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At least 1,477 records · Page 82

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NASA NTRS1992

Evaluation and application of the Baldwin-Lomax turbulence model in two-dimensional, unsteady, compressible boundary layers with and without separation in engine inlets

The Baldwin-Lomax model is used in many CFD codes because it is quick and easy to implement. In this paper, we discuss implementing the Baldwin-Lomax turbulence model for both steady and unsteady compressible flows. In addition, these flows may be either separated or attached. In order to apply this turbulence model to flows which may be subjected to these conditions, certain modifications should be made to the original Baldwin-Lomax model. We discuss these modifications and determine whether the Baldwin-Lomax model is a viable turbulence model that produces reasonably accurate results for high speed flows that can be found in engine inlets.

Sakowski, Barbara↗

NASA NTRS1992

Evaluation and application of the Baldwin-Lomax turbulence model in two-dimensional, unsteady, compressible boundary layers with and without separation in engine inlets

Sakowski, Barbara↗

NASA NTRS1993

Results of Low Power Deicer tests on a swept inlet component in the NASA Lewis Icing Research Tunnel

Tests were conducted under a USAF/NASA Low Power Deicer program on two expulsive technologies to examine system performance on hardware representative of a modern aircraft part. The BF Goodrich Electro-Expulsive Deicing System and Pneumatic Impulse Ice Protection system were installed on a swept, compound curve, engine inlet component with varying leading edge radius, and tested through a range of icing and system operating conditions in the NASA Lewis Icing Research Tunnel. A description of the experimental procedure and results, including residual ice thickness, shed ice particle size, and changes in system energy/pressure characteristics are presented.

Bond, Thomas H.↗

NASA NTRS1993

Foreign object impact assessment of a high-Mach engine inlet

The paper presents a foreign-object damage (FOD) assessment of the engine inlet region of a high-velocity aircraft design. Of particular interest is the potential for damage to the active cooling system, which is an integral part of the external skin. A variety of impactors is considered, including hailstones, birds, and rocks of various sizes. An explicit finite element solution, which permits detailed modeling of inelastic material behavior and other nonlinearities, is used for the simulations. The paper also discusses some general aspects of foreign object impact modeling and explicit solution techniques.

Thomson, D. E.↗

NASA NTRS1993

A three-dimensional pressure flux-split RNS application to sub/supersonic flow in inlets and ducts

The reduced Navier-Stokes (RNS) formulation is combined with a pressure based flux split procedure for the computational analysis of three dimensional flow. Only lowest order diffusion, that is required to satisfy the no-slip boundary condition is retained in this approximation. The governing equations are appropriately differenced such that the physical boundary conditions are closely coupled and combined to form a closed discrete system. No numerical or characteristic type boundary conditions are required. A global pressure relaxation procedure, i.e. multi-sweep PNS, is considered. A sparse matrix direct solver is applied to the crossplane solution. At the outflow boundary only the pressure or pressure gradient is prescribed. The resulting formulation has been applied to the solution of three-dimensional subsonic/supersonic internal flows in inlets and ducts.

Khosla, P. K.↗

NASA NTRS1990

NASA. Lewis Research Center/IFMD inlet duct and nozzle high speed validation experiments

Personnel from NASA Lewis Research Center spoke on the subject of inlet duct and nozzle high speed validation experiments to include crossing shocks and boundary layer interaction, unsteady shock/boundary layer interactions, and vortex generators. Attention was also paid to the subjects of high speed mixing and transition ducts. Specific application was made to the NASP hypermixing concepts.

Hingst, Warren↗

NASA NTRS1948

Bibliography of NACA and Other Reports on Air Inlets and Internal Flows

At the suggestion of the Subcommittee on Internal Flow, the National Advisory Committee for Aeronautics has assembled this bibliography on air inlet and internal flow data for aircraft. It was felt that the purpose of this publication, to furnish the engineer and researcher with a listing of theoretical and experimental studies related directly and indirectly to the field of internal flows, could be served best by presenting some of the more pertinent and interesting references. Thus this bibliography is not complete in its listing of internal flow reports from the NACA or from other sources. The works referenced here are, for the most part, those published before January 1948. To aid the investigator in finding references, the bibliography has been indexed according to subject matter. NACA reports and reports from other sources are grouped separately. The subject breakdown used may be found in the table of contents on the preceding pages.

Bibliographies and indexes↗

NASA NTRS1947

Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets

Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Jones, William L.↗

NASA NTRS1950

Performance of 4600-pound-thrust centrifugal-flow-type turbojet engine with water-alcohol injection at inlet

An experimental investigation of the effects of injecting a water-alcohol mixture of 2:1 at the compressor inlet of a centrifugal-flow type turbojet engine was conducted in an altitude test chamber at static sea-level conditions and at an altitude of 20,000 feet with a flight Mach number of 0.78 with an engine operating at rated speed. The net thrust was augmented by 0.16 for both flight conditions with a ratio of injected liquid to air flow of 0.05. Further increases in the liquid-air ratio did not give comparable increases in thrust.

Glasser, Philip W↗

NASA NTRS1950

Correlation of Forced-convection Heat-transfer Data for Air Flowing in Smooth Platinum Tube with Long-approach Entrance at High Surface and Inlet-air Temperatures

A heat-transfer investigation was conducted with air in an electrically heated platinum tube with long-approach entrance, inside diameter of 0.525 inch, and effective heat-transfer length of 24 inches over ranges of Reynolds number up to 320,000, average inside-tube-wall temperature up to 3053 degrees R, and inlet-air temperature up to 1165 degrees R. Correlation of data by the conventional Nusselt relation resulted in separation of data with tube-wall temperature. Good correlation was obtained, however, by use of a modified Reynolds number.

Desmon, Leland G↗

NASA NTRS1951

Secondary flows in annular cascades and effects on flow in inlet guide vanes

Qualitative discussion is presented of the general nature of secondary flows in stationary annular cascades with thin wall boundary layers and radial design variation of circulation. Deviations from ideal mean outlet flows (based on blade-element performance) exist in potential-flow region of vanes because of conditions imposed by end-wall boundaries, displacement of wall boundary layers toward blade suction surfaces, and irrotationality requirement. As a consequence of existence of nonuniform radial flow across blade spacing, it may not generally be possible to obtain an arbitrarily specified design variation of the turning angle along the radial height of a blade row. Quantitative turning angle corrections due to effects of secondary flows in axial-flow compressor inlet guide vanes were obtained from induced deflections of a superimposed vortex system in conjunction with an empirically determined correlation factor.

CASCADES, EXPERIMENT↗

NASA NTRS1953

Analytical evaluation of effect of equivalence ratio inlet-air temperature and combustion pressure on performance of several possible ram-jet fuels

The results of an analytical investigation of the theoretical air specific impulse performance and adiabatic combustion temperatures of several possible ram-jet fuels over a range of equivalence ratios, inlet-air temperatures, and combustion pressures, is presented herein. The fuels include octane-1, 50-percent-magnesium slurry, boron, pentaborane, diborane, hydrogen, carbon, and aluminum. Thermal effects from high combustion temperatures were found to effect considerably the combustion performance of all the fuels. An increase in combustion pressure was beneficial to air specific impulse at high combustion temperatures. The use of these theoretical data in engine operation and in the evaluation of experimental data is described.

COMBUSTION - TURBINE ENGINES↗

NASA NTRS1954

Experimental investigation at Mach numbers 1.88, 3.16, and 3.83 of pressure drag of wedge diverters simulating boundary-layer-removal systems for side inlets

The pressure drag coefficients for a number of simulated wedge diverter boundary-layer-removal systems for side inlets were obtained at Mach numbers of 1.88m 3.16, and 3.83. Wedge included angle, wedge height, and wedge axial position relative to the splitter plate were independently varied for unswept and swept splitter plate configurations immersed in the turbulent boundary layer of a flat plate. Friction and total drag coefficients for several configurations were also obtained at Mach number 3.16

Piercy, Thomas G↗

NASA NTRS1956

Investigation of the air-flow regulation characteristics of a translating-spike inlet with two oblique shocks from Mach 1.6 to 2.0

The pressure recovery of an axially symmetric translating-spike inlet was essentially the same as for a single cone with the same total angle. In order to match a turbojet engine over the Mach range of 1.6 to 2.0, the spike translation must be larger than it is for a single cone.

Nettles, J C↗

NASA NTRS1956

Performance Characteristics of an Axial-flow Transonic Compressor Operating up to Tip Relative Inlet Mach Number of 1.34

Performance data are presented for a transonic axial-flow compressor rotor designed to operate at a tip speed of 1300 ft/sec with maximum relative tip Mach number of 1.37. The compressor had an inlet diameter of 16 inches, a hub-tip diameter ratio of 0.5 and design specific weight flow of 31.1 (lb/sec/(sq ft frontal area). Experimental values of relative total-pressure-loss coefficient were considerably higher than the assumed values. This disparity, hub choking, and application of the simple radial-equilibrium concept are discussed. The data of this report are used to extend previously presented correlation plots of compressor design parameters to higher Mach numbers.

Creagh, John W R↗

NASA NTRS1957

Experimental evaluation of methods for improving diffuser-exit total-pressure profiles for a side-inlet model at Mach number 1.91

Distortion effects of the following model additions or changes were studied: (1) constant-area mixing lengths at the end of the subsonic diffuser, (2) screens with systematic variations in mesh and flow-area blockage, (3) bleeds slots on the inlet cowl, and (4) a simulated compressor hub. Data at a diffuser-exit Mach number of 0.36 are compared with data of a similar study at Mach 0.20. At the higher Mach number, the same length of constant-area mixing was less effective in reducing total-pressure distortions, while the use of screens was more effective.

Klann, John L↗

NASA NTRS1928

The comparative performance of an aviation engine at normal and high inlet air temperatures

This report presents some results obtained during an investigation to determine the effect of high inlet air temperature on the performance of a Liberty 12 aviation engine. The purpose of this investigation was to ascertain, for normal service carburetor adjustments and a fixed ignition advance, the relation between power and temperature for the range of carburetor air temperatures that may be encountered when supercharging to sea level pressure at altitudes of over 20,000 feet and without intercooling when using plain aviation gasoline and mixtures of benzol and gasoline. The results show that for the conditions of test, both the brake and indicated power decrease with increase in air temperature at a faster rate than given by the theoretical assumption that power varies inversely as the square root of the absolute temperature. On a brake basis, the order of the difference in power for a temperature difference of 120 degrees F. Is 3 to 5 per cent. The observed relation between power and temperature when using the 30-70 blend was found to be linear. But, although these differences are noted, the above theoretical assumption may be considered as generally applicable except where greater precision over a wide range of temperatures is desired, in which case it appears necessary to test the particular engine under the given conditions. (author)

Gardiner, Arthur W↗

NASA NTRS1940

Effect of Fuel-Air Ratio, Inlet Temperature, and Exhaust Pressure on Detonation

An accurate determination of the end-gas condition was attempted by applying a refined method of analysis to experimental results. The results are compared with those obtained in Technical Report no. 655. The experimental technique employed afforded excellent control over the engine variables and unusual cyclic reproducibility. This, in conjunction with the new analysis, made possible the determination of the state of the end-gas at any instant to a fair degree of precision. Results showed that for any given maximum pressure the maximum permissible end-gas temperature increased as the fuel-air ratio was increased. The tendency to detonate was slightly reduced by an increase in residual gas content resulting from an increase in exhaust backpressure with inlet pressure constant.

Taylor, E S↗