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This study combines both experiments and computations to investigate pressure-side bleed on the trailing edge of a turbine blade. Realistic engine conditions are considered with a lip thickness to slot height ratio of 0.9 and mainstream Mach numbers of 0.7 at the coolant injection point expanding to sonic conditions at the exit plane of the test section. The purpose of this study is to understand the complex physics of pressure-side bleed, in particular, the unusual behavior that occurs with increasing blowing ratio. Experimentally, it is shown that as the blowing ratio increases, the film cooling effectiveness at a point near the end of the test section increases for blowing ratios less than 0.8, while decreasing over the range of blowing ratios from 1.0 through 1.25. For blowing ratios higher than 1.25, effectiveness increases. This phenomenon has been repeated experimentally for many years without being fully understood. Parts I and II of this paper describe the mechanism responsible for the unusual experimental results. This mechanism is unsteady vortex shedding. Experimental results are from a row of jets with the use of foreign gas injection that simulates the engine conditions that would be seen by the pressure side of an airfoil with pressure-side bleed. These results consist of the pressure distribution due to the nozzle and the effectiveness along the test surface downstream of the injection site. The computational model is designed to replicate the experimental setup. High-quality grids, high-order discretization schemes, and an advanced turbulence model are employed to ensure that the computational results can be used to explain the complex physics of transonic pressure-side bleed film cooling. The grid consists of 2.2 million cells and a high-quality, unstructured, multi-topology, super-block mesh with the resolution of the viscous sub-layer and y+ < 1 on all surfaces. The simulations are fully converged and grid-independent. Effects of blowing ratio are examined, with blowing ratio ranging from 0.5 to 2.0 and a density ratio of 1.52. The geometry consists of not only the transonic mainstream flow and the jet, but also the creeping plenum flow. As a result of the significant lip thickness to slot height ratio, it is shown that unsteady effects are the dominant mechanism in the physics of pressure-side bleed film cooling.

This study examines the unsteady transonic pressure-side bleed film cooling on the trailing edge of a turbine blade and resolves the key mechanism responsible for the unusual relationship between film cooling effectiveness and increasing blowing ratio. This study is meant to show that unsteadiness is the key mechanism causing the unexpected results seen in the experiments. It is believed that this unsteadiness is highly dependent on the ratio of the lip thickness to slot height and the shedding frequencies of the passage and coolant vortices, which depend on blowing ratio. For low blowing ratio, hot passage flow has the dominant vortices. For high blowing ratio, coolant flow has the dominant vortices. For intermediate blowing ratio, the vortices have the potential to interact and cause the unusual behavior seen in pressure-side bleed film cooling. On the basis of these observations, experiments were repeated with pressure probes used to acquire the shedding frequencies at the effectiveness measurement location, which showed that unsteadiness was indeed present. Realistic engine conditions are considered with lip thickness to slot height ratio of 0.9 and mainstream Mach numbers of 0.7 at the coolant injection point and expanding to sonic conditions at the exit plane of the test section. Numerical results are from a 2-D mid-plane cut of the original geometry and a full-pitch 3-D model. Computations use high quality grids, high order discretization schemes, and an advanced turbulence model. The 3-D grid consists of 4.4 million cells and a high quality, unstructured, multi-topology mesh with resolution of the viscous sublayer and y+ < 1 on all surfaces. The simulations are fully converged, time accurate, and grid-independent. A novel methodology is used to introduce unsteadiness into the simulations. Effects of blowing ratio are examined, where blowing ratio is equal to 1.0 for 3-D and ranges from 0.3 to 1.5 for 2-D with a density ratio of 1.52. By performing an unsteady simulation, the unusual relationship between the effectiveness and blowing ratio is demonstrated in an unsteady framework.

State-of-the-art experimental and computational techniques are used to study film cooling on the suction and pressure surfaces of a modern turbine blade under realistic engine conditions. Measured data and predicted results are compared for coolant jets injected through a row of three fundamentally different configurations: (1) Compound-angle round (CAR) holes; (2) Axial shaped holes (ASH); and (3) Compound-angle shaped holes (CASH). Experiments employ a single-passage cascade for validation-quality adiabatic film effectiveness measurements using a gas analysis technique. Computations use a novel combination of geometry and grid generation techniques, discretization scheme, turbulence modeling, and numerical solvers to evaluate a “best practice” standard for use in the gas turbine industry. The gridding procedure uses a super-block, multi-topology, unstructured/adaptive, non-conformal, near-wall resolved mesh to accurately capture all of the mean flow features of the 3-D jet-in-crossflow interaction. The effects of blowing ratio (M) are examined, with M = 1.0, 1.5, and 2.0 on the suction surface and M = 1.5, 3.0, and 4.5 on the pressure surface. All simulations are run with a density ratio of 1.52. The simulations model the three-way coupling between a transonic blade passage flow, subsonic film-hole flow, and creeping plenum flow; high pressure gradients; high rates of curvature; and large strain-rates found in actual engines. Computed results are compared to experimental data in terms of aerodynamic loading and spanwise-averaged adiabatic effectiveness on the blade surfaces in order to validate the computational methodology for this class of problems and to explain the mechanisms responsible for the performance of CAR, ASH, and CASH configurations.Copyright © 2002 by ASME

The performance and physics of film cooling with compound-angle shaped holes on a modern high-pressure turbine airfoil is studied in detail using state-of-the-art computational simulations. Computations model high-speed single-airfoil-passage cascade experiments, and computational results show good agreement with experimental data. Evaluation of physics includes examination of flow features and adiabatic effectiveness. The blowing ratios (M) simulated on the pressure surface (PS) of the blade are 1.5, 3.0, and 4.5, with a single density ratio of 1.52. On the pressure surface the dominant mechanism affecting coolant behavior is vorticity, which increasingly tucks hot crossflow under the coolant as the blowing ratio increases. Thus at high blowing ratios, a lower percentage of the coolant provides thermal protection for the blade until the vortices dissipate far downstream. Also, the vortex structures cause large lateral temperature gradients despite the lateral motion of the flow induced by the compound-angle injection. The dominance of vorticity can be attributed to poor diffusion of the coolant inside the diffuser of the film hole. On the suction surface (SS), the simulated blowing ratios are 1.0, 1.5, and 2.0, with a single density ratio of 1.52. Pressure gradients normal to the SS result in the flow pushing the coolant onto the blade. Also, vorticity is less dominant since diffusion of coolant inside the film hole is better due to low blowing ratios and due to a hole metering section that is almost 3 times longer than that of the PS hole. Hot crossflow ingestion into the film hole is observed at M = 2.0. Ingested crossflow causes heating of the surface inside the hole that extends down to the end of the hole metering section, where the surface temperatures are approximately equal to an average of the crossflow and coolant temperatures. These results demonstrate the inadequacy of 1-D, empirical design tools and demonstrate the need for a validated CFD-based film cooling methodology.

The performance and physics of film cooling with compound-angle shaped holes on a modern high-pressure turbine airfoil is studied in detail using state-of-the-art computational simulations. Computations model high-speed single-airfoil-passage cascade experiments, and computational results show good agreement with experimental data. Evaluation of physics includes examination of flow features and adiabatic effectiveness. The blowing ratios (M) simulated on the pressure surface (PS) of the blade are 1.5, 3.0, and 4.5, with a single density ratio of 1.52. On the pressure surface the dominant mechanism affecting coolant behavior is vorticity, which increasingly tucks hot crossflow under the coolant as the blowing ratio increases. Thus at high blowing ratios, a lower percentage of the coolant provides thermal protection for the blade until the vortices dissipate far downstream. Also, the vortex structures cause large lateral temperature gradients despite the lateral motion of the flow induced by the compound-angle injection. The dominance of vorticity can be attributed to poor diffusion of the coolant inside the diffuser of the film hole. On the suction surface (SS), the simulated blowing ratios are 1.0, 1.5, and 2.0, with a single density ratio of 1.52. Pressure gradients normal to the SS result in the flow pushing the coolant onto the blade. Also, vorticity is less dominant since diffusion of coolant inside the film hole is better due to low blowing ratios and due to a hole metering section that is almost 3 times longer than that of the PS hole. Hot crossflow ingestion into the film hole is observed at M = 2.0. Ingested crossflow causes heating of the surface inside the hole that extends down to the end of the hole metering section, where the surface temperatures are approximately equal to an average of the crossflow and coolant temperatures. These results demonstrate the inadequacy of 1-D, empirical design tools and demonstrate the need for a validated CFD-based film cooling methodology.Copyright © 2002 by ASME

A well-tested computational methodology and high-quality data from a companion experimental study are used to analyze the physics of axial-injected, shaped-hole film cooling on the pressure and suction surfaces of a modern high-pressure turbine blade. Realistic engine conditions, including transonic flow, high turbulence levels, and a nominal density ratio of 1.52, are used to examine blowing ratios of 1.0, 1.5, and 2.0 on the suction surface (SS) and 1.5, 3.0, and 4.5 on the pressure surface (PS). SS results show excellent film-cooling performance with the hole shaping, but massive hot crossflow ingestion is found using similar hole shaping on the PS. Primary mechanisms governing the near and far-field cooling effectiveness and crossflow ingestion are identified, including: (1) the nature of the coolant entry into the film hole; (2) location of hole shaping relative to major coolant flow characteristics; and (3) susceptibility of low-momentum fluid to pressure gradients. Changes in blowing ratio, while not introducing new physical mechanisms, significantly alter the extent to which the mechanisms already present affect the flow. These effects are highly non-linear for both SS and PS geometries, highlighting the inadequacy of one-dimensional design practices and the potential usefulness of CFD as a predictive tool.Copyright © 2002 by ASME

A single passage test model has been designed to simulate the mainstream aerodynamics between two adjacent turbine airfoils and to measure the film cooling effectiveness from coolant injection on the pressure and suction sides of the airfoils. Film cooling tests were run on the model using a gas concentration/mass transfer technique with a foreign gas as the coolant to match density ratio. Aspects of the design and test are discussed including the use of a two-dimensional inviscid flow analysis to design boundary layer bleeds upstream of the pressure- and suction-side airfoil surfaces. Results of two- and three-dimensional viscous flow analyses that were used to evaluate various design features including inlet bellmouth, boundary layer bleeds, adjustable tailboards and model backpressure are presented. Aerodynamic and film cooling effectiveness test measurements made with the model will show that the model flow field can be controlled to match results from a previous thermal cascade test.Copyright © 1995 by ASME