During on-wing time of a jet engine deterioration occurs and leads to decreasing component efficiencies. This results in increasing Exhaust Gas Temperature (EGT) and Specific Fuel Consumption (SFC). Thereby, the condition of the High Pressure Compressor (HPC) has a comparatively large influence on these parameters defining overall engine performance. This can be explained by a changing flow field in the HPC due to geometric deviations which may occur during operation. The geometries are influenced by erosion which results in thinner airfoils, changed leading- and trailing edge geometries, shortened airfoils and increasing tip clearance. The objective for future maintenance strategies is to determine the influence of the different wear mechanisms. This can be done by experiments and Computational Fluid Dynamics (CFD). Because of the high number of different mechanisms and locations of deterioration, CFD-calculations seem to be necessary to give a detailed view of the influence of deterioration. In this regard, the following paper will present a developed procedure for analyzing, manipulating and meshing HPC-blades to simulate their flow. Therefore, three different software-routines for the mentioned steps will be shown and explained. Additionally, an example blade will pass through the process and will be manipulated for max. deviation caused by deterioration. At the end, a CFD-calculation of these blades will be carried out and analyzed for its aerodynamic behavior.

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In this paper flow field calculations for an entire two shaft industrial gas turbine will be described. This method is based on individual through flow codes for axial compressors and air-cooled expansion turbines developed by the authors which are automatically coupled using simple combustion and secondary flow models connecting compressor and turbine flow paths. With this approach the complete quasi 3-D flow field from compressor inlet to turbine exit can be solved simultaneously (flange-to-flange). Details are explained in this paper. The through-flow computation for the analysis of cooled axial multistage turbines is fed by air from the compressor bleeds which are part of the through flow model of the compressor. The through-flow methods are based on a stream function approach and a finite element solution procedure. They include high-fidelity loss and deviation models with improved correlations. Advanced radial mixing and endwall boundary layer models are applied to simulate 3-D flow effects. For air-cooled turbine analysis, various types of cooling air injection were encompassed: film cooling, trailing edge injection and disc/endwall coolant flow. Compressor and turbine flow path computations were extensively validated individually and published by the authors. Predicted gas turbine operating points of MAN’s MGT-gas turbine will be compared with results of the 3-D Navier- Stokes solver TBLOCK which was run for both compressor and turbines individually using the boundary conditions derived from the present analysis. The focus is on the comparison of mean data and radial distributions at inlet and outlet stations as well as planes between individual stages and blade rows. They will be compared with measured data at MAN’s gas turbine test rig which were obtained in the turn of a prototype telemetry test campaign. It will be demonstrated that the new method presented is an essential and quick tool for overall gas turbine design and matching of the gas turbine components.

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