yolanda R.; Locke, Randy J.; Anderson, Robert C.; DeGroot, Wilhelmus A.

    2001-01-01

    Researchers at the NASA Glenn Research Center, in collaboration with industry, are reducing gas turbine engine emissions by studying visually the air-fuel interactions and combustion processes in combustors. This is especially critical for next generation engines that, in order to be more fuel-efficient, operate at higher temperatures and pressures than the current fleet engines. Optically based experiments were conducted in support of the Ultra-Efficient Engine Technology program in Glenn's unique, world-class, advanced subsonic combustion rig (ASCR) facility. The ASCR can supply air and jet fuel at the flow rates, temperatures, and pressures that simulate the conditions expected in the combustors of high-performance, civilian aircraft engines. In addition, this facility is large enough to support true sectors ("pie" slices of a full annular combustor). Sectors enable one to test true shapes rather than rectangular approximations of the actual hardware. Therefore, there is no compromise to actual engine geometry. A schematic drawing of the sector test stand is shown. The test hardware is mounted just upstream of the instrumentation section. The test stand can accommodate hardware up to 0.76-m diameter by 1.2-m long; thus sectors or small full annular combustors can be examined in this facility. Planar (two-dimensional) imaging using laser-induced fluorescence and Mie scattering, chemiluminescence, and video imagery were obtained for a variety of engine cycle conditions. The hardware tested was a double annular sector (two adjacent fuel injectors aligned radially) representing approximately 15 of a full annular combustor. An example of the two-dimensional data obtained for this configuration is also shown. The fluorescence data show the location of fuel and hydroxyl radical (OH) along the centerline of the fuel injectors. The chemiluminescence data show C2 within the total observable volume. The top row of this figure shows images obtained at an engine low

  17. Catalytic combustion of alcohols for microburner applications

    NASA Astrophysics Data System (ADS)

    Behrens, Douglas A.; Lee, Ivan C.; Waits, C. Michael

    The combustion of energy dense liquid fuels in a catalytic micro-combustor, whose temperatures can be used in energy conversion devices, is an attractive alternative to cumbersome batteries. To miniaturize the reactor, an evaporation model was developed to calculate the minimum distance required for complete droplet vaporization. By increasing the ambient temperature from 298 to 350 K, the distance required for complete evaporation of a 6.5 μm droplet decreases from 3.5 to 0.15 cm. A platinum mesh acted as a preliminary measurement and demonstrated 75% conversion of ethanol. We then selected a more active rhodium-coated alumina foam with a larger surface area and attained 100% conversion of ethanol and 95% conversion of 1-butanol under fuel lean conditions. Effluent post-combustion gas analysis showed that varying the equivalence ratio results in three possible modes of operation. A regime of high carbon selectivity for CO 2 occurs at low equivalence ratios and corresponds to complete combustion with a typical temperature of 775 K that is ideal for PbTe thermoelectric energy conversion devices. Conversely for equivalence ratios greater than 1, carbon selectivity for CO 2 decreases as hydrogen, olefin and paraffin production increases. By tuning the equivalence ratio, we have shown that a single device can combust completely for thermoelectric applications, operate as a fuel reformer to produce hydrogen gas for fuel cells or perform as a bio-refinery for paraffin and olefin synthesis.

  18. Catalytic distillation process

    DOEpatents

    Smith, Jr., Lawrence A.

    1982-01-01

    A method for conducting chemical reactions and fractionation of the reaction mixture comprising feeding reactants to a distillation column reactor into a feed zone and concurrently contacting the reactants with a fixed bed catalytic packing to concurrently carry out the reaction and fractionate the reaction mixture. For example, a method for preparing methyl tertiary butyl ether in high purity from a mixed feed stream of isobutene and normal butene comprising feeding the mixed feed stream to a distillation column reactor into a feed zone at the lower end of a distillation reaction zone, and methanol into the upper end of said distillation reaction zone, which is packed with a properly supported cationic ion exchange resin, contacting the C.sub.4 feed and methanol with the catalytic distillation packing to react methanol and isobutene, and concurrently fractionating the ether from the column below the catalytic zone and removing normal butene overhead above the catalytic zone.

  19. Catalytic distillation process

    DOEpatents

    Smith, L.A. Jr.

    1982-06-22

    A method is described for conducting chemical reactions and fractionation of the reaction mixture comprising feeding reactants to a distillation column reactor into a feed zone and concurrently contacting the reactants with a fixed bed catalytic packing to concurrently carry out the reaction and fractionate the reaction mixture. For example, a method for preparing methyl tertiary butyl ether in high purity from a mixed feed stream of isobutene and normal butene comprising feeding the mixed feed stream to a distillation column reactor into a feed zone at the lower end of a distillation reaction zone, and methanol into the upper end of said distillation reaction zone, which is packed with a properly supported cationic ion exchange resin, contacting the C[sub 4] feed and methanol with the catalytic distillation packing to react methanol and isobutene, and concurrently fractionating the ether from the column below the catalytic zone and removing normal butene overhead above the catalytic zone.

  20. Evolution of catalytic function

    NASA Technical Reports Server (NTRS)

    Joyce, G. F.

    1993-01-01

    An RNA-based evolution system was constructed in the laboratory and used to develop RNA enzymes with novel catalytic function. By controlling the nature of the catalytic task that the molecules must perform in order to survive, it is possible to direct the evolving population toward the expression of some desired catalytic behavior. More recently, this system has been coupled to an in vitro translation procedure, raising the possibility of evolving protein enzymes in the laboratory to produce novel proteins with desired catalytic properties. The aim of this line of research is to reduce darwinian evolution, the fundamental process of biology, to a laboratory procedure that can be made to operate in the service of organic synthesis.

  1. Catalytic distillation structure

    DOEpatents

    Smith, L.A. Jr.

    1984-04-17

    Catalytic distillation structure is described for use in reaction distillation columns, and provides reaction sites and distillation structure consisting of a catalyst component and a resilient component intimately associated therewith. The resilient component has at least about 70 volume % open space and is present with the catalyst component in an amount such that the catalytic distillation structure consists of at least 10 volume % open space. 10 figs.

  2. Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine

    NASA Technical Reports Server (NTRS)

    Childs, J. Howard; McCafferty, Richard J.

    1948-01-01

    A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical

  3. Inorganic contents reduction in cellulignin catalytic fuel by HGMS

    NASA Astrophysics Data System (ADS)

    Pereira, Maria Luiza G.; Rodrigues, Durval; Conte, Rosa A.; Pinatti, Daltro G.; Shigue, Carlos Y.

    2004-08-01

    Cellulignin is a catalytic fuel obtained by acidic prehydrolysis of biomass (wood, agricultural and forest residues, and the organic matter of municipal solid waste--MOL). In order to be used as a solid fuel in gas turbines with external combustor, the ash content of the cellulignin has to be as low as possible due to the corrosion effect of the Na, K, Mg, Ca and other inorganics on the turbine blades. The main objective of the work was to apply the HGMS to the cellulignin and observe the efficiency of the separation in terms of the reduction of the inorganic constituent concentrations. The separation efficiency was measured by the analysis of the cellulignin samples before and after the magnetic separation. The elements Al, Ca, Fe, K, Mg, Mn, Na, P, S, Si and Zn were determined by ICP-OES. The results showed a concentration reduction for all the elements after the HGMS, allowing this technology to be applied to the cellulignin.

  4. Single annular combustor: Experimental investigations of aerodynamics, dynamics and emissions

    NASA Astrophysics Data System (ADS)

    Mohammad, Bassam Sabry

    The present work investigates the aerodynamics, dynamics and emissions of a Single Cup Combustor Sector. The Combustor resembles a real Gas Turbine Combustor with primary, secondary and dilution zones (also known as fuel rich dome combustor). The research is initiated by studying the effect of the combustor front end geometry on the flow field. Two different exit configurations (one causes a sudden expansion to the swirling flow while the other causes a gradual expansion), installed in a dump combustor, are tested using LDV. The results reveal that the expanding surface reduces the turbulence activities, eliminates the corner recirculation zone and increases the length of the CRZ appreciably. An asymmetry in the flow field is observed due to the asymmetry of the expanding surface. To study the effect of chamber geometry on the flow field, the dome configuration is tested in the combustor sector with the primary dilution jets blocked. The size of the CRZ is reduced significantly (40% reduction in the height). With active primary jets, the CRZ is reconstructed in 3D by conducting several PIV measurements off-center. The confinement appears to significantly influence the shape of the CRZ such that the area ratio is similar for both the confinement and the CRZ (approximately 85%). The primary jets considerably contribute to the heat release process at high power conditions. Also, the primary jets drastically impact the flow field structure. Therefore, the parameters influencing the primary jets are studied using PIV (pressure drop, jets size, off-centering, interaction with convective cooling air, jet blockage and fuel injection). This study is referred to as a jet sensitivity study. The results indicate that the primary jets can be used effectively in controlling the flow field structure. A pressure drop of 4.3% and 7.6% result in similar flows with no noticeable effect on the size of the CRZ and the four jets wake regions. On the other hand, the results show that the

  5. 40 CFR 60.56a - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste combustor operating practices. 60.56a Section 60.56a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) STANDARDS OF PERFORMANCE FOR NEW STATIONARY SOURCES Standards of Performance for Municipal Waste Combustors...

  6. Topping combustor development for second-generation pressurized fluidized bed combined cycles

    SciTech Connect

    Domeracki, W.F.; Dowdy, T.E.; Bachovchin, D.M.

    1994-08-01

    A project team consisting of Foster Wheeler Development Corp. Westinghouse Electric Corp., Gilbert/Commonwealth and the Institute of Gas Technology, are developing a Second Generation Pressurized Fluidized Bed System. Foster Wheeler is developing a carbonizer (a partial gasifier) and a pressurized fluidized bed combustor. Both these units operate a nominal 1600{degrees}F (870{degrees}C) for optimal sulfur capture. Since this temperature is well below the current combustion turbine combustor outlet operating temperature of 2350{degrees}F (1290{degrees}C) to reach commercialization, a topping combustor and hot gas cleanup (HGCU) equipment must be developed. Westinghouse is participating in the development of the high temperature gas cleanup equipment and the topping combustor. This paper concentrates on the design and test of the topping combustor. The topping combustor in this cycle must utilize a low heating value syngas from the carbonizer at approximately 1600{degrees}F (870{degrees}C) and 150 to 210 psi (1.0 to 1.4 MPa). The syngas entering the topping combustor has been previously cleaned of particulates and alkali by the hot gas cleanup (HGCU) system. It also contains significant fuel bound nitrogen present as ammonia and other compounds. The fuel-bound nitrogen is significant because it will selectively convert to NO{sub x} if the fuel is burned under the highly oxidizing conditions of standard combustion turbine combustors.

  7. 40 CFR 60.53a - Standard for municipal waste combustor organics.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... Municipal Waste Combustors for Which Construction is Commenced After December 20, 1989 and on or Before September 20, 1994 § 60.53a Standard for municipal waste combustor organics. (a) (b) On and after the date... 40 Protection of Environment 6 2011-07-01 2011-07-01 false Standard for municipal waste...

  8. 40 CFR 60.52a - Standard for municipal waste combustor metals.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standard for municipal waste combustor metals. 60.52a Section 60.52a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR... September 20, 1994 § 60.52a Standard for municipal waste combustor metals. (a) On and after the date...

  9. 40 CFR Table 3 to Subpart Cb of... - Municipal Waste Combustor Operating Guidelines

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Municipal Waste Combustor Operating... and Compliance Times for Large Municipal Waste Combustors That are Constructed on or Before September 20, 1994 Pt. 60, Subpt. Cb, Table 3 Table 3 to Subpart Cb of Part 60—Municipal Waste...

  10. 40 CFR 60.56a - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 6 2011-07-01 2011-07-01 false Standards for municipal waste combustor operating practices. 60.56a Section 60.56a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) STANDARDS OF PERFORMANCE FOR NEW STATIONARY SOURCES Standards of Performance for Municipal Waste Combustors...

  11. 40 CFR 60.53a - Standard for municipal waste combustor organics.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standard for municipal waste combustor organics. 60.53a Section 60.53a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR... September 20, 1994 § 60.53a Standard for municipal waste combustor organics. (a) (b) On and after the...

  12. Second generation PFBC systems research and development: Phase 2, Topping combustor testing at UTSI

    SciTech Connect

    Johanson, N.R.; Foote, J.P.

    1992-12-01

    This report describes a second generation pressurized fluidized bed combustion (PFBC) power plant. The topping combustor testing is briefly described. The topping combustor burns low BTU gas produced from substoichiometric combustion of coal in a pressurized carbonizer. Char produced is burned in a PFBC.

  13. Effect of Fuel Variables on Carbon Formation in Turbojet-Engine Combustors

    NASA Technical Reports Server (NTRS)

    Jonash, Edmund R; Wear, Jerrold D; Cook, William P

    1958-01-01

    Report presents the results of an investigation of the effects of fuel properties and of a number of fuel additives on combustion-chamber carbon deposition and exhaust-gas smoke formation in a single tubular turbojet-engine combustor. Limited tests were conducted with a number of the fuels in several full-scale turbojet engines to verify single-combustor data.

  14. Isothermal flow in a gas turbine combustor — a benchmark experimental study

    NASA Astrophysics Data System (ADS)

    Koutmos, P.; McGuirk, J. J.

    1989-04-01

    An experimental investigation of the three-dimensional flow field within a water model of a can-type gas turbine combustion chamber is presented. Flow visualisation demonstrated that internal flow patterns simulated closely those expected in real combustors. The combustor comprised a swirl driven primary zone, annulus fed primary and dilution jets and an exit contraction nozzle. LDA measurements of the three mean velocity components and corresponding turbulence intensities were obtained to map out the flow development throughout the combustor. Besides providing information to aid understanding of the complex flow events inside combustors, the data are believed to be of sufficient quantity and quality to act as a benchmark test case for the assessment of the predictive accuracy of computational models for gas-turbine combustors.

  15. A Numerical and an Experimental Study for Optimization of a Small Annular Combustor

    NASA Astrophysics Data System (ADS)

    Iki, Norihiko; Gruber, Andrea; Yoshida, Hiro

    The small annular combustor of a micro gas turbine fueled with methane is investigated experimentally and numerically in order to improve the overall efficiency of the small engine. The CFD analysis of the tiny combustor relies on a low Reynolds number turbulence model coupled to the Eddy Dissipation Concept (EDC) and provides important insight about the turbulent flow pattern, flame shape, position and optimal flame anchoring. For the experimental observation, a model combustor, representing 120 degrees of the original annular combustor, is fabricated, which enables us to visualize internal flow. The burning area in the combustion chamber moves to downstream with increase of air flow rate. At full-load, some fuel remains at the combustion chamber exit. Moreover, temperatures are measured and compared with the numerical simulations. The results shown here will form the basis for future optimization of the micro gas turbine with minimal or no increase in combustor pressure loss.

  16. The large-amplitude combustion oscillation in a single-side expansion scramjet combustor

    NASA Astrophysics Data System (ADS)

    Ouyang, Hao; Liu, Weidong; Sun, Mingbo

    2015-12-01

    The combustion oscillation in scramjet combustor is believed not existing and ignored for a long time. Compared with the flame pulsation, the large-amplitude combustion oscillation in scramjet combustor is indeed unfamiliar and difficult to be observed. In this study, the specifically designed experiments are carried out to investigate this unusual phenomenon in a single-side expansion scramjet combustor. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The obtained results show that the large-amplitude combustion oscillation can exist in scramjet combustor, which is not occasional and can be reproduced. Under the given conditions of this study, moreover, the large-amplitude combustion oscillation is regular and periodic, whose principal frequency is about 126 Hz. The proceeding of the combustion oscillation is accompanied by the transformation of the flame-holding pattern and combustion mode transition between scramjet mode combustion and ramjet mode combustion.

  17. The effects of controlling vortex formation on the performance of a dump combustor

    SciTech Connect

    Mcmanus, K.R.

    1990-01-01

    The use of flow control methods to improve the performance of air breathing combustors, specifically with respect to controlling both volumetric energy release and combustion instability, is investigated. The flow control techniques were chosen for their ability to control vortex formation near the inlet of the combustor. Periodic spanwise forcing of the inlet boundary layer was used to control the spanwise vortex shedding process in the combustor. Delta wing vortex generators and vortex generator jets were used to introduce streamwise vorticity into the flow. These control strategies were applied separately and together to a laboratory-scale dump combustor, and the effects on combustor performance were determined. The effect of spanwise forcing on both nonreacting and reacting flowfields was to modulate the formation of spanwise coherent vortex structures just downstream of the flow separation. In the nonreacting flowfield, the shear layer spreading rate was increased when forcing was applied. In the reacting flow, forcing caused a modulation of the flame structure.

  18. Fuel properties effect on the performance of a small high temperature rise combustor

    NASA Technical Reports Server (NTRS)

    Acosta, Waldo A.; Beckel, Stephen A.

    1989-01-01

    The performance of an advanced small high temperature rise combustor was experimentally determined at NASA-Lewis. The combustor was designed to meet the requirements of advanced high temperature, high pressure ratio turboshaft engines. The combustor featured an advanced fuel injector and an advanced segmented liner design. The full size combustor was evaluated at power conditions ranging from idle to maximum power. The effect of broad fuel properties was studied by evaluating the combustor with three different fuels. The fuels used were JP-5, a blend of Diesel Fuel Marine/Home Heating Oil, and a blend of Suntec C/Home Heating Oil. The fuel properties effect on the performance of the combustion in terms of pattern factor, liner temperatures, and exhaust emissions are documented.

  19. The 3-D CFD modeling of gas turbine combustor-integral bleed flow interaction

    NASA Technical Reports Server (NTRS)

    Chen, D. Y.; Reynolds, R. S.

    1993-01-01

    An advanced 3-D Computational Fluid Dynamics (CFD) model was developed to analyze the flow interaction between a gas turbine combustor and an integral bleed plenum. In this model, the elliptic governing equations of continuity, momentum and the k-e turbulence model were solved on a boundary-fitted, curvilinear, orthogonal grid system. The model was first validated against test data from public literature and then applied to a gas turbine combustor with integral bleed. The model predictions agreed well with data from combustor rig testing. The model predictions also indicated strong flow interaction between the combustor and the integral bleed. Integral bleed flow distribution was found to have a great effect on the pressure distribution around the gas turbine combustor.

  20. Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

    NASA Technical Reports Server (NTRS)

    Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

    2015-01-01

    Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at highpressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NOx emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8%.

  1. Investigation of low NOx staged combustor concept in high-speed civil transport engines

    NASA Technical Reports Server (NTRS)

    Nguyen, Hung Lee; Bittker, David A.; Niedzwiecki, Richard W.

    1989-01-01

    Levels of exhaust emissions due to high temperatures in the main combustor of high-speed civil transport (HSCT) engines during supersonic cruise are predicted. These predictions are based on a new combustor design approach: a rich burn/quick quench/lean burn combustor. A two-stage stirred reactor model is used to calculate the combustion efficiency and exhaust emissions of this novel combustor. A propane-air chemical kinetics model is used to simulate the fuel-rich combustion of jet fuel. Predicted engine exhaust emissions are compared with available experimental test data. The effect of HSCT engine operating conditions on the levels of exhaust emissions is also presented. The work described in this paper is a part of the NASA Lewis Research Center High-Speed Civil Transport Low NO(x) Combustor program.

  2. The pollution reduction technology program for can-annular combustor engines - Description and results

    NASA Technical Reports Server (NTRS)

    Roberts, R.; Fiorentino, A. J.; Diehl, L.

    1976-01-01

    Pollutant reduction and performance characteristics were determined for three successively more advanced combustor concepts. Program Element I consisted of minor modifications to the current production JT8D combustor and fuel system to evaluate means of improved fuel preparation and changes to the basic airflow distribution. Element II addressed versions of the two-staged Vorbix (vortex burning and mixing) combustor and represented a moderate increase in hardware complexity and difficulty of development. The concept selected for Element III employed vaporized fuel as a means of achieving minimum emission levels and represented the greatest difficulty of development and adaptation to the JT8D engine. Test results indicate that the Element I single-stage combustors were capable of dramatic improvement in idle pollutants. The multistage combustors evaluated in Program Elements II and III simultaneously reduced CO, THC and NOx emissions, but were unable to satisfy the current 1979 EPA standards.

  3. Study of research and development requirements of small gas-turbine combustors

    NASA Technical Reports Server (NTRS)

    Demetri, E. P.; Topping, R. F.; Wilson, R. P., Jr.

    1980-01-01

    A survey is presented of the major small-engine manufacturers and governmental users. A consensus was undertaken regarding small-combustor requirements. The results presented are based on an evaluation of the information obtained in the course of the study. The current status of small-combustor technology is reviewed. The principal problems lie in liner cooling, fuel injection, part-power performance, and ignition. Projections of future engine requirements and their effect on the combustor are discussed. The major changes anticipated are significant increases in operating pressure and temperature levels and greater capability of using heavier alternative fuels. All aspects of combustor design are affected, but the principal impact is on liner durability. An R&D plan which addresses the critical combustor needs is described. The plan consists of 15 recommended programs for achieving necessary advances in the areas of liner thermal design, primary-zone performance, fuel injection, dilution, analytical modeling, and alternative-fuel utilization.

  4. Development of a topping combustor for advanced concept pressurized fluidized-bed combustion systems

    SciTech Connect

    Domeracki, W.F.; Dowdy, T.E.; Bachovchin, D.

    1995-11-01

    A project team consisting of Foster Wheeler Development Corporation, Westinghouse Electric Corporation, Gilbert/Commonwealth and the Institute of Gas Technology, are developing a Second Generation Pressurized Fluidized Bed System. Foster Wheeler is developing a carbonizer (a partial gasifier) and a pressurized fluidized bed combustor. Both these units operate at a nominal 1600{degrees}F (870{degrees}C) for optimal sulfur capture. Since this temperature is well below the current combustion turbine combustor outlet operating temperature of 2350{degrees}F (1290{degrees}C), to reach commercialization, a topping combustor and hot gas cleanup (HGCU) equipment must be developed. Westinghouse`s efforts are focused on the development of the high temperature gas cleanup equipment and the topping combustor. This paper concentrates on the design and test of the topping combustor, which must use a low heating value syngas from the carbonizer at approximately 1600{degrees}F and 150 to 210 psi.

  5. Parametric study of flame radiation characteristics of a tubular-can combustor

    NASA Technical Reports Server (NTRS)

    Humenik, F. M.; Claus, R. W.; Neely, G. M.

    1983-01-01

    A series of combustor tests were conducted with a tubular-can combustor to study flame radiation characteristics and effects with parametric variations in combustor operating conditions. Two alternate combustor assemblies using a different fuel nozzle were compared. Spectral and total radiation detectors were positioned at three stations along the length of the combustor can. Data were obtained for a range of pressures from 0.34 to 2.07 MPa (50 to 300 psia), inlet temperatures from 533 to 700K (500 to 800 F), for Jet A (13.9 deg hydrogen) and ERBS (12.9% hydrogen) fuels, and with fuel-air ratios nominally from 0.008 to 0.021. Spectral radiation data, total radiant heat flux data, and liner temperature data are presented to illustrate the flame radiation characteristics and effects in the primary, secondary, and tertiary combustion zones.

  6. Effect of Fuel Injection and Mixing Characteristics on Pulse-Combustor Performance at High-Pressure

    NASA Technical Reports Server (NTRS)

    Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

    2014-01-01

    Recent calculations of pulse-combustors operating at high-pressure conditions produced pressure gains significantly lower than those observed experimentally and computationally at atmospheric conditions. The factors limiting the pressure-gain at high-pressure conditions are identified, and the effects of fuel injection and air mixing characteristics on performance are investigated. New pulse-combustor configurations were developed, and the results show that by suitable changes to the combustor geometry, fuel injection scheme and valve dynamics the performance of the pulse-combustor operating at high-pressure conditions can be increased to levels comparable to those observed at atmospheric conditions. In addition, the new configurations can significantly reduce the levels of NOx emissions. One particular configuration resulted in extremely low levels of NO, producing an emission index much less than one, although at a lower pressure-gain. Calculations at representative cruise conditions demonstrated that pulse-combustors can achieve a high level of performance at such conditions.

  7. Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

    NASA Technical Reports Server (NTRS)

    Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

    2015-01-01

    Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at high pressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NO(x) emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8.

  8. Comment on tangential wave motion in unstable combustors

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.

    1993-01-01

    An attempt is made to gain insight into the injection boundary instability mechanism as captured by a simple computational tool developed for liquid rocket engines. Considerable evidence is generated supporting a hypothesis that destructive tangential-mode instabilities in liquid-fuel combustors are driven by combustion processes occurring at the propellant-wetted injector face. The driving action is found to be directly analogous to a burning solid propellant surface mechanism, responsible for solid rocket motor instabilities. Since the proposed instability mechanism has not been fundamentally confirmed, experimental research is strongly recommended.

  9. Refractory experience in circulating fluidized bed combustors, Task 7

    SciTech Connect

    Vincent, R.Q.

    1989-11-01

    This report describes the results of an investigation into the status of the design and selection of refractory materials for coal-fueled circulating fluidized-bed combustors. The survey concentrated on operating units in the United States manufactured by six different boiler vendors: Babcock and Wilcox, Combustion Engineering, Foster Wheeler, Keeler Dorr-Oliver, Pyropower, and Riley Stoker. Information was obtained from the boiler vendors, refractory suppliers and installers, and the owners/operators of over forty units. This work is in support of DOE's Clean Coal Technology program, which includes circulating fluidized-bed technology as one of the selected concepts being evaluated.

  10. Nondestructive Evaluation of Ceramic Matrix Composite Combustor Components

    NASA Technical Reports Server (NTRS)

    Sun, Jiangang G.; Verrilli, Michael J.; Stephan, Robert R.; Barnett, Terry R.; Ojard, Greg C.

    2003-01-01

    Combustor liners fabricated from a SiC/SiC composite (silicon carbide fibers in a silicon carbide matrix) were nondestructively interrogated before and after combustion rig testing by x-ray, ultrasonic, and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured liners, and from coupons machined from the components after combustion exposure. Thermography indications correlated with reduced material properties obtained after rig testing. The thermography indications in the SiC/SiC liners were delaminations and damaged fiber tows, as determined through microstructural examinations.

  11. Particulate exhaust emissions from an experimental combustor. [gas turbine engine

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Ingebo, R. D.

    1975-01-01

    The concentration of dry particulates (carbon) in the exhaust of an experimental gas turbine combustor was measured at simulated takeoff operating conditions and correlated with the standard smoke-number measurement. Carbon was determined quantitatively from a sample collected on a fiberglass filter by converting the carbon in the smoke sample to carbon dioxide and then measuring the volume of carbon dioxide formed by gas chromatography. At a smoke of 25 (threshold of visibility of the smoke plume for large turbojets) the carbon concentration was 2.8 mg carbon/cu m exhaust gas, which is equivalent to an emission index of 0.17 g carbon/kg fuel.

  12. Nondestructive Evaluation of Ceramic Matrix Composite Combustor Components

    NASA Technical Reports Server (NTRS)

    Sun, J. G.; Verrilli, M. J.; Stephan, R.; Barnett, T. R.; Ojard, G.

    2003-01-01

    Combustor liners fabricated from a SiC/SiC composite were nondestructively interrogated before and after combustion rig testing by X-ray, ultrasonic and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured liners, and from coupons machined from the components after combustion exposure. Thermography indications were found to correlate with reduced material properties obtained after rig testing. The thermography indications in the SiC/SiC liners were delaminations and damaged fiber tows, as determined through microstructural examinations. [copyright] 2003 American Institute of Physics

  13. Laser-Based Diagnostic Measurements of Low Emissions Combustor Concepts

    NASA Technical Reports Server (NTRS)

    Hicks, Yolanda R.

    2011-01-01

    This presentation provides a summary of primarily laser-based measurement techniques we use at NASA Glenn Research Center to characterize fuel injection, fuel/air mixing, and combustion. The report highlights using Planar Laser-Induced Fluorescence, Particle Image Velocimetry, and Phase Doppler Interferometry to obtain fuel injector patternation, fuel and air velocities, and fuel drop sizes and turbulence intensities during combustion. We also present a brief comparison between combustors burning standard JP-8 Jet fuel and an alternative fuels. For this comparison, we used flame chemiluminescence and high speed imaging.

  14. Calculation of CO concentration for liquid fueled gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Patil, P. B.; Sichel, M.; Nicholls, J. A.

    1978-01-01

    The extensive computation time required for the numerical integration of the differential equations obtained in studies of the concentrations of pollutants emitted by gas turbine combustors, can be reduced significantly by assuming the heat releasing hydrocarbon reactions to be in local equilibrium. In determining the CO and NO concentrations produced in spray combustion, it is, therefore, tempting to use the same local equilibrium assumption in order to simplify the kinetic calculations. An investigation of the validity of the local equilibrium assumption is conducted by using a simple analytical model, and then by actually carrying out the kinetic and local equilibrium calculations for typical case.

  15. Computations of Emissions Using a 3-D Combustor Program

    NASA Technical Reports Server (NTRS)

    Srivatsa, S. K.

    1983-01-01

    A general 3-D combustor performance program developed by Garrett was extended to predict soot and NOx emissions. The soot formation and oxidation rates were computed by quasi-global models, taking into account the influence of turbulence. Radiation heat transfer was computed by the six-flux radiation mode. The radiation properties include the influence of CO2 and H2O in addition to soot. NOx emissions were computed from a global four-step hydrocarbon oxidation scheme and a set of rate-controlled reactions involving radicals and nitrogen oxides.

  16. Device for improved air and fuel distribution to a combustor

    DOEpatents

    Laster, Walter R.; Schilp, Reinhard

    2016-05-31

    A flow conditioning device (30, 50, 70, 100, 150) for a can annular gas turbine engine, including a plurality of flow elements (32, 34, 52, 54, 72, 74, 102) disposed in a compressed air flow path (42, 60, 80, 114, 122) leading to a combustor (12), configured such that relative adjustment of at least one flow directing element (32, 52, 72, 110) with respect to an adjacent flow directing element (34, 54, 74, 112, 120) during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path (42, 60, 80, 114, 122).

  17. Performance of a swirl-can combustor at idle conditions

    NASA Technical Reports Server (NTRS)

    Niedzwiecki, R. W.; Trout, A. M.; Mularz, E. J.

    1972-01-01

    Test results of a full annulus swirl-can combustor operated at simulated engine idle conditions indicated that significant improvements in combustion efficiency and accompanying reductions in pollutants could be realized with radial scheduling of fuel. Test conditions were an inlet air temperature of 478 K, a pressure of 4 atmospheres, and a reference velocity of 26 meters per second. At a fuel-air ratio of 0.008, radial scheduling of fuel increased combustion efficiency from less than 50 percent with no scheduling to nearly 100 percent and produced emission index values of unburned hydrocarbons and carbon monoxide of 10 and 40, respectively.

  18. Catalytic coherence transformations

    NASA Astrophysics Data System (ADS)

    Bu, Kaifeng; Singh, Uttam; Wu, Junde

    2016-04-01

    Catalytic coherence transformations allow the otherwise impossible state transformations using only incoherent operations with the aid of an auxiliary system with finite coherence that is not being consumed in any way. Here we find the necessary and sufficient conditions for the deterministic and stochastic catalytic coherence transformations between a pair of pure quantum states. In particular, we show that the simultaneous decrease of a family of Rényi entropies of the diagonal parts of the states under consideration is a necessary and sufficient condition for the deterministic catalytic coherence transformations. Similarly, for stochastic catalytic coherence transformations we find the necessary and sufficient conditions for achieving a higher optimal probability of conversion. We thus completely characterize the coherence transformations among pure quantum states under incoherent operations. We give numerous examples to elaborate our results. We also explore the possibility of the same system acting as a catalyst for itself and find that indeed self-catalysis is possible. Further, for the cases where no catalytic coherence transformation is possible we provide entanglement-assisted coherence transformations and find the necessary and sufficient conditions for such transformations.

  19. Catalytic nanoporous membranes

    DOEpatents

    Pellin, Michael J; Hryn, John N; Elam, Jeffrey W

    2013-08-27

    A nanoporous catalytic membrane which displays several unique features Including pores which can go through the entire thickness of the membrane. The membrane has a higher catalytic and product selectivity than conventional catalysts. Anodic aluminum oxide (AAO) membranes serve as the catalyst substrate. This substrate is then subjected to Atomic Layer Deposition (ALD), which allows the controlled narrowing of the pores from 40 nm to 10 nm in the substrate by deposition of a preparatory material. Subsequent deposition of a catalytic layer on the inner surfaces of the pores reduces pore sizes to less than 10 nm and allows for a higher degree of reaction selectivity. The small pore sizes allow control over which molecules enter the pores, and the flow-through feature can allow for partial oxidation of reactant species as opposed to complete oxidation. A nanoporous separation membrane, produced by ALD is also provided for use in gaseous and liquid separations. The membrane has a high flow rate of material with 100% selectivity. Also provided is a method for producing a catalytic membrane having flow-through pores and discreet catalytic clusters adhering to the inside surfaces of the pores.

  20. Active Combustion Control for Aircraft Gas-Turbine Engines-Experimental Results for an Advanced, Low-Emissions Combustor Prototype

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Kopasakis, George; Saus, Joseph R.; Chang, Clarence T.; Wey, Changlie

    2012-01-01

    Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. Mitigation of instabilities is an enabling technology for these low-emissions combustors. NASA Glenn Research Center s prior activity has demonstrated active control to suppress a high-frequency combustion instability in a combustor rig designed to emulate an actual aircraft engine instability experience with a conventional, rich-front-end combustor. The current effort is developing further understanding of the problem specifically as applied to future lean-burning, very low-emissions combustors. A prototype advanced, low-emissions aircraft engine combustor with a combustion instability has been identified and previous work has characterized the dynamic behavior of that combustor prototype. The combustor exhibits thermoacoustic instabilities that are related to increasing fuel flow and that potentially prevent full-power operation. A simplified, non-linear oscillator model and a more physics-based sectored 1-D dynamic model have been developed to capture the combustor prototype s instability behavior. Utilizing these models, the NASA Adaptive Sliding Phasor Average Control (ASPAC) instability control method has been updated for the low-emissions combustor prototype. Active combustion instability suppression using the ASPAC control method has been demonstrated experimentally with this combustor prototype in a NASA combustion test cell operating at engine pressures, temperatures, and flows. A high-frequency fuel valve was utilized to perturb the combustor fuel flow. Successful instability suppression was shown using a dynamic pressure sensor in the combustor for controller feedback. Instability control was also shown with a pressure feedback sensor in the lower temperature region upstream of the combustor. It was also demonstrated that the controller can prevent the instability from occurring while combustor operation was transitioning from a stable, low-power condition to

  1. Catalytic hydrotreating process

    DOEpatents

    Karr, Jr., Clarence; McCaskill, Kenneth B.

    1978-01-01

    Carbonaceous liquids boiling above about 300.degree. C such as tars, petroleum residuals, shale oils and coal-derived liquids are catalytically hydrotreated by introducing the carbonaceous liquid into a reaction zone at a temperature in the range of 300.degree. to 450.degree. C and a pressure in the range of 300 to 4000 psig for effecting contact between the carbonaceous liquid and a catalytic transition metal sulfide in the reaction zone as a layer on a hydrogen permeable transition metal substrate and then introducing hydrogen into the reaction zone by diffusing the hydrogen through the substrate to effect the hydrogenation of the carbonaceous liquid in the presence of the catalytic sulfide layer.

  2. Catalytic membranes beckon

    SciTech Connect

    Caruana, C.M.

    1994-11-01

    Chemical engineers here and abroad are finding that the marriage of catalysts and membranes holds promise for faster and more specific reactions, although commercialization of this technology is several years away. Catalytic membrane reactors (CMRs) combine a heterogeneous catalyst and a permselective membrane. Reactions performed by CMRs provide higher yields--sometimes as much as 50% higher--because of better reaction selectivity--as opposed to separation selectivity. CMRs also can work at very high temperatures, using ceramic materials that would not be possible with organic membranes. Although the use of CMRs is not widespread presently, the development of new membranes--particularly porous ceramic and zeolite membranes--will increase the potential to improve yields of many catalytic processes. The paper discusses ongoing studies, metal and advanced materials for membranes, the need for continued research, hydrogen recovery from coal-derived gases, catalytic oxidation of sulfides, CMRs for water purification, and oxidative coupling of methane.

  3. Flow and Emissions Characteristics of Multi-Swirler Combustor

    NASA Astrophysics Data System (ADS)

    Gutmark, Ephraim; Li, Guoqiang

    2003-11-01

    Modern industrial gas-turbine spray combustors feature multiple swirlers and distributed fuel injection for rapid mixing and stabilization. The flow field of this combustor, the related combustion characteristics and their control are discussed. The velocity flow field downstream of a Triple Annular Research Swirler (TARS) was characterized. Multiple combinations of swirlers were tested in cold flow under atmospheric conditions with and without confining combustion chamber. The experiments showed that a central recirculation zone (CTRZ), an annular jet with internal and external shear layers dominated the flow field downstream of TARS. Compared to unconfined case, flow with confined tube showed an enlarged CTRZ region and a recirculation region in the expansion corner with reduced concentration of turbulence intensity in the jet region. TARS also produced low emissions of NOx and CO. Measurements were performed to study the effects of several factors, including swirler combinations, exhaust nozzle size, air assist for fuel atomization and mixing length on NOx and CO emissions and combustion instability. The data showed that emissions and stability depend on the combination of several of these factors.

  4. Performance of a second-generation PFB pilot plant combustor

    SciTech Connect

    Conn, R.; Van Hook, J.; Robertson, A.; Bonk, D.

    1995-07-01

    Second-generation pressurized fluidized bed combustion (PFBC) plants promise higher efficiency with lower costs of electricity and lower stack emissions. With a conventional reheat steam cycle and a 3% sulfur Pittsburgh No. 8 coal, a 45% efficiency (HHV of coal basis) and a cost of electricity 20% lower than that of a pulverized-coal-fired plant with stack gas scrubbing are being projected. This advanced plant concept incorporates three major steps: carbonization, circulating fluidized bed combustion and topping combustion. Foster Wheeler Development Corporation has constructed and operated a second-generation PFB pilot plant at the Foster Wheeler research facility (the John Blizard Research Center) in Livingston, New Jersey. Results of the pilot plant combustor portion of the test program supporting the development of this new type of plant are presented. The fuels evaluated in this test program included several char-sorbent residues produced in a pressurized carbonizer pilot plant and their parent coals. The data confirmed the viability of the PFB combustor concept in terms of both combustion and emissions performance.

  5. Lean premixed flames for low NO{sub x} combustors

    SciTech Connect

    Sojka, P.; Tseng, L.; Bryyjak, J.

    1995-12-31

    The overall objectives of the research at Purdue are to: obtain a reduced mechanism description of high pressure NO formation chemistry using experiments and calculations for laminar lean premixed methane air flames, develop a statistical model of turbulence NO chemistry interactions using a Bunsen type jet flame, and utilize the high pressure chemistry and turbulence models in a commercial design code, then evaluate its predictions using data from an analog gas turbine combustor. Work to date has resulted in the following achievements: spatially resolved measurements of NO in high-pressure high-temperature flat flames, plus evaluation of the influence of flame radiation on the measured temperature profile; measurements of temperature and velocity PDFs for a turbulent methane/air flame were obtained for the first time, under operating conditions which allow their study in the distributed regimes, and the increase in EINO{sub x} with equivalence ratio predicted using a chemical kinetics model; and simulation of non-reacting combustor flow fields from ambient to elevated pressure and temperature conditions and comparison of those results with experimental velocity profiles.

  6. Design of thermal protection system for 8 foot HTST combustor

    NASA Technical Reports Server (NTRS)

    Moskowitz, S.

    1973-01-01

    The combustor in the 8-foot high temperature structures tunnel at the NASA-Langley Research Center has encountered cracking over a period of 50-250 tunnel tests within a limited range of the required operating envelope. A program was conducted which analyzed the failed combustor liner hardware and determined that the mechanism of failure was vibratory fatigue. A vibration damper system using wave springs located axially between the liner T-bar and the liner support was designed as an intermediate solution to extend the life of the current two-pass regenerative air-cooled liner. The effects of liner wall thickness, cooling air passage height, stiffener ring geometry, reflective coatings, and liner material selection were investigated for these designs. Preliminary layout design arrangements including the external water-cooling system requirements, weight estimates, installation requirements and preliminary estimates of manufacturing costs were prepared for the most promissing configurations. A state-of-the-art review of thermal barrier coatings and an evaluation of reflective coatings for the gasside surface of air-cooled liners are included.

  7. Fabrication of strain-isolated ceramic coated combustor components

    NASA Technical Reports Server (NTRS)

    Rutter, S.

    1985-01-01

    The use of strain-isolated ceramic coated material to produce an AGT1500 combustor scroll-shaped transition duct which requires no air for film cooling is investigated. The scroll receives the exhaust of the can-style combustor liner and turns it into the annular inlet of the high pressure gas producer turbine nozzle. Strain-isolation of plasma sprayed thermal barrier coating is achieved by placing a compliant pad between the structural base metal and the ceramic coating. The compliant pad is brazed to the metal structure. In order to achieve a good braze bond, the strain-isolating compliant pad and base metal must be closely matched in shape and tightly fixtured for joining. The complex geometry of the AGT1500 scroll makes it impractical to attack pads to the supporting structure in its finished shape. Instead the pads are brazed to flat stock and post-formed into scroll sections. While test samples were successfully post-formed, plasma sprayed, and subjected to cyclic heating, the forming of full scale parts by normal methods resulted in tearing of the Hastelloy-X base metal because of embrittlement by the braze material. Several solutions were explored which finally resulted in the successful forming of full scale scroll parts.

  8. Evaluation of fuel preparation systems for lean premixing- prevaporizing combustors

    SciTech Connect

    Dodds, W.J.; Ekstedt, .E.E.

    1986-04-01

    A series of tests was conducted to provide data for the design of premixing-prevaporizing fuel-air mixture preparation systems for aircraft gas turbine engine combustors. Fifteen configurations of four different fuel-air mixture preparation system design concepts were evaluated to determine fuel-air mixture uniformity at the system exit over a range of conditions representative of cruise operation for a modern commercial turbofan engine. Operating conditions, including pressure, temperature, fuel-air ratio and velocity had no clear effect on mixture uniformity in systems which used low-pressure fuel injectors. However, performance of systems using pressure atomizing fuel nozzles and large-scale mixing devices was shown to be sensitive to operating conditions. Variations in system design variables were also evaluated and correlated. Mixture uniformity improved with increased system length, pressure drop, and number of fuel injection points per unit area. A premixing system compatible with the combustor envelope of a typical combustion system and capable of providing mixture nonuniformity (standard deviation/mean) below 15% over a typical range of cruise operating conditions was demonstrated.

  9. Investigation of the transient fuel preburner manifold and combustor

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See; Chen, Yen-Sen; Farmer, Richard C.

    1989-01-01

    A computational fluid dynamics (CFD) model with finite rate reactions, FDNS, was developed to study the start transient of the Space Shuttle Main Engine (SSME) fuel preburner (FPB). FDNS is a time accurate, pressure based CFD code. An upwind scheme was employed for spatial discretization. The upwind scheme was based on second and fourth order central differencing with adaptive artificial dissipation. A state of the art two-equation k-epsilon (T) turbulence model was employed for the turbulence calculation. A Pade' Rational Solution (PARASOL) chemistry algorithm was coupled with the point implicit procedure. FDNS was benchmarked with three well documented experiments: a confined swirling coaxial jet, a non-reactive ramjet dump combustor, and a reactive ramjet dump combustor. Excellent comparisons were obtained for the benchmark cases. The code was then used to study the start transient of an axisymmetric SSME fuel preburner. Predicted transient operation of the preburner agrees well with experiment. Furthermore, it was also found that an appreciable amount of unburned oxygen entered the turbine stages.

  10. Ceramic composite liner material for gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Ercegovic, D. B.; Walker, C. L.; Norgren, C. T.

    1984-01-01

    The application of ceramics to gas turbine combustor liners to reduce liner metal temperature was studied in an experiment in which yttria-stabilized zirconia plasma was sprayed on compliant metal substrates exposed to near stoichiometric combustion. The strain isolation pad materials chosen were Hoskins Alloy 875 and BRUNSLLOY 534 Fiber Metal of 0.25 and 0.38 cm thicknesses and 35 and 45 percent density levels. Combustor screening tests of all specimens showed no evidence of deterioration or failure. Specimens exposed to flame temperatures in excess of 2100 K were convectively or convective-transpiration cooled and were evaluated in a 10 cm sq flame tube at inlet air temperature of 533 K and pressure of 0.5 MPa. The results suggest the superiority of a system composed of the Hoskins Alloy 875 compliant pad with 0.25 cm thickness and 35 percent density coupled with a NiCrAlY bond coat and a 8 percent Y2O3-ZrO2 ceramic top coat of 0.19 cm thickness.

  11. Three Dimensional CFD Analysis of the GTX Combustor

    NASA Technical Reports Server (NTRS)

    Steffen, C. J., Jr.; Bond, R. B.; Edwards, J. R.

    2002-01-01

    The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. ScramJet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel massflow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.

  12. Vortex generation and mixing in three-dimensional supersonic combustors

    NASA Technical Reports Server (NTRS)

    Riggins, D. W.; Vitt, P. H.

    1993-01-01

    The generation and evolution of the flow vorticity established by instream injector ramps in a high Mach number/high enthalpy scramjet combustor flow-field are described in detail for a number of computational cases. Classical fluid dynamic circulation is presented for these cases in order to clarify the spatial distribution and convection of the vorticity. The ability of the simulations to accurately represent Stokes Law of circulation is discussed and shown. In addition, the conservation of swirl (effectively the moment-of-momentum theorem) is presented for these flows. The impact of both turbulent diffusion and the vortex/ramp non-uniformity on the downstream mixing rate is clearly illustrated. A correlation over the length of the combustor between fuel-air mixing and a parameter called the vortex stirring length is demonstrated. Finally, computational results for a representative ramp injector are compared with experimental data. Influence of the stream vorticity on the effective turbulent Prandtl number used in the simulation is discussed.

  13. Auto-ignition system for premixed gas turbine combustors

    SciTech Connect

    Mumford, S.E.

    1993-08-24

    In a gas turbine power plant having at least one combustor, the combustor is described comprising a pilot section and a main burn section, the pilot section, having a recirculation zone and comprising inlet air vents for providing inlet air into the recirculation zone, the main burn section being downstream and adjacent to the pilot section, a method for establishing a diffusion flame in the pilot section, comprising the steps of: injecting a flow of natural gas fuel into the pilot section; providing inlet air into the recirculation zone through the inlet air vents; operating the pilot section in a premix mode where the natural gas fuel mixes with the inlet air without combustion in the pilot section; and injecting a flow of liquid fuel into the recirculation zone of the pilot section just prior to the end of operation of the pilot section in the premix mode, the liquid fuel having an auto-ignition temperature less than the temperature of the inlet air, whereby the liquid fuel is auto-ignited in the recirculation zone and the natural gas fuel bursts into flame to establish the diffusion flame in the pilot section.

  14. Detecting deterministic nature of pressure measurements from a turbulent combustor

    NASA Astrophysics Data System (ADS)

    Tony, J.; Gopalakrishnan, E. A.; Sreelekha, E.; Sujith, R. I.

    2015-12-01

    Identifying nonlinear structures in a time series, acquired from real-world systems, is essential to characterize the dynamics of the system under study. A single time series alone might be available in most experimental situations. In addition to this, conventional techniques such as power spectral analysis might not be sufficient to characterize a time series if it is acquired from a complex system such as a thermoacoustic system. In this study, we analyze the unsteady pressure signal acquired from a turbulent combustor with bluff-body and swirler as flame holding devices. The fractal features in the unsteady pressure signal are identified using the singularity spectrum. Further, we employ surrogate methods, with translational error and permutation entropy as discriminating statistics, to test for determinism visible in the observed time series. In addition to this, permutation spectrum test could prove to be a robust technique to characterize the dynamical nature of the pressure time series acquired from experiments. Further, measures such as correlation dimension and correlation entropy are adopted to qualitatively detect noise contamination in the pressure measurements acquired during the state of combustion noise. These ensemble of measures is necessary to identify the features of a time series acquired from a system as complex as a turbulent combustor. Using these measures, we show that the pressure fluctuations during combustion noise has the features of a high-dimensional chaotic data contaminated with white and colored noise.

  15. Detecting deterministic nature of pressure measurements from a turbulent combustor.

    PubMed

    Tony, J; Gopalakrishnan, E A; Sreelekha, E; Sujith, R I

    2015-12-01

    Identifying nonlinear structures in a time series, acquired from real-world systems, is essential to characterize the dynamics of the system under study. A single time series alone might be available in most experimental situations. In addition to this, conventional techniques such as power spectral analysis might not be sufficient to characterize a time series if it is acquired from a complex system such as a thermoacoustic system. In this study, we analyze the unsteady pressure signal acquired from a turbulent combustor with bluff-body and swirler as flame holding devices. The fractal features in the unsteady pressure signal are identified using the singularity spectrum. Further, we employ surrogate methods, with translational error and permutation entropy as discriminating statistics, to test for determinism visible in the observed time series. In addition to this, permutation spectrum test could prove to be a robust technique to characterize the dynamical nature of the pressure time series acquired from experiments. Further, measures such as correlation dimension and correlation entropy are adopted to qualitatively detect noise contamination in the pressure measurements acquired during the state of combustion noise. These ensemble of measures is necessary to identify the features of a time series acquired from a system as complex as a turbulent combustor. Using these measures, we show that the pressure fluctuations during combustion noise has the features of a high-dimensional chaotic data contaminated with white and colored noise. PMID:26764769

  16. NYU-DOE Pressurized Fluidized Bed Combustor Facility

    SciTech Connect

    Zakkay, V.; Kolar, A.; Sellakumar, K.; Srinivasaragavan, S.; Miller, G.; Panunzio, S.; Joseph, A.; Sundaresan, C.

    1983-01-01

    New York University (NYU), under a Department of Energy (DOE) Contract, has designed and constructed a sub-pilot scale Pressurized Fluidized-Bed Combustor (PFBC) Facility at the Antonio Ferri Laboratories, Westbury, Long Island. The basic feature of this Experimental Research Facility is a well-instrumented, 30-inch diameter coal combustor capable of operating up to 10 atm and provided with a liberal number of ports, making it a versatile unit for study of fundamental in-bed phenomena. Additionally, the overall design features make it a flexible facility for solving a variety of industrial research problems. The main objectives of the facility are two-fold: (1) to perform research in important areas of Pressurized Fluidized-Bed Combustion like low-grade fuel combustion under pressure; and (2) to provide the PFBC community with a experimental research tool for basic and applied research in order to accelerate the commercialization of this technology. New York University will initially test the facility of burning low-grade fuels under pressure. During the test program, emphasis will be placed on burning North Dakota lignite under pressures up to 7 atm. The performance of lignite with regard to its feeding, combustion efficiency, sulfur adsorption and sorbent requirements will be investigated. This report describes the various systems of the PFBC facility and operating procedures, and presents an outline of the test program planned for the facility. Other details are provided in the Equipment and Maintenance Manual, Test Program and Data Acquisition Manual, and Training Manual.

  17. Modeling the performance of the whole tree energy combustor

    SciTech Connect

    Bryden, K.M.; Ragland, K.W.; Ostlie, L.D.

    1994-12-31

    A computational model has been developed for the deep, fixed bed whole-tree combustor/gasifier which includes drying, pyrolysis, heat and mass transfer, char and volatile reactions. The lowest portion of the fuel bed is an oxidizing region and the remainder of the fuel bed acts as a gasification and drying region. Heat release rates, temperature profiles, CO, CO{sub 2}, H{sub 2}O, hydrocarbon, and O{sub 2} profiles, and solid and gas velocity profiles are calculated as a function of the fuel properties and air inputs. Model results are compared to pilot scale tests of the whole-tree combustor/gasifier. The effect of operating parameters such as underfire air flow (2-12 ft/s), fuel moisture content (15-45%), and fuel bed height (8-16 ft) on the fuel feed rate, heat release rate, temperature and velocity profiles, and overfire air requirements is presented and discussed. Heat release rates range from 1x10{sup 6} to 4.5x10{sup 6} Btu/hr-ft{sup 2} depending on the conditions.

  18. Catalytic Combustion for Ultra-Low NOx Hydrogen Turbines

    SciTech Connect

    Etemad, Shahrokh; Baird, Benjamin; Alavandi, Sandeep

    2011-06-30

    Precision Combustion, Inc., (PCI) in close collaboration with Solar Turbines, Incorporated, has developed and demonstrated a combustion system for hydrogen fueled turbines that reduces NOx to low single digit level while maintaining or improving current levels of efficiency and eliminating emissions of carbon dioxide. Full scale Rich Catalytic Hydrogen (RCH1) injector was developed and successfully tested at Solar Turbines, Incorporated high pressure test facility demonstrating low single digit NOx emissions for hydrogen fuel in the range of 2200F-2750F. This development work was based on initial subscale development for faster turnaround and reduced cost. Subscale testing provided promising results for 42% and 52% H2 with NOx emissions of less than 2 ppm with improved flame stability. In addition, catalytic reactor element testing for substrate oxidation, thermal cyclic injector testing to simulate start-stop operation in a gas turbine environment, and steady state 15 atm. operation testing were performed successfully. The testing demonstrated stable and robust catalytic element component life for gas turbine conditions. The benefit of the catalytic hydrogen combustor technology includes capability of delivering near-zero NOx without costly post-combustion controls and without requirement for added sulfur control. In addition, reduced acoustics increase gas turbine component life. These advantages advances Department of Energy (DOE’s) objectives for achievement of low single digit NOx emissions, improvement in efficiency vs. postcombustion controls, fuel flexibility, a significant net reduction in Integrated Gasification Combined Cycle (IGCC) system net capital and operating costs, and a route to commercialization across the power generation field from micro turbines to industrial and utility turbines.

  19. Advances in measurements and simulation of gas-particle flows and coal combustion in burners/combustors

    NASA Astrophysics Data System (ADS)

    Zhou, L. X.

    2009-02-01

    Innovative coal combustors were developed, and measurement and simulation of gas-particle flows and coal combustion in such combustors were done in the Department of Engineering Mechanics, Tsinghua University. LDV/PDPA measurements are made to understand the behavior of turbulent gas-particle flows in coal combustors. Coal combustion test was done for the non-slagging cyclone coal combustor. The full two-fluid model developed by the present author was used to simulate turbulent gas-particle flows, coal combustion and NOx formation. It is found by measurements and simulation that the optimum design can give large-size recirculation zones for improving the combustion performance for all the combustors. The combustion test shows that the nonslagging coal combustor can burn 3-5mm coal particles with good combustion efficiency and low NO emission. Simulation in comparison with experiments indicates that the swirl number can significantly affect the NO formation in the swirl coal combustor.

  20. Hollow fiber catalytic membranes

    SciTech Connect

    Ma, Yi Hua; Moser, W.; Shelekhin, A.; Pien, Shyhing

    1993-09-01

    The objective of the present research is to investigate the possibility of the enhancement of the H{sub 2}S thermal decomposition in the IGCC system by employing the hollow fiber catalytic membrane reactor. To accomplish the objective, the following major components in the analysis of the high temperature membrane reactor must be investigated: high-temperature stability of the porous glass membrane; catalytic properties of MoS{sub 2} and of the porous glass membrane; catalytic decomposition of H{sub 2}S in a packed bed reactor; catalytic decomposition of 100%, 8.6%, and 1.1% H{sub 2}S gas mixtures in the membrane reactor. The study has been shown that the conversion of the H{sub 2}S can be increased in the packed bed membrane reactor compared to the equilibrium conversion on the shell side. The development of a mathematical model for the proposed process is in progress. The model will enable optimization of the H{sub 2}S decomposition. These conditions include selectivity factors and pressure drop across the membrane.

  1. Monolithic catalytic igniters

    NASA Technical Reports Server (NTRS)

    La Ferla, R.; Tuffias, R. H.; Jang, Q.

    1993-01-01

    Catalytic igniters offer the potential for excellent reliability and simplicity for use with the diergolic bipropellant oxygen/hydrogen as well as with the monopropellant hydrazine. State-of-the-art catalyst beds - noble metal/granular pellet carriers - currently used in hydrazine engines are limited by carrier stability, which limits the hot-fire temperature, and by poor thermal response due to the large thermal mass. Moreover, questions remain with regard to longevity and reliability of these catalysts. In this work, Ultramet investigated the feasibility of fabricating monolithic catalyst beds that overcome the limitations of current catalytic igniters via a combination of chemical vapor deposition (CVD) iridium coatings and chemical vapor infiltration (CVI) refractory ceramic foams. It was found that under all flow conditions and O2:H2 mass ratios tested, a high surface area monolithic bed outperformed a Shell 405 bed. Additionally, it was found that monolithic catalytic igniters, specifically porous ceramic foams fabricated by CVD/CVI processing, can be fabricated whose catalytic performance is better than Shell 405 and with significantly lower flow restriction, from materials that can operate at 2000 C or higher.

  2. Catalytic coal liquefaction process

    DOEpatents

    Garg, D.; Sunder, S.

    1986-12-02

    An improved process for catalytic solvent refining or hydroliquefaction of non-anthracitic coal at elevated temperatures under hydrogen pressure in a solvent comprises using as catalyst a mixture of a 1,2- or 1,4-quinone and an alkaline compound, selected from ammonium, alkali metal, and alkaline earth metal oxides, hydroxides or salts of weak acids. 1 fig.

  3. Catalytic coal liquefaction process

    DOEpatents

    Garg, Diwakar; Sunder, Swaminathan

    1986-01-01

    An improved process for catalytic solvent refining or hydroliquefaction of non-anthracitic coal at elevated temperatures under hydrogen pressure in a solvent comprises using as catalyst a mixture of a 1,2- or 1,4-quinone and an alkaline compound, selected from ammonium, alkali metal, and alkaline earth metal oxides, hydroxides or salts of weak acids.

  4. Combustion Dynamics in Multi-Nozzle Combustors Operating on High-Hydrogen Fuels

    SciTech Connect

    Santavicca, Dom; Lieuwen, Tim

    2013-09-30

    Actual gas turbine combustors for power generation applications employ multi-nozzle combustor configurations. Researchers at Penn State and Georgia Tech have extended previous work on the flame response in single-nozzle combustors to the more realistic case of multi-nozzle combustors. Research at Georgia Tech has shown that asymmetry of both the flow field and the acoustic forcing can have a significant effect on flame response and that such behavior is important in multi-flame configurations. As a result, the structure of the flame and its response to forcing is three-dimensional. Research at Penn State has led to the development of a three-dimensional chemiluminescence flame imaging technique that can be used to characterize the unforced (steady) and forced (unsteady) flame structure of multi-nozzle combustors. Important aspects of the flame response in multi-nozzle combustors which are being studied include flame-flame and flame-wall interactions. Research at Penn State using the recently developed three-dimensional flame imaging technique has shown that spatial variations in local flame confinement must be accounted for to accurately predict global flame response in a multi-nozzle can combustor.

  5. Development of a new method for improving load turndown in fluidized bed combustors: Final report

    SciTech Connect

    Brown, R.C.

    1988-12-01

    The objective of this research was to investigate a new concept in fluidized bed design that improves load turndown capability. This improvement is accomplished by independently controlling heat transfer and combustion in the combustor. The design consists of two fluidized beds, one central and one annular. The central bed serves as the combustion bed. The annular bed is fluidized separately from the combustion bed and its level of fluidization determine the overall heat transfer rate from the combustion bed to the surrounding water jacket. Early theoretical considerations suggested a load turndown exceeding ten was possible for this design. This research consisted of three major phases: development of a computational model to predict heat transfer in the two-bed combustor, heat transfer measurements in hot-and-cold flow models of the combustor, and combustion tests in an optimally designed combustor. The computation model was useful in selecting the design of the combustor. Annular bed width and particle sizes were chosen with the aid of the model. The heat transfer tests were performed to determine if the existing correlations for fluidized bed heat transfer coefficients were sufficiently accurate for high aspect ratio fluidized beds (such as the annular bed in the combustor). Combustion tests were performed in an optimally designed combustor. Three fuel forms were used: double screened, crushed coal, coal-water-limestone mixtures (CWLM), and coal-limestone briquettes. 18 refs., 30 figs., 8 tabs.

  6. Experimental and Computational Study of Trapped Vortex Combustor Sector Rig With Tri-Pass Diffuser

    NASA Technical Reports Server (NTRS)

    Hendricks, R. C.; Shouse, D. T.; Roquernore, W. M.; Burrus, D. L.; Duncan, B. S.; Ryder, R. C.; Brankovic, A.; Liu, N.-S.; Gallagher, J. R.; Hendricks, J. A.

    2004-01-01

    The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL modes of combustion. The present work describes the operational principles of the TVC, and extends diffuser velocities toward choked flow and provides system performance data. Performance data include EINOx results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable compared to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

  7. Wide range operation of advanced low NOx combustors for supersonic high-altitude aircraft gas turbines

    NASA Technical Reports Server (NTRS)

    Roberts, P. B.; Fiorito, R. J.

    1977-01-01

    An initial rig program tested the Jet Induced Circulation (JIC) and Vortex Air Blast (VAB) systems in small can combustor configurations for NOx emissions at a simulated high altitude, supersonic cruise condition. The VAB combustor demonstrated the capability of meeting the NOx goal of 1.0 g NO2/kg fuel at the cruise condition. In addition, the program served to demonstrate the limited low-emissions range available from the lean, premixed combustor. A follow-on effort was concerned with the problem of operating these lean, premixed combustors with acceptable emissions at simulated engine idle conditions. Various techniques have been demonstrated that allow satisfactory operation on both the JIC and VAB combustors at idle with CO emissions below 20 g/kg fuel. The VAB combustor was limited by flashback/autoignition phenomena at the cruise conditions to a pressure of 8 atmospheres. The JIC combustor was operated up to the full design cruise pressure of 14 atmospheres without encountering an autoignition limitation although the NOx levels, in the 2-3 g NO2/kg fuel range, exceeded the program goal.

  8. Experimental and Computational Study of Trapped Vortex Combustor Sector Rig with Tri-pass Diffuser

    NASA Technical Reports Server (NTRS)

    Hendricks, Robert C.; Shouse, D. T.; Roquemore, W. M.; Burrus, D. L.; Duncan, B. S.; Ryder, R. C.; Brankovic, A.; Liu, N.-S.; Gallagher, J. R.; Hendricks, J. A.

    2001-01-01

    The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL (Rich burn/Quick mix/Lean burn) modes of combustion. The present work describes the operational principles of the TVC, and provides detailed performance data on a configuration featuring a tri-pass diffusion system. Performance data include EINOx (NO(sub x) emission index) results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable in comparison to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

  9. Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

    NASA Technical Reports Server (NTRS)

    Brown, William S.

    1991-01-01

    In order to increase the performance of a high performance, advanced expander-cycle engine combustor, higher chamber pressures are required. In order to increase chamber pressure, more heat energy is required to be transferred to the combustor coolant circuit fluid which drives the turbomachinery. This requirement was fulfilled by increasing the area exposed to the hot-gas by using combustor ribs. A previous technology task conducted 2-d hot air and cold flow tests to determine an optimum rib height and configuration. In task C.5 a combustor calorimeter was fabricated with the optimum rib configuration, 0.040 in. high ribs, in order to determine their enhancing capability. A secondary objective was to determine the effects of mixture ratio changers on the enhancement during hot-fire testing. The program used the Rocketdyne Integrated Component Evaluator (ICE) reconfigured into a thrust chamber only mode. The test results were extrapolated to give a projected enhancement from the ribs for a 16 in. long cylindrical combustor at 15 Klb nominal thrust level. The hot-gas wall ribs resulted in a 58 percent increase in heat transfer. When projected to a full size 15K combustor, it becomes a 46 percent increase. The results of those tests, a comparison with previous 2-d results, the effects of mixture ratio and combustion gas flow on the ribs and the potential ramifications for expander cycle combustors are detailed.

  10. Catalytic pyrolysis of wood biomass in an auger reactor using calcium-based catalysts.

    PubMed

    Veses, A; Aznar, M; Martínez, I; Martínez, J D; López, J M; Navarro, M V; Callén, M S; Murillo, R; García, T

    2014-06-01

    Wood catalytic pyrolysis using calcium-based materials was studied in an auger reactor at 450°C. Two different catalysts, CaO and CaO·MgO were evaluated and upgraded bio-oils were obtained in both cases. Whilst acidity and oxygen content remarkable decrease, both pH and calorific value increase with respect to the non-catalytic test. Upgrading process was linked to the fact that calcium-based materials could not only fix the CO2-like compounds but also promoted the dehydration reactions. In addition, process simulation demonstrated that the addition of these catalysts, especially CaO, could favour the energetic integration since a lowest circulation of heat carrier between combustor and auger reactor should be needed. An energy self-sustained system was obtained where thermal energy required for biomass drying and for pyrolysis reaction was supplied by non-condensable gas and char combustion, respectively. PMID:24759640

  11. NONEQUILIBRIUM SULFUR CAPTURE & RETENTION IN AN AIR COOLED SLAGGING COAL COMBUSTOR

    SciTech Connect

    Bert Zauderer

    2003-04-21

    Calcium oxide injected in a slagging combustor reacts with the sulfur from coal combustion to form sulfur-bearing particles. The reacted particles impact and melt in the liquid slag layer on the combustor wall by the centrifugal force of the swirling combustion gases. Due to the low solubility of sulfur in slag, it must be rapidly drained from the combustor to limit sulfur gas re-evolution. Prior analyses and laboratory scale data indicated that for Coal Tech's 20 MMBtu/hour, air-cooled, slagging coal combustor slag mass flow rates in excess of 400 lb/hr should limit sulfur re-evolution. The objective of this 42-month project was to validate this sulfur-in-slag model in a group of combustor tests. A total of 36 days of testing on the combustor were completed during the period of performance of this project. This was more that double the 16 test days that were required in the original work statement. The extra tests were made possible by cost saving innovations that were made in the operation of the combustor test facility and in additional investment of Coal Tech resources in the test effort. The original project plan called for two groups of tests. The first group of tests involved the injection of calcium sulfate particles in the form of gypsum or plaster of Paris with the coal into the 20 MMBtu/hour-combustor. The second group of tests consisted of the entire two-step process, in which lime or limestone is co-injected with coal and reacts with the sulfur gas released during combustion to form calcium sulfate particles that impact and dissolve in the slag layer. Since this sulfur capture process has been validated in numerous prior tests in this combustor, the primary effort in the present project was on achieving the high slag flow rates needed to retain the sulfur in the slag.

  12. Mitsunobu Reactions Catalytic in Phosphine and a Fully Catalytic System

    PubMed Central

    Buonomo, Joseph A; Aldrich, Courtney C

    2015-01-01

    The Mitsunobu reaction is renowned for its mild reaction conditions and broad substrate tolerance, but has limited utility in process chemistry and industrial applications due to poor atom economy and the generation of stoichiometric phosphine oxide and hydrazine by-products that complicate purification. A catalytic Mitsunobu reaction using innocuous reagents to recycle these by-products would overcome both of these shortcomings. Herein we report a protocol that is catalytic in phosphine (1-phenylphospholane) employing phenylsilane to recycle the catalyst. Integration of this phosphine catalytic cycle with Taniguchi’s azocarboxylate catalytic system provided the first fully catalytic Mitsunobu reaction. PMID:26347115

  13. Mitsunobu Reactions Catalytic in Phosphine and a Fully Catalytic System.

    PubMed

    Buonomo, Joseph A; Aldrich, Courtney C

    2015-10-26

    The Mitsunobu reaction is renowned for its mild reaction conditions and broad substrate tolerance, but has limited utility in process chemistry and industrial applications due to poor atom economy and the generation of stoichiometric phosphine oxide and hydrazine by-products that complicate purification. A catalytic Mitsunobu reaction using innocuous reagents to recycle these by-products would overcome both of these shortcomings. Herein we report a protocol that is catalytic in phosphine (1-phenylphospholane) employing phenylsilane to recycle the catalyst. Integration of this phosphine catalytic cycle with Taniguchi's azocarboxylate catalytic system provided the first fully catalytic Mitsunobu reaction. PMID:26347115

  14. Computation of the flow field in an annular gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Cline, Michael C.; Deur, John M.; Micklow, Gerald J.; Harper, Michael R.; Kundu, Krishna P.

    1993-01-01

    The KIVA-II code was modified to calculate the 3D flow field in a typical annular gas turbine combustor. The airblast fuel nozzle, cooling baffle, cooling slots, primary and dilution jets, and effusion cooling (bleed) pads were accounted for in this calculation. The turbulence and combustion were modeled using the k-epsilon model and laminar Arrhenius kinetics, respectively. The fuel was modeled as an evaporating liquid spray. The results illustrate the complicated flow fields present in such combustors. From the results obtained to date it appears that the modified KIVA-II code can be used to study the effects of different annular combustor designs and operating conditions.

  15. The impact of emission standards on the design of aircraft gas turbine engine combustors

    NASA Technical Reports Server (NTRS)

    Rudey, R. A.

    1976-01-01

    The advent of environmental standards for controlling aircraft gas turbine engine emissions has led to a reevaluation of combustor design techniques. Effective emission control techniques have been identified and a wide spectrum of potential applications for these techniques to existing and advanced engines are being considered. Results from advanced combustor concept evaluations and from fundamental experiments are presented and discussed and comparisons are made with existing EPA emission standards and recommended levels for high altitude cruise. The impact that the advanced low emission concepts may impose on future aircraft engine combustor designs and related engine components is discussed.

  16. Effects of fuel nozzle design on performance of an experimental annular combustor using natural gas fuel

    NASA Technical Reports Server (NTRS)

    Wear, J. D.; Schultz, D. F.

    1972-01-01

    Tests of various fuel nozzles were conducted with natural gas fuel in a full-annulus combustor. The nozzles were designed to provide either axial, angled, or radial fuel injection. Each fuel nozzle was evaluated by measuring combustion efficiency at relatively severe combustor operating conditions. Combustor blowout and altitude ignition tests were also used to evaluate nozzle designs. Results indicate that angled injection gave higher combustion efficiency, less tendency toward combustion instability, and altitude relight characteristics equal to or superior to those of the other fuel nozzles that were tested.

  17. Quantitative characterization of a nonreacting, supersonic combustor flowfield using unified, laser-induced iodine fluorescence

    NASA Technical Reports Server (NTRS)

    Fletcher, D. G.; Mcdaniel, J. C.

    1989-01-01

    A calibrated, nonintrusive optical technique, laser-induced iodine fluorescence (LIIF) was used to quantify the steady, compressible flowfield of a nonreacting, supersonic combustor. The combustor was configured with single and staged, transverse-air injection into a supersonic-air freestream behind a rearward-facing step. Pressure, temperature, two-velocity components, and injectant mole fraction were measured with high spatial resolution in the three-dimensional flowfields. These experimental results provide a benchmark set of data for validation of computational fluid dynamic (CFD) codes being developed to model supersonic combustor flowfields.

  18. CFD Investigation of Pollutant Emission in Can-Type Combustor Firing Natural Gas, LNG and Syngas

    NASA Astrophysics Data System (ADS)

    Hasini, H.; Fadhil, SSA; Mat Zian, N.; Om, NI

    2016-03-01

    CFD investigation of flow, combustion process and pollutant emission using natural gas, liquefied natural gas and syngas of different composition is carried out. The combustor is a can-type combustor commonly used in thermal power plant gas turbine. The investigation emphasis on the comparison of pollutant emission such in particular CO2, and NOx between different fuels. The numerical calculation for basic flow and combustion process is done using the framework of ANSYS Fluent with appropriate model assumptions. Prediction of pollutant species concentration at combustor exit shows significant reduction of CO2 and NOx for syngas combustion compared to conventional natural gas and LNG combustion.

  19. Development of an analytical model to assess fuel property effects on combustor performance

    NASA Technical Reports Server (NTRS)

    Sutton, R. D.; Troth, D. L.; Miles, G. A.; Riddlebaugh, S. M.

    1987-01-01

    A generalized first-order computer model has been developed in order to analytically evaluate the potential effect of alternative fuels' effects on gas turbine combustors. The model assesses the size, configuration, combustion reliability, and durability of the combustors required to meet performance and emission standards while operating on a broad range of fuels. Predictions predicated on combustor flow-field determinations by the model indicate that fuel chemistry, as defined by hydrogen content, exerts a significant influence on flame retardation, liner wall temperature, and smoke emission.

  20. Quiet Clean Short-haul Experimental Engine (QCSEE). Double-annular clean combustor technology development report

    NASA Technical Reports Server (NTRS)

    Bahr, D. W.; Burrus, D. L.; Sabla, P. E.

    1979-01-01

    A sector combustor technology development program was conducted to define an advanced double annular dome combustor sized for use in the quiet clean short haul experimental engine (QCSEE). A design which meets the emission goals, and combustor performance goals of the QCSEE engine program was developed. Key design features were identified which resulted in substantial reduction in carbon monoxide and unburned hydrocarbon emission levels at ground idle operating conditions, in addition to very low nitric oxide emission levels at high power operating conditions. Their significant results are reported.

  1. Flame dynamics in a micro-channeled combustor

    SciTech Connect

    Hussain, Taaha; Balachandran, Ramanarayanan; Markides, Christos N.

    2015-01-22

    The increasing use of Micro-Electro-Mechanical Systems (MEMS) has generated a significant interest in combustion-based power generation technologies, as a replacement of traditional electrochemical batteries which are plagued by low energy densities, short operational lives and low power-to-size and power-to-weight ratios. Moreover, the versatility of integrated combustion-based systems provides added scope for combined heat and power generation. This paper describes a study into the dynamics of premixed flames in a micro-channeled combustor. The details of the design and the geometry of the combustor are presented in the work by Kariuki and Balachandran [1]. This work showed that there were different modes of operation (periodic, a-periodic and stable), and that in the periodic mode the flame accelerated towards the injection manifold after entering the channels. The current study investigates these flames further. We will show that the flame enters the channel and propagates towards the injection manifold as a planar flame for a short distance, after which the flame shape and propagation is found to be chaotic in the middle section of the channel. Finally, the flame quenches when it reaches the injector slots. The glow plug position in the exhaust side ignites another flame, and the process repeats. It is found that an increase in air flow rate results in a considerable increase in the length (and associated time) over which the planar flame travels once it has entered a micro-channel, and a significant decrease in the time between its conversion into a chaotic flame and its extinction. It is well known from the literature that inside small channels the flame propagation is strongly influenced by the flow conditions and thermal management. An increase of the combustor block temperature at high flow rates has little effect on the flame lengths and times, whereas at low flow rates the time over which the planar flame front can be observed decreases and the time of

  2. Development of a pressure gain combustor for improved cycle efficiency

    SciTech Connect

    Gemmen, R.S.; Richards, G.A.; Janus, M.C.

    1994-09-01

    This paper presents results from an experimental research program attempting to improve the thermodynamic efficiencies of gas-turbine combustors. An elementary thermodynamic analysis shows that the thermodynamic cycle efficiencies of gas turbines can be significantly improved by using unsteady combustion that achieves quasi-constant-volume combustion. The ability to produce the so-called pressure gain via this process has already been demonstrated by others for pressures less than 3 atmospheres. This paper presents experimental results for pressures up to 11 atmospheres, compares certain process parameters to a numerical simulation, and briefly examines the problem of scale-up. Results of pollutant measurements over the 2--11 atmospheric range of operation are also included.

  3. Municipal waste combustor operator training program: Course manual. Final report

    SciTech Connect

    Beard, J.T.; Lanier, W.S.; Lee, S.Y.

    1993-04-01

    The Course Manual, addresses the training needs of municipal waste combustor (MWC) operators. The training program focuses on the knowledge required by operators for understanding the basis for proper operation and maintenance of MWC`s with particular emphasis on the aspects of combustion which are important for environmental control. The training program includes general introductory material relative to municipal solid waste (MSW) treatment and MSW as a fuel. The bulk of the program addresses the principles of good combustion. The potential sources of air pollution emissions and their control are discussed. Instrumentation, automatic control systems, control room operations and practices, and the troubleshooting of upsets are presented. Special system considerations are included: water teatment, electrical theory, and turbines and generators. Finally, risk management procedures such as preventive maintenance and safety considerations are addressed.

  4. Numerical investigation of chemically reacting flows in ramjet dump combustors

    NASA Technical Reports Server (NTRS)

    Hsieh, Kwang-Chung; Liu, Jong-Shang

    1989-01-01

    The time-dependent Navier-Stokes equations, including second-order turbulence model, are numerically integrated by using four-stage Runge-Kutta scheme to predict the steady-state supersonic flow structures in ramjet dump combustors. The formulation is derived for reacting flows with finite-rate chemistry. In the present study, it is firstly attempted to assess the accuracy of existing high-order turbulence model in supersonic flows. The comparison shows reasonable agreement between calculated and measured data in terms of velocity distributions. It is indicated that a modified constant C-mu for calculating turbulent eddy viscosity is needed in the supersonic flow regime and the adaptive meshing is preferred to capture the recirculation zone. In the reacting flow calculation, the results from a test case of hydrogen and air combustion at premixed conditon show that the rearward facing step is able to increase flow residence time and stabilize the flame in supersonic flows.

  5. The application of lime sorbents in municipal waste combustors

    SciTech Connect

    Benson, L.; Licata, A.

    1998-07-01

    Lime is the sorbent most utilized to control acid gas emissions from Municipal Waste Combustors (MWCs) throughout the world. Line is safe, economical, and easy to handle. In addition to acid gas controls, lime has been demonstrated to reduce mercury and dioxin emissions when used in spray dryers. Lime also has applications in controlling the leachability of heavy metals from MWC ash. Although lime is used throughout the industry, the authors see many misapplications and misunderstandings of this technology. They have seen the wrong type of silos used as well as the wrong size silos. Slaking is a major problem for some plants because they use the wrong water and lime products. This paper will discuss the selection criteria and economics for lime handling and feeding systems with design data. Definitions and the chemistry of lime will be presented to enable design engineers to better prepare systems specifications. This paper will be beneficial to plants planning to upgrade to the MACT standards.

  6. A Simplified Model for Detonation Based Pressure-Gain Combustors

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2010-01-01

    A time-dependent model is presented which simulates the essential physics of a detonative or otherwise constant volume, pressure-gain combustor for gas turbine applications. The model utilizes simple, global thermodynamic relations to determine an assumed instantaneous and uniform post-combustion state in one of many envisioned tubes comprising the device. A simple, second order, non-upwinding computational fluid dynamic algorithm is then used to compute the (continuous) flowfield properties during the blowdown and refill stages of the periodic cycle which each tube undergoes. The exhausted flow is averaged to provide mixed total pressure and enthalpy which may be used as a cycle performance metric for benefits analysis. The simplicity of the model allows for nearly instantaneous results when implemented on a personal computer. The results compare favorably with higher resolution numerical codes which are more difficult to configure, and more time consuming to operate.

  7. Fluidized bed combustor and removable windbox and tube assembly therefor

    DOEpatents

    DeFeo, Angelo; Hosek, William S.

    1981-01-01

    A fluidized bed combustor comprises a housing having a chamber therein with a top having a discharge for the gases which are generated in the chamber and a bottom with a discharge for heated fluid. An assembly is arranged in the lower portion of the chamber and the assembly includes a lower plate which is mounted on a support flange of the housing so that it is spaced from the bottom of the chamber and defines a fluid plenum between it and the bottom of the chamber for the discharge of heated fluid. The assembly includes a heat exchanger inlet plenum having tubes therethrough for the passage of fluidizer air and a windbox above the heat exchanger plenum which has a distributor plate top wall. A portion of the chamber above the top wall defines a fluidized bed.

  8. Fluidized bed combustor and removable windbox and tube assembly therefor

    DOEpatents

    DeFeo, Angelo; Hosek, William

    1983-01-01

    A fluidized bed combustor comprises a housing having a chamber therein with a top having a discharge for the gases which are generated in the chamber and a bottom with a discharge for heated fluid. An assembly is arranged in the lower portion of the chamber and the assembly includes a lower plate which is mounted on a support flange of the housing so that it is spaced from the bottom of the chamber and defines a fluid plenum between it and the bottom of the chamber for the discharge of heated fluid. The assembly includes a heat exchanger inlet plenum having tubes therethrough for the passage of fluidizer air and a windbox above the heat exchanger plenum which has a distributor plate top wall. A portion of the chamber above the top wall defines a fluidized bed.

  9. Performance of a high efficiency advanced coal combustor

    SciTech Connect

    Toqan, M.A.; Paloposki, T.; Yu, T.; Teare, J.D.; Beer, J.M. )

    1989-12-01

    Under contract from DOE-PETC, Combustion Engineering, Inc. undertook the lead-role in a multi-task R D program aimed at development of a new burner system for coal-based fuels; the goal was that this burner system should be capable of being retrofitted in oil- or gas-fired industrial boilers, or usable in new units. In the first phase of this program a high efficiency advanced coal combustor was designed jointly by CE and MIT. Its burner is of the multiannular design with a fixed shrouded swirler in the center immediately surrounding the atomizer gun to provide the primary act,'' and three further annuli for the supply of the secondary air.'' The degree of rotation (swirl) in the secondary air is variable. The split of the combustion air into primary and secondary air flows serves the purpose of flame stabilization and combustion staging, the latter to reduce NO{sub x} formation.

  10. Modeling of a Sequential Two-Stage Combustor, Supplement

    NASA Technical Reports Server (NTRS)

    Hendricks, R. C.; Liu, N.-S.; Gallagher, J. R.; Ryder, R. C.; Brankovic, A.; Hendricks, J. A.

    2005-01-01

    A sequential two-stage, natural gas fueled power generation combustion system is modeled to examine the fundamental aerodynamic and combustion characteristics of the system. The modeling methodology includes CAD-based geometry definition, and combustion computational fluid dynamics analysis. Graphical analysis is used to examine the complex vortical patterns in each component, identifying sources of pressure loss. The simulations demonstrate the importance of including the rotating high-pressure turbine blades in the computation, as this results in direct computation of combustion within the first turbine stage, and accurate simulation of the flow in the second combustion stage. The direct computation of hot-streaks through the rotating high-pressure turbine stage leads to improved understanding of the aerodynamic relationships between the primary and secondary combustors and the turbomachinery.

  11. Role of fuel chemical properties on combustor radiative heat load

    NASA Technical Reports Server (NTRS)

    Rosfjord, T. J.

    1984-01-01

    In an attempt to rigorously study the fuel chemical property influence on combustor radiative heat load, United Technologies Research Center (UTRC) has conducted an experimental program using 25 test fuels. The burner was a 12.7-cm dia cylindrical device fueled by a single pressure-atomizing injector. Fuel physical properties were de-emphasized by selecting injectors which produced high-atomized, and hence rapidly-vaporizing sprays. The fuels were specified to cover the following wide ranges of chemical properties; hydrogen, 9.1 to 15- (wt) pct; total aromatics, 0 to 100 (vol) pct; and naphthalene, 0 to 30 (vol) pct. They included standard fuels, specialty products and fuel blends. Fuel naphthalene content exhibited the strongest influence on radiation of the chemical properties investigated. Smoke point was a good global indicator of radiation severity.

  12. Role of fuel chemical properties on combustor radiative heat load

    NASA Technical Reports Server (NTRS)

    Rosfjord, T. J.

    1984-01-01

    In an attempt to rigorously study the fuel chemical property influence on combustor radiative heat load, UTRC has conducted an experimental program using 25 test fuels. The burner was a 12.7-cm dia cylindrical device fueled by a single pressure-atomizing injector. Fuel physical properties were de-emphasized by selecting injectors which produced highly-atomized, and hence rapidly-vaporizing sprays. The fuels were specified to cover the following wide ranges of chemical properties: hydrogen, 9.1 to 15- (wt) pct; total aromatics, 0 to 100 (vol) pct; and naphthalene, 0 to 30 (vol) pct. They included standard fuels, specialty products and fuel blends. Fuel naphthalene content exhibited the strongest influence on radiation of the chemical properties investigated. Smoke point was a good global indicator of radiation severity.

  13. Prediction of swirling reacting flow in ramjet combustors

    NASA Technical Reports Server (NTRS)

    Lilley, D. G.; Samples, J. W.; Rhode, D. L.

    1981-01-01

    Numerical computations have been undertaken for a basic two-dimensional axisymmetric flowfield which is similar to that found in conventional gas turbine and ramjet combustors. A swirling flow enters a larger chamber via a sudden or gradual expansion. The calculation method involves a staggered grid system for axial and radial velocities, a line relaxation procedure for efficient solution of the equations, a two-equation turbulence energy-turbulence dissipation rate turbulence model, a stairstep boundary representation of the expansion flow, and realistic accommodation of swirl effects. The results include recirculation zone characterization and predicted mean streamline patterns. Predictions with and without chemical reaction are obtained. An associated isothermal experimental flow study is providing a useful data base. Successful outcomes of the work can be incorporated into the more combustion- and hardware-oriented activities of industrial concerns.

  14. Gas turbine combustor exit piece with hinged connections

    DOEpatents

    Charron, Richard C.; Pankey, William W.

    2016-04-26

    An exit piece (66) with an inlet throat (67) that conducts a combustion gas flow (36A) in a path (82) from a combustor (63) to an annular chamber (68) that feeds the first blade section (37) of a gas turbine (26). The exit piece further includes an outlet portion (69) that forms a circumferential segment of the annular chamber. The outlet portion interconnects with adjacent outlet portions by hinges (78A, 78B, 80A, 80B). Each hinge may have a hinge axis (82A, 82B) parallel to a centerline (21) of the turbine. Respective gas flows (36A) are configured by an assembly (60) of the exit pieces to converge on the feed chamber (68) into a uniform helical flow that drives the first blade section with minimal circumferential variations in force.

  15. Modeling of a Sequential Two-Stage Combustor

    NASA Technical Reports Server (NTRS)

    Hendricks, R. C.; Liu, N.-S.; Gallagher, J. R.; Ryder, R. C.; Brankovic, A.; Hendricks, J. A.

    2005-01-01

    A sequential two-stage, natural gas fueled power generation combustion system is modeled to examine the fundamental aerodynamic and combustion characteristics of the system. The modeling methodology includes CAD-based geometry definition, and combustion computational fluid dynamics analysis. Graphical analysis is used to examine the complex vortical patterns in each component, identifying sources of pressure loss. The simulations demonstrate the importance of including the rotating high-pressure turbine blades in the computation, as this results in direct computation of combustion within the first turbine stage, and accurate simulation of the flow in the second combustion stage. The direct computation of hot-streaks through the rotating high-pressure turbine stage leads to improved understanding of the aerodynamic relationships between the primary and secondary combustors and the turbomachinery.

  16. Near-zero emissions combustor system for syngas and biofuels

    SciTech Connect

    Yongho, Kim; Rosocha, Louis

    2010-01-01

    A multi-institutional plasma combustion team was awarded a research project from the DOE/NNSA GIPP (Global Initiative for Prolifereation Prevention) office. The Institute of High Current Electronics (Tomsk, Russia); Leonardo Technologies, Inc. (an American-based industrial partner), in conjunction with the Los Alamos National Laboratory are participating in the project to develop novel plasma assisted combustion technologies. The purpose of this project is to develop prototypes of marketable systems for more stable and cleaner combustion of syngas/biofuels and to demonstrate that this technology can be used for a variety of combustion applications - with a major focus on contemporary gas turbines. In this paper, an overview of the project, along with descriptions of the plasma-based combustors and associated power supplies will be presented. Worldwide, it is recognized that a variety of combustion fuels will be required to meet the needs for supplying gas-turbine engines (electricity generation, propulsion), internal combustion engines (propulsion, transportation), and burners (heat and electricity generation) in the 21st Century. Biofuels and biofuel blends have already been applied to these needs, but experience difficulties in modifications to combustion processes and combustor design and the need for flame stabilization techniques to address current and future environmental and energy-efficiency challenges. In addition, municipal solid waste (MSW) has shown promise as a feedstock for heat and/or electricity-generating plants. However, current combustion techniques that use such fuels have problems with achieving environmentally-acceptable air/exhaust emissions and can also benefit from increased combustion efficiency. This project involves a novel technology (a form of plasma-assisted combustion) that can address the above issues. Plasma-assisted combustion (PAC) is a growing field that is receiving worldwide attention at present. The project is focused on

  17. Control of a fluidized bed combustor using fuzzy logic

    SciTech Connect

    Koffman, S.J.; Brown, R.C.; Fullmer, R.R.

    1996-01-01

    Fuzzy logic--an artificial intelligence technique--can be employed to exploit the wealth of information human experts have learned about complex systems while attempting to control them. This information is usually of a qualitative nature that is unusable by rigid conventional control techniques. Fuzzy logic, uses as a control method, manipulates linguistically expressed, heuristic knowledge from a human expert to derive control actions for a described system. As an alternative approach to classical controls, fuzzy logic is examined for start-up control and normal regulation of a bubbling fluidized bed combustor. To validate the fuzzy logic approach, the fuzzy controller is compared to a classical proportional and integral (PI) controller, commonly used in industrial applications, designed by Ziegler-Nichols tuning.

  18. Erosivity of particles in circulating fluidized bed combustors

    SciTech Connect

    Levy, A.V.; Wang, B.Q.; Geng, G.Q. ); Mack, W. )

    1989-01-01

    The metal wastage of superheater tubes in the convection pass region of circulating fluidized bed combustors (CFBC) is a current problem. An investigation was carried out to determine what kinds of metal loss rates and mechanisms occurred when the various types of particles in CFBC's were used as the erodent in a laboratory blast nozzle tester. The laboratory tests were compared to in-service exposures. A loss mechanism was established that was based on metal losses and the observed microstructures of E-C surfaces. It was additionally determined that more angular and larger size particles have a greater level of erosivity and that the particles must be strong enough not to shatter upon impact in order to be erosive. Favorable comparisons between laboratory and in-service loss mechanisms are reported.

  19. Combustion of hydrogen in an experimental trapped vortex combustor

    NASA Astrophysics Data System (ADS)

    Wu, Hui; Chen, Qin; Shao, Weiwei; Zhang, Yongliang; Wang, Yue; Xiao, Yunhan

    2009-09-01

    Combustion performances of pure hydrogen in an experimental trapped vortex combustor have been tested under different operating conditions. Pressure fluctuations, NOx emissions, OH distributions, and LBO (Lean Blow Out) were measured in the tests. Results indicate that the TVC test rig has successfully realized a double vortex construction in the cavity zone in a wide range of flow conditions. Hydrogen combustion in the test rig has achieved an excellent LBO performance and relatively low NOx emissions. Through comparison of dynamic pressure data, OH fluctuation images, and NOx emissions, the optimal operating condition has been found out to be Φp =0.4, fuel split =0.4, and primary air/fuel premixed.

  20. Parametric Design of Injectors for LDI-3 Combustors

    NASA Technical Reports Server (NTRS)

    Ajmani, Kumud; Mongia, Hukam; Lee, Phil

    2015-01-01

    Application of a partially calibrated National Combustion Code (NCC) for providing guidance in the design of the 3rd generation of the Lean-Direct Injection (LDI) multi-element combustion configuration (LDI-3) is summarized. NCC was used to perform non-reacting and two-phase reacting flow computations on several LDI-3 injector configurations in a single-element and a five-element injector array. All computations were performed with a consistent approach for mesh-generation, turbulence, spray simulations, ignition and chemical kinetics-modeling. Both qualitative and quantitative assessment of the computed flowfield characteristics of the several design options led to selection of an optimal injector LDI- 3 design that met all the requirements including effective area, aerodynamics and fuel-air mixing criteria. Computed LDI-3 emissions (namely, NOx, CO and UHC) will be compared with the prior generation LDI- 2 combustor experimental data at relevant engine cycle conditions.