The Experts below are selected from a list of 243 Experts worldwide ranked by ideXlab platform

Anthony J Strazisar - One of the best experts on this subject based on the ideXlab platform.

  • The Effect of Ultrapolish on a Transonic Axial Rotor
    Journal of Turbomachinery, 2012
    Co-Authors: William B. Roberts, Scott A. Thorp, Patricia S. Prahst, Anthony J Strazisar
    Abstract:

    Back-to-back testing was done using NASA fan rotor 67 in the Glenn Research Center W8 Axial Compressor Test Facility. The rotor was baseline tested with a normal industrial root-mean-square (RMS) surface finish of 0.5 μm to 0.6 μm (20 microinches to 24 microinches) at 60, 80, and 100% of design speed. At design speed the tip Relative Mach Number was 1.38. The blades were then removed from the facility and ultrapolished to a surface finish of 0.125 μm (5 microinch) or less and retested. At 100% speed near the design point, the ultrapolished blades showed approximately 0.3% to 0.5% increase in adiabatic efficiency. The difference was greater near maximum flow. Due to increased Relative measurement error at 60 and 80% speed, the performance difference between the normal and ultrapolished blades was indeterminate at these speeds.

Antonio Colella - One of the best experts on this subject based on the ideXlab platform.

  • Aerodesign and Performance Analysis of a Radial Transonic Impeller for a 9:1 Pressure Ratio Compressor
    Journal of Turbomachinery, 2008
    Co-Authors: Salvatore Colantuoni, Antonio Colella
    Abstract:

    Copyright © 1992 by ASME. The aerodynamic design of a centrifugal compressor for technologically advanced small aeroengines requires more and more the use of sophisticated computational tools in order to meet the goals successfully at minimum cost development. The objective of the present work is the description of the procedure adopted to design a transonic impeller having 1.31 Relative Mach Number at the inducer tip, 45° back-swept exit blade angle, and a tip speed of 636 m/s. The optimization of the blade shape has been done analyzing the aerodynamic flowfield by extensive use of a quasi-3d code and a fully 3D Euler solver based on a time-marching approach and a finite volume discretization. Testing has been done on the impeller-only configuration, using a compressor rig that simulates a real engine hardware, i.e. having an S-shape air-intake. The overall performance of the impeller are presented and discussed.

  • aerodesign and performance analysis of a radial transonic impeller for a 9 1 pressure ratio compressor
    Journal of Turbomachinery-transactions of The Asme, 1993
    Co-Authors: Salvatore Colantuoni, Antonio Colella
    Abstract:

    The aerodynamic design of a centrifugal compressor for technologically advanced small aeroengines requires more and more the use of sophisticated computational tools in order to meet the goals successfully at minimum cost development. The objective of the present work is the description of the procedure adopted to design a transonic impeller having 1.31 Relative Mach Number at the inducer tip, 45 deg back-swept exit blade angle, and a tip speed of 636 m/s. The optimization of the blade shape has been done by analyzing the aerodynamic flowfield by extensive use of a quasi-three-dimensional code and a fully three-dimensional Euler solver based on a time-marching approach and a finite volume discretization. Testing has been done on the impeller-only configuration, using a compressor rig that simulates real engine hardware, i.e. having an S-shaped air-intake. The overall performance of the impeller is presented and discussed.

Walter F. O'brien - One of the best experts on this subject based on the ideXlab platform.

  • An Improved Streamline Curvature Approach for Off-Design Analysis of Transonic Axial Compression Systems
    Volume 5: Turbo Expo 2002 Parts A and B, 2002
    Co-Authors: Keith M. Boyer, Walter F. O'brien
    Abstract:

    A streamline curvature throughflow numerical approach is assessed and modified to better approximate the flow fields of transonic axial compression systems. Improvements in total pressure loss modeling are implemented, central to which is a physics-based shock model, to ensure accurate and reliable off-design performance prediction. The new model accounts for shock geometry changes, with shock loss estimated as a function of inlet Relative Mach Number, blade section loading (flow turning), solidity, leading edge radius, and suction surface profile. Data from a single-stage, isolated rotor provide the basis for experimental comparisons. Improved performance prediction is shown. The importance of properly accounting for shock geometry and loss changes with operating conditions is demonstrated.© 2002 ASME

  • Application of an Improved Streamline Curvature Approach to a Modern, Two-Stage Transonic Fan: Comparison With Data and CFD
    Volume 5: Turbo Expo 2002 Parts A and B, 2002
    Co-Authors: Keith M. Boyer, Walter F. O'brien
    Abstract:

    A streamline curvature method with improvements to key loss models is applied to a two-stage, low aspect ratio, transonic fan with design tip Relative Mach Number of approximately 1.65. Central to the improvements is the incorporation of a physics-based shock model. The attempt here is to capture the effects of key flow phenomena Relative to the off-design performance of the fan. A quantitative analysis regarding solution sensitivities to model parameters that influence the key phenomena over a wide range of operating conditions is presented. Predictions are compared to performance determined from overall and interstage measurements, as well as from a three-dimensional, steady, Reynolds-averaged Navier-Stokes method applied across the first rotor. Overall and spanwise comparisons demonstrate that the improved model gives reasonable performance trending and generally accurate results. The method can be used to provide boundary conditions to higher-order solvers, or implemented within novel approaches using the streamline curvature method to explore complex engine-inlet integration issues, such as time-variant distortion.Copyright © 2002 by ASME

William B. Roberts - One of the best experts on this subject based on the ideXlab platform.

  • The Effect of Ultrapolish on a Transonic Axial Rotor
    Journal of Turbomachinery, 2012
    Co-Authors: William B. Roberts, Scott A. Thorp, Patricia S. Prahst, Anthony J Strazisar
    Abstract:

    Back-to-back testing was done using NASA fan rotor 67 in the Glenn Research Center W8 Axial Compressor Test Facility. The rotor was baseline tested with a normal industrial root-mean-square (RMS) surface finish of 0.5 μm to 0.6 μm (20 microinches to 24 microinches) at 60, 80, and 100% of design speed. At design speed the tip Relative Mach Number was 1.38. The blades were then removed from the facility and ultrapolished to a surface finish of 0.125 μm (5 microinch) or less and retested. At 100% speed near the design point, the ultrapolished blades showed approximately 0.3% to 0.5% increase in adiabatic efficiency. The difference was greater near maximum flow. Due to increased Relative measurement error at 60 and 80% speed, the performance difference between the normal and ultrapolished blades was indeterminate at these speeds.

Wang Hui - One of the best experts on this subject based on the ideXlab platform.

  • NUMERICAL SIMULATION OF SHOCK SYSTEMS IN CONVERGINGDIVERGING NOZZLE ROTOR TURBINE STAGE WITH DIFFERENT OUTLET STATIC PRESSURE
    Journal of Engineering Thermophysics, 2009
    Co-Authors: Wang Hui
    Abstract:

    Detailed numerical simulations have been carried out to investigate the shock systems in a converging-diverging nozzle rotor turbine stage.Near the design point,along the rotor span, the contours of the Relative Mach Number near the rotor trailing edge are different,and the strength of the fishtail shock and its echo wave on the suction surface of the adjacent rotor is also different. A group of weak pressure wave produces from the suction surface after about 60%axial chord of the rotor,where there has a concave wall.On the 87.5%span of the rotor,the weak pressure wave will develop to an oblique shock.The strength both of the fishtail shock and its echo wave varies with the increasing of the outlet static pressure of the turbine stage,generally which weakens, but on the 87.5%span of the rotor,the oblique shock developed by the weak pressure wave is intensified.

  • INFLUENCE OF ROTOR TIP CLEARANCE ON PERFORMANCE OF VANELESS COUNTER-ROTATING TURBINE
    Journal of Engineering Thermophysics, 2006
    Co-Authors: Wang Hui
    Abstract:

    A detailed unsteady numerical simulation has been carried out to investigate the influence of the rotor tip clearance on the performance of a 1+1/2 counter-rotating turbine (vaneless counter-rotating turbine (VCRT)). Through analyzing the total parameters of performance, the distribution of pitch-averaged outlet air angle, and pitch-averaged outlet Relative Mach Number and the loadings at different span etc. in VCRT, it can be found that, the tip clearance of the rotors can impose a clear influence on the VCRT's total performance, and the bigger of the clearance, the more of the influence. When the width of the tip clearance is equal to 4.5% of the rotor span, the flow in tip clearance is supersonic in the most area, which is needed to pay more attention in designing VCRT.