The sealing of the stator-rotor gap and rotor-platform cooling are vital to the safe operation of the high-pressure turbine. Contrary to the experience in subsonic turbines, this paper demonstrates the potential to improve the efficiency in transonic turbines at certain rim seal rates. Two types of cooling techniques were investigated: purge gas ejected out of the cavity between the stator rim and the rotor disk, and cooling at the rotor-platform. The tests were carried out in a full annular stage fed by a compression tube at M2is=1.1, Re=1.1×106, and at temperature ratios reproducing engine conditions. The stator outlet was instrumented to allow the aerothermal characterization of the purge flow. The rotor blade was heavily instrumented with fast-response pressure sensors and double-layer thin film gauges. The tests are coupled with numerical calculations performed using the ONERA’s code ELSA. The results indicate that the stator-rotor interaction is significantly affected by the stator-rim seal, both in terms of heat transfer and pressure fluctuations. The flow exchange between the rotor disk cavity and the mainstream passage is mainly governed by the vane trailing edge shock patterns. The purge flow leads to the appearance of a large coherent vortex structure on the suction side of the blade, which enhances the overall heat transfer coefficient due to the blockage effect created. The impact of the platform cooling is observed to be restricted to the platform, with negligible effects on the blade suction side. The platform cooling results in a clear attenuation of pressure pulsations at some specific locations. The experimental and computational fluid dynamics results show an increase in the turbine performance compared with the no rim seal case. A detailed loss breakdown analysis helped to identify the shock loss as the major loss source. The presented results should help designers improve the protection of the rotor platform while minimizing the amount of coolant used.

The accurate heat transfer prediction of film‐cooled blades is a key issue for the aerothermal turbine design. For this purpose, advanced numerical methods have been developed at Snecma Moteurs. The goal of this paper is the assessment of a three‐dimensional Navier‐Stokes solver, based on the ONERA CANARI‐COMET code, devoted to the steady aerothermal computations of film‐cooled blades. The code uses a multidomain approach to discretize the blade to blade channel with overlapping structured meshes for the injection holes. The turbulence closure is done by means of either Michel mixing length model or Spalart‐Allmaras one transport equation model. Computations of thin 3D slices of three film‐cooled nozzle guide vane blades with multiple injections are performed. Aerothermal predictions are compared to experiments carried out by the von Karman Institute. The behavior of the turbulence models is discussed, and velocity and temperature injection profiles are investigated.