Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
1999-04-05
2004-07-13
Verdier, Christopher (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C416S09000A, C416S092000, C416S191000, C416S192000
Reexamination Certificate
active
06761534
ABSTRACT:
TECHNICAL FIELD
This invention relates to a cooling circuit for a gas turbine bucket and tip shroud, using air from the gas turbine compressor.
BACKGROUND OF THE INVENTION
Gas turbine bucket tip shrouds are subject to creep damage due to the combination of high temperature and centrifugally induced bending stresses. In U.S. Pat. No. 5,482,435, there is described a concept for cooling the shroud of a gas turbine bucket, but the cooling design relies on air dedicated to cooling the shroud. Other cooling arrangements for bucket airfoils or fixed nozzle vanes are disclosed in U.S. Pat. Nos. 5,480,281; 5,391,052 and 5,350,277.
BRIEF SUMMARY OF THE INVENTION
This invention utilizes spent cooling air exhausted from the airfoil itself for cooling the associated tip shroud of the bucket. Specifically, the invention seeks to reduce the likelihood of gas turbine tip shroud creep damage while minimizing the cooling flow required for the bucket airfoil and shroud. Thus, the invention proposes the use of air already used for cooling the bucket airfoil, but still at a lower temperature than the gas in the turbine flowpath, for cooing the tip shroud. This more efficient use of cooling air has the dual advantage of tip shroud cooling with minimal degradation of performance.
In one exemplary embodiment of the invention, leading and trailing groups of cooling passages extend radially within the blade or airfoil. Each group of holes communicates with a common chamber or plenum in the tip shroud. Spent cooling air from the radial cooling passages thus flows into the tip shroud plenum, and then exits through passages from the plenum into the hot gas path. The plenum extends throughout the tip shroud, substantially from front-to-back and side-to-side, lying substantially in the plane of the shroud. The cooling air exits into the hot gas path via passages extending from the plenum to the peripheral edges of the tip shroud. Some cooling air may also be exhausted through one or more metering holes in the top surface of the tip shroud.
In a second exemplary embodiment, two discrete plenums are provided on the tip shroud, one for each of the group or set of leading cooling holes and the group or set of trailing cooling holes. A cover is provided for each plenum, extending above the tip shroud top surface. Here again, cooling air exhausts through passages extending from the plenums to the peripheral edges of the tip shroud, and, optionally, through one or more metering holes in the covers.
In its broader aspects, therefore, the invention relates to an open cooling circuit for a gas turbine airfoil and associated tip shroud including a first group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a leading edge of the airfoil; a second group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a trailing edge of the airfoil; a common plenum in the tip shroud in direct communication with the first and second group of cooling holes; and a plurality of exhaust holes extending from the plenum, through the tip shroud and opening along a peripheral edge of the tip shroud.
In another aspect, the invention relates to an open cooling circuit for a gas turbine airfoil and associated tip shroud comprising a first group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a leading edge of the airfoil; a second group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a trailing edge of the airfoil; a pair of plenums in the tip shroud, each in communication with one of the first and second groups of cooling holes; a plurality of exhaust holes extending from the pair of plenums, through the tip shroud and opening along a peripheral edge of the tip shroud.
In still another aspect, the invention relates to a method of cooling a gas turbine airfoil and associated tip shroud comprising a) providing radial holes in the airfoil and supplying cooling air to the radial holes; b) channeling the cooling air to a plenum in the tip shroud and c) passing the cooling air from the plenum and through the tip shroud.
Additional objects and advantages of the invention will become apparent from the detailed description which follows.
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General Electric Company
Nixon & Vanderhye P.C.
Verdier Christopher
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