An Analytical Shock-Loss Model Fully Integrated to a Streamline Curvature Method for Axial-Flow Compressors

Abstract

Abstract The constant and increasing demand to obtain more accurate turbomachinery performance prediction in the design and analysis process has led to the development of higher fidelity flow field models. Despite extensive flow field information can be collected from 3-D RANS numerical simulations, the computational cost is expensive in terms of time and resources, especially if they are used as solvers within a design-optimisation framework. In contrast, 2-D throughflow methods, such as streamline curvature (SLC), provide an acceptable flow solution in minutes. The use of modern and advanced-design transonic axial-flow compressors and fans has been expanding due to their high shock-induced single-stage pressure ratios while being light, compact and robust. Transonic-flow analysis in blading is complex due to the shock structures involved and associated phenomena. Previous 2-D SLC tools have failed to replicate the real compressible-flow physics, assuming and oversimplifying the shock-system shape and location. The situation aggravates, when the assumed overall shock configuration applies only for design point at unstarted operations, requiring of empirical correlations to estimate the shock-loss coefficient for off-design operations. The overall compressor performance prediction is thence highly-dependent on the shock modelling quality. For this reason, a physics-based shock -structure and -loss model was developed and implemented into an existing in-house 2-D SLC compressor performance simulator to enhance the aerodynamic prediction in transonic axial-flow compressors. The novel shock-loss model is fully coupled to the 2-D SLC software, for which a blade-element-layout method was adapted to obtain the profile geometry definition. The analytical shock-loss model possesses the capability to operate at started and unstarted passages utilizing an iterative-solution method to position the choke-induced passage-shock. A significant contribution of the new shock-loss model is the solution of the relative total-pressure loss for the entire blade span, comprising the inlet relative subsonic supercritical and supersonic regions. In this manner, shock losses were determined throughout the blade span and for various off-design operating conditions, including those at choking. 2-D SLC simulations were conducted for the NASA Rotor 67 Fan to validate the models accordingly against test-rig data and verify against previous model estimations and 3-D CFD results. The analytical shock - structure and -loss model improved the shock-loss prediction between 40–50% with respect of the state-of-the-art models and showed satisfactory agreement against measured data within 0.6% at the blade tip and 0.3% at mid-span sections.