The effects of heat treatment (H1 and H2) on the creep behavior and microstructures of Mar-M247 at 871 °C/250 MPa are studied. The results show that the as-cast microstructure is composed of eutectics, γ phase, γ’ phase, MC and M₂₃C₆ carbides, while new M₆C appears in heat-treated microstructures, indicating the transformation of carbides after heat treatments. The Mar-M247 is excellent, with over 1500 h of creep life, and H1 is 48% higher than H2. The addition of post-brazing and diffusion heat treatment in H2 is detrimental to creep resistance; the two steps promoted the transformation of MC into M₂₃C₆ in advance. The increase and coarsening of M₂₃C₆ would consume more γ-phase-forming elements, weakening the solution strengthening at grain boundaries. As a result, the resistance of the grain boundary and γ/γ’ interface to dislocation motion is significantly reduced, leading to the cracks’ initiation and propagation along the grain boundaries.

A multi-objective, aerodynamic optimization of the last stage turbine blade of an 100Hz industrial gas turbine is presented. We aim at maximizing the stage’s efficiency, while at the same time meeting mechanical integrity. The optimization procedure was build, during the optimization process the mass flow rate and stage meridian were kept constant to obtain the required stage loading without changing the operating point. The optimized geometry are chosen taking into account aerodynamic constraints as well as mechanical constrains, such as centrifugal stress and bending stress. The SST turbulence model was used to calculate the flow field of the original and optimized stage, respectively. The geometry of the original and optimized blade was compared. The stage efficiency of the design point and under different operate conditions were compared. Stage and blade efficiency along span was analyzed in details. The blade loading at root mid top span of the original and optimized blade were discussed. Detailed flow field analysis shows that, after optimization, the peak Mach number of the root span profile was decreased from 1.22 to 1.09, which improves the flow field. Profile loss and end wall loss were also reduced. After optimization, maximum centrifugal stress and the maximum air bending stress were decreased which leads to improved blade life and safety. The number of blades and b/h were suggested for the last stage blade.

This study aims to enhance the understanding of film cooling performance in an actual turbine vane by investigating influencing factors and developing more precise numerical prediction methods. Pressure sensitive paint (PSP) testing and Reynolds-Averaged Navier–Stokes (RANS) simulations were conducted. The findings indicate that the current design blowing ratio of S1 holes (0.89) is too high, resulting in poor film cooling effectiveness. However, the blowing ratios of P3 (0.78) and P4 (0.69) holes are relatively low, suggesting that increasing the coolant flow could improve the film cooling effectiveness. It is not recommended to design an excessively low blowing ratio on the suction surface, as this can lead to poor wall adherence downstream of the film holes. A slight increase in turbulence intensity enhances the film covering effect, particularly on the suction surface. Additionally, a novel superposition method for multirow fan-shaped film cooling holes on an actual turbine vane is proposed, exhibiting better agreement with experimental data. Compared with experimental results, the numerical predictions tend to underestimate the film cooling effectiveness with the examined k-ε-based viscosity turbulence models and Reynolds stress turbulence models, while the SST demonstrates relatively higher accuracy owing to its hybrid k-ω/k-ε formulation that better resolves near-wall physics and separation flows characteristic of turbine cooling configurations. This study contributes to the advancement of turbine vane thermal analysis and design in engineering applications.

The low-speed rectangular exit wind tunnel with sectioned contraction is widely used. The secondary flow vortices are found at the contraction exit, which would lead to the non-uniform boundary layer and influence the aerodynamic experiment accuracy. In this paper, the contraction of the impeller mechanical plane cascade experimenter is taken as the research object. Under the