This paper presents a CFD study of a transonic highpressure 1-stage turbine that includes the blade upstream disk cavity. The emphasis of the analysis was to understand and quantify the impact of the blade leading edge platform shape on the flow interaction between the upstream disk cavity flow and the gaspath mainstream flow. Two blade platform shapes were analyzed: a recessed and a raised leading edge shape. The results presented include steadystate and transient simulations in order to describe the flow interaction and quantify the impact on stage efficiency. A sensitivity analysis on the amount of cavity flow was performed to investigate the impact on secondary losses (interpreted by entropy generation) and stage efficiency. It was found that the blade leading edge platform shape and cavity flow amount affected the blade hub passage vortex structure and location. At the nominal engine condition, the raised leading edge platform shape showed an improvement in stage efficiency. It also showed a reduced sensitivity of stage efficiency due to cavity flow amount.

This paper describes the design and performance of a high work single stage research turbine with a pressure ratio of 5.0, a stage loading of 2.2 and cooled stator and rotor. Tests were carried out in a cold flow rig and as part of a gas generator facility. The performance of the turbine was assessed, through measurements of reaction, rotor exit conditions and efficiency, with and without airfoil cooling. The measured cooled efficiency in the cold rig was 79.9%, which, after correcting for temperature and measuring plane location, matched reasonably well the efficiency of 81.5% in the gas generator test. The effect of cooling, as measured in the cold rig, was to reduce the turbine efficiency by 2.1%. A part load turbine map was obtained at 100, 110 and 118% design speed and at 3.9, 5.0 and 6.0 pressure ratio. The influence of speed and the limit load pattern for transonic turbines are discussed. The effect of the downstream measuring distance on the calculated efficiency was determined using three different locations. An efficiency drop of 3.2% was measured between the rotor trailing edge plane and a distance four chords downstream.

This work summarizes the results of the CFD analyses to investigate the effect of the geometrical parameters for a typical coverplate-disk cavity and blade broach system also known as the blade cooling flow supply system. A turbofan high pressure turbine was used as the test vehicle for this investigation. The main objective was to explore potential improvements in engine SFC (aerodynamic performance) by reducing the parasitic work while minimizing the impact on the factors that affect the durability of the turbine blades; feed pressure, temperature, and mass flow. Various tangential on-board injection (TOBI) blade cooling flow supply systems were considered: i) Phase 1 compared the radial TOBI and axial TOBI, ii) Phase 2 compared coverplate-disk cavity shapes, and iii) Phase 3 compared blade broach shapes. The in-house CFD code NS3D was used for the analyses. Compared to the radial TOBI, the axial TOBI has a positive impact on the parasitic work (lower) and blade feed temperature (lower) while it has a negative impact on the blade feed pressure (lower). Further, the coverplate-disk cavity shapes investigated had no significant impact on the parasitic work, blade feed pressure, and blade feed temperature. The CFD solutions show that the major portion of the parasitic work is due to flow turning at the broach entrance. Finally, reducing the blade broach cross-section by sloping up the lower wall has no significant impact on the parasitic work and blade feed temperature but a negative impact on the blade feed pressure and mass flow. Modifying the broach pressure side wall shape is preferred among the blade broach geometries investigated. Future work to improve the CFD analysis consists of performing unsteady analyses to better capture the vortex flow in the blade broach, and including upstream stationary components with either iterative boundary condition modeling or an unsteady multi-stage approach.

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This paper describes the design and performance of a high work single-stage research turbine with a pressure ratio of 5.0, a stage loading of 2.2, and cooled stator and rotor. Tests were carried out in a cold flow rig and as part of a gas generator facility. The performance of the turbine was assessed, through measurements of reaction, rotor exit conditions and efficiency, with and without airfoil cooling. The measured cooled efficiency in the cold rig was 79.9 percent, which, after correcting for temperature and measuring plane location, matched reasonably well the efficiency of 81.5 percent in the gas generator test. The effect of cooling, as measured in the cold rig, was to reduce the turbine efficiency by 2.1 percent. A part-load turbine map was obtained at 100, 110, and 118 percent design speed and at 3.9, 5.0, and 6.0 pressure ratio. The influence of speed and the limit load pattern for transonic turbines are discussed. The effect of the downstream measuring distance on the calculated efficiency was determined using three different locations. An efficiency drop of 3.2 percent was measured between the rotor trailing edge plane and a distance four chords downstream.