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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 649 records · Page 36
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

Conceptual Aerodynamic Design of a Tail-Cone Thruster System Under Axi-Symmetric Inlet Distortion

This paper presents a conceptual design of a tail-cone thruster system which is operating under an axisymmetric inlet distortion. An effort to realize the targeted fuel burn saving that was proposed in NASA's STARC_ABL aircraft design is made through a CFD (Computational Fluid Dynamics)-based design approach. This method employs three iterative steps to exploit the CFD tools until the design requirements are met: a quasi-2D through-flow model to design the fan/EGV (Exit Guide Vane), a 3-D RANS (Reynolds Averaging Navier-Stokes) simulation of the single blade row to account for the inlet/fan and the EGV/nozzle interaction, and a 3-D RANS simulation of the airframe with a propulsor installed - propulsion airframe integration (PAI). The design requirements which include the thrust, and shaft power of the propulsor are matched throughout the evaluations coming from two CFD domains, i.e., the turbo-machinery and the PAI. During the switch between these different computational domains, the inlet and exit profiles are matched via the correction factors of the body-force model. The present tail-cone thruster (TCT) aerodynamic design leverages a low-pressure ratio fan (FPR=1.2 to approximately 1.25) of which the camber-line angles are predicted by a quasi-2D through-flow model. The quasi-2D model is derived to analyze the radially distorted flow resulting from the ingested boundary layer at the inlet. It also estimates the appropriate velocity vectors of the metal angles of the fan and EGV which is subjected to different types of vortex at the fan exit. The baseline geometry is revisited and its internal flow-path and exhaust cone are redesigned to illustrate the strong correlation among the components of the propulsor in the PAI domain. The peak efficiency point of the fan/EGV with respect to the blade counts, also known as solidity, and rotational speed is chosen for the cruise condition via parametric studies. The corresponding performance maps are presented. The resulting performance metrics of the new conceptual design of the BLI (Boundary Layer Ingestion) propulsor are analyzed and compared with these of the baseline in the PAI aspect. Finally, ideas of the CFD based design of a BLI propulsor are discussed based on the observations drawn from the numerical results.

Inlet↗
NASA NTRS1952

A Low-speed Investigation of an Annular Transonic Air Inlet

Low-speed wind-tunnel tests were conducted as preliminary steps in the study of fuselage-air-inlet arrangements believed suitable for use at transonic speeds. The forward part of the model consisted of an NACA 1-85-050 cowling located at the base of the long protruding fuselage nose designed to maintain substream surface velocities everywhere ahead of the entrance and thereby to avoid or minimize adverse boundary-layer-shock interaction effects up to low supersonic speeds. Pressure-recovery and surface-pressure measurements are presented for the model with three fuselage nose shapes for ranges of angle of attack and inlet-velocity ratio useful for high-speed flight

AIR INLETS, SCOOPS↗
DOE OSTI2024

Energy and Environmental Analyses of a Solar–Gas Turbine Combined Cycle with Inlet Air Cooling

Sensitivity to ambient air temperatures, consuming a large amount of fuel, and wasting a significant amount of heat dumped into the ambient atmosphere are three major challenges facing gas turbine power plants. This study was conducted to simultaneously solve all three aforementioned GT problems using solar energy and introducing a new configuration that consists of solar preheating and inlet-air-cooling systems. In this study, air was preheated at a combustion chamber inlet using parabolic trough collectors. Then, inlet air to the compressor was cooled by these collectors by operating an absorption cooling cycle. At the design point conditions, this novel proposed integration resulted in a 6.87% relative increase in generated power and a 10.53% relative decrement in fuel consumption, achieving a 19.45% relative increment in the plant’s thermal efficiency. This was accompanied by a reduction of 0.026 kg/s, 4.2 kg/s, and 0.278 kg/s in CO 2 , CO, and NOx emissions, respectively. Finally, spider diagrams were employed to assess the impact of the operating parameters on the overall system’s performance and its associated environmental implications.

14 SOLAR ENERGY↗
NASA NTRS1972

Experimental study of flow distribution and pressure loss with circumferential inlet and outlet manifolds

Water flow tests with circumferential inlet and outlet manifolds were conducted to determine factors affecting fluid distribution and pressure losses. Various orifice sizes and manifold geometries were tested over a range of flow velocities. With inlet manifolds, flow distribution was related directly to orifice discharge coefficients. A correlation indicated that nonuniform distribution resulted when the velocity head ratio at the orifice was not in the range of constant discharge coefficient. With outlet manifolds, nonuniform flow was related to static pressure variations along the manifold. Outlet manifolds had appreciably greater pressure losses than comparable inlet manifolds.

Dittrich, R. T.↗
NASA NTRS1972

Experimental dynamic response of a two-dimensional, Mach 2.7, mixed compression inlet

A test program was conducted on a two-dimensional supersonic inlet. Internal disturbances in diffuser exit mass flow were produced by oscillating overboard bypass doors. Open-loop dynamic responses of shock position, throat exit and diffuser exit static pressures are presented. The steady-state and dynamic coupling between ducts were also obtained. The experimental results from the two-dimensional inlet are compared to results from a similar size axisymmetric inlet and also to a transfer function synthesis program.

Baumbick, R. J.↗
NASA NTRS1972

Sonic inlet development for turbofan engines

The static program has shown to date that very large noise reductions can be achieved by the sonic inlet concept with realistic inlet length and good aerodynamic performance. It is also shown that different inlet concepts produce substantially different results.

Klujber, F.↗
NASA NTRS1972

Tests of a mixed compression axisymmetric inlet with large transonic mass flow at Mach numbers 0.6 to 2.65

A 38.8-cm (15.28-in.) capture diameter model of a mixed-compression axisymmetric inlet system with a translating cowl was designed and tested. The internal contours, designed for Mach number 2.65, provided a throat area of 59 percent of the capture area when the cowl was retracted for transonic operation. Other model features included a boundary-layer removal system, vortex generators, an engine airflow bypass system, cowl support struts, and rotating rakes at the engine face. All tunnel testing was conducted at a tunnel total pressure of about 1 atm (a unit Reynolds number of about 8.53 million/m at Mach number 2.65) at angles of attack from 0 deg to 4 deg. Results for the following were obtained: total-pressure recovery and distortion at the engine face as a function of bleed mass-flow ratio, the effect of bleed and vortex generator configurations on pressure recovery and distortion, inlet tolerance to unstart due to changes in angle of attack or Mach number, surface pressure distributions, boundary-layer profiles, and transonic additive drag. At Mach number 2.65 and with the best bleed configurations, maximum total pressure recovery at the engine face ranged from 91 to 94.5 percent with bleed mass-flow ratios from 4 to 9 percent, respectively, and total-pressure distortion was less than 10 percent. At off-design supersonic Mach numbers above 1.70, maximum total-pressure recoveries and corresponding bleed mass flows were about the same as at Mach number 2.65, with about 10 to 15 percent distortion. In the transonic Mach number range, total pressure recovery was high (above 96 percent) and distortion was low (less than 15 percent) only when the inlet mass-flow ration was reduced 0.02 to 0.06 from the maximum theoretical value (0.590 at Mach number 1.0).

Smeltzer, D. B.↗
NASA NTRS1973

Preliminary investigation of inlet ingestion of a wing tip vortex

An inlet-coldpipe assembly was placed in a Mach 0.4 stream to ingest the tip vortex of a forward mounted wing. The strongest vortex was produced by a wing angle of attack of 11 degrees. The vortex displayed a tangential velocity of 57 percent of local stream velocity prior to entering the inlet, and a tangential velocity of 25 percent of local velocity at the simulated compressor-face. The total-pressure profiles measured by standard compressor-face rakes were changed by the presence of the vortex only at the highest tested inlet mass-flow ratios.

Mitchell, G. A.↗
NASA NTRS1973

Analytic and experimental performance of two isentropic mixed compression axisymmetric inlets at Mach numbers 0.8 to 2.65

A mixed compression axisymmetric inlet model with a capture diameter of 50 cm was tested at Mach numbers ranging from 0.8 to 2.65 at 0 deg angle of attack and a constant total pressure of approximately 1 atm. Analytical methods accounting for the effects of both viscous and inviscid flows and incorporating empirical bleed discharge coefficients were used in the procedure for designing the inlet contours and the bleed system. Experimental results are compared with analytic predictions and are also compared with results from earlier tests of an inlet with the same internal contours but with a bleed system developed by cut and try methods in the wind tunnel. With the bleed configuration predicted by the design procedure, maximum total pressure recovery at the engine face at the design Mach number of 2.65 was 93 percent, with a total pressure distortion less than 10 percent. Corresponding bleed mass flow was approximately 7.5 percent, which was about 1.3 percent less than predicted. At lower supersonic Mach numbers, pressure recovery and bleed were generally lower and distortion generally higher.

Smeltzer, D. B.↗
NASA NTRS1973

ERTS-1 observations of sea surface circulation and sediment transport, Cook Inlet, Alaska

Cook Inlet is a large tide-dominated estuary in southern Alaska. Highly turbid streams enter the upper inlet, providing an excellent tracer for circulation in the lower inlet. MSS 4 and 5 images both can be used in this area to plot sediment and pollutant trajectories, areas of (probable) commercial fish concentration, and the entire circulation regime.

Wright, F. F.↗
NASA NTRS1973

Investigation of noise suppression by sonic inlets for turbofan engines. Volume 1: Program summary

Results of a program for sonic inlet technology development are presented. This program includes configuration and mechanical design selection of concepts, aerodynamic design description of the models, and results of test evaluation. Several sonic inlet concepts were tested and compared for aerodynamic and acoustic performance. Results of these comparative evaluations are presented. Near-field measurements were taken inside several of the inlet models. Results of these tests are discussed with respect to the effect of Mach number gradients on noise attenuation and rotor shock wave attenuation, and boundary layer effects on noise propagation. The test facilities and experimental techniques employed are described briefly.

Klujber, F.↗
NASA NTRS1973

Afterburner performance of film-vaporizing V-gutters for inlet temperatures up to 1255 K

Combustion tests of five variations of an integral, spray-bar - flameholder combination were conducted in a 0.49-m-diameter duct. Emphasis was on low levels of augmentation. Fuel impinged on guide plates, mixed with a controlled amount of inlet air, vaporized, and was guided into the V-gutter wake. Combustor length was 0.92 m. Good performance was demonstrated at fuel-air ratios less than 0.025 for inlet temperatures of 920 to 1255 K. Maximum combustion efficiency occured in the vicinity of fuel-air ratios of 0.02 and was 92 to 100 percent, depending on the inlet temperature. Lean blowout fuel-air ratios were in the vicinity of 0.005. Improvements in rich-limit blowout resulted from enlarging the guide-flow passageway areas. Other means of extending the operating range are suggested. A simplified afterburner concept for application to advanced engines is described.

Branstetter, J. R.↗
NASA NTRS1973

Experimental investigation of boundary layers in an axisymmetric, Mach 2.5, mixed-compression inlet

An experimental investigation has been made of the boundary layer flow in a supersonic axisymmetric inlet. The inlet was designed for Mach 2.5 and had a 60 percent internal supersonic area contraction. Porous bleed regions were incorporated into both the cowl and centerbody for boundary-layer control. Five bleed configurations were studied; these included bleed regions upstream of, across, and downstream of the experimentally determined shock-impingement points. The tests were conducted at tunnel unit Reynolds numbers of 8.2 million and 2.7 million per meter. Static-pressure measurements were made along the internal inlet surfaces. In addition to locating the position of the shocks, these measurements also showed local regions of expansion resulting from the suction bleed. Boundary-layer profiles were measured with translating pitot pressure probes located on both sides of the shock-impingement points on the cowl and centerbody. These measurements showed that the boundary-layer profiles were distorted in the vicinity of the terminal shock.

Hingst, W. R.↗
NASA NTRS1973

Calculation procedures for potential and viscous flow solutions for engine inlets

The method and basic elements of computer solutions for both potential flow and viscous flow calculations for engine inlets are described. The procedure is applicable to subsonic conventional (CTOL), short-haul (STOL), and vertical takeoff (VTOL) aircraft engine nacelles operating in a compressible viscous flow. The calculated results compare well with measured surface pressure distributions for a number of model inlets. The paper discusses the uses of the program in both the design and analysis of engine inlets, with several examples given for VTOL lift fans, acoustic splitters, and for STOL engine nacelles. Several test support applications are also given.

Albers, J. A.↗
NASA NTRS1972

Bleed system design technology for supersonic inlets.

A boundary layer bleed system design procedure for supersonic inlets, with emphasis on the selection of bleed hole geometry, is described. Available experimental bleed hole performance data, coupled with bleed drag calculations, show that holes with shallow inclination are superior to holes normal to the surface in terms of overall inlet performance. Recent test results from large-scale inlet models indicate that bleed hole size, bleed hole length, and boundary layer velocity profile upstream of the bleed region are important parameters in the design of an effective and efficient bleed system.

Syberg, J.↗
NASA NTRS1972

Inlet shear heating in elastohydrodynamic lubrication.

In elastohydrodynamic lubrication, the oil film thickness of rollers is controlled by the rate at which the oil is drawn into the conjunction of the disks by the moving surfaces of the rollers. The theory often assumes isothermal conditions in the inlet, although it can be shown that the maximum shear rate often exceeds 1,000,000 per sec, even in pure rolling. A theoretical analysis is presented for the oil temperature rise in the inlet of rollers, and the result is applied to predict the consequent film thickness. It is found that thermal effects on film thickness are only negligible at low rolling speeds. A comparison with experiment supports the conclusion that the thinning of the film thickness below that predicted by isothermal theory is substantially explained by inlet shear heating of the lubricant.

Greenwood, J. A.↗
NASA NTRS1973

Application of compressibility correction to calculation of flow in inlets.

An application of the compressibility correction developed by Lieblein and Stockman (1972) to the calculation of flow in axisymmetric inlets is described. The results with experimental data from wind tunnel model tests are compared. The configuration is a conventional subsonic inlet with a NACA series one external cowl shape and a two-to-one ellipse internal lip. It is shown that the compressibility correction gives a relatively good approximation to the internal compressible flow behavior, and thus should be useful in the design and analysis of engine nacelle inlets.

Albers, J. A.↗
NASA NTRS1973

Comparison of predicted and measured low-speed performance of two 51 centimeter-diameter inlets at incidence angle

Theoretical and experimental internal flow characteristics of two 51-cm-diameter inlets are compared. Theoretical flow characteristics along the inlet surface were obtained from an axisymmetric potential flow and boundary layer analysis. The experimental data were obtained from low-speed tests of a high-bypass-ratio turbofan engine simulator. Comparisons between calculated internal surface pressure distributions and experimental data are presented for a free-system velocity of 45 m/sec and for incidence angles from 0 deg to 50 deg. Analysis of boundary layer separation on the inlet lip at incidence angle is the major emphasis of this report. Theoretical boundary layer shape factors, skin friction coefficients, and velocity profiles in the boundary layer are presented, along with the location of the transition region. Theoretical and experimental separation locations are also discussed.

Albers, J. A.↗