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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. Transient catalytic combustor model

    NASA Astrophysics Data System (ADS)

    Tien, J. S.

    1981-05-01

    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.

  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. 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.

  6. 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.

  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. Advanced Catalytic Combustors for Low Pollutant Emissions

    DTIC Science & Technology

    1979-11-01

    concepts were selected for further design efforts. Results of the Phase I design effort indicate that catalytic combustion is a promising means for...L VAMD. Recent efforta to develop fuel-air carburetion concepts fol use in gas turbine catalytic combustion systems, which are summarized in...Radial/Axial Parallel-Staged combustor shown in F’gure 10 (Con- cept 6) is essentially two separate combustion systems in parallel. In this design concept

  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 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.

  13. 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.

  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. 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.

  16. 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

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. Performance of a catalytic reactor at simulated gas turbine combustor operating conditions

    NASA Technical Reports Server (NTRS)

    Anderson, D. N.; Tacina, R. R.; Mroz, T. S.

    1975-01-01

    The performance of a catalytic reactor 12 cm in diameter and 17 cm long was evaluated at simulated gas turbine combustor operating conditions using premixed propane and air. Inlet temperatures of 600 and 800 K, pressures of 3 and 6 atm, and reference velocities of 9 to 30 m/s were tested. Data were taken for equivalence ratios as high as 0.43. The operating range was limited on the low-temperature side by very poor efficiency; the minimum exit temperature for good performance ranged from 1400 to 1600 K depending on inlet conditions. As exit temperatures were raised above this minimum, emissions of unburned hydrocarbons decreased, carbon monoxide emissions became generally less than 1 g CO/kg fuel, and nitrogen oxides were less than about 0.1 g NO2/kg fuel.

  7. 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.

  8. 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.

  9. Catalytic combustion: an investigation of combustor geometry effects. Shelton Energy Research report No. 8272R

    SciTech Connect

    Shelton, J.; Graeser, L.

    1982-01-01

    A Riteway Model 37 wood stove was modified to accept a catalytic combustor. Twenty-six metal plates coated with a platinum/palladium catalyst were assembled into one, two or three layers (26, 13 and 9 plates per layer). The stove's energy efficiency was measured without the catalyst and with the catalyst in each of its three geometries. A combination of room calorimetry and flue gas heat loss measurements was used for these determinations. The one-layer catalyst increased the unit's combustion efficiency by 11 percentage points to 95 percent at a power output of 35,000 Btu/h. Increases of 8 and 5 percentage points were recorded for the two- and three-layer geometries, respectively. In order to investigate relative creosote plugging rates in the event of catalyst failure or low temperature operation, two samples of each substrate geometry were installed in the flues of six identical stoves and cool fires were burned for four days. Total creosote accumulation was greatest for the sparser geometries, but it was less hazardous - the pressure drop for a given flow was less because of the wider spacing. Total plugging is a definite possibility in relatively short time periods for all geometries. Two passive bypass systems were investigated to detour flue gases around a total obstruction of the catalyst. One of these recirculated flue gas during normal stove operation, but both alleviated the possible safety hazard of smoke spillage. Finally, a theory of catalyst design was developed and tested; it predicted the experimental combustion efficiencies within two percentage points.

  10. 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%.

  11. Fuel cell system combustor

    DOEpatents

    Pettit, William Henry

    2001-01-01

    A fuel cell system including a fuel reformer heated by a catalytic combustor fired by anode and cathode effluents. The combustor includes a turbulator section at its input end for intimately mixing the anode and cathode effluents before they contact the combustors primary catalyst bed. The turbulator comprises at least one porous bed of mixing media that provides a tortuous path therethrough for creating turbulent flow and intimate mixing of the anode and cathode effluents therein.

  12. Fuel rich catalytic comustion: The first stage of a two-stage combustor

    NASA Technical Reports Server (NTRS)

    Brabbs, T. A.; Olson, S. L.

    1984-01-01

    An experimental program demonstrated that fuel-rich catalytic combustion can be accomplished soot free as long as the combustion temperature is less than the temperature at the rich limit of combustion. Although soot was not measured directly, three pieces of data strongly suggest that it was not present: (1) the product gases were completely transparent and produced no radiation characteristic of soot, (2) measured reaction temperatures followed closely those calculated for equilibrium with no soot present, and (3) over 99 percent of the carbon was accounted for in the measured reaction products. Data for two catalyst configurations were taken along with gas samples at two locations downstream of the catalyst bed.

  13. Rich catalytic injection

    DOEpatents

    Veninger, Albert

    2008-12-30

    A gas turbine engine includes a compressor, a rich catalytic injector, a combustor, and a turbine. The rich catalytic injector includes a rich catalytic device, a mixing zone, and an injection assembly. The injection assembly provides an interface between the mixing zone and the combustor. The injection assembly can inject diffusion fuel into the combustor, provides flame aerodynamic stabilization in the combustor, and may include an ignition device.

  14. 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.

  15. 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.

  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. 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.

  18. High pressure combustor for generating steam downhole

    SciTech Connect

    Retallick, W.B.

    1983-08-09

    A catalytic combustor for generating a mixture of steam and combustion gas is located downhole in oil well, so that the gas mixture can be injected directly into the oil reservoir to displace heavy oils from the reservoir. There can be a single stage of catalytic combustion, or there can be a stage of thermal combustion followed by a catalytic stage. In either case the purpose of the catalyst is drive the combustion to completion so that the gas mixture contains no soot that would plug the reservoir.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Variable volume combustor

    DOEpatents

    Ostebee, Heath Michael; Ziminsky, Willy Steve; Johnson, Thomas Edward; Keener, Christopher Paul

    2017-01-17

    The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Composite Matrix Experimental Combustor

    DTIC Science & Technology

    1994-04-01

    Preliminary (Macro) Combustor Design ............................. 28 4.1 Preliminary Design Study-Early Concept Combustion System ............. 28 4.2...provided in Appendix B. 4.1 PRELIMINARY DESIGN STUDY-EARLY CONCEPT COMBUSTION SYSTEM The preliminary design effort resulted in the selection of the early...overall flowpath. The concept I combustor is a compact, annular, reverse-flow design incorporating a single row of primary combustion air holes and a

  9. 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.

  10. Combustor liner support assembly

    NASA Technical Reports Server (NTRS)

    Halila, Ely E. (Inventor)

    1994-01-01

    A support assembly for a gas turbine engine combustor includes an annular frame having a plurality of circumferentially spaced apart tenons, and an annular combustor liner disposed coaxially with the frame and including a plurality of circumferentially spaced apart tenons circumferentially adjoining respective ones of the frame tenons for radially and tangentially supporting the liner to the frame while allowing unrestrained differential thermal radial movement therebetween.

  11. 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.

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

    SciTech Connect

    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.

  13. Advanced Low-Emissions Catalytic Combustor Program

    DTIC Science & Technology

    1981-06-01

    Program Goals 5 3.2 Reference Engine Description 6 4.0 GENERAL AEROTHERMODYNAMIC AND MECHANICAL DESIGN STUDIES 11 4.1 Combustion Concepts 11 4.2...Pressure Loss Variation with Reactor Length 29 22 Preliminary Flowpath for Concept 2 31 23 Dependence of Idle Combustion Efficiency on Equivalence Ratio...Controlled Variable Geometry 52 38 Concept 4: Combustion Stability Margin Dependence on Cruise Conditions for Engine Speed-Controlled Variable

  14. 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.

  15. 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.

  16. Combustor burner vanelets

    DOEpatents

    Lacy, Benjamin [Greer, SC; Varatharajan, Balachandar [Loveland, OH; Kraemer, Gilbert Otto [Greer, SC; Yilmaz, Ertan [Albany, NY; Zuo, Baifang [Simpsonville, SC

    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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. Combustor and method for distributing fuel in the combustor

    SciTech Connect

    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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Multi-port dump combustor

    SciTech Connect

    Dale, L. A.; Grenleski Jr., S. E.; Keirsey, J. L.; Stevens, C. E.

    1985-09-10

    A four-ported dump combustor is designed for use with a ramjet engine and provides high combustion efficiency and pressure recovery for length-to-diameter (L/D) ratios of between 1.3 and 4.4, over a range of operating conditions.

  14. 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.

  15. 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.

  16. 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.

  17. Integrated CFD modeling of gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Fuller, E. J.; Smith, C. E.

    1993-01-01

    3D, curvilinear, multi-domain CFD analysis is becoming a valuable tool in gas turbine combustor design. Used as a supplement to experimental testing. CFD analysis can provide improved understanding of combustor aerodynamics and used to qualitatively assess new combustor designs. This paper discusses recent advancements in CFD combustor methodology, including the timely integration of the design (i.e. CAD) and analysis (i.e. CFD) processes. Allied Signal's F124 combustor was analyzed at maximum power conditions. The assumption of turbulence levels at the nozzle/swirler inlet was shown to be very important in the prediction of combustor exit temperatures. Predicted exit temperatures were compared to experimental rake data, and good overall agreement was seen. Exit radial temperature profiles were well predicted, while the predicted pattern factor was 25 percent higher than the harmonic-averaged experimental pattern factor.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Slagging retrofit pulsed coal combustor: Final report

    SciTech Connect

    Not Available

    1987-01-01

    A concept for a novel form of slagging retrofit pulsed coal combustor was tested in the laboratory. The combustor is based on controlled use of a form of high pressure amplitude combustion instability. The approach adopted was to resolve, in single pulse experiments, the basic technical issues arising in the development of the combustor. In a cold flow device, the issues of coal spatial distribution were addressed and a combustor and solids disperser configuration was developed to give uniform coal distribution in the combustor. Single pulse ignition experiments were conducted to determine the pressure rise in combustor, pressure rise-decay times, and coal conversion a function of various operating variables. Coal injection, flame propagation, and blowdown times leading to potential combustor size reduction of three times over steady flow combustors were demonstrated. The results give high pressure exhaust leading to potentially improved downstream heat transfer and reduced boiler size. Finally, zero-, one-, and two-dimensional mathematical models were developed in support of the experiments and also to provide design capability. 11 refs., 43 figs.

  3. 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.

  4. Combustor with non-circular head end

    DOEpatents

    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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Rich burn combustor technology at Pratt and Whitney

    NASA Technical Reports Server (NTRS)

    Lohmann, Robert P.; Rosfjord, T. J.

    1992-01-01

    The topics covered include the following: near term objectives; rich burn quick quench combustor (RBQC); RBQC critical technology areas; cylindrical RBQQ combustor rig; modular RBQQ combustor; cylindrical rig objectives; quench zone mixing; noneffusive cooled liner; variable geometry requirements; and sector combustor rig.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Experimental clean combustor program: Noise study

    NASA Technical Reports Server (NTRS)

    Sofrin, T. G.; Riloff, N., Jr.

    1976-01-01

    Under a Noise Addendum to the NASA Experimental Clean Combustor Program (ECCP) internal pressure fluctuations were measured during tests of JT9D combustor designs conducted in a burner test rig. Measurements were correlated with burner operating parameters using an expression relating farfield noise to these parameters. For a given combustor, variation of internal noise with operating parameters was reasonably well predicted by this expression but the levels were higher than farfield predictions and differed significantly among several combustors. For two burners, discharge stream temperature fluctuations were obtained with fast-response thermocouples to allow calculation of indirect combustion noise which would be generated by passage of the temperature inhomogeneities through the high pressure turbine stages of a JT9D turbofan engine. Using a previously developed analysis, the computed indirect combustion noise was significantly lower than total low frequency core noise observed on this and several other engines.

  9. 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.

  10. Application of Gas Analysis to Combustor Research

    NASA Technical Reports Server (NTRS)

    Hibbard, R. R.; Evans, Albert

    1959-01-01

    The performance of turbine-engine combustors usually is given in terms of operating limits and combustion efficiency. The latter property is determined most often by measuring the increase in enthalpy across the combustor through the use of thermocouples. This investigation was conducted to determine the ability of gas-analytical techniques to provide additional information about combustor performance. Gas samples were taken at the outlet and two upstream stations and their compositions determined. In addition to over-all combustion efficiency, estimates of local fuel-air ratios, local combustion efficiencies, and heat-release rates can be made. Conclusions can be drawn concerning the causes of combustion inefficiency and may permit corrective design changes to be made more intelligently. The purpose of this investigation was not to present data for a given combustor but rather to show the types and value of additional information that can be gained from gas-analytical data.

  11. Operational Characteristics of an Ultra Compact Combustor

    DTIC Science & Technology

    2014-03-27

    recirculation conditions. If fuel is mixed into core flows too quickly or escapes into cold spots of a combustor liner , then it will remain...of fuel like a cylinder in a cross flow as it circulates about the cavity. The impact of the flow upon these cylindrical plumes creates a stagnation...combustor liner or the entrainment of cavity flow into core flow within a UCC. Therefore, the test matrix in Table 3.5 was built to highlight changes to

  12. Hypermixer Pylon Fuel Injection for Scramjet Combustors

    DTIC Science & Technology

    2008-09-11

    wall cavities, intrusive geometries like pylons and struts , or a combination thereof. The type of fuel injection strategy largely depends on the...be affixed to combustor walls or placed on intru- sive geometries like struts or pylons. Wall injection comes with an inherent combustor airflow...only one wall . Using in-stream struts and pylons as fueling devices has been a common practice in scramjet design. Much research continues to be

  13. Multiducted Inlet Combustor Research and Development.

    DTIC Science & Technology

    1982-10-01

    qualitative data from the multi-ducted inlet combustor configuration for flow analysis and matematical modeling purposes. The major portion of the support...data on multi-ducted inlet combustor configurations. These efforts will provide the information necessary to perform flow field analysis and aid in the...instrumentation, test program, data reduction, data presentation, flow field analysis and math modelling efforts, and conclusions and recommendations. SECTION 2

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Performance of low-Btu fuel gas turbine combustors

    SciTech Connect

    Bevan, S.; Bowen, J.H.; Feitelberg, A.S.; Hung, S.L.; Lacey, M.A.; Manning, K.S.

    1995-11-01

    This reports on a project to develop low BTU gas fuel nozzle for use in large gas turbine combustors using multiple fuel nozzles. A rich-quench-lean combustor is described here which reduces the amount of NO{sub x} produced by the combustion of the low BTU gas. The combustor incorporates a converging rich stage combustor liner, which separates the rich stage recirculation zones from the quench stage and lean stage air.

  5. Catalytic Reforming

    SciTech Connect

    Little, D.M.

    1985-01-01

    Don Little's Catalytic Reforming deals exclusively with reforming. With the increasing need for unleaded gasoline, the importance of this volume has escalated since it combines various related aspects of reforming technology into a single publication. For those with no practical knowledge of catalytic reforming, the chemical reactions, flow schemes and how the cat reformer fits into the overall refinery process will be of interest. Contents include: Catalytic reforming in refinery processing: How catalytic reformers work - chemical reactions; Process design; The catalyst, process variables and unit operation; Commercial processes; BTX operation; Feed preparation; naphtha hydrotreating and catalytic reforming; Index.

  6. 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.

  7. Multi-Ducted Inlet Combustor Research and Development.

    DTIC Science & Technology

    1983-11-01

    of a reactor or combustor as defined in equation (1) is the combustor volume divided by the fluid flow rate through the combustor. Therefore, for a...Development Laboratories, Inc., Costa Mesa, California, March, 1983. 3. 0. Levenspiel , Chemical Reaction Engineering, John Wiley and Sons, 1962. 59 •rac v £98 kg3-ඃ-,162-;8b

  8. 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.

  9. 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.

  10. 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.

  11. Flow establishment in a generic scramjet combustor

    NASA Astrophysics Data System (ADS)

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

    1990-10-01

    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.

  12. Laser diagnostics on a hypersonic combustor

    NASA Technical Reports Server (NTRS)

    Taylor, David J.; Oldenborg, R. C.; Tiee, J. J.; Northam, G. Burton; Antcliff, Richard R.; Cutler, Andrew D.; Jarrett, O.; Smith, M. W.

    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.

  13. 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.

  14. Variable volume combustor with aerodynamic support struts

    DOEpatents

    Ostebee, Heath Michael; Johnson, Thomas Edward; Stewart, Jason Thurman; Keener, Christopher Paul

    2017-03-07

    The present application provides 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 number of support struts supporting the fuel nozzles and providing the flow of fuel therethrough. The support struts may include an aerodynamic contoured shape so as to distribute evenly a flow of air to the micro-mixer fuel nozzles.

  15. 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.

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

    DTIC Science & Technology

    1994-06-01

    shock tests were performed to addition, there were 16 static pressure taps assess the durability of the CME combustor . The throughout the combustor rig...Scheme for Small Gas Turbine Combustors ," AIAA Paper No. 7. M.D. Paskin, H.C. Mongia , and W.A. Acosta, 90-2158, July 1990. "An Efficient Liner Cooling...for Combustor Applications," NASA CR-168103, April 1983. 8. M.D. Paskin and H.C. Mongia , "Composite Matrix Experimnental Combustor ," NASA CR- 5. D.B

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  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. 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.

  8. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc; Bowles, Jeffrey V.

    1989-01-01

    The Wave Combustor is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture and thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter, lighter engine compared to the scramjet. This engine, which is called the Oblique Detonation Wave Engine (ODWE), can then be utilized to provide a smaller, lighter vehicle or to provide a higher payload capability for a given vehicle weight. An analysis of the performance of a conceptual trans-atmospheric vehicle powered by an ODWE is given here.

  9. 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.

  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. Laser Diagnostic System Validation and Ultra-Compact Combustor Characterization

    DTIC Science & Technology

    2008-03-01

    run 15 Specified V, S same as last run 16 Chapman - Jouguet detonation of last-run mixture 17 One of the above at specified composition 18...turbofan engine (Ref. 7). Since a traditional combustor is too big to fit between turbine stages, the idea of an ultra-compact combustor emerged...obtained along with an enhanced combustion efficiency of about 18 99% (Ref. 12). Compared to traditional swirl stabilized combustors, TVC

  14. 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.

  15. Porous Media Combustors for Clean Gas Turbine Engines

    DTIC Science & Technology

    2007-11-02

    emissions , no cooling requirement for the! combustor itself and the potential to operate free from combustion- induced noise. The reduced combustion...that the combustor operates in a "super-adiabatic" mode, with low emissions . Intrinsic pressure loss is within values, commonly accepted for propulsion...principles for low emissions turbulent flame gas turbine combustors are well established. The preferred strategy remains lean burn, often with staging to

  16. 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).

  17. 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.

  18. 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.

  19. NASA/GE advanced low emissions combustor program

    NASA Technical Reports Server (NTRS)

    Ekstedt, E. E.; Fear, J. S.

    1987-01-01

    The Advanced Low Emissions Combustor Program consisted of the design and testing of advanced combustor concepts utilizing lean, premixed, prevaporized fuel and variable geometry. The objective was to evaluate the potential of these combustor systems to provide very low pollutant emissions levels, superior performance and high durability relative to contemporary combustor designs. Four full annular combustor concepts were designed and fabricated for a 30:1 pressure ratio high bypass turbofan engine. The four full annular combustors with active variable geometry were tested at pressures up to approximately 0.7 MPa with Jet A fuel. The two most promising concepts were also tested in a high pressure sector combustor test rig capable of operation at the maximum engine pressures. The high pressure sector combustor tests were conducted with Jet A and a fuel with reduced hydrogen content. Results of the sector combustor tests are presented in this paper. The potential for very low emissions with premixed fuel was demonstrated. However, autoignition or flashback within the premixing systems was encountered at high pressures. Further development effort is required to address this problem area.

  20. 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.

  1. Apparatus and method for cooling a combustor cap

    SciTech Connect

    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.

  2. Numerical Analysis and Optimization of the Ultra Compact Combustor

    DTIC Science & Technology

    2005-03-01

    Armstrong, Jason M, “Effect of Equivalence Ratio on G-Loading on In-situ Mea- surements of Chemiluminescence in an Ultra Compact Combustor,” M.S. thesis...S., “Experimental and Computational Study of Trapped Vortex Combustor Sector Rig with Tri-Pass Diffuser,” NASA/TM–2004-212507, Jan 2004. 10. Heywood

  3. Variable volume combustor with nested fuel manifold system

    SciTech Connect

    McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward; Ostebee, Heath Michael

    2016-09-13

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles, a fuel manifold system in communication with the micro-mixer fuel nozzles to deliver a flow of fuel thereto, and a linear actuator to maneuver the micro-mixer fuel nozzles and the fuel manifold system.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. Premixing and flash vaporization in a two-stage combustor

    SciTech Connect

    Sjoblom, B.G.A.

    1982-01-01

    A double recirculation zone two-stage combustor fitted with airblast atomizers has been investigated in a previous work. This paper describes further tests with premixing tubes in the secondary combustion zone. Flash vaporization was employed to ensure complete vaporization of the secondary fuel, which was heated to 600K by the combustor inlet air. 9 refs.

  10. Variable volume combustor with pre-nozzle fuel injection system

    SciTech Connect

    Keener, Christopher Paul; Johnson, Thomas Edward; McConnaughhay, Johnie Franklin; Ostebee, Heath Michael

    2016-09-06

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system.

  11. 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

  12. 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.

  13. 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.

  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. System and method for reducing combustion dynamics in a combustor

    SciTech Connect

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

    2016-11-29

    A system for reducing combustion dynamics in a combustor includes an end cap that extends radially across the combustor and includes an upstream surface axially separated from a downstream surface. A combustion chamber is downstream of the end cap, and tubes extend from the upstream surface through the downstream surface. Each tube provides fluid communication through the end cap to the combustion chamber. The system further includes means for reducing combustion dynamics in the combustor. A method for reducing combustion dynamics in a combustor includes flowing a working fluid through tubes that extend axially through an end cap that extends radially across the combustor and obstructing at least a portion of the working fluid flowing through a first set of the tubes.

  16. L-star pulsed coal combustor for residential space heating

    SciTech Connect

    Not Available

    1989-03-01

    This quarter, substantial improvement in the coal carbon conversion was achieved. Specifically, for a scaled-down version of the residential combustor, coal carbon conversions exceeding 97 percent were realized, when utilizing methane as carrier gas for the coal. Design changes include insulation of the combustor, introduction of a flame holder, combustion air preheat and presence of an obstructing plate at the combustor exhaust port. Only the first two changes contributed towards substantial improvement in coal conversion. In addition, monitoring of CH{sub 4} concentration in the exhaust gases gave a real time indication of the combustor performance. Finally, the results of experiments performed in this quarter contributed to design changes that have led to a combustor that has achieved the program goal of > 99 percent conversion of coal carbon. 5 figs., 2 tabs.

  17. 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.

  18. 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.

  19. 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.

  20. Combined fluidized bed retort and combustor

    DOEpatents

    Shang, Jer-Yu; Notestein, John E.; Mei, Joseph S.; Zeng, Li-Wen

    1984-01-01

    The present invention is directed to a combined fluidized bed retorting and combustion system particularly useful for extracting energy values from oil shale. The oil-shale retort and combustor are disposed side-by-side and in registry with one another through passageways in a partition therebetween. The passageways in the partition are submerged below the top of the respective fluid beds to preclude admixing or the product gases from the two chambers. The solid oil shale or bed material is transported through the chambers by inclining or slanting the fluidizing medium distributor so that the solid bed material, when fluidized, moves in the direction of the downward slope of the distributor.

  1. Radial inlet guide vanes for a combustor

    DOEpatents

    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.

  2. 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.

  3. Catalytic combustion of heavy partially-vaporized fuels

    NASA Technical Reports Server (NTRS)

    Rosfjord, T. J.

    1980-01-01

    An experimental program to demonstrate efficient catalytic combustion of fuel-lean and fuel-rich mixtures of residual fuel and air, and to assess the influence of incomplete fuel vaporization on the performance of a catalytic reactor is being conducted. A 7.5-cm diameter catalytic reactor was designed and will be tested over a matrix of conditions representative of a gas turbine combustor inlet. For each of three test phases, two series of tests with a uniform but poorly vaporized (less than 50 percent) mixture of No. 6 fuel oil and air will be performed. In the first series, the non-vaporized fuel will be contained in a spray of droplets with a Sauter Mean Diameter (SMD) less than 30 microns. In the second series, the non-vaporized fuel will be characterized by a spray SMD approximately equal to 100 microns. The designs of the fuel injection system and the catalytic reactor are described in this paper.

  4. Laser-Induced Fluorescence and Synthetic Jet Fuel Analysis in the Ultra Compact Combustor

    DTIC Science & Technology

    2009-12-01

    flame temperatures, emissions and other characteristics. 6 II. Theory and Previous Research II.1 Standard Gas Turbine Engine Combustor A...losses occur in standard combustors resulting in decreased efficiencies and increased emissions . Combustors create heat and entropy (S) in the...and lowering of the amount of harmful emissions produced. Conventional combustor designs are limited by the fact that combustion reactions require

  5. ULTRA-LOW NOX ADVANCED VORTEX COMBUSTOR

    SciTech Connect

    Ryan G. Edmonds; Robert C. Steele; Joseph T. Williams; Douglas L. Straub; Kent H. Casleton; Avtar Bining

    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.

  6. 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.

  7. 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.

  8. 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.

  9. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Bowles, Jeffrey V.; Adelman, Henry G.; Cambier, Jean-Luc

    1989-01-01

    A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

  10. Investigation of combustion instability in ramjet combustors

    SciTech Connect

    Reuter, D.M.

    1988-01-01

    This research is concerned with investigation of the mechanisms responsible for the driving of longitudinal instabilities in dump-type ramjet combustors. In particular, the coupling between the core flame which is stabilized at the entrance of the combustor and the longitudinal acoustic field was studied. The time-dependent structure of premixed V-shaped flames was experimentally examined using pressure measurements, space- and time-resolved C-H radical radiation measurements, high-speed shadow cine photography, and laser-Doppler velocimetry. The investigation revealed that the acoustic energy to sustain the instability is mainly supplied by the oscillatory heat release from the flame. Based on this finding, a model was developed that is capable of predicting the acoustic pressure spectrum from measured heat-release rates. Furthermore, it was shown that the periodic heat-release rates largely result from periodic changes in the flame surface area caused by acoustically triggered symmetric vortex shedding in the wake of the flame holders. Lastly, experiments were conducted that used this mechanism to show the suppression of instabilities at the fundamental acoustic mode by staggering multiple flames so that the unsteady heat release fields destructively interfere with one another.

  11. 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.

  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. Emission characteristics of a liquid spray sudden expansion combustor using computational fluid dynamics

    NASA Astrophysics Data System (ADS)

    Rodriguez, Daniel

    A sudden expansion combustor (SUE) is analyzed using computation fluid dynamics (CFD). CO emissions and NOx emissions are computed for various operating conditions of the SUE combustor using a can type and an annular type geometrical configurations. The goal of this thesis is to see if the SUE combustor is a viable alternative to conventional combustors which utilize swirlers. It is found that for the can type combustor the NO x emissions were quite low compared to other combustor types but the CO emissions were fairly high. The annular combustor shows better CO emissions compared to the can type, but the CO emissions are still high compared to other combustors. Emissions can be improved by providing better mixing in the primary combustion zone. The SUE combustor design needs to be further refined in order for it to be a viable alternative to conventional combustors with swirlers.

  14. 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.

  15. Preliminary investigation of a two-zone swirl flow combustor

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    The effect of full-annular swirling-flow on a flow-zone combustor design is investigated. Swirl flow angles of 25, 35, and 45 degrees were investigated in a combustor design envelope typical of those used in modern engines. The two-zone combustor had 24 pilot-zone fuel injectors and 24 main-fuel injectors located in the centerbody between the pilot and swirl passage. Combustor performance was determined at idle, and two parametric 589 K inlet temperature conditions. Combustor performance was highest with the 45 degree swirl vane design; at the idle condition, combustion efficiency was 99.5 percent. The 45 degree swirl vane also produced the lowest pattern factor of the three angles and showed a combustor lean blowout limit below a 0.001 fuel-air ratio. Combustor total pressure drop varied from a low of 4.6 percent for the 25 degree swirl to a high of 4.9 percent for the 45 degree swirl.

  16. 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

  17. 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.

  18. 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.

  19. Ramjet-Mode Operation in a Combined Cycle Engine Combustor

    NASA Astrophysics Data System (ADS)

    Kato, Kanenori; Kudo, Kenji; Murakami, Atsuo; Tani, Kouichiro; Kanda, Takeshi

    A rocket-ramjet combined-cycle engine was tested in ramjet-mode. The combustor model had two rockets in the combustor section. They were used as an igniter in this operation mode. In the preliminary tests, the downstream combustion ramjet-mode was demonstrated with a 1.4-degree of divergent duct condition. In this study, the upstream and downstream combustion ramjet-mode operations were applied to the combined cycle engine model with large angle of divergent duct condition. In the case of upstream combustion ramjet-mode, the combustion condition at the exit of the combustor showed high combustion efficiency.

  20. 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.

  1. 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.

  2. 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

  3. 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

  4. Oxy-combustor operable with supercritical fluid

    DOEpatents

    Brun, Klaus; McClung, Aaron M.; Owston, Rebecca A.

    2017-04-04

    An oxy-combustor is provided which comprises a combustion vessel including at least one solid fuel slurry inlet port, at least one oxygen inlet port and at least one supercritical fluid inlet port, wherein the combustion vessel is operable at an operating pressure of at least 1,100 psi; an interior of the combustion vessel comprises a combustion chamber and a supercritical fluid infusion chamber surrounding at least a part of the combustion chamber, the supercritical fluid infusion chamber and the combustion chamber are separated by a porous liner surrounding the combustion chamber, and the supercritical infusion chamber is located between the porous liner and an outer casing of the combustion vessel.

  5. 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.

  6. 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.

  7. 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.

  8. CFD Code Development for Combustor Flows

    NASA Technical Reports Server (NTRS)

    Norris, Andrew

    2003-01-01

    During the lifetime of this grant, work has been performed in the areas of model development, code development, code validation and code application. For model development, this has included the PDF combustion module, chemical kinetics based on thermodynamics, neural network storage of chemical kinetics, ILDM chemical kinetics and assumed PDF work. Many of these models were then implemented in the code, and in addition many improvements were made to the code, including the addition of new chemistry integrators, property evaluation schemes, new chemistry models and turbulence-chemistry interaction methodology. Validation of all new models and code improvements were also performed, while application of the code to the ZCET program and also the NPSS GEW combustor program were also performed. Several important items remain under development, including the NOx post processing, assumed PDF model development and chemical kinetic development. It is expected that this work will continue under the new grant.

  9. 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.

  10. A clean coal combustion technology-slagging combustors

    SciTech Connect

    Chang, S. L.; Berry, G. F.

    1989-03-01

    Slagging combustion is an advanced clean coal technology technique characterized by low NOx and SOx emission, high combustion efficiency, high ash removal, simple design and compact size. The design of slagging combustors has operational flexibility for a wide range of applications, including retrofitting boilers, magnetohydrodynamic combustors, coal-fired gas turbines, gasifiers and hazardous waste incinerators. In recent years, developers of slagging combustors have achieved encouraging progress toward the commercialization of this technology. Although there is a diversity of technical approaches among the developers, they all aim for a compact design of pulverized coal combustion with high heat release and sub-stoichiometric combustion regimes of operation to suppress NOx formation, and most aim to capture sulfur by using sorbent injection in the combustor. If the present pace toward commercialization continues, retrofitting boilers of sizes ranging from 20 to 250 MMBtu/hr (5.9 to 73 MWt) may be available for commercial use in the 1990's. 18 refs., 2 figs.

  11. 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.

  12. 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.

  13. Combustor materials requirements and status of ceramic matrix composites

    NASA Technical Reports Server (NTRS)

    Hecht, Ralph J.; Johnson, Andrew M.

    1992-01-01

    The HSCT combustor will be required to operate with either extremely rich or lean fuel/air ratios to reduce NO(x) emission. NASA High Speed Research (HSR) sponsored programs at Pratt & Whitney (P&W) and GE Aircraft Engines (GEAE) have been studying rich and lean burn combustor design approaches which are capable of achieving the aggressive HSCT NO(x) emission goals. In both of the combustor design approaches under study, high temperature (2400-3000 F) materials are necessary to meet the HSCT emission goals of 3-8 gm/kg. Currently available materials will not meet the projected requirements for the HSCT combustor. The development of new materials is an enabling technology for the successful introduction to service of the HSCT.

  14. Catalytic reforming

    SciTech Connect

    Aldag, A.W. Jr.

    1986-01-28

    This patent describes a process for the catalytic reforming of a feedstock which contains at least one reformable organic compound. The process consists of contacting the feedstock under suitable reforming conditions with a catalyst composition selected from the group consisting of a catalyst. The catalyst essentially consists of zinc oxide and a spinel structure alumina. Another catalyst consists essentially of a physical mixture of zinc titanate and a spinel structure alumina in the presence of sufficient added hydrogen to substantially prevent the formation of coke. Insufficient zinc is present in the catalyst composition for the formation of a bulk zinc aluminate.

  15. Flame Driving of Longitudinal Instabilities in Liquid Fueled Dump Combustors

    DTIC Science & Technology

    1988-10-01

    instabilities These instabilities are characterized by either low frequency (i.e., rumble) or high frequency (i.e., screech ) pressure and velocity...in Lhe combusTor . In contrast, the high frequency screech occurs when one of the tangential acoustic modes of the combustor is excited (e.g., see...result in excessive vibrational loads on the system. On the other hand screech type instabilities result in an increase in heat transfer rates to

  16. Combustion Control in Industrial Multi-Swirl Stabilized Spray Combustor

    DTIC Science & Technology

    2005-08-21

    driven mechanism for thermo-acoustic combustion dynamics can be categorized into two groups according to Mongia et al. (2003). First category, which...CONTRACT NUMBER Combustion Control in Industrial Multi-Swirl Stabilized Spray Combustor 02PR12898-01 5b. GRANT NUMBER N00014-02-1-0756 5c. PROGRAM ELEMENT...ABSTRACT The focus of this study is to investigate the emission characteristics and combustion dynamics of multiple swirl spray combustors either in

  17. Variable volume combustor with an air bypass system

    DOEpatents

    Johnson, Thomas Edward; Ziminsky, Willy Steve; Ostebee, Heath Michael; Keener, Christopher Paul

    2017-02-07

    The present application provides a combustor for use with flow of fuel and a flow of air in a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles positioned within a liner and an air bypass system position about the liner. The air bypass system variably allows a bypass portion of the flow of air to bypass the micro-mixer fuel nozzles.

  18. Analysis of Flow Migration in an Ultra-Compact Combustor

    DTIC Science & Technology

    2011-03-01

    into the high- pressure tur- bine rotor while presenting a uniform temperature across the turbine blades. Several numerical parameter studies have been... pressure losses through the combustor section. As a result of these investigations a 0.75m diameter UCC combustor design has been developed along with a...hybrid turning vane which replaces the last compressor vane and high- pressure turbine vane. Furthermore, the issue of cooling the hybrid vane in the

  19. 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.

  20. 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.

  1. 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.

  2. Characterization of supersonic mixing in a nonreacting Mach 2 combustor

    SciTech Connect

    Hollo, S.D.; Mcdaniel, J.C.; Hartfield, R.J., JR. )

    1992-01-01

    Planar measurements of the injection mole fraction distribution and the velocity field within a nonreacting model SCRAMJET combustor have been made using laser-induced iodine fluorescence. The combustor geometry investigated in this work is staged transverse injection of air into a Mach 2 freestream. A complete three-dimensional survey of the injectant mole fraction distribution has been generated and a single planar velocity measurement has been completed. The measurements reveal the dramatic effect of streamwise vortices on the mixing of the injectant in the near field of the injectors, as well as the rapid mixing generated by staging two field injectors. Analysis of the downstream decay of the maximum injectant mole fraction in this and other nonreacting combustor geometries indicates that the relative rate of injectant mixing well downstream of the injectors is independent of combustor geometry, combustor Mach number, and injectant molecular weight. Mixing within this region of the combustor is dominated by turbulent diffusion within the injectant plume. The transition of the dominant mixing mechanism, from vortex-driven mixing in the near field to turbulent diffusion in the far field, was found to occur in the region between 10 and 20 jet diameters downstream of the injectors. 22 refs.

  3. 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.

  4. 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

  5. Comparison of perovskite and hexaaluminate-type catalysts for CO/H{sub 2}-fueled gas turbine combustors

    SciTech Connect

    Cristiani, C.; Groppi, G.; Forzatti, P.

    1996-12-31

    In this work the results of catalytic activity tests in CH{sub 4}, CO and H{sub 2} combustion over perovskite (LaCoO{sub 3}, LaMnO{sub 3} and LaFeO{sub 3}) and hexaaluminate-type (BaMnAl{sub 11}O{sub 19}, Sr{sub 0.8}La{sub 0.2}MnAl{sub 11}O{sub 19}, and BaFeAl{sub 11}O{sub 19}) systems are compared in order to investigate the potential of such materials as catalysts for syngas fueled combustors for gas turbines. Perovskites-type catalysts are shown to be the most active systems in the combustion of all the investigated fuels but to suffer from thermal stability problems that constrain their use in high temperature applications. Mn-substituted hexaaluminates have been shown to be more active by orders of magnitude in CO-H{sub 2} combustion than in CH{sub 4} combustion. Scale up of the activity data by mathematical modelling has demonstrated the potential of such catalysts in meeting the operating requirements of syngas fueled catalytic combustors.

  6. Catalytic reactor

    SciTech Connect

    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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. Effect of model selection on combustor performance and stability using ROCCID. [Rocket Combustor Interactive Design

    NASA Technical Reports Server (NTRS)

    Giuliani, James E.; Klem, Mark D.

    1992-01-01

    The ROCket Combustor Interactive Design (ROCCID) methodology is an interactive computer program that combines previously developed combustion analysis models to calculate the combustion performance and stability of liquid rocket engines. Test data from a 213 kN (48,000 lbf) Liquid Oxygen (LOX)/RP-1 combustor with a O-F-O (oxidizer-fuel-oxidizer) triplet injector were used to characterize the predictive capabilities of the ROCCID analysis models for this injector/propellant configuration. Thirteen combustion performance and stability models have been incorporated into ROCCID, and ten of them, which have options for triplet injectors, were examined in this study. Calculations using different combinations of analysis models, with little or no anchoring, were carried out on a test matrix of operating conditions matching those of the test program. Results of the computer analyses were compared to test data, and the ability of the model combinations to correctly predict combustion stability or instability was determined. For the best model combination(s), sensitivity of the calculations to fuel drop size and mixing efficiency was examined. Error in the stability calculations due to uncertainty in the pressure interaction index (N) was examined. The recommended model combinations for this O-F-O triplet LOX/RP-1 configuration are proposed.

  16. 40 CFR Table 3 to Subpart Cb of... - Municipal Waste Combustor Operating Guidelines

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ...-derived fuel mixed fuel-fired combustor 150 4 Spreader stoker coal/refuse-derived fuel mixed fuel-fired... conversion 250 c 24 Spreader stoker fixed floor refuse-derived fuel-fired combustor/100 percent coal...

  17. 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

  18. 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.

  19. Operating Characteristics of a Fluidic Premixed Dump Combustor

    NASA Astrophysics Data System (ADS)

    Ahmed, Kareem; Carr, Zakery; Forliti, David

    2007-11-01

    A transverse slot jet issuing into a channel flow has been shown to develop a large-scale recirculation zone. The current work involves both reacting and nonreacting flow studies of a fluidic dump combustor that utilizes a transverse slot jet in a planar channel flow. The motivation is to develop low thrust penalty flame holding methodologies that increase thrust and improve fuel economy. The reacting flow studies addressed the stabilization limits and combustion phenomena observed for the fluidic dump combustor. The fluidic stream consists of a mixture of methane fuel and air at an equivalence ratio matching that of the main combustor flow. A wall-mounted V-gutter was also studied to provide a comparison to a more traditional flame holder. The fluidic dump combustor has slightly degraded stabilization performance in terms of lean and rich blowout limits compared to the V-gutter. It also observed both stable and oscillatory combustion at different operating conditions. The combustion efficiency is higher for the fluidic dump combustor. The effect of the size of the slot jet was also explored.

  20. Mount assembly for porous transition panel at annular combustor outlet

    NASA Technical Reports Server (NTRS)

    Sweeney, Ralph B. (Inventor); Verdouw, Albert J. (Inventor)

    1980-01-01

    A gas turbine engine combustor assembly of annular configuration has outer and inner walls made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween and each outer and inner wall including a transition panel of porous metal defining a combustor assembly outlet supported by a combustor mount assembly including a stiffener ring having a side undercut thereon fit over a transition panel end face; and wherein an annular weld joins the ring to the end face to transmit exhaust heat from the end face to the stiffener ring for dissipation from the combustor; a combustor pilot member is located in axially spaced, surrounding relationship to the end face and connector means support the stiffener ring in free floating relationship with the pilot member to compensate for both radial and axial thermal expansion of the transition panel; and said connector means includes a radial gap for maintaining a controlled flow of coolant from outside of the transition panel into cooling relationship with the stiffener ring and said weld to further cool the end face against excessive heat build-up therein during flow of hot gas exhaust through said outlet.

  1. 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.

  2. 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.

  3. 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.

  4. Performance and Operability of a Dual Cavity Flame Holder in a Supersonic Combustor

    DTIC Science & Technology

    2009-06-01

    identified for each run by analyzing the pressure tap readings near the center of the combustor in the top side cavity. This tap was assumed the one with...last pressure tap in the combustor and dividing it by the lowest pressure found at the beginning of the isolator. The combustor exit pressure ratios...fitted with a quartz window in the combustor sidewall for one run night. This allowed flame emission images to be captured through digital and high speed

  5. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... combustor unit; (3) Procedures for receiving, handling, and feeding municipal solid waste; (4) Municipal... proper combustion air supply levels; (6) Procedures for operating the municipal waste combustor unit...) Procedures for handling ash; (10) Procedures for monitoring municipal waste combustor unit emissions; and...

  6. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... combustor unit; (3) Procedures for receiving, handling, and feeding municipal solid waste; (4) Municipal... proper combustion air supply levels; (6) Procedures for operating the municipal waste combustor unit...) Procedures for handling ash; (10) Procedures for monitoring municipal waste combustor unit emissions; and...

  7. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Standards of Performance for Large Municipal Waste Combustors for Which Construction is Commenced After... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... Standards for municipal waste combustor operator training and certification. (a) No later than the date...

  8. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... municipal waste combustor operating practices. (a) On and after the date on which the initial...

  9. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste combustor... municipal waste combustor operating practices. (a) On and after the date on which the initial...

  10. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 7 2012-07-01 2012-07-01 false Standards for municipal waste combustor... municipal waste combustor operating practices. (a) On and after the date on which the initial...

  11. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

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

  12. 40 CFR 60.55b - Standards for municipal waste combustor fugitive ash emissions.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste combustor... municipal waste combustor fugitive ash emissions. (a) On and after the date on which the initial...

  13. 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.

  14. 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.

  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.

    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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  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. 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.

  4. Variable volume combustor with aerodynamic fuel flanges for nozzle mounting

    DOEpatents

    McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward; Ostebee, Heath Michael

    2016-09-20

    The present application provides 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 number of support struts supporting the fuel nozzles and for providing the flow of fuel therethrough. The fuel injection system also may include a number of aerodynamic fuel flanges connecting the micro-mixer fuel nozzles and the support struts.

  5. 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.

  6. 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.

  7. Nondestructive evaluation of ceramic matrix composite combustor components.

    SciTech Connect

    Sun, J. G.; Verrilli, M. J.; Stephan, R.; Barnett, T. R.; Ojard, G.

    2002-11-08

    Combustor liners fabricated from a SiC/SiC composite were nondestructively interrogated before and after combustion rig testing. The combustor liners were inspected 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. Microstructural examination of the SiC/SiC liners revealed the thermography indications to be delaminations and damaged fiber tows.

  8. 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.

  9. 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.

  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. 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.

  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. 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.

  14. 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.

  15. 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.

  16. DEMONSTRATION BULLETIN: CIRCULATING BED COMBUSTOR - OGDEN ENVIRONMENTAL SERVICES, INC.

    EPA Science Inventory

    An evaluation of the Ogden Environmental Services (OES) circulating bed combustor (CBC) technology was carried out under the superfund Innovative Technology Evaluation (SITE) Program to determine its applicabilitY as an on-site treatment method for waste site cleanups, and more s...

  17. 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.

  18. 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 dynamcis 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.

  19. 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.

  20. Assessment, development, and application of combustor aerothermal models

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    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.

  1. Pylon Fuel Injector Design for a Scramjet Combustor (Postprint)

    DTIC Science & Technology

    2007-07-01

    Combustor by Application of Laser Diagnostics,” AIAA Paper 2002-5203, Oct. 2002. 22Sunami, T., Magre, P., Bresson , A., Grisch, F., Orain, M., and Kodera, M...Asymmetric Nozzles,” AIAA Paper 96-0200, Jan. 1996. 25Haimovitch, Y., Gartenberg, E., Roberts , A., and Northam, G., “Effects of Internal Nozzle

  2. 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.

  3. 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...

  4. Improved Controllers For Heaters In Toxic-Gas Combustors

    NASA Technical Reports Server (NTRS)

    Wishard, James; Lamb, James; Fortier, Edward; Velasquez, Hugo; Waltman, Doug

    1995-01-01

    Commercial electronic proportional controllers installed in place of mechanical power controllers for electric heaters in toxic-gas combustors at NASA's Jet Propulsion Laboratory. Designed to maintain temperature of heater at preset value by turning power fully on or fully off when temperature falls below or rises above that value, respectively. Solid-state power controllers overcome deficiencies of mechanical power controllers.

  5. Coanda injection system for axially staged low emission combustors

    DOEpatents

    Evulet, Andrei Tristan [Clifton Park, NY; Varatharajan, Balachandar [Cincinnati, OH; Kraemer, Gilbert Otto [Greer, SC; ElKady, Ahmed Mostafa [Niskayuna, NY; Lacy, Benjamin Paul [Greer, SC

    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.

  6. MHD coal combustor technology. Final report, phase II

    SciTech Connect

    Not Available

    1980-09-01

    The design, performance, and testing of a 20-MW coal combustor for scaleup to 50 MW for use in an MHD generator are described. The design incorporates the following key features: (1) a two-stage combustor with an intermediate slag separator to remove slag at a low temperture, thus minimizing enthalpy losses required for heating and vaporizing the slag; (2) a first-stage pentad (four air streams impinging on one coal stream) injector design with demonstrated efficient mixing, promoting high carbon burnout; (3) a two-section first-stage combustion chamber; the first stage using a thin slag-protected refractory layer and the second section using a thick refractory layer, both to minimize heat losses; (4) a refractory lining in the slag separator to minimize heat losses; (5) a second-stage combustor, which provided both de-swirl of the combustion products exiting from the slag separator and simple mixing of the vitiated secondary air and seed; (6) a dense-phase coal feed system to minimize cold carrier gas entering the first-stage combustors; (7) a dry seed injection system using pulverized K/sub 2/CO/sub 3/ with a 1% amorphous, fumed silicon dioxide additive to enhance flowability, resulting in rapid vaporization and ionization and ensuring maximum performance; and (8) a performance evaluation module (PEM) of rugged design based on an existing, successfully-fired unit. (WHK)

  7. 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...

  8. 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...

  9. 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.

  10. 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...

  11. 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.

  12. 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

  13. An Experimental Study of a Catalytic Combustor for an Expendable Turbojet Engine

    DTIC Science & Technology

    1978-03-01

    Air mass flow was measured by static taps on either side of the 1.8 in. diameter orifice connected to a mercury manometer . two inch gate valve...flow was determined on the external air supply test stand by measuring the differential pressure across a 1.8 in. diameter orifice with a mercury ... manometer . On ’-he engine, three static pressure taps located on the bellmouth inlet measured inlet vacuum with a pressure transducer. Calibration curves

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Performance of an annual combustor designed for a low-cost turbojet engine

    NASA Technical Reports Server (NTRS)

    Fear, J. S.

    1973-01-01

    Performance tests were conducted on a combustor designed for use in a low-cost turbojet engine. Low-cost features included the use of very inexpensive simplex fuel nozzles and combustor liners of perforated sheet material. Combustion efficiencies at the altitude-cruise and sea-level design points were approximately 94 and 96 percent, respectively. The combustor isothermal total-pressure loss was 8.8 percent at the altitude-cruise-condition diffuser-inlet Mach number of 0.335. The combustor-exit temperature pattern factor was less than 0.3 at the altitude-cruise, sea-level-cruise, and sea-level-static design conditions. The combustor-exit average radial temperature profiles at all conditions were in very good agreement with the design profile. The intense mixing required because of the very high combustor heat-release rate had an adverse effect on ignition capability at altitude windmilling design conditions.

  3. 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)

  4. Driving of combustor oscillations by gaseous propellant injectors

    NASA Technical Reports Server (NTRS)

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.

    1979-01-01

    Measurements of reactive admittances that describe the capabilities of the combustion processes associated with coaxial gaseous fuel injectors to amplify combustor oscillations are presented. These admittances are needed in (1) stability analyses of rocket motors and gas turbine combustors; (2) for the evaluation of 'driving' capabilities of injectors; and (3) for checking the applicability of a theoretical model. The modified standing-wave technique was used to determine the admittances of the combustion processes in coaxial injectors utilizing air-acetylene mixtures. The measurements indicated that (1) coaxial injectors can sustain combustion instabilities over wide frequency ranges; (2) the maximum driving capability of an injector decreases with an increased equivalence ratio; (3) the frequency at which maximum driving is observed decreases with the increased equivalence ratio; and (4) the characteristic combustion time of an injector design decreases with increased frequency.

  5. 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.

  6. 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).

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Reduction of nitric oxide emissions from a combustor

    SciTech Connect

    Craig, R.A.; Pritchard, H.O.

    1980-05-27

    A turbojet combustor and method for controlling nitric oxide emissions is provided by employing successive combustion zones wherein 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, which usually result in large emissions of nitric oxide in a primary combustion zone, 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 emissions of nitric oxide.

  12. 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.

  13. Computational investigation on combustion instabilities in a rocket combustor

    NASA Astrophysics Data System (ADS)

    Yuan, Lei; Shen, Chibing

    2016-10-01

    High frequency combustion instability is viewed as the most challenging task in the development of Liquid Rocket Engines. In this article, results of attempts to capture the self-excited high frequency combustion instability in a rocket combustor are shown. The presence of combustion instability was demonstrated using point measurements, along with Fast Fourier Transform analysis and instantaneous flowfield contours. A baseline case demonstrates a similar wall heat flux profile as the associated experimental case. The acoustic oscillation modes and corresponding frequencies predicted by current simulations are almost the same as the results obtained from classic acoustic analysis. Pressure wave moving back and forth across the combustor was also observed. Then this baseline case was compared against different fuel-oxidizer velocity ratios. It predicts a general trend: the smaller velocity ratio produces larger oscillation amplitudes than the larger one. A possible explanation for the trend was given using the computational results.

  14. 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.

  15. Investigation of a small solid fuel ramjet combustor

    SciTech Connect

    Levy, Yeshayahou; Gany, Alon; Zvuloni, Roni

    1989-06-01

    Experimental and analytical investigations of a small solid-fuel ramjet (SFRJ) combustor were conducted. A static test system with a 25-kW electrical air heater simulated the air temperature and pressure encountered in flight at a Mach number of 3 at sea level. The transparent polymethylmethacrylate fuel used in the tests permitted continuous video photography, revealing the local fuel-regression-rate behavior and the instantaneous ignition and combustion phenomena. The results demonstrated high combustion efficiency and indicated peculiar local and average fuel-regression-rate correlations. The analysis indicated that the specific conditions resulting from the low Reynolds number range in small SFRJ motors, in contrast to large combustors, enhance the effect of the sudden-expansion heat-transfer regime relative to the boundary-layer regime. 14 refs.

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

    DOEpatents

    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.

  17. 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.

  18. 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.

  19. Systems and methods for preventing flashback in a combustor assembly

    SciTech Connect

    Johnson, Thomas Edward; Ziminsky, Willy Steve; Stevenson, Christian Xavier

    2016-04-05

    Embodiments of the present application include a combustor assembly. The combustor assembly may include a combustion chamber, a first plenum, a second plenum, and one or more elongate air/fuel premixing injection tubes. Each of the elongate air/fuel premixing injection tubes may include a first length at least partially disposed within the first plenum and configured to receive a first fluid from the first plenum. Moreover, each of the elongate air/fuel premixing injection tubes may include a second length disposed downstream of the first length and at least partially disposed within the second plenum. The second length may be formed of a porous wall configured to allow a second fluid from the second plenum to enter the second length and create a boundary layer about the porous wall.

  20. N+2 Advanced Low NOx Combustor Technology Final Report

    NASA Technical Reports Server (NTRS)

    Herbon, John; Aicholtz, John; Hsieh, Shih-Yang; Viars, Philip; Birmaher, Shai; Brown, Dan; Patel, Nayan; Carper, Doug; Cooper, Clay; Fitzgerald, Russell

    2017-01-01

    In accordance with NASAs technology goals for future subsonic vehicles, this contract identified and developed new combustor concepts toward meeting N+2 generation (2020) LTO (landing and take-off) NOx emissions reduction goal of 75 from the standard adopted at Committee on Aviation Environmental Protection 6 (CAEP6). Based on flame tube emissions, operability, and autoignition testing, one concept was down selected for sector testing at NASA. The N+2 combustor sector successfully demonstrated 75 reduction for LTO NOx (vs. CAEP6) and cruise NOx (vs. 2005 B777-200 reference) while maintaining 99.9 cruise efficiency and no increase in CO and HC emissions.The program also developed enabling technologies for the combustion system including ceramic matrix composites (CMC) liner materials, active combustion control concepts, and laser ignition for improved altitude relight.

  1. 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.

  2. 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.

  3. Numerical Investigation of a Model Scramjet Combustor Using DDES

    NASA Astrophysics Data System (ADS)

    Shin, Junsu; Sung, Hong-Gye

    2017-04-01

    Non-reactive flows moving through a model scramjet were investigated using a delayed detached eddy simulation (DDES), which is a hybrid scheme combining Reynolds averaged Navier-Stokes scheme and a large eddy simulation. The three dimensional Navier-Stokes equations were solved numerically on a structural grid using finite volume methods. An in-house was developed. This code used a monotonic upstream-centered scheme for conservation laws (MUSCL) with an advection upstream splitting method by pressure weight function (AUSMPW+) for space. In addition, a 4th order Runge-Kutta scheme was used with preconditioning for time integration. The geometries and boundary conditions of a scramjet combustor operated by DLR, a German aerospace center, were considered. The profiles of the lower wall pressure and axial velocity obtained from a time-averaged solution were compared with experimental results. Also, the mixing efficiency and total pressure recovery factor were provided in order to inspect the performance of the combustor.

  4. Heat Transfer Experiments on a Pulse Detonation Driven Combustor

    DTIC Science & Technology

    2011-03-01

    steps that need to take place before such a hybrid is successfully developed. PDEs obtain their increased efficiency by means of detonation , a pressure...combustion in the Brayton cycle. A PDE utilizes detonations , which offer much higher pressures at the site of fuel ignition, generating less...HEAT TRANSFER EXPERIMENTS ON A PULSE DETONATION DRIVEN COMBUSTOR THESIS Nicholas C. Longo, Captain, USAF AFIT/GAE/ENY/11-M18

  5. 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.

  6. System for tuning a combustor of a gas turbine

    SciTech Connect

    Hughes, Michael John

    2016-12-27

    A system for tuning a combustor of a gas turbine includes a flow sleeve having an annular main body. The main body includes an upstream end, a downstream end, an inner surface and an outer surface. A cooling channel extends along the inner surface of the main body. The cooling channel extends at least partially between the downstream end and the upstream end of the main body.

  7. 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.

  8. Quiet Clean Short-haul Experimental Engine (QCSEE) clean combustor test report

    NASA Technical Reports Server (NTRS)

    1975-01-01

    A component pressure test was conducted on a F101 PFRT combustor to evaluate the emissions levels of this combustor design at selected under the wing and over the wing operating conditions for the quiet clean short haul experimental engine (QCSEE). Emissions reduction techniques were evaluated which included compressor discharge bleed and sector burning in the combustor. The results of this test were utilized to compare the expected QCSEE emissions levels with the emission goals of the QCSEE engine program.

  9. Progress in Applying Tunable Diode Laser Absorption Spectroscopy to Scramjet Isolators and Combustors

    DTIC Science & Technology

    2010-05-01

    AFRL-RZ-WP-TP-2010-2146 PROGRESS IN APPLYING TUNABLE DIODE LASER ABSORPTION SPECTROSCOPY TO SCRAMJET ISOLATORS AND COMBUSTORS Michael S... COMBUSTORS 5a. CONTRACT NUMBER IN HOUSE 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 61102F 6. AUTHOR(S) Michael S. Brown, Skip Williams...paths -- scramjet engines in particular. In this report we summarize our progress in applying TDLAS to scramjet isolators and combustors . 15

  10. Diode Laser-Based Detection of Combustor Instabilities with Application to a Scramjet Engine

    DTIC Science & Technology

    2010-02-01

    combustion instability has focused primarily on subsonic turbulent combustors [7]. NOx emission regulations on gas turbines have driven the use of...Optical Access Laser Beam Location ermocouple eholder) Static Pressure Taps (Open circles)holders Combustor w, direct-connect scramjet flowpath. 4.8 s...AFRL-RZ-WP-TP-2010-2055 DIODE LASER-BASED DETECTION OF COMBUSTOR INSTABILITIES WITH APPLICATION TO A SCRAMJET ENGINE (POSTPRINT) Campbell

  11. Ignition and Flameholding in a Supersonic Combustor by an Electrical Discharge Combined with a Fuel Injector

    DTIC Science & Technology

    2014-01-01

    1 Ignition and Flameholding in a Supersonic Combustor by an Electrical Discharge Combined with a Fuel Injector K. V. Savelkin 1 , D. A...presents the results of an experimental study of supersonic combustor operation enhanced by an electrical discharge. A novel scheme of plasma assisted...experimental combustor with the cross section of 72 mm (width)  60 mm (height) and length of 600 mm operates at a Mach number of M=2, initial stagnation

  12. Performance of a short combustor at high altitudes using hydrogen fuel

    NASA Technical Reports Server (NTRS)

    Sivo, Joseph N; Fenn, David B

    1956-01-01

    Performance characteristics of a 16-inch annular-type combustor installed in a full-scale engine using gaseous-hydrogen fuel were obtained at simulated altitudes from 66,000 to 86,000 feet at a flight Mach number of 0.8. Combustion efficiencies of 86 percent were obtained at 86,000 feet (combustor pressure, 420 lb/sq ft abs). Combustor blowout was not encountered during the investigation.

  13. 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.

  14. 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.

  15. 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).

  16. 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

  17. 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.

  18. 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.

  19. 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.

  20. Massively parallel computation of three-dimensional scramjet combustor

    NASA Astrophysics Data System (ADS)

    Zheng, Z. H.; Le, J. L.

    Recent progress of computational study of scramjet combustor has been described in Refs 1-3. However, detailed flow properties, especially the lateral properties and the sidewall effects are not considered. In this paper, a parallel simulation of an experimental dual-mode scramjet combustor configuration is presented, considering the jet-to-jet symmetry and the full-duct modeling. Turbulence is modeled with the k-ɛ two-equation turbulence model and a 7-specie, 8-equation kinetics model is used to model hydrogen/air combustion. The conservation form of the Navier-Stokes equations with finite-rate chemistry reactions is solved using a diagonal implicit finite-volume method. For the two cases, the three-dimension flow-fields with equivalence ratio Φ=0.0 and 0.35 have been respectively simulated on the COW and MPP. Wall pressure comparisons between CFD and experiments (CARDC and NAL) show fair agreement for the jet-to-jet case. For the full-duct modeling, more detailed flow properties are obtained. The fuelpenetrating heights of the injectors are different because of the effects of the sidewall boundary layer and the shock wave in the combustor. According to numerical results, if adjusting the locations of the injectors, the combustion efficiency could be improved.

  1. Combustor cap having non-round outlets for mixing tubes

    SciTech Connect

    Hughes, Michael John; Boardman, Gregory Allen; McConnaughhay, Johnie Franklin; Arguinzoni, Carlo Antonio

    2016-12-27

    A system includes a a combustor cap configured to be coupled to a plurality of mixing tubes of a multi-tube fuel nozzle, wherein each mixing tube of the plurality of mixing tubes is configured to mix air and fuel to form an air-fuel mixture. The combustor cap includes multiple nozzles integrated within the combustor cap. Each nozzle of the multiple nozzles is coupled to a respective mixing tube of the multiple mixing tubes. In addition, each nozzle of the multiple nozzles includes a first end and a second end. The first end is coupled to the respective mixing tube of the multiple mixing tubes. The second end defines a non-round outlet for the air-fuel mixture. Each nozzle of the multiple nozzles includes an inner surface having first and second portions, the first portion radially diverges along an axial direction from the first end to the second end, and the second portion radially converges along the axial direction from the first end to the second end.

  2. Numerical investigation of a hydrogen-fueled scramjet combustor at flight conditions

    NASA Astrophysics Data System (ADS)

    Rabadan, E.; Weigand, B.

    2013-03-01

    Numerical investigations of a hydrogen-fueled scramjet combustor at cruise flight conditions of Mach 8 at an altitude of 30 km have been performed. Two combustor configurations were investigated: a single-stage combustor with a central strut injector and a two-staged combustor combining the central strut and wall-mounted ramp injectors. These numerical simulations are aimed to study the flow structure, supersonic mixing, autoignition, and combustion for the present combustor configurations. A turbulent flow from a separate intake calculation was used as inflow condition for the combustor. A better performance for the two-staged combustor configuration was observed. The combination of the central strut injection together with the wall-ramp injection improved the turbulent mixing and, consequently, the combustion process. Inside the supersonic combustion chamber, the autoignition zone occurred downstream of the injectors, and combustion takes place accompanied by high heat release and pressure rise. As the equivalence ratio was increased, the combustion became stronger causing an upstream displacement of the shock train producing different pressure variations. For the two-staged combustor configuration, the location of the autoignition zone was found to appear further upstream compared to single-stage combustor. Mixing was improved by addition of the second-stage injection. The influence of the wall temperature was also investigated showing an effect on the combustion pressure rise and the length and location of the shock train.

  3. 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.

  4. Preliminary investigation of the performance of a single tubular combustor at pressure up to 12 atmospheres

    NASA Technical Reports Server (NTRS)

    Wear, Jerrold D; Butze, Helmut F

    1954-01-01

    The effects of combustor operation at conditions representative of those encountered in high pressure-ratio turbojet engines or at high flight speeds on carbon deposition, exhaust smoke, and combustion efficiency were studied in a single tubular combustor. Carbon deposition and smoke formation tests were conducted over a range of combustor-inlet pressures from 33 to 173 pounds per square inch absolute and combustor reference velocities from 78 to 143 feet per second. Combustion efficiency tests were conducted over a range of pressures from 58 to 117 pounds per square inch absolute and velocities from 89 to 172 feet per second.

  5. Advanced low NO/x/ combustors for supersonic high-altitude aircraft gas turbines

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    A test rig program was conducted with the objective of evaluating and minimizing the exhaust emissions, in particular NO(x), of two advanced aircraft combustor concepts at a simulated, high-altitude cruise condition. The two combustor designs, both members of the lean-reaction, pre-mixed family, 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 0.127-m 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(x) level of 1.1 gm NO2/kg fuel with essentially 100 percent combustion efficiency at the simulated cruise combustor condition of 507 kPa, 833 K inlet pressure and temperature, respectively and 1778 K outlet temperature on Jet-A1 fuel. In addition, emissions data were obtained at low combustor inlet pressure and temperatures that indicate the potential performance at engine off-design conditions.

  6. 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.

  7. 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.

  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. 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.

  10. 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)

  11. 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

  12. Flame structure and stabilization in miniature liquid film combustors

    NASA Astrophysics Data System (ADS)

    Pham, Trinh Kim

    Liquid-fueled miniature combustion systems can be promising portable power devices when high specific power and long operation duration are required. A uniquely viable fueling option for small scale combustion is to introduce the liquid fuel as a film on the combustor walls. As one example of such systems, this dissertation characterizes 1-cm-diameter tubular combustors fed by liquid fuel films, and seeks to identify the mechanisms by which flames are stabilized within them. Early experimental work demonstrates that flame behavior is dependent upon steadiness in fuel and air injection and in geometric symmetry and uniformity. Significant discoveries in later work include the impact of direct strain on the flame by the airflow, the fact that no local recirculation zone appears to exist for stabilization as was previously believed, and that the film thickness, uniformity, and location directly affect the flame's characteristics and stability. A gradient in film thickness is required for stable operation, and this requirement may explain why the combustor maintains overall rich conditions. Initial numerical simulations of two-dimensional cold and reacting flows in a simplified model of the combustor yields flame shape and flow field results that do not match experiments in the burning case, therefore suggesting that local turbulence in the fuel injection region provides the necessary degree of mixing. A three-dimensional model of the combustor is needed if reacting flows are to be simulated accurately. It was also found that thermal conduction from the chamber exit to the chamber base plays an important role in fuel vaporization and the stability of the flame. Consequently, flames cannot be sustained in quartz and other transparent but thermally insulating materials for the selected geometry, so observation of the flame's entire structure cannot be accomplished without either the addition of other flameholding elements or the employment of a more thermally conductive

  13. Large eddy simulation of soot evolution in an aircraft combustor

    NASA Astrophysics Data System (ADS)

    Mueller, Michael E.; Pitsch, Heinz

    2013-11-01

    An integrated kinetics-based Large Eddy Simulation (LES) approach for soot evolution in turbulent reacting flows is applied to the simulation of a Pratt & Whitney aircraft gas turbine combustor, and the results are analyzed to provide insights into the complex interactions of the hydrodynamics, mixing, chemistry, and soot. The integrated approach includes detailed models for soot, combustion, and the unresolved interactions between soot, chemistry, and turbulence. The soot model is based on the Hybrid Method of Moments and detailed descriptions of soot aggregates and the various physical and chemical processes governing their evolution. The detailed kinetics of jet fuel oxidation and soot precursor formation is described with the Radiation Flamelet/Progress Variable model, which has been modified to account for the removal of soot precursors from the gas-phase. The unclosed filtered quantities in the soot and combustion models, such as source terms, are closed with a novel presumed subfilter PDF approach that accounts for the high subfilter spatial intermittency of soot. For the combustor simulation, the integrated approach is combined with a Lagrangian parcel method for the liquid spray and state-of-the-art unstructured LES technology for complex geometries. Two overall fuel-to-air ratios are simulated to evaluate the ability of the model to make not only absolute predictions but also quantitative predictions of trends. The Pratt & Whitney combustor is a Rich-Quench-Lean combustor in which combustion first occurs in a fuel-rich primary zone characterized by a large recirculation zone. Dilution air is then added downstream of the recirculation zone, and combustion continues in a fuel-lean secondary zone. The simulations show that large quantities of soot are formed in the fuel-rich recirculation zone, and, furthermore, the overall fuel-to-air ratio dictates both the dominant soot growth process and the location of maximum soot volume fraction. At the higher fuel

  14. 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

  15. Planar Laser-Induced Fluorescence Imaging of OH in a Supersonic Combustor Fueled with Ethylene and Methane (Postprint)

    DTIC Science & Technology

    2010-02-01

    the facility is instrumented with multiple pressure taps and thermocouples for evaluating combustor performance. Combustor run times are typically 30...AFRL-RZ-WP-TP-2010-2056 PLANAR LASER-INDUCED FLUORESCENCE IMAGING OF OH IN A SUPERSONIC COMBUSTOR FUELED WITH ETHYLENE AND METHANE...FLUORESCENCE IMAGING OF OH IN A SUPERSONIC COMBUSTOR FUELED WITH ETHYLENE AND METHANE (POSTPRINT) 5a. CONTRACT NUMBER In-house 5b. GRANT NUMBER 5c

  16. 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.

  17. 40 CFR 62.14104 - Requirements for municipal waste combustor operating practices.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... accordance with 5 U.S.C. 552(a) and 1 CFR part 51. You may obtain a copy from the American Society of... Before September 20, 1994 § 62.14104 Requirements for municipal waste combustor operating practices. (a... with the municipal waste combustor operating practice requirements listed in 40 CFR 60.53b(b) and...

  18. 40 CFR 62.14104 - Requirements for municipal waste combustor operating practices.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... accordance with 5 U.S.C. 552(a) and 1 CFR part 51. You may obtain a copy from the American Society of... Before September 20, 1994 § 62.14104 Requirements for municipal waste combustor operating practices. (a... with the municipal waste combustor operating practice requirements listed in 40 CFR 60.53b(b) and...

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

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 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 and... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standard for municipal waste...

  20. 40 CFR 60.56a - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 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. (a... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste...

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

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 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 and... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standard for municipal waste...

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

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 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. (a... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste...

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

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 8 2011-07-01 2011-07-01 false Municipal Waste Combustor Operating... POLLUTANTS Federal Plan Requirements for Large Municipal Waste Combustors Constructed on or Before September 20, 1994 Pt. 62, Subpt. FFF, Table 3 Table 3 to Subpart FFF of Part 62—Municipal Waste...

  4. 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...

  5. 40 CFR Table 3 to Subpart Fff of... - Municipal Waste Combustor Operating Requirements

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 9 2013-07-01 2013-07-01 false Municipal Waste Combustor Operating... POLLUTANTS Federal Plan Requirements for Large Municipal Waste Combustors Constructed on or Before September 20, 1994 Pt. 62, Subpt. FFF, Table 3 Table 3 to Subpart FFF of Part 62—Municipal Waste...

  6. 40 CFR 60.53a - Standard for municipal waste combustor organics.

    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 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...

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

    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 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...

  8. 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...

  9. 40 CFR 60.53a - Standard for municipal waste combustor organics.

    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 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...

  10. 40 CFR 60.53a - Standard for municipal waste combustor organics.

    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 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...

  11. Combustion stability with baffles, absorbers and velocity sensitive combustion. [liquid propellant rocket combustors

    NASA Technical Reports Server (NTRS)

    Mitchell, C. E.

    1980-01-01

    Analytical and computational techniques were developed to predict the stability behavior of liquid propellant rocket combustors using damping devices such as acoustic liners, slot absorbers, and injector face baffles. Models were developed to determine the frequency and decay rate of combustor oscillations, the spatial and temporal pressure waveforms, and the stability limits in terms of combustion response model parameters.

  12. 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.

  13. 40 CFR 60.52a - Standard for municipal waste combustor metals.

    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 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...

  14. 40 CFR 60.52a - Standard for municipal waste combustor metals.

    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 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...

  15. 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...

  16. 40 CFR 60.52a - Standard for municipal waste combustor metals.

    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 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...

  17. 40 CFR 60.52a - Standard for municipal waste combustor metals.

    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 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...

  18. 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.

  19. A conceptual design of shock-eliminating clover combustor for large scale scramjet engine

    NASA Astrophysics Data System (ADS)

    Sun, Ming-bo; Zhao, Yu-xin; Zhao, Guo-yan; Liu, Yuan

    2017-01-01

    A new concept of shock-eliminating clover combustor is proposed for large scale scramjet engine to fulfill the requirements of fuel penetration, total pressure recovery and cooling. To generate the circular-to-clover transition shape of the combustor, the streamline tracing technique is used based on an axisymmetric expansion parent flowfield calculated using the method of characteristics. The combustor is examined using inviscid and viscous numerical simulations and a pure circular shape is calculated for comparison. The results showed that the combustor avoids the shock wave generation and produces low total pressure losses in a wide range of flight condition with various Mach number. The flameholding device for this combustor is briefly discussed.

  20. 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.

  1. 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.

  2. 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%.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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

  11. Fundamental Aspects of the Aerodynamics of Turbojet Engine Combustors

    NASA Technical Reports Server (NTRS)

    Barrere, M.

    1978-01-01

    Aerodynamic considerations in the design of high performance combustors for turbojet engines are discussed. Aerodynamic problems concerning the preparation of the fuel-air mixture, the recirculation zone where primary combustion occurs, the secondary combustion zone, and the dilution zone were examined. An aerodynamic analysis of the entire primary chamber ensemble was carried out to determine the pressure drop between entry and exit. The aerodynamics of afterburn chambers are discussed. A model which can be used to investigate the evolution of temperature, pressure, and rate and efficiency of combustion the length of the chamber was developed.

  12. 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.

  13. 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.

  14. An Investigation of Flame Stability in a Coaxial Combustor

    DTIC Science & Technology

    1979-01-01

    appreciation to Dr, Harold E. Wright, Chairman of iry advisory conuittee, for his initial suggestion that I should enter the Doctoral program at AFIT...OF FLAME STABILITY IN A COAXIAL DUMP COMBUSTOR ZE by Edward T. Curran, Ph.D. Dr. Harold E. Wright, Advisor S 4 An experimental investigation of the...lenoth of the RZ zone and Urelas TU r - rl Urel is the relative flow velocity. Thus UreI = Um + U where Um is the downstream velocity of the flow

  15. Device for improved air and fuel distribution to a combustor

    SciTech Connect

    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).

  16. An efficient computational tool for ramjet combustor research

    SciTech Connect

    Vanka, S.P.; Krazinski, J.L.; Nejad, A.S.

    1988-01-01

    A multigrid based calculation procedure is presented for the efficient solution of the time-averaged equations of a turbulent elliptic reacting flow. The equations are solved on a non-orthogonal curvilinear coordinate system. The physical models currently incorporated are a two equation k-epsilon turbulence model, a four-step chemical kinetics mechanism, and a Lagrangian particle tracking procedure applicable for dilute sprays. Demonstration calculations are presented to illustrate the performance of the calculation procedure for a ramjet dump combustor configuration. 21 refs., 9 figs., 2 tabs.

  17. Particle Sizing in a Fuel-Rich Ramjet Combustor.

    DTIC Science & Technology

    1983-08-01

    COVERED Particle Sizing in a Fuel-Rich Ramjet Combustor Technical Memorandum 6 PERFORMING ORG. REPORT NUMBER 7. AIJTHORII CONTRACT OR GRANT NUMBER~s...R. Turner and R. A. Murphy N00024-83-C-S3Ol 9. PERFORMING ORGANIZATION NAME & ADDRESS 10. PROGRAM ELEMENT. PROJECT. TASK The Johns Hopkins University... Analyi , of t)op- pier Signal Characteristics for a Cross-tean I aser Doppler Ve- locimcier." 4ppI. Opt.. 14. 2177 (1975). In the present configuration

  18. 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

  19. 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.

  20. 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.

  1. Numerical Simulation of a Small-Scale Mild Combustor

    NASA Astrophysics Data System (ADS)

    Veríssimo, A.; Oliveira, R.; Coelho, P. J.; Costa, M.

    2012-11-01

    This work reports numerical simulations of a small-scale cylindrical combustor operating in the mild combustion regime. Preheated air is supplied by a central nozzle, while the fuel (methane) is injected through 16 holes placed equidistantly in a circumference concentric with the air nozzle. The calculations were carried out using the commercial code Ansys-Fluent. Turbulence was modelled using the realizable k-epsilon model. Two different combustion models were employed, namely the eddy dissipation concept and the joint composition pdf transport model. In both cases, a chemical mechanism comprising 13 transported species and 73 chemical reactions was used, as well as a global single-step reaction. A thorough comparison of the predictions obtained using the pdf transport model and the eddy dissipation concept with detailed experimental data is presented. Both models are able to accurately predict the temperature and the O2 and CO2 molar fractions over most of the combustor, but the temperature field is overestimated in the vicinity of the burner. Discrepancies are found in the prediction of the CO molar fraction, particularly when the eddy dissipation concept is used.

  2. Performance of a second-generation PFB pilot plant combustor

    SciTech Connect

    Bonk, D.L.; Conn, R.; Van Hook, J.; Robertson, A.

    1995-04-01

    Second-generation on pressurized fluidized bed combustion (PFBC) plants promise higher efficiency with lower costs of electricity and lower stack emissions. With a conventional reheat cycle and a 3-percent sulfur Pittsburgh No. 8 coal, a 45-percent efficiency (HHV of coal basis) and a cost of electricity {approximately}20 percent 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.

  3. 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.

  4. 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.

  5. 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 mass flow 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

    NASA Astrophysics Data System (ADS)

    Brown, William S.

    1991-12-01

    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.

  13. 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.

  14. 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.

  15. 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.

  16. Sensing and dynamics of lean blowout in a swirl dump combustor

    NASA Astrophysics Data System (ADS)

    Thiruchengode, Muruganandam

    This thesis describes an investigation on the blowout phenomenon in gas turbine combustors. The combustor primarily used for this study was a swirl- and dump-stabilized, atmospheric pressure device, which did not exhibit dynamic combustion instabilities. The first part of the thesis work concentrated on finding a sensing methodology to be able to predict the onset of approach of combustor blowout using optical methods. Temporary extinction-reignition events that occurred prior to blowout were found to be precursor events to blowout. A threshold based method was developed to identify these events in the time-resolved sensor output. The number and the average length of each event were found to increase as the LBO limit (fuel-air ratio) is approached. This behavior is used to predict the proximity to lean blowout. In the second part of this study, the blowout sensor was incorporated into a control system that monitored the approach of blowout and then actuated an alternate mechanism to stabilize the combustor near blowout. Enhanced stabilization was achieved by redirecting a part of the main fuel to a central preinjection pilot injection. The sensing methodology, without modification, was effective for the combustor with pilot stabilization. An event based control algorithm for controlling the combustor from blowing out was also developed in this study. The control system was proven to stabilize the combustor even when the combustor loading was rapidly changed. The final part of this study focused on understanding the physical mechanisms behind the precursor events. High speed movies of flame chemiluminescence and laser sheet scattering from oil droplets seeded into the reactants were analyzed to explain the physical processes that cause the extinction and the reignition of the combustor during a precursor event. A physical model for coupling of the fluid dynamics of vortex breakdown and combustion during precursor and blowout events is proposed. This model of blowout

  17. Three-dimensional predictions of reactive turbulent recirculating flow of a cylindrical MHD type combustor

    NASA Astrophysics Data System (ADS)

    Lee, J. J.

    1990-01-01

    A computational procedure is employed to predict the axisymmetric reactive turbulent recirculating flow-field within a cylindrical MHD combustor. The procedure used in the analysis is an extended version of the three-dimensional Combustor Performance Program developed at the Garrett Turbine Engine Company. The separated flow created by four liquid fuel nozzles and 148 oxidizer holes passing through an injector plate is examined. Numerical results for flow, heat/mass transfer and combustion are presented to describe these complex three-dimensional interactions. The detailed analysis achieved by the numerical model is useful for evaluating combustor performance and in the interpretation of laboratory test data.

  18. 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.

  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. Engine-Scale Combustor Rig Designed, Fabricated, and Tested for Combustion Instability Control Research

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Breisacher, Kevin J.

    2000-01-01

    Low-emission combustor designs are prone to combustor instabilities. Because active control of these instabilities may allow future combustors to meet both stringent emissions and performance requirements, an experimental combustor rig was developed for investigating methods of actively suppressing combustion instabilities. The experimental rig has features similar to a real engine combustor and exhibits instabilities representative of those in aircraft gas turbine engines. Experimental testing in the spring of 1999 demonstrated that the rig can be tuned to closely represent an instability observed in engine tests. Future plans are to develop and demonstrate combustion instability control using this experimental combustor rig. The NASA Glenn Research Center at Lewis Field is leading the Combustion Instability Control program to investigate methods for actively suppressing combustion instabilities. Under this program, a single-nozzle, liquid-fueled research combustor rig was designed, fabricated, and tested. The rig has many of the complexities of a real engine combustor, including an actual fuel nozzle and swirler, dilution cooling, and an effusion-cooled liner. Prior to designing the experimental rig, a survey of aircraft engine combustion instability experience identified an instability observed in a prototype engine as a suitable candidate for replication. The frequency of the instability was 525 Hz, with an amplitude of approximately 1.5-psi peak-to-peak at a burner pressure of 200 psia. The single-nozzle experimental combustor rig was designed to preserve subcomponent lengths, cross sectional area distribution, flow distribution, pressure-drop distribution, temperature distribution, and other factors previously found to be determinants of burner acoustic frequencies, mode shapes, gain, and damping. Analytical models were used to predict the acoustic resonances of both the engine combustor and proposed experiment. The analysis confirmed that the test rig

  1. 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.

  2. 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.

  3. 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.

  4. Design and preliminary results of a fuel flexible industrial gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Novick, A. S.; Troth, D. L.; Yacobucci, H. G.

    1981-01-01

    The design characteristics are presented of a fuel tolerant variable geometry staged air combustor using regenerative/convective cooling. The rich/quench/lean variable geometry combustor is designed to achieve low NO(x) emission from fuels containing fuel bound nitrogen. The physical size of the combustor was calculated for a can-annular combustion system with associated operating conditions for the Allison 570-K engine. Preliminary test results indicate that the concept has the potential to meet emission requirements at maximum continuous power operation. However, airflow sealing and improved fuel/air mixing are necessary to meet Department of Energy program goals.

  5. 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

  6. Flame dynamics in a micro-channeled combustor

    NASA Astrophysics Data System (ADS)

    Hussain, Taaha; Markides, Christos N.; Balachandran, Ramanarayanan

    2015-01-01

    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

  7. 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.

  8. 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.

  9. Gas turbine combustor exit piece with hinged connections

    SciTech Connect

    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.

  10. 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.

  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. 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.

  13. 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.

  14. Determination of decay coefficients for combustors with acoustic absorbers

    NASA Technical Reports Server (NTRS)

    Mitchell, C. E.; Espander, W. R.; Baer, M. R.

    1972-01-01

    An analytical technique for the calculation of linear decay coefficients in combustors with acoustic absorbers is presented. Tuned circumferential slot acoustic absorbers were designed for the first three transverse modes of oscillation, and decay coefficients for these absorbers were found as a function of backing distance for seven different chamber configurations. The effectiveness of the absorbers for off-design values of the combustion response and acoustic mode is also investigated. Results indicate that for tuned absorbers the decay coefficient increases approximately as the cube of the backing distance. For most off-design situations the absorber still provides a damping effect. However, if an absorber designed for some higher mode of oscillation is used to damp lower mode oscillations, a driving effect is frequently found.

  15. Turbulent transport measurements in a model of GT-combustor

    NASA Astrophysics Data System (ADS)

    Chikishev, L. M.; Gobyzov, O. A.; Sharaborin, D. K.; Lobasov, A. S.; Dulin, V. M.; Markovich, D. M.; Tsatiashvili, V. V.

    2016-10-01

    To reduce NOx formation modern industrial power gas-turbines utilizes lean premixed combustion of natural gas. The uniform distribution of local fuel/air ratio in the combustion chamber plays one of the key roles in the field of lean combustion to prevent thermo-acoustic pulsations. Present paper reports on simultaneous Particle Image Velocimetry and acetone Planar Laser Induced Fluorescence measurements in a cold model of GT-combustor to investigate mixing processes which are relevant to the organization of lean premixed combustion. Velocity and passive admixture pulsations correlations were measured to verify gradient closer model, which is often used in Reynolds-Averaged Navier-Stokes (RANS) simulation of turbulent mixing.

  16. Nondestructive evaluation of environmental barrier coatings in CFCC combustor liners.

    SciTech Connect

    Sun, J. G.; Benz, J.; Ellingson, W. A.; Kimmel, J. B.; Price, J. R.; Energy Technology; Solar Turbines, Inc

    2007-01-01

    Advanced combustor liners fabricated of SiC/SiC continuous fiber-reinforced ceramic composite (CFCC) and covered with environmental barrier coatings (EBCs) have been successfully tested in Solar Turbines Inc. field engines. The primary goal for the CFCC/EBC liners is to reach a 30,000-h lifetime. Because the EBCs, when applied on the hot surfaces of liners, protect the underlying CFCC from oxidation damage, their performance is critical in achieving the lifetime goal. To determine CFCC/EBC liner condition and assess operating damage, the liners were subjected to nondestructive evaluation (NDE) during various processing stages, as well as before and after the engine test. The NDE techniques included pulsed infrared thermal imaging, air-coupled ultrasonic scanning, and X-ray computerized tomography. It was found that EBC damage and spallation depend on the condition of the CFCC material. The NDE results and correlations with destructive examination are discussed.

  17. 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.

  18. Use of natural gas to reduce emissions in waste combustors

    SciTech Connect

    Abbasi, H.A.; Khinkis, M.J.; Dunnette, R.; Nakazato, Kunihiro

    1992-12-31

    The Institute of Gas Technology (IGT), together with industrial partners, is developing a technology that utilizes natural gas to reduce air pollutant emissions from municipal waste combustors (MWCs). This natural gas injection technology is termed METHANE de-NOX{sup sm}. The results of field evaluation tests carried out at a 90 tonne/day MWC in 1991 show simultaneous reductions of 60% in nitrogen oxides and 50% in carbon monoxide with natural gas injection equal to 15% of total waste heat input. Excess air requirements were also reduced by 40% thus increasing the overall waste-to-energy plant efficiency. This approach is now being combined with injection of sorbents to also reduce the emissions of hydrochloric acid, sulfur oxides, dioxins, and furans. This paper describes the overall system design, the results of field evaluations to date and the schedule for sorbent injection trials.

  19. Use of natural gas to reduce emissions in waste combustors

    SciTech Connect

    Abbasi, H.A.; Khinkis, M.J. ); Dunnette, R. ); Nakazato, Kunihiro )

    1992-01-01

    The Institute of Gas Technology (IGT), together with industrial partners, is developing a technology that utilizes natural gas to reduce air pollutant emissions from municipal waste combustors (MWCs). This natural gas injection technology is termed METHANE de-NOX[sup sm]. The results of field evaluation tests carried out at a 90 tonne/day MWC in 1991 show simultaneous reductions of 60% in nitrogen oxides and 50% in carbon monoxide with natural gas injection equal to 15% of total waste heat input. Excess air requirements were also reduced by 40% thus increasing the overall waste-to-energy plant efficiency. This approach is now being combined with injection of sorbents to also reduce the emissions of hydrochloric acid, sulfur oxides, dioxins, and furans. This paper describes the overall system design, the results of field evaluations to date and the schedule for sorbent injection trials.

  20. 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.

  1. Numerical simulation of pressure oscillations in a ramjet combustor

    NASA Technical Reports Server (NTRS)

    Jou, Wen-Huei; Menon, Suresh

    1989-01-01

    Large-eddy simulations of compressible cold flow in a ramjet combustor configuration were performed. The objectives were to investigate the mechanisms for vortex-acoustic interaction in such a device and to develop a simulation method that can be extended to cases with combustion heat release to study combustion instability. From the simulation, it was found that the separated shear layer rolls up into concentrated vortices that merge to form large coherent structures. These vortices interact with the choked nozzle downstream to produce an axial acoustic dipole. The spectrum of the pressure fluctuation at the base of the backward-facing step shows that there are two types of oscillations: an acoustic resonant mode and a vortex-acoustic coupled mode. Based on the flow physics observed in the simulations, a simple one-dimensional model for the vortex-acoustic coupled mode was proposed. The eigenvalue problem based on this model was solved to obtain the frequency of the coupled mode.

  2. Turbulent combustion modelling in a side dump ramjet combustor

    SciTech Connect

    Montazel, X.; Samaniego, J.M.; Lacas, F.; Poinsot, T.; Candel, S. Laboratoire d'Energetique Moleculaire et Macroscopique-Combustion, Chatenay-Malabry )

    1992-07-01

    Experimental and theoretical studies are reported on the turbulent flow in a side dump rectangular combustor. Hot wire velocimetry was used to measure the axial and transverse mean velocity and turbulence intensity components in the nonreacting case. Flow oscillations in nonreacting case are determined from spectral analysis of the signal. In the presence of combustion, gas analysis (CO, CO2) was carried out in the chamber to obtain the temperature field. Light emission from CH radicals provides additional information which is interpreted as mean rate of heat release per unit volume. Regions where combustion takes place are identified. Measurements are compared with calculations performed with a combustion model relying on the flamelet concept and using a transport equation for the surface density. Good agreement is achieved between the model predictions and experimental results. 26 refs.

  3. O and temperature in a hydrocarbon-fueled scramjet combustor

    NASA Astrophysics Data System (ADS)

    Goldenstein, C. S.; Schultz, I. A.; Spearrin, R. M.; Jeffries, J. B.; Hanson, R. K.

    2014-09-01

    The design and demonstration of a two-color tunable diode laser sensor for measurements of temperature and H2O in an ethylene-fueled model scramjet combustor are presented. This sensor probes multiple H2O transitions in the fundamental vibration bands near 2.5 μm that are up to 20 times stronger than those used by previous near-infrared H2O sensors. In addition, two design measures enabled high-fidelity measurements in the nonuniform flow field. (1) A recently developed calibration-free scanned-wavelength-modulation spectroscopy spectral-fitting strategy was used to infer the integrated absorbance of each transition without a priori knowledge of the absorption lineshape and (2) transitions with strengths that scale near-linearly with temperature were used to accurately determine the H2O column density and the H2O-weighted path-averaged temperature from the integrated absorbance of two transitions.

  4. Direct measurements of skin friction in a scramjet combustor

    NASA Technical Reports Server (NTRS)

    Deturris, D. J.; Schetz, J. A.; Hellbaum, R. F.

    1990-01-01

    A direction experimental measurement was made of the skin friction in the turbulent boundary layer of three-dimensional supersonic combustor flows. A floating element configuration was designed to detect two components of the small tangential shear forces due to the flow passing over a nonobtrusive instrument. A cantilevered sensing head extending from a very sensitive piezoresistive displacement transducer simultaneously measured displacement in the axial and transverse directions to the flow. The small overall deflection means that no self-nulling is required. Consistently repeatable output from the gage during testing show the skin friction with supersonic combustion is higher than for a corresponding noncombusting flow. The results are probably accurate to within 10 percent for the mean streamwise component.

  5. Turbulent Radiation Effects in HSCT Combustor Rich Zone

    NASA Technical Reports Server (NTRS)

    Hall, Robert J.; Vranos, Alexander; Yu, Weiduo

    1998-01-01

    A joint UTRC-University of Connecticut theoretical program was based on describing coupled soot formation and radiation in turbulent flows using stretched flamelet theory. This effort was involved with using the model jet fuel kinetics mechanism to predict soot growth in flamelets at elevated pressure, to incorporate an efficient model for turbulent thermal radiation into a discrete transfer radiation code, and to couple die soot growth, flowfield, and radiation algorithm. The soot calculations used a recently developed opposed jet code which couples the dynamical equations of size-class dependent particle growth with complex chemistry. Several of the tasks represent technical firsts; among these are the prediction of soot from a detailed jet fuel kinetics mechanism, the inclusion of pressure effects in the soot particle growth equations, and the inclusion of the efficient turbulent radiation algorithm in a combustor code.

  6. Turbulent combustion modelling in a side dump ramjet combustor

    NASA Astrophysics Data System (ADS)

    Montazel, X.; Samaniego, J. M.; Lacas, F.; Poinsot, T.; Candel, S.

    1992-07-01

    Experimental and theoretical studies are reported on the turbulent flow in a side dump rectangular combustor. Hot wire velocimetry was used to measure the axial and transverse mean velocity and turbulence intensity components in the nonreacting case. Flow oscillations in nonreacting case are determined from spectral analysis of the signal. In the presence of combustion, gas analysis (CO, CO2) was carried out in the chamber to obtain the temperature field. Light emission from CH radicals provides additional information which is interpreted as mean rate of heat release per unit volume. Regions where combustion takes place are identified. Measurements are compared with calculations performed with a combustion model relying on the flamelet concept and using a transport equation for the surface density. Good agreement is achieved between the model predictions and experimental results.

  7. 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

  8. Formation of polychlorinated dibenzo-p-dioxins and dibenzofurans on secondary combustor/boiler ash from a rotary kiln burning hazardous waste.

    PubMed

    Addink, R; Altwicker, E R

    2004-10-18

    Ash from the secondary combustor/boiler of a rotary kiln burning hazardous chemical waste was tested in the laboratory for its potential to form polychlorinated dibenzo-p-dioxins and dibenzofurans (PCDD/F). The ash contained only a small quantity of "native" PCDD/F, i.e., formed on the ash in the facility. However, it produced a considerable amount of these compounds when heated in 10% O(2)/N(2) under "de novo" conditions, i.e., with residual carbon (present on the ash as result of incomplete combustion) as the only organic material. The ash yielded PCDD/F for up to 90 min; gave PCDD/F yields proportional to the amount of ash used in the reaction bed; and displayed an optimum temperature range for formation (397-548 degrees C) higher than seen for most municipal solid waste (MSW) fly ashes. The role of copper and iron as catalytic material on the ash is discussed.

  9. Experimental study on premixed CH{sub 4}/air mixture combustion in micro Swiss-roll combustors

    SciTech Connect

    Zhong, Bei-Jing; Wang, Jian-Hua

    2010-12-15

    Excess enthalpy combustion is a promising approach to stabilize flame in micro-combustors. Using a Swiss-roll combustor configuration, excess enthalpy combustion can be conveniently achieved. In this work, three types of Swiss-roll combustors with double spiral-shaped channels were designed and fabricated. The combustors were tested using methane/air mixtures of various equivalence ratios. Both temperature distributions and extinction limits were determined for each combustor configuration at different methane mass flow rates. Results indicate that the Swiss-roll combustors developed in the current study greatly enhance combustion stability in center regions of the combustors. At the same time, excess enthalpy combustors of the Swiss-roll configuration significantly extend the extinction limits of methane/air mixtures. In addition, the effects of combustor configurations and thermal insulation arrangements on temperature distributions and extinction limits were evaluated. With heat losses to the environment being significant, the use of thermal insulations further enhances the flame stability in center regions of the Swiss-roll combustors and extends flammable ranges. (author)

  10. Enabling Advanced Modeling and Simulations for Fuel-Flexible Combustors

    SciTech Connect

    Pitsch, Heinz

    2010-05-31

    The overall goal of the present project is to enable advanced modeling and simulations for the design and optimization of fuel-flexible turbine combustors. For this purpose we use a high fidelity, extensively-tested large-eddy simulation (LES) code and state-of-the-art models for premixed/partially-premixed turbulent combustion developed in the PI's group. In the frame of the present project, these techniques are applied, assessed, and improved for hydrogen enriched premixed and partially premixed gas-turbine combustion. Our innovative approaches include a completely consistent description of flame propagation; a coupled progress variable/level set method to resolve the detailed flame structure, and incorporation of thermal-diffusion (non-unity Lewis number) effects. In addition, we have developed a general flamelet-type transformation holding in the limits of both non-premixed and premixed burning. As a result, a model for partially premixed combustion has been derived. The coupled progress variable/level method and the general flamelet transformation were validated by LES of a lean-premixed low-swirl burner that has been studied experimentally at Lawrence Berkeley National Laboratory. The model is extended to include the non-unity Lewis number effects, which play a critical role in fuel-flexible combustor with high hydrogen content fuel. More specifically, a two-scalar model for lean hydrogen and hydrogen-enriched combustion is developed and validated against experimental and direct numerical simulation (DNS) data. Results are presented to emphasize the importance of non-unity Lewis number effects in the lean-premixed low-swirl burner of interest in this project. The proposed model gives improved results, which shows that the inclusion of the non-unity Lewis number effects is essential for accurate prediction of the lean-premixed low-swirl flame.

  11. Enabling Advanced Modeling and Simulations for Fuel-Flexible Combustors

    SciTech Connect

    Heinz Pitsch

    2010-05-31

    The overall goal of the present project is to enable advanced modeling and simulations for the design and optimization of fuel-flexible turbine combustors. For this purpose we use a high-fidelity, extensively-tested large-eddy simulation (LES) code and state-of-the-art models for premixed/partially-premixed turbulent combustion developed in the PI's group. In the frame of the present project, these techniques are applied, assessed, and improved for hydrogen enriched premixed and partially premixed gas-turbine combustion. Our innovative approaches include a completely consistent description of flame propagation, a coupled progress variable/level set method to resolve the detailed flame structure, and incorporation of thermal-diffusion (non-unity Lewis number) effects. In addition, we have developed a general flamelet-type transformation holding in the limits of both non-premixed and premixed burning. As a result, a model for partially premixed combustion has been derived. The coupled progress variable/level method and the general flamelet tranformation were validated by LES of a lean-premixed low-swirl burner that has been studied experimentally at Lawrence Berkeley National Laboratory. The model is extended to include the non-unity Lewis number effects, which play a critical role in fuel-flexible combustor with high hydrogen content fuel. More specifically, a two-scalar model for lean hydrogen and hydrogen-enriched combustion is developed and validated against experimental and direct numerical simulation (DNS) data. Results are presented to emphasize the importance of non-unity Lewis number effects in the lean-premixed low-swirl burner of interest in this project. The proposed model gives improved results, which shows that the inclusion of the non-unity Lewis number effects is essential for accurate prediction of the lean-premixed low-swirl flame.

  12. Integral cylindrical cyclone and loopseal assembly for CFB combustors

    SciTech Connect

    Alliston, M.G.; Wu, S.; Sinn, B.T.; Lato, K.

    1995-12-31

    The cyclone solids separator has been one of the main features characteristically associated with many Circulating Fluidized Bed combustor designs. Most designs incorporating the cyclone coflector also include a pressure sealing device through which the solids collected by the cyclone are returned to the combustor. The cyclone and the loopseal have typically been two separate components, and various designs have arisen for each component comprising refractory, water-cooled, or steam cooled equipment. The two obvious advantages of the use of water-cooled or steam-, cooled equipment are the reduction of refractory requirements and the conversion of these components into boiler-useful devices. The creation of these devices from tube and header assemblies has followed more-or-less artful paths, the directions of which have generally been based upon the goal of maintaining the specific geometric forms that had previously been used for CFB cyclones and loopseals. The assembly described in this paper was instead developed based upon the goals of (a) ease of manufacturing, (b) elimination of problematic components, (c) maximizing the boiler-usefulness of the surfaces created, and (d) adding potential process benefits. The resulting embodiment of these goals was a cylindrical vessel in which the upper portion forms the solids separator and the lower section is divided into chambers to form the loopseal. This report describes the details of the Integral Cylindrical Cyclone and Loopseal (ICCL), and also describes its performance as demonstrated in a CFB cold flow model incorporating the new apparatus. The commercial application of the ICCL in a 1.2 MM lb steam/hr CFB boiler as well as its potential application in reheat CFB units are also discussed.

  13. 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.

  14. 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.

  15. 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.

  16. System and method for reducing combustion dynamics and NO.sub.x in a combustor

    DOEpatents

    Uhm, Jong H.; Johnson, Thomas Edward

    2015-11-20

    A system for reducing combustion dynamics and NO.sub.x in a combustor includes a tube bundle that extends radially across at least a portion of the combustor, wherein the tube bundle comprises an upstream surface axially separated from a downstream surface. A shroud circumferentially surrounds the upstream and downstream surfaces. A plurality of tubes extends through the tube bundle from the upstream surface through the downstream surface, wherein the downstream surface is stepped to produce tubes having different lengths through the tube bundle. A method for reducing combustion dynamics and NO.sub.x in a combustor includes flowing a working fluid through a plurality of tubes radially arranged between an upstream surface and a downstream surface of an end cap that extends radially across at least a portion of the combustor, wherein the downstream surface is stepped.

  17. Investigation of soot and carbon formation in small gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Rosfjord, T. J.

    1982-01-01

    An investigation of hardware configurations which attempt to minimize carbon and soot-production without sacrificing performance in small gas turbine combustors was conducted. Four fuel injectors, employing either airblast atomization, pressure atomization, or fuel vaporization techniques were combined with nozzle air swirlers and injector sheaths. Eight configurations were screened at sea-level takeoff and idle test conditions. Selected configurations were focused upon in an attempt to quantify the influence of combustor pressure, inlet temperature, primary zone operation, and combustor loading on soot and carbon formation. Cycle tests were also performed. It was found that smoke emission levels depended on the combustor fluid mechanics, the atomization quality of the injector and the fuel hydrogen content.

  18. On the modelling of scalar and mass transport in combustor flows

    NASA Technical Reports Server (NTRS)

    Nikjooy, M.; So, R. M. C.

    1989-01-01

    Results are presented of a numerical study of swirling and nonswirling combustor flows with and without density variations. Constant-density arguments are used to justify closure assumptions invoked for the transport equations for turbulent momentum and scalar fluxes, which are written in terms of density-weighted variables. Comparisons are carried out with measurements obtained from three different axisymmetric model combustor experiments covering recirculating flow, swirling flow, and variable-density swirling flow inside the model combustors. Results show that the Reynolds stress/flux models do a credible job of predicting constant-density swirling and nonswirling combustor flows with passive scalar transport. However, their improvements over algebraic stress/flux models are marginal. The extension of the constant-density models to variable-density flow calculations shows that the models are equally valid for such flows.

  19. MERCURY CONTROL IN MUNICIPAL WASTE COMBUSTORS AND COAL-FIRED UTILITIES

    EPA Science Inventory

    Control of mercury (Hg) emissions from municipal waste combustors (MWCs) and coal-fired utilities has attracted attention due to current and potential regulations. Among several techniques evaluated for Hg control, dry sorbent injection (primarily injection of activated carbon) h...

  20. Effect of fuel injector type on performance and emissions of reverse-flow combustor

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

    The combustion process in a reverse-flow combustor suitable for a small gas turbine engine was investigated to evaluate the effect of fuel injector type on performance and emissions. Fuel injector configurations using pressure-atomizing, spill-flow, air blast, and air-assist techniques were compared and evaluated on the basis of performance obtained in a full-scale experimental combustor operated at inlet conditions corresponding to takeoff, cruise, low power, and idle and typical of a 16:1-pressure-ratio turbine engine. Major differences in combustor performance and emissions characteristics were experienced with each injector type even though the aerodynamic configuration was common to most combustor models. Performance characteristics obtained with the various fuel injector types could not have been predicted from bench-test injector spray characteristics. The effect of the number of operating fuel injectors on performance and emissions is also presented.

  1. A variable-geometry combustor used to study primary and secondary zone stoichiometry

    NASA Technical Reports Server (NTRS)

    Briehl, D.; Schultz, D. F.; Ehlers, R. C.

    1983-01-01

    A combustion program is underway to evaluate fuel quality effects on gas turbine combustors. A rich-lean variable geometry combustor design was chosen to evaluate fuel quality effects over a wide range of primary and secondary zone equivalence ratios at simulated engine operating conditions. The first task of this effort, was to evaluate the performance of the variable geometry combustor. The combustor incorporates three stations of variable geometry to control primary and secondary zone equivalence ratio and overall pressure loss. Geometry changes could be made while a test was in progress through the use of remote control actuators. The primary zone liner was water cooled to eliminate the concern of liner durability. Emissions and performance data were obtained at simulated engine conditions of 80 percent and full power. Inlet air temperature varied from 611 to 665K, inlet total pressure varied from 1.02 to 1.24 MPa, reference velocity was a constant 1400 K.

  2. Large Municipal Waste Combustors (LMWC): New Source Performance Standards (NSPS) and Emissions Guidelines

    EPA Pesticide Factsheets

    Learn about the NSPS, emission guidelines and compliance times for large municipal waste combustors (MWC) by reading the rule summary, rule history and the federal register citations and supporting documents

  3. A chemical reactor network for oxides of nitrogen emission prediction in gas turbine combustor

    NASA Astrophysics Data System (ADS)

    Hao, Nguyen Thanh

    2014-06-01

    This study presents the use of a new chemical reactor network (CRN) model and non-uniform injectors to predict the NOx emission pollutant in gas turbine combustor. The CRN uses information from Computational Fluid Dynamics (CFD) combustion analysis with two injectors of CH4-air mixture. The injectors of CH4-air mixture have different lean equivalence ratio, and they control fuel flow to stabilize combustion and adjust combustor's equivalence ratio. Non-uniform injector is applied to improve the burning process of the turbine combustor. The results of the new CRN for NOx prediction in the gas turbine combustor show very good agreement with the experimental data from Korea Electric Power Research Institute.

  4. Design of Combustor for Long-range Ram-jet Engine and Performance of Rectangular Analog

    NASA Technical Reports Server (NTRS)

    Rayle, Warren D; Koch, Richard G

    1954-01-01

    The report describes the design of a piloted combustor intended for a ram-jet engine of long flight range. The unit comprises a large annular basket of V-type cross-section, the inner surface of which is slotted and bent into small V-gutters. At the trailing edge of the basket, eight V-gutters are used to propagate the flame into the main stream. A rectangular analog of this combustor was tested at air-flow conditions corresponding to those that might be obtained during cruise. At these conditions, combustion efficiencies of as much as 90 percent were calculated for the combustor at the design equivalence ratio of 0.52. The performance of the unit was relatively insensitive to mounting and flow variables; the greatest effect on efficiency was that of the manner and location of the fuel injection. A full-scale version of this combustor has been designed for a 48-inch-diameter engine.

  5. Flow aerodynamics modeling of an MHD swirl combustor - Calculations and experimental verification

    NASA Technical Reports Server (NTRS)

    Gupta, A. K.; Beer, J. M.; Louis, J. F.; Busnaina, A. A.; Lilley, D. G.

    1981-01-01

    The paper describes a computer code for calculating the flow dynamics of a constant-density flow in the second-stage trumpet shaped nozzle section of a two-stage MHD swirl combustor for application to a disk generator. The primitive pressure-velocity variable, finite-difference computer code has been developed for the computation of inert nonreacting turbulent swirling flows in an axisymmetric MHD model swirl combustor. The method and program involve a staggered grid system for axial and radial velocities, and a line relaxation technique for the efficient solution of the equations. The code produces as output the flow field map of the nondimensional stream function, axial and swirl velocity. It was found that the best location for seed injection to obtain a uniform distribution at the combustor exit is in the central location for seed injected at the entrance to the second stage combustor.

  6. Impacts of broadened-specification fuels on aircraft turbine engine combustors

    SciTech Connect

    Bahr, D.W.

    1982-11-01

    In the future, aircraft turbine engines may be required to accommodate fuels with lower hydrogen contents, fluidity/volatility characteristics, and thermal stability ratings than those of present-day jet fuels. The key results of several investigations conducted to quantify the impacts of such fuel property changes on the performance, pollutant emission, and durability characteristics of aircraft turbine engine combustors are reviewed. These results were obtained in both component and engine tests of the current production combustors used in the CF6-6, CF6-50, F101/CFM56, and J79/CJ805 engines. Based on these results, it is concluded that the most significant concerns associated with such fuel property changes are decreased combustor life, decreased ground starting/altitude relight capabilities, and increased fuel injector gumming/coking tendencies. The combustor design improvements that appear to be needed to permit the use of jet fuels of significantly lower quality than present day fuels are also reviewed.

  7. Coherent Anti-Stokes Raman Spectroscopic Thermometry in a Supersonic Combustor

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

    An experiment has been conducted to acquire data for the validation of computational fluid dynamics codes used in the design of supersonic combustors. The flow in a supersonic combustor, consisting of a diverging duct with a single downstream-angled wail injector, is studied. Combustor entrance Mach number is 2 and enthalpy nominally corresponds to Mach 7 flight. The primary measurement technique is coherent anti-Stokes Raman spectroscopy, but 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 to minimize systematic errors. Temperature maps are obtained at several planes in the flow for a case in which the combustor is piloted by injecting fuel upstream of the main injector and one case in which it is not piloted. Boundary conditions and uncertainties are characterized.

  8. Transient analysis of a pulsed detonation combustor using the numerical propulsion system simulation

    NASA Astrophysics Data System (ADS)

    Hasler, Anthony Scott

    The performance of a hybrid mixed flow turbofan (with detonation tubes installed in the bypass duct) is investigated in this study and compared with a baseline model of a mixed flow turbofan with a standard combustion chamber as a duct burner. Previous studies have shown that pulsed detonation combustors have the potential to be more efficient than standard combustors, but they also present new challenges that must be overcome before they can be utilized. The Numerical Propulsion System Simulation (NPSS) will be used to perform the analysis with a pulsed detonation combustor model based on a numerical simulation done by Endo, Fujiwara, et. al. Three different cases will be run using both models representing a take-off situation, a subsonic cruise and a supersonic cruise situation. Since this study investigates a transient analysis, the pulse detonation combustor is run in a rig setup first and then its pressure and temperature are averaged for the cycle to obtain quasi-steady results.

  9. Simulation of a Single-Element Lean-Direct Injection Combustor Using Arbitrary Polyhedral Mesh

    NASA Technical Reports Server (NTRS)

    Wey, Thomas; Liu, Nan-Suey

    2012-01-01

    This paper summarizes procedures of generating the arbitrary polyhedral mesh as well as presents sample results from its application to the numerical solution of a single-element LDI combustor using a preliminary version of the new OpenNCC.

  10. Experimental Investigation of Reacting Flow Characteristics in a Dual-Mode Scramjet Combustor

    NASA Astrophysics Data System (ADS)

    Shi, Deyong; Song, Wenyan; Ye, Jingfeng; Tao, Bo; Wang, Yanhua

    2016-06-01

    In this work, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS), respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y-direction, i.e. the transverse direction.

  11. Effects of Burning Alternative Fuel in a 5-Cup Combustor Sector

    NASA Technical Reports Server (NTRS)

    Tacina, K. M.; Chang, C. T.; Lee, C.-M.; He, Z.; Herbon, J.

    2015-01-01

    A goal of NASA's Environmentally Responsible Aviation (ERA) program is to develop a combustor that will reduce the NOx emissions and that can burn both standard and alternative fuels. To meet this goal, NASA partnered with General Electric Aviation to develop a 5-cup combustor sector; this sector was tested in NASA Glenn's Advanced Subsonic Combustion Rig (ASCR). To verify that the combustor sector was fuel-flexible, it was tested with a 50-50 blend of JP-8 and a biofuel made from the camelina sativa plant. Results from this test were compared to results from tests where the fuel was neat JP-8. Testing was done at three combustor inlet conditions: cruise, 30% power, and 7% power. When compared to burning JP-8, burning the 50-50 blend did not significantly affect emissions of NOx, CO, or total hydrocarbons. Furthermore, it did not significantly affect the magnitude and frequency of the dynamic pressure fluctuations.

  12. Small gas turbine combustor study: Fuel injector performance in a transpiration-cooled liner

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

    The effect of fuel injection technique on the performance of an advanced reverse flow combustor liner constructed of Lamilloy (a multilaminate transpiration type material) was determined. Performance and emission levels are documented over a range of simulated flight conditions using simplex pressure atomizing, spill return, and splash cone airblast injectors. A parametric evaluation of the effect of increased combustor loading with each of the fuel injector types is obtained.

  13. Flowpath Design of a Three-Tube Valve-Less Pulse Detonation Combustor

    DTIC Science & Technology

    2009-09-01

    geometry. 15. NUMBER OF PAGES 77 14. SUBJECT TERMS PDE Pulse Detonation Engine Side Dump Combustor Recirculation Zone 16. PRICE CODE 17...INTENTIONALLY LEFT BLANK 1 I. BACKGROUND A. INTRODUCTION Multi-tube Pulse Detonation Engines ( PDEs ) represent a significant opportunity for reducing...rocket, for example. PDE Envelope (representative) 3 B. ENGINE OPERATION PDEs , of which the Pulse Detonation Combustor is the operational core, are

  14. Design and Evaluation of a Single-Inlet Pulse Detonation Combustor

    DTIC Science & Technology

    2011-06-01

    Kilogram/second m/s Meters/ second N Nitrogen NPS Naval Postgraduate School O Oxygen PDC Pulse Detonation Combustion PDE Pulse Detonation Engine...EVALUATION OF A SINGLE-INLET PULSE DETONATION COMBUSTOR by Danny Soria June 2011 Thesis Advisor: Christopher M. Brophy Second Reader: Garth V...COVERED Master’s Thesis 4. TITLE AND SUBTITLE Design and Evaluation of a Single-Inlet Pulse Detonation Combustor 6. AUTHOR(S) Danny Soria 5

  15. A Portable Thermoelectric Power Generator Based on a Microfabricated Silicon Combustor with Low Resistance to Flow

    DTIC Science & Technology

    2010-12-14

    stainless steel device has been used to combust hydrogen, propane [6, 15 , 16], and methanol [1 7]. The combustor was integrated with a commercial...22] and wet chemical etch [14,23-25] techniques. Steel [14, 16, 23-27] and ceramic [13, 28] combustors have also been investigated, The choice of...intended for use over a wide temperature range is no simple task. In this work, a silicon combustion zone was integrated with steel fluidi c conduits

  16. The impact of emissions standards on the design of aircraft gas turbine engine combustors

    NASA Technical Reports Server (NTRS)

    Rudey, R. A.

    1976-01-01

    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.

  17. Small gas turbine combustor study - Fuel injector performance in a transpiration-cooled liner

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

    The effect of fuel injection technique on the performance of an advanced reverse flow combustor liner constructed of Lamilloy (a multilaminate transpiration type material) was determined. Performance and emission levels are documented over a range of simulated flight conditions using simplex pressure atomizing, spill return, and splash cone airblast injectors. A parametric evaluation of the effect of increased combustor loading with each of the fuel injector types is obtained.

  18. Combustor having mixing tube bundle with baffle arrangement for directing fuel

    DOEpatents

    Hughes, Michael John; McConnaughhay, Johnie Franklin

    2016-08-23

    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, and a plurality of tubes extend from the upstream surface through the downstream surface to provide fluid communication through the tube bundle. A barrier extends radially inside the tube bundle between the upstream and downstream surfaces, and a baffle extends axially inside the tube bundle between the upstream surface and the barrier.

  19. 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.

  20. Effect of combustor geometry and fuel injection scheme on the combustion process in a supersonic flow

    NASA Astrophysics Data System (ADS)

    Cai, Zun; Wang, Zhenguo; Sun, Mingbo; Bai, Xue-Song

    2016-12-01

    The combustion process in a hydrogen fueled scramjet combustor with a rearwall-expansion cavity was investigated numerically under inflow conditions of Ma=2.52 with stagnation pressure P0=1.6 Mpa and stagnation temperature T0=1486 K. The numerical solver was first evaluated for supersonic reactive flows in a similar combustor configuration where experimental data is available. Wall-pressure distribution was compared with the experiments, and grid independency analysis and chemical mechanism comparison were conducted. The numerical results showed fairly good agreements with the available experimental data under supersonic combustion conditions. Then the numerical solver was used to study the effects of combustor geometry, fuel injection scheme and injection equivalence ratio on the combustion process. It was found that under the same fuel injection condition, the combustor configuration with a rearwall-expansion cavity is in favor of the supersonic combustion mode and present better ability of thermal choking prevention than the other combustor configurations. For the rearwall-expansion cavity combustor, the supersonic flow field was found to be sensitive to the injector position and injection scheme, but not highly sensitive to the injection pressure. Besides, rearwall-expansion cavity with the combined fuel injection scheme (with an injection upstream the cavity and a direct injection on the rear wall) is an optimized injection scheme during the flame stabilization process.

  1. The Coal Tech Advanced Cyclone Combustor Demonstration Project: A DOE assessment

    SciTech Connect

    Not Available

    1993-05-01

    The objective of this project was to demonstrate a technology for retrofitting oil/gas designed boilers, and conventional pulverized coal-fired boilers, by using the patented air-cooled slagging coal combustor in place of oil/gas/coal burners. The project aimed to utilize coals with a wide range of sulfur contents and to achieve efficient combustion under fuel-rich conditions. The three performance goals of the combustor were to limit emissions of SO{sub 2} and NO{sub x}, while maintaining maximum sulfur retention in the slag removed from the combustor. A slagging cyclone combustor is a high-temperature device in which a high-velocity swirling gas is used to burn crushed or pulverized coal. The key novel feature of this combustor is the use of air cooling. This is accomplished by using a ceramic liner which is cooled by secondary air and maintained at a temperature high enough to keep the slag in a liquid, free-flowing state. The arrangement also promotes slag retention in the combustor, an important feature for retrofitting in boilers designed for oil/gas.

  2. Investigation of heat transfer and combustion in the advanced fluidized bed combustor (FBC)

    SciTech Connect

    Lee, S.W.

    1999-09-01

    This technical report summarizes the research conducted and progress achieved during the period from April 1, 1998 to June 30, 1998. The numerical simulation was continued to determine the concentration distribution of the gas species, heat flux and heat transfer coefficients in the hot combustor model. The different gas concentration profiles showed the gas mixing characteristics along the combustor height. The center zone of the combustor has a relatively high methane mass concentration. The injection of secondary air squeezes the uprising flue gas and methane that causes the fuel-lean zone near the secondary air nozzles. The carbon dioxide concentration increased with the increasing of the combustor height. The peak concentration of oxygen remains at the combustor wall because of the secondary injection. The heat flux on the wall of the upper chamber is much higher than that of the lower chamber. It is believed that the heat flux is affected by the designed strong swirl and secondary air injection. The heat transfer coefficient changes along the combustor height were also affected by the multiple secondary air injection. The numerical simulation results could verify the predictions of the experimental results. It is a quite similar trend of the heat transfer coefficient changes based on the combustion test results.

  3. Unstructured LES of Reacting Multiphase Flows in Realistic Gas Turbine Combustors

    NASA Technical Reports Server (NTRS)

    Ham, Frank; Apte, Sourabh; Iaccarino, Gianluca; Wu, Xiao-Hua; Herrmann, Marcus; Constantinescu, George; Mahesh, Krishnan; Moin, Parviz

    2003-01-01

    As part of the Accelerated Strategic Computing Initiative (ASCI) program, an accurate and robust simulation tool is being developed to perform high-fidelity LES studies of multiphase, multiscale turbulent reacting flows in aircraft gas turbine combustor configurations using hybrid unstructured grids. In the combustor, pressurized gas from the upstream compressor is reacted with atomized liquid fuel to produce the combustion products that drive the downstream turbine. The Large Eddy Simulation (LES) approach is used to simulate the combustor because of its demonstrated superiority over RANS in predicting turbulent mixing, which is central to combustion. This paper summarizes the accomplishments of the combustor group over the past year, concentrating mainly on the two major milestones achieved this year: 1) Large scale simulation: A major rewrite and redesign of the flagship unstructured LES code has allowed the group to perform large eddy simulations of the complete combustor geometry (all 18 injectors) with over 100 million control volumes; 2) Multi-physics simulation in complex geometry: The first multi-physics simulations including fuel spray breakup, coalescence, evaporation, and combustion are now being performed in a single periodic sector (1/18th) of an actual Pratt & Whitney combustor geometry.

  4. A feasibility study of oil shale fired pulse combustors with applications to oil shale retorting

    SciTech Connect

    Morris, G.J.; Johnson, E.K.; Zhang, G.Q.; Roach, R.A.

    1992-07-01

    The results of the experimental investigation performed to determine the feasibility of using pulverized Colorado oil shale to fuel a bench scale pulse combustor reveal that oil shale cannot sustain pulsations when used alone as fuel. Trace amounts of propane mixed with the oil shale enabled the pulsations, however. Up to 80% of the organic material in the oil shale was consumed when it was mixed with propane in the combustor. Beyond the feasibility objectives, the operating conditions of the combustor fuel with propane and mixtures of oil shale and propane were characterized with respect to pulsation amplitude and frequency and the internal combustor wall temperature over fuel lean and fuel rich stoichiometries. Maximum pressure excursions of 12.5 kPa were experienced in the combustor. Pulsation frequencies ranged from 50 to nearly 80 Hz. Cycle resolved laser Doppler anemometry velocities were measured at the tail pipe exit plane. Injecting inert mineral matter (limestone) into the pulse combustor while using propane fuel had only a slight effect on the pulsation frequency for the feed rates tested.

  5. A Design Methodology for Rapid Implementation of Active Control Systems Across Lean Direct Injection Combustor Platforms

    NASA Technical Reports Server (NTRS)

    Baumann, William T.; Saunders, William R.; Vandsburger, Uri; Saus, Joseph (Technical Monitor)

    2003-01-01

    The VACCG team is comprised of engineers at Virginia Tech who specialize in the subject areas of combustion physics, chemical kinetics, dynamics and controls, and signal processing. Currently, the team's work on this NRA research grant is designed to determine key factors that influence combustion control performance through a blend of theoretical and experimental investigations targeting design and demonstration of active control for three different combustors. To validiate the accuracy of conclusions about control effectiveness, a sequence of experimental verifications on increasingly complex lean, direct injection combustors is underway. During the work period January 1, 2002 through October 15, 2002, work has focused on two different laboratory-scale combustors that allow access for a wide variety of measurements. As the grant work proceeds, one key goal will be to obtain certain knowledge about a particular combustor process using a minimum of sophisticated measurements, due to the practical limitations of measurements on full-scale combustors. In the second year, results obtained in the first year will be validated on test combustors to be identified in the first quarter of that year. In the third year, it is proposed to validate the results at more realistic pressure and power levels by utilizing the facilities at the Glenn Research Center.

  6. NASA Project Develops Next-Generation Low-Emissions Combustor Technologies

    NASA Technical Reports Server (NTRS)

    Lee, Chi-Ming; Chang, Clarence T.; Herbon, John T.; Kramer, Stephen K.

    2013-01-01

    NASA's Environmentally Responsible Aviation (ERA) Project is working with industry to develop the fuel flexible combustor technologies for a new generation of low-emissions engine targeted for the 2020 timeframe. These new combustors will reduce nitrogen oxide (NOx) emissions to half of current state-of-the-art (SOA) combustors, while simultaneously reducing noise and fuel burn. The purpose of the low NOx fuel-flexible combustor research is to advance the Technology Readiness Level (TRL) and Integration Readiness Level (IRL) of a low NOx, fuel flexible combustor to the point where it can be integrated in the next generation of aircraft. To reduce project risk and optimize research benefit NASA chose to found two Phase 1 contracts. The first Phase 1 contracts went to engine manufactures and were awarded to: General Electric Company, and Pratt & Whitney Company. The second Phase 1 contracts went to fuel injector manufactures Goodrich Corporation, Parker Hannifin Corporation, and Woodward Fuel System Technology. In 2012, two sector combustors were tested at NASA's ASCR. The results indicated 75% NOx emission reduction below the 2004 CAEP/6 regulation level.

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

    NASA Technical Reports Server (NTRS)

    Locke, Randy J.; Anderson, Robert C.; Zaller, Michelle M.; Hicks, Yolanda R.

    1998-01-01

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

  8. Catalytic efficiency of designed catalytic proteins

    PubMed Central

    Korendovych, Ivan V; DeGrado, William F

    2014-01-01

    The de novo design of catalysts that mimic the affinity and specificity of natural enzymes remains one of the Holy Grails of chemistry. Despite decades of concerted effort we are still unable to design catalysts as efficient as enzymes. Here we critically evaluate approaches to (re)design of novel catalytic function in proteins using two test cases: Kemp elimination and ester hydrolysis. We show that the degree of success thus far has been modest when the rate enhancements seen for the designed proteins are compared with the rate enhancements by small molecule catalysts in solvents with properties similar to the active site. Nevertheless, there are reasons for optimism: the design methods are ever improving and the resulting catalyst can be efficiently improved using directed evolution. PMID:25048695

  9. 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.

  10. 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.

  11. Experimental/Computational Studies of Combined Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors

    DTIC Science & Technology

    2010-10-01

    Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors 5a. CONTRACT NUMBER 5b. GRANT NUMBER FA9550-04-1-0387 5c...development of new tunable diode-laser diagnostics for supersonic combustors . The scientific results obtained from this MURI program were disseminated widely...Inlets and Scramjet Combustors TABLE OF CONTENTS TOPIC PAGE 1. COVER PAGE / ABSTRACT i 1. SUMMARY OF ACTIVITIES 1 2. PERSONNEL SUPPORTED 2 3

  12. Initiation Mechanisms of Low-loss Swept-ramp Obstacles for Deflagration to Detonation Transition in Pulse Detonation Combustors

    DTIC Science & Technology

    2009-12-01

    MECHANISMS OF LOW-LOSS SWEPT-RAMP OBSTACLES FOR DEFLAGRATION TO DETONATION TRANSITION IN PULSE DETONATION COMBUSTORS by Charles B. Myers IV...TITLE AND SUBTITLE Initiation Mechanisms of Low-loss Swept-ramp Obstacles for Deflagration to Detonation Transition in Pulse Detonation Combustors 6...DISTRIBUTION CODE A 13. ABSTRACT (maximum 200 words) In order to enhance the performance of pulse detonation combustors (PDCs), an efficient

  13. Experimental/Computational Studies of Combined-Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors

    DTIC Science & Technology

    2010-05-22

    Scramjet Combustors 5b. GRANT NUMBER FA9550-04-1-0387 5c. PROGRAM ELEMENT NUMBER 61103F 6. AUTHOR(S) 5d. PROJECT NUMBER 5094...technique for compressible jets-in-crossflow, and development of new tunable diode-laser diagnostics for supersonic combustors . The scientific results...Computational Studies of Combined-Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors TABLE OF CONTENTS

  14. NONEQUILIBRIUM SULFUR CAPTURE AND RETENTION IN AN AIR COOLED SLAGGING COAL COMBUSTOR

    SciTech Connect

    Dr. Bert Zauderer

    1999-03-15

    Calcium oxide injected in a slagging combustor reacts with the sulfur from coal combustion to form sulfur-bearing particles. They are deposited on the liquid slag layer on the combustor wall. Due to the low solubility of sulfur in slag, slag must be rapidly drained from the combustor to limit sulfur gas re-evolution. Analysis indicated that slag mass flow rates in excess of 400 lb/hr should limit sulfur re-evolution. The objective of this 42-month project was to perform a series of tests to determine the factors that control the retention of the sulfur in the slag. 36 days of testing on the combustor were completed prior to the end of this reporting period, 12/31/98. This compares with 16 tests required in the original project plan. Combustor tests in early 1997 with high (37%) ash, Indian coal confirmed that high slag mass flow rates of about 500 lb/hr resulted in retention in the slag of up to 20% of the injected sulfur content mineral matter. To further increase the slag flow rate, rice husks, which contain 20% ash, and rice husk char, which contain 70% ash, were co-fired with coal in the combustor. A series of 13 combustor tests were performed in fourth quarter of 1997 and a further 6 tests were performed in January 1998 and in the summer of 1998. The test objective was to achieve slag flow rates between 500 and 1,000 lb/hr. Due to the very low bulk density of rice husk, compared to pulverized coal, almost the entire test effort focused on developing methods for feeding the rice husks into combustor. In the last test of December 1997, a peak mineral matter, injection rate of 592 lb/hr was briefly achieved by injection of coal, rice husk char, gypsum, and limestone into the combustor. However, no significant sulfur concentration was measured in the slag removed from the combustor. The peak injection rate reached with biomass in the 1997 tests was 310 lb/hr with rice husk, and 584 lb/hr with rice husk char.

  15. Numerical investigation of oblique detonation waves for a shcramjet combustor

    NASA Astrophysics Data System (ADS)

    Fusina, Giovanni

    Research in hypersonic airbreathing propulsion strives to provide an efficient, cost-effective alternative to rocket propulsion for space transportation systems. The supersonic combustion ramjet (scramjet) is one of the most common hypersonic airbreathing propulsion concepts, but its massive combustor could have detrimental effects on its efficiency. The shock-induced combustion ramjet (shcramjet) overcomes this drawback by using standing oblique detonation waves (coupled shock-combustion fronts) as a means of nearly instantaneous heat addition. A numerical investigation of standing oblique detonation waves for their use in shcramjet propulsion is the main purpose of this investigation. The laminar, two-dimensional Navier-Stokes equations coupled with non-equilibrium hydrogen/air combustion models based on chemical kinetics are used to represent the physical system. The combustion models are incorporated into an in-house computational fluid dynamics solver based on a shock-capturing scheme by Yee scheme and an approximate factorization algorithm with a dual-time stepping technique to regain time-accuracy. The solver is validated with experimental data found in the literature. A time accurate simulation of the formation of a standing oblique detonation wave (ODW) near the Chapman Jouguet (minimum entropy) condition yields a non-oscillatory, stable structure. The stability of the ODW to inhomogeneities in the oncoming fuel/air mixture is assessed through other time-accurate simulations by artificially introducing small disturbances consisting of pure air just upstream of the ODW structure. The ODW is shown to be resilient to these disturbances: an upstream displacement is observed followed by the return of the ODW to its original position. Steady-state simulations are carried out to determine the effect of laminar boundary layers on ODW structures and properties above and below the Chapman-Jouguet point. A comparison with analogous inviscid simulations shows that the

  16. Non-intrusive measurements in a rocket engine combustor

    NASA Astrophysics Data System (ADS)

    Farhangi, S.; Gylys, V. T.; Jensen, R. J.

    1993-07-01

    In recent years analytical tools to characterize combustor flow have been developed in order to support design. To facilitate anchoring of combustion related physical models and the CFD codes in which they are incorporated, considerable development and application of non-intrusive combustion diagnostic capabilities has occurred. Raman spectroscopy can be used to simultaneously detect all polyatomic molecules present in significant concentrations and to determine gas temperature. This is because all molecules possess a distinct temperature dependent Raman spectrum. A multi-point diagnostic system for non-intrusive temperature and species profiling in rocket engines has been developed at Rocketdyne. In the present effort, the system has been undergoing validation for application to rocket engine component testing. A 4 inch diameter windowed combustor with a coaxial gas-gas injector was chosen for this series of validation experiments. Initially an excimer-pumped tunable dye laser and later a solid state Nd-Yag laser served as excitation sources. The Raman signal was dispersed by a monochromator and detected by a gated, intensified Charged Coupled Device (CCD) array. Experiments were carried out prior to each series of hot fire tests to ensure that the Raman signal detected was due to a spontaneous rather than a stimulated Raman emission process. Over sixty hot fire tests were conducted during the first series of tests with the excimer/dye laser. All hot fire testing was at a mixture ratio of 0.5 and chamber pressures of approximately 100 and approximately 300 psia. The Raman spectra of hydrogen, water vapor and oxygen recorded during single element hot fire tests were reduced and analyzed. A significant achievement was the attainment of single shot Raman spectra in cold flow tests. Unfortunately, the single shot signal-to-noise ratio deteriorated to an unacceptable level during the hot fire testing. Attempts to obtain temperature data from the hydrogen Q1-branch

  17. Non-intrusive measurements in a rocket engine combustor

    NASA Technical Reports Server (NTRS)

    Farhangi, S.; Gylys, V. T.; Jensen, R. J.

    1993-01-01

    In recent years analytical tools to characterize combustor flow have been developed in order to support design. To facilitate anchoring of combustion related physical models and the CFD codes in which they are incorporated, considerable development and application of non-intrusive combustion diagnostic capabilities has occurred. Raman spectroscopy can be used to simultaneously detect all polyatomic molecules present in significant concentrations and to determine gas temperature. This is because all molecules possess a distinct temperature dependent Raman spectrum. A multi-point diagnostic system for non-intrusive temperature and species profiling in rocket engines has been developed at Rocketdyne. In the present effort, the system has been undergoing validation for application to rocket engine component testing. A 4 inch diameter windowed combustor with a coaxial gas-gas injector was chosen for this series of validation experiments. Initially an excimer-pumped tunable dye laser and later a solid state Nd-Yag laser served as excitation sources. The Raman signal was dispersed by a monochromator and detected by a gated, intensified Charged Coupled Device (CCD) array. Experiments were carried out prior to each series of hot fire tests to ensure that the Raman signal detected was due to a spontaneous rather than a stimulated Raman emission process. Over sixty hot fire tests were conducted during the first series of tests with the excimer/dye laser. All hot fire testing was at a mixture ratio of 0.5 and chamber pressures of approximately 100 and approximately 300 psia. The Raman spectra of hydrogen, water vapor and oxygen recorded during single element hot fire tests were reduced and analyzed. A significant achievement was the attainment of single shot Raman spectra in cold flow tests. Unfortunately, the single shot signal-to-noise ratio deteriorated to an unacceptable level during the hot fire testing. Attempts to obtain temperature data from the hydrogen Q1-branch

  18. Characterization and Simulation of the Thermoacoustic Instability Behavior of an Advanced, Low Emissions Combustor Prototype

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Paxson, Daniel E.

    2008-01-01

    Extensive research is being done toward the development of ultra-low-emissions combustors for aircraft gas turbine engines. However, these combustors have an increased susceptibility to thermoacoustic instabilities. This type of instability was recently observed in an advanced, low emissions combustor prototype installed in a NASA Glenn Research Center test stand. The instability produces pressure oscillations that grow with increasing fuel/air ratio, preventing full power operation. The instability behavior makes the combustor a potentially useful test bed for research into active control methods for combustion instability suppression. The instability behavior was characterized by operating the combustor at various pressures, temperatures, and fuel and air flows representative of operation within an aircraft gas turbine engine. Trends in instability behavior versus operating condition have been identified and documented, and possible explanations for the trends provided. A simulation developed at NASA Glenn captures the observed instability behavior. The physics-based simulation includes the relevant physical features of the combustor and test rig, employs a Sectored 1-D approach, includes simplified reaction equations, and provides time-accurate results. A computationally efficient method is used for area transitions, which decreases run times and allows the simulation to be used for parametric studies, including control method investigations. Simulation results show that the simulation exhibits a self-starting, self-sustained combustion instability and also replicates the experimentally observed instability trends versus operating condition. Future plans are to use the simulation to investigate active control strategies to suppress combustion instabilities and then to experimentally demonstrate active instability suppression with the low emissions combustor prototype, enabling full power, stable operation.

  19. Characterization and Simulation of Thermoacoustic Instability in a Low Emissions Combustor Prototype

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Paxson, Daniel E.

    2008-01-01

    Extensive research is being done toward the development of ultra-low-emissions combustors for aircraft gas turbine engines. However, these combustors have an increased susceptibility to thermoacoustic instabilities. This type of instability was recently observed in an advanced, low emissions combustor prototype installed in a NASA Glenn Research Center test stand. The instability produces pressure oscillations that grow with increasing fuel/air ratio, preventing full power operation. The instability behavior makes the combustor a potentially useful test bed for research into active control methods for combustion instability suppression. The instability behavior was characterized by operating the combustor at various pressures, temperatures, and fuel and air flows representative of operation within an aircraft gas turbine engine. Trends in instability behavior vs. operating condition have been identified and documented. A simulation developed at NASA Glenn captures the observed instability behavior. The physics-based simulation includes the relevant physical features of the combustor and test rig, employs a Sectored 1-D approach, includes simplified reaction equations, and provides time-accurate results. A computationally efficient method is used for area transitions, which decreases run times and allows the simulation to be used for parametric studies, including control method investigations. Simulation results show that the simulation exhibits a self-starting, self-sustained combustion instability and also replicates the experimentally observed instability trends vs. operating condition. Future plans are to use the simulation to investigate active control strategies to suppress combustion instabilities and then to experimentally demonstrate active instability suppression with the low emissions combustor prototype, enabling full power, stable operation.

  20. Suppression of Combustor-Inlet Interaction in a Scramjet Engine under Mach 4 Flight Conditions

    NASA Astrophysics Data System (ADS)

    Kobayashi, Kan; Kanda, Takeshi; Tomioka, Sadatake; Tani, Kouichiro; Sakuranaka, Noboru; Mitani, Tohru

    A sidewall-compression-type scramjet engine was tested under Mach 4 flight conditions. The tested engine had an inlet, a constant cross-sectional area isolator, a constant cross-sectional area combustor, a diverging combustor, and an internal nozzle. In a previous study under the same flight conditions, the maximum thrust increment using fuel injection within the constant-area combustor was 1,380 N at an equivalence ratio of 0.31, and further fuel injection resulted in combustor-inlet interaction (designated as CII). To suppress the CII in the present study, we attempted (1) two-stage fuel injection within the constant-area combustor and the diverging combustor and (2) a boundary layer bleed on the top wall. The former was to suppress heat release around the first-stage fuel injectors in the constant-area combustor, and the latter was to decrease interaction length by decreasing the boundary layer thickness on the top wall. In the case of two-stage fuel injection, the maximum thrust increment was 2,230 N at an equivalence ratio of 0.63. In the case of the boundary layer bleed, on the other hand, the maximum thrust increment was 2,300 N at an equivalence ratio of 0.66. Thus, two-stage fuel injection and boundary layer bleed led to 62% and 67% higher maximum thrust increments than that obtained in the previous study, respectively. Finally, both the two-stage fuel injection and boundary layer bleed were applied simultaneously to obtain the best thrust performance, and the maximum thrust increment was 2,560 N at an equivalence ratio of 0.95. As a result, we obtained an 86% higher maximum thrust increment than that in the previous study. The thrust achievement factor, which was defined as the ratio of the maximum thrusts obtained from experiment and theoretical prediction, under this condition was estimated as 70%.