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 relatively large surface area of the platform of the gas turbine blades being exposed directly to the hot, mainstream gas, it is vital to efficiently cool this region of the blades. This region is particularly difficult to protect due to the strong secondary flows developed at the airfoil junction (formation of the leading edge horseshoe vortex) and circumferentially across the blade passage (strengthening passage vortex moving from the pressure side to the suction side of the passage). Over the past decade, researchers and engine designers have attempted to combat the enhanced heat transfer to the blade platform by implementing both frontside and backside novel cooling techniques. This paper presents a review of platform cooling technology ranging from frontside film cooling via stator-rotor purge flow, mid-passage purge flow, and discrete film holes to backside cooling achieved via impinging jet arrays or cooling channels. To gain a full understanding of state-of-the-art cooling technology, recent patents, journal articles, and conference proceedings are included in this review.

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.

This work presents numerical simulation results of a single stage axial turbine consisting of a nozzle and squealer tipped rotor. The VLES method is a hybrid URANS/LES method based on the standard k-ω SST and Coherent Structure LES turbulence models. The simulations were performed at the stage aerodynamic design point (ADP) and the results were validated against high-quality steady experimental data acquired at the University of Notre Dame’s axial transonic research turbine (TRT) facility. Along with the experimental validation, the VLES simulation results were further compared to those predicted using URANS highlighting the benefits of VLES compared to traditional predictive methods. All simulations were performed using a RANS-type grid density to highlight the efficiency of the VLES method and improved performance prediction.