Endwall surface film cooling effectiveness was measured on a turbine vane endwall surface using the pressure-sensitive paint (PSP) technique. A double staggered row of holes and a single row of discrete slots were used to supply film cooling in front of the nozzle cascade leading edges. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to be 0.5 to 3.0 percent of the mainstream mass flow. The free-stream Reynolds number was about 283,000 and Mach number was about 0.11. The free-stream turbulence intensity was kept at 6.0 percent for all the tests, measured by a thermal anemometer. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the endwall surface. Film effectiveness distributions were measured on a flat endwall surface for five different mass flow rates. The film effectiveness increased nonlinearly with mass flow rate, indicating a strong interference between the cooling jets and the endwall secondary flows. At lower mass flow ratios, the secondary flow dominated the near wall flow field, resulting in a low film effectiveness. At higher mass flow ratios, the cooling jet momentum dominated the near wall flow field, resulting in a higher film effectiveness. The comparison between hole injection and slot injection was also made.

An experimental study with the application of a custom designed heat flux probe and a modified Kiel probe to measure the turbulence intensity along the turbine gas path downstream of the combustor was performed in an industrial gas turbine engine. The probes were calibrated in a hot-cascade facility test rig which closely emulates engine operating conditions, particularly Re number and Mach number. Correlations based on the test results were compared with the data available in the literature. Indirect turbulence measurements performed in gas turbine engines show that the turbulence intensity at the combustor exit for different combustors can vary from 8% to 12% and the turbulence intensities (combined mainstream and wake generated turbulence) downstream from first and second stage blades can reach 12% to 18%. The results show that the unsteady wakes generated from upstream blade trailing edge play a major role in the turbulence augmentation on downstream nozzle heat transfer. A similar effect is expected in the turbine blade leading edge with a wake turbulence generated by the trailing edge of the upstream vane. An established correlation allows the prediction of a combined turbulence intensity along the gas path when initial combustor mainstream turbulence intensity and engine design parameters are defined (rotating speed, number of airfoils, trailing edge diameter, mainstream velocity, etc.).

Using the pressure sensitive paint (PSP) technique, film cooling effectiveness was measured on a turbine vane pressure surface, with a four-row showerhead cooling hole configuration and a single row of holes on the pressure side. Nitrogen gas was used to simulate film cooling flow providing an oxygen concentration map corresponding to an effectiveness map by the mass transfer analogy. Three showerhead coolant injection angles (45°, 90°, and 135°) were studied and two pressure side injection angles (20° and 40°) for cylindrical holes and a 40° angle for shaped hole were studied. In addition, studies were performed on three combinations of shower head and pressure side injections. Film effectiveness was measured for each of the cases at three blowing ratios. The pressure sensitive paint (PSP) was used to indicate oxygen concentration and was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The results indicate that 45° spanwise angle injection provides best film coverage for the shower head injections. For pressure side injections, the 20° cylindrical hole injection results in the highest effectiveness values and the shaped hole improves film effectiveness immediately downstream from the injection point. The film effectiveness for three combined injections and the interaction between showerhead injection and the pressure side injection are also presented and discussed.

Endwall surface film cooling effectiveness was measured on a turbine vane endwall surface using the pressure sensitive paint (PSP) technique. A double staggered row of holes and a single row of discrete slots were used to supply film cooling in front of the nozzle cascade leading edges. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to be 0.5 to 3.0% of the mainstream mass flow. The freestream Reynolds number was about 283000 and Mach number was about 0.11. The freestream turbulence intensity was kept at 6.0% for all the tests, measured by a thermal anemometer. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the endwall surface. Film effectiveness distributions were measured on a flat endwall surface for five different mass flow rates. The film effectiveness increased nonlinearly with mass flow rate, indicating a strong interference between the cooling jets and the endwall secondary flows. At lower mass flow ratios, the secondary flow dominated the near wall flow field, resulting in a low film effectiveness. At higher mass flow ratios, the cooling jet momentum dominated the near wall flow field, resulting in a higher film effectiveness. The comparison between hole injection and slot injection was also made.

Detailed knowledge of local external heat transfer around a turbine airfoil is critical for an accurate prediction of metal temperature and component life. This paper discusses the results of measurements of the local gas side heat transfer coefficient distribution which were performed in an annular gas turbine vane hot-cascade test facility and provides comparisons with analytical predictions. The steady state tests were performed at simulated engine operating conditions. The tests were conducted on an internally cooled airfoil with the intent to provide close to uniform coolant temperature at the mid-span of the airfoil and also to obtain a high heat flux through the wall, minimizing uncertainty. Internal coolant side heal transfer coefficients were measured through calibration tests outside the cascade by a wire heater. A calibrated infrared pyrometer was used to provide detailed airfoil surface local temperatures and the corresponding external heat transfer coefficients. The airfoil was instrumented with a number of thermocouples for both calibration and temperature measurement. A 12% freestream turbulence was generated by a turbulence grid upstream of the test cascade. The static pressure measured over the test vane agreed well with invicid predictions. Heat transfer coefficients were measured and compared to a TEXSTAN boundary layer code prediction. The influence of Reynolds number and Mach number on the airfoil external heat transfer coefficient were also studied and is presented in the paper.