Turbine blade with impingement cavity cooling including pin fins
a technology of impingement cavity and turbine blade, which is applied in the direction of blade accessories, engine fuctions, machines/engines, etc., can solve the problems of localized hot spots and the likelihood of failure, and achieve the effect of increasing reducing the flow rate of cooling air
Inactive Publication Date: 2013-04-04
SIEMENS AG
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Benefits of technology
The patent describes a turbine blade cooling system that includes an air supply passage and two cooling circuits. One cooling circuit is a serpentine path with a first pass that extends radially to the blade tip wall and receives cooling air from the supply passage. The other cooling circuit is a second flow of air that includes the upstream and discharge impingement cavities. The pin fins in the upstream impingement cavity restrict the flow of cooling air through the second circuit, which increases the flow rate of cooling air through the first circuit. The technical effect is improved cooling efficiency for the turbine blade.
Problems solved by technology
In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
However, some configurations of flow circuits comprising the flow channels may prevent some areas of the turbine blade from being adequately cooled, which may result in the formation of localized hot spots.
Method used
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[0019]In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
[0020]Reference is now made to FIG. 1 which shows a turbine blade 2 according to aspects of the present invention. The turbine blade 2 is one of a plurality of turbine blades that are circumferentially aligned in a gas turbine engine. The turbine blade 2 comprises an airfoil 4 extending radially outwardly from a mounting location such as a rotor disk of a rotor assembly. The airfoil 4 includes an airfoil outer wall 6 extending in a span-wise direction radially outwardly from a blade platform 8 and blade root 10. The airfoil outer wall 6 ...
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A turbine blade including an airfoil having a cavity defined between pressure and suction side walls, and first and second cooling circuits located within the cavity. The second cooling circuit includes a first upstream impingement cavity, a second upstream impingement cavity and a discharge impingement cavity. A plurality of pin fins are located within each of the impingement cavities, extending between the pressure and suction side walls to provide cooling to the pressure and suction side walls, and to restrict flow through the second cooling circuit and to direct an increased flow of cooling fluid into the first cooling circuit.
Description
FIELD OF THE INVENTION[0001]This invention is directed generally to turbine blades and, more particularly, to a turbine blade having cooling circuits for conducting a cooling fluid through an airfoil of the blade to provide an improved thermal balance between parts of the airfoil.BACKGROUND OF THE INVENTION[0002]A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a rotor assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the rotor assembly, to rotate.[0003]Typical turbine combustor configurations expose turbine blade assemblies ...
Claims
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Login to View More IPC IPC(8): F01D5/18
CPCF05D2260/201F01D5/186
Inventor JIANG, NAN
Owner SIEMENS AG