Effects of pressure side film cooling hole placement and condition on adiabatic film cooling effectiveness characteristics of a transonic turbine blade tip

Abstract

The effects of film cooling hole placement location along the upper pressure side of a transonic squealer are considered. The thermal performance of four different film cooling configurations; B1, B2, B3 and B4, are considered using the University of Alabama in Huntsville's SS/TS/WT (supersonic/transonic/wind tunnel) experimental facility and a simulated turbine blade row using a linear cascade. Surface-varying results are provided for both the squealer blade tip surface, and for the upper pressure side of the squealer blade. These results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged adiabatic film cooling effectiveness distributions. Because local static pressure variations and gradients vary in a significant manner at different blade locations, coolant accumulations near surfaces (and the associated surface thermal protection and film effectiveness distributions) produced by the different film cooling configurations are vastly different along the squealer blade tip surface. The B1 and B3 film cooling configurations show that substantial magnitudes of adiabatic film cooling effectiveness are present along the pressure side rim and the suction side rim. Also present is a considerable region of locally increased film cooling effectiveness and a wider region of surface protection within the squealer recess region for the B2 film cooling configuration. The contributions are unique because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.

Introduction

Turbine blade tip regions and sections near the tip are highly vulnerable to thermal loading and aerodynamic loss issues. These regions are also often the most challenging to cool in order to maintain acceptable temperature levels and acceptable temperature gradients. Results from a number of studies consider tip gap flows and associated heat transfer characteristics for low speed, incompressible flow environments. However, because modern turbine engines often operate with transonic flows, subsonic tip gap data are not sufficient for the design of such transonic turbine blades with film cooling. Bunker [1] presents a review of associated turbine blade tip aerodynamics and heat transfer characteristics. In regard to early investigations, Moore and Tilton [2] provide the first evidence of the existence of shock waves within high speed tip gap flows. Additional confirmation is provided by Moore and Elward [3] using water table experimental data. In a later investigation, Wheeler et al. [4] compare surface heat transfer of plane blade tips with subsonic and transonic conditions. These investigators indicate that higher speeds give reduced surface heat transfer relative to low speed tests, partially because of a reduction in local turbulence production as a result of local flow acceleration. Zhang et al. [5] demonstrate that shock wave reflections are present within tip gap flows, which create large local variations in static pressure and surface heat flux.

O'Dowd et al. [6] investigate a winglet tip using a linear cascade and transient infrared thermography. The investigators indicate that the complex geometry of a winglet give more chaotic shock wave reflections compared to plane tip arrangements. O'Dowd et al. [7] also investigate the aerothermal performance of a cooled winglet tip. According to these investigators, the winglet with film cooling injection outperforms the uncooled winglet in terms of loss coefficient magnitudes. Wheeler and Saleh [8] consider the effects of slot cooling flow on the tip leakage flow rates for both plane and squealer geometries with transonic conditions. Results indicate that slot coolant injection has the potential to reduce aerodynamic loss. In addition, the plane blade tip outperforms the squealer blade tip in terms of both loss and turning at cooling mass flows above 2 percent of mainstream mass flow values. In a numerical study, Wang et al. [9] compare a flat blade tip, a squealer blade tip, and a partial squealer blade tip using three-dimensional conjugate heat transfer analysis. Results show that replacing the squealer with a flat geometry in the rear transonic portion of the blade gives improved aerodynamic performance. The cooled flat tip additionally outperformes the squealer in regard to losses associated with over tip leakage mass flow. Kim et al. [10] optimize squealer geometry with respect to aerothermal performance with cooled and transonic conditions. According to the investigators, varying cavity depth has little effect on the aerodynamic loss, however, modifying the front radius and aft blend radius influence the pressure loss. Arisi et al. [11] investigate a ribbed squealer blade with two dusting holes upstream of mid-cavity ribs. Using oil flow visualizations, the investigators show that the presence of the ribs enhances recirculation and locally chaotic flows. The investigators also demonstrate that the ribs enhance heat transfer and act to block the film coolant from advecting downstream. Ma et al. [12] describe an experimental and numerical study to investigate interactions between squealer blade tip cooling flow and transonic over tip leakage flow. Observed are large local heat transfer variations in between film cooling holes due to vortical structures which originate near each cooling hole exit.

Considering transonic turbine blade tip flows, Hofer and Arts [13] examine the effects of cooling on aerodynamic performance of a full squealer (with a full suction side squealer rim) and a partial squealer (with a partial pressure side squealer rim). The film cooling configuration consists of 16 pressure side film cooling holes and 4 tip camber line dust holes. The partial squealer rim exhibits higher overall loss coefficients with a tip gap of 1.75 percent, relative to the blade chord length. Oil flow visualizations evidence over tip leakage flows which are vastly different for the two configurations. Using computational fluid dynamics, Naik et al. [14] investigate the geometry and film cooling arrangement of Hofer and Arts. Naik et al. show complex interactions from the upper pressure side and the dusting hole flows between the cavity and the casing wall. According to these investigators, the film cooling from camberline-located dusting holes partially impinges upon the casing surface. Also noted are different heat transfer and effectiveness distributions for the squealer and partial squealer configurations along the leading edge and trailing edge regions. Also using computational fluid dynamics analysis, Zhou et al. [15] demonstrate that combined pressure side coolant with tip coolant reduces the overall heat load for transonic squealer blade tips. With simulated endwall motion, the heat transfer within the recess and upper suction side is increased, as film cooling effectiveness is reduced overall. Collopy et al. [16] present data which illustrate the effects of tip gap on transonic squealer blades with pressure side film cooling. Results show that the surface thermal protection is generally improved for a tip gap of 0.8 mm, relative to a larger tip gap of 1.4 mm, due to a decrease in overall heat transfer coefficient and local increases in adiabatic film cooling effectiveness values. Note that the Collopy et al. results are only provided for a single film cooling configuration B1, considering the effects of tip gap, whereas the results within the present paper are provided for four different upper pressure side film hole configurations, but for only one tip gap value.

Engine manufacturers require reliable data regarding the heat transfer and aerodynamic performance and characteristics of turbine blade tips which operate with transonic flow conditions. This is because tip gap flow structure and heat transfer characteristics vary drastically from one geometric configuration to another. This is also because, currently, there is a lack of understanding of film cooling protection for transonic turbine blade tips because the majority of research efforts from the past have focused upon subsonic flow conditions and results. Within the present investigation, spatially-resolved and spatially-averaged surface adiabatic film cooling effectiveness distributions are presented for blowing ratios ranging from 0.42 to 3.20 for four different film cooling configurations. These results are unique and significant because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider and compare the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.