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At least 469 records · Page 26

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NASA NTRS1986

Performance studies on an axial flow compressor stage

A low-speed, medium loaded axial flow compressor stage is studied experimentally and theoretically. The flow compressor facility, composed of an inlet guide vane row, a rotor blade row, and a stator blade row, and the principles of the streamline curvature method (SCM) and the Douglas-Neumann cascade program are described. The radial distribution of the flow properties, the rotor blade static pressure distribution, and the lift coefficient and relative flow angle derived experimentally and theoretically are compared. It is determined that there is good correlation between the experimental flow properties and the SCM data, the Douglas-Neumann cascade program and experimental rotor blade static pressure data, and the experimental and theoretical lift coefficients only in the midspan region. Modifications to the SCM and the Douglas-Neumann cascade program in order to improve their accuracy are discussed.

Sitaram, N.↗

NASA NTRS1987

Enhancing compressor distortion tolerance by asymmetric stator control

This paper describes a preliminary study of a method for increasing compressor tolerance in circumferential distortion through asymmetric vane stagger control. The changes in vane stagger about the circumference are used to increase compressor pressure rise and flow capability over part of the annulus and decrease these over the rest. The average pressure rise and flow thus remain roughly constant but the velocity distortion is reduced (and can even be made zero). Calculations are carried out to show the magnitude of the vane motions that must be used, and it is found that significant increases in distoriton attenuation can be achieved with realistic vane swings. In addition, the asymmetric vane stagger scheme leads to a substantial decrease in stall pressure rise loss due to inlet total pressure distortion. This type of control strategy appears worthy of further investigation, both theoretical and experimental.

Chen, G. T.↗

NASA NTRS1988

Experimental Vibration Damping Characteristics of the Third-stage Rotor of a Three-stage Transonic Axial-flow Compressor

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynamic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall preformance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.

Newman, Frederick A.↗

NASA NTRS1988

Experimental Determination of Aerodynamic Damping in a Three-Stage Transonic Axial-Flow Compressor

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gauge output power spectra. The combined damping consists of aerodynamic and structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given equivalent speed, equivalent mass flow, and pressure ratio while structural and mechanical damping are assumed to be constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third stage rotor blade aerodynamic damping is presented and discussed for 70, 80, 90, and 100 percent design equivalent speed. The compressor overall performance and experimental Campbell diagrams for the third stage rotor blade row are also presented.

Newman, Frederick A.↗

NASA NTRS1987

Real-time data compressor for Eos-class missions

A conceptual design for a real-time VLSI compressor capable of processing rate up to one gigabit per second is presented. This scheme is capable of providing a three-to-one distortion-free data reduction factor to both the High Resolution Imaging Spectrometer and processed SAR imaging data. The design uses a VLSI parallel/piplined architecture capable of processing at a real time rate. The design consists of a parallel array of VLSI compressor modules. Each module is built on a single customized VLSI chip using existing state-of-the-art semiconductor technology.

Lee, Jun-Ji↗

NASA NTRS1988

Experimental vibration damping characteristics of the third-stage rotor of a three-stage transonic axial-flow compressor

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynanmic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall performance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.

Newman, Frederick A.↗

NASA NTRS1988

Design point variation of 3-D loss and deviation for axial compressor middle stages

The available data on middle-stage research compressors operating near design point are used to derive simple empirical models for the spanwise variation of three-dimensional viscous loss coefficients for middle-stage axial compressor blading. The models make it possible to quickly estimate the total loss and deviation across the blade span when the three-dimensional distribution is superimposed on the two-dimensional variation calculated for each blade element. It is noted that extrapolated estimates should be used with caution since the correlations have been derived from a limited data base.

Roberts, William B.↗

NASA NTRS1988

Effect of stage loading on endwall flows in an axial flow compressor rotor

This paper reports results from investigations conducted to determine the effect of stage loading on endwall flows in a low speed axial flow compressor. These investigations consisted of two sets of measurements. The first set consisted of radial transverse of flow properties at the rotor inlet and exit, at five flow coefficients. These measurements are used to determine the boundary layer integral parameters. The displacement thicknesses at the rotor hub and tip agree reasonably well with Smith's (1970) correlation for multistage axial compressors. The second set consisted of measurements of static pressures on the rotor blade at four flow coefficients. From these measurements, lift coefficient is determined. Also, loss of lift coefficient near the tip is calculated, and is attributed mainly to the tip leakage flows.

Sitaram, N.↗

NASA NTRS1988

Experimental investigation of the performance of a supersonic compressor cascade

Supersonic cascade wind tunnel results are presented for a linear, supersonic compressor cascade derived from the near-tip section of a high-throughflow axial flow compressor rotor over the inlet Mach number range of 1.30-1.71. Laser anemometry was used to obtain flow-velocity measurements showing the wave pattern in the entrance region. Attention is given to the unique-incidence relationship for this cascade, which relates the supersonic inlet Mach number to the inlet flow direction. An empirical correlation is obtained for the influence of the independent parameters of back pressure, axial velocity density ratio, and blade element performance.

Tweedt, T. L.↗

NASA NTRS1988

Experimental investigation of the performance of a supersonic compressor cascade

Results are presented from an experimental investigation of a linear, supersonic, compressor cascade tested in the supersonic cascade wind tunnel facility at the DFVLR in Cologne, Federal Republic of Germany. The cascade design was derived from the near-tip section of a high-through-flow axial flow compressor rotor with a design relative inlet Mach number of 1.61. Test data were obtained over a range of inlet Mach numbers from 1.23 to 1.71, and a range of static pressure ratios and axial-velocity-density ratios (AVDR) at the design inlet condition. Flow velocity measurements showing the wave pattern in the cascade entrance region were obtained using a laser transit anemometer. From these measurements, some unique-incidence conditions were determined, thus relating the supersonic inlet Mach number to the inlet flow direction. The influence of static pressure ratio and AVDR on the blade passage flow and the blade-element performance is described, and an empirical correlation is used to show the influence of these two (independent) parameters on the exit flow angle and total-pressure loss for the design inlet condition.

Tweedt, D. L.↗

NASA NTRS1989

A prediction of 3-D viscous flow and performance of the NASA low-speed centrifugal compressor

A prediction of the 3-D turbulent flow in the NASA Low-Speed Centrifugal Compressor Impeller has been made. The calculation was made for the compressor design conditions with the specified uniform tip clearance gap. The predicted performance is significantly worse than that predicted in the NASA design study. This is explained by the high tip leakage flow in the present calculation and by the different model adopted for tip leakage flow mixing. The calculation gives an accumulation for high losses in the shroud/pressure-side quadrant near the exit of the impeller. It also predicts a region of meridional backflow near the shroud wall. Both of these flow features should be extensive enough in the NASA impeller to allow detailed flow measurements, leading to improved flow modelling. Recommendations are made for future flow studies in the NASA impeller.

Moore, John↗

NASA NTRS1988

The design and development of transonic multistage compressors

The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.

Ball, C. L.↗

NASA NTRS1989

High stability design for new centrifugal compressor

It is essential that high-performance centrifugal compressors be free of subsynchronous vibrations. A new high-performance centrifugal compressor has been developed by applying the latest rotordynamics knowledge and design techniques: (1) To improve the system damping, a specially designed oil film seal was developed. This seal attained a damping ratio three times that of the conventional design. The oil film seal contains a special damper ring in the seal cartridge. (2) To reduce the destabilizing effect of the labyrinth seal, a special swirl canceler (anti-swirl nozzle) was applied to the balance piston seal. (3) To confirm the system damping margin, the dynamic simulation rotor model test and the full load test applied the vibration exciting test in actual load conditions.

Kanki, H.↗

NASA NTRS1989

Two-dimensional computations of multi-stage compressor flows using a zonal approach

A clear understanding of the fluid dynamics associated with rotor/stator configurations can be very helpful when optimizing the performance of turbomachinery. In this study, a two-dimensional, implicit, thin-layer, Navier-Stokes zonal approach has been used to investigate the flow within a 2 1/2-stage compressor. Relative motion between the rotor and stator airfoils is made possible with the use of systems of patched and overlaid grids that move with respect to each other. The treatment of multistage turbomachines with arbitrary numbers of airfoils per row is made possible by the use of a flexible database system. Results in the form of instantaneous pressure and entropy contours and time-averaged pressures are presented for the 2 1/2-stage compressor. Time-averaged pressures and pressure amplitudes for a single-stage turbine configuration are also presented. The numerical results compare well with experimental data.

Gundy-Burlet, Karen L.↗

NASA NTRS1989

L2F and LDV velocimetry measurement and analysis of the 3-D flow field in a centrifugal compressor

The flow field in the Purdue Research Centrifugal Compressor is studied using a laser two-focus (L2F) velocimeter. L2F data are obtained which quantify: (1) the compressor inlet flow field; (2) the steady-state velocity field in the impeller blade passages; and (3) the flow field in the radial diffuser. The L2F data are compared with both laser Doppler velocimetry (LDV) data and predictions from three-dimensional inviscid and viscous flow models. In addition, a model is developed to calculate the effect on the measurement volume geometry of refraction by curved windows. Finally, the advantages and disadvantages of using the L2F for turbomachinery measurements is discussed in terms of measurement accuracy, ease of use, including sample time per correlated event and the ability to make measurements in regions of high noise due to stray radiation from wall reflections.

Fagan, John R., Jr.↗

NASA NTRS1990

Temporally and spatially resolved flow in a two-stage axial compressor. Part 2: Computational assessment

Fluid dynamics of turbomachines are complicated due to aerodynamic interactions between rotors and stators. It is necessary to understand the aerodynamics associated with these interactions in order to design turbomachines that are both light and compact as well as reliable and efficient. The current study uses an unsteady, thin-layer Navier-Stokes zonal approach to investigate the unsteady aerodynamics of a multi-stage compressor. Relative motion between rotors and stators is made possible by use of systems of patched and overlaid grids. Results have been computed for a 2 1/2-stage compressor configuration. The numerical data compares well with experimental data for surface pressures and wake data. In addition, the effect of grid refinement on the solution is studied.

Gundy-Burlet, Karen L.↗

NASA NTRS1990

WINCOF-I code for prediction of fan compressor unit with water ingestion

The PURDUE-WINCOF code, which provides a numerical method of obtaining the performance of a fan-compressor unit of a jet engine with water ingestion into the inlet, was modified to take into account: (1) the scoop factor, (2) the time required for the setting-in of a quasi-steady distribution of water, and (3) the heat and mass transfer processes over the time calculated under 2. The modified code, named WINCOF-I was utilized to obtain the performance of a fan-compressor unit of a generic jet engine. The results illustrate the manner in which quasi-equilibrium conditions become established in the machine and the redistribution of ingested water in various stages in the form of a film out of the casing wall, droplets across the span, and vapor due to mass transfer.

Murthy, S. N. B.↗

NASA NTRS1991

Concentric Regenerative Sorption Compressor

Proposed sorption compressor for cryogenic system features four sorption units. Concentric sorption units alternately connected and disconnected via timed gas-gap thermal switches. During given half cycle, one of C/Kr sorption units operates on waste heat from one of PCO sorption units, which in turn operates partially on waste heat from other PCO sorption unit. Required power equals that of mechanical compressor. Vibrates less, and more reliable because it contains only few moving parts actuated relatively infrequently.

Jones, Jack A.↗