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 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.

  3. 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.

  4. Hydrocarbon Constituents of T-56 Combustor Exhaust

    DTIC Science & Technology

    1975-04-01

    May 1973 to determine feasibility of exhaust cryotrapping and to establish analytical methodology for identifying individual hydrocarbon compounds (3...midtemperature setting was at about 166°C, and 33 psig, corresponding to the conditions of many moderate-pressure ratio engines (1). Finally the high -temperature...condition of 2040C and 50 psig simulates many newer high -pressure-ratio engines, like the F-101 and F-100. Table 2 lists typical military engines and

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. 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.

  10. Spray flame and exhaust jet characteristics of a pressurized swirl combustor

    NASA Astrophysics Data System (ADS)

    Linck, Martin Brendan

    This work describes an investigation of swirl-stabilized flames, created in a combustor featuring co-annular swirling airflows, under unenclosed, enclosed, and submerged conditions. A centrally-located fuel nozzle, which uses air-assist atomization, creates a methanol fuel spray. This approach provides great control over fuel spray properties in a compact geometry. Factors affecting the structure of the flames, including the effect of the central atomization air jet, are investigated using three-dimensional particle image velocimetry, direct imaging, and phase-Doppler particle analysis techniques. Exhaust jet temperatures are measured. The dynamic events affecting two-phase exhaust jets from the combustor under submerged conditions are examined using high-speed cinematography and sound spectrum analysis. It is found that the structures of the flames examined, which feature low overall equivalence ratios, are closely linked to the features of the air flowfield in the combustor. Swirl numbers of flows emerging from twisted-vane swirl assemblies are characterized. The structure of the flow is affected by the swirl configuration, but does not depend heavily on the Reynolds number. The central atomization air jet (with or without fuel) reshapes the recirculation region in the swirling flow and has a significant, controllable effect on the structure of the airflow and flame. The effect is the same for nonreacting and reacting flows. In one unique case, the central atomization air interacts with the swirling flow to create two recirculation regions and a lifted flame. The lifted flame is more compact than similar non-lifted flames. The twin-fluid atomization approach is shown to provide effective atomization over a wide range of operating conditions. The two-phase interaction of the exhaust jet is found to depend on the pressure drop over the exhaust nozzle. The dynamic behavior of the exhaust jet is buoyancy-driven at low pressure drops, and is affected by complex

  11. 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.

  12. 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.

  13. 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...

  14. 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.

  15. Desensitizing Flame Structure and Exhaust Emissions to Flow Parameters in an Ultra-Compact Combustor

    DTIC Science & Technology

    2012-03-22

    14 Figure 2.9: CFD computed radial temperature (K) at UCC exit plane showing potential for multiple peak temperatures...21 x Figure 2.12: Changes in jet plume trajectories based on injection angle ................... 22 Figure...operate above stoichiometric fuel-to-air conditions in order to obtain an appropriate heat release in the combustor to produce the necessary work from

  16. 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.

  17. 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.

  18. 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.

  19. 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.

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

    SciTech Connect

    Not Available

    1989-03-01

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

  1. 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.

  2. 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.

  3. 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.

  4. Evaluation of Probe Sampling Versus Optical in situ Measurements of Nitric Oxide Concentrations in a Jet Engine Combustor Exhaust

    DTIC Science & Technology

    1977-01-01

    ATM. ~IAUST CAL TANKS-- I , . . / SOURCI COMBUSTOR BSORPTION CELL 2 A SPECTROMETER SPECTROMETER Figure 1. ,--CO / ANALYZER SYSTEM CxHy...3.0 RESONANCE ABSORPTION METHOD 3.1 INTRODUCTION The use of absorption spectroscopy for species density measurements may take on many forms

  5. 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.

  6. Numerical Evaluation of an Ejector-Enhanced Resonant Pulse Combustor with a Poppet Inlet Valve and a Converging Exhaust Nozzle

    NASA Technical Reports Server (NTRS)

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

    2016-01-01

    A computational investigation of a pressure-gain combustor system for gas turbine applications is presented. The system consists of a valved pulse combustor and an ejector, housed within a shroud. The study focuses on two enhancements to previous models, related to the valve and ejector components. First, a new poppet inlet valve system is investigated, replacing the previously used reed valve configuration. Secondly, a new computational approach to approximating the effects of choked turbine inlet guide vanes present immediately downstream of the Ejector-Enhanced Resonant Pulse Combustor (EERPC) is investigated. Instead of specifying a back pressure at the EERPC exit boundary (as was done in previous studies) the new model adds a converging-diverging (CD) nozzle at the exit of the EERPC. The throat area of the CD nozzle can be adjusted to obtain the desired back pressure level and total mass flow rate. The results presented indicate that the new poppet valve configuration performs nearly as well as the original reed valve system, and that the addition of the CD nozzle is an effective method to approximate the exit boundary effects of a turbine present downstream of the EERPC. Furthermore, it is shown that the more acoustically reflective boundary imposed by a nozzle as compared to a constant pressure surface does not significantly affect operation or performance.

  7. Diesel engine catalytic combustor system. [aircraft engines

    NASA Technical Reports Server (NTRS)

    Ream, L. W. (Inventor)

    1984-01-01

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

  8. Reevaluation of nitric oxide concentration in exhaust of jet engines and combustors. Final report 1 October 1979-30 September 1980

    SciTech Connect

    Few, J.D.; Lowry, H.S. III

    1981-08-01

    Several measurements of nitric oxide (NO) concentrations in the exhaust of jet engines and combustors by probe sampling and ultraviolet (UV) resonance absorption have been reported. Generally, the optically determined concentrations have exceeded the probe determined values appreciably. The theoretical basis for the determination of the NO concentration from absorption measurements has recently been reexamined, and the values previously reported were redetermined. Upon reexamination, the parameter a (the ratio of collisional plus natural half-width to the Doppler half-width) was found to obey the relationship a = C P/T sub 0 to the 1.5 power where P is the pressure in atmospheres, T is the temperature in K, and C is a constant. The value of C was found to be 3.81 (+ or - .40) x 10,000 k to the 1.5 power/atm. The major conclusion of previous studies remains unchanged, namely, that optically determined NO concentrations may be a factor of 5 larger than probe determined values or may be the same, depending on the combustion gas flow conditions and/or probe design. However, certain optically determined NO concentrations varied by as much as 46 percent from previously reported values because of alterations made to the theory and parameters.

  9. 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.

  10. 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.

  11. 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.

  12. Sorbent performance in fluidized-bed combustors: The effect of calcination on the sulfation rate and calcium utilization

    NASA Astrophysics Data System (ADS)

    Romans, David Earl

    This investigation was conducted to determine if greater amounts of SOsb2 can be removed by naturally occurring sorbents from coal-fired, atmospheric, fluidized-bed combustors if a greater understanding of the calcination process during simultaneous calcination and sulfation is ascertained. A bench-scale, fluidized-bed reactor, ca. 68 insp3 (1,114 cmsp3) of total volume, 1.5 inI.D., utilizing 100 mesh U.S. Standard Sieve stainless steel screens for a frit, was successfully operated and found to be capable of producing the necessary, repeatable data for addressing the hypothesis. Electronic mass flow meters were used to deliver a synthetic gas mixture consisting of Nsb2,\\ Osb2,\\ COsb2, and SOsb2. On-line gas analyzers were used to measure the real-time effluent concentrations of the latter three species. Gas compositions usually consisted of 4 vol% Osb2; 2,000 vppm SOsb2; 0-40 vol% COsb2; and the balance Nsb2. Three limestones with particles sizes ranging from 150 to 1,000 mum were used along with sized quartz for bed material. All experiments were performed at atmospheric pressure with temperatures generally ranging from 850 to 900sp°C. Calcination and sulfation occur simultaneously with respect to time and both the sulfation rate and calcium utilization generally increase with increasing calcination rates. COsb2-induced sintering decreases the calcine surface area while simultaneously increasing the mean diameter pore size; the significant reduction in developed surface area negates any perceived benefit of larger mean diameter pores for increased sorbent performance. Additionally, the use of elevated concentrations of COsb2 in an attempt to minimize the development of an impermeable CaSOsb4 rim via the slow release of COsb2 through the porous solid is generally not a viable method for increasing the sorbent performance (sulfation rate and calcium utilization) since intra-particle, thermally-induced fractures provide access paths for CaO located towards the

  13. 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.

  14. Slagging retrofit pulsed coal combustor: Final report

    SciTech Connect

    Not Available

    1987-01-01

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

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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

  1. Organic Compounds in Turbine Combustor Exhaust.

    DTIC Science & Technology

    1975-09-01

    used were JP—4, JP— 5, JP—8, DEL—5, and meter. The nominal flow was 500 cc/mm at 21.1°C andF isooctane . The JP—4 was tested at inlet pressure. of 760...willS and high—pressure—ratio (PR) idle respectively. ~~~~~~~~ not be trapped and concentrated are those with suffi— isooctane , JP—5, and DEL—5 (an...increasing inlet pressures of JP—5 and JP—5 + Ferrocene at 75 psig 15, 33, and 50 psig; 2) the various fuels (JP—4 , JP—5, S JP—8, isooctane , and DEL—S and

  2. 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.

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

    NASA Astrophysics Data System (ADS)

    Rodriguez, Daniel

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

  4. 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.

  5. 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.

  6. A Simplified Model for Detonation Based Pressure-Gain Combustors

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2010-01-01

    A time-dependent model is presented which simulates the essential physics of a detonative or otherwise constant volume, pressure-gain combustor for gas turbine applications. The model utilizes simple, global thermodynamic relations to determine an assumed instantaneous and uniform post-combustion state in one of many envisioned tubes comprising the device. A simple, second order, non-upwinding computational fluid dynamic algorithm is then used to compute the (continuous) flowfield properties during the blowdown and refill stages of the periodic cycle which each tube undergoes. The exhausted flow is averaged to provide mixed total pressure and enthalpy which may be used as a cycle performance metric for benefits analysis. The simplicity of the model allows for nearly instantaneous results when implemented on a personal computer. The results compare favorably with higher resolution numerical codes which are more difficult to configure, and more time consuming to operate.

  7. 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.

  8. 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.

  9. 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.

  10. Fuel cell system combustor

    DOEpatents

    Pettit, William Henry

    2001-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Variable volume combustor

    DOEpatents

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

    2017-01-17

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

  19. 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.

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

    NASA Technical Reports Server (NTRS)

    Wear, Jerrold D; Butze, Helmut F

    1954-01-01

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

  1. 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.

  2. NASA/GE advanced low emissions combustor program

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

  3. 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.

  4. Transient catalytic combustor model

    NASA Astrophysics Data System (ADS)

    Tien, J. S.

    1981-05-01

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

  5. Combustor liner cooling system

    DOEpatents

    Lacy, Benjamin Paul; Berkman, Mert Enis

    2013-08-06

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

  6. Dual-Mode Combustor

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  7. 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)

  8. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 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...

  9. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 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...

  10. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 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...

  11. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 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...

  12. 14 CFR 23.1123 - Exhaust system.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 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...

  13. 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.

  14. Composite Matrix Experimental Combustor

    DTIC Science & Technology

    1994-04-01

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

  15. 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.

  16. Combustor liner support assembly

    NASA Technical Reports Server (NTRS)

    Halila, Ely E. (Inventor)

    1994-01-01

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

  17. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    A test rig program was conducted with the objective of evaluating and minimizing the exhaust emissions, in particular NO(x), of two advanced aircraft combustor concepts at a simulated, high-altitude cruise condition. The two combustor designs, both members of the lean-reaction, pre-mixed family, are known as the Jet Induced Circulation (JIC) combustor and the Vortex Air Blast (VAB) combustor and were rig tested in the form of reverse flow can combustors in the 0.127-m size range. Various configuration modifications were applied to each of the initial JIC and VAB combustor model designs in an effort to reduce the emissions levels. The VAB combustor demonstrated a NO(x) level of 1.1 gm NO2/kg fuel with essentially 100 percent combustion efficiency at the simulated cruise combustor condition of 507 kPa, 833 K inlet pressure and temperature, respectively and 1778 K outlet temperature on Jet-A1 fuel. In addition, emissions data were obtained at low combustor inlet pressure and temperatures that indicate the potential performance at engine off-design conditions.

  5. 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.

  6. 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.

  7. 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.

  8. Combustor burner vanelets

    DOEpatents

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

    2012-02-14

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

  9. 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.

  10. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  11. Design and development test of a gas turbine combustor for air-blow Lurgi coal gas fuel

    SciTech Connect

    Beebe, K.W.; Ye, L.J.

    1985-01-01

    A heavy-duty industrial gas turbine combustion system has been designed and laboratory tested for use with low heating value coal gas produced by an air-blown Lurgi coal gasifier. The design fuel has a nominal lower heating value of 4.21 MJ/Nm/sup 3/ (107 BTU/SCF). The combustor design utilizes high-swirl fuel and air injection to provide rapid fuel/air mixing and a stable flame front, and is physically interchangeable with the conventional fuel combustion system for the General Electric Model MS5001 gas turbine. Full-pressure, full-scale tests of the new combustion system have been conducted at the General Electric Gas Turbine Development Laboratory in Schenectady, New York, USA, for the Shanghai Power Plant Equipment Research Institute (SPPERI) of the Peoples Republic of China. Simulated clean low heating value coal gas fuel with a composition specified by SPPERI was used for these tests. Laboratory test reults are presented for important combustor operating parameters, including exhaust emissions, combustion efficiency, exhaust temperature profile, dynamic pressure, and metal temperature distribution.

  12. Power Generation Efficiency of Photovoltaics and a SOFC-PEFC Combined Micro-grid with Time Shift Utilization of the SOFC Exhaust Heat

    NASA Astrophysics Data System (ADS)

    El-Sayed, Abeer Galal; Obara, Shin'ya

    In this study, the combined system of a solid-oxide fuel cell (SOFC) and a proton-exchange membrane fuel cell (PEFC) is developed. The proposed system consists of a SOFC-PEFC combined system and a photovoltaic system (PV) as the energy supply to a micro-grid. The exhaust heat of the SOFC is used for the steam reforming of the bio-ethanol gas with time shift utilization of the exhaust heat of the SOFC in optional time. The SOFC-PEFC combined system with the PV was introduced in a micro-grid of 30 residences in Sapporo, Japan. The operation plan of the system has three cases: without solar power, with 50% and with 100% of solar output power. Moreover, three types of system operation of using the SOFC independent operation, PEFC independent operation and SOFC-PEFC combined system are used to supply the demand side. A comparative study between the types of system operation is presented. The power generation efficiency is investigated for different load patterns: average load pattern, compressed load pattern and extended load pattern. This paper reported that the power generation efficiencies of the proposedsystem in consideration of these load patterns are 27% to 48%.

  13. 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.

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

    NASA Technical Reports Server (NTRS)

    1975-01-01

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

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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

  20. 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.

  1. 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

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Segmented annular combustor

    DOEpatents

    Reider, Samuel B.

    1979-01-01

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

  8. 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.

  9. Experimental clean combustor program, phase 1

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  10. Operating Characteristics of a Fluidic Premixed Dump Combustor

    NASA Astrophysics Data System (ADS)

    Ahmed, Kareem; Carr, Zakery; Forliti, David

    2007-11-01

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

  11. 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.

  12. 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.

  13. 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

  14. 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.

  15. 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.

  16. Combustor and method for distributing fuel in the combustor

    SciTech Connect

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

    2016-04-26

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

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

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

    NASA Technical Reports Server (NTRS)

    Cutrone, M. B.

    1981-01-01

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

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  8. 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.

  9. 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.

  10. 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.

  11. Pollution emissions from single swirl-can combustor modules at parametric test conditions

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Exhaust pollutant emissions were measured from single swirl-can combustor modules operating over a pressure range of 69 to 276 N/sq cm (100 to 400 psia), over a fuel-air ratio range of 0.01 to 0.04, at an inlet air temperature of 733 K (860 F), and at a constant reference velocity of 23.2 m/sec). Many swirl-can module designs were evaluated; the 11 most promising designs exhibited oxides of nitrogen emission levels lower than that from conventional gas-turbine combustors. Although these single module test results are not necessarily indicative of the performance characteristics of a large array of modules, the results are very promixing and offer a number of module designs that should be tested in a full combustor.

  12. 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.

  13. Multi-port dump combustor

    SciTech Connect

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

    1985-09-10

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

  14. Premixed Prevaporized Combustor Technology Forum

    NASA Technical Reports Server (NTRS)

    1979-01-01

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

  15. 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.

  16. 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.

  17. 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.

  18. Heat Transfer Experiments on a Pulse Detonation Driven Combustor

    DTIC Science & Technology

    2011-03-01

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

  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. 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.

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

    NASA Astrophysics Data System (ADS)

    Hasler, Anthony Scott

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

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

    DOEpatents

    Wijmans, Johannes G [Menlo Park, CA; Merkel, Timothy C [Menlo Park, CA; Baker, Richard W [Palo Alto, CA

    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.

  3. Integrated CFD modeling of gas turbine combustors

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  4. Variable residence time vortex combustor

    DOEpatents

    Melconian, Jerry O.

    1987-01-01

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

  5. Investigation of the mechanism in RIJKE pulse combustors with tangential air and fuel injection. Progress report, August 1, 1992--January 31, 1993

    SciTech Connect

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

    1993-02-01

    This report summarizes the accomplishments of DOE Contract No. DE-AS04-85AL31881. This three year investigation started in August 1989 and its objective was to elucidate the mechanisms that control the driving of pulsations in the liquid fuel burning, Rijke type, pulse combustor developed under a preceding DOE contracts. It was demonstrated in that contract that the developed Rijke type pulse combustor can burn a variety of light and heavy liquid fuel oils with high combustion efficiencies while using low excess air, which produces high thermal efficiencies. Since the elucidation of the driving mechanism in the Rijke pulse combustor required the use of optical diagnostics (e.g., radiation measurements), it was decided to perform these investigations in a Rijke pulse combustor that burned propane instead of a liquid fuel in order to avoid difficulties that are often encountered due to the presence of liquid droplets in the combustion region. Consequently, an effort was made to develop a Rijke pulse combustor that is similar to the one developed in the preceding program and demonstrated similar performance characteristics. Such a pulse combustor was developed in the early phases of this program. The developed experimental setup was provided with capabilities for measuring steady combustor temperature distributions, the characteristics of the excited pressure oscillations, the exhaust flow composition, the characteristics of the flow field and the reaction rates. This pulse combustor consists of a cylindrical tube that is attached to a decoupling chamber at each end. Fuel and air are supplied via a tangential air/fuel injection system that is located at a distance of L/4 from the combustor entrance, where L is the combustor length. Part of the combustor tube, where combustion occurs, is water cooled. This section is also equipped with flat quartz windows to permit optical diagnostics.

  6. 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.

  7. Combustor with multistage internal vortices

    DOEpatents

    Shang, Jer Yu; Harrington, R.E.

    1987-05-01

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

  8. Vertical combustor for particulate refuse

    NASA Astrophysics Data System (ADS)

    Chung, P. M.; Carlson, L.

    1981-03-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  10. 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.

  11. Driving of combustor oscillations by gaseous propellant injectors

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  13. 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.

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

    SciTech Connect

    Not Available

    1993-05-01

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

  15. 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.

  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

    DOEpatents

    Kim, Won -Wook; McMahan, Kevin Weston

    2015-09-29

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

  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. Capacity control of power stations by O 2/H 2 rocket combustor technology

    NASA Astrophysics Data System (ADS)

    Sternfeld, Ing. H. J.

    1995-10-01

    The concept of a hydrogen/oxygen spinning reserve system is described. The novel component of this concept is a socalled hydrogen/oxygen steam generator derived from modern H 2/O 2 rocket combustor technology. With the HYDROSS-project the DLR and German power plant industries as well as electric utilities have converted the rocket combustor technology to a power plant component. The status of the project as well as technical problems encountered with the conversion are described. Finally, future options for utilizing H 2/O 2 steam generator technology for stand-by and peak-load power plants are discussed.

  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. 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.

  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. 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.

  5. Rich burn combustor technology at Pratt and Whitney

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  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. 14 CFR 23.1125 - Exhaust heat exchangers.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes the following apply: (a... suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;...

  8. 14 CFR 23.1125 - Exhaust heat exchangers.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes the following apply: (a... suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;...

  9. 14 CFR 23.1125 - Exhaust heat exchangers.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes the following apply: (a... suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;...

  10. 14 CFR 23.1125 - Exhaust heat exchangers.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes the following apply: (a... suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;...

  11. 14 CFR 23.1125 - Exhaust heat exchangers.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Exhaust System § 23.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes the following apply: (a... suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;...

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  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. 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.

  20. 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.

  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. 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.

  3. 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.

  4. 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.

  5. Experimental clean combustor program: Noise study

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  6. 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.

  7. Application of Gas Analysis to Combustor Research

    NASA Technical Reports Server (NTRS)

    Hibbard, R. R.; Evans, Albert

    1959-01-01

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

  8. 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.

  9. Operational Characteristics of an Ultra Compact Combustor

    DTIC Science & Technology

    2014-03-27

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

  10. Hypermixer Pylon Fuel Injection for Scramjet Combustors

    DTIC Science & Technology

    2008-09-11

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

  11. Multiducted Inlet Combustor Research and Development.

    DTIC Science & Technology

    1982-10-01

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

  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. 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.

  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. 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.

  16. Rolling contact mounting arrangement for a ceramic combustor

    DOEpatents

    Boyd, Gary L.; Shaffer, James E.

    1995-01-01

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

  17. Rolling contact mounting arrangement for a ceramic combustor

    DOEpatents

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

    1995-10-17

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

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

    NASA Astrophysics Data System (ADS)

    Mueller, Michael E.; Pitsch, Heinz

    2013-11-01

    -to-air ratio, the maximum soot volume fraction is found inside the recirculation zone; at the lower fuel-to-air ratio, turbulent fluctuations in the mixture fraction promote the oxidation of soot inside the recirculation zone and suppress the accumulation of a large soot volume fraction. Downstream, soot exits the combustor in intermittent fuel-rich pockets that are not mixed during the injection of dilution air and subsequent secondary fuel-lean combustion. At the higher fuel-to-air ratio, the frequency of these fuel-rich pockets is increased, leading to higher soot emissions from the combustor. Quantitatively, the soot emissions from the combustor are overpredicted by about 50%, which is a substantial improvement over previous works utilizing RANS to predict such emissions. In addition, the ratio between the two fuel-to-air ratios predicted by LES compares very favorably with the experimental measurements. Furthermore, soot growth is dominated by an acetylene-based pathway rather than an aromatic-based pathway, which is usually the dominant mechanism in nonpremixed flames. This finding is the result of the interactions between the hydrodynamics, mixing, chemistry, and soot in the recirculation zone and the resulting residence times of soot at various mixture fractions (compositions), which are not the same in this complex recirculating flow as in nonpremixed jet flames.

  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. Frequency Domain Analysis of a Gas Fired Mechanically Valved Pulse Combustor

    NASA Astrophysics Data System (ADS)

    Neumeier, Yedidia

    1993-01-01

    This study presents an analysis of a mechanically valved pulse combustor in the frequency domain. The pulse combustor is treated as a feedback system. The forward branch of the system consists of the acoustic resonator, while the feedback loop consists of the combustion heat release process and the heat loss through the pulse combustor walls. The model of the acoustic resonator is based upon an analysis of the conservation equation for the combustion chamber energy oscillations and the tail pipe acoustics. The present study utilizes a phenomenological approach to model the periodic heat release process in the combustor. The model accounts for the mixing of cold reactants with hot products, flame extinction due to high velocities during the injection time, and the kinetics of the chemical reaction. The results of this study indicate that a high periodic heat transfer must take place in the combustion zone of mechanically valved pulse combustors. The intensity of this oscillating heat transfer is an order of magnitude larger than has been observed until the present study. The results imply that the unsteady heat transfer contributes vital feedback that stabilizes the limit cycle of the pulse combustor. The study was extended to include preliminary experimental evidence for the existence of a very high wall heat losses in the combustion zone, as predicted by the model. No model of pulse combustor that fails to account for unsteady heat transfer in the combustion zone can be considered fully accurate. A calibration procedure is developed to optimize the model prediction. This calibration process is oriented to provide an accurate prediction of the important parameters; namely, the pressure level and the operating frequency, at the cost of less accuracy in predicting relatively less important parameters. The model developed in the study has been used to develop a software package application for design and analysis of pulse combustors. The use of the software for pulse

  1. An investigation of the effects of smoke suppressant fuel additives on engine and test cell exhaust gas opacities. Final report for 1981

    SciTech Connect

    Thornburg, D.W.; Darnell, T.R.; Netzer, D.W.

    1982-05-01

    Tests were conducted in a one-eighth scale turbojet test cell with a ramjet type combustor to investigate the effects of fuel additives on smoke reduction. Particle size and mass concentrations were determined at the engine and stack exhausts using three wavelength optical detector systems. Particulate samples were also collected at the engine exhaust and analyzed with a scanning electron microscope. Combustor temperature and fuel additives were found to significantly affect particulate mass concentrations emitted from the engine while particle size appeared to be unaffected. No significant changes in the particulate size or mass occurred from the engine exhaust to the stack exhaust. The optical determination of exhaust mean particulate size/mass concentration with three wavelength optical detector systems appears to be reasonably accurate technique for evaluating the effects of engine and test cell operating conditions and fuel composition changes on the emitted particulates.

  2. 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.

  3. 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

  4. Flame dynamics in a micro-channeled combustor

    NASA Astrophysics Data System (ADS)

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

    2015-01-01

    The increasing use of Micro-Electro-Mechanical Systems (MEMS) has generated a significant interest in combustion-based power generation technologies, as a replacement of traditional electrochemical batteries which are plagued by low energy densities, short operational lives and low power-to-size and power-to-weight ratios. Moreover, the versatility of integrated combustion-based systems provides added scope for combined heat and power generation. This paper describes a study into the dynamics of premixed flames in a micro-channeled combustor. The details of the design and the geometry of the combustor are presented in the work by Kariuki and Balachandran [1]. This work showed that there were different modes of operation (periodic, a-periodic and stable), and that in the periodic mode the flame accelerated towards the injection manifold after entering the channels. The current study investigates these flames further. We will show that the flame enters the channel and propagates towards the injection manifold as a planar flame for a short distance, after which the flame shape and propagation is found to be chaotic in the middle section of the channel. Finally, the flame quenches when it reaches the injector slots. The glow plug position in the exhaust side ignites another flame, and the process repeats. It is found that an increase in air flow rate results in a considerable increase in the length (and associated time) over which the planar flame travels once it has entered a micro-channel, and a significant decrease in the time between its conversion into a chaotic flame and its extinction. It is well known from the literature that inside small channels the flame propagation is strongly influenced by the flow conditions and thermal management. An increase of the combustor block temperature at high flow rates has little effect on the flame lengths and times, whereas at low flow rates the time over which the planar flame front can be observed decreases and the time of

  5. Performance of low-Btu fuel gas turbine combustors

    SciTech Connect

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

    1995-11-01

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

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

    NASA Technical Reports Server (NTRS)

    Emmerling, J. J.

    1975-01-01

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

  7. Multi-Ducted Inlet Combustor Research and Development.

    DTIC Science & Technology

    1983-11-01

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

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Flow establishment in a generic scramjet combustor

    NASA Astrophysics Data System (ADS)

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

    1990-10-01

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

  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. 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.

  15. Numerical prediction of low frequency combustion instability in a model ramjet combustor

    SciTech Connect

    Shang, H.M.; Chen, Y.S.; Shih, M.S.; Farmer, R.C.

    1996-12-31

    A numerical analysis has been conducted for low-frequency combustion instability in a model ramjet combustor. The facility is two-dimensional, and is comprised of a long inlet duct, a dump combustor chamber, and an exhaust nozzle. The experiments observed that the combustor pressure oscillation under the particular operating condition did not have much cycle-to-cycle variation. The main resonant frequency occurs at about 65 Hz for this case. In the numerical analysis, a time accurate Computational Fluid Dynamics (CFD) code with a pressure-correction algorithm is used, and the combustion process was modeled with a single step chemistry model and a modified eddy breakup model. A high-order upwind scheme with flux limiter is used for convection terms. The convergence of the linear algebraic equations is accelerated through a preconditioned conjugate gradient matrix solver. The numerical predictions show that the flame oscillates in the combustion chamber at the calculation condition and are justified by the experimental schlieren photographs. The numerical analyses correctly predict the chamber pressure oscillation frequency is over-predicted compared with the experimental data. The discrepancy can be explained by the simplified turbulence and combustion model used in this study, and the uncertainty of the inlet boundary conditions.

  16. 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.

  17. 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.

  18. Laser diagnostics on a hypersonic combustor

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  19. Low NO.sub.x combustor

    DOEpatents

    Taylor, Jack R.

    1987-01-01

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

  20. Variable volume combustor with aerodynamic support struts

    DOEpatents

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

    2017-03-07

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and providing the flow of fuel therethrough. The support struts may include an aerodynamic contoured shape so as to distribute evenly a flow of air to the micro-mixer fuel nozzles.

  1. Micro-combustor for gas turbine engine

    DOEpatents

    Martin, Scott M.

    2010-11-30

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

  2. CONTINUED DEVELOPMENT OF THE ROTARY COMBUSTOR FOR REFIRING PULVERIZED COAL BOILERS

    SciTech Connect

    Murray F. Abbott; Jamal B. Mereb; Simon P. Hanson; Michael J. Virr

    2000-11-01

    were higher those targeted for the device which were 100 ppmv. This was assumed to be because of the aforementioned temperature spiking. The primary operating problem remains control of the fluid-bed temperature. Although improvements were made, steam flow control was manual, and very coarse. To accomplish this will require finer control of the steam flow to the rotary drum air plenum, and development of an algorithm for automatic control using the Moore APACS{trademark}. This is the recommended succeeding step in the development of the Rotary Combustor for industrial or utility use.

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

    DTIC Science & Technology

    1994-06-01

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

  4. System and method for controlling a combustor assembly

    DOEpatents

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

    2013-03-05

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

  5. Eye and respiratory irritants in jet engine exhaust.

    PubMed

    Miyamoto, Y

    1986-11-01

    It has been noted that eye and respiratory irritation frequently occurred in the ground crews and pilots working on the field behind an aircraft with a low smoke combustor (LSC) engine. This study was attempted to analyze the exhaust sampled at about 50 m behind the LSC J79 engines at idle power setting by means of a high performance liquid chromatography (HPLC) technique. Nine kinds of lower aliphatic carbonyl compound (seven aldehydes and two ketones) were identified. The concentration of formaldehyde was the highest among them, showing the value above the threshold reported by previous investigators. Concentration of NOx was simultaneously measured by a gas detector tube in the same sample. The exhaust of a conventional J79 engine, which has rarely caused irritation, was also analyzed by the same technique and the results were compared. It was concluded that formaldehyde plays a major role in causing irritation.

  6. Effect of Fuel on Performance of a Single Combustor of an I-16 Turbojet Engine at Simulated Altitude Conditions

    NASA Technical Reports Server (NTRS)

    Zettle, Eugene V; Bolz, Ray E; Dittrich, R T

    1947-01-01

    As part of a study of the effects of fuel composition on the combustor performance of a turbojet engine, an investigation was made in a single I-16 combustor with the standard I-16 injection nozzle, supplied by the engine manufacturer, at simulated altitude conditions. The 10 fuels investigated included hydrocarbons of the paraffin olefin, naphthene, and aromatic classes having a boiling range from 113 degrees to 655 degrees F. They were hot-acid octane, diisobutylene, methylcyclohexane, benzene, xylene, 62-octane gasoline, kerosene, solvent 2, and Diesel fuel oil. The fuels were tested at combustor conditions simulating I-16 turbojet operation at an altitude of 45,000 feet and at a rotor speed of 12,200 rpm. At these conditions the combustor-inlet air temperature, static pressure, and velocity were 60 degrees F., 12.3 inches of mercury absolute, and 112 feet per second respectively, and were held approximately constant for the investigation. The reproducibility of the data is shown by check runs taken each day during the investigation. The combustion in the exhaust elbow was visually observed for each fuel investigated.

  7. Active Control of High-Frequency Combustor Instability Demonstrated

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Chang, Clarence T.

    2003-01-01

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

  8. Combustor for a low-emissions gas turbine engine

    DOEpatents

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

    2000-01-01

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

  9. Characterization of nonlinear heat release-acoustic interactions in gas turbine combustors

    NASA Astrophysics Data System (ADS)

    Bellows, Benjamin D.

    This thesis describes an experimental investigation of the flame transfer function between flow disturbances and heat release oscillations in lean, premixed combustors. This research effort was motivated by the fact that modern gas turbines, operating fuel lean to minimize exhaust emissions, are susceptible to self-excited combustion oscillations. These instabilities generally occur when the unsteady combustion process couples with the acoustic modes of the combustion chamber. The resultant flow and structural vibrations can substantially reduce hot section part life. As such, avoiding operating regimes where high dynamics occur often requires operating at lower power outputs and/or higher pollutant emissions than the turbine is otherwise capable. This work demonstrated nonlinearities in the chemiluminescence response at large amplitude velocity oscillations in a turbulent, swirling flame. It is observed that the nonlinear flame response can exhibit a variety of behaviors, both in the shape of the response curve and the forcing amplitude at which nonlinearity is first observed depending on the operating conditions of the combustor. The phase between the flow oscillations and heat release is also seen to have substantial amplitude dependence. In addition, the interactions between the fundamental frequency and the higher and subharmonics of the measured signals can significantly influence the flame as well as the frequency response of the system. The nonlinear flame dynamics are governed by different mechanisms in different frequency and flowrate regimes. Three mechanisms, vortex rollup, unsteady flame liftoff, and parametric instability, are identified to influence the nonlinear flame response in these combustors. Analysis of the results shows that the mechanisms responsible for nonlinearity m the flame response are influenced by the Strouhal number, the mean velocity at the combustor dump plane, and the ratio of the oscillating velocity amplitude to the laminar

  10. Flashback Arrestor for LPP, Low NOx Combustors

    NASA Technical Reports Server (NTRS)

    Kraemer, Gil; Lee, Chi-Ming

    1998-01-01

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

  11. Stably operating pulse combustor and method

    DOEpatents

    Zinn, B.T.; Reiner, D.

    1990-05-29

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

  12. Steam Reformer With Fibrous Catalytic Combustor

    NASA Technical Reports Server (NTRS)

    Voecks, Gerald E.

    1987-01-01

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

  13. Stably operating pulse combustor and method

    DOEpatents

    Zinn, Ben T.; Reiner, David

    1990-01-01

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

  14. Near-zero emissions combustor system for syngas and biofuels

    SciTech Connect

    Yongho, Kim; Rosocha, Louis

    2010-01-01

    A multi-institutional plasma combustion team was awarded a research project from the DOE/NNSA GIPP (Global Initiative for Prolifereation Prevention) office. The Institute of High Current Electronics (Tomsk, Russia); Leonardo Technologies, Inc. (an American-based industrial partner), in conjunction with the Los Alamos National Laboratory are participating in the project to develop novel plasma assisted combustion technologies. The purpose of this project is to develop prototypes of marketable systems for more stable and cleaner combustion of syngas/biofuels and to demonstrate that this technology can be used for a variety of combustion applications - with a major focus on contemporary gas turbines. In this paper, an overview of the project, along with descriptions of the plasma-based combustors and associated power supplies will be presented. Worldwide, it is recognized that a variety of combustion fuels will be required to meet the needs for supplying gas-turbine engines (electricity generation, propulsion), internal combustion engines (propulsion, transportation), and burners (heat and electricity generation) in the 21st Century. Biofuels and biofuel blends have already been applied to these needs, but experience difficulties in modifications to combustion processes and combustor design and the need for flame stabilization techniques to address current and future environmental and energy-efficiency challenges. In addition, municipal solid waste (MSW) has shown promise as a feedstock for heat and/or electricity-generating plants. However, current combustion techniques that use such fuels have problems with achieving environmentally-acceptable air/exhaust emissions and can also benefit from increased combustion efficiency. This project involves a novel technology (a form of plasma-assisted combustion) that can address the above issues. Plasma-assisted combustion (PAC) is a growing field that is receiving worldwide attention at present. The project is focused on

  15. Film Cooling Flow Effects on Post-Combustor Trace Chemistry

    NASA Technical Reports Server (NTRS)

    Wey, Thomas; Liu, Nan-Suey

    2003-01-01

    Film cooling injection is widely applied in the thermal design of turbomachinery, as it contributes to achieve higher operating temperature conditions of modern gas turbines, and to meet the requirements for reliability and life cycles. It is a significant part of the high-pressure turbine system. The film cooling injection, however, interacts with the main flow and is susceptible to have an influence on the aerodynamic performance of the cooled components, and through that may cause a penalty on the overall efficiency of the gas turbine. The main reasons are the loss of total pressure resulting from mixing the cooling air with mainstream and the reduction of the gas stagnation temperature at the exit of the combustion chamber to a lower value at the exit of nozzle guide vane. In addition, the impact of the injected air on the evolution of the trace species of the hot gas is not yet quite clear. This work computationally investigates the film cooling influence on post-combustor trace chemistry, as trace species in aircraft exhaust affect climate and ozone.

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

    NASA Astrophysics Data System (ADS)

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

    2016-10-01

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

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

    SciTech Connect

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

    1992-12-31

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

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

    SciTech Connect

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

    1992-01-01

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

  19. Short review on heat recovery from exhaust gas

    NASA Astrophysics Data System (ADS)

    Jaber, Hassan; Khaled, Mahmoud; Lemenand, Thierry; Ramadan, Mohamad

    2016-07-01

    The increasing growth of energy demand leads to issues associated with energy demand reduction and propose new energy efficient solutions. Heat recovery consists the most promising solution especially in regions where renewable energy resources are not available. That is why the domain of heat recovery has shown a tremendous improvement during the recent years. On the other hand, few works have been dedicated to heat recovery from exhaust gas. This paper presents a review on heat recovery from exhaust gas. The authors propose to classify exhaust gas heat recovery systems within three different classifications that are exhaust gas temperature, utilized equipment and recovery purposes.

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

    SciTech Connect

    2009-11-01

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

  1. National Combustion Code, a Multidisciplinary Combustor Design System, Will Be Transferred to the Commercial Sector

    NASA Technical Reports Server (NTRS)

    Steele, Gynelle C.

    1999-01-01

    The NASA Lewis Research Center and Flow Parametrics will enter into an agreement to commercialize the National Combustion Code (NCC). This multidisciplinary combustor design system utilizes computer-aided design (CAD) tools for geometry creation, advanced mesh generators for creating solid model representations, a common framework for fluid flow and structural analyses, modern postprocessing tools, and parallel processing. This integrated system can facilitate and enhance various phases of the design and analysis process.

  2. Analysis on Impulse Characteristics of PDRE with Exhaust Measurements

    NASA Astrophysics Data System (ADS)

    Hu, Hong-bo; Weng, Chun-sheng; Lv, Xiao-jing; Li, Ning

    2014-06-01

    The exhaust characteristics related to impulse was investigated in a pulse detonation rocket engine (PDRE) by tunable diode laser absorption sensing system. The instantaneous parameters of temperature, velocity and pressure were obtained for exhaust at engine exit. Analysis on impulse characteristics based on control volume of the PDRE was conducted for a single operation circle with experimental results. It was concluded that the impulse (3.26 N·s) achieved by exhaust measurements was in agreement with that (3.09 N·s) by a load cell. The impulse caused by exhaust momentum experienced an extremely sharp ascending, a steep rising and a slow increment in sequence. The exhausts during the sharp ascending and steep rising were under expansion with high mass weighted average temperature (>1266 K), so there was a possible promotion for exhausts utilizing.

  3. Measuring Carbon Monoxide in Auto Exhaust by Gas Chromatography.

    ERIC Educational Resources Information Center

    Jaffe, Dan; Herndon, Scott

    1995-01-01

    Presents a simple and reliable technique using commonly available equipment for monitoring carbon monoxide in automobile exhaust. The experiment utilizes a gas chromatograph and a thermal conductivity detector (TCD). (DDR)

  4. Systems Characterization of Combustor Instabilities With Controls Design Emphasis

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2004-01-01

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

  5. Advanced composite combustor structural concepts program

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  6. Fuel property effects in stirred combustors

    NASA Technical Reports Server (NTRS)

    1980-01-01

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

  7. Advanced Catalytic Combustors for Low Pollutant Emissions

    DTIC Science & Technology

    1979-11-01

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

  8. Laser Diagnostic System Validation and Ultra-Compact Combustor Characterization

    DTIC Science & Technology

    2008-03-01

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

  9. Low NOx Fuel Flexible Combustor Integration Project Overview

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  10. Porous Media Combustors for Clean Gas Turbine Engines

    DTIC Science & Technology

    2007-11-02

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

  11. Advanced liner-cooling techniques for gas turbine combustors

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  12. High-temperature durability considerations for HSCT combustor

    NASA Technical Reports Server (NTRS)

    Jacobson, Nathan S.

    1992-01-01

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

  13. Rapid-quench axially staged combustor

    DOEpatents

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

    1999-01-01

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

  14. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

    W. R. Laster; E. Anoshkina

    2008-01-31

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

  15. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

    Laster, W. R.; Anoshkina, E.

    2008-01-31

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

  16. Catalytic Combustor for Fuel-Flexible Turbine

    SciTech Connect

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

    2006-03-31

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

  17. Hyperventilation and exhaustion syndrome.

    PubMed

    Ristiniemi, Heli; Perski, Aleksander; Lyskov, Eugene; Emtner, Margareta

    2014-12-01

    Chronic stress is among the most common diagnoses in Sweden, most commonly in the form of exhaustion syndrome (ICD-10 classification - F43.8). The majority of patients with this syndrome also have disturbed breathing (hyperventilation). The aim of this study was to investigate the association between hyperventilation and exhaustion syndrome. Thirty patients with exhaustion syndrome and 14 healthy subjects were evaluated with the Nijmegen Symptom Questionnaire (NQ). The participants completed questionnaires about exhaustion, mental state, sleep disturbance, pain and quality of life. The evaluation was repeated 4 weeks later, after half of the patients and healthy subjects had engaged in a therapy method called 'Grounding', a physical exercise inspired by African dance. The patients reported significantly higher levels of hyperventilation as compared to the healthy subjects. All patients' average score on NQ was 26.57 ± 10.98, while that of the healthy subjects was 15.14 ± 7.89 (t = -3.48, df = 42, p < 0.001). The NQ scores correlated strongly with two measures of exhaustion (Karolinska Exhaustion Scale KES r = 0.772, p < 0.01; Shirom Melamed Burnout Measure SMBM r = 0.565, p < 0.01), mental status [Hospital Anxiety and Depression Score (HADS) depression r = 0.414, p < 0.01; HADS anxiety r = 0.627, p < 0.01], sleep disturbances (r = -0.514, p < 0.01), pain (r = -.370, p < 0.05) and poor well-being (Medical Outcomes Survey Short Form 36 questionnaire- SR Health r = -0.529, p < 0.05). In the logistic regression analysis, the variance in the scores from NQ were explained to a high degree (R(2) = 0.752) by scores in KES and HADS. The brief Grounding training contributed to a near significant reduction in hyperventilation (F = 2.521, p < 0.124) and to significant reductions in exhaustion scores and scores of depression and anxiety. The conclusion is that hyperventilation is common in exhaustion syndrome patients and that it can be reduced by systematic physical therapy

  18. Controlled pilot oxidizer for a gas turbine combustor

    DOEpatents

    Laster, Walter R.; Bandaru, Ramarao V.

    2010-07-13

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

  19. Apparatus and method for cooling a combustor cap

    SciTech Connect

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

    2014-04-29

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

  20. Stagnation point reverse flow combustor

    NASA Technical Reports Server (NTRS)

    Zinn, Ben T. (Inventor); Neumeier, Yedidia (Inventor); Seitzman, Jerry M. (Inventor); Jagoda, Jechiel (Inventor); Weksler, Yoav (Inventor)

    2008-01-01

    A method for combusting a combustible fuel includes providing a vessel having an opening near a proximate end and a closed distal end defining a combustion chamber. A combustible reactants mixture is presented into the combustion chamber. The combustible reactants mixture is ignited creating a flame and combustion products. The closed end of the combustion chamber is utilized for directing combustion products toward the opening of the combustion chamber creating a reverse flow of combustion products within the combustion chamber. The reverse flow of combustion products is intermixed with combustible reactants mixture to maintain the flame.

  1. Exhaust bypass flow control for exhaust heat recovery

    DOEpatents

    Reynolds, Michael G.

    2015-09-22

    An exhaust system for an engine comprises an exhaust heat recovery apparatus configured to receive exhaust gas from the engine and comprises a first flow passage in fluid communication with the exhaust gas and a second flow passage in fluid communication with the exhaust gas. A heat exchanger/energy recovery unit is disposed in the second flow passage and has a working fluid circulating therethrough for exchange of heat from the exhaust gas to the working fluid. A control valve is disposed downstream of the first and the second flow passages in a low temperature region of the exhaust heat recovery apparatus to direct exhaust gas through the first flow passage or the second flow passage.

  2. Hybrid Exhaust Component

    NASA Technical Reports Server (NTRS)

    Pelletier, Gerard D. (Inventor); Logan, Charles P. (Inventor); McEnerney, Bryan William (Inventor); Haynes, Jeffrey D. (Inventor)

    2015-01-01

    An exhaust includes a wall that has a first composite material having a first coefficient of thermal expansion and a second composite material having a second coefficient of the thermal expansion that is less than the first coefficient of thermal expansion.

  3. Diesel engine exhaust

    Integrated Risk Information System (IRIS)

    Diesel engine exhaust ; CASRN N.A . Human health assessment information on a chemical substance is included in the IRIS database only after a comprehensive review of toxicity data , as outlined in the IRIS assessment development process . Sections I ( Health Hazard Assessments for Noncarcinogenic Ef

  4. Co-combustion of agricultural residues with coal in a fluidized bed combustor.

    PubMed

    Ghani, W A W A K; Alias, A B; Savory, R M; Cliffe, K R

    2009-02-01

    Power generation from biomass is an attractive technology that utilizes agricultural residual waste. In order to explain the behavior of biomass-fired fluidized bed incinerator, biomass sources from agricultural residues (rice husk and palm kernel) were co-fired with coal in a 0.15m diameter and 2.3m high fluidized bed combustor. The combustion efficiency and carbon monoxide emissions were studied and compared with those for pure coal combustion. Co-combustion of a mixture of biomass with coal in a fluidized bed combustor designed for coal combustion increased combustion efficiency up to 20% depending upon excess air levels. Observed carbon monoxide levels fluctuated between 200 and 900 ppm with the addition of coal. It is evident from this research that efficient co-firing of biomass with coal can be achieved with minimal modifications to existing coal-fired boilers.

  5. Computations of spray, fuel-air mixing, and combustion in a lean-premixed-prevaporized combustor

    NASA Technical Reports Server (NTRS)

    Dasgupta, A.; Li, Z.; Shih, T. I.-P.; Kundu, K.; Deur, J. M.

    1993-01-01

    A code was developed for computing the multidimensional flow, spray, combustion, and pollutant formation inside gas turbine combustors. The code developed is based on a Lagrangian-Eulerian formulation and utilizes an implicit finite-volume method. The focus of this paper is on the spray part of the code (both formulation and algorithm), and a number of issues related to the computation of sprays and fuel-air mixing in a lean-premixed-prevaporized combustor. The issues addressed include: (1) how grid spacings affect the diffusion of evaporated fuel, and (2) how spurious modes can arise through modelling of the spray in the Lagrangian computations. An upwind interpolation scheme is proposed to account for some effects of grid spacing on the artificial diffusion of the evaporated fuel. Also, some guidelines are presented to minimize errors associated with the spurious modes.

  6. NOx results from two combustors tested on medium BTU coal gas

    NASA Technical Reports Server (NTRS)

    Sherlock, T. P.; Carl, D. E.; Vermes, G.; Schwab, J.; Notardonato, J. J.

    1982-01-01

    The results of tests of two combustor configurations using coal gas from a 25 ton/day fluidized bed coal gasifier are reported. The trials were run with a ceramic-lined, staged rich/lean burner and an integral, all metal multiannular swirl burner (MASB) using a range of temperatures and pressures representative of industrial turbine inlet conditions. A lean mixture was examined at 104, 197, and 254 Btu/Scf, yielding NO(x) emissions of 5, 20, and 70 ppmv, respectively. The MASB was employed only with a gas rated at 220-270 Btu/Scf, producing 80 ppmv NO(x) at rated engine conditions. The results are concluded to be transferrable to current machines. Further tests on the effects of gas composition, the scaling of combustors to utility size, and the development of improved wall cooling techniques and variable geometry are indicated.

  7. Numerical Analysis and Optimization of the Ultra Compact Combustor

    DTIC Science & Technology

    2005-03-01

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

  8. Variable volume combustor with nested fuel manifold system

    SciTech Connect

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

    2016-09-13

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

  9. Combustor nozzle for a fuel-flexible combustion system

    DOEpatents

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

    2011-03-22

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

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

    NASA Technical Reports Server (NTRS)

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

    1957-01-01

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

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

    DOEpatents

    Abreu, Mario E.; Kielczyk, Janusz J.

    2001-01-01

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

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

    SciTech Connect

    Sjoblom, B.G.A.

    1982-01-01

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

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

    SciTech Connect

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

    2016-09-06

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

  14. Rayleigh/Raman/LIF measurements in a turbulent lean premixed combustor

    SciTech Connect

    Nandula, S.P.; Pitz, R.W.; Barlow, R.S.; Fiechtner, G.J.

    1995-12-31

    Much of the industrial electrical generation capability being added worldwide is gas-turbine engine based and is fueled by natural gas. These gas-turbine engines use lean premixed (LP) combustion to meet the strict NO{sub x} emission standards, while maintaining acceptable levels of CO. In conventional, diffusion flame gas turbine combustors, large amount of NO{sub x} forms in the hot stoichiometric zones via the Zeldovich (thermal) mechanism. Hence, lean premixed combustors are rapidly becoming the norm, since they are specifically designed to avoid these hot stoichiometric zones and the associated thermal NO{sub x}. However, considerable research and development are still required to reduce the NO{sub x} levels (25-40 ppmvd adjusted to 15% O{sub 2} with the current technology), to the projected goal of under 10 ppmvd by the turn of the century. Achieving this objective would require extensive experiments in LP natural gas (or CH{sub 4}) flames for understanding the combustion phenomena underlying the formation of the exhaust pollutants. Although LP combustion is an effective way to control NO{sub x}, the downside is that it increases the CO emissions. The formation and destruction of the pollutants (NO{sub x} and CO) are strongly affected by the fluid mechanics, the finite-rate chemistry, and their (turbulence-chemistry) interactions. Hence, a thorough understanding of these interactions is vital for controlling and reducing the pollutant emissions. The present research is contributing to this goal by providing a detailed nonintrusive laser based data set with good spatial and temporal resolutions of the pollutants (NO and CO) along with the major species, temperature, and OH. The measurements reported in this work, along with the existing velocity data on a turbulent LP combustor burning CH{sub 4}, would provide insight into the turbulence-chemistry interactions and their effect on pollutant formation.

  15. Rayleigh/Raman/LIF measurements in a turbulent lean premixed combustor

    SciTech Connect

    Nandula, S.P.; Pitz, R.W.; Barlow, R.S.

    1995-10-01

    Much of the industrial electrical generation capability being added worldwide is gas-turbine engine based and is fueled by natural gas. These gas-turbine engines use lean premixed (LP) combustion to meet the strict NO{sub x} emission standards, while maintaining acceptable levels of CO. In conventional, diffusion flame gas turbine combustors, large amount of NO{sub x} forms in the hot stoichiometric zones via the Zeldovich (thermal) mechanism. Hence, lean premixed combustors are rapidly becoming the norm, since they are specifically designed to avoid these hot stoichiometric zones and the associated thermal NO, However, considerable research and development are still required to reduce the NO{sub x} levels (25-40 ppmvd adjusted to 15% O{sub 2} with the current technology), to the projected goal of under 10 ppmvd by the turn of the century. Achieving this objective would require extensive experiments in LP natural gas (or CH{sub 4}) flames for understanding the combustion phenomena underlying the formation of the exhaust pollutants. Although LP combustion is an effective way to control NO{sub x}, the downside is that it increases the CO emissions. The formation and destruction of the pollutants (NO{sub x} and CO) are strongly affected by the fluid mechanics, the finite-rate chemistry, and their (turbulence-chemistry) interactions. Hence, a thorough understanding of these interactions is vital for controlling and reducing the pollutant emissions. The present research is contributing to this goal by providing a detailed nonintrusive laser based data set with good spatial and temporal resolutions of the pollutants (NO and CO) along with the major species, temperature, and OH. The measurements reported in this work, along with the existing velocity data on a turbulent LP combustor burning CH{sub 4}, would provide insight into the turbulence-chemistry interactions and their effect on pollutant formation.

  16. Measurement of alkali vapors in PFBC exhaust. Final report

    SciTech Connect

    Lee, S.H.D.; Swift, W.M.

    1994-01-01

    Under the auspices of the US Department of Energy through Morgantown Energy Technology Center, laboratory-scale studies were conducted to develop a regenerable activated-bauxite adsorbent (RABA) for use in an in situ regenerable activated-bauxite sorber alkali monitor (RABSAM). The RABSAM is a sampling probe that does not require a high-temperature/high-pressure sampling line for reliable measurement of alkali vapor in the exhaust of pressurized fluidized-bed combustor (PFBC). The RABA can be generated from the commercial grade activated bauxite by deactivating (or reacting) clay impurities in activated bauxite with NaCl or LiCl vapor. Under the atmospheric deactivation process, however, only a partial deactivation of clay impurities is achieved, probably due to limited access of NaCl or LiCl vapor into micropores of activated bauxite. Because LiCl vapor chemically reacts with alumina substrate of activated bauxite, resulting in pore enlargement, loss of pore surface area, and a decrease in the subsequent NaCl-vapor sorption capacity of the RABA, NaCl is a more suitable deactivation agent than LiCl vapor. In a simulated PFBC exhaust environment, the RABA behaves similarly to fresh activated bauxite in capturing NaCl vapor from the simulated PFBC exhaust. Based on results of this work, we recommend generating chemically and thermally stable RABA by deactivating clay impurities of commercial grade activated bauxite with NaCl or KCl vapor under simulated PFBC exhaust environment, that is, high-temperature, high-pressure, and high concentrations of NaCl or KCl vapor in simulated PFBC exhaust compositions.

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

    NASA Technical Reports Server (NTRS)

    DeLaat, John C.; Chang, Clarence T.

    2002-01-01

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

  18. Heating and Efficiency Comparison of a Fischer-Tropsch (FT) Fuel, JP-8+100, and Blends in a Three-Cup Combustor Sector

    NASA Technical Reports Server (NTRS)

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

    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 MIL-DTL-83133F(2008) or ASTM D 7566-Annex 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 aF (533 K), 125 psia (0.86 MPa) at 625 aF (603 K), 175 psia (1.21 MPa) at 725 aF (658 K), and 225 psia (1.55 MPa) at 790 aF (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 (% 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. In general, 100% SPK-FT fuel and blends with JP-8+100 produce less particulates and less smoke and have lower thermal impact on combustor hardware.

  19. Coal desulfurization in a rotary kiln combustor

    SciTech Connect

    Cobb, J.T. Jr.

    1992-09-11

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

  20. Partially integrated exhaust manifold

    SciTech Connect

    Hayman, Alan W; Baker, Rodney E

    2015-01-20

    A partially integrated manifold assembly is disclosed which improves performance, reduces cost and provides efficient packaging of engine components. The partially integrated manifold assembly includes a first leg extending from a first port and terminating at a mounting flange for an exhaust gas control valve. Multiple additional legs (depending on the total number of cylinders) are integrally formed with the cylinder head assembly and extend from the ports of the associated cylinder and terminate at an exit port flange. These additional legs are longer than the first leg such that the exit port flange is spaced apart from the mounting flange. This configuration provides increased packaging space adjacent the first leg for any valving that may be required to control the direction and destination of exhaust flow in recirculation to an EGR valve or downstream to a catalytic converter.

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

    SciTech Connect

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

    2016-11-29

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

  2. Combustor air flow control method for fuel cell apparatus

    DOEpatents

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

    2001-01-01

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

  3. Combined fluidized bed retort and combustor

    DOEpatents

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

    1984-01-01

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

  4. High pressure combustor for generating steam downhole

    SciTech Connect

    Retallick, W.B.

    1983-08-09

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

  5. Radial inlet guide vanes for a combustor

    DOEpatents

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

    2013-02-12

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

  6. Testing and Characterization of CMC Combustor Liners

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Verrilli, Michael J.

    2003-01-01

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

  7. Low NOx heavy fuel combustor concept program addendum: Low/mid heating value gaseous fuel evaluation

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

    The combustion performance of a rich/quench/lean (RQL) combustor was evaluated when operated on low and mid heating value gaseous fuels. Two synthesized fuels were prepared having lower heating values of 10.2 MJ/cu m. (274 Btu/scf) and 6.6 MJ/cu m (176 Btu/scf). These fuels were configured to be representative of actual fuels, being composed primarily of nitrogen, hydrogen, carbon monoxide, and carbon dioxide. A liquid fuel air assist fuel nozzle was modified to inject both of the gaseous fuels. The RQL combustor liner was not changed from the configuration used when the liquid fuels were tested. Both gaseous fuels were tested over a range of power levels from 50 percent load to maximum rated power of the DDN Model 570-K industrial gas turbine engine. Exhaust emissions were recorded for four power level at several rich zone equivalence ratios to determine NOx sensitivity to the rich zone operating point. For the mid Btu heating value gas, ammonia was added to the fuel to simulate a fuel bound nitrogen type gaseous fuel. Results at the testing showed that for the low heating value fuel NOx emissions were all below 20 ppmc and smoke was below a 10 smoke number. For the mid heating value fuel, NOx emissions were in the 50 to 70 ppmc range with the smoke below a 10 smoke number.

  8. A Modified Through-Flow Wave Rotor Cycle with Combustor Bypass Ducts

    NASA Technical Reports Server (NTRS)

    Paxson Daniel E.; Nalim, M. Razi

    1998-01-01

    A wave rotor cycle is described which avoids the inherent problem of combustor exhaust gas recirculation (EGR) found in four-port, through-flow wave rotor cycles currently under consideration for topping gas turbine engines. The recirculated hot gas is eliminated by the judicious placement of a bypass duct which transfers gas from one end of the rotor to the other. The resulting cycle, when analyzed numerically, yields an absolute mean rotor temperature 18% below the already impressive value of the conventional four-port cycle (approximately the turbine inlet temperature). The absolute temperature of the gas leading to the combustor is also reduced from the conventional four-port design by 22%. The overall design point pressure ratio of this new bypass cycle is approximately the same as the conventional four-port cycle. This paper will describe the EGR problem and the bypass cycle solution including relevant wave diagrams. Performance estimates of design and off-design operation of a specific wave rotor will be presented. The results were obtained using a one-dimensional numerical simulation and design code.

  9. Factors affecting cleanup of exhaust gases from a pressurized, fluidized-bed coal combustor

    NASA Technical Reports Server (NTRS)

    Rollbuhler, R. J.; Kobak, J. A.

    1980-01-01

    The cleanup of effluent gases from the fluidized-bed combustion of coal is examined. Testing conditions include the type and feed rate of the coal and the sulfur sorbent, the coal-sorbent ratio, the coal-combustion air ratio, the depth of the reactor fluidizing bed, and the technique used to physically remove fly ash from the reactor effluent gases. Tests reveal that the particulate loading matter in the effluent gases is a function not only of the reactor-bed surface gas velocity, but also of the type of coal being burnt and the time the bed is operating. At least 95 percent of the fly ash particules in the effluent gas are removed by using a gas-solids separator under controlled operating conditions. Gaseous pollutants in the effluent (nitrogen and sulfur oxides) are held within the proposed Federal limits by controlling the reactor operating conditions and the type and quantity of sorbent material.

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

    SciTech Connect

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

    1992-12-01

    Regarding PAH preparation, our efforts this quarter were directed at identifying the operating conditions in the laboratory furnace that clearly resolve primary devolatilization from secondary pyrolysis. In previous studies, pristine tars from primary devolatilization were prepared in the radiant coal flow experiment with a furnace length of only 5 cm. Tars at various extents of secondary pyrolysis were generated in a hot-zone of 12.5 cm. For this project we decided to try to prepare both kinds of PAH in the 12.5 cm hot zone. Preliminary screening surveys showed that, among the dime possible ways to compensate for the longer furnace tube, only lowering the fumace temperature is really effective. Lowering the loading from its already low value of 300 No./cm[sup 3] is inconsequential, and raising the gas velocity to lower residence times in the longer hot zones causes deposition problems. The coal jet undergoes the transition to turbulence in less than 12 cm for entrainment velocities greater than 1 cm/s. By the end of the quarter, it was clear that the best value for primary devolatilization studies will fall within the range of 1 150 to 1400 K. But at present we are still not completely satisfied with the resolution of primary devolatilization. As intended, the tars collected at short residence times are completely soluble in tetrahydrofuran (THF), indicating that no soot is present so no secondary pyrolysis has occurred. But at residence times long enough to attain ultimate primary devolatilization yields, some soot is present. We will make our final choice early in the next quarter, then begin to collect an inventory of primary devoatilization tars for subsequent detailed characterization.

  11. Spectral radiance measurements and calculated soot concentrations along the length of an experimental combustor

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    Radiometric data were obtained over a range of parametric test conditions at three positions along the length of an experimental combustor segment corresponding to the primary, intermediate, and dilution zones. The concentration of soot entrained in the combustion gases was calculated by a technique using spectral radiance measurements. Tests were conducted primarily with Jet A fuel, although limited data were taken with two fuels having higher aromatic content, diesel oil number 2 and a blend of 40 percent tetralin in Jet A fuel. Radiometric observation of the combustion gases indicated that the maximum total radiance peaked at the intermediate zone, which was located immediately upstream of the dilution holes. Soot concentrations calculated from optical measurements in the dilution zone compared favorably with those obtained by in situ gas sampling at the exhaust. The total radiance increased with the higher aromatic content fuels.

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

    DTIC Science & Technology

    2009-12-01

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

  13. Effects of porous insert on flame dynamics in a lean premixed swirl-stabilized combustor

    NASA Astrophysics Data System (ADS)

    Brown, Marcus; Agrawal, Ajay; Allen, James; Kornegay, John

    2016-11-01

    In this study, we investigated different methods of determining the effect a porous insert has on flame dynamics during lean premixed combustion. A metallic porous insert is used to mitigate instabilities in a swirl-stabilized combustor. Thermoacoustic instabilities are seen as negative consequences of lean premixed combustion and eliminating them is the motivation for our research. Three different diagnostics techniques with high-speed Photron SA5 cameras were used to monitor flame characteristics. Particle image velocimetry (PIV) was used to observe vortical structures and recirculation zones within the combustor. Using planar laser induced fluorescence (PLIF), we were able to observe changes in the reaction zones during instabilities. Finally, utilizing a color high-speed camera, visual images depicting a flame's oscillations during the instability were captured. Using these monitoring techniques, we are able to support the claims made in previous studies stating that the porous insert in the combustor significantly reduces the thermoacoustic instability. Funding for this research was provided by the NSF REU site Grant EEC 1358991 and NASA Grant NNX13AN14A.

  14. Nonintrusive laser-induced imaging for speciation and patternation in high-pressure gas turbine combustors

    NASA Astrophysics Data System (ADS)

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

    1999-10-01

    The next generation of ga turbine combustors for aerospace applications will be required to meet increasingly stringent constraints on fuel efficiency, noise abatement, and emissions. The power plants being designed to meet these constraints will operate at extreme conditions of temperature and pressure, thereby generating unique challenges to the previously employed diagnostic methodologies. Current efforts at NASA Glenn Research Center GRC utilize optically accessible, high-pressure flametubes and sector combustor rigs to probe, via advanced nonintrusive laser techniques, the complex flowfields encountered in advanced combustor designs. The fuel-air mixing process is of particular concern for lowering NOx emissions generated in lean, premixed engine concepts. Using planar laser-induced fluorescence we have obtained real- time, detailed imaging of the fuel spray distribution for a number of fuel injectors over a wide range of operational conditions that closely match those expected in the proposed propulsion systems. Using a novel combination of planar imaging of fuel fluorescence and computational analysis that allows an examination of the flowfield from any perspective, we have produced spatially and temporally resolved fuel-air distribution maps. These maps provide detailed insight into the fuel injection process at actual conditions never before possible, thereby greatly enhancing the evaluation of fuel injector performance and combustion phenomena.

  15. Testing of DLR C/C-SiC for HIFiRE 8 Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael

    2013-01-01

    Ceramic Matrix Composites (CMCs) have been proposed for hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a C/C and a C/C-SiC material system fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for HIFiRE 8, a joint Australia / AFRL hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kPa. Flat panels of the DLR C/C and the C/C-SiC were tested in the NASA Langley Direct Connect Rig (DCR) at Mach 5 and Mach 6 enthalpy for several minutes. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used in some of the tests to increase the surface temperature of the C/C-SiC panel for approximately 350degF. The final C/C-SiC panel was tested for 3 cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

  16. ULTRA-LOW NOX ADVANCED VORTEX COMBUSTOR

    SciTech Connect

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

    2006-05-01

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

  17. Ultra-Low NOx Advanced Vortex Combustor

    SciTech Connect

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

    2006-05-01

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

  18. Error Reduction Program. [combustor performance evaluation codes

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  19. Investigation of combustion instability in ramjet combustors

    SciTech Connect

    Reuter, D.M.

    1988-01-01

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

  20. Low NO.sub.x multistage combustor

    DOEpatents

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

    2000-01-01

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

  1. Combustor Computations for CO2-Neutral Aviation

    NASA Technical Reports Server (NTRS)

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

    2011-01-01

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

  2. Drying Milk With Boiler Exhaust

    NASA Technical Reports Server (NTRS)

    Broussard, M. R.

    1984-01-01

    Considerable energy saved in powdered-milk industry. Only special requirement boiler fired with natural gas or other clean fuel. Boiler flue gas fed to spray drier where it directly contacts product to be dried. Additional heat supplied by auxillary combustor when boiler output is low. Approach adaptable to existing plants with minimal investment because most already equipped with natural-gas-fired boilers.

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

    NASA Astrophysics Data System (ADS)

    Lo, Kai-Fan; Wong, Shwin-Chung

    2012-09-01

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

  4. Combustor technology for future small gas turbine aircraft

    NASA Technical Reports Server (NTRS)

    Lyons, Valerie J.; Niedzwiecki, Richard W.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  6. Achieving improved cycle efficiency via pressure gain combustors

    SciTech Connect

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

    1995-04-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

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

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

    DOEpatents

    Lieuwen, Tim C.; Nair, Suraj

    2006-08-15

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

  10. Induction time effects in pulse combustors

    SciTech Connect

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

    1999-04-09

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

  11. Advanced Low Emissions Subsonic Combustor Study

    NASA Technical Reports Server (NTRS)

    Smith, Reid

    1998-01-01

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

  12. Exhaust gas ignition

    SciTech Connect

    1996-04-01

    This article describes a system developed for rapid light-off of underbody catalysts that has shown potential to meet Euro Stage III emissions targets and to be more cost-effective than some alternatives. Future emissions legislation will require SI engine aftertreatment systems to approach full operating efficiency within the first few seconds after starting to reduce the high total-emissions fraction currently contributed by the cold phase of driving. A reduction of cold-start emissions during Phase 1 (Euro) or Bag 1 (FTP), which in many cases can be as much as 80% of the total for the cycle, has been achieved by electrical heating of the catalytic converter. But electrically heated catalyst (EHC) systems require high currents (100--200 A) to heat the metallic substrate to light-off temperatures over the first 15--20 seconds. Other viable approaches to reducing cold-start emissions include use of a fuel-powered burner upstream of the catalyst. However, as with EHC, the complexity of parts and the introduction of raw fuel into the exhaust system make this device unsatisfactory. Still another approach, an exhaust gas ignition (EGI) system, was first demonstrated in 1991. The operation of a system developed by engineers at Ford Motor Co., Ltd., Cambustion Ltd., and Tickford Ltd. is described here.

  13. Oxy-combustor operable with supercritical fluid

    DOEpatents

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

    2017-04-04

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

  14. Numerical Analysis of the SCHOLAR Supersonic Combustor

    NASA Technical Reports Server (NTRS)

    Rodriguez, Carlos G.; Cutler, Andrew D.

    2003-01-01

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

  15. Liquid rocket combustor computer code development

    NASA Technical Reports Server (NTRS)

    Liang, P. Y.

    1985-01-01

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

  16. Low NOx heavy fuel combustor concept program

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  17. CFD Code Development for Combustor Flows

    NASA Technical Reports Server (NTRS)

    Norris, Andrew

    2003-01-01

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

  18. Application of Chimera Grid Scheme to Combustor Flowfields at all Speeds

    NASA Technical Reports Server (NTRS)

    Yungster, Shaye; Chen, Kuo-Huey

    1997-01-01

    A CFD method for solving combustor flowfields at all speeds on complex configurations is presented. The approach is based on the ALLSPD-3D code which uses the compressible formulation of the flow equations including real gas effects, nonequilibrium chemistry and spray combustion. To facilitate the analysis of complex geometries, the chimera grid method is utilized. To the best of our knowledge, this is the first application of the chimera scheme to reacting flows. In order to evaluate the effectiveness of this numerical approach, several benchmark calculations of subsonic flows are presented. These include steady and unsteady flows, and bluff-body stabilized spray and premixed combustion flames.

  19. Mercury emissions from municipal solid waste combustors

    SciTech Connect

    Not Available

    1993-05-01

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

  20. A clean coal combustion technology-slagging combustors

    SciTech Connect

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

    1989-03-01

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

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

    NASA Technical Reports Server (NTRS)

    1982-01-01

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

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

    DOEpatents

    Chalfant, Robert W.; Clingerman, Bruce J.

    2002-01-01

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

  3. Combustor materials requirements and status of ceramic matrix composites

    NASA Technical Reports Server (NTRS)

    Hecht, Ralph J.; Johnson, Andrew M.

    1992-01-01

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

  4. Jet engine exhaust emissions of high altitude commercial aircraft projected to 1990

    NASA Technical Reports Server (NTRS)

    Grobman, J.; Ingebo, R. D.

    1974-01-01

    Projected minimum levels of engine exhaust emissions that may be practicably achievable for future commercial aircraft operating at high-altitude cruise conditions are presented. The forecasts are based on:(1) current knowledge of emission characteristics of combustors and augmentors; (2) the status of combustion research in emission reduction technology; and (3) predictable trends in combustion systems and operating conditions as required for projected engine designs that are candidates for advanced subsonic or supersonic commercial aircraft fueled by either JP fuel, liquefied natural gas, or hydrogen. Results are presented for cruise conditions in terms of both an emission index (g constituent/kg fuel) and an emission rate (g constituent/hr).

  5. Exhaust gas sensors

    SciTech Connect

    Hiller, J.; Miree, T.J.

    1997-02-09

    The automotive industry needed a fast, reliable, under-the-hood method of determining nitrogen oxides in automobile exhaust. Several technologies were pursued concurrently. These sensing technologies were based on light absorption, electrochemical methods, and surface mass loading. The Y-12 plant was selected to study the methods based on light absorption. The first phase was defining the detailed technical objectives of the sensors--this was the role of the automobile companies. The second phase was to develop prototype sensors in the laboratories--the national laboratories. The final phase was testing of the prototype sensors by the automobile industries. This program was canceled a few months into what was to be a three-year effort.

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

    DTIC Science & Technology

    1988-10-01

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

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

    DTIC Science & Technology

    2005-08-21

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

  8. Variable volume combustor with an air bypass system

    DOEpatents

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

    2017-02-07

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

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

    DTIC Science & Technology

    2011-03-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Ajmani, Kumud

    2015-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Ajmani, Kumud

    2015-01-01

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

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

    SciTech Connect

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

    1992-01-01

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

  14. Automotive Fuel and Exhaust Systems.

    ERIC Educational Resources Information Center

    Irby, James F.; And Others

    Materials are provided for a 14-hour course designed to introduce the automotive mechanic to the basic operations of automotive fuel and exhaust systems incorporated on military vehicles. The four study units cover characteristics of fuels, gasoline fuel system, diesel fuel systems, and exhaust system. Each study unit begins with a general…

  15. Turbocharged engine with exhaust purifier

    SciTech Connect

    Tadokoro, T.; Matsuda, I.; Okimoto, H.

    1986-09-23

    The patent described a control system for an automobile engine having intake and exhaust systems for respectively conducting intake gases to and exhaust gases from the engine, which comprises, in combination: a turbocharger including a turbine disposed in the exhaust system and adapted to be driven by the flow of the exhaust gases therethrough and a blower disposed in the intake system and drivingly connected with the turbine for supercharging the intake gases; and exhaust purifying device disposed in the exhaust system downstream of the turbine with respect to the direction of flow of the exhaust gases; a regulating means for varying the effective cross-section of a portion of the exhaust system leading to the turbine; a control means for controlling the regulating means in dependence on an operating condition of the engine, the control means causing the regulating means to decrease the effective cross-section during a low speed operating condition, but to increase the effective cross-section during a high speed operating condition of the engine.

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

    NASA Technical Reports Server (NTRS)

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

    1948-01-01

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

  17. Fuel Flexible Gas Turbine Combustor Flametube Facility Upgraded

    NASA Technical Reports Server (NTRS)

    Little, James E.; Nemets, Steve A.; Tornabene, Robert T.; Smith, Timothy D.; Frankenfeld, Bruce J.

    2004-01-01

    In fiscal year 2003, test cell 23 of the Research Combustion Laboratory (RCL 23) at the NASA Glenn Research Center was upgraded with the addition of gaseous hydrogen as a working propellant and the addition of a 450-psig air-supply system. Test flexibility was further enhanced by upgrades to the facility control systems. RCL 23 can now test with gaseous hydrogen flow rates up to 0.05 lbm/sec and jet fuel flow rates up to 0.62 lbm/sec. Research airflow rates up to 3 lbm/sec are possible with the 450-psig supply system over a range of inlet temperatures. Nonvitiated, heated air is supplied from a shell and tube heat exchanger. The maximum nonvitiated facility air temperature is 1100 F at 1.5 lbm/sec. Research-section exhaust temperatures are limited to 3200 F because of material and cooling capacity limits. A variety of support systems are available depending on the research hardware configuration. Test section ignition can be provided via either a hydrogen air torch system or an electronic spark system. Emissions measurements are obtained with either pneumatically or electromechanically actuated gas sample probes, and the electromechanical system allows for radial measurements at a user-specified axial location for measurement of emissions profiles. Gas analysis data can be obtained for a variety of species, including carbon monoxide (CO), carbon dioxide (CO2), nitrogen oxides (NO and NOx), oxygen (O2), unburnt hydrocarbons, and unburnt hydrogen. Facility control is accomplished with a programmable logic control system. Facility operations have been upgraded to a system based on graphical user interface control screens. A data system is available for real-time acquisition and monitoring of both measurements in engineering units and performance calculations. The upgrades have made RCL 23 a highly flexible facility for research into low emissions gas turbine combustor concepts, and the flame tube configuration inherently allows for a variety of fuel nozzle

  18. Further exhaust emission control for two-stroke engines

    SciTech Connect

    Sato, Kazuo; Nakano, Masamitsu; Ukawa, Haruo; Inaga, Hisashi

    1994-09-01

    Two-stroke engines are being utilized in large numbers as small utility, lawn and garden equipment engines. The following two subjects were examined with regards to exhaust emission control. The first subject was to compare the theoretical values of a combustion model simulation with the experimentally measured values of the base line emission of two-stroke volume. The second subject was to examine the emission conformability to the 1995 and 1999 California Air Resources Board (CARB) exhaust emission regulations California Regulations for 1995 and Later Utility and Lawn and Garden Equipment Engine, adopted at March 20, 1992, amended, at November 3, 1993. in two-stroke engines with various combinations between various fuels, fuel supply systems and scavenging systems. For this subject it was determine;that the emission control systems based on the lean combustion can be used to meet the 1995 CARB exhaust emission regulations. However, it was also concluded that to meet the 1999 CARB exhaust emission regulations, various emission control systems with various combinations regarding such parameters as fuels, scavenging systems and exhaust systems must be used. 27 refs., 20 figs., 4 tabs.

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

    DOEpatents

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

    2013-09-10

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

  20. A study of automobile exhaust noise preferences

    NASA Astrophysics Data System (ADS)

    Haire, Jay B.; Carney, Melinda J.; Cheenne, Dominique J.

    2005-04-01

    A study was conducted to investigate the relationship between preferences in automobile exhaust noise and the demographic factors of a listening jury. Noise samples of four different vehicles were recorded at idle as well as at 3000 RPM, and 1/3 octave sound spectra were acquired simultaneously. The recordings were presented to the jury using headphones and a preference survey was administered. Zwicker loudness was computed for all samples. Demographic factors such as gender, age, current and future vehicle ownership, were correlated to listening preferences, and unforeseen results were found, especially in regards to sport utility vehicles (SUV).

  1. Diesel exhaust-gas purification system

    SciTech Connect

    Doherty, B.J.

    1982-07-01

    The design of a diesel exhaust gas purification system is presented. It will provide 2000 scfm of dry, anerobic gas (essentially nitrogen) for use in air drilling operations where drill pipe corrosion is a problem, such as geothermal applications. The system is operable in the field and may be transported via highways. It will operate at ambient temperatures up to 110/sup 0/F and requires no water - diesel fuel is used to combust excess oxygen and to generate electricity for the system. Gas production costs, including capital amortization, operations, fuel and maintenance (for reasonable utilization) are about $1.50/1000 scf.

  2. Forecast of jet engine exhaust emissions for future high altitude commercial aircraft

    NASA Technical Reports Server (NTRS)

    Grobman, J.; Ingebo, R. D.

    1974-01-01

    Projected minimum levels of engine exhaust emissions that may be practicably achievable for future commercial aircraft operating at high altitude cruise conditions are presented. The forecasts are based on: (1) current knowledge of emission characteristics of combustors and augmentors; (2) the current status of combustion research in emission reduction technology; (3) predictable trends in combustion systems and operating conditions as required for projected engine designs that are candidates for advanced subsonic or supersonic commercial aircraft. Results are presented for cruise conditions in terms of an emission index, g pollutant/kg fuel. Two sets of engine exhaust emission predictions are presented: the first, based on an independent NASA study and the second, based on the consensus of an ad hoc committee composed of industry, university, and government representatives. The consensus forecasts are in general agreement with the NASA forecasts.

  3. Core Noise: Overview of Upcoming LDI Combustor Test

    NASA Technical Reports Server (NTRS)

    Hultgren, Lennart S.

    2012-01-01

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

  4. Melt Infiltrated Ceramic Matrix Composites for Shrouds and Combustor Liners of Advanced Industrial Gas Turbines

    SciTech Connect

    Gregory Corman; Krishan Luthra; Jill Jonkowski; Joseph Mavec; Paul Bakke; Debbie Haught; Merrill Smith

    2011-01-07

    This report covers work performed under the Advanced Materials for Advanced Industrial Gas Turbines (AMAIGT) program by GE Global Research and its collaborators from 2000 through 2010. A first stage shroud for a 7FA-class gas turbine engine utilizing HiPerComp{reg_sign}* ceramic matrix composite (CMC) material was developed. The design, fabrication, rig testing and engine testing of this shroud system are described. Through two field engine tests, the latter of which is still in progress at a Jacksonville Electric Authority generating station, the robustness of the CMC material and the shroud system in general were demonstrated, with shrouds having accumulated nearly 7,000 hours of field engine testing at the conclusion of the program. During the latter test the engine performance benefits from utilizing CMC shrouds were verified. Similar development of a CMC combustor liner design for a 7FA-class engine is also described. The feasibility of using the HiPerComp{reg_sign} CMC material for combustor liner applications was demonstrated in a Solar Turbines Ceramic Stationary Gas Turbine (CSGT) engine test where the liner performed without incident for 12,822 hours. The deposition processes for applying environmental barrier coatings to the CMC components were also developed, and the performance of the coatings in the rig and engine tests is described.

  5. Determination of an acoustic reflection coefficient at the inlet of a model gas turbine combustor for power generation

    NASA Astrophysics Data System (ADS)

    Song, W. J.; Cha, D. J.

    2017-01-01

    A phenomenon that potentially influences the reliability of power generation systems is the presence of thermo-acoustic oscillations in the combustion chamber of a land- based gas turbine. To develop specific measures that prevent the instability, it is essential to predict and/or evaluate the underlying physics of the thermo-acoustics, which requires the acoustic boundary condition at the exit of the burner, that is, at the inlet of the combustor. Here we report a procedure for calculating acoustic reflection coefficients at the burner exit by utilizing two microphone method (TMM) for dynamic pressure signals. The procedure has been verified by comparing its results with reported ones and further successfully employed to determine the acoustic boundary condition of the burner of a partially-premixed model gas turbine combustor.

  6. External combustor for gas turbine engine

    DOEpatents

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

    1991-01-01

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

  7. Analysis of Regen Cooling in Rocket Combustors

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

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

  8. Alternate-Fueled Combustor-Sector Emissions

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  9. Fluidized bed combustor and tube construction therefor

    DOEpatents

    De Feo, Angelo; Hosek, William

    1981-01-01

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

  10. Tube construction for fluidized bed combustor

    DOEpatents

    De Feo, Angelo; Hosek, William

    1984-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Giuliani, James E.; Klem, Mark D.

    1992-01-01

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

  12. 40 CFR Table 3 to Subpart Cb of... - Municipal Waste Combustor Operating Guidelines

    Code of Federal Regulations, 2011 CFR

    2011-07-01

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

  13. Computational Simulation of Acoustic Modes in Rocket Combustors

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

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

  14. Ceramic composite liner material for gas turbine combustors

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

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

    DTIC Science & Technology

    2009-06-01

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

  16. Characteristics of a Steadily Operating Metal Combustor

    DTIC Science & Technology

    1976-08-01

    exhaust duct were insulated with approximately a 5 cm layer of Fiberfrax Lo-Con ceramic fiber blanket manufactured by the Carborundum Company. 4.2.2 Fuel...fuel tank were insulated with approximately a 5 cm layer of Fiberfrax Lo-Con blanket. 4.2.3 Product Collection Tank. The product collection tank was...parasitic heat loss from this region. 57 The exterior surfaces of the reaction chamber were insulated with approximately a 5 cm layer of Fiberfrax Lo-Con

  17. Coupled Lagrangian Monte Carlo PDF-CFD computation of gas turbine combustor flowfields with finite-rate chemistry

    SciTech Connect

    Tolpadi, A.K.; Hu, I.Z.; Correa, S.M.; Burrus, D.L.

    1997-07-01

    A coupled Lagrangian Monte Carlo Probability Density Function (PDF)-Eulerian Computational Fluid Dynamics (CFD) technique is presented for calculating steady three-dimensional turbulent reacting flow in a gas turbine combustor. PDF transport methods model turbulence-combustion interactions more accurately than conventional turbulence models with an assumed shape PDF. The PDF transport equation was solved using a Lagrangian particle tracking Monte Carlo (MC) method. The PDF modeled was over composition only. This MC module has been coupled with CONCERT, which is a fully elliptic three-dimensional body-fitted CFD code based on pressure correction techniques. In an earlier paper, this computational approach was described, but only fast chemistry calculations were presented in a typical aircraft engine combustor. In the present paper, reduced chemistry schemes were incorporated into the MC module that enabled the modeling of finite rate effects in gas turbine flames and therefore the prediction of CO and NO{sub x} emissions. With the inclusion of these finite rate effects, the gas temperatures obtained were also more realistic. Initially, a two scalar scheme was implemented that allowed validation against Raman data taken in a recirculation bluff body stabilized CO/H{sub 2}/N{sub 2}-air flame. Good agreement of the temperature and major species were obtained. Next, finite rate computations were performed in a single annular aircraft engine combustor by incorporating a simple three scalar reduced chemistry scheme for Jet A fuel. This three scalar scheme was an extension of the two scalar scheme for CO/H{sub 2}/N{sub 2} fuel. The solutions obtained using the present approach were compared with those obtained using the fast chemistry PDF transport approach as well as the presumed shape PDF method. The calculated exhaust gas temperature using the finite rate model showed the best agreement with measurements made by a thermocouple rake.

  18. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

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

  19. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

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

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

    Code of Federal Regulations, 2014 CFR

    2014-07-01

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

  1. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

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

  2. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

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

  3. 40 CFR 60.53b - Standards for municipal waste combustor operating practices.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

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

  4. 40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

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

  5. 40 CFR 60.55b - Standards for municipal waste combustor fugitive ash emissions.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

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

  6. Application of Combined Laser-Induced Ignition and Plasma Spectroscopy on a Partially Premixed Combustor

    NASA Astrophysics Data System (ADS)

    Okai, Keiichi; Zimmer, Laurent; Kurosawa, Yoji

    Combined Laser Induced Ignition and Plasma Spectroscopy has a potential to give in situ measurement of fuel equivalence ratio in the laser ignition event at an exact time and location of plasma initiation. This method is an extension of a non-intuitive gaseous-composite measurement technique, i.e. Laser Induced Plasma Spectroscopy, to simultaneous ignition utilizing the same plasma source both for gaseous composite measurements and ignition. Calibration methods so far are not sufficient in terms of accuracy. The present study utilizes a new calibration technology without energy measurement under the ignition case. The example dealt in this study is a single injector combustor which provides a time-varying non-uniform fuel equivalence ratio under fuel lean cases. The result showed that the present calibration method is valid for the present measurement of the fuel equivalence ratio, and that the data is very useful to understand the ignition process of the configuration given.

  7. Numerical Simulation of Dual-Mode Scramjet Combustors

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

  8. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2007-01-01

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

  9. National Combustion Code: A Multidisciplinary Combustor Design System

    NASA Technical Reports Server (NTRS)

    Stubbs, Robert M.; Liu, Nan-Suey

    1997-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  11. Combustor with two stage primary fuel assembly

    DOEpatents

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

    2000-01-01

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

  12. Development of pressurized coal partial combustor

    SciTech Connect

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

    1995-12-31

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

  13. Parametric performance analysis of steam-injected gas turbine with a thermionic-energy-converter-lined combustor

    NASA Technical Reports Server (NTRS)

    Choo, Y. K.; Burns, R. K.

    1982-01-01

    The performance of steam-injected gas turbines having combustors lined with thermionic energy converters (STIG/TEC systems) was analyzed and compared with that of two baseline systems; a steam-injected gas turbine (without a TEC-lined combustor) and a conventional combined gas turbine/steam turbine cycle. Common gas turbine parameters were assumed for all of the systems. Two configurations of the STIG/TEC system were investigated. In both cases, steam produced in an exhaust-heat-recovery boiler cools the TEC collectors. It is then injected into the gas combustion stream and expanded through the gas turbine. The STIG/TEC system combines the advantage of gas turbine steam injection with the conversion of high-temperature combustion heat by TEC's. The addition of TEC's to the baseline steam-injected gas turbine improves both its efficiency and specific power. Depending on system configuration and design parameters, the STIG/TEC system can also achieve higher efficiency and specific power than the baseline combined cycle.

  14. Parametric performance analysis of steam-injected gas turbine with a thermionic-energy-converter-lined combustor

    NASA Astrophysics Data System (ADS)

    Choo, Y. K.; Burns, R. K.

    1982-02-01

    The performance of steam-injected gas turbines having combustors lined with thermionic energy converters (STIG/TEC systems) was analyzed and compared with that of two baseline systems; a steam-injected gas turbine (without a TEC-lined combustor) and a conventional combined gas turbine/steam turbine cycle. Common gas turbine parameters were assumed for all of the systems. Two configurations of the STIG/TEC system were investigated. In both cases, steam produced in an exhaust-heat-recovery boiler cools the TEC collectors. It is then injected into the gas combustion stream and expanded through the gas turbine. The STIG/TEC system combines the advantage of gas turbine steam injection with the conversion of high-temperature combustion heat by TEC's. The addition of TEC's to the baseline steam-injected gas turbine improves both its efficiency and specific power. Depending on system configuration and design parameters, the STIG/TEC system can also achieve higher efficiency and specific power than the baseline combined cycle.

  15. Effect of fuel zoning and fuel nozzle design on pollution emissions at ground idle conditions for a double-annular ram-induction combustor

    NASA Technical Reports Server (NTRS)

    Clements, T. R.

    1973-01-01

    An exhaust emission survey was conducted on a double-annular ram induction combustor at simulated ground idle conditions. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. The emission levels of total hydrocarbon (THC), carbon monoxide, carbon dioxide, and nitric oxide were measured. The effects of fuel zoning, fuel nozzle design, and operating conditions (inlet temperature and reference Mach number) on the level of these emissions were determined. At an overall combustor fuel/air ratio of 0.007, fuel zoning reduced THC emissions by a factor of 5 to 1. The reduction in THC emissions is attributed to the increase in local fuel/air ratio provided by the fuel zoning. An alternative method of increasing fuel/air ratio would be to operate with larger-than-normal compressor overboard bleed; however, analysis on this method indicated an increase in idle fuel consumption of 20 percent. The use of air-atomizing nozzles reduced the THC emissions by 2 to 1.

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

    DOEpatents

    Clingerman, Bruce J.; Mowery, Kenneth D.

    2002-01-01

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

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

    DOEpatents

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

    2016-09-20

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

  18. Variable volume combustor with center hub fuel staging

    DOEpatents

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

    2016-08-23

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

  19. Numerical Simulations of Static Tested Ramjet Dump Combustor

    NASA Astrophysics Data System (ADS)

    Javed, Afroz; Chakraborty, Debasis

    2016-06-01

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

  20. Nondestructive evaluation of ceramic matrix composite combustor components.

    SciTech Connect

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

    2002-11-08

    Combustor liners fabricated from a SiC/SiC composite were nondestructively interrogated before and after combustion rig testing. The combustor liners were inspected by X-ray, ultrasonic and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured liners, and from coupons machined from the components after combustion exposure. Thermography indications were found to correlate with reduced material properties obtained after rig testing. Microstructural examination of the SiC/SiC liners revealed the thermography indications to be delaminations and damaged fiber tows.

  1. Stagnation point reverse flow combustor for a combustion system

    NASA Technical Reports Server (NTRS)

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

    2007-01-01

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

  2. Effect of water injection on nitric oxide emissions of a gas turbine combustor burning natural gas fuel

    NASA Technical Reports Server (NTRS)

    Marchionna, N. R.; Diehl, L. A.; Trout, A. M.

    1973-01-01

    The effect of direct water injection on the exhaust gas emissions of a turbojet combustor burning natural gas fuel was investigated. The results are compared with the results from similar tests using ASTM Jet-A fuel. Increasing water injection decreased the emissions of oxides of nitrogen (NOX) and increased the emissions of carbon monoxide and unburned hydrocarbons. The greatest percentage decrease in NOX with increasing water injection was at the lowest inlet-air temperature tested. The effect of increasing inlet-air temperature was to decrease the effect of the water injection. The reduction in NOX due to water injection was almost identical to the results obtained with Jet-A fuel. However, the emission indices of unburned hydrocarbons, carbon monoxide, and percentage nitric oxide in NOX were not.

  3. Design and fabrication of a meso-scale stirling engine and combustor.

    SciTech Connect

    Echekki, Tarek (Sandia National Laboratories, Livermore, CA); Haroldsen, Brent L. (Sandia National Laboratories, Livermore, CA); Krafcik, Karen L. (Sandia National Laboratories, Livermore, CA); Morales, Alfredo Martin; Mills, Bernice E.; Liu, Shiling; Lee, Jeremiah C. (Sandia National Laboratories, Livermore, CA); Karpetis, Adionos N. (Sandia National Laboratories, Livermore, CA); Chen, Jacqueline H. (Sandia National Laboratories, Livermore, CA); Ceremuga, Joseph T. (Sandia National Laboratories, Livermore, CA); Raber, Thomas N.; Hekmuuaty, Michelle A.

    2005-05-01

    Power sources capable of supplying tens of watts are needed for a wide variety of applications including portable electronics, sensors, micro aerial vehicles, and mini-robotics systems. The utility of these devices is often limited by the energy and power density capabilities of batteries. A small combustion engine using liquid hydrocarbon fuel could potentially increase both power and energy density by an order of magnitude or more. This report describes initial development work on a meso-scale external combustion engine based on the Stirling cycle. Although other engine designs perform better at macro-scales, we believe the Stirling engine cycle is better suited to small-scale applications. The ideal Stirling cycle requires efficient heat transfer. Consequently, unlike other thermodynamic cycles, the high heat transfer rates that are inherent with miniature devices are an advantage for the Stirling cycle. Furthermore, since the Stirling engine uses external combustion, the combustor and engine can be scaled and optimized semi-independently. Continuous combustion minimizes issues with flame initiation and propagation. It also allows consideration of a variety of techniques to promote combustion that would be difficult in a miniature internal combustion engine. The project included design and fabrication of both the engine and the combustor. Two engine designs were developed. The first used a cylindrical piston design fabricated with conventional machining processes. The second design, based on the Wankel rotor geometry, was fabricated by through-mold electroforming of nickel in SU8 and LIGA micromolds. These technologies provided the requisite precision and tight tolerances needed for efficient micro-engine operation. Electroformed nickel is ideal for micro-engine applications because of its high strength and ductility. A rotary geometry was chosen because its planar geometry was more compatible with the fabrication process. SU8 lithography provided rapid

  4. Odor intensity and characterization of jet exhaust and chemical analytical measurements

    NASA Technical Reports Server (NTRS)

    Kendall, D. A.; Levins, P. L.

    1973-01-01

    Odor and chemical analyses were carried out on the exhaust samples from a J-57 combustor can operated over a range of inlet conditions, and with several fuel types and nozzle modifications. The odor characteristics and total intensity of odor for each exhaust were determined over a range of dilutions to allow for a least squares determination of the intensity at 1,000 to 1 dilutions. Analytical measures included the concentration of total hydrocarbons and the concentrations of aromatic organic species and oxygenated organic species from collected samples which were taken concurrently. A correlation was found between the concentration of the odorous oxygenated fraction and the total intensity of aroma. Inlet operating conditions and nozzle modifications which increase the efficiency of combustion as measured by exhaust gas analyses reduce the odor intensity and the quantity of oxygenates in the exhaust. The type of fuel burned altered the intensity of odor in relation to the quantity of oxygenates produced and, in some instances, changed the odor character.

  5. Development of a vortex combustor (VC) for space/water heating applications (combustion tests). Final report

    SciTech Connect

    Fu, T.T.; Nieh, S.

    1990-11-01

    This is the final report for Interagency Agreement DE-AI22-87PC79660 on ``Combustion Test`` for vortex combustor (VC) development for commercial applications. The work culminated in the successful demonstration of a 2 MB/H proof-of-concept (POC) model firing coal-water fuel (CWF). This development is concerned with a new concept in combustion, and was a general lack of relevant information. The work therefore began (in addition to the companion cold flow modeling study) with the design and test of two subscale models (0.15 and 0.3 MB/H) and one full scale model (3 MB/H) to obtain the needed information. With the experience gained, the 2 MB/H POC model was then designed and demonstrated. Although, these models were designed somewhat differently from one another, they all performed well and demonstrated the superiority of the concept. In summary, test results have shown that VC can be fired on several coal fuels (CWF, dry ultrafine coal, utility grind pulverized coal) at high combustion efficiency (>99%), high firing intensity (up to 0.44 MB/H-ft{sup 3}), and at temperatures sufficiently low or dry ash removal. The combustion process is completed totally inside the combustor. Conventional combustion enhancement techniques such as: preheating (air and/or fuel), pre-combustion, and post combustion are not needed.

  6. Development of a vortex combustor (VC) for space/water heating applications (combustion tests)

    SciTech Connect

    Fu, T.T. ); Nieh, S. . Combustion and Multiphase Flows Lab.)

    1990-11-01

    This is the final report for Interagency Agreement DE-AI22-87PC79660 on Combustion Test'' for vortex combustor (VC) development for commercial applications. The work culminated in the successful demonstration of a 2 MB/H proof-of-concept (POC) model firing coal-water fuel (CWF). This development is concerned with a new concept in combustion, and was a general lack of relevant information. The work therefore began (in addition to the companion cold flow modeling study) with the design and test of two subscale models (0.15 and 0.3 MB/H) and one full scale model (3 MB/H) to obtain the needed information. With the experience gained, the 2 MB/H POC model was then designed and demonstrated. Although, these models were designed somewhat differently from one another, they all performed well and demonstrated the superiority of the concept. In summary, test results have shown that VC can be fired on several coal fuels (CWF, dry ultrafine coal, utility grind pulverized coal) at high combustion efficiency (>99%), high firing intensity (up to 0.44 MB/H-ft[sup 3]), and at temperatures sufficiently low or dry ash removal. The combustion process is completed totally inside the combustor. Conventional combustion enhancement techniques such as: preheating (air and/or fuel), pre-combustion, and post combustion are not needed.

  7. Industrial-scale demonstration of a new sorbent reactivation technology for fluidized bed combustors.

    PubMed

    Anthony, Edward J; McCleave, Robert; Gandolfi, Eduardo; Wang, Jinsheng

    2003-10-01

    To minimize the disposal of highly reactive spent sorbent from a fluidized bed combustor, a new method for reactivation has been developed. The method consists of grinding the spent ash in a rotary mill, hydrating the ash with an excess of water, and mixing the wet ground ash with dry solids to absorb the excess water. The mixing process eliminates the formation of a concrete-like product that normally results as wet fluidized bed combustor ash ages. Pilot-scale combustion trials proved to be successful, and the process was scaled up using a 35MWt utility boiler at Purdue University. The test lasted for 3 days and resulted in net reduction of limestone sorbent use of 18%. The results generated in this work have been used to develop an economic evaluation for a 165MWe circulating fluidized bed (CFB) boiler, which projects significant savings due to reduction of limestone supply and ash disposal costs. The evaluation also suggests that the process is cost competitive with other processes, albeit that those processes have not been demonstrated at industrial scale. Furthermore, it also has the potential to make a small net reduction in CO(2) emissions, due to reduced limestone usage.

  8. Combustion of various types of residues in a circulating fluidized bed combustor

    SciTech Connect

    Gulyurtlu, I.; Frade, E.; Lopes, H.; Figueiredo, F.; Cabrita, I.

    1997-12-31

    Combustion studies of different wastes alone or mixed were undertaken in an atmospheric circulating fluidized bed. The combustor was operated over a temperature range of 700 to 1,000 C. The residues studied included biomass, tyres, and oil sludges. The main parameters that were investigated are (1) where and how to feed residues, (2) the ratios of amounts of residues when they are burned mixed, (3) air staging, and (4) excess air levels along the riser. The main conclusions are: (1) a large variation in the combustion efficiency was observed depending on the fuel, its particle size and where the fuel was fed in the combustor, (2) in the case of the use of mixture of wastes the utilization of biomass residues as the base fuel increased their combustion efficiency and resulted in very stable combustion conditions, (3) temperature appeared to be still very influential in reducing the levels of unburned carbon and hydrocarbons released from residues, and (4) the air staging in the freeboard improved combustion efficiency by enhancing the combustion of volatiles released from residues in the riser and produced lower emissions of both NO{sub x}. Particles collected in the cyclone and those removed from the bed were also analyzed to determine the levels of heavy metals. When oil sludge was added, high amounts of very fine particles of heavy metals were observed in the combustion gases.

  9. Non-Intrusive Laser-Induced Imaging for Speciation and Patternation in High Pressure Gas Turbine Combustors

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

    The next generation of was turbine combustors for aerospace applications will be required to meet increasingly stringent constraints on fuel efficiency, noise abatement, and emissions. The power plants being designed to meet these constraints will operate at extreme conditions of temperature and pressure, thereby generating unique challenges to the previously employed diagnostic methodologies. Current efforts at NASA Glenn Research Center (GRC) utilize optically accessible, high pressure flametubes and sector combustor rigs to probe, via advanced nonintrusive laser techniques, the complex flowfields encountered in advanced combustor designs. The fuel-air mixing process is of particular concern for lowering NO(x) emissions generated in lean, premixed engine concepts. Using planar laser-induced fluorescence (PLIF) we have obtained real-time, detailed imaging of the fuel spray distribution for a number of fuel injector over a wide range of operational conditions that closely match those expected in the proposed propulsion systems. Using a novel combination of planar imaging, of fuel fluorescence and computational analysis that allows an examination of the flowfield from any perspective, we have produced spatially and temporally resolved fuel-air distribution maps. These maps provide detailed insight into the fuel injection at actual conditions never before possible, thereby greatly enhancing the evaluation of fuel injector performance and combustion phenomena.

  10. Thermal Barrier and Protective Coatings to Improve the Durability of a Combustor Under a Pulse Detonation Engine Environment

    NASA Technical Reports Server (NTRS)

    Ghosn, Louis J.; Zhu, Dongming

    2008-01-01

    Pulse detonation engine (PDE) concepts are receiving increasing attention for future aeronautic propulsion applications, due to their potential thermodynamic cycle efficiency and higher thrust to density ratio that lead to the decrease in fuel consumption. But the resulting high gas temperature and pressure fluctuation distributions at high frequency generated with every detonation are viewed to be detrimental to the combustor liner material. Experimental studies on a typical metal combustion material exposed to a laser simulated pulse heating showed extensive surface cracking. Coating of the combustor materials with low thermal conductivity ceramics is shown to protect the metal substrate, reduce the thermal stresses, and hence increase the durability of the PDE combustor liner material. Furthermore, the temperature fluctuation and depth of penetration is observed to decrease with increasing the detonation frequency. A crack propagation rate in the coating is deduced by monitoring the variation of the coating apparent thermal conductivity with time that can be utilized as a health monitoring technique for the coating system under a rapid fluctuating heat flux.

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

    DOEpatents

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

    2013-12-24

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

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

    NASA Technical Reports Server (NTRS)

    Srivatsa, S. K.

    1982-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  14. Ventilation Exhaust Power Recovery Design

    NASA Astrophysics Data System (ADS)

    Yandell, Jeremy

    2012-11-01

    Due to the expense of designing ductwork and exhaust fans to meet the exact desired flow rate for building exhaust, there is wasted energy that is unrecovered when exhausted to the atmosphere. By designing a small diameter wind turbine the kinetic energy in the exhaust stream can be recovered and power provided back into the building. Unlike large scale commercial wind turbines that must be designed to provide power from a large range of wind speeds and directions, this smaller scale turbine can be optimized for a single constant wind speed with no variation in direction. The critical component is to prevent backpressure feeding through the system and increasing the load on the exhaust fan. This design project began with the theoretical airfoil and blade design, followed by modeling the system in fluid dynamics software, a full CAD design was created and modified for the selected manufacturing process, prototype creation and testing will be completed both in a wind tunnel and in a real environment, and the completed data will be compared with theoretical and computational results. Note: There is a patent pending for this design and concept.

  15. Testing of DLR C/C-SiC and C/C for HIFiRE 8 Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael K.

    2014-01-01

    Ceramic Matrix Composites (CMCs) have been proposed for use as lightweight hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a carbon/carbon (C/C) and a carbon/carbon-silicon carbide (C/C-SiC) material fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for Hypersonic International Flight Research Experimentation (HIFiRE) 8, a joint Australia / Air Force Research Laboratory hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kilopascals. Flat panels of the DLR C/C and C/C-SiC materials were installed downstream of a hydrogen-fueled, dual-mode scramjet combustor and tested for several minutes at conditions simulating flight at Mach 5 and Mach 6. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used during some of the tests to reduce heat loss from the back surface and thus increase the heated surface temperature of the C/C-SiC panel approximately 177 C (350 F). The final C/C-SiC panel was tested for three cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

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

    DOEpatents

    Hosek, William S.; Garruto, Edward J.

    1984-01-01

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

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

    EPA Science Inventory

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

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Gregory, Peyton B.; Holland, Anne D.

    1992-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

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

    DTIC Science & Technology

    2007-07-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

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

    EPA Science Inventory

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

  5. Improved Controllers For Heaters In Toxic-Gas Combustors

    NASA Technical Reports Server (NTRS)

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

    1995-01-01

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

  6. Catalytic combustor for integrated gasification combined cycle power plant

    DOEpatents

    Bachovchin, Dennis M.; Lippert, Thomas E.

    2008-12-16

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

  7. Coanda injection system for axially staged low emission combustors

    DOEpatents

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

    2012-05-15

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

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

    SciTech Connect

    Not Available

    1980-09-01

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

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

    EPA Science Inventory

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

  10. COMBUSTION CONTROL OF ORGANIC EMISSIONS FROM MUNICIPAL WASTE COMBUSTORS

    EPA Science Inventory

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

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

    EPA Science Inventory

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

  12. Assessment, development and application of combustor aerothermal models

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Dodds, W. J.

    1979-01-01

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

  14. Hydrogen Fuel Capability Added to Combustor Flametube Rig

    NASA Technical Reports Server (NTRS)

    Frankenfield, Bruce J.

    2003-01-01

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

  15. Low Emissions RQL Flametube Combustor Component Test Results

    NASA Technical Reports Server (NTRS)

    Holdeman, James D.; Chang, Clarence T.

    2001-01-01

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

  16. Mixed exhaust flow supersonic jet engine and method

    SciTech Connect

    Klees, G.W.

    1993-06-08

    A method of operating a supersonic jet engine installation is described comprising (a) providing an engine having a variable area air inlet means and an outlet to discharge engine exhaust; (b) providing a secondary air passageway means; (c) receiving ambient air in the air inlet means and providing the ambient air as primary air to the engine inlet and secondary air to the secondary air passageway means; (d) providing a mixing section having an inlet portion and an exit portion, utilizing the mixing section in directing the exhaust from the engine to primary convergent/divergent exit passageway segments, where the exhaust is discharged at supersonic velocity as primary flow components, and directing secondary air flow from the secondary air passageway means to secondary exit passageway segments which are interspersed with the primary segments and from which the secondary air is discharged at subsonic velocity as secondary flow components; and (e) providing an exhaust section to receive the primary and secondary flow components in a mixing region and causing the primary and secondary flow components to mix to create a supersonic mixed flow, the exhaust section having a variable area final nozzle through which the mixed flow is discharged.

  17. LES Investigation of Core Noise Mechanisms inside a Combustor-Nozzle System

    NASA Astrophysics Data System (ADS)

    O'Brien, Jeffrey; Bake, Friedrich; Kim, Jeonglae; Ihme, Matthias

    2016-11-01

    The aim of the work is to expand knowledge of core noise physics through the study of a representative aviation-type combustor with converging-diverging nozzle attached to the exhaust. First, a fully compressible LES of the entire flowpath is performed and validated against experimental measurements. From this calculation, the time history of the flow is sampled in a plane near the nozzle entrance to construct a library of representative fluctuations that are potential precursors to the direct & indirect noise observed at the nozzle outlet. This data is then used as an inflow for a series of separate nozzle simulations in which fluctuations in pressure, temperature ("hot spots"), and mixture composition are imposed separately to isolate their effect on the emitted noise. This methodology allows quantitative investigation of core-noise physics that lower-order models do not, including: the effect of non-linearity of high-amplitude perturbations, superposition of forcing types, the impact of the spatial structure of the perturbations, and the restriction to low-frequency perturbations and calorically perfect gas assumption. The calculations also represent the first time variations in mixture composition have been shown to induce downstream noise in a high-fidelity, 3D simulation.

  18. Combustion generated noise in gas turbine combustors. [engine noise/noise reduction

    NASA Technical Reports Server (NTRS)

    Strahle, W. C.; Shivashankara, B. N.

    1974-01-01

    Experiments were conducted to determine the noise power and spectra emitted from a gas turbine combustor can exhausting to the atmosphere. Limited hot wire measurements were made of the cold flow turbulence level and spectra within the can. The fuels used were JP-4, acetone and methyl alcohol burning with air at atmospheric pressure. The experimental results show that for a fixed fuel the noise output is dominated by the airflow rate and not the fuel/air ratio. The spectra are dominated by the spectra of the cold flow turbulence spectra which were invariant with airflow rate in the experiments. The effect of fuel type on the noise power output was primarily through the heat of combustion and not the reactivity. A theory of combustion noise based upon the flame radiating to open surroundings is able to reasonably explain the observed results. A thermoacoustic efficiency for noise radiation as high as .00003 was observed in this program for JP-4 fuel. Scaling rules are presented for installed configurations.

  19. Automotive Fuel and Exhaust Systems.

    ERIC Educational Resources Information Center

    Marine Corps Inst., Washington, DC.

    This correspondence course, originally developed for the Marine Corps, is designed to provide mechanics with an understanding of the construction, operation, malfunction, diagnosis, maintenance, and repair of the fuel and exhaust systems used in automobiles. The course contains five study units covering fundamentals of gasoline engine fuel…

  20. Reduction of gaseous pollutant emissions from gas turbine combustors using hydrogen-enriched jet fuel

    NASA Technical Reports Server (NTRS)

    Clayton, R. M.

    1976-01-01

    Recent progress in an evaluation of the applicability of the hydrogen enrichment concept to achieve ultralow gaseous pollutant emission from gas turbine combustion systems is described. The target emission indexes for the program are 1.0 for oxides of nitrogen and carbon monoxide, and 0.5 for unburned hydrocarbons. The basic concept utilizes premixed molecular hydrogen, conventional jet fuel, and air to depress the lean flammability limit of the mixed fuel. This is shown to permit very lean combustion with its low NOx production while simulataneously providing an increased flame stability margin with which to maintain low CO and HC emission. Experimental emission characteristics and selected analytical results are presented for a cylindrical research combustor designed for operation with inlet-air state conditions typical for a 30:1 compression ratio, high bypass ratio, turbofan commercial engine.

  1. Predictions and measurements of isothermal flowfields in axisymmetric combustor geometries. Ph.D. Thesis. Final Report

    NASA Technical Reports Server (NTRS)

    Rhodes, D. L.; Lilley, D. G.

    1985-01-01

    Numerical predictions, flow visualization experiments and time-mean velocity measurements were obtained for six basic nonreacting flowfields (with inlet swirl vane angles of 0 (swirler removed), 45 and 70 degrees and sidewall expansion angles of 90 and 45 degrees) in an idealized axisymmetric combustor geometry. A flowfield prediction computer program was developed which solves appropriate finite difference equations including a conventional two equation k-epsilon eddy viscosity turbulence model. The wall functions employed were derived from previous swirling flow measurements, and the stairstep approximation was employed to represent the sloping wall at the inlet to the test chamber. Recirculation region boundaries have been sketched from the entire flow visualization photograph collection. Tufts, smoke, and neutrally buoyant helium filled soap bubbles were employed as flow tracers. A five hole pitot probe was utilized to measure the axial, radial, and swirl time mean velocity components.

  2. Optimization of limestone sizing for CFB combustors: Results of pilot plant and bench-scale testing

    SciTech Connect

    Alliston, M.; Edvardsson, C.; Wu, S.; Probst, S.

    1994-12-31

    A grant to study the performance of limestones in a Circulating Fluidized Bed Combustor was obtained in 1991 from the Pennsylvania Energy Development Authority (PEDA) by Tampella Power Corporation (TPC). The overall objective of this PEDA project was to carry out a systematic pilot plant tests at TPC`s pilot plant in Williamsport, Pennsylvania, in systematic order to identify ways of improving sulfur capture and limestone utilization through better control of the size distribution and residence time of the limestone particles in the furnace. It was also an objective to determine if bench scale testing could be of value in predicting CFB sorbent behavior. The pilot plant and bench test results were incorporated into an empirical Correlation which accounts for the size distribution and residence time of solids in CFB boiler.

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

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2004-01-01

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

  4. 46 CFR 169.609 - Exhaust systems.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... Electrical Internal Combustion Engine Installations § 169.609 Exhaust systems. Engine exhaust installations... separate from the engine cooling system, a suitable warning device must be provided to indicate a...

  5. 46 CFR 169.609 - Exhaust systems.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... Electrical Internal Combustion Engine Installations § 169.609 Exhaust systems. Engine exhaust installations... separate from the engine cooling system, a suitable warning device must be provided to indicate a...

  6. 46 CFR 169.609 - Exhaust systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... Electrical Internal Combustion Engine Installations § 169.609 Exhaust systems. Engine exhaust installations... separate from the engine cooling system, a suitable warning device must be provided to indicate a...

  7. 46 CFR 169.609 - Exhaust systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... Electrical Internal Combustion Engine Installations § 169.609 Exhaust systems. Engine exhaust installations... separate from the engine cooling system, a suitable warning device must be provided to indicate a...

  8. 46 CFR 169.609 - Exhaust systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... Electrical Internal Combustion Engine Installations § 169.609 Exhaust systems. Engine exhaust installations... separate from the engine cooling system, a suitable warning device must be provided to indicate a...

  9. 14 CFR 25.1123 - Exhaust piping.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Exhaust System § 25.1123 Exhaust piping. For powerplant... corrosion resistant, and must have provisions to prevent failure due to expansion by operating...

  10. 14 CFR 25.1123 - Exhaust piping.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Exhaust System § 25.1123 Exhaust piping. For powerplant... corrosion resistant, and must have provisions to prevent failure due to expansion by operating...

  11. 14 CFR 25.1123 - Exhaust piping.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Exhaust System § 25.1123 Exhaust piping. For powerplant... corrosion resistant, and must have provisions to prevent failure due to expansion by operating...

  12. 14 CFR 25.1123 - Exhaust piping.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Exhaust System § 25.1123 Exhaust piping. For powerplant... corrosion resistant, and must have provisions to prevent failure due to expansion by operating...

  13. 14 CFR 25.1123 - Exhaust piping.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Exhaust System § 25.1123 Exhaust piping. For powerplant... corrosion resistant, and must have provisions to prevent failure due to expansion by operating...

  14. Engine Would Recover Exhaust Energy More Efficiently

    NASA Technical Reports Server (NTRS)

    Dimpelfeld, Philip M.

    1993-01-01

    Exhaust energy used for supercharging and extra shaft power. Flow of exhaust apportioned by waste gate to meet demand of turbocharger, and portion not fed to turbocharger sent to power-recovery turbine. Expected to increase fuel efficiency.

  15. 49 CFR 325.91 - Exhaust systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... sound reduction, such as exhaust gas leaks or alteration or deterioration of muffler elements, (small traces of soot on flexible exhaust pipe sections shall not constitute a violation of this subpart);...

  16. 40 CFR 1065.130 - Engine exhaust.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ....130 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR POLLUTION CONTROLS... exhaust tubing that has either a wall thickness of less than 2 mm or is air gap-insulated to minimize... balance of fuel, intake air, and exhaust according to § 1065.655 to verify exhaust system integrity....

  17. 40 CFR 1065.130 - Engine exhaust.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ....130 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR POLLUTION CONTROLS... exhaust tubing that has either a wall thickness of less than 2 mm or is air gap-insulated to minimize... balance of fuel, intake air, and exhaust according to § 1065.655 to verify exhaust system integrity....

  18. 40 CFR 1065.130 - Engine exhaust.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ....130 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR POLLUTION CONTROLS... exhaust tubing that has either a wall thickness of less than 2 mm or is air gap-insulated to minimize..., intake air, and exhaust according to § 1065.655 to verify exhaust system integrity. (f)...

  19. 40 CFR 1065.130 - Engine exhaust.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ....130 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR POLLUTION CONTROLS... exhaust tubing that has either a wall thickness of less than 2 mm or is air gap-insulated to minimize... balance of fuel, intake air, and exhaust according to § 1065.655 to verify exhaust system integrity....

  20. 40 CFR 1065.130 - Engine exhaust.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ....130 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR POLLUTION CONTROLS... exhaust tubing that has either a wall thickness of less than 2 mm or is air gap-insulated to minimize... balance of fuel, intake air, and exhaust according to § 1065.655 to verify exhaust system integrity....

  1. 46 CFR 128.320 - Exhaust systems.

    Code of Federal Regulations, 2013 CFR

    2013-10-01

    ... 46 Shipping 4 2013-10-01 2013-10-01 false Exhaust systems. 128.320 Section 128.320 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) OFFSHORE SUPPLY VESSELS MARINE ENGINEERING: EQUIPMENT AND SYSTEMS Main and Auxiliary Machinery § 128.320 Exhaust systems. No diesel-engine exhaust...

  2. 46 CFR 128.320 - Exhaust systems.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 46 Shipping 4 2014-10-01 2014-10-01 false Exhaust systems. 128.320 Section 128.320 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) OFFSHORE SUPPLY VESSELS MARINE ENGINEERING: EQUIPMENT AND SYSTEMS Main and Auxiliary Machinery § 128.320 Exhaust systems. No diesel-engine exhaust...

  3. 46 CFR 128.320 - Exhaust systems.

    Code of Federal Regulations, 2011 CFR

    2011-10-01

    ... 46 Shipping 4 2011-10-01 2011-10-01 false Exhaust systems. 128.320 Section 128.320 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) OFFSHORE SUPPLY VESSELS MARINE ENGINEERING: EQUIPMENT AND SYSTEMS Main and Auxiliary Machinery § 128.320 Exhaust systems. No diesel-engine exhaust...

  4. 46 CFR 128.320 - Exhaust systems.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 46 Shipping 4 2012-10-01 2012-10-01 false Exhaust systems. 128.320 Section 128.320 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) OFFSHORE SUPPLY VESSELS MARINE ENGINEERING: EQUIPMENT AND SYSTEMS Main and Auxiliary Machinery § 128.320 Exhaust systems. No diesel-engine exhaust...

  5. 46 CFR 128.320 - Exhaust systems.

    Code of Federal Regulations, 2010 CFR

    2010-10-01

    ... 46 Shipping 4 2010-10-01 2010-10-01 false Exhaust systems. 128.320 Section 128.320 Shipping COAST GUARD, DEPARTMENT OF HOMELAND SECURITY (CONTINUED) OFFSHORE SUPPLY VESSELS MARINE ENGINEERING: EQUIPMENT AND SYSTEMS Main and Auxiliary Machinery § 128.320 Exhaust systems. No diesel-engine exhaust...

  6. Segregated exhaust SOFC generator with high fuel utilization capability

    DOEpatents

    Draper, Robert; Veyo, Stephen E.; Kothmann, Richard E.

    2003-08-26

    A fuel cell generator contains a plurality of fuel cells (6) in a generator chamber (1) and also contains a depleted fuel reactor or a fuel depletion chamber (2) where oxidant (24,25) and fuel (81) is fed to the generator chamber (1) and the depleted fuel reactor chamber (2), where both fuel and oxidant react, and where all oxidant and fuel passages are separate and do not communicate with each other, so that fuel and oxidant in whatever form do not mix and where a depleted fuel exit (23) is provided for exiting a product gas (19) which consists essentially of carbon dioxide and water for further treatment so that carbon dioxide can be separated and is not vented to the atmosphere.

  7. Sixth annual coal preparation, utilization, and environmental control contractors conference

    SciTech Connect

    Not Available

    1990-01-01

    A conference was held on coal preparation, utilization and environmental control. Topics included: combustion of fuel slurries; combustor performance; desulfurization chemically and by biodegradation; coal cleaning; pollution control of sulfur oxides and nitrogen oxides; particulate control; and flue gas desulfurization. Individual projects are processed separately for the databases. (CBS).

  8. Exhaust emission control and diagnostics

    DOEpatents

    Mazur, Christopher John; Upadhyay, Devesh

    2006-11-14

    A diesel engine emission control system uses an upstream oxidation catalyst and a downstream SCR catalyst to reduce NOx in a lean exhaust gas environment. The engine and upstream oxidation catalyst are configured to provide approximately a 1:1 ratio of NO to NO2 entering the downstream catalyst. In this way, the downstream catalyst is insensitive to sulfur contamination, and also has improved overall catalyst NOx conversion efficiency. Degradation of the system is determined when the ratio provided is no longer near the desired 1:1 ratio. This condition is detected using measurements of engine operating conditions such as from a NOx sensor located downstream of the catalysts. Finally, control action to adjust an injected amount of reductant in the exhaust gas based on the actual NO to NO2 ratio upstream of the SCR catalyst and downstream of the oxidation catalyst.

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

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Kopasakis, George

    2005-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2008-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Fear, J. S.

    1973-01-01

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

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

    DOEpatents

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

    2009-10-20

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

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

    SciTech Connect

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

    2012-11-20

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

  16. Fuel-Flexible Gas Turbine Combustor Flametube Facility

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

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

  17. Self-regulating fuel staging port for turbine combustor

    DOEpatents

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

    2014-07-08

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

  18. Reduction of nitric oxide emissions from a combustor

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Cohen, J. D.

    1983-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

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

    NASA Technical Reports Server (NTRS)

    McCarty, Bob; Tomondi, Chris; McGinley, Ray

    2004-01-01

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

  2. Reduction of nitric oxide emissions from a combustor

    SciTech Connect

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

    1980-05-27

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

  3. Computational investigation on combustion instabilities in a rocket combustor

    NASA Astrophysics Data System (ADS)

    Yuan, Lei; Shen, Chibing

    2016-10-01

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

  4. Investigation of a small solid fuel ramjet combustor

    SciTech Connect

    Levy, Yeshayahou; Gany, Alon; Zvuloni, Roni

    1989-06-01

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

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

    DOEpatents

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

    2015-09-01

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

  6. A mathematical model for jet engine combustor pollutant emissions

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

  7. Conceptual model of turbulent flameholding for scramjet combustors

    NASA Technical Reports Server (NTRS)

    Huber, P. W.

    1980-01-01

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

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

    SciTech Connect

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

    2016-04-05

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

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

    NASA Technical Reports Server (NTRS)

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

    2017-01-01

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

  10. Numerical Study of Low Emission Gas Turbine Combustor Concepts

    NASA Technical Reports Server (NTRS)

    Yang, Song-Lin

    2002-01-01

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

  11. System for supporting bundled tube segments within a combustor

    DOEpatents

    Melton, Patrick Benedict

    2016-03-01

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

  12. Numerical Investigation of a Model Scramjet Combustor Using DDES

    NASA Astrophysics Data System (ADS)

    Shin, Junsu; Sung, Hong-Gye

    2017-04-01

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

  13. Space shuttle exhaust cloud properties

    NASA Technical Reports Server (NTRS)

    Anderson, B. J.; Keller, V. W.

    1983-01-01

    A data base describing the properties of the exhaust cloud produced by the launch of the Space Transportation System and the acidic fallout observed after each of the first four launches was assembled from a series of ground and aircraft based measurements made during the launches of STS 2, 3, and 4. Additional data were obtained from ground-based measurements during firings of the 6.4 percent model of the Solid Rocket Booster at the Marshall Center. Analysis indicates that the acidic fallout is produced by atomization of the deluge water spray by the rocket exhaust on the pad followed by rapid scavening of hydrogen chloride gas aluminum oxide particles from the Solid Rocket Boosters. The atomized spray is carried aloft by updrafts created by the hot exhaust and deposited down wind. Aircraft measurements in the STS-3 ground cloud showed an insignificant number of ice nuclei. Although no measurements were made in the column cloud, the possibility of inadvertent weather modification caused by the interaction of ice nuclei with natural clouds appears remote.

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  15. System for tuning a combustor of a gas turbine

    SciTech Connect

    Hughes, Michael John

    2016-12-27

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

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

    DTIC Science & Technology

    2010-05-01

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

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

    DTIC Science & Technology

    2010-02-01

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

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

    DTIC Science & Technology

    2014-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Sivo, Joseph N; Fenn, David B

    1956-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Davoudzadeh, Farhad; Liu, Nan-Suey

    2005-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

  4. Sectoral combustor for burning low-BTU fuel gas

    DOEpatents

    Vogt, Robert L.

    1980-01-01

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

  5. Massively parallel computation of three-dimensional scramjet combustor

    NASA Astrophysics Data System (ADS)

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

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

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

    SciTech Connect

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

    2016-12-27

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

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

    NASA Astrophysics Data System (ADS)

    Rabadan, E.; Weigand, B.

    2013-03-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Moorhead, Paul E.

    1988-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1975-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

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

    DOEpatents

    Clingerman, Bruce J.; Chalfant, Robert W.

    2002-01-01

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

  13. Performance Cycle Analysis of a Two-Spool, Separate-Exhaust Turbofan With Interstage Turbine Burner

    NASA Technical Reports Server (NTRS)

    Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

    2005-01-01

    This paper presents the performance cycle analysis of a dual-spool, separate-exhaust turbofan engine, with an Interstage Turbine Burner serving as a secondary combustor. The ITB, which is located at the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet engine propulsion. A detailed performance analysis of this engine has been conducted for steady-state engine performance prediction. A code is written and is capable of predicting engine performances (i.e., thrust and thrust specific fuel consumption) at varying flight conditions and throttle settings. Two design-point engines were studied to reveal trends in performance at both full and partial throttle operations. A mission analysis is also presented to assure the advantage of saving fuel by adding ITB.

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

    SciTech Connect

    Richter, W.

    1981-02-01

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

  15. The 3D rocket combustor acoustics model

    NASA Technical Reports Server (NTRS)

    Priem, Richard J.; Breisacher, Kevin J.

    1992-01-01

    The theory and procedures for determining the characteristics of pressure oscillations in rocket engines with prescribed burning rate oscillations are presented. Analyses including radial and hub baffles and absorbers can be performed in one, two, and three dimensions. Pressure and velocity oscillations calculated using this procedure are presented for the SSME to show the influence of baffles and absorbers on the burning rate oscillations required to achieve neutral stability. Comparisons are made between the results obtained utilizing 1-D, 2-D, and 3-D assumptions with regards to capturing the physical phenomena of interest and computational requirements.

  16. Optical Diagnosis of Gas Turbine Combustors Being Conducted

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  17. Flame structure and stabilization in miniature liquid film combustors

    NASA Astrophysics Data System (ADS)

    Pham, Trinh Kim

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

  18. Predictive models of circulating fluidized bed combustors: SO[sub 2] sorption in the CFB loop

    SciTech Connect

    Gidaspow, D.; Therdthianwong, A. . Dept. of Chemical Engineering)

    1993-02-01

    The overall objective of this investigation is to develop experimentally verified models for circulating fluidized bed (CFB) combustors. Sorption of S0[sub 2] with calcined limestone was studied in a PYROFLOW type CFB loop at conditions approximating those found in a CFB combustor. Initially the CFB loop contained 150 micron CaO particles of a density of 3.3 g/cm[sup 3] and air at 1143[degrees]K and 3.25 atm. Atzero time, air containing 600 ppm SO[sub 2], was introduced into the riser bottom at 1143[degrees]K. The effect of gas velocity, sorbent inventory and inlet pressure on the sorption of SO[sub 2], were studied isothermally by running our hydrodynamic code with the S0[sub 2] sorption conservation of species equation. At a velocity of 5m/sec., reported to be a typical velocity by PYROPOWER, there is reasonably good S0[sub 2] removal. At 10 m/sec the S0[sub 2] removal is poor. The best SO[sub 2], removal is for a velocity of 5 m/s and a high bed inventory, initial bed height, H = 9m. Most of the S0[sub 2] is removed in the first two meters of the reactor. However, the S0[sub 2] removal is not complete at the bed outlet. This is due to mixing. At the left wall of the reactor (wall opposite the solids inlet) the S0[sub 2] removal was poor due to gas bypassing caused by the asymmetrical solids inlet. Simulation of the PYROPOWER loop with a symmetrical inlet gave us an order of magnitude improvement over the conventional PYROPOWER system. These results demonstrate the practical utility of the predictive model that we have developed over the last three years.

  19. Nitration of polynuclear aromatic hydrocarbons in coal combustors and exhaust streams. Quarterly report, January 1, 1992--March 31, 1992

    SciTech Connect

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

    1992-12-01

    Regarding PAH preparation, our efforts this quarter were directed at identifying the operating conditions in the laboratory furnace that clearly resolve primary devolatilization from secondary pyrolysis. In previous studies, pristine tars from primary devolatilization were prepared in the radiant coal flow experiment with a furnace length of only 5 cm. Tars at various extents of secondary pyrolysis were generated in a hot-zone of 12.5 cm. For this project we decided to try to prepare both kinds of PAH in the 12.5 cm hot zone. Preliminary screening surveys showed that, among the dime possible ways to compensate for the longer furnace tube, only lowering the fumace temperature is really effective. Lowering the loading from its already low value of 300 No./cm{sup 3} is inconsequential, and raising the gas velocity to lower residence times in the longer hot zones causes deposition problems. The coal jet undergoes the transition to turbulence in less than 12 cm for entrainment velocities greater than 1 cm/s. By the end of the quarter, it was clear that the best value for primary devolatilization studies will fall within the range of 1 150 to 1400 K. But at present we are still not completely satisfied with the resolution of primary devolatilization. As intended, the tars collected at short residence times are completely soluble in tetrahydrofuran (THF), indicating that no soot is present so no secondary pyrolysis has occurred. But at residence times long enough to attain ultimate primary devolatilization yields, some soot is present. We will make our final choice early in the next quarter, then begin to collect an inventory of primary devoatilization tars for subsequent detailed characterization.

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

    NASA Technical Reports Server (NTRS)

    Childs, J. Howard; McCafferty, Richard J.

    1948-01-01

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