Science.gov

Sample records for combustors utilizing exhaust

  1. 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.

  2. Exhaust gas emissions of a vortex breakdown stabilized combustor

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  3. Combustor exhaust-emissions and blowout-limits with diesel number 2 and jet A fuels utilizing air-atomizing and pressure atomizing nozzles

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Experimental tests with diesel number 2 and Jet A fuels were conducted in a combustor segment to obtain comparative data on exhaust emissions and blowout limits. An air-atomizing nozzle was used to inject the fuels. Tests were also made with diesel number 2 fuel using a pressure-atomizing nozzle to determine the effectiveness of the air-atomizing nozzle in reducing exhaust emissions. Test conditions included fuel-air ratios of 0.008 to 0.018, inlet-air total pressures and temperatures of 41 to 203 newtons per square centimeter and 477 to 811 K, respectively, and a reference velocity of 21.3 meters per second. Smoke number and unburned hydrocarbons were twice as high with diesel number 2 as with Jet A fuel. This was attributed to diesel number 2 having a higher concentration of aromatics and lower volatility than Jet A fuel. Oxides of nitrogen, carbon monoxide, and blowout limits were approximately the same for the two fuels. The air-atomizing nozzle, as compared with the pressure-atomizing nozzle, reduced oxides-of-nitrogen by 20 percent, smoke number by 30 percent, carbon monoxide by 70 percent, and unburned hydrocarbons by 50 percent when used with diesel number 2 fuel.

  4. 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.

  5. 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.

  6. 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.

  7. TRW Advanced Slagging Coal Combustor Utility Demonstration

    SciTech Connect

    Not Available

    1989-01-01

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

  8. Odor intensity and characterization studies of exhaust from a turbojet engine combustor

    NASA Technical Reports Server (NTRS)

    Butze, H. F.; Kendall, D. A.

    1973-01-01

    Sensory odor tests of the exhaust from a turbojet combustor operating at simulated idle conditions were made by a human panel sniffing diluted exhaust gas. Simultaneously, samples of undiluted exhaust gas were collected on adsorbent substrates, subsequently removed by solvent flushing, and analyzed chemically by liquid chromatographic methods. The concentrations of the principal malodorous species, the aromatic (unburned fuel-related) and the oxygenated (partially burned fuel) fractions, as determined chromatographically, correlated well with the intensity of the odor as determined by sniffing. Odor intensity increased as combustion efficiency decreased. Combustor modifications which increased combustion efficiency decreased odor intensity.

  9. Treatment of power utilities exhaust

    DOEpatents

    Koermer, Gerald

    2012-05-15

    Provided is a process for treating nitrogen oxide-containing exhaust produced by a stationary combustion source by the catalytic reduction of nitrogen oxide in the presence of a reductant comprising hydrogen, followed by ammonia selective catalytic reduction to further reduce the nitrogen oxide level in the exhaust.

  10. Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

    PubMed Central

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

  11. Ignition and flame stabilization of a strut-jet RBCC combustor with small rocket exhaust.

    PubMed

    Hu, Jichao; Chang, Juntao; Bao, Wen

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

  12. Effects of prevaporized fuel on exhaust emissions of an experimental gas turbine combustor

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    Effects of fuel vaporization on the exhaust emission levels of oxides of nitrogen (NOX), carbon monoxide, total hydrocarbons, and smoke number were obtained in an experimental turbojet combustor segment. Two fuel injector types were used in which liquid ASTM A-1 jet fuel and vapor propane fuel were independently controlled to simulate varying degrees of vaporization. Tests were conducted over a range of inlet-air temperatures from 478 to 700 K (860 to 1260 R), pressures from 4 to 20 atmospheres, and combustor reference velocities from 15.3 to 27.4 m/sec (50 to 90 ft/sec). Converting from liquid to complete vapor fuel resulted in NOX reductions as much as 22 percent and smoke number reductions up to 51 percent.

  13. A Gas Chromatograph/Mass Spectrometer System for UltraLow-Emission Combustor Exhaust Studies

    NASA Technical Reports Server (NTRS)

    Brabbs, Theodore A.; Wey, Chowen Chou

    1996-01-01

    A gas chromatograph (GC)/mass spectrometer (MS) system that allows the speciation of unburnt hydrocarbons in the combustor exhaust has been developed at the NASA Lewis Research Center. Combustion gas samples are withdrawn through a water-cooled sampling probe which, when not in use, is protected from contamination by a high-pressure nitrogen purge. The sample line and its connecting lines, filters, and valves are all ultraclean and are heated to avoid condensation. The system has resolution to the parts-per-billion (ppb) level.

  14. Detection of very large ions in aircraft gas turbine engine combustor exhaust: charged small soot particles?

    NASA Astrophysics Data System (ADS)

    Wilhelm, S.; Haverkamp, H.; Sorokin, A.; Arnold, F.

    Small electrically charged soot particles (CSP) present in the exhaust of a jet aircraft engine combustor have been detected by a Large Ion Mass Spectrometer and quantitatively measured by an Ion Mobility Analyzer. The size and concentration measurements which took place at an aircraft gas-turbine engine combustor test-rig at the ground covered different combustor conditions (fuel flow=FF, fuel sulphur content=FSC). At the high-pressure turbine stage of the engine, CSP-diameters were mostly around 6 nm and CSP-concentrations reached up to 4.8×10 7 cm -3 (positive and negative) corresponding to a CSP-emission index ECSP=2.5×10 15 CSP kg -1 fuel burnt. The ECSP increased with FF but did not increase with FSC. The latter indicates that sulphur was not a major component of the large ions. Possible CSP-sources and CSP-sinks as well as CSP-roles are discussed.

  15. Effect of primary-zone equivalence ratio and hydrogen addition on exhaust emission in a hydrocarbon-fueled combustor

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

    The effects of reducing the primary-zone equivalence ratio on the exhaust emission levels of oxides of nitrogen, carbon monoxide, and unburned hydrocarbons in experimental hydrocarbon-fueled combustor segments at simulated supersonic cruise and idle conditions were investigated. In addition, the effects of the injection of hydrogen fuel (up to 4 percent of the total weight of fuel) on the stability of the hydrocarbon flame and exhaust emissions were studied and compared with results obtained without hydrogen addition.

  16. Effect of operating conditions on the exhaust emissions from a gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Briehl, D.; Papathakos, L.; Strancar, R. J.

    1972-01-01

    Exhaust concentrations of total unburned hydrocarbons, carbon monoxide, and nitric oxide were measured from a single J-57 combustor liner installed in a 30 diameter test section. Tests were conducted over a range of inlet total pressures from 1 to 20 atmospheres, inlet total temperatures from 310 to 590 K, reference velocities from 8 to m/sec, and fuel-air ratios from 0.004 to 0.015. Most of the data were obtained using ASTM A-1 fuel; however, a limited number of tests was performed with natural gas fuel. Combustion efficiency and emission levels are correlated with operating conditions. Sampling error at operating conditions for which combustion efficiency was below about 90 percent resulted in abnormally low readings for hydrocarbon emissions.

  17. 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...

  18. Combustor exhaust emissions with air-atomizing splash-groove fuel injectors burning Jet A and Diesel number 2 fuels

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Air-atomizing, splash-groove injectors were shown to improve primary-zone fuel spreading and reduce combustor exhaust emissions for Jet A and diesel number 2 fuels. With Jet A fuel large-orifice, splash-groove injectors the oxides-of-nitrogen emission index was reduced, but emissions of carbon monoxide, unburned hydrocarbons, or smoke were unaffected. Small-orifice, splash-groove injectors did not reduce oxides of nitrogen, but reduced the smoke number and carbon monoxide and unburned-hydrocarbon emission indices. With diesel number 2 fuel, the small-orifice, splash-groove injectors reduced oxides of nitrogen by 19 percent, smoke number by 28 percent, carbon monoxide by 75 percent, and unburned hydrocarbons by 50 percent. Smoke number and unburned hydrocarbons were twice as high with diesel number 2 as with Jet A fuel. Combustor blowout limits were similar for diesel number 2 and Jet A fuels.

  19. Exhaust pollutant emissions from swirl-can combustor module arrays at parametric test conditions

    NASA Technical Reports Server (NTRS)

    Mularz, E. J.; Wear, J. D.; Verbulecz, P. W.

    1975-01-01

    Improved designs of swirl-can combustor modules were tested using seven-module arrays in a combustor. The combustor was operated over a pressure range of 69 to 207 N/sq cm, a fuel-air ratio range of 0.015 to 0.046, at a constant inlet air temperature of 733 K, and at reference velocities of 23.9 and 30.6 m/sec. The three designs tested performed with high combustion efficiency at all conditions tested and exhibited oxides of nitrogen emissions substantially lower than that of conventional gas turbine combustors. A correlating parameter used to extrapolate oxides of nitrogen emissions to full power or takeoff conditions for large commercial turbofan engines predicts oxides of nitrogen emissions somewhat higher than those specified in the 1979 government emissions standards.

  20. TRW Advanced Slagging Coal Combustor Utility Demonstration. Fourth Quarterly progress report, August 1989--October 1989

    SciTech Connect

    Not Available

    1989-12-31

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

  1. Effect of exhaust gas recirculation on emissions from a flame-tube combustor using Liquid Jet A fuel

    NASA Technical Reports Server (NTRS)

    Marek, C. J.; Tacina, R. R.

    1976-01-01

    The effects of uncooled exhaust gas recirculation as an inert diluent on emissions of oxides of nitrogen (NO + NO2) and on combustion efficiency were investigated. Ratios of recirculated combustion products to inlet airflow were varied from 10 to 80 percent by using an inlet air ejector nozzle. Liquid Jet A fuel was used. The flame-tube combustor was 10.2 cm in diameter. It was operated with and without a flameholder present. The combustor pressure was maintained constant at 0.5 MPa. The equivalence ratio was varied from 0.3 to 1.0. The inlet air temperature was varied from 590 to 800 K, and the reference velocity from 10 to 30 m/sec. Increasing the percent recirculation from 10 to 25 had the following effects: (1) the peak NOx emission was decreased by 37 percent, from 8 to 5 g NO2/kg fuel, at an inlet air temperature of 590 K and a reference velocity of 15 m/sec; (2) the combustion efficiency was increased, particularly at the higher equivalence ratios; and (3) for a high combustion efficiency of greater than 99.5 percent, the range of operation of the combustor was nearly doubled in terms of equivalence ratio. Increasing the recirculation from 25 to 50 percent did not change the emissions significantly.

  2. Experimental clean combustor program, phase 1. [aircraft exhaust/gas analysis - gas turbine engines

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    A program of screening three low emission combustors for conventional takeoff and landing, by testing and analyzing thirty-two configurations is presented. Configurations were tested that met the emission goals at idle operating conditions for carbon monoxide and for unburned hydrocarbons (emission index values of 20 and 4, respectively). Configurations were also tested that met a smoke number goal of 15 at sea-level take-off conditions. None of the configurations met the goal for oxides of nitrogen emissions at sea-level take-off conditions. The best configurations demonstrated oxide of nitrogen emission levels that were approximately 61 percent lower than those produced by the JT9D-7 engine, but these levels were still approximately 24 percent above the goal of an emission index level of 10. Additional combustor performance characteristics, including lean blowout, exit temperature pattern factor and radial profile, pressure loss, altitude stability, and altitude relight characteristics were documented. The results indicate the need for significant improvement in the altitude stability and relight characteristics. In addition to the basic program for current aircraft engine combustors, seventeen combustor configurations were evaluated for advanced supersonic technology applications. The configurations were tested at cruise conditions, and a conceptual design was evolved.

  3. Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J. (Inventor); Yankowich, Paul R. (Inventor)

    2006-01-01

    A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.

  4. Utilization of LPG and gasoline engine exhaust emissions by microalgae.

    PubMed

    Taştan, Burcu Ertit; Duygu, Ergin; Ilbaş, Mustafa; Dönmez, Gönül

    2013-02-15

    The effect of engine exhaust emissions on air pollution is one of the greatest problems that the world is facing today. The study focused on the effects of realistic levels of engine exhaust emissions of liquid petroleum gas (LPG) and gasoline (GSN) on Phormidium sp. and Chlorella sp. Multi parameters including pH, different medial compositions, fuel types, flow rates and biomass concentrations were described in detail. Effects of some growth factors such as triacontanol (TRIA) and salicylic acid (SA) have also been tested. The maximum biomass concentration of Phormidium sp. reached after 15 days at 0.36 and 0.15 g/L initial biomass concentrations were found as 1.160 g/L for LPG emission treated cultures and 1.331 g/L for GSN emission treated cultures, respectively. The corresponding figures were 1.478 g/L for LPG emission treated cultures and 1.636 g/L for GSN emission treated cultures at 0.65 and 0.36 g/L initial Chlorella sp. biomass concentrations. This study highlights the significance of using Phormidium sp. and Chlorella sp. for utilization of LPG and GSN engine exhaust emissions by the help of growth factors. PMID:23298742

  5. 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.

  6. 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.

  7. Nitration of polynuclear aromatic hydrocarbons in coal combustors and exhaust streams

    SciTech Connect

    Yu, L.; Dadamio, J.; Hildemann, L.; Niska, S.

    1993-02-01

    Our efforts this quarter were directed at preparing PAH samples at well-controlled extents of primary devolatilization. In the long (12.5 cm) hot zone, Preliminary studies with a Pit. No. 8 hvA bituminous coal showed that a furnace temperature of 1380 K yields the best resolution of primary devolatilization from secondary pyrolysis. There is no soot at all in the PAH samples collected during the first half of tar evolution. Although some soot is present in the total sample of all tar from primary devolatilization it never amounts to more than 10 % of the aerosol and is usually less than 5 %. Also, the maximum weight loss observed with the new operating conditions is nearly 50 wt. % daf, which compares favorably to our previous value of 55 wt. % for this coal. Regarding PAH analysis during this quarter, the assembly of the gravity-flow column chromatographic system for prefractionating the coal tar was completed. Standard compounds spanning a wide range of polarities and molecular weights were selected and purchased. These compounds were utilized to begin testing and refining the prefractionation procedure.

  8. A hardened CARS system utilized for temperature measurements in a supersonic combustor

    NASA Technical Reports Server (NTRS)

    Antcliff, Richard R.; Smith, Michael W.; Jarrett, Olin, Jr.; Northam, G. Burton; Cutler, Andrew D.

    1991-01-01

    A coherent anti-Stokes Raman scattering (CARS) system has been hardened for utilization in a NASA Langley supersonic combustion test cell that can obtain temperature cross sections of the flow at three locations. This system is remotely operated and environmentally protected. Measurements were obtained in a scramjet combustor model consisting of a rearward-facing step, followed by an expansion duct. The freestream conditions were Mach 2, with static pressure that ranged from 0.8 to 1.9 atm, and a static temperature of about 800 K. Data for two different flow conditions were obtained that provided a comparison between reacting and nonreacting mixing of injected hydrogen fuel with the combustion-heated supersonic stream.

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

    SciTech Connect

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

    1994-07-01

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

  10. 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.

  11. A Hardened CARS System Utilized for Temperature Measurements in a Supersonic Combustor

    NASA Technical Reports Server (NTRS)

    Antcliff, Richard R.; Smith, Michael W.; Jarret, Olin, Jr.; Northam, G. Burton; Cutler, Andrew D.; Taylor, David J.

    1990-01-01

    A coherent anti-Stokes Raman scattering (CARS) system has been hardened for use in a NASA Langley supersonic combustion test cell. The system can obtain temperature cross sections of the flow at three locations. The system is environmentally protected and remotely operated. Measurements were made in a scram-jet combustor model consisting of a rear- ward-facing step, followed by an expansion duct. The duct is nominally 4 feet in length. The free stream conditions were Mach 2, with static pressure which ranged from 0.8 to 1.9 atm, and a static temperature of approximately 800K. Three vertical slots were machined into each side of the duct to allow optical access. The CARS system utilized a planar BOXCARS beam arrangement. This arrangement allowed the laser beams to pass through the vertical slots in the tunnel. Translation stages were utilized to move the focussing volume within the tunnel. These stages allowed complete cross sections to be obtained at each slot location. A fiber optic carried the signal to a remotely located monochrometer and reticon detector.Data for two different flow conditions were taken at each of the three slot locations. These two conditions provided a comparison between reacting and non-reacting mixing of injected hydrogen fuel with the combustion heated supersonic stream.

  12. 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.

  13. High-pressure combustor exhaust emissions with improved air-atomizing and conventional pressure-atomizing fuel nozzles

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    A high-pressure combustor segment 0.456 meter (18 in.) long with a maximum cross section of 0.153 by 0.305 meter (6 by 12 in.) was tested with specially designed air-atomizing and conventional pressure-atomizing fuel nozzles at inlet-air temperatures of 340 to 755 k (610 deg to 1360 R), reference velocities of 12.4 to 26.1 meters per second (41 to 86 ft/sec), and fuel-air ratios of 0.008 to 0.020. Increasing inlet-air pressure from 4 to 20 atmospheres generally increased smoke number and nitric oxide, but decreased carbon monoxide and unburned hydrocarbon concentrations with air-atomizing and pressure-atomizing nozzles. Emission indexes for carbon monoxide and unburned hydrocarbons were lower at 4, 10, and 20 atmospheres, and nitric oxide emission indexes were lower at 10 and 20 atmospheres with air-atomizing than with pressure-atomizing nozzles.

  14. Nitration of polynuclear aromatic hydrocarbons in coal combustors and exhaust streams: Final report, September 1, 1991--September 30, 1994

    SciTech Connect

    Yu, L.; Cho, S.; Hildemann, L.; Niksa, S.

    1995-02-01

    The objectives of this three-year project were to (1) identify the conditions which promote the nitration of PAH during primary combustion, reburning, hot gas cleanup, and particulate removal; and (2) investigate the potential relationship between NOx abatement and PAH nitration. Meeting the objectives of this program involved two broad tasks: (1) Preparing the polynuclear aromatic hydrocarbons (PAH) under closely monitored pulverized fuel (p. f.) firing conditions; and, (2) analyzing the PAH samples to monitor extents of nitration, ring number distribution, etc. A novel coal flow reactor burning actual coal products that operates over the domains of heating rates, temperatures, fuel-equivalence ratios, and residence times in utility boilers was used to generate the coal tar samples. The distribution of products obtained from primary, secondary, and oxidative pyrolysis of two coal types, Pittsburgh No. 8 and Dietz, were analyzed, with emphasis on the nitrogen-containing species generated. The coal tax samples collected from the coal flow reactor were fractionated based on their size and polarity using gravity flow column chromatography. After examining how the sample fractionation depended on the coal type and pyrolysis conditions, the relatively nonpolar fraction was further analyzed via high performance liquid chromatography, to characterize the ring number distribution of the polycyclic aromatic compounds (PAC) present. Finally, gas chromatographic techniques were utilized to measure the amount of nitrogen-containing PAC present, and to investigate how much of these nitrogen-containing species consist of nitro-PAH.

  15. Carbohydrate supercompensation and muscle glycogen utilization during exhaustive running in highly trained athletes.

    PubMed

    Madsen, K; Pedersen, P K; Rose, P; Richter, E A

    1990-01-01

    Three female and three male highly trained endurance runners with mean maximal oxygen uptake (VO2max) values of 60.5 and 71.5 ml.kg-1.min-1, respectively, ran to exhaustion at 75%-80% of VO2max on two occasions after an overnight fast. One experiment was performed after a normal diet and training regimen (Norm), the other after a diet and training programme intended to increase muscle glycogen levels (Carb). Muscle glycogen concentration in the gastrocnemius muscle increased by 25% (P less than 0.05) from 581 mmol.kg-1 dry weight, SEM 50 to 722 mmol.kg-1 dry weight, SEM 34 after Carb. Running time to exhaustion, however, was not significantly different in Carb and Norm, 77 min, SEM 13 vs 70 min, SEM 8, respectively. The average glycogen concentration following exhaustive running was 553 mmol.kg-1 dry weight, SEM 70 in Carb and 434 mmol.kg-1 dry weight, SEM 57 in Norm, indicating that in both tests muscle glycogen stores were decreased by about 25%. Periodic acid-Schiff staining for semi-quantitative glycogen determination in individual fibres confirmed that none of the fibres appeared to be glycogen-empty after exhaustive running. The steady-state respiratory exchange ratio was higher in Carb than in Norm (0.92, SEM 0.01 vs 0.89, SEM 0.01; P less than 0.05). Since muscle glycogen utilization was identical in the two tests, the indication of higher utilization of total carbohydrate appears to be related to a higher utilization of liver glycogen. We have concluded that glycogen depletion of the gastrocnemius muscle is unlikely to be the cause of fatigue during exhaustive running at 75%-80% of VO2max in highly trained endurance runners. Furthermore, diet- and training-induced carbohydrate super-compensation does not appear to improve endurance capacity in such individuals. PMID:2079068

  16. Nitration of polynuclear aromatic hydrocarbons in coal combustors and exhaust streams. Quarterly report, April 1, 1992--June 30, 1992

    SciTech Connect

    Yu, L.; Dadamio, J.; Hildemann, L.; Niska, S.

    1993-02-01

    Our efforts this quarter were directed at preparing PAH samples at well-controlled extents of primary devolatilization. In the long (12.5 cm) hot zone, Preliminary studies with a Pit. No. 8 hvA bituminous coal showed that a furnace temperature of 1380 K yields the best resolution of primary devolatilization from secondary pyrolysis. There is no soot at all in the PAH samples collected during the first half of tar evolution. Although some soot is present in the total sample of all tar from primary devolatilization it never amounts to more than 10 % of the aerosol and is usually less than 5 %. Also, the maximum weight loss observed with the new operating conditions is nearly 50 wt. % daf, which compares favorably to our previous value of 55 wt. % for this coal. Regarding PAH analysis during this quarter, the assembly of the gravity-flow column chromatographic system for prefractionating the coal tar was completed. Standard compounds spanning a wide range of polarities and molecular weights were selected and purchased. These compounds were utilized to begin testing and refining the prefractionation procedure.

  17. Fuel cell system with combustor-heated reformer

    DOEpatents

    Pettit, William Henry

    2000-01-01

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

  18. 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.

  19. 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.

  20. Direct heating surface combustor

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  1. Effect of flame stabilizer design on performance and exhaust pollutants of a two-row 72-module swirl-can combustor

    NASA Technical Reports Server (NTRS)

    Biaglow, J. A.; Trout, A. M.

    1976-01-01

    A test program was conducted to evaluate the effects of four flame stabilizer designs on the performance and gaseous pollutant levels of an experimental full-annular swirl-can combustor. Combustor operating parameters, including inlet-air temperature, reference velocity, and fuel-air ratio, were set to simulate conditions in a 30:1 pressure ratio engine. Combustor inlet total pressure was held constant at 6 atm due to the facility limit. Combustor performance and gaseous pollutant levels were strongly affected by the geometry and resulting total pressure loss of the four flame stabilizer designs investigated. The addition of shrouds to two designs produced an 18 to 22% decrease in the combustion chamber pressure loss and thus resulted in doubling the exit temperature pattern factor and up to 42% higher levels of oxides of nitrogen. A previously developed oxides of nitrogen correlating parameter agreed with each model within an emission index of plus or minus 1 but was not capable of correlating all models together.

  2. Effect of Flame Stabilizer Design on Performance and Exhaust Pollutants of a Two-Row Swirl-Can Combustor Operated to Near-Stoichiometric Conditions

    NASA Technical Reports Server (NTRS)

    Biaglow, James A.; Trout, Arthur M.

    1977-01-01

    Emissions and performance characteristics were determined for two full annulus modular combustors operated to near stoichiometric fuel air ratios. The tests were conducted to obtain stoichiometric data at inlet air temperatures from 756 to 894 K and to determine the effects of a flat plate circular flame stabilizer with upstream fuel injection and a contraswirl flame stabilizer with downstream fuel injection. Levels of unburned hydrocarbons were below 0.50 gram per kilogram of fuel for both combustors and thus there was no detectable difference in the two methods of fuel injection. The contraswirl flame stabilizer did not produce the level of mixing obtained with a flat plate circular flame stabilizer. It did produce higher levels of oxides of nitrogen, which peaked at a fuel air ratio of 0.037. For the flat plate circular flame stabilizer, oxides of nitrogen emission levels were still increasing with fuel air ratio to the maximum tested value of 0.045.

  3. Electrically charged small soot particles in the exhaust of an aircraft gas-turbine engine combustor: comparison of model and experiment

    NASA Astrophysics Data System (ADS)

    Sorokin, A.; Arnold, F.

    The emission of electrically charged soot particles by an aircraft gas-turbine combustor is investigated using a theoretical model. Particular emphasis is placed on the influence of the fuel sulfur content (FSC). The model considers the production of primary "combustion" electrons and ions in the flame zone and their following interaction with molecular oxygen, sulfur-bearing molecules (e.g. O 2, SO 2, SO 3, etc.) and soot particles. The soot particle size distribution is approximated by two different populations of mono-dispersed large and small soot particles with diameters of 20-30 and 5-7 nm, respectively. The effect of thermal ionization of soot and its interaction with electrons and positive and negative ions is included in the model. The computed positive and negative chemiion (CI) concentrations at the combustor exit and relative fractions of small neutral and charged soot particles were found to be in satisfactory agreement with experimental data. The results show that the FSC indeed may influence the concentration of negative CI at low fuel flow into combustor. Importantly the simulation indicates a very efficient mutual interaction of electrons and ions with soot particles with a large effect on both ion and charged soot particle concentrations. This result may be interpreted as a possible indirect effect of FSC on the growth and size distribution of soot particles.

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

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1975-01-01

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

  5. Energy conservation in fruit dehydrators utilizing recirculation of exhaust air and heat-recovery heat exchangers. Final report

    SciTech Connect

    Groh, J.E.; Thompson, T.L.

    1981-12-01

    Dehydration of fruit in the United States is often done by means of a tunnel dehydrator utilizing large quantities of fossil fuel. Existing dehydrators have been designed to operate with maximum product through-put and with little regard for energy efficiency. By incorporating dampers for air recirculation and thermal energy recovery equipment on the exhaust air, the energy required in dehydration was cut by over 40%, satisfying the original objectives of the program. A commercial dehydrator tunnel was modified by installing a heat recovery heat exchanger and an exhaust air recirculation damper. Another tunnel was equipped with the exhaust air recirculation damper only. A third tunnel was unmodified. These three tunnels of a 24 tunnel facility were equipped with individual natural gas meters to measure energy consumption. The energy consumption of the heat exchanger equipped tunnel normally amounted to approximately 40% of the unmodified tunnel during raisin production.

  6. 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.

  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. Segmented combustor

    NASA Technical Reports Server (NTRS)

    Halila, Ely E. (Inventor)

    1994-01-01

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

  9. 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.

  10. 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.

  11. Pollutant emissions from and within a model gas turbine combustor at elevated pressures and temperatures

    NASA Technical Reports Server (NTRS)

    Drennan, S. A.; Peterson, C. O.; Khatib, F. M.; Sowa, W. A.; Samuelsen, G. S.

    1993-01-01

    Conventional and advanced gas turbine engines are coming under increased scrutiny regarding pollutant emissions. This, in turn, has created a need to obtain in-situ experimental data at practical conditions, as well as exhaust data, and to obtain the data in combustors that reflect modern designs. The in-situ data are needed to (1) assess the effects of design modifications on pollutant formation, and (2) develop a detailed data base on combustor performance for the development and verification of computer modeling. This paper reports on a novel high pressure, high temperature facility designed to acquire such data under controlled conditions and with access (optical and extractive) for in-situ measurements. To evaluate the utility of the facility, a model gas turbine combustor was selected which features practical hardware design, two rows of jets (primary and dilution) with four jets in each row, and advanced wall cooling techniques with laser drilled effusive holes. The dome is equipped with a flat-vaned swirler with vane angles of 60 degrees. Data are obtained at combustor pressures ranging from 2 to 10 atmospheres of pressure, levels of air preheat to 427 C, combustor reference velocities from 10.0 to 20.0 m/s, and an overall equivalence ratio of 0.3. Exit plane and in-situ measurements are presented for HC, O2, CO2, CO, and NO(x). The exit plane emissions of NO(x) correspond to levels reported from practical combustors and the in-situ data demonstrate the utility and potential for detailed flow field measurements.

  12. Evaluation of Energy Saving Characteristics of a High-Efficient Cogeneration System Utilizing Gas Engine Exhaust Heat

    NASA Astrophysics Data System (ADS)

    Pak, Pyong Sik

    A high efficiency cogeneration system (CGS) utilizing high temperature exhaust gas from a gas engine is proposed. In the proposed CGS, saturated steam produced in the gas engine is superheated with a super heater utilizing regenerative burner and used to drive a steam turbine generator. The heat energy is supplied by extracting steam from the steam turbine and turbine outlet low-temperature steam. Both of the energy saving characteristics of the proposed CGS and a CGS constructed by using the original gas engine (GE-CGS) were investigated and compared, by taking a case where energy for office buildings was supplied by the conventional energy systems. It was shown that the proposed CGS has energy saving rate of 24.5%, higher than 1.83 times, compared with that of the original GE-CGS.

  13. 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.

  14. Platform for a Hydrocarbon Exhaust Gas Sensor Utilizing a Pumping Cell and a Conductometric Sensor

    PubMed Central

    Biskupski, Diana; Geupel, Andrea; Wiesner, Kerstin; Fleischer, Maximilian; Moos, Ralf

    2009-01-01

    Very often, high-temperature operated gas sensors are cross-sensitive to oxygen and/or they cannot be operated in oxygen-deficient (rich) atmospheres. For instance, some metal oxides like Ga2O3 or doped SrTiO3 are excellent materials for conductometric hydrocarbon detection in the rough atmosphere of automotive exhausts, but have to be operated preferably at a constant oxygen concentration. We propose a modular sensor platform that combines a conductometric two-sensor-setup with an electrochemical pumping cell made of YSZ to establish a constant oxygen concentration in the ambient of the conductometric sensor film. In this paper, the platform is introduced, the two-sensor-setup is integrated into this new design, and sensing performance is characterized. Such a platform can be used for other sensor principles as well. PMID:22423212

  15. Platform for a hydrocarbon exhaust gas sensor utilizing a pumping cell and a conductometric sensor.

    PubMed

    Biskupski, Diana; Geupel, Andrea; Wiesner, Kerstin; Fleischer, Maximilian; Moos, Ralf

    2009-01-01

    Very often, high-temperature operated gas sensors are cross-sensitive to oxygen and/or they cannot be operated in oxygen-deficient (rich) atmospheres. For instance, some metal oxides like Ga(2)O(3) or doped SrTiO(3) are excellent materials for conductometric hydrocarbon detection in the rough atmosphere of automotive exhausts, but have to be operated preferably at a constant oxygen concentration. We propose a modular sensor platform that combines a conductometric two-sensor-setup with an electrochemical pumping cell made of YSZ to establish a constant oxygen concentration in the ambient of the conductometric sensor film. In this paper, the platform is introduced, the two-sensor-setup is integrated into this new design, and sensing performance is characterized. Such a platform can be used for other sensor principles as well. PMID:22423212

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

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  17. Studies on Effective Utilization of SOFC Exhaust Heat Using Thermoelectric Power Generation Technology

    NASA Astrophysics Data System (ADS)

    Terayama, Takeshi; Nagata, Susumu; Tanaka, Yohei; Momma, Akihiko; Kato, Tohru; Kunii, Masaru; Yamamoto, Atsushi

    2013-07-01

    Solid oxide fuel cells (SOFCs) are being researched around the world. In Japan, a compact SOFC system with rated alternative current (AC) power of 700 W has become available on the market, since the base load electricity demand for a standard home is said to be less than 700 W AC. To improve the generating efficiency of SOFC systems in the 700-W class, we focused on thermoelectric generation (TEG) technology, since there are a lot of temperature gradients in the system. Analysis based on simulations indicated the possibility of introducing thermoelectric generation at the air preheater, steam generator, and exhaust outlet. Among these options, incorporating a TEG heat exchanger comprising multiple CoSb3/SiGe-based TEG modules into the air preheater had potential to produce additional output of 37.5 W and an improvement in generating efficiency from 46% to 48.5%. Furthermore, by introducing thermoelectric generation at the other two locations, an increase in maximum output of more than 50 W and generating efficiency of 50% can be anticipated.

  18. 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.

  19. 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.

  20. Power plant including an exhaust gas recirculation system for injecting recirculated exhaust gases in the fuel and compressed air of a gas turbine engine

    DOEpatents

    Anand, Ashok Kumar; Nagarjuna Reddy, Thirumala Reddy; Shaffer, Jason Brian; York, William David

    2014-05-13

    A power plant is provided and includes a gas turbine engine having a combustor in which compressed gas and fuel are mixed and combusted, first and second supply lines respectively coupled to the combustor and respectively configured to supply the compressed gas and the fuel to the combustor and an exhaust gas recirculation (EGR) system to re-circulate exhaust gas produced by the gas turbine engine toward the combustor. The EGR system is coupled to the first and second supply lines and configured to combine first and second portions of the re-circulated exhaust gas with the compressed gas and the fuel at the first and second supply lines, respectively.

  1. 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.

  2. 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.

  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

    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.

  4. 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.

  5. Enhancing understanding of the operation of the dynamic containment combustor through CFD modelling

    SciTech Connect

    Lanicek, L.; Jicha, M.; Alizadeh, S.; Strykowski, P.J.

    1999-07-01

    The dynamic Containment Combustor (DCC) is an Ultra Low NOx burner, consisting of two counter-flowing, co-swirling, annular feeds of premixed gaseous methane-air mixture into a cylindrical chamber. The exhaust is located on the axis, at one end of the combustor. A cylindrical flame with a radius about the size of the exhaust port is established along the length of the chamber allowing the combustor walls to operate at cool conditions. In order to provide Ultra Low NOx operation with suitable turn down ratios, stability of operation and high combustion efficiencies, in the absence of detailed experimental measurements, CFD simulation of a simplified 2D model has been utilized to enhance understanding of the behavior of the combustor operation. A suitable choice of combustion model constants were made. Based on nominal design conditions of 0.03kg/s mass flow rate, equivalence ratio of 0.8 and rear mass fraction of 10%, extensive parametric tests were conducted. The effect of equivalence ratio (0.4--1.1), rear flow split (7--14%), front flow angle (40deg--70deg) and rear flow angle (10deg--80deg) on burner performance characteristics are reported here. The simulations show that with appropriate adjustments to the constants of the Eddy-Dissipation model, some of the main flow field features, like the position of the flame sheet and the expected shapes of axial and radial velocity profiles have been successfully predicted. The simulations have also highlighted the disadvantage by using a fast chemistry combustion model. Future work needs to address this issue by using a combustion model which includes chemical kinetic effects.

  6. Evaluation of pilot-scale pulse-corona-induced plasma device to remove NO{sub x} from combustion exhausts from a subscale combustor and from a hush house at Nellis AFB, Nevada. Final report, August 1994--January 1997

    SciTech Connect

    Haythornthwaite, S.M.; Durham, M.D.; Anderson, G.L.; Rugg, D.E.

    1997-05-01

    Jet engine test cells (JETCs) are used to test-fire new, installed, and reworked jet engines. Because JETCs have been classified as stationary sources of pollutant emissions, they are subject to possible regulation under Title 1 of the Clean Air Act (CAA) as amended in 1990. In Phase 1 of the Small Business Innovation Research (SBIR) program, a novel NOx-control approach utilizing pulsed-corona-induced plasma successfully showed 90% removal of NOx in the laboratory. The objective of Phase 2 was to reproduce the laboratory-scale results in a pilot-scale system. The technology was successfully demonstrated at pilot scale in the field, on a slipstream of JETC flue gas at Nellis Air Force Base. Based on the field data, cost projections were made for a system to treat the full JETC exhaust. The technology efficiently converted NO into ONO, and a wet scrubber was required to achieve the treatment goal of 50-percent removal and destruction of NOx. The plasma simultaneously removes hydrocarbons from the flue gas stream. This project demonstrated that pulse-corona-induced plasma technology is scalable to practical industrial dimensions.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  8. 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.

  9. 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.

  10. 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.

  11. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  13. 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.

  14. 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.

  15. 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.

  16. Investigation of the mechanism in Rijke pulse combustors with tangential air and fuel injection. Final report

    SciTech Connect

    Zinn, B.T.; Jagoda, J.I.; Daniel, B.R.; Bai, T.

    1993-03-01

    To study the mechanisms that control the operation of this combustor, an experimental setup is developed with access for detailed optical measurements. Propane is employed as fuel because the absence of liquid drops and combustion generated particulates in the combustion region significantly simplifies the optical diagnostics. The experimental techniques utilized include acoustic pressure measurements, space and time resolved radiation measurements, steady temperature measurements, exhaust flow chemical analysis, high speed video and intensified images of the reacting flow field by a computer based CCD camera imaging system. Flow visualization by the imaging system and the results from radiation intensity distribution measurements suggest that the periodic combustion processes caused by periodic vortex shedding and impingement provide the energy required to sustain the pressure oscillations. High radiation intensity occurs during a relatively short period of time and is in phase with the pressure oscillations, indicating that Rayleigh`s criterion is satisfied. Periodic variations of the air and fuel flow rates and, consequently, the air/fuel ratio of the reacting mixture inside the combustor appear to be another mechanism that contributes to the occurrence of periodic combustion and heat release processes. The presence of this mechanism has been uncovered by acoustic pressure measurements that revealed the presence of traveling pressure waves inside the air and fuel feed lines. These traveling waves produce periodic fuel and air feed rates which, in turn, result in periodic combustion and heat release processes within the combustor.

  17. 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.

  18. 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.

  19. 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.

  20. Experimental clean combustor program: Diesel no. 2 fuel addendum, phase 3

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    A CF6-50 engine equipped with an advanced, low emission, double annular combustor was operated 4.8 hours with No. 2 diesel fuel. Fourteen steady-state operating conditions ranging from idle to full power were investigated. Engine/combustor performance and exhaust emissions were obtained and compared to JF-5 fueled test results. With one exception, fuel effects were very small and in agreement with previously obtained combustor test rig results. At high power operating condition, the two fuels produced virtually the same peak metal temperatures and exhaust emission levels. At low power operating conditions, where only the pilot stage was fueled, smoke levels tended to be significantly higher with No. 2 diesel fuel. Additional development of this combustor concept is needed in the areas of exit temperature distribution, engine fuel control, and exhaust emission levels before it can be considered for production engine use.

  1. 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)

  2. 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.

  3. SLUDGE COMBUSTOR USING SWIRL AND ACTIVE COMBUSTION CONTROL

    EPA Science Inventory

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

  4. 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.

  5. 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.

  6. 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.

  7. Pollution technology program, can-annular combustor engines

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  8. 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.

  9. 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.

  10. Multifuel evaluation of rich/quench/lean combustor

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. Dust control considerations for deep-cut mining when utilizing exhaust ventilation and a scrubber. Report of investigations/1996

    SciTech Connect

    Colinet, J.F.; Jankowski, R.A.

    1996-05-01

    The U.S. Bureau of Mines conducted a series of laboratory tests to investigate the effectiveness of using a flooded-bed scrubber with exhaust ventilation in deep-cut faces of up to 12.2 m in length. An experimental test program to determine the impact on respirable dust levels resulting from changes in face airflow, curtain setback distance, operator positioning, and operating parameters of the external spray system on the miner was completed. Gravimetric sampling was conducted in the immediate return and at three sampling locations on the off-curtain side of the entry. Statistically significant differences in dust levels on the order of 0.5 to 1.2 mg/cu m were observed between specific sampling locations and changes in several test parameters. Several of the statistically significant relationships were found at the inby operator position, which is the least desirable of the operator locations that were tested.

  15. Combustor burner vanelets

    DOEpatents

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

    2012-02-14

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

  16. 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.

  17. 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.

  18. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Exhaust system. 23.1123 Section 23.1123... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1123 Exhaust system. (a) Each exhaust system must be fireproof and corrosion-resistant, and must have means...

  19. Development of the control and ignition systems on a high pressure gas turbine combustor

    NASA Astrophysics Data System (ADS)

    Valdez, Carlos Alejandro

    The ignition and control systems of a laboratory scale high-pressure gas turbine combustor were developed in the present work. This work provides a detailed description of the design, development and testing of the remote control system developed for a High Pressure Gas Turbine Combustor (HPTC). The combustor has the capability to operate at pressures up to 1.5 MPa and temperatures up to 2400 K. It is also designed for a maximum air and fuel flow rates of 81.93 g/s and 35.77 g/s respectively. The fuel used will be CH4 for the early experiments but it is designed to operate using a mixture of H2-CO with a hydrogen fuel composition variation of up to 30 percent. The HPTC also has optical accessibility capabilities in its combustion chamber with a converging nozzle that restricts the exhaust flow. It also has three circular ports that can be used as instrumentation ports to obtain real time data from the combustion chamber. LabVIEW was used as the controlling interface for the user. A detailed outline of the LabVIEW programming is also described. LabVIEW controlled the proportional valves (ball valves), and solenoid valves; it also provided the user with data from mass flow meters as well as pressure transducers. Both proportional and solenoid valves are 1.91 cm and can withstand pressures of up to 1551 kPa. Thermal mass flow meters were used to obtain the flow in the lines with a range from 200-1000 L/min with an accuracy of 1.5 percent. Pressure transducers with a range from 0 to 2068 kPa were also positioned on the lines in order to know the line pressures. The ignition system design, development and testing is also described with its integration to the High Pressure Gas Turbine Combustor. A modified spark plug was used to provide the igniter with an ignition source. A diffusion flame was used to ignite the main line using methane as the fuel that utilizes the air in the combustion chamber as the oxidizer. Testing included a functional test of the equipment, and

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Combustor technology for future aircraft

    NASA Technical Reports Server (NTRS)

    Tacina, Robert R.

    1990-01-01

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

  6. Atmospheric scavenging exhaust

    NASA Technical Reports Server (NTRS)

    Fenton, D. L.; Purcell, R. Y.

    1977-01-01

    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. The airborne HCl concentration varied from 0.2 to 10.0 ppm and the raindrop sizes tested included 0.55 mm, 1.1 mm, and 3.0 mm. Two chambers were used to conduct the experiments. A large, rigid walled, spherical chamber stored the exhaust constituents while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique employed. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity.

  7. 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

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

    NASA Technical Reports Server (NTRS)

    Riggins, David

    1998-01-01

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

  9. Investigation on the flame dynamics of meso-combustors

    NASA Astrophysics Data System (ADS)

    Ahmed, Mahbub

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

  10. 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.

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

    NASA Technical Reports Server (NTRS)

    Smith, K. O.

    1979-01-01

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

  12. 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.

  13. 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.

  14. Analytical and experimental evaluations of the effect of broad property fuels on combustors for commercial aircraft gas turbine engines

    NASA Technical Reports Server (NTRS)

    Smith, A. L.

    1980-01-01

    The impacts of broad property fuels on the design, performance, durability, emissions, and operational characteristics of current and advanced combustors for commercial aircraft gas turbine engines were studied. The effect of fuel thermal stability on engine and airframe fuel system was evaluated. Tradeoffs between fuel properties, exhaust emissions, and combustor life were also investigated. Results indicate major impacts of broad property fuels on allowable metal temperatures in fuel manifolds and injector support, combustor cyclic durability, and somewhat lesser impacts on starting characteristics, lightoff, emissions, and smoke.

  15. Analytical and experimental evaluations of the effect of broad property fuels on combustors for commercial aircraft gas turbine engines

    NASA Technical Reports Server (NTRS)

    Smith, A. L.

    1980-01-01

    Analytical and experimental studies were conducted in three contract activities funded by the National Aeronautics and Space Administration, Lewis Research Center, to assess the impacts of broad property fuels on the design, performance, durability, emissions and operational characteristics of current and advanced combustors for commercial aircraft gas turbine engines. The effect of fuel thermal stability on engine and airframe fuel system was evaluated. Trade-offs between fuel properties, exhaust emissions and combustor life were also investigated. Results indicate major impacts of broad property fuels on allowable metal temperatures in fuel manifolds and injector support, combustor cyclic durability and somewhat lesser impacts on starting characteristics, lightoff, emissions and smoke.

  16. Experimental clean combustor program, phase 3: Noise measurement addendum. [CF6-50 high bypass turbofan engine noise

    NASA Technical Reports Server (NTRS)

    Doyle, V. L.

    1978-01-01

    The acoustic characteristics of the double annular combustor in a CF6-50 high bypass turbofan engine were investigated. Internal fluctuating pressure measurements were made in the combustor region and in the core exhaust. The transmission loss across the turbine and nozzle was determined from the measurements and compared to previous component results and present theory. The primary noise source location in the combustor was investigated. Spectral comparisons of test rig results were made with the engine results. The measured overall power level was compared with component and engine correlating parameters.

  17. 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.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

  1. 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.

  2. Annular vortex combustor

    DOEpatents

    Nieh, Sen; Fu, Tim T.

    1992-01-01

    An apparatus for burning coal water fuel, dry ultrafine coal, pulverized l and other liquid and gaseous fuels including a vertically extending outer wall and an inner, vertically extending cylinder located concentrically within the outer wall, the annnular space between the outer wall and the inner cylinder defining a combustion chamber and the all space within the inner cylinder defining an exhaust chamber. Fuel and atomizing air are injected tangentially near the bottom of the combustion chamber and secondary air is introduced at selected points along the length of the combustion chamber. Combustion occurs along the spiral flow path in the combustion chamber and the combined effects of centrifugal, gravitational and aerodynamic forces cause particles of masses or sizes greater than the threshold to be trapped in a stratified manner until completely burned out. Remaining ash particles are then small enough to be entrained by the flue gas and exit the system via the exhaust chamber in the opposite direction.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  4. 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

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

    SciTech Connect

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

    1994-08-01

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

  6. Experimental clean combustor program, phase 2

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  7. 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.

  8. 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.

  9. Combustor and method for distributing fuel in the combustor

    DOEpatents

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

    2016-04-26

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

  10. Compact Laser-Based Sensors for Monitoring and Control of Gas Turbine Combustors

    NASA Technical Reports Server (NTRS)

    Hanson, Ronald K.; Jeffries, Jay B.

    2003-01-01

    Research is reported on the development of sensors for gas turbine combustor applications that measure real-time gas temperature using near-infrared water vapor absorption and concentration in the combustor exhaust of trace quantities of pollutant NO and CO using mid-infrared absorption. Gas temperature is extracted from the relative absorption strength of two near-infrared transitions of water vapor. From a survey of the water vapor absorption spectrum, two overtone transitions near 1800 nm were selected that can be rapidly scanned in wavelength by injection current tuning a single DFB diode laser. From the ratio of the absorbances on these selected transitions, a path-integrated gas temperature can be extracted in near-real time. Demonstration measurements with this new temperature sensor showed that combustor instabilities could be identified in the power spectrum of the temperature versus time record. These results suggest that this strategy is extremely promising for gas turbine combustor control applications. Measurements of the concentration of NO and CO in the combustor exhaust are demonstrated with mid-infrared transitions using thermo-electrically cooled, quantum cascade lasers operating near 5.26 and 4.62 microns respectively. Measurements of NO are performed in an insulated exhaust duct of a C2H4-air flame at temperatures of approximately 600 K. CO measurements are performed above a rich H2-air flame seeded with CO2 and cooled with excess N2 to 1150 K. Using a balanced ratiometric detection technique a sensitivity of 0.36 ppm-m was achieved for NO and 0.21 ppm-m for CO. Comparisons between measured and predicted water-vapor and CO2 interference are discussed. The mid-infrared laser quantum cascade laser technology is in its infancy; however, these measurements demonstrate the potential for pollutant monitoring in exhaust gases with mid-IR laser absorption.

  11. 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.

  12. 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.

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

    SciTech Connect

    Rodriguez, Jose L.

    2015-09-15

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

  14. Pulse combustor with controllable oscillations

    DOEpatents

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

    1992-01-01

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

  15. Pulse combustor with controllable oscillations

    SciTech Connect

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

    1991-12-31

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

  16. Assessment of analytical and experimental techniques utilized in conducting plume technology tests 575 and 593. [exhaust flow simulation (wind tunnel tests) of scale model Space Shuttle Orbiter

    NASA Technical Reports Server (NTRS)

    Baker, L. R.; Sulyma, P. R.; Tevepaugh, J. A.; Penny, M. M.

    1976-01-01

    Since exhaust plumes affect vehicle base environment (pressure and heat loads) and the orbiter vehicle aerodynamic control surface effectiveness, an intensive program involving detailed analytical and experimental investigations of the exhaust plume/vehicle interaction was undertaken as a pertinent part of the overall space shuttle development program. The program, called the Plume Technology program, has as its objective the determination of the criteria for simulating rocket engine (in particular, space shuttle propulsion system) plume-induced aerodynamic effects in a wind tunnel environment. The comprehensive experimental program was conducted using test facilities at NASA's Marshall Space Flight Center and Ames Research Center. A post-test examination of some of the experimental results obtained from NASA-MSFC's 14 x 14-inch trisonic wind tunnel is presented. A description is given of the test facility, simulant gas supply system, nozzle hardware, test procedure and test matrix. Analysis of exhaust plume flow fields and comparison of analytical and experimental exhaust plume data are presented.

  17. 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.

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

    NASA Technical Reports Server (NTRS)

    Bianco, Jean

    1995-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  20. 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.

  1. Modeling a Transient Catalytic Combustor

    NASA Technical Reports Server (NTRS)

    Tien, J. S.

    1985-01-01

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

  2. 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.

  3. 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.

  4. Analytical fuel property effects: Small combustors

    NASA Technical Reports Server (NTRS)

    Cohen, J. D.

    1984-01-01

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

  5. 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.

  6. 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.

  7. Performance and emission characteristics of swirl-can combustors to near-stoichiometric fuel-air ratio

    NASA Technical Reports Server (NTRS)

    Diehl, L. A.; Trout, A. M.

    1976-01-01

    Emissions and performance characteristics were determined for two full annular swirl-can combustors operated to near stoichiometric fuel-air ratio. Test condition variations were as follows: combustor inlet-air temperatures, 589, 756, 839, and 894 K; reference velocities, 24 to 37 meters per second; inlet pressure, 62 newtons per square centimeter; and fuel-air ratios, 0.015 to 0.065. The combustor average exit temperature and combustor efficiency were calculated from the combustor exhaust gas composition. For fuel-air ratios greater than 0.04, the combustion efficiency decreased with increasing fuel-air ratios in a near-linear manner. Increasing the combustor inlet air temperature tended to offset this decrease. Maximum oxides of nitrogen emission indices occurred at intermediate fuel-air ratios and were dependent on combustor design. Carbon monoxide levels were extremely high and were the primary cause of poor combustion efficiency at the higher fuel-air ratios. Unburned hydrocarbons were low for all test conditions. For high fuel-air ratios SAE smoke numbers greater than 25 were produced, except at the highest inlet-air temperatures.

  8. The E3 combustors: Status and challenges. [energy efficient turbofan engines

    NASA Technical Reports Server (NTRS)

    Sokolowski, D. E.; Rohde, J. E.

    1981-01-01

    The design, fabrication, and initial testing of energy efficient engine combustors, developed for the next generation of turbofan engines for commercial aircraft, are described. The combustor designs utilize an annular configuration with two zone combustion for low emissions, advanced liners for improved durability, and short, curved-wall, dump prediffusers for compactness. Advanced cooling techniques and segmented construction characterize the advanced liners. Linear segments are made from castable, turbine-type materials.

  9. Emissions of nitrogen oxides from an experimental hydrogen-fueled gas turbine combustor

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

    The effect of operating variables of a hydrogen fueled combustor on exhaust concentrations of total oxides of nitrogen was determined at inlet-air temperature levels up to 810 K, pressure of 414,000N/sa m, and reference velocity of 21.3 m/sec. The combustor, which was originally designed for hydrocarbon fuel produced a NO(x) concentration of 380 ppm with hydrogen at 810 K inlet-air temperature. A reduction in NO(x) of about 30 % was obtained by modification to a lean or rich primary zone. The lowest NO(x) levels obtained with hydrogen were equivalent to those of the reference combustor burning hydrocarbon fuels.

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

    SciTech Connect

    Tolpadi, A.K.

    1995-10-01

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

  11. Exhaust emission control apparatus

    SciTech Connect

    Eng, J.W.

    1991-09-24

    This patent describes an exhaust control apparatus for muffling noise and treating odors and pollutants, including solid particulate and gases in the exhaust of an internal combustion engine. It comprises an exhaust inlet tube for receiving the exhaust generated by an internal combustion engine; a cyclone barrier concentrically surrounding the exhaust inlet tube, a ring cavity between the cyclone tube and exhaust inlet tube defining a cyclone chamber in which the exhaust is treated; means for directing the exhaust from the exhaust inlet tube into the cyclone chamber; electrode means having small openings through which the exhaust passes to enter the cyclone chamber, the electrode means generating electrostatic forces which charge the solid particulate in the exhaust, ionize air and generate ozone in the cyclone chamber near the electrode; means for injecting air into the cyclone chamber causing centrifugal flow of the air and the exhausted within the cyclone chamber and increasing a dwell time of the exhaust within the cyclone chamber.

  12. 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.

  13. Modelling of furnaces and combustors

    SciTech Connect

    Kahil, E.E.

    1985-01-01

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

  14. TRW utility demonstration unit

    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. During this report period, activity continued to address the total program funding shortfall. Ideas and responsibilities for further evaluation have been put forward to reduce the shortfall. In addition, an effort aimed at gaining additional program sponsorships, was initiated.

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

    NASA Technical Reports Server (NTRS)

    Karchmer, A. M.

    1983-01-01

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

  16. 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.

  17. Rectangular capture area to circular combustor scramjet engine

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.

    1978-01-01

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

  18. 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.

  19. 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.

  20. Gas separation process using membranes with permeate sweep to remove CO.sub.2 from gaseous fuel combustion exhaust

    DOEpatents

    Wijmans Johannes G.; Merkel, Timothy C.; Baker, Richard W.

    2012-05-15

    A gas separation process for treating exhaust gases from the combustion of gaseous fuels, and gaseous fuel combustion processes including such gas separation. The invention involves routing a first portion of the exhaust stream to a carbon dioxide capture step, while simultaneously flowing a second portion of the exhaust gas stream across the feed side of a membrane, flowing a sweep gas stream, usually air, across the permeate side, then passing the permeate/sweep gas back to the combustor.

  1. 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.

  2. 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.

  3. Combustor modelling for scramjet engines

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  4. Optical and probe determination of soot concentrations in a model gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Eckerle, W. A.; Rosfjord, T. J.

    1986-01-01

    An experimental program was conducted to track the variation in soot loading in a generic gas turbine combustor. The burner is a 12.7-cm dia cylindrical device consisting of six sheet-metal louvers. Determination of soot loading along the burner length is achieved by measurement at the exit of the combustor and then at upstream stations by sequential removal of liner louvers to shorten burner length. Alteration of the flow field approaching and within the shortened burners is minimized by bypassing flow in order to maintain a constant linear pressure drop. The burner exhaust flow is sampled at the burner centerline to determine soot mass concentration and smoke number. Characteristic particle size and number density, transmissivity of the exhaust flow, and local radiation from luminous soot particles in the exhaust are determined by optical techniques. Four test fuels are burned at three fuel-air ratios to determine fuel chemical property and flow temperature influences. Particulate concentration data indicate a strong oxidation mechanism in the combustor secondary zone, though the oxidation is significantly affected by flow temperature. Soot production is directly related to fuel smoke point.

  5. 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.

  6. 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.

  7. Combustor with multistage internal vortices

    DOEpatents

    Shang, Jer Y.; Harrington, Richard E.

    1989-01-01

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

  8. Critical Propulsion Components. Volume 3; Exhaust Nozzle

    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 Propulsion System Studies, Volume 2: Combustor, Volume 3: Exhaust Nozzle, and Volume 4: Inlet and Fan/Inlet Acoustic Team.

  9. The Applicability of Jet-Shear-Layer Mixing and Effervescent Atomization for Low-NO(x) Combustors

    NASA Technical Reports Server (NTRS)

    Colantonio, R. O.

    1998-01-01

    An investigation has been conducted to develop appropriate technologies for a low-NO(x), liquid-fueled combustor. The combustor incorporates an effervescent atomizer used to inject fuel into a premixing duct. Only a fraction of the combustion air is used in the premixing process. This fuel-rich mixture is introduced into the remaining combustion air by a rapid jet-shear-layer mixing process involving radial fuel-air jets impinging on axial air jets in the primary combustion zone. Computational modeling was used as a tool to facilitate a parametric analysis appropriate to the design of an optimum low-NO(x) combustor. A number of combustor configurations were studied to assess the key combustor technologies and to validate the three-dimensional modeling code. The results from the experimental testing and computational analysis indicate a low-NO(x) potential for the jet-shear-layer combustor. Key features found to affect NOx emissions are the primary combustion zone fuel-air ratio, the number of axial and radial jets, the aspect ratio and radial location of the axial air jets, and the radial jet inlet hole diameter. Each of these key parameters exhibits a low-NO(x) point from which an optimized combustor was developed Also demonstrated was the feasibility of utilizing an effervescent atomizer for combustor application. Further developments in the jet-shear-layer mixing scheme and effervescent atomizer design promise even lower NO(x) with high combustion efficiency.

  10. 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.

  11. 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.

  12. 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.

  13. Development of topping combustor for advanced concept pressurized fluidized-bed combustion

    SciTech Connect

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

    1994-10-01

    The objective of this program is to develop a topping combustor to operate in a Second-Generation Pressurized Fluidized Bed (PFBC) Combined Cycle power generation system. The combustor must be able to: lightoff with a high heating value fuel and compressor discharge air to heat the fluidized bed(s) and provide power for PFBC and carbonizer off-line; operate with 1,600 F oxygen depleted air from the PFBC and high heating value fuel to handle carbonizer off-line conditions; ramp up to 100% carbonizer syngas firing (normal operation) by firing a blend of decreasing high heating value fuel and increasing low heating value syngas; utilize the vitiated air, at temperatures up to 1,600 F for as much cooling of the metal combustor as possible, thus minimizing the compressor bypass air needed for combustor cooling; provide an acceptance exit temperature pattern at the desired burner outlet temperature (BOT); minimize the conversion of fuel bound nitrogen (FBN) present in the syngas to NO{sub x}; and have acceptably high combustion efficiency, and low emissions of carbon monoxide, UHC, etc. This paper reports the results of tests of a 14 inch diameter topping combustor with a modified fuel-rich zone conducted in June 1993, design of an 18 inch diameter topping combustor to be tested in June 1994 and afterwards, and results of a 50% scale cold flow model which has been built and tested.

  14. 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.

  15. Combustor oscillation pressure stabilizer

    SciTech Connect

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

    1996-12-31

    In accordance with the objective of the present invention, the active control of unsteady combustion induced oscillations in a combustion chamber fired by a suitable fuel and oxidizer mixture, such as a hydrocarbon fuel and air mixture, is provided by restructuring and moving the position of the main flame front and thereby increasing the transport time and displacing the pressure wave further away from the in-phase relationship with the periodic heat release. The restructuring and repositioning of the main flame are achieved by utilizing a pilot flame which is pulsed at a predetermined frequency corresponding to less than about one-half the frequency of the combustion oscillation frequency with the duration of each pulse being sufficient to produce adequate secondary thermal energy to restructure the main flame and thereby decouple the heat release from the acoustic coupling so as to lead to a reduction in the dynamic pressure amplitude. The pulsating pilot flame produces a relatively small and intermittently existing flame front in the combustion zone that is separate from the oscillating main flame front but which provides the thermal energy necessary to effectively reposition the location of the oscillating main flame front out of the region in the combustion zone where the acoustic coupling can occur with the main flame and thereby effectively altering the oscillation-causing phase relationship with the heat of combustion.

  16. 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.

  17. Combustor with non-circular head end

    SciTech Connect

    Kim, Won -Wook; McMahan, Kevin Weston

    2015-09-29

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

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Flow and Emissions Characteristics of Multi-Swirler Combustor

    NASA Astrophysics Data System (ADS)

    Gutmark, Ephraim; Li, Guoqiang

    2003-11-01

    Modern industrial gas-turbine spray combustors feature multiple swirlers and distributed fuel injection for rapid mixing and stabilization. The flow field of this combustor, the related combustion characteristics and their control are discussed. The velocity flow field downstream of a Triple Annular Research Swirler (TARS) was characterized. Multiple combinations of swirlers were tested in cold flow under atmospheric conditions with and without confining combustion chamber. The experiments showed that a central recirculation zone (CTRZ), an annular jet with internal and external shear layers dominated the flow field downstream of TARS. Compared to unconfined case, flow with confined tube showed an enlarged CTRZ region and a recirculation region in the expansion corner with reduced concentration of turbulence intensity in the jet region. TARS also produced low emissions of NOx and CO. Measurements were performed to study the effects of several factors, including swirler combinations, exhaust nozzle size, air assist for fuel atomization and mixing length on NOx and CO emissions and combustion instability. The data showed that emissions and stability depend on the combination of several of these factors.

  5. Fabrication of strain-isolated ceramic coated combustor components

    NASA Technical Reports Server (NTRS)

    Rutter, S.

    1985-01-01

    The use of strain-isolated ceramic coated material to produce an AGT1500 combustor scroll-shaped transition duct which requires no air for film cooling is investigated. The scroll receives the exhaust of the can-style combustor liner and turns it into the annular inlet of the high pressure gas producer turbine nozzle. Strain-isolation of plasma sprayed thermal barrier coating is achieved by placing a compliant pad between the structural base metal and the ceramic coating. The compliant pad is brazed to the metal structure. In order to achieve a good braze bond, the strain-isolating compliant pad and base metal must be closely matched in shape and tightly fixtured for joining. The complex geometry of the AGT1500 scroll makes it impractical to attack pads to the supporting structure in its finished shape. Instead the pads are brazed to flat stock and post-formed into scroll sections. While test samples were successfully post-formed, plasma sprayed, and subjected to cyclic heating, the forming of full scale parts by normal methods resulted in tearing of the Hastelloy-X base metal because of embrittlement by the braze material. Several solutions were explored which finally resulted in the successful forming of full scale scroll parts.

  6. 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.

  7. 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.

  8. Solid Fuel Ramjet Combustor Design

    NASA Astrophysics Data System (ADS)

    Krishnan, S.; George, Philmon

    1998-03-01

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

  9. 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.

  10. 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.

  11. Heat Exhaustion, First Aid

    MedlinePlus

    ... rashes clinical tools newsletter | contact Share | Heat Exhaustion, First Aid A A A Heat exhaustion signs and symptoms ... specific to the other stages of heat illness. First Aid Guide Use a combination of the following measures ...

  12. Gas turbine combustor stabilization by heat recirculation

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  13. 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.

  14. 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.

  15. Chaos in an imperfectly premixed model combustor.

    PubMed

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

    2015-02-01

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

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Heat transfer in circulating fluidized bed combustor

    SciTech Connect

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

    1999-07-01

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

  1. 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.

  2. A flow calorimeter for determining combustion efficiency from residual enthalpy of exhaust gases

    NASA Technical Reports Server (NTRS)

    Evans, Albert; Hibbard, Robert R

    1954-01-01

    A flow calorimeter for determining the combustion efficiency of turbojet and ram-jet combustors from measurement of the residual enthalpy of combustion of the exhaust gas is described. Briefly, the calorimeter catalytically oxidizes the combustible constituents of exhaust-gas samples, and the resultant temperature rise is measured. This temperature rise is related to the residual enthalpy of combustion of the sample by previous calibration of the calorimeter. Combustion efficiency can be calculated from a knowledge of the residual enthalpy of the exhaust gas and the combustor input enthalpy. An accuracy of +-0.2 Btu per cubic foot was obtained with prepared fuel-air mixtures, and the combustion efficiencies of single turbojet combustors measured by both the flow-calorimeter and heat-balance methods compared within 3 percentage units. Flow calorimetry appears to be a suitable method for determining combustion efficiencies at high combustor temperatures where ordinary thermocouples cannot be used. The method is fundamentally more accurate than heat-balance methods at high combustion efficiencies and can be used to verify near-100-percent efficiency data.

  3. Coherence between internal and external noise generated by gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Strahle, W. C.; Muthukrishnan, M.; Neale, D. H.

    1977-01-01

    Experiments and analysis on a gas turbine combustor unit are reported with a view in mind to separate propagated acoustic power from non-propagating 'pseudo-sound'. Analytically, it is suggested that a transition frequency will exist below which the interior pressure fluctuations are non-propagating, whereas above this frequency, of the order of 100 Hz, the noise is dominated by propagating acoustic waves. Coherence measurements are reported which show this concept to be borne out experimentally. Coherence between interior and exterior microphones is measured over a wide range of experimental conditions for a gas turbine combustor exhausting directly to the atmosphere. The purpose is to show that below a certain frequency, measurements of interior noise are not indicative of combustion noise ultimately propagating from an engine.

  4. 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.

  5. Introducing the VRT gas turbine combustor

    NASA Astrophysics Data System (ADS)

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

    1990-07-01

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

  6. 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.

  7. 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.

  8. Scramjet including integrated inlet and combustor

    SciTech Connect

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

    1992-02-04

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

  9. Prediction of soot and thermal radiation in a model gas turbine combustor burning kerosene fuel spray at different swirl levels

    NASA Astrophysics Data System (ADS)

    Ghose, Prakash; Patra, Jitendra; Datta, Amitava; Mukhopadhyay, Achintya

    2016-05-01

    Combustion of kerosene fuel spray has been numerically simulated in a laboratory scale combustor geometry to predict soot and the effects of thermal radiation at different swirl levels of primary air flow. The two-phase motion in the combustor is simulated using an Eulerian-Lagragian formulation considering the stochastic separated flow model. The Favre-averaged governing equations are solved for the gas phase with the turbulent quantities simulated by realisable k-ɛ model. The injection of the fuel is considered through a pressure swirl atomiser and the combustion is simulated by a laminar flamelet model with detailed kinetics of kerosene combustion. Soot formation in the flame is predicted using an empirical model with the model parameters adjusted for kerosene fuel. Contributions of gas phase and soot towards thermal radiation have been considered to predict the incident heat flux on the combustor wall and fuel injector. Swirl in the primary flow significantly influences the flow and flame structures in the combustor. The stronger recirculation at high swirl draws more air into the flame region, reduces the flame length and peak flame temperature and also brings the soot laden zone closer to the inlet plane. As a result, the radiative heat flux on the peripheral wall decreases at high swirl and also shifts closer to the inlet plane. However, increased swirl increases the combustor wall temperature due to radial spreading of the flame. The high incident radiative heat flux and the high surface temperature make the fuel injector a critical item in the combustor. The injector peak temperature increases with the increase in swirl flow mainly because the flame is located closer to the inlet plane. On the other hand, a more uniform temperature distribution in the exhaust gas can be attained at the combustor exit at high swirl condition.

  10. Nonequilibrium sulfur capture and retention in an air cooled slagging coal combustor. Quarterly technical progress report, 1996

    SciTech Connect

    Zauderer, B.

    1996-11-01

    The objective of this 24 month project is to determine the degree of sulfur retention in slag in a full scale cyclone coal combustor with sulfur capture by calcium oxide sorbent injection into the combustor. This sulfur capture process consists of two steps: Capture of sulfur with calcined calcium oxide followed by impact of the reacted sulfur-calcium particles on the liquid slag lining the combustor. The sulfur bearing slag must be removed within several minutes from the combustor to prevent re-evolution of the sulfur from the slag. To accomplish this requires slag mass flow rates in the range of several 100 lb/hr. To study this two step process in the combustor, two groups of tests are being implemented. In the first group, calcium sulfate in the form of gypsum, or plaster of Paris, was injected in the combustor to determine sulfur evolution from slag. In the second group, the entire process is tested with limestone and/or calcium hydrate injected into the combustor. This entire effort consists of a series of up to 16 parametric tests in a 20 MMtu/hr slagging, air cooled, cyclone combustor. During the present quarterly reporting period ending September 30,1996, three tests in this project were implemented, bringing the total tests to 5. In addition, a total of 10 test days were completed during this quarter on the parallel project that utilizes the same 20 MMtu/hr combustor. The results of that project, especially those related to improved slagging performance, have a direct bearing on this project in assuring proper operation at the high slag flow rates that may be necessary to achieve high sulfur retention in slag.

  11. Evaluation of Federal Aviation Administration ion engine exhaust sampling rake

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    A FAA exhaust emissions rake was tested in the Experimental Clean Combustor Program, Phase 3 to permit comparison of the values of gaseous emissions and smoke measured by the FAA rake with those measured with the NASA Pratt and Whitney Aircraft (P and WA) rake used in the Phase 3 Experimental Clean Combustor Program and with station seven probes. The results showed that the levels of CO, THC, NOx and smoke measured by the FAA and NASA/P and WA rakes agree well at high power, but that CO emissions measured by the FAA rake were approximately 10 percent higher than those measured by the NASA/P and WA rake at low power.

  12. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2009-01-01

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

  14. 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.

  15. Combustor kinetic energy efficiency analysis of the hypersonic research engine data

    NASA Astrophysics Data System (ADS)

    Hoose, K. V.

    1993-11-01

    A one-dimensional method for measuring combustor performance is needed to facilitate design and development scramjet engines. A one-dimensional kinetic energy efficiency method is used for measuring inlet and nozzle performance. The objective of this investigation was to assess the use of kinetic energy efficiency as an indicator for scramjet combustor performance. A combustor kinetic energy efficiency analysis was performed on the Hypersonic Research Engine (HRE) data. The HRE data was chosen for this analysis due to its thorough documentation and availability. The combustor, inlet, and nozzle kinetic energy efficiency values were utilized to determine an overall engine kinetic energy efficiency. Finally, a kinetic energy effectiveness method was developed to eliminate thermochemical losses from the combustion of fuel and air. All calculated values exhibit consistency over the flight speed range. Effects from fuel injection, altitude, angle of attack, subsonic-supersonic combustion transition, and inlet spike position are shown and discussed. The results of analyzing the HRE data indicate that the kinetic energy efficiency method is effective as a measure of scramjet combustor performance.

  16. Emissions Prediction and Measurement for Liquid-Fueled TVC Combustor with and without Water Injection

    NASA Technical Reports Server (NTRS)

    Brankovic, A.; Ryder, R. C., Jr.; Hendricks, R. C.; Liu, N.-S.; Shouse, D. T.; Roquemore, W. M.

    2005-01-01

    An investigation is performed to evaluate the performance of a computational fluid dynamics (CFD) tool for the prediction of the reacting flow in a liquid-fueled combustor that uses water injection for control of pollutant emissions. The experiment consists of a multisector, liquid-fueled combustor rig operated at different inlet pressures and temperatures, and over a range of fuel/air and water/fuel ratios. Fuel can be injected directly into the main combustion airstream and into the cavities. Test rig performance is characterized by combustor exit quantities such as temperature and emissions measurements using rakes and overall pressure drop from upstream plenum to combustor exit. Visualization of the flame is performed using gray scale and color still photographs and high-frame-rate videos. CFD simulations are performed utilizing a methodology that includes computer-aided design (CAD) solid modeling of the geometry, parallel processing over networked computers, and graphical and quantitative post-processing. Physical models include liquid fuel droplet dynamics and evaporation, with combustion modeled using a hybrid finite-rate chemistry model developed for Jet-A fuel. CFD and experimental results are compared for cases with cavity-only fueling, while numerical studies of cavity and main fueling was also performed. Predicted and measured trends in combustor exit temperature, CO and NOx are in general agreement at the different water/fuel loading rates, although quantitative differences exist between the predictions and measurements.

  17. Second-generation PFBC systems research and development, Phase 2 topping combustor development

    SciTech Connect

    Domeracki, W.F.; Dowdy, T.E.; Bachovchin, D.; Foote, J.; Pillsbury, P.W.; Bouvier, B.U.; Muller, K.F.

    1993-09-01

    The use of a Circulating Pressurized Fluidized Bed Combustor (CPFBC) as the primary combustion system for a combustion turbine requires transporting compressor air to the CPFBC and vitiated air/fuel gas back to the turbine. In addition, the topping combustion system must be located in the returning vitiated airflow path. The conventional fuel system and turbine center section require major changes for the applications. The combustion zone of the Westinghouse 501F turbine currently in production cannot contain the topping combustion system within the main structural pressure shell. Although the pressure casing can be enlarged both radially and longitudinally to accommodate the topping combustor system, the integrity and rigidity of the main shell would be significantly affected and, it could introduce rotor dynamics problems and preclude shipping the unit assembled. The currently favored configuration, which utilizes two topping combustor assemblies, one on each side of the unit, is shown in Figure 1. Half of the vitiated air from the CPFBC enters each of the internal plenum chambers in which the topping combustors are mounted. Fuel gas enters the assembly via the fuel nozzles at the head end of the combustor. Combustion occurs, and the products of combustion are ducted into the main shell for distribution to the first-stage turbine vanes. Compressor discharge air leaves the main shell, flowing around the annular duct into adjacent combustion shells. The air flows around the vitiated air plenums and leaves each combustion assembly via nozzles and is ducted to the CPFBC and carbonizer.

  18. An efficient liner cooling scheme for advanced small gas turbine combustors

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    A joint Army/NASA program was conducted to design, fabricate, and test an advanced, small gas turbine, reverse-flow combustor utilizing a compliant metal/ceramic (CMC) wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CMC cooling technique and then demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F burner outlet temperature. The CMC 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 benefits of improved efficiency, reduced emissions, and lower smoke levels. The program was divided into four tasks. Task 1 defined component materials and localized design of the composite wall structure in conjunction with development of basic design models for the analysis of flow and heat transfer through the wall. Task 2 included implementation of the selected materials and validated design models during combustor preliminary design. Detail design of the selected combustor concept and its refinement with 3D aerothermal analysis were completed in Task 3. Task 4 covered detail drawings, process development and fabrication, and a series of burner rig tests. The purpose of this paper is to provide details of the investigation into the fundamental flow and heat transfer characteristics of the CMC wall structure as well as implementation of the fundamental analysis method for full-scale combustor design.

  19. Exhaust Nozzle Materials Development for the High Speed Civil Transport

    NASA Technical Reports Server (NTRS)

    Grady, J. E.

    1999-01-01

    The United States has embarked on a national effort to develop the technology necessary to produce a Mach 2.4 High Speed Civil Transport (HSCT) for entry into service by the year 2005. The viability of this aircraft is contingent upon its meeting both economic and environmental requirements. Two engine components have been identified as critical to the environmental acceptability of the HSCT. These include a combustor with significantly lower emissions than are feasible with current technology, and a lightweight exhaust nozzle that meets community noise standards. The Enabling Propulsion Materials (EPM) program will develop the advanced structural materials, materials fabrication processes, structural analysis and life prediction tools for the HSCT combustor and low noise exhaust nozzle. This is being accomplished through the coordinated efforts of the NASA Lewis Research Center, General Electric Aircraft Engines and Pratt & Whitney. The mission of the EPM Exhaust Nozzle Team is to develop and demonstrate this technology by the year 1999 to enable its timely incorporation into HSCT propulsion systems.

  20. 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.

  1. 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.

  2. Exhaust gas purification device

    SciTech Connect

    Fujiwara, H.; Hibi, T.; Sayo, S.; Sugiura, Y.; Ueda, K.

    1980-02-19

    The exhaust gas purification device includes an exhaust manifold , a purification cylinder connected with the exhaust manifold through a first honey-comb shaped catalyst, and a second honeycomb shaped catalyst positioned at the rear portion of the purification cylinder. Each catalyst is supported by steel wool rings including coarse and dense portions of steel wool. The purification device further includes a secondary air supplying arrangement.

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  4. Preliminary comparison of theory and experiment for a conical, pressurized-fluidized-bed coal combustor

    NASA Technical Reports Server (NTRS)

    Patch, R. W.

    1979-01-01

    A published model was used for a comparison of theory with an actual combustor burning caking bituminous coal and using limestone to reduce sulfur dioxide emission. Theoretical bed pressure drop was in good agreement with experiment. The burnable carbon elutriated was not in agreement with experiment, at least partly because the exhaust port was apparently below the transport disengaging height. The observed nitrogen oxides emission rate was about half the theoretical value. There was order-or-magnitude agreement of sulfur dioxide emission rates.

  5. Combustor bulkhead heat shield assembly

    SciTech Connect

    Zeisser, M.H.

    1990-06-19

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

  6. 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

  7. Altitude Wind Tunnel Investigation of the Performance of Compressor, Combustor, and Turbine Components of Prototype J47D (RX1-1) Turbojet Engine

    NASA Technical Reports Server (NTRS)

    Farley, John M

    1951-01-01

    As a portion of an over-all performance investigation of the prototype J47D (RX-1) turbojet engine, performance of the compressor, combustor, and turbine components has been determined in the Lewis altitude wind tunnel over a range of altitude from 5000 to 55,000 feet and at flight Mach numbers from 0.19 to 0.92. Investigations were conducted with the engine operating on an electronic control schedule and slow with a two-lever control system by which fuel flow and exhaust-nozzle area could be controlled separately. Two combustor configurations were investigated.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  9. Immune Exhaustion and Transplantation.

    PubMed

    Sanchez-Fueyo, A; Markmann, J F

    2016-07-01

    Exhaustion of lymphocyte function through chronic exposure to a high load of foreign antigen is well established for chronic viral infection and antitumor immunity and has been found to be associated with a distinct molecular program and characteristic cell surface phenotype. Although exhaustion has most commonly been studied in the context of CD8 viral responses, recent studies indicate that chronic antigen exposure may affect B cells, NK cells and CD4 T cells in a parallel manner. Limited information is available regarding the extent of lymphocyte exhaustion development in the transplant setting and its impact on anti-graft alloreactivity. By analogy to the persistence of a foreign virus, the large mass of alloantigen presented by an allograft in chronic residence could provide an ideal setting for exhausting donor-reactive T cells. The extent of T cell exhaustion occurring with various allografts, the kinetics of its development, whether exhaustion is influenced positively or negatively by different immunosuppressants, and the impact of exhaustion on graft survival and tolerance development remains a fertile area for investigation. Harnessing or encouraging the natural processes of exhaustion may provide a novel means to promote graft survival and transplantation tolerance. PMID:26729653

  10. Duplex tab exhaust nozzle

    NASA Technical Reports Server (NTRS)

    Gutmark, Ephraim Jeff (Inventor); Martens, Steven (nmn) (Inventor)

    2012-01-01

    An exhaust nozzle includes a conical duct terminating in an annular outlet. A row of vortex generating duplex tabs are mounted in the outlet. The tabs have compound radial and circumferential aft inclination inside the outlet for generating streamwise vortices for attenuating exhaust noise while reducing performance loss.

  11. Diesel engine exhaust oxidizer

    SciTech Connect

    Kammel, R.A.

    1992-06-16

    This patent describes a diesel engine exhaust oxidizing device. It comprises: an enclosure having an inlet for receiving diesel engine exhaust, a main flow path through the enclosure to an outlet of the enclosure, a by-ass through the enclosure, and a microprocessor control means.

  12. Combustor development for automotive gas turbines

    SciTech Connect

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

    1983-09-01

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

  13. Atmospheric scavenging of solid rocket exhaust effluents

    NASA Technical Reports Server (NTRS)

    Fenton, D. L.; Purcell, R. Y.

    1978-01-01

    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

  14. Exhaust purification apparatus

    SciTech Connect

    Shinzawa, M.; Ushimura, S.

    1987-05-05

    An exhaust purification apparatus is described for use in an internal combustion engine having an exhaust conduit through which exhaust particles are discharged together with exhaust gas to the atmosphere. Included is an outer shell having an inlet connected to the exhaust conduit and an outlet connected to the atmosphere. The outer shell contains a trap element and a regenerative burner located upstream of the trap element, the regenerative burner comprising: a cylindrical hollow member fixed to the liner and extending within a combustion chamber to define an evaporation chamber, a glow plug for igniting the mixture supplied into the evaporated chamber when actuated; and a control unit responsive to a regeneration requirement for actuating the glow plug and supplying an air-fuel mixture into the evaporation chamber through the mixture conduit.

  15. Development of a dry ultra-low NO{sub x} double swirler staged gas turbine combustor

    SciTech Connect

    Sato, H.; Mori, M.; Nakamura, T.

    1998-01-01

    This paper describes the development of an ultra-low NO{sub x} gas turbine combustor for cogeneration systems. The combustor, called a double swirler staged combustor, utilizes three-staged premixed combustion for low NO{sub x} emission. The unique feature of the combustor is its tertiary premix nozzles located downstream of the double swirler premixing nozzles around the combustor liner. Engine output is controlled by simply varying the fuel gas flow, and therefore employs no complex variable geometries for air flow control. Atmospheric combustion tests have demonstrated the superior performance of the combustor. NO{sub x} level is maintained at less than 3 ppm (O{sub 2} = 15%) over the range of engine output between 50 and 100%. Assuming the general relationship that NO{sub x} emission is proportional to the square root of operating pressure, the NO{sub x} level is estimated at less than 9 ppm (O{sub 2} = 15%) at the actual pressure of 0.91 MPa (abs.). Atmospheric tests have also shown high combustion efficiency; more than 99.9% over the range of engine output between 60 and 100%. Emissions of CO and UHC are maintained at 0 and 1 ppm (O{sub 2} = 15%), respectively, at the full engine load.

  16. 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.

  17. Combustor development for automotive gas turbines

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  18. Flow establishment in a generic scramjet combustor

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  19. LDV measurements in an annular combustor model

    NASA Astrophysics Data System (ADS)

    Barron, Dean A.

    1986-08-01

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

  20. 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.

  1. 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.

  2. Optical Detection Of Flameout In A Combustor

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

  3. 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.

  4. Evaluation of an experimental short-length annular combustor: One-side-entry dilution airflow concept

    NASA Technical Reports Server (NTRS)

    Humenik, F. M.; Biaglow, J. A.

    1973-01-01

    A test program was conducted to evaluate an experimental short-length annular combustor that uses a one-side-entry dilution airflow concept. The combustor design features scoops on the outer liner for controlling the primary- and secondary-zone airflow distribution. Combustor inlet total pressures were limited to 62 N/sq cm (90 psia) with inlet-air temperatures from 590 K (600 F) to 890 K (1150 F). At a diffuser inlet Mach number of 0.25, the exit temperature pattern factor was 0.44 with an average exit temperature of 1436 K (2124 F) and a total pressure loss of 4.3 percent. At a diffuser inlet Mach number of 0.31, the exit temperature pattern factor was reduced to 0.29 with an average exit temperature of 1450 K (2151 F) and a total pressure loss of 6.1 percent. Nominal combustion efficiencies of 100 percent were obtained with the ASTM A-1 fuel. Exhaust gas emissions, smoke, and altitude relight data are included with exit-temperature profiles and distribution patterns.

  5. Results of the NASA/General Electric Experimental Clean Combustor Program

    NASA Technical Reports Server (NTRS)

    Gleason, C. C.; Niedzwiecki, R. W.

    1976-01-01

    The NASA/General Electric Experimental Clean Combustor Program is a multi-year, major contract effort. Primary program objectives are the generation of technology for development of advanced commercial CTOL engines with lower exhaust emissions than current aircraft and, demonstrations of this technology in a full-scale CF6-50C engine in 1976. This paper describes pollution and performance goals, Phase I and II test results and Phase III hardware, pollution sampling techniques and test plans. Pollution results are presented in emission index and Environmental Protection Agency 1979 Standard Parameters (EPAP). Best results were obtained with a double annular combustor concept. This concept, which incorporates multistage burning, produced EPAP values extrapolated to CF6-50C engine conditions for CO, HC, and NOx of 3.3, 0.3 and 4.5, respectively. These represent respective CO, HC and NOx percentage reductions of 69, 93 and 42%, compared to current CF6-50 engine values. The combustor also met development engine performance requirements.

  6. Experimental investigation of the low NO/sub x vortex airblast annular combustor

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  7. The experimental clean combustor program: Description and status to November 1975

    NASA Technical Reports Server (NTRS)

    Niedzwiecki, R. W.

    1975-01-01

    The generation of technology was studied for the development of advanced commercial CTOL aircraft engines with lower exhaust emissions than current aircraft. The program is in three phases. Phase 1, already completed, consisted of screening tests of low pollution combustor concepts. Phase 2, currently in progress, consists of test rig refinement of the most promising combustor concepts. Phase 2 test results are reported. Phase 3, also currently in progress, consists of incorporating and evaluating the best combustors as part of a complete engine. Engine test plans and pollution sampling techniques are described in this report. Program pollution goals, specified at engine idle and take-off conditions, are idle emission index value of 20 and 4 for carbon monoxide (CO) and total unburned hydrocarbons (THC), respectively, and at take-off are an oxides of nitrogen (NOx) emission index level of 10 and a smoke number of 15. Pollution data were obtained at all engine operating conditions. Results are presented in terms of emission index and also in terms of the Environmental Protection Agency's 1979 Standards Parameter.

  8. Oxidation of low calorific value gases -- Applying optimization techniques to combustor design

    SciTech Connect

    Gemmen, R.S.

    1998-07-01

    The design of an optimal air-staged combustor for the oxidation of a low calorific value gas mixture is presented. The focus is on the residual fuel emitted from the anode of a molten carbonate fuel-cell. Both experimental and numerical results are presented. The simplified numerical model considers a series of plug-flow-reactor sections, with the possible addition of a perfectly-stirred-reactor. The parameter used for optimization, Z, is the sum of fuel-component molar flow rates leaving a particular combustor section. An optimized air injection profile is one that minimizes Z for a given combustor length and inlet condition. Since a mathematical proof describing the significance of global interactions remains lacking, the numerical model employs both a Local optimization procedure and a Global optimization procedure. The sensitivity of Z to variations in the air injection profile and inlet temperature is also examined. The results show that oxidation of the anode exhaust gas is possible with low pollutant emissions.

  9. Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

    NASA Technical Reports Server (NTRS)

    Gregory, John W.; Straight, David M.

    1961-01-01

    Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

  10. Predicting and Preventing Incipient Flameout in Combustors

    NASA Technical Reports Server (NTRS)

    Puster, Richard Lee

    2003-01-01

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