Numerical Investigations of Tangential Direction Cooling in the Intermediate Pressure Cylinder (IPC)

2015 ◽  
Author(s):  
Liangliang Liu ◽  
Daiwei Zhou ◽  
Hao Liu ◽  
Xiaocheng Zhu ◽  
Zhaohui Du
Keyword(s):  
Mach Number ◽  
Prandtl Number ◽  
Turbulence Intensity ◽  
Internal Flow ◽  
Tangential Direction ◽  
Cooling Effect ◽  
Turbulent Prandtl Number ◽  
Cooling Effectiveness ◽  
Intermediate Pressure ◽  
Pressure Cylinder

The cooling protection of the hot end is the key technique for ultra-supercritical steam turbine to cool the hot components. In this paper, the internal flow mechanism and cooling characteristics in the tangential direction cooling channel in the intermediate pressure cylinder (IPC) are investigated using the method of computational fluid dynamics. Various turbulence models and mesh sizes are evaluated (the k-ε model, the k-ω model, the SST model). Also, calculated results of the nonrotating bottom wall are compared with the experimental results in a previous research to show the reliability of the CFD program used. Detailed predictions of the contours of velocity, pressure and temperature are carried out. The focus of this study is to investigate the effects of inlet Mach number, turbulent Prandtl number, rotational wall speed, and inlet turbulence intensity on the cooling effectiveness of the vortex channel. The inlet Mach number, the turbulent Prandtl number, the rotational wall speed and the inlet turbulence intensity varied from 0.315 to 0.512, from 0.71 to 1.0, from 0 to 3000rpm, and from 1% to 10%, respectively. Results show that the parameters have different effects on the cooling effectiveness of the vortex cooling chamber. The cooling effect of the vortex channel increases with increasing inlet Mach number and rotational wall speed, while decreases with the increasing turbulent Prandtl number. As the inlet turbulence intensity increases, the cooling effect increases firstly, and then decreases.

Numerical Investigation of Gas Turbine Combustor Liner Film Cooling Slots

2018 ◽  
Author(s):  
Firat Kiyici ◽  
Ahmet Topal ◽  
Ender Hepkaya ◽  
Sinan Inanli
Keyword(s):  
Heat Transfer ◽  
Prandtl Number ◽  
Surface Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Turbulent Prandtl Number ◽  
Surface Cooling ◽  
Gas Turbine Combustor ◽  
Cooling Effectiveness ◽  
Combustor Liner

A numerical study, based on experimental work of Inanli et al. [1] is conducted to understand the heat transfer characteristics of film cooled test plates that represent the gas turbine combustor liner cooling system. Film cooling tests are conducted by six different slot geometries and they are scaled-up model of real combustor liner. Three different blowing ratios are applied to six different geometries and surface cooling effectiveness is determined for each test condition by measuring the surface temperature distribution. Effects of geometrical and flow parameters on cooling effectiveness are investigated. In this study, Conjugate Heat Transfer (CHT) simulations are performed with different turbulence models. Effect of the turbulent Prandtl Number is also investigated in terms of heat transfer distribution along the measurement surface. For this purpose, turbulent Prandtl number is calculated with a correlation as a function of local surface temperature gradient and its effect also compared with the constant turbulent Prandtl numbers. Good agreement is obtained with two-layered k–ϵ with modified Turbulent Prandtl number.

scholarly journals Numerical Simulation and Aerothermal Physics of Leading Edge Film Cooling

1998 ◽  
Author(s):  
A. Chernobrovkin ◽  
B. Lakshminarayana
Keyword(s):  
Numerical Simulation ◽  
Mach Number ◽  
Numerical Investigation ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Leading Edge ◽  
Length Scale ◽  
Cooling Effectiveness ◽  
Turbulence Length ◽  
Turbulence Length Scale

A viscous flow solver based on the Runge-Kutta scheme has been modified for the numerical investigation of the aerothermal field due to the leading edge film cooling at a compound angle. An existing code has been modified to incorporate multi-block capabilities. Good agreement with the measured data has been achieved. Results of the numerical investigation have been used to analyze the vortex structure associated with the coolant jet-freestream interaction to understand the contribution of different vortices on the cooling effectiveness and aerothermal losses. Two counter-rotating vortices generated by the interaction between the mainflow and the coolant jet have been found to have a major influence in decreasing the cooling efficiency through strong entrainment of the hot fluid. Numerical simulation was carried out to investigate the influence of the inlet Mach number, inlet turbulence intensity, and length scale on the aerothermal field due to the leading edge film cooling. Variation of the inlet Mach number leads to a minor modification of the cooling effectiveness, and this is predominantly caused by the modified pressure gradient. Increased turbulence intensity has profound effect on the cooling near the leading edge. Adiabatic effectiveness downstream of the second row of coolant holes is less sensitive to a change in turbulence intensity. Results of the numerical simulation indicate that the turbulence length scale has a significant effect on the accuracy of the numerical prediction of film cooling. Not only the inlet turbulence intensity but also the turbulence length scale should be accurately set to achieve a reliable numerical prediction of the heat and mass transfer due to film cooling.

Effect of Inlet Mach Number on Aerothermodynamics of Sealing Leakage Flow Cooling on Endwall

2020 ◽  
Author(s):  
Hongyu Gao ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Cooling Effect ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Low Mach Number ◽  
Blowing Ratio ◽  
Passage Vortex ◽  
Adiabatic Cooling ◽  
Cooling Design

Abstract The gas turbine has an assembly gap between the combustion chamber and the first stage vane. The coolant air discharge from the gap can prevent the ingestion of the high temperature gas. This leakage flow also provides a cooling coverage on the vane endwall. Taking the cooling effect of the leakage flow on the endwall into consideration is very important for an efficient cooling design. In this paper, the cooling effect of leakage flow on endwall is studied by means of experimental and numerical methods. The study included slots at 30°, 45°, and 60° angles, and six blowing ratios of 0.3, 0.6, 1.0, 1.4, 1.7, and 2.0. The experiment and numerical calculation are conducted under the condition that the inlet Mach number is 0.125 and the outlet Mach number is 0.72, which is close to the working Mach number of the real gas turbine. Under the same slot inclination and blowing ratio, the distribution of endwall adiabatic cooling effectiveness is more nonuniform under the condition of near-real engine Mach number. This is because the passage vortex is weaker under the low Mach number condition, and the leakage flow has a better wall attachment effect. In terms of the spanwise average of endwall adiabatic cooling effectiveness, when the blowing ratio is small, the adiabatic cooling effectiveness is lower under the condition of near-real engine Mach number than that under the condition of low Mach number, but the opposite is true under the condition of large blowing ratio. This is because under the condition of large blowing ratio, the turbulence is stronger under the condition of Mach number of near-real engine. With the reduction of blowing ratio, the turbulent kinetic energy weakens more strongly. In the studied cases, there is a critical blowing ratio of 1.0, and the total endwall cooling adiabatic cooling effectiveness is not significantly affected by the Mach number when it is smaller than M1.0. The average adiabatic cooling effectiveness of the endwall under the condition of near-real engine Mach number is about 7% lower than that under the condition of low Mach number. It means that the experimental results of leakage flow cooling obtained under the condition of low inlet Mach number need to be corrected by a correction factor, which may be less than 1 to make it engine relevant.

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