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CMC NOZZLE SEALING ARRANGEMENT

IP.com Disclosure Number: IPCOM000247720D
Publication Date: 2016-Sep-29
Document File: 8 page(s) / 170K

Publishing Venue

The IP.com Prior Art Database

Abstract

The invention relates to gas turbine engines. Specifically, the invention relates to a sealing arrangement for controlling airflow between shroud segments and axially adjacent engine members.

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Page 01 of 8

CMC NOZZLE SEALING ARRANGEMENT

TECHNICAL FIELD

    The invention relates to gas turbine engines. Specifically, the invention relates to a sealing arrangement for controlling airflow between shroud segments and axially adjacent engine members.

BACKGROUND

    Gas turbine engines typically include a multistage axial flow low pressure compressor and a multistage axial flow high pressure compressor which supplies high pressure air to a combustor. The compressors include stages of stationary components referred to as stators and stages of rotational components, which add work to the system, referred to as rotors.

    A portion of compressed high pressure air supplied to the combustor is mixed with fuel, ignited, and utilized to generate hot combustion gases which flow further downstream to one of the multistage flowpaths. Particularly, the combustion gases flow through one or more turbine stages which extract energy from the hot gases to power the rotors in the compressors and provide other useful work.

    One turbine stage downstream from the combustor is commonly referred to as a turbine nozzle stage and includes a plurality of circumferentially spaced vanes that extend in a radial direction with respect to a central axis of the turbine engine. The vanes extend between an outer band and an inner band that assist in maintaining axial and radial positioning of the vanes and define a flowpath for the combustion gases.


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    Downstream from the nozzle stage is another rotor stage that includes a plurality of circumferentially spaced rotor blades that extend radially outward from a rotor disk and are surrounded by an annular shroud which also defines a flowpath for the combustion gases. Adjacent ends of the turbine nozzle outer band and rotor shroud are spaced apart by a gap to facilitate assembly as well as to accommodate differential thermal expansion and contraction that occurs during operation of the engine. The gap, however, also is a potential leakage path for compressed air.

    Particularly, a portion of the compressed air may be extracted from the high pressure compressor for turbine section cooling, airframe pressurization, anti-icing, and other uses. For example, for turbine section cooling, a portion of the extracted air is channeled through the nozzle vanes. The extracted air is at a higher pressure than the combustion gases and is channeled to a cavity formed between the outer wall of the outer band and the stator casing. The extracted air naturally seeks to move from the cavity to the combustion gas flowpath formed by an inside face of the outer band and the rotor shroud. Therefore, the gap between the outer band and shroud must be properly sealed, otherwise the high pressure extracted air would leak through the gap, or an interface, to the lower pressure flowpath.

    It is known to use a leaf seal to seal the interfaces in turbine engines. An example of a leaf seal is disclosed by U.S. Patent No. 5,118,120 to Drerup et al. For example, i...