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At least 595 records · Page 33

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NASA NTRS1983

Flame radiation and liner heat transfer in a tubular-can combustor

Heat transfer within a combuster were examined. Total and spectral flame radiation in a tubular can combustor at a series of parametric operating conditions was measured. Radiation measurements were taken for a range of inlet air pressures from 0.34 to 2.0 MPa, inlet air temperatures from 533 to 700 K, with two different fuels, Jet-A and ERBS. Measurements of liner temperatures combined with the parametric radiation results allowed a calculation of the combustor liner heat loads. Flame emissivity was determined from the spectral measurements.

Claus, R. W.↗

NASA NTRS1984

A review of NASA combustor and turbine heat transfer research

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.

Rudey, R. A.↗

NASA NTRS1984

Combustor flame flashback

Flashback, a problem that occurs in premixed-prevaporized combustors, is the upstream propagation of the flame from the combustor into the premixing tubes. Not only does flashback change the combustion process from premixed burning to diffusion burning, thus creating more pollutants, but it also inflicts considerable damage to the fuel injector, premixing tube and other equipment upstream. The conditions at which flashback occurs in steady burning and the mechanism that causes flashback in both steady and transient flow are studied. The equivalence ratio at which flashback occurs is being measured for inlet temperatures of 600-950 K, premixer wall temperatures of 450-1050 K and premixer velocities of 40-80 ft/s. These data are presented.

Proctor, M. P.↗

NASA NTRS1984

Analytical fuel property effects: Small combustors

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

Cohen, J. D.↗

NASA NTRS1984

A theoretical model for the cross spectra between pressure and temperature downstream of a combustor

A theoretical model developed to calculate pressure-temperature cross spectra, pressure spectra, temperature spectra and pressure cross spectra in a ducted combustion system is presented. The model assumes the presence of a fluctuating-volumetric-heat-release-rate disk source and takes into account the spatial distribution of the steady-state volumetric-heat flux. Using the model, pressure, velocity, and temperature perturbation relationships can be obtained. The theoretical results show that, at a given air mass flow rate, the calculated pressure-temperature cross spectra phase angle at the combustor exit depends on the model selected for the steady-state volumetric-heat flux in the combustor. Using measurements of the phase angle, an appropriate source region model was selected. The model calculations are compared with the data. The comparison shows good agreement and indicates that with the use of this model the pressure-temperature cross spectra measurements provide useful information on the physical mechanisms active at the combustion noise source.

Miles, J. H.↗

NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

The application of ceramics to gas turbine combustor liners to reduce liner metal temperature was studied in an experiment in which yttria-stabilized zirconia plasma was sprayed on compliant metal substrates exposed to near stoichiometric combustion. The strain isolation pad materials chosen were Hoskins Alloy 875 and BRUNSLLOY 534 Fiber Metal of 0.25 and 0.38 cm thicknesses and 35 and 45 percent density levels. Combustor screening tests of all specimens showed no evidence of deterioration or failure. Specimens exposed to flame temperatures in excess of 2100 K were convectively or convective-transpiration cooled and were evaluated in a 10 cm sq flame tube at inlet air temperature of 533 K and pressure of 0.5 MPa. The results suggest the superiority of a system composed of the Hoskins Alloy 875 compliant pad with 0.25 cm thickness and 35 percent density coupled with a NiCrAlY bond coat and a 8 percent Y2O3-ZrO2 ceramic top coat of 0.19 cm thickness.

Ercegovic, D. B.↗

NASA NTRS1984

Calculations of turbulent mass transport in a bluff-body diffusion-flame combustor

Experimental measurements of turbulent mass transit in a bluff-body diffusion-flame combustor (using CO2 instead of fuel) are analyzed to evaluate the accuracy of physical models used in flow computations. The data of Lightman and Magill (1981), Lightman et al. (1983), and Roquemore et al. (1983) are used to calculate apparent turbulent Schmidt numbers (TSN) for a series of flow conditions by a modified TEACH viscous-flow code. The modeling principles, calculation grid, and boundary conditions are discussed, and the results are presented in graphs comparing calculated and measured values. It is shown that models such as the two-equation (K - epsilon) model which use a single value of the TSN are inappropriate for conditions typical of gas-turbine combustors: local TSN variations due to turbulence behavior must be taken into account in improved models.

Sturgess, G. J.↗

NASA NTRS1984

Flame radiation and linear heat transfer in a tubular can combustor

Heat transfer within a combuster were examined. Total and spectral flame radiation in a tubular can combustor at a series of parameteric operating conditions was measured. Radiation measurements were taken for a range of inlet air pressures from 0.34 to 2.0 MPa, inlet air temperatures from 533 to 700 K, with two different fuels, Jet-A and ERBS. Measurements of linear temperatures combined with the parametric radiation results allowed a calculation of the combustor linear heat loads. Flame emissivity was determined from the spectral measurements. Previously announced in STAR as N84-13188

Claus, R. W.↗

NASA NTRS1985

Modeling a Transient Catalytic Combustor

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

Tien, J. S.↗

NASA NTRS1985

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

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

Riddlebaugh, S. M.↗

NASA NTRS1985

Combustor turbulence

The turbulence entering the turbine is produced in the combustor. High turbulence levels from the combustor can alter the location of the transition point on the turbine vane. The dynamics of turbulence and the progress being made in computing the flow are discussed.

Marek, C. J.↗

NASA NTRS1985

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

Riddlebaugh, S. M.↗

NASA NTRS1984

Concentration distributions in cylindrical combustors

Experimental studies have been conducted with the aim to obtain a better understanding of the fluid dynamics of mixing in gas turbine combustors, and solid fuel ramjet combustors subject to spin. The present investigation represents a continuation of studies conducted by So et al. (1984). It is also concerned with the verification of some conclusions reported by Ahmed et al. (1984). Attention is given to the experimental facility and instrumentation, centerline concentration measurements, mean concentration profiles, and a comparison of concentration and axial velocity results in the case of swirling flow.

Ahmed, S. A.↗

NASA NTRS1986

Toward improved durability in advanced combustors and turbines - Progress in prediction of thermomechanical loads

NASA is sponsoring the Turbine Engine Hot Section Technology (HOST) Project to address the need for improved durability in advanced combustors and turbines. Analytical and experimental activities aimed at more accurate prediction of the aerothermal environment, the thermomechanical loads, the material behavior and structural responses to such loading, and life predictions for high temperature cyclic operation have been underway for several years and are showing promising results. Progress is reported in the development of advanced instrumentation and in the improvement of combustor aerothermal and turbine heat transfer models that will lead to more accurate prediction of themomechanical loads.

Sokolowski, D. E.↗

NASA NTRS1987

Computer Program for Flow in a Combustor

New version of TEACH computer program developed specifically for use in analysis of subsonic, swirling, reacting, turbulent flow in cylindrical bluff-body research combustor. Combustor design widely used as research tool for mathematical modeling of gas turbines and for development of diagnostic instrumentation.

Chiappetta, Louis M.↗

NASA NTRS1986

Toward improved durability in advanced combustors and turbines: Progress in the prediction of thermomechanical loads

Sokolowski, Daniel E.↗

NASA NTRS1986

Numerical prediction of cold turbulent flow in combustor configurations with different centerbody flame holders

A numerical study of cold turbulent flow in a combustor containing a centerbody flame holder is presented. The axisymmetric Navier-Stokes equations incorporating a k-epsilon turbulence model were solved in a nonorthogonal curvilinear coordinate system. The finite volume method, applied to a staggered grid system, was used to discretize the differential equations. These finite differenced equations were then solved iteratively utilizing the SIMPLE algorithm described by Patankar (1980). Solutions were obtained for two combustor duct geometries with various centerbody flame holders which included a disk, a cone, and a sphere. The extent of mixing due to these bodies was evaluated. A comparison with previously obtained experimental data yielded moderately good agreement.

Yung, C. N.↗

NASA NTRS1987

Three-dimensional numerical predictions of the flow behind a rearward-facing-step in a supersonic combustor

A CFD code was developed to compute the mixing and combustion of hydrogen fuel in the turbulent flow fields of scramjets. The code is based on the complete Reynold's time-averaged three-dimensional Navier-Stokes equations, and chemical species equations. The code includes a global finite rate hydrogen-air chemical reaction model and the Baldwin-Lomax (1978) turbulence model. The validity of this CFD code was demonstrated by comparing computed results from the code with existing data for three configurations, each containing an important feature of the scramjet combustor flow. The code was then applied to the reacting flow field in a complex combustor model in which hydrogen fuel was injected transversely into a supersonic hot airstream through several fuel injectors.

Uenishi, K.↗