Effect of Rotation on Heat Transfer in AR = 2:1 and AR = 4:1 Channels Connected by a Series of Crossover Jets

Detailed heat transfer measurements using transient liquid crystal thermography were performed on a novel cooling design covering the mid-chord and trailing edge region of a typical gas turbine blade under rotation. The test section comprised of two channels with aspect ratio (AR) of 2:1 and 4:1, where the coolant was fed into the AR = 2:1 channel. Rib turbulators with a pitch-to-rib height ratio (p/e) of 10 and rib height-to-channel hydraulic diameter ratio (e/Dh) of 0.075 were placed in the AR = 2:1 channel at 60° relative to flow direction. The coolant after entering this section was routed to the AR = 4:1 section through a set of crossover jets. The 4:1 section had a realistic trapezoidal shape that mimics the trailing edge of an actual gas turbine blade. The pin fins were arranged in a staggered array with a center-to-center spacing of 2.5 times pin diameter. The trailing edge section consisted of radial and cutback exit holes for flow exit. Experiments were performed for Reynolds number of 20,000 at Rotation numbers (Ro) of 0, 0.1 and 0.14. The channel averaged heat transfer coefficient on trailing side was ~28% (AR = 2:1) and ~7.6% (AR = 4:1) higher than the leading side for Ro = 0.1. It is shown that the combination of crossover jets and pin-fins can be an effective method for cooling wedge shaped trailing edge channels over axial cooling flow designs.

Life Prediction of Thermal Barrier Coated C3X Gas Turbine Blade of CMSX-4 Material

The drive to higher efficient gas turbine engines and improved performance is attained by increasing turbine inlet temperature. This lead to the use of advanced material, multi-layered thermal barrier coatings (TBCs) and effective cooling technique in a gas turbine blade system. The main objective is to predict and compare the life between an uncoated blade developed by NASA named C3X and a tri-layered thermal barrier coated C3X blade working under four different high temperature operating conditions. The geometry of the uncoated blade was modelled using Catia software. Three layers of coatings i.e., top coat, bond coat and thermally grown oxide with suitable thickness were generated by the mesh offset technique which was applied to an uncoated blade to model the coated blade. Thereafter steady-state 3D conjugate heat transfer analysis (CFD) with k-ε turbulence model by Ansys Cfx was performed to obtain temperature, pressure and heat transfer coefficient distribution on the surface of the blade. Firstly, this CFD analysis was performed using stainless steel as substrate material, then validated with experimental values and lastly, the same simulation model was applied to Nickel based super-alloy CMSX-4 material. The next step was carrying out transient uncoupled heat transfer, thermal stress, mechanical stress and sequentially coupled thermo-mechanical stress analyses using Abaqus for a flight length of 5000 seconds. At last, the creep-fatigue interaction life of the blade was computed by ductility exhaustion concept with morrow mean stress correction method using Fesafe/Turbolife software. After carrying out above mentioned processes for both uncoated and coated blade, an effective comparison was made. A significant decrease in Temperature (up to 127 K) and Thermo-Mechanical Stress (up to 263 MPa) and a significant increase in Fatigue-Creep Life (up to 16.5 times) was observed when TBCs were applied. The result shows that the thermal load was more severe than the mechanical load. The maximum thermo-mechanical stress was found at the trailing edge and fixed portion of the blade.

Effect of Mainstream Velocity on the Heat Transfer Coefficient of Gas Turbine Blade Tips

This study experimentally investigated the effects of cascade inlet velocity on the distribution and the level of the heat transfer coefficient on a gas turbine blade tip. The tests were conducted in a transient turbine test facility at Korea Aerospace University, and three cascade inlet velocities—30, 60, and 90 m/s—were considered. The heat transfer coefficient was measured using the transient IR camera technique with a linear regression method, and both the squealer and plane tips were investigated. The results showed that the overall averaged heat transfer coefficient was generally proportional to the inlet velocity. As the inlet velocity is increased from 30 m/s to 60 m/s and 90 m/s, the heat transfer coefficient increased by 11.4% and 25.0% for plane tip, and 26.6% and 64.1% for squealer tip, respectively. However, the heat transfer coefficient near the leading edge of the squealer tip and the reattachment region of the plane tip was greatly affected by the cascade inlet velocity. Therefore, heat transfer experiments for a gas turbine blade tip should be performed under engine simulating conditions.