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公开(公告)号：US20180313364A1
公开(公告)日：2018-11-01
申请号：US15499731
申请日：2017-04-27
Applicant: General Electric Company
Inventor： Curtis William Moeckel , Matthew Ford Adam , Rudolf Konrad Selmeier , Michael Anthony Thomas , Steven Mitchell Taylor , Anton Streit
Abstract: A compressor bleed slot apparatus includes: an annular compressor casing; a stator vane row including a plurality of stator vanes disposed inside the compressor casing; a blade row mounted for rotation about a centerline axis inside the compressor casing, axially downstream of the stator row; a bleed slot passing through the compressor casing, the bleed slot having an inlet and an outlet and extending along a slot axis, wherein the bleed slot is bounded by inboard and outboard walls defined within the compressor casing, the inboard and outboard walls diverging from each other in a downstream direction relative to the bleed slot; and a plurality of turning vanes disposed in the slot, the turning vanes configured to reduce a tangential velocity of airflow through the bleed slot.
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公开(公告)号：US10221859B2
公开(公告)日：2019-03-05
申请号：US15018126
申请日：2016-02-08
Applicant: General Electric Company
Inventor： Curtis William Moeckel , Peter John Wood , Matthew Ford Adam , Eric Andrew Falk , Mark Joseph Stecher
IPC: F04D29/32
Abstract: An airfoil for a compressor blade of a gas turbine engine has a chord between a leading edge and a trailing edge and a span between a root and a tip. The airfoil can further include a reduction in local chord from about 75% span to the tip for about a 5% reduction in local solidity. The airfoil can have decreasing sweep angles for the leading edge and the trailing edge from 50% span to the tip and can have decreasing leading and trailing edge dihedral angles from 50% span to the tip.
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公开(公告)号：US20170227016A1
公开(公告)日：2017-08-10
申请号：US15018126
申请日：2016-02-08
Applicant: General Electric Company
Inventor： Curtis William Moeckel , Peter John Wood , Matthew Ford Adam , Eric Andrew Falk , Mark Joseph Stecher
IPC: F04D29/32
CPC classification number: F04D29/324 , F01D5/141 , F05D2240/303 , F05D2240/304 , Y02T50/673
Abstract: An airfoil for a compressor blade of a gas turbine engine has a chord between a leading edge and a trailing edge and a span between a root and a tip. The airfoil can further include a reduction in local chord from about 75% span to the tip for about a 5% reduction in local solidity. The airfoil can have decreasing sweep angles for the leading edge and the trailing edge from 50% span to the tip and can have decreasing leading and trailing edge dihedral angles from 50% span to the tip.
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