Peigang Yan

Conjugate Heat Transfer Analysis of a High Pressure Air-Cooled Gas Turbine Vane

The ANSYS-CFX software is used to simulate NASA-Mark II high pressure air-cooled gas turbine. The work condition is Run 5411 which have transition flow characteristics. The different turbulence models are adopted to solve conjugate heat transfer problem of this three-dimensional turbine blade. Comparing to the experimental results, k-ω-SST-γ-θ turbulence model results are more accurate and can simulate accurately the flow and heat transfer characteristics of turbine with transition flow characteristics. But k-ω-SST-γ-θ turbulence model overestimates the turbulence kinetic energy of blade local region and makes the heat transfer coefficient higher. It causes that local region temperature of suction side is higher. In this paper, the compiled code adopts the B-L algebra model and simulates the same computation model. The results show that the results of B-L model are accurate besides it has 4% temperature error in the suction side transition region. In addition, different turbulence characteristic boundary conditions of turbine inner-cooling passages are given and K-ω-SST-γ-θ turbulence model is adopted in order to obtain the effect of turbulence characteristic boundary conditions for the conjugate heat transfer computation results. The results show that the turbulence characteristic boundary conditions of turbine inner-cooling passages have a great effect on the conjugate heat transfer results of high pressure gas turbine. ANSYS is applied to analysis the thermal stress of Mark II blade which has ten radial cooled passages and the results of Von Mises stress show that the temperature gradient results have a great effect on the results of blade thermal stress.Copyright

Coupled BEM and FDM Conjugate Analysis of a Three-Dimensional Air-Cooled Turbine Vane

A coupled boundary element method (BEM) and finite difference method (FDM) are applied to solve conjugate heat transfer problem of a three-dimensional air-cooled turbine blade. A loosely coupled strategy is adopted, in which each set of field equations is solved to provide boundary conditions for the other. In the fluid region, computation code (HIT-NS CODE) adopts the FDM to solve the Navier-Stokes equations. In the solid region, the BEM is adopted to resolve the conduction heat transfer equations. An iterated convergence criterion is the continuity of temperature and heat flux at the fluid-solid interface. The solid heat transfer computation code (3D-BEM CODE) is validated by comparing with the results of an analytic solution and the results of commercial code, the results from 3D-BEM CODE have a good agreement with the analytic solution and commercial code results. The BEM uses a weighted residual method to make the Laplace equation convert into a surface integral equation and the surface integral equation is discretized. The BEM avoids the complicated mesh needed in other computation methods and saves the computation time. In addition, the BEM has the characteristic of a combination of an analytic and a discrete solution. So the BEM solutions of heat conduction problems are more accurate. The results of the coupling computation code (HIT-NS-3DBEM CODE) have a good agreement with the experimental results. The adiabatic condition result is different from the results of experiment and code calculation. So the results from conjugate heat transfer analysis are more accurate and they are closer to realistic thermal environment of turbines. Four turbulence models are applied: K-epsilon model, K-omega model, K-omega (SST-Gamma Theta) model, and B-L model adopted by computation code. Different turbulence models gives different the results of vane wall temperature. Comparing the four turbulence models, the different turbulence models can exactly simulate the flow field, but they can not give exact values for the heat conduction simulation in the boundary layer. The result of K-Omega (SST-Gamma Theta) turbulence model is closer to the experimental data.Copyright

Conjugate Heat Transfer Numerical Validation and PSE Analysis of Transonic Internally-Cooled Turbine Cascade

Conjugate heat transfer numerical simulation of a transonic internal cooled turbine vane is carried out with a third-order accuracy TVD (Total Variation Diminishing) scheme and multi-block structured grids using the code developed in this paper. Comparison between results of commercial CFD codes with several turbulence models and those of this code show that it is incorrect of computational codes to predict the thermal boundary layer with traditional turbulence models, and that the turbulence models considering transitional phenomenon is able to acquire better accurate heat transfer in thermal boundary layer despite of certain deficiencies yet. The predicted distributions of aerodynamic parameters agree well with the experiments except for the temperature and heat transfer coefficient over the profile surface, which are largely different from the measured data. Results by Star-CD with V2 -F model meet the same problem. Results by the code of this paper are close to those by CFX with K-ω-SST-ML transition model. Adopting transition model of Menter & Langtry (Shear-Stess-Transition model) gives the best results by adjusting the transition onset momentum thickness Reynolds number and the inlet viscosity ratio, especially for the mid part of the suction side where the validation accuracy is of a serious shortage by traditional turbulence model. It is proved in this paper that commercial codes have the ability to simulate transition process and thermal boundary layer accurately, but the designer’s experience is of utmost importance. Robust transition turbulence model should be developed further. PSE stability analysis equation and e-N prediction method are both integrated in this paper. It is concluded that boundary layer stability analysis by PSE method can be applied to the prediction of the transition onset without too many experiences and is able to define the empirical turbulence parameters for the accurate simulation of thermal boundary layer for any airfoil, such as the very important transition onset momentum thickness Reynolds number.Copyright