The current work explains the development and the validation of a low-fidelity numerical tool for the fast simulation of axial compressors under clean and distorted inlet flow conditions. The focus is to get a preliminary assessment of the surge margin loss due to the presence of a total pressure non-uniformity at the Aerodynamic Interface Plane (AIP) section. The code here presented, written in C++, is divided in two parts: a meanline solver for the calculation of the compressor performance under clean inlet conditions, and an inlet flow distortion module for the evaluation of the stability margin reduction. Experimental data and CFD numerical results allowed the tuning of the empirical models and a complete validation of the code.  
The 1D subsonic/transonic meanline solver calculates the thermodynamics and the performance of the  
machine, operating under uniform inlet flow conditions, for different spanwise positions. The solver has been validated considering the experimental datasets of four test cases, both in stage and isolated-rotor configurations: HP9 subsonic stage, NASA S37 transonic stage, NASA R37 transonic rotor and NASA R67 transonic fan. The results have been calculated as performance maps and spanwise distributions.  
Two different calculation strategies allow the evaluation of the impact of inlet flow distortion, in terms of a reduction of the surge margin. The first method implemented follows the ARP1420 Gas Turbine Engine Inlet Flow Distortion Guidelines, which give both a parametrization of the total pressure field at the AIP and an evaluation of the stability margin variation through sets of empirical correlations. The second method is to run a parallel compressor routine, which allows a more accurate prediction of the surge margin loss relying on the meanline solver calculation. The tuning of the outlet static pressure allows to calibrate the parallel compressor model for different distortion patterns. The validation of this module of the code has been conducted considering the numerical results of a full annulus CFD simulation of NASA R37 test case.  
The compressible solver of ANSYS CFX has been adopted. A full annulus 36M elements structured mesh has been validated against experimental data, for the available speed lines. After that, a distortion pattern has been imposed as inlet boundary condition. A 180° circumferential pressure distortion has been selected. The CFD results of the surge line have been compared with the numerical output of the code. The validity of the proposed low-fidelity methodology for the assessment of surge margin loss has been confirmed.


Meanline Calculation of Surge Margin Loss due to Inlet Flow Distortion

Division / Sponsor:
International Gas Turbine Institute (IGTI)

Paper Proceedings: https://asmedigitalcollection.asme.org/GT/GT2020/volume/84058

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