Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough

Abstract

A gas turbine combustor liner, including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical.

Description

BACKGROUND OF THE INVENTION[0001]The present invention relates generally to a liner for a gas turbine engine combustor and, in particular, to the configuration of the cooling holes utilized in a multihole cooling scheme for such liner.[0002]Combustor liners are generally used in the combustion section of a gas turbine engine located between the compressor and turbine sections of the engine, although such liners may also be used in the exhaust sections of aircraft engines that employ augmentors. Combustors generally include an exterior casing and an interior combustor where fuel is burned to produce a hot gas at an intensely high temperature (e.g., 3000° F. or even higher). To prevent this intense heat from damaging the combustor case and the surrounding engine before it exits to a turbine, a heat shield or combustor liner is provided in the interior of the combustor.[0003]Various liner designs have been disclosed in the art having different types of cooling schemes. One example of a...

AI Technical Summary

Benefits of technology

[0006]In accordance with a first exemplary embodiment of the invention, a gas turbine combustor liner is disclosed as including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical.

[0007]In a second exemplary embodiment of the invention, a gas turbine combustor liner is disclosed as including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end

Problems solved by technology

In particular, each cooling hole includes a first portion having a substantially uniform diameter through said liner and a second portion having a non-uniform diameter through said liner.

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