SEARCH · Engineering Papers

Results for “INLET FLOW”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 199 records · Page 11

All sources NASA NTRS DOE OSTI

NASA NTRS2003

Enhanced Mixing in a Rectangular Duct

An experimental investigation of the mixing of non-reacting opposed rows of jets injected normal to a confined rectangular crossflow has been conducted. Planar Mie-scattering was used to measure the time-average concentration distribution of the jet fluid in planes perpendicular to the duct axis. Particular emphasis was placed on the study of closely spaced orifice configurations applicable to the mixing zone of an RQL combustor. Baseline studies were performed of mixing under "ideal" conditions, i.e., plenum fed jets injecting into a crossflow uniform in velocity and turbulence intensity. In addition, more practical ("non-ideal") issues encountered during hardware design were also studied. As in other studies, mixing effectiveness, determined using a spatial unmixedness parameter based on the variance of mean jet concentration distributions, was found to be optimum when the spacing-to-duct-height ratio was inversely proportional to the square root of the jet-to-mainstream momentum-flux ratio. This relationship is suitable for design under ideal flow conditions. Inlet flow boundary conditions of the jet and approach flow (mainstream) were found to strongly influence mixing performance, but no attempt was made to determine optimum performance under non-ideal conditions. The tests performed do offer some guidance as to expected mixing behavior for several common variables likely to be imposed by hardware constraints. Additionally, in this study it was found that for rows of orifices with opposite centerlines inline, mixing was similar for blockages up to 89 percent (previous crossflow mixing studies concerned with dilution zone configurations, blockages were typically less than 50 percent). Lower levels of unmixedness were obtained as a function of downstream location when axial injection length was minimized. Mixing may be enhanced if orifice centerlines of opposed rows are staggered, but blockage must be =50 percent in this configuration. Round hole and "square" orifice shapes had similar performance. Other variations in orifice shape did not substantially augment overall mixing performance. Furthermore an isothermal mixing data set was generated and used by CFDRC as input to a NO(x) inference code.

Liscinsky, D. S.↗

NASA NTRS1967

Analytical studies of aspect ratio and curvature variations for axial flow compressor inlet stages under high loading

Analyses of aspect ratio and curvature variations for axial flow compressor inlet stages under loading

AXIAL FLOW COMPRESSOR↗

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 NTRS1988

A preliminary design study of supersonic through-flow fan inlets

From Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic through-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drag and size and weight.

Barnhart, Paul J.↗

NASA NTRS1988

A preliminary design study of supersonic through-flow fan inlets

From Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic through-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one-dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drag and size and weight.

Barnhart, Paul J.↗

NASA NTRS1986

Calculation of external-internal flow fields for mixed-compression inlets

Supersonic inlet flows with mixed external-internal compressions were computed using a combined implicit-explicit (Beam-Warming-Steger/MacCormack) method for solving the three-dimensional unsteady, compressible Navier-Stokes equations in conservation form. Numerical calculations were made of various flows related to such inlet operations as the shock-wave intersections, subsonic spillage around the cowl lip, and inlet started versus unstarted conditions. Some of the computed results were compared with wind tunnel data.

Chyu, W. J.↗

NASA NTRS1987

Calculation of external-internal flow fields for mixed-compression inlets

Chyu, W. J.↗

NASA NTRS1989

The automation of an inlet mass flow control system

The automation of a closed-loop computer controlled system for the inlet mass flow system (IMFS) developed for a wind tunnel facility at Langley Research Center is presented. This new PC based control system is intended to replace the manual control system presently in use in order to fully automate the plug positioning of the IMFS during wind tunnel testing. Provision is also made for communication between the PC and a host-computer in order to allow total animation of the plug positioning and data acquisition during the complete sequence of predetermined plug locations. As extensive running time is programmed for the IMFS, this new automated system will save both manpower and tunnel running time.

Supplee, Frank↗

NASA NTRS1955

Preliminary Results of the Determination of Inlet-Pressure Distortion Effects on Compressor Stall and Altitude Operating Limits of the J57-P-1 Turbojet Engine

During an investigation of the J57-P-1 turbojet engine in the Lewis altitude wind tunnel, effects of inlet-flow distortion on engine stall characteristics and operating limits were determined. In addition to a uniform inlet-flow profile, the inlet-pressure distortions imposed included two radial, two circumferential, and one combined radial-circumferential profile. Data were obtained over a range of compressor speeds at an altitude of 50,000 and a flight Mach number of 0.8; in addition, the high- and low-speed engine operating limits were investigated up to the maximum operable altitude. The effect of changing the compressor bleed position on the stall and operating limits was determined for one of the inlet distortions. The circumferential distortions lowered the compressor stall pressure ratios; this resulted in less fuel-flow margin between steady-state operation and compressor stall. Consequently, the altitude operating Limits with circumferential distortions were reduced compared with the uniform inlet profile. Radial inlet-pressure distortions increased the pressure ratio required for compressor stall over that obtained with uniform inlet flow; this resulted in higher altitude operating limits. Likewise, the stall-limit fuel flows required with the radial inlet-pressure distortions were considerably higher than those obtained with the uniform inlet-pressure profile. A combined radial-circumferential inlet distortion had effects on the engine similar to the circumferential distortion. Bleeding air between the two compressors eliminated the low-speed stall limit and thus permitted higher altitude operation than was possible without compressor bleed.

Wallner, L. E.↗

NASA NTRS1985

Numerical simulation of self-excited oscillations in a ramjet inlet-diffuser flow

This paper describes numerical simulations of self-excited oscillations in a two-dimensional transonic inlet-diffuser flow by solving the Navier-Stokes equations with a two-equation turbulence model. The calculated amplitudes of oscillations for the terminal shock and the velocity fields compare well with experimental measurements; however, the predicted frequency of oscillations is about 50 percent higher. The formation of a pair of downstream-traveling, counter-rotating vortices at each cycle of velocity fluctuations, as reported experimentally, is vividly revealed by the numerical results.

Hsieh, T.↗

NASA NTRS2000

Investigation of Flow Separation in a Transonic-fan Linear Cascade Using Visualization Methods

An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on two visualization techniques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to simulate these flows numerically.

Lepicovsky, Jan↗

NASA NTRS1975

Two-stage, low noise advanced technology fan. 4: Aerodynamic final report

A two-stage research fan was tested to provide technology for designing a turbofan engine for an advanced, long range commercial transport having a cruise Mach number of 0.85 -0.9 and a noise level 20 EPNdB below current requirements. The fan design tip speed was 365.8m/sec (1200ft/sec);the hub/tip ratio was 0.4; the design pressure ratio was 1.9; and the design specific flow was 209.2 kg/sec/sq m(42.85lbm/sec/sq ft). Two fan-versions were tested: a baseline configuration, and an acoustically treated configuration with a sonic inlet device. The baseline version was tested with uniform inlet flow and with tip-radial and hub-radial inlet flow distortions. The baseline fan with uniform inlet flow attained an efficiency of 86.4% at design speed, but the stall margin was low. Tip-radial distortion increased stall margin 4 percentage points at design speed and reduced peak efficiency one percentage point. Hub-radial distortion decreased stall margin 4 percentage points at all speeds and reduced peak efficiency at design speed 8 percentage points. At design speed, the sonic inlet in the cruise position reduced stall margin one percentage point and efficiency 1.5 to 4.5 percentage points. The sonic inlet in the approach position reduced stall margin 2 percentage points.

Harley, K. G.↗

NASA NTRS1979

Combined pressure and temperature distortion effects on internal flow of a turbofan engine

The flow characteristics obtained experimentally for the compression of a 2-spool TF30-P-3 turbofan engine operating with 180 grad combined pressure and temperature distortion in the inlet flow are presented. The analytical model (Mazzawy and Banks, 1976), 'tuned' during Lewis testing, was used for pretest predictions of the effects that these distortions would have on the engine flow characteristics and the limiting distortion values. The effect of inlet flow distortion on the performance of the engine is discussed, including: (1) the flow between a screen mounted in the inlet duct and the inlet guide vanes; (2) the flow through the compression system; and (3) the effects of the combined distortion and its orientation on the compressor stability limits. It is concluded that the model used in this program was capable of predicting the effects of total pressure, total temperature and combined total pressure-total temperature distortions in terms of flow profiles, inlet flow angles and attenuation of the distortions through the compressor system. It was also capable of predicting the trends of the limiting values experienced with various orientations of the combined distortions.

Braithwaite, W. M.↗

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 NTRS1972

Analysis of the flow field generated near an aircraft engine operating in reverse thrust

A computer solution is developed to the exhaust gas reingestion problem for aircraft operating in the reverse thrust mode on a crosswind-free runway. The computer program determines the location of the inlet flow pattern, whether the exhaust efflux lies within the inlet flow pattern or not, and if so, the approximate time before the reversed flow reaches the engine inlet. The program is written so that the user is free to select discrete runway speeds or to study the entire aircraft deceleration process for both the far field and cross-ingestion problems. While developed with STOL applications in mind, the solution is equally applicable to conventional designs. The inlet and reversed jet flow fields involved in the problem are assumed to be noninteracting. The nacelle model used in determining the inlet flow field is generated using an iterative solution to the Neuman problem from potential flow theory while the reversed jet flow field is adapted using an empirical correlation from the literature. Sample results obtained using the program are included.

Ledwith, W. A., Jr.↗

NASA NTRS1974

Potential and viscous flow prediction in V/STOL propulsion system inlets

Highlights of the method of analysis of inlet flow are given. To indicate the accuracy of the method, several comparisons with experiments for different V/STOL inlet configurations and various operating conditions are given. Applications to inlet design and analysis are then discussed. A summary of current efforts is given, and areas of possible future work are indicated.

Stockman, N. O.↗

NASA NTRS1975

Potential and viscous flow prediction in V/STOL propulsion system inlets

Highlights of the method of analysis of inlet flow are given. To indicate the accuracy of the method, several comparisons with experiment for different V/STOL inlet configurations and various operating conditions are given. Two applications to inlet design and analysis are then discussed. A summary of current efforts is given, and finally areas of possible future work are indicated.

Stockman, N. O.↗

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.↗