In an axial flow compressor, the presence of separated flow near the hub-end of a stator would result in an overall loss in the performance. In the present paper, stator hub-stall is attempted to be eliminated for a high hub-tip ratio (0.8) axial flow compressor stage consisting of a highly loaded rotor and stator. Numerical and experimental studies on an untreated straight stator (straight-stacked, twisted) blade show a large vortex near its hub. The large vortex is attempted to be reduced by modifying the present blade by (i) providing an additional twist at the hub-end of the stator blade (ii) providing a hub-clearance (a cantilevered blade fixed from the casing). The straight (untreated) stator, hub-end-bend version and the hub-clearance version are studied for two different rotor-tip clearances. Detailed computational analysis of the variation of hub-clearance at a fixed rotor-tip clearance is also carried out. Throughout the study, experiments are carried out on the above discussed different stator (untreated & hub-treated) configurations, in combination with the same rotor, at a fixed rotor-tip clearance. The studies show that the flow conditions are improved near the hub of the highly loaded stator blade both by the hub-end-bend design and by the hub-clearance provided.
Reynolds Averaged Navier-Stokes CFD analysis is carried out for a twin spool turbine of a typical small gas turbine engine, using commercial CFD solvers by employing SST turbulence model to understand the 3D flow field. In this work, detailed performance characterization at design and off-design speeds of both the high pressure (HP) and the low pressure (LP) stages is carried out using two approaches. In the First approach, individual HP and LP turbine stages are analyzed separately (case 1) with well-defined inlet and outlet boundary conditions and, in the second approach HP and LP turbine stages are analyzed together (case 2). NUMECA-AutoGrid is used to generate a good quality mesh with y+ around one. RANS CFD simulations are carried out for design speed, using both ANSYS-CFX and NUMECA“s FINE/Turbo solver and compared for conformity of the CFD analysis results. Further ANSYS-CFX is used for the detailed flow simulations for design and off-design speeds. The turbine parameters such as mass flow function, specific work function and total-to-total adiabatic efficiency of the HP and LP turbines are compared for case 1 and case 2. From case 1 & case 2 analysis, it is observed that, LP turbine stage is capable of allowing higher mass flow than required, but HP turbine stage limits the mass flow. Cascade testing of HP turbine mean section profiles has been carried out and compared with CFD analysis results.
Aero-thermodynamic and mechanical design of a single stage axial turbine stage has been carried out for small gas turbine engine in Propulsion Division, CSIR-NAL. From the engine design configuration extract, it is envisaged that the single stage axial gas turbine operating close to 50500 rpm and at an elevated temperature of 1095K would meet the power requirement of mixed flow compressor of 385kW. This paper presents the aero-thermodynamic, mechanical design and analysis of a single stage highly loaded axial turbine stage with a stage loading coefficient of 1.45 and a flow coefficient of 0.67. The mean-line and detailed 3D aero-thermodynamic design is carried out using commercially available dedicated turbomachinery design codes Axial® and Axcent™ of Concepts NREC. The number of blades of the rotor and stator are 50 & 19 respectively. The turbine stage has undergone a series of design improvements. The final configuration of single stage turbine is analyzed using commercially available RANS CFD software ANSYS-CFX™ and NUMECAFINE™/Turbo flow solver. The design is carried out by aiming 88% total-to-total efficiency. Detailed 3D-RANS CFD analysis of the turbine shows that, the design requirements of turbine are achieved with enhanced efficiency of 90%. Mechanical design & analysis of the turbine stage is carried out using ANSYS-Mechanical™ software. Nimonic-90 material is selected for fabrication. Detailed non-linear steady thermal-structural analysis is carried out for both stator assembly and rotor BLISK. Burst margin of rotor disk is estimated to be around 63% at design speed.
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