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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 793 records · Page 44
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Core compressor exit stage study. Volume 2: Data and performance report for the baseline configuration

The objective of the program is to develop rear stage blading designs that have lower losses in their endwall boundary layer regions. The overall technical approach in this efficiency improvement program utilized General Electric's Low Speed Research Compressor as the principal investigative tool. Tests were conducted in two ways: using four identical stages of blading so that test data would be obtained in a true multistage environment and using a single stage of blading so that comparison with the multistage test results could be made.

Wisler, D. C.↗
NASA NTRS1980

Program to develop sprayed, plastically deformable compressor shroud seal materials

A study of fundamental rub behavior for 10 dense, sprayed materials and eight current compressor clearance materials was conducted. A literature survey of a wide variety of metallurgical and thermophysical properties was conducted and correlated to rub behavior. Based on the results, the most promising dense rub material was Cu-9A1. Additional studies on the effects of porosity, incursion rate, blade solidity, and ambient temperature were carried out on aluminum bronze (Cu-9Al-1Fe) with and without a 515B Feltmetal underlayer. A further development effort was conducted to assess the property requirements of a porous, aluminum bronze, seal material. Strength, thermal cycle capabilities, erosion and oxidation resistance, machinability, and abradability at several porosity levels were examined.

Schell, J. D.↗
NASA NTRS1981

The boundary layer on compressor cascade blades

The flow field about an airfoil in cascade at a Reynolds number of 5 x 10 to the 5th power is described. Hot wire and laser anemometry are combined with flow visualization techniques in order to obtain detailed flow data (e.g., boundary layer profiles, points of separation, and the transition zone) on a cascade of relatively highly loaded blades. Benchmark data is provided for the evaluation of current and future predictive models, in this way aiding in the compressor design process.

Deutsch, S.↗
NASA NTRS1981

Semi-actuator disk theory for compressor choke flutter

A mathematical anaysis predict the unsteady aerodynamic utilizing semi actuator theory environment for a cascade of airfoils harmonically oscillating in choked flow was developed. A normal shock is located in the blade passage, its position depending on the time dependent geometry, and pressure perturbations of the system. In addition to shock dynamics, the model includes the effect of compressibility, interblade phase lag, and an unsteady flow field upstream and downstream of the cascade. Calculated unsteady aerodynamics were compared with isolated airfoil wind tunnel data, and choke flutter onset boundaries were compared with data from testing of an F100 high pressure compressor stage.

Micklow, J.↗
NASA NTRS1979

Abradable compressor and turbine seals, volume 2

The applications and advantages of abradable coatings as gas path seals in a general aviation turbofan engine were investigated. Abradable materials were evaluated for the high pressure radial compressor and the axial high and low pressure turbine shrouds.

Sundberg, D. V.↗
NASA NTRS1981

Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 1: Volute-manifold and stator performance

The aerodynamic performance of the inlet manifold and stator assembly of the compressor drive turbine was experimentally determined with cold air as the working fluid. The investigation included measurements of mass flow and stator-exit fluid torque as well as radial surveys of total pressure and flow angle at the stator inlet and annulus surveys of total pressure and flow angle at the stator exit. The stator-exit aftermixed flow conditions and overall stator efficiency were obtained and compared with their design values and the experimental results from three other stators. In addition, an analysis was made to determine the constituent aerodynamic losses that made up the stator kinetic energy loss.

Roelke, R. J.↗
NASA NTRS1981

End wall flows in rotors and stators of a single stage compressor

A solution of the flow in the rotor end wall region, including the effects of tip clearance flow, is presented. A method for leakage flow measurement at the tip of a compressor rotor blade is discussed. Measurements are given for a rotor hub wall boundary layer.

Govindam, T. R.↗
NASA NTRS1980

Structure and decay characteristics of turbulence in the near- and far-wake of a moderately loaded compressor rotor-blade

The wake of a turbomachinery rotor-blade is turbulent, highly three-dimensional, and nonisotropic with appreciable curvature in the trailing-edge and near-wake regions. The characteristics of the turbulence vary considerably with radius, blade loading, free-stream turbulence, Reynolds number, and the rotor-blade geometry. This paper is concerned with the turbulence properties of a moderately loaded compressor blade, particularly near the blade trailing-edge. The tangential variation of the axial, tangential and radial intensities and stresses across the wake, as well as their decay characteristics were measured with a tri-axial hot-wire probe in the rotor frame of reference. The decay of intensities and stresses were found to be very rapid in the trailing-edge and near-wake regions and slow in the far-wake region. The effects of inlet-guide-vane and the hub-wall boundary layers on the rotor wake turbulence spectra are also discussed. Similarity rules for the three components of intensity are also derived and presented in this paper.

Ravindranath, A.↗
NASA NTRS1981

Effects of solid particles suspended in fluid flow through an axial flow compressor stage

An approximate method for calculating the flow properties of gas-particle mixture flowing over blades in a cascade is studied. Using an analytical method, the solid particle trajectory and location of collisions between the solid particles and the blade surfaces are determined. In addition an experimental investigation of the trajectories and velocities of solid particles suspended in a fluid passing through an axial flow compressor cascade was performed. The cascade blades were made of 2024 aluminum alloy and the solid particles used were quartz sand with average diameter of 165 microns. The following parameters were investigated: blade pressure distribution, and total pressure loss coefficient, both depending on the degree of erosion. In addition, a theoretical estimation of the blade erosion is compared with the experimental data.

Tabakoff, W.↗
NASA NTRS1981

Rotor wake mixing effects downstream of a compressor rotor

An experimental study of rotor wake was conducted in the trailing-edge and near-wake regions of a moderately loaded compressor rotor blade using a rotating triaxial hot-wire probe in a rotating frame of reference. The flow-field was surveyed very close to the trailing-edge as well as inside the annulus- and hub-wall boundary layers. The large amount of data acquired during this program has been analyzed to discern the decay effects as well as the spanwise variation of three components of velocity, three components of intensities and three components of shear stresses. The data set also include extensive information on the variation of the flow properties downstream. The other derived quantities include wake momentum thickness and deviation angles at various spanwise and downstream locations. These data are presented and interpreted, with emphasis on the downstream mixing as well as endwall-wake interaction effects.

Ravindranath, A.↗
NASA NTRS1981

Turbulence characteristics in the annulus-wall boundary layer and wake mixing region of a compressor rotor exit

The flow at the exit of a turbomachinery rotor near the annulus wall is three-dimensional with a varying degree of turbulence and nonisotropic characteristics. The characteristics of the turbulence vary considerably with radial and axial location, blade loading, Reynolds number, free-stream turbulence, and the rotor blade geometry. This paper is concerned with the turbulence properties near the annulus wall region of a moderately loaded compressor blade. The variation of the streamwise, normal, and radial intensities as well as the stresses across the wake, its profile and decay characteristics, were determined with a tri-axial hot-wire probe in the stationary frame of reference. The turbulence characteristics in the annulus-wall region of the rotor exit were found to be highly complex due to the interaction of the blade wake and the annulus wall.

Davino, R.↗
NASA NTRS1981

Core compressor exit stage study. Volume 5: Design and performance report for the Rotor C/Stator B configuration

Rear stage blading designs that have lower losses in their endwall boundary layer regions were developed. The design of rotor-C and the performance results for rotor-C running with stator B are described. A low speed research compressor is utilized as the principal investigative tool. Four identical stages of blading are used to obtained data in a true multistage environment.

Wisler, D. C.↗
NASA NTRS1982

The use of optimization techniques to design controlled diffusion compressor blading

A method for automating compressor blade design using numerical optimization, and applied to the design of a controlled diffusion stator blade row is presented. A general purpose optimization procedure is employed, based on conjugate directions for locally unconstrained problems and on feasible directions for locally constrained problems. Coupled to the optimizer is an analysis package consisting of three analysis programs which calculate blade geometry, inviscid flow, and blade surface boundary layers. The optimizing concepts and selection of design objective and constraints are described. The procedure for automating the design of a two dimensional blade section is discussed, and design results are presented.

Sanger, N. L.↗
NASA NTRS1982

Comparison of two and three dimensional flow computations with laser anemometer measurements in a transonic compressor rotor

A procedure for using an efficient axisymmetric code to generate downstream pressure input for more costly Euler codes is discussed. Two and three dimensional inviscid solutions for the flow within a transonic axial compressor rotor at design speed are compared to laser anemometer measurements at maximum flow and near stall operating points. Computational details of the 2-D axisymmetric stream function solution and the 3-D full Euler solution are described. Relative Mach number contours, shock location, and shock strength as measured and as predicted by the 3-D code are compared. Downstream of the rotor the inviscid computations agree with each other but predict higher pressure ratios than those measured. Euler codes require a downstream pressure as input. Since that pressure controls the computed mass flow and shock system, it must be consistent with an inviscid solution.

Chima, R. V.↗
NASA NTRS1981

Computer program for aerodynamic and blading design of multistage axial-flow compressors

A code for computing the aerodynamic design of a multistage axial-flow compressor and, if desired, the associated blading geometry input for internal flow analysis codes is presented. Compressible flow, which is assumed to be steady and axisymmetric, is the basis for a two-dimensional solution in the meridional plane with viscous effects modeled by pressure loss coefficients and boundary layer blockage. The radial equation of motion and the continuity equation are solved with the streamline curvature method on calculation stations outside the blade rows. The annulus profile, mass flow, pressure ratio, and rotative speed are input. A number of other input parameters specify and control the blade row aerodynamics and geometry. In particular, blade element centerlines and thicknesses can be specified with fourth degree polynomials for two segments. The output includes a detailed aerodynamic solution and, if desired, blading coordinates that can be used for internal flow analysis codes.

Crouse, J. E.↗
NASA NTRS1982

Three dimensional flow field inside compressor rotor, including blade boundary layers

The flow in a turbomachinery blade passage has a predominant flow direction. The viscous diffusion in the streamwise direction is usually small and the elliptic influence is transmitted upstream through the pressure field. Starting with a guessed pressure field, it is possible to converge on the full elliptic solution by iterating between a parabolic solution and an iteration of the pressure field. The main steps of the calculation are given. The blade boundary layers which are three dimensional with laminar, transitional, turbulent, and separation zones are investigated. The kinetic energy is analyzed, and the dissipation equation is presented. Measurements were made of the three dimensional flow inside an axial flow compressor passage.

Pouagare, M.↗
NASA NTRS1982

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.63 and 1.78, respectively, and with design pressure ratio of 1.82

The overall and blade-element performance of a transonic compressor stage is presented over the stable operating flow range for speeds from 50 to 100 percent of design. The stage was designed for a pressure ratio of 1.82 at a flow 20.2 kg/sec and a tip speed of 455 m/sec. At design speed the stage achieved a peak efficiency of 0.821 at a pressure ratio of 1.817. The stage stall margin at design speed based on conditions at stall and peak efficiency was about 11 percent.

Moore, R. D.↗
NASA NTRS1982

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.63 and 1.77, respectively, and with design pressure ratio of 2.05

The overall and blade-element performance of a transonic compressor stage is presented over the stable operating range for speeds from 50 to 100 percent of design. The stage was designed for a pressure ratio of 2.05 at a flow of 20.2 kg/sec and a tip speed of 455 m/sec. At design speed the rotor and stage achieved peak efficiencies of 0.849 and 0.831, respectively, at the minimum flow condition. The stage stall point occurred at a flow higher than the design flow.

Moore, R. D.↗