Cooling of modern gas turbine blades and vanes is crucial, as they are exposed to extremely high temperatures. The present investigation makes a three-dimensional conjugate heat transfer analysis of internal convection of a ribbed gas turbine vane-blade system under rotational conditions. Simulations have been performed for a single turbine stage, including stationary vanes and rotating blades. The performance of rib-roughened coolant channels is compared to that of smooth channels for different conditions. Flow turbulence is resolved using the Shear Stress Transport (SST k-omega) model with automatic wall function. The compressibility effects in the external hot-gas flow and the rotational effect-induced complex flow patterns inside the rib-roughened coolant channels have been analyzed. The Coriolis force, caused by the rotation of blades, affects the internal flow structures within the smooth and ribbed coolant channels. Ribbed channels significantly enhance local heat transfer by disrupting Coriolis-induced vortex structures and promoting secondary flow. The flow and heat transfer characteristics within coolant channels vary at different blade positions. A transition from supersonic flow at the leading edge to the transonic flow at the trailing edge of the suction side of rotor blades is observed. Compared to the smooth channels, the ribbed channels achieve a blade temperature reduction of up to 191.8 K and a 49.9% increase in the Nusselt number, with a moderate pressure penalty. The study highlights the critical role of rib geometry and rotational effects in optimizing the internal cooling performance of turbine blades for high-temperature applications.

For effective heat dissipation from turbine blades, consistent efforts have been made to explore several passive heat transfer enhancement techniques in the internal coolant channels of modern gas turbine blades. Various researchers have studied different types of turbulators and vortex generators, like ribs, dimples, and protrusions inside the channel. The present work investigates the thermal performances of a rectangular-shaped cooling channel of a 4-aspect ratio with multiple arrays of compound rib dimples arranged on top and bottom walls for realistic conditions. Numerical simulations have been performed for two types of ribs, i.e., 45 degrees V-rib and broken V-rib with the spherical dimple structures for Reynolds numbers ranging from 20,000 to 80,000. The effects of rib height and dimple depth on flow and heat transfer characteristics are presented for different combinations of rib-dimple structures. The compound rib structures result in the formation of longitudinal vortices, recirculating vortices, and counter-rotating vortices in different scales associated with flow recirculation, flow separation, and flow reattachment. Consequently, significant turbulence and heat transfer improvements are observed in specific regions of the coolant channels. The maximum heat transfer enhancement factor of 2.46, i.e., the ratio of heat transfer rate with rib structure and heat transfer rate in smooth channels, is observed for the configuration of 45 degrees V rib, 1.5 mm height with spherical dimple of 4.0 mm depth. Similarly, a 45 degrees V rib with a height of 1.0 mm and a spherical dimple of 2.0 mm depth produces the maximum thermal performance factor of 1.205.

The present study numerically investigates the flow and heat transfer characteristics in a roughened coolant channel with a 4:1 aspect ratio pertinent to modern gas turbine (GT) blade cooling. Three structural configurations: hybrid V-ribs with spherical-dimples, standalone spherical-dimples, and standalone V-ribs are analyzed for various Reynolds numbers. Simulations are performed to investigate the effects of different geometric arrangements and operational variables relevant to the actual GT blade channels used in aviation industries. The standalone spherical-dimples generate recirculating vortices, while standalone V-ribs produce weak longitudinal vortices. In contrast, the hybrid V-rib and spherical-dimple configuration develops stronger longitudinal vortex pairs, enhancing turbulence mixing and promoting more uniform heat transfer. The highest heat transfer zones shift, with the hybrid configuration achieving a normalized Nusselt number of 2.55–8.40

The present work reports a conjugate heat transfer analysis of a turbine blade exposed to high-temperature compressible gas flow and the convection cooling inside the blade. A nickel-based superalloy material CMSX-4 with better mechanical and anticorrosive properties has been introduced for blade materials, and grooved channels are proposed for heat transfer enhancement in internal convection. Each channel contains nine mini-grooves having groove-depth to channel-diameter ratio in the range of 0.08–1.12. Three prominent turbulence models, namely, k-ε, k-ω shear stress transport (SST), and γ-θ transition SST, are used to capture the flow turbulence in a transonic boundary layer flow. Simulations have been performed for actual operating conditions of turbine blades with a wall-to-gas temperature ratio of 0.84 and an inlet-to-outlet pressure ratio of 1.69. The inlet Reynolds number is 5.3 × 105 for the hot gas region, and for coolant flow, the Reynolds number varies from 16 000 to 70 000. The Mach number reaches to a maximum value of 1.14 in the external hot gas flow. Boundary layer transition and wake flow from nearby blades affect the flow in the suction side of the blade. The incorporation of scalable wall function improves the performance of the k-ε turbulence model. Compared to the smooth channel, a 25 K reduction in the average blade surface temperature and 27.3% enhancement in the Nusselt number in blade cooling are obtained for the grooved cooling channel.

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.