Chiyuki Nakamata

Cooling Performance of an Integrated Impingement and Pin Fin Cooling Configuration

The cooling configuration adopted in this study integrates impingement cooling and pin fin cooling devices into one body, the aim being enhancement of the effective heat transfer area. The purpose of the study is to confirm improvement of cooling effectiveness for two different pin density configurations by experimental verification. Experiments were conducted in similar conditions to actual engines using large-scaled flat-plate specimens manufactured by a new rapid prototype casting technique. The results were compared with predictions by one-dimensional analysis adopting the fin efficiency theory. Although the coarse pin density, one pin in a unit area, shows good agreement with the prediction, the fine pin density, four pins in the unit area, was overpredicted. It was found by numerical analysis that heat transfer of the new pin geometry did not increase, so that its surface area increased. CFD-aided prediction was proposed and validated with two specimen’s data.Copyright

Film Cooling Hole Shape Optimization Using Proper Orthogonal Decomposition

Film cooling is a very effective cooling method for protecting the turbine blades exposed to hot gas from the heat. Since its cooling effectiveness is highly dependent on the shape of the hole, a wide variety of concepts and design parameters regarding hole shapes have been researched. However, there are no well-defined ways to determine the optimum shape of a film cooling hole.The CFD is a powerful tool for film cooling hole optimization. But with the number of parameters that define the film cooling hole shapes being so numerous, analytical optimization with CFD often requires computational resources that are unrealistic for the average design environment. Accordingly, for CFD to be effective in the optimization process, it is necessary to reduce the number of computations or shorten the calculation time per computation.In order to solve this problem, this paper presents a novel approach of applying 3D-POD (3D-Proper Orthogonal Decomposition) to the optimization of film cooling holes. POD is one of the most important component analysis methods and has the potential to reduce the number of parameters.From the computation results, a solution group was made by the RSM (Response Surface Method) and assessment functions, i.e., film cooling effectiveness, heat transfer coefficient, mixing loss, concentration of stress and robustness were considered first. In the end, however, considering the sensitivity of each objective function, the optimal hole shapes were obtained with only the film effectiveness being evaluated.In the following sections, this method and its results are described in detail.Copyright

Simulations of Multiphase Particle Deposition on a Showerhead With Staggered Film-Cooling Holes

The demand for cleaner, more efficient energy has driven the motivation for improving the performance standards for gas turbines. Increasing the combustion temperature is one way to get the best possible performance from a gas turbine. One problem associated with increased combustion temperatures is that particles ingested in the fuel and air become more prone to deposition with an increase in turbine inlet temperature. Deposition on aero-engine turbine components caused by sand particle ingestion can impair turbine cooling methods and lead to reduced component life. It is necessary to understand the extent to which particle deposition affects turbine cooling in the leading edge region of the nozzle guide vane where intricate showerhead cooling geometries are utilized. For the current study, wax was used to dynamically simulate multiphase particle deposition on a large scale showerhead cooling geometry. The effects of deposition development, coolant blowing ratio, and particle temperature were tested. Infrared thermography was used to quantify the effects of deposition on cooling effectiveness. Although deposition decreased with an increase in coolant blowing ratio, results showed that reductions in cooling effectiveness caused by deposition increased with an increase in blowing ratio. Results also showed that effectiveness reduction increased with an increase in particle temperature. Reductions in cooling effectiveness reached as high as 36% at M ¼1.0. [DOI: 10.1115/1.4004757]

Effects of Surface Geometry on Film Cooling Performance at Airfoil Trailing Edge

Cooling at the trailing edge of a gas turbine airfoil is one of the most difficult problems because of its thin shape, high thermal load from both surfaces, hard-to-cool geometry of narrow passages, and at the same time demand for structural strength. In this study, the heat transfer coefficient and film cooling effectiveness on the pressure-side cutback surface was measured by a transient infrared thermography method. Four different cutback geometries were examined: two smooth cutback surfaces with constant-width and converging lands (base and diffuser cases) and two roughened cutback surfaces with transverse ribs and spherical dimples. The Reynolds number of the main flow defined by the mean velocity and two times the channel height was 20,000, and the blowing ratio was varied among 0.5, 1.0, 1.5, and 2.0. The experimental results clearly showed spatial variation of the heat transfer coefficient and the film cooling effectiveness on the cutback and land top surfaces. The cutback surface results clearly showed periodically enhanced heat transfer due to the periodical surface geometry of ribs and dimples. Generally, the increase of the blowing ratio increased both the heat transfer coefficient and the film cooling effectiveness. Within the present experimental range, the dimple surface was a favorable cutback-surface geometry because it gave the enhanced heat transfer without deterioration of the high film cooling effectiveness.

Computational Predictions of Endwall Film Cooling for a Turbine Nozzle Vane With an Asymmetric Contoured Passage

This paper presents results of a computational study for the endwall film cooling of an annular nozzle cascade employing a circumferentially asymmetric contoured passage. The investigated geometrical parameters and the flow conditions are set consistent with a generic modern HP-turbine nozzle. Rows of cylindrical film cooling holes on the contoured endwall are arranged with a design practice for the ordinary axisymmetric endwall. The solution domain, which includes the mainflow, cooling hole paths, and the coolant plenum, is discretized in the RANS equations with the realizable k-epsilon model. The calculated flow field shows that the pressure gradients across the passage between the pressure and the suction side are reduced with the asymmetric endwall, and consequently, the rolling up of the inlet boundary layer into the passage vortex is delayed and the separation line has moved further downstream. With the asymmetric endwall, because of the effective suppression of the secondary flow, more uniform film coverage is achieved especially in the rear part of the passage and the laterally averaged effectiveness is also significantly improved in this region. The closer inspection of the calculated thermal field reveals that, with the asymmetric passage, the coolant ejected from the holes are less deflected by the secondary vortices, and it attaches better to the endwall in this rear part.Copyright

Squealer Tip Heat Transfer With Film Cooling

Heat transfer coefficients and film cooling effectiveness values were obtained numerically on a film cooled 2-D gas turbine blade tip model featuring a cutback squealer. In addition, pressure distributions were obtained at 50% and 98% spans. The calculations were performed for a single blade with periodic boundary conditions imposed along the two mid-passage boundaries formed by the adjacent blades. The calculations were performed with the realizable k-e turbulence model and non-equilibrium wall function using 1.1 million elements. The numerical results are obtained for 4 blowing ratios and for Reynolds number based on axial chord and inlet velocity of 75,000. Limited experimental measurements of the blade pressure distributions and the uncooled tip heat transfer coefficients were performed for validation of the numerical results. The experiments were conducted in a six-blade low-speed wind tunnel cascade at a Reynolds number of 75,000. The heat transfer experiment involved a transient infrared thermography technique. Experimental heat transfer coefficients were extracted using a transient technique. The predicted pressure distributions agree very well with the measurements while the heat transfer coefficient predictions show qualitative agreement. From the numerical results, it can be seen that as the blowing ratio is increased, larger regions of film cooling effectiveness were seen with higher effectiveness values between the camber line and suction side. Heat transfer coefficients were largest near the leading edge for all cases.Copyright

Heat Transfer and Pressure Measurements in a Lattice-Cooled Trailing Edge of a Turbine Airfoil

An experimental study of the heat transfer distribution and pressure drop through a converging lattice-matrix structure has been performed. This structure represents a gas turbine blade trailing-edge cooling passage. Stationary tests were performed on a scaled up model under three Reynolds numbers (24000<Re<60000). To obtain the wall temperature, the narrow band liquid crystal technique was used, and the heat transfer coefficient value was obtained using the transient method. It’s found that the Nusselt number ratio (Nu/Nu0 ) is around 4–5, comparing to the channel flow of similar hydraulic diameter and Re, for the whole lattice-matrix structure. Under the impingement and turning areas, the ratio can be as high as 7–8. Pressure data are taken throughout the lattice structure following the flow direction. The pressure drop increases with Reynolds number and as a result there is a decrease in the thermal performance factor at higher Reynolds number. In the present study thermal performance factor is found to be around 1–1.2. For comparison, pin fin based trailing edge configuration has a typical thermal performance factor of 0.7 to 0.85 under the same Reynolds numbers.Copyright

Experimental Study on Racetrack-Shaped Holes Impingement Cooling With Bump Type Roughening Element

Cooling effectiveness of an impingement cooling with array of racetrack-shaped impingement holes is investigated. Two types of specimens are investigated. One is a plain target plate and the other is a plate roughened with bump type elements. Sensitivity of relative location of bump to impingement hole on the cooling effectiveness is also investigated.Experiments are conducted under three different mainflow Reynolds numbers ranging from 2.6×105 to 4.7×105, with four different cooling air Reynolds numbers for each main flow condition. The cooling air Reynolds numbers are in the range from 1.2×103 to 1.3×104.Copyright

Heat Transfer and Pressure Drop in a Lattice Channel With Bleed Holes

A lattice structure for internal cooling with coolant bleeds is investigated experimentally. The lattice configuration provides a serpentine complex flow passage where the flow takes multiple twists and turns with impingement before exiting the coolant flow channel. The combination of impingement, tortuous flow path and turbulators are expected to provide high heat transfer coefficients. In this study, measurements of heat transfer coefficient and total pressure drop were performed for a constant cross-section lattice geometry with bleed holes at the end of the passage as the flow exits. A transient liquid crystal technique was used for the measurements. Stationary tests were performed for four Reynolds number (5500<Re<22000) in a lattice structure with two inlet channels. The data indicated high heat transfer coefficients at locations corresponding to the impingement sites (with the peak Nu/Nu0 ranging from 8–9 at the lowest Re and 3–4 at the highest Re). The spanwise-averaged Nu/Nu0 ratios showed a rapid asymptotic development and shows constant values beyond about 10 sub-channel hydraulic diameters. Channel averaged Nu/Nu0 values are obtained in the range 2.25–3.1. Pressure drop measurements were made, and are combined with the Nu/Nu0 values to produce a TPF. These values are in the range of 1–1.6 with the higher values exceeding TPF’s of turbulated and pin-fin channels.Copyright

Heat Transfer Enhancement due to Combination of Dimples, Protrusions, and Ribs in Narrow Internal Passage of Gas Turbine Blade

Internal convective cooling of gas-turbine airfoil is essential because turbine inlet temperature becomes higher for pursuing higher thermal efficiency. For higher cooling performance, heat transfer is often enhanced by installing ribs and/or pin-fins in the internal passage. In this study, in order to enhance heat transfer, the combination of spherical dimples, cylindrical protrusions, and transverse square ribs was applied to one wall of a narrow passage. As for the cylindrical protrusions, two different diameter cases were examined. The heat transfer enhancement was measured by a transient infrared thermography method for the Reynolds number of 2,000, 6,000, and 10,000. The pressure loss was also measured in the experiments, and RANS simulation was performed to give a rationale for the experimental results. The present results clearly showed the spatial variation of the local Nusselt number: the high Nusselt number was observed on the rib top-surface and also near the leading edge on the protrusion top-surface. In addition, the areas around the dimple’s trailing-edge on the oblique line connecting the neighbor dimples showed moderately enhanced heat transfer. When two different protrusion-diameter cases were compared, both the mean Nusselt number and the friction factor were similarly higher in the larger protrusion case than the smaller protrusion case, and therefore the larger protrusion case was more effective in cooling even when the pressure loss was taken into account.Copyright