| dc.description.abstract | Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane in a more authentic configuration because of the complicated flow characteristics and intensive heat load that comes from the exit of the combustor. With the three-dimensional vane geometry and the corresponding wind tunnel design, a more realistic flow field can thus be simulated. The annular sector cascades were therefore selected and the Pressure-sensitive paint (PSP) technique was used to study the film cooling effectiveness on the endwall and vane’s suction surface. On the other hand, PSP technique was also applied on a linear cascade to study the effect of upstream leakage injection angle on various endwall film cooling designs.
In the first part of the study, full-scale turbine vanes were used to construct a three-vane annular sector cascade. Two kinds of film cooling designs are attained through the layback fan-shaped and cylindrical holes dispersed on the vanes and endwalls. This study targets the film cooling effectiveness comparison on the inner endwall of turbine vane. Tests were performed under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09. Coolant-to-mainstream mass flow ratios (2%, 3%, 4%) and density ratios (1.0, 1.5) were examined. The results provide the gas turbine engine designer a direct performance comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.
The second part of the study is focused on the film cooling on the suction surface of a turbine vane under transonic flow condition. The experiments were performed in a five-vane annular sector cascade. The exit Mach numbers were varied from 0.7, 0.9, and 1.1. Density ratios, ranging from 1.0, 1.5 to 2.0; three blowing ratios ranging from 0.7, 1.0, 1.6 for the suction surface were studied. Three rows of radial-angle cylindrical holes around the leading edge and two rows of compound-angle shaped holes on are the suction surface of the test vane. The results demonstrate the mix effects of coolant amount, density ratio, and exit Mach number on the film cooling as well as its interaction with the potential shock wave.
The third part of the study is to investigate the effects of upstream leakage injection angle on film cooling effectiveness of a turbine vane endwall with various film-hole designs. To simulate the upstream leakage flow between the combustor and the turbine vane, three potential leakage injection angles were studied: 30 degrees, 40 degrees, and 50 degrees. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, mid-chord row, and downstream row. Experiments were conducted in a four-passage linear cascade at the exit isentropic Mach number 0.5 corresponding to the inlet Reynolds number 380,000 based on turbine vane axial chord length. For a fair performance comparison, the coolant was fixed at mass flow ratio 1% and density ratio 1.5. The results provide the gas turbine designers a valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream leakage injection angles. | en |
