Axial compressor

About: Axial compressor is a research topic. Over the lifetime, 12035 publications have been published within this topic receiving 127766 citations.

An Improved Streamline Curvature Approach for Off-Design Analysis of Transonic Axial Compression Systems

Abstract: : A streamline curvature (SLC) throughflow numerical model was assessed and modified to better approximate the flow fields of highly transonic fans typical of military fighter applications. Specifically, improvements in total pressure loss modeling were implemented to ensure accurate and reliable off-design performance prediction. The assessment was made relative to the modeling of key transonic flow field phenomena, and provided the basis for improvements, central to which was the incorporation of a physics-based shock loss model. 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. Other improvements included incorporation of loading effects on the tip secondary loss model, use of radial blockage factors to model tip leakage effects, and an improved estimate of the blade section incidence at which minimum loss occurs. Data from a single-stage, isolated rotor and a two-stage, advanced-design (low aspect ratio, high solidity) fan provided the basis for experimental comparisons. The two-stage fan was the primary vehicle used to verify the present work. Results from a three-dimensional, steady, Reynolds-averaged Navier-Stokes model of the first rotor of the two-stage fan were also used to compare with predicted performance from the improved SLC representation. In general, the effects of important flow phenomena relative to off-design performance of the fan were adequately captured. These effects included shock loss, secondary flow, and spanwise mixing. Most notably, the importance of properly accounting for shock geometry and loss changes with operating conditions was clearly demonstrated. The majority of the increased total pressure loss with loading across the important first-stage tip region was shown to be the result of increased shock loss, even at part-speed.

Design of Industrial Axial Compressor Blade Sections for Optimal Range and Performance

Abstract: A design system for the blade sections of industrial axial compressors has been developed. The method combines a parametric geometry definition method, a powerful blade-to-blade flow solver (MISES) and an optimization technique (breeder genetic algorithm) with an appropriate fitness function. Particular effort has been devoted to the design of the fitness function for this application which includes non-dimensional terms related to the required performance at design and off-design operating points. It has been found that essential aspects of the design (such as the required flow turning, or mechanical constraints) should not be part of the fitness function, but need to be treated as so-called “killer” criteria in the genetic algorithm. Finally, it has been found worthwhile to examine the effect of the weighting factors of the fitness function to identify how these affect the performance of the sections. The system has been tested on the design of a repeating stage for the middle stages of an industrial axial compressor. The resulting profiles show an increased operating range compared to an earlier design using NACA65 profiles.Copyright © 2003 by ASME

Method and apparatus for predicting and stabilizing compressor stall

Abstract: A method and apparatus for predicting the onset of stall in an axial flow compressor and stabilizing stall so as to achieve the maximum pressure rise to the combustor. Stall in an axial flow compressor is predicted by obtaining pressure or axial velocity measurements at a plurality of circumferential angles around a compressor stage, fourier transforming the measurements with respect to the circumferential angles to obtain a time history of the zeroth, first and higher modes, and determining amplitude gain at twice the rotor frequency of the transfer function from the product of the zeroth and first modes to the first mode, with an increase in amplitude gain depicting a right skewed compressor going into stall and a decrease in amplitude gain depicting a left skewed compressor going into stall. Rotating stall and surge in the compressor is stabilized so that the pressure rise to the combustor can be maximized by creating an oscillatory motion of the pressure and feeding said oscillatory motion back to vary the set point of an extremum seeking controller, wherein an internal rotating stall and surge feedback loop in said controller that stabilizes said compressor is closed and wherein an outer extremum seeking loop in said controller which maximizes the pressure rise is closed.

Optimization of a stator blade using response surface method in a single-stage transonic axial compressor

Abstract: This article describes the shape optimization of a stator blade in a single-stage transonic axial compressor. The blade optimization has been performed using response surface method and three-dimensional Navier - Stokes analysis. Thin-layer approximation is introduced to the Navier - Stokes equations, and an explicit Runge - Kutta scheme is used to solve the gov- erning equations. Two geometric design variables of the stator blade, which are used to define a stacking line, are introduced to increase an adiabatic efficiency. D-optimal design is employed to reduce the number of evaluation points for response surface. With the optimization of the stator blade, the adiabatic efficiency is successfully improved when compared with that of the reference shape of the stator with straight stacking line. Positive stacking line, which bends on blade pressure side, effectively suppresses the flow separation on the blade suction surface of the stator.