ABILITY OF A POPULAR TURBULENCE MODEL TO CAPTURE CURVATURE EFFECTS: A FILM COOLING TEST CASE

ABSTRACT Computations are performed in conjunction with code validation quality experiments found in the open literature to specifically address the usage of popular two-equation eddy viscosity models in day-to-day gas turbine applications. In such simulations many features such as pressure gradients, curvature effects are present. The present work is focused on testing a popular turbulence model to resolve film cooling on curved surfaces. A systematic computational methodology has been employed in order to minimize numerical errors and evaluate the performance of a popular turbulence model. The test cases were examined for a single row of holes, blowing rates ranging from 1 to 2.5, isolated effects of convex and concave curvature on film cooling, density ratio close to 2, and an injection angle of 35°. Key aspects of the study include: (1) extremely dense, high quality, multi-block, multi-topology grid involving over 3 million finite volumes; (2) higher order discretization; (3) turbulence model with two-layer near-wall treatment; (4) strict convergence criteria; and (5) grid independence. A fully-implicit, pressurecorrection Navier-Stokes solver is used to obtain all the solutions. Results for adiabatic cooling effectiveness are compared with measurements in order to document the: (1) Range of applicability of the present modeling capability; and (2) Possible reasons for discrepancies. The data shows that the computations predicted the effects of curvature on mean flow, however effect on turbulence field is not captured. A clear set of recommendations is provided for future treatments of this class of problems

TRANSONIC PASSAGE TURBINE BLADE TIP CLEARANCE WITH SCALLOPED SHROUD: PART III -HEAT TRANSFER IN ENGINE CONFIGURATION

ABSTRACT This work presents a numerical study that was done to investigate the heat transfer characteristics of a transonic turbine blade with a scalloped shroud operating at realistic engine conditions typical of those found in a large scale, land-based gas turbine. The geometry under investigation was an infinite, linear cascade composed of the same blade and shroud design used in an experimental test rig by the research sponsor. This simulation was run for varying nominal tip clearances of 20, 80, and 5.08 mm. For each of these clearances, the simulation was run with and without the scrubbing effects of the outer casing, resulting in a total of six cases that could be used to determine the influence of tip clearance and relative casing motion on heat transfer. A high quality grid (ranging from approximately 10-12 million finite volumes depending on tip clearance) with y + for first layer cells at or below 1.0 everywhere was used to resolve the flow down to the viscous sublayer. The "realizable" k-ε turbulence model was used for all cases. A constant wall heat flux was imposed on all the surrounding surfaces to obtain heat transfer data. Results produced include a full map of heat transfer coefficients for the suction and pressure surfaces of the blade as well as the tip, shroud, and outer casing for every case. Physical mechanisms responsible for the final heat transfer outcome for all six cases are documented

Impact of Film-Cooling Jets on Turbine Aerodynamic Losses

ABSTRACT This paper documents a computational investigation of the aerodynamic impact of filmcooling on a linear turbine airfoil cascade. The simulations were for single row injection on both the pressure and suction surfaces, downstream of the leading edge region. The cases match experimental efforts prey iouslydocumented in the open literature. Results were obtained for density ratio equal to 1.0 and 2.0, and a blowing ratio range from 0.91 to 6.6. The domain included the passage flow as well as the film-hole and blade interior. The simulation used a dense, high-quality. unstructured hybrid-topology grid, comprised of hexahedra, tetrahedra. prisms and pyramids. The processing was performed with a pressurecorrection solution procedure and a second-order discretization scheme. Turbulence closure was obtained using standard, RNG, and "realizable" k-E models, as well as a Reynolds stress model. Results were compared to experimental data in terms of total pressure loss downstream of the blade row. Flow mechanisms responsible for the variation of aerodynamic losses due to suction and pressure surface coolant injection are documented. The results demonstrate that computational methods can be used to accurately predict losses on film-cooled airfoils

PHYSICS OF HOT CROSSFLOW INGESTION IN FILM COOLING

ABSTRACT Computational fluid dynamics (CFD) is used to isolate the flow physics responsible for hot crossflow ingestion, a phenomenon that can cause failure of a film cooled gas turbine component. In the gas turbine industry, new compound-angle shaped hole (CASH) geometries are currently being developed to decrease the heat transfer coefficient and increase the adiabatic effectiveness on film cooled surfaces. These new CASH geometries can have unexpected flow patterns that result in hot crossflow ingestion at the film hole. This investigation examines a 15° forward-diffused cylindrical film hole injected streamwise at 35° with a compound angle of 60° (FDIFF60) and with a length-to-diameter ratio (UD) of 4.0. Qualitative and quantitative aspects of computed results agreed well with measurements, thus lending credibility to predictions. The FDIFF60 configuration is a good representative of a typical CASH geometry, and produces flow mechanisms that are characteristic of CASH film cooling. FDIFF60 has been shown to have impressive downstream film cooling performance, while simultaneously having undesirable ingestion at the film hole. In addition to identifying the Physical mechanisms driving ingestion, this paper documents the effects on ingestion of the blowing ratio, the density ratio, and the film hole Reynolds number over realistic gas turbine ranges of 0.5 to 1.88, 1.6 to 2.0, and 17,350 to 70,000. respectively. The results of this study show that hot crossflow ingestion is caused by a combination of coolant blockage at the film hole exit plane and of crossflow boundary layer vorticity that has been re-oriented streamwise by the presence of jetting coolant: Ingestion results when this re-oriented vorticity passes over the blocked region of the film hole. The density ratio and the film hole Reynolds number do not have a significant effect on ingestion over the ranges studied, but the blowing ratio has a surprising non-linear effect. Another important result of this study is that the blockage of coolant hampers convection and allows diffusion to transfer heat into the film hole even when ingestion is not present. This produces both an undesirable temperature gradient and high temperature level on the film hole wall itself. Lessons learned about the physics of ingestion are generalized to arbitrary CASH configurations. The systematic computational methodology currently used has been previously documented and has become a standard for ensuring accurate results. The methodology includes exact modeling of flow physics, proper modeling of the geometry including the crossflow, plenum, and film hole regions, a high quality mesh for grid independent results, second order discretization, and the two-equation k-c turbulence model with generalized wall functions. The steady, Reynolds-averaged Navier-Stokes equations are solved using a fully-elliptical and fullyimplicit pressure-correction solver with multi-block unstructured and adaptive grid capability and with multi-grid convergence acceleration. NOMENCLATUR

Film Cooling on a Modern HP Turbine Blade: Part IV -- Compound-Angle Shaped Holes

ABSTRACT The performance and physics of film cooling with compoundangle 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. NOMENCLATURE CASH compound-angle shaped holes C P pressure coefficient = (P-P ¥ )/(r ¥ v ¥ 2 /2

Film Cooling on a Modern HP Turbine Blade: Part II -- Compound-Angle Round Holes

ABSTRACT A well-tested computational methodology and a companion experimental study are used to analyze the physics of compoundangle, cylindrical-hole film cooling on the pressure and suction surfaces of a modern high-pressure turbine airfoil. A single-passage cascade (SPC) is used to model the blade passage flow experimentally and computationally. 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). The predicted results agree with experimental trends, and differences are explained in terms of known deficiencies in the turbulence treatment. The mean-flow physics downstream of coolant injection are influenced primarily by a single dominant vortex that entrains coolant and mainstream fluid, and by the effect of convex (SS) or concave (PS) curvature on the coolant jet