Gas Turbine Engines operate at temperatures higher than current material temperature limits. This necessitates cooling the metal through internal or external means and/ or protecting the metal with coatings that have higher material limits. Film cooling is one of the major technologies allowing today’s gas turbines to operate at extremely high turbine inlet temperatures, consequently higher power density, and extend the cooled components life. Film cooling is a technique where a coolant is blown over the surface exposed to hot gas and a film of low temperature gas is maintained that protects the metal surface from the hot gas. The application of effective film-cooling techniques provides the first and best line of defense for hot gas path surfaces against the onslaught of extreme heat fluxes, serving to directly reduce the incident convective heat flux on the surface. The effectiveness of film cooling methods depends on the blowing ratio, shape of the cooling holes, and geometrical parameters such as the area ratio and diffusion angle. Film cooling is performed almost exclusively through the use of discrete holes. The holes can be of round or other shaped. A detailed study of the literature shows that the fan shaped has higher effectiveness when compared to other shapes. In this study a number of cooling hole shapes are evaluated numerically using the Computational Fluid Dynamics (CFD) tool ANSYS-CFX-11.0 with the objective of improving cooling effectiveness under a favorable pressure gradient main flow. In order to delineate the effects of shape from that of diffusion, a constant area ratio is assumed first. In the next set of analyses the effect of hole exit diffusion is considered. Results are presented in terms of surface temperatures and adiabatic effectiveness at three different blowing ratios for the different film cooling hole shapes analyzed. Comparison is made with reference to the fan shaped film cooling hole with forward and lateral angles of 10/10/10 degree respectively. Hole shapes that show improvement over the fan shaped hole are identified and optimized.

The flow field in axial gas turbines is driven by strong unsteady interactions between stationary and moving components. While time-averaged measurements can highlight many important flow features, developing a deeper understanding of the complicated flows present in high-speed turbomachinery requires time-accurate measurements that capture this unsteady behavior. Toward this end, time-accurate measurements are presented for a fully cooled transonic high-pressure turbine stage operating at design-corrected conditions. The turbine is run in a short-duration blowdown facility with uniform, radial, and hot streak vane-inlet temperature profiles as well as various amounts of cooling flow. High-frequency response surface pressure and heat-flux instrumentation installed in the rotating blade row, stator vane row, and stationary outer shroud provide detailed measurements of the flow behavior for this stage. Previous papers have reported the time-averaged results from this experiment, but this paper focuses on the strong unsteady phenomena that are observed. Heat-flux measurements from double-sided heat-flux gauges (HFGs) cover three spanwise locations on the blade pressure and suction surfaces. In addition, there are two instrumented blades with the cooling holes blocked to isolate the effect of just blade cooling. The stage can be run with the vane and blade cooling flow either on or off. High-frequency pressure measurements provide a picture of the unsteady aerodynamics on the vane and blade airfoil surfaces, as well as inside the serpentine coolant supply passages of the blade. A time-accurate computational fluid dynamics (CFD) simulation is also run to predict the blade surface pressure and heat-flux, and comparisons between prediction and measurement are given. It is found that unsteady variations in heat-flux and pressure are stronger at low to midspan and weaker at high span, likely due to the impact of secondary flows such as the tip leakage flow. Away from the tip, it is seen that the unsteady fluctuations in pressure and heat-flux are mostly in phase with each other on the suction side, but there is some deviation on the pressure side. The flow field is ultimately shown to be highly three-dimensional, as the movement of high heat transfer regions can be traced in both the chord and spanwise directions. These measurements provide a unique picture of the unsteady flow physics of a rotating turbine, and efforts to better understand and model these time-varying flows have the potential to change the way we think about even the time-averaged flow characteristics.

With the advent of fast computers and availability of less costly memory resources, computational fluid dynamics (CFD) has emerged as a powerful tool for the design and analysis of flow and heat transfer of high pressure turbine stages. CFD gives an insight in to flow patterns that are difficult, expensive or impossible to study using experimental techniques. However, the application of CFD depends on its accuracy and reliability. This requires the CFD code to be validated with laboratory measurements to ensure its predictive capacity. In the continual effort to improve analysis and design techniques, Honeywell has been investigating in the use of CFD to predict the aerodynamic performance of a high pressure turbine. Reynolds Averaged Navier Stokes (RANS), unsteady models like detached eddy simulation (DES), large eddy simulation (LES), and Scale Adaptive Simulation (SAS) are used to predict the aerodynamic performance of a high pressure turbine. Mixing plane approach is used to address the flow data transport across the stationary interface in RANS simulation. The film holes on blade surface and end walls for all the analysis are modeled by using actual film hole modeling technique. The validation is accomplished with the test results of a high pressure turbine, Energy Efficient Engine (E3). The aerodynamic performance data at design point, typical off-design points are taken as test cases for the validation study. One dimensional performance parameters such as corrected mass flow rate, total pressure ratio, cycle efficiency, and two dimensional spanwise distributions of total pressure, total temperature and flow angle that are obtained from CFD results are compared with test data. Streamlines and flow field results at different measurement planes are presented to understand the aerodynamic behavior.

This paper presents results of a study of the conjugate heat transfer (CHT) to calculate the metal temperature for a film-cooled gas turbine blade. ANSYS CFX14.0 code was selected as the computational fluid dynamic (CFD) tool to perform the CHT simulation. The two-equation SST turbulence model with automatic wall treatment was employed. The main flow inlet and exit boundary conditions were deduced from a multi-blade-row CFD code, Fine/Turbo by NUMECA. A core engine test operated at the maximum power condition. Thermocouples were used to validate the blade metal temperature calculations. The blade temperature comparison between test data and CHT predictions was in good agreement except at the suction side near the leading edge region. The pressure, temperature and Mach number distributions for blade internal and external flows were presented and examined. The streamline contours of the film flows on the pressure side and suction side were plotted and used to visualize the cooling effectiveness. In order to evaluate the influence of the turbulence model, the thermal results of four additional turbulence models (SA, RNG, K-ε, and SST with transition control) were compared to the test data. The SST model is suggested to be the appropriate turbulence model for the film-cooled blade temperature calculation in this study.

An experiment is performed using a cooled transonic high-pressure turbine stage operating at design-corrected conditions. Pressure measurements are taken at several locations within the forward purge cavity between the high-pressure stator and rotor, as well as on the blade platforms and vane inner endwalls. Double-sided Kapton heat-flux gauges are installed on the upper surface of the rotor blade platform (open to the hot gas path flow) and underneath the platform (exposed to coolant and leakage flow). The blade airfoil and purge flow cooling are supplied by the same flow circuit and must be varied together, but the influence of the airfoil cooling has previously been shown to be negligible in the platform region flow of interest to this study. A separate cooling circuit supplies the aft purge flow between the rotor and downstream components. The vane cooling holes have been blocked off for this experiment to simplify analysis. In order to determine the effect of the purge flow on the blade aerodynamics and heat transfer, the forward and aft cooling flow rates are varied independently. Both time-averaged and time-accurate results are presented for the pressure and heat-flux data to illustrate the complex interactions between the purge cavity flow structures and the external flow. Time-accurate data are presented using both Fast-Fourier Transforms (FFTs) to identify driving frequencies and ensemble average plots to highlight the impact of different wake shapes.