Dissertation
Dissertation > Aviation, aerospace > Aviation > Aero-engine ( propulsion system ) > Engine theory > Aero-engine gas mechanics

Numerical Simulation of Film Cooling in Turbine Cascade with Non-Axisymmetric Endwall Method

Author JiangYanMing
Tutor HuangHongYan
School Harbin Institute of Technology
Course Power Machinery and Engineering
Keywords Film cooling Blowing ratio Non-axisymmetric endwall Numerical simulation
CLC V231.3
Type Master's thesis
Year 2008
Downloads 93
Quotes 0
Download Dissertation

The turbine inlet temperature is increasing continuously, and it has already exceeded the tolerance of the turbine blade material. So it has become the focus which will reduce the temperature of the turbine blade of current research in each country. The objective of modern aero turbine designer and fabricant is how to design blade cooling structure in order to increase the cooling effect. The endwall losses contribute significantly to the overall losses in modern turbo machinery, especially when aerodynamic load increases and aspect ratio decreases. The 3-D non-axisymmetric endwall warp design method can choose the position, shape and height of the profile endwall freely by means of controlling the streamline curvature at different positions in the cascade passage. Non-axisymmetric endwall can control the endwall static pressure distribution and weaken the intensity of the secondary flow at the endwall, so as to improve the performance of the turbomachinery.This paper designs the cooling method of the blade and the endwall in the straight cascade firstly, two cooling cavities with separate sources, three rows of cooling holes are staggered on the leading edge of the blade, and other were placed along streamwise at low endwall. The CFX was employed to solve the film cooling under the different blowing ratio, and then analyzed the characteristic of the cooling injection flow field, temperature distribution on the surface, and the structure of the injection flow field and so on. When the blowing ratio reduce, the cooling air injection near the leading edge performs just as film overlay the surface of the blade, the effect limited in the boundary layer with lower loss; when the blowing ratio increase, back flow area may be exist; the cooling air may penetrate the main flow boundary layer when the blowing ratio increase further, this will affect the flow field of the mainstream.The film cooling effectiveness decreases with the increasing of the blowing ratio in the scale. But when the cooling air ratio decrease excessively, the cooling air is endowed with the ability to traverse the surface boundary layer, then the cooling air is difficult to attach to the surface of the blade near the hole, which result in the decline of the film cooling effectiveness. Three types of non-axisymmetric endwall are designed in the condition of blowing ratio M=1.5. This paper investigates the influence of flow parameters, such as the static pressure, the energy losses, the losses of the total pressure which are important to the cascade aerodynamic performance, under the non-axisymmetric endwall method was adopted. By using the non-axisymmetric endwall method that is lowered near the pressure side and raised near the suction side design in this paper, the static pressure of the endwall will increase, the cross pressure gradient from the pressure side to suction side decreased, however, the energy losses near the endwall may not be reduced, as a result, the cross pressure gradient are not the unique reason to promote the passage vortex production. By research further, we found that the reasonable arrangement of the cooling holes on the endwall may blow away the lower energy fluid that gathered near the hub of the blade which will benefit the flow.

Related Dissertations
More Dissertations