公开(公告)号：US20200325781A1

公开(公告)日：2020-10-15

申请号：US16380288

申请日：2019-04-10

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Chao ZHANG ,  Michael PAPPLE ,  Marc TARDIF

IPC: F01D5/18

Abstract: A method of reducing creep in an internally cooled turbine blade, comprising: providing a radially extending intermediate wall to continuously join a localized high stress zone of a concave side wall and a convex side wall in an intermediate cooling air channel through the blade. The intermediate wall distributes stress from the localized zone to a zone of lower stress to balance the creep inducing stress and temperature more evenly.

公开(公告)号：US20160251963A1

公开(公告)日：2016-09-01

申请号：US14633503

申请日：2015-02-27

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Marc TARDIF ,  Domenico DI FLORIO ,  Aldo ABATE

IPC: F01D5/10

CPC classification number: F01D5/10 ,  F01D5/22 ,  F01D5/3007 ,  F05D2220/32 ,  F05D2260/96

Abstract: A rotor blade vibration damper for a gas turbine engine includes an elongated damper body including a top portion extending longitudinally between a front end and a rear end. The top portion has a width defined between spaced apart lateral sides and is substantially flat between the front and rear ends and between the lateral sides such as to define a longitudinal plane within which the top portion lies. A front tab extends downwardly from the front end of the top portion relative to the longitudinal plane. The rear end of the top portion is flat and generally contained in the longitudinal plane. A pair of lateral tabs extends downwardly from each of said lateral sides of the top portion relative to the longitudinal plane.

Abstract translation: 用于燃气涡轮发动机的转子叶片减振器包括细长的阻尼器主体，其包括在前端和后端之间纵向延伸的顶部。 顶部具有限定在间隔开的横向侧面之间的宽度，并且在前端和后端之间以及横向侧之间基本上是平坦的，以便限定顶部部分所在的纵向平面。 前突片相对于纵向平面从顶部的前端向下延伸。 顶部的后端是平的，并且通常包含在纵向平面中。 一对横向突片相对于纵向平面从顶部的所述横向侧边中的每一个向下延伸。

公开(公告)号：US20200166004A1

公开(公告)日：2020-05-28

申请号：US16202788

申请日：2018-11-28

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Assaf FARAH ,  Marc TARDIF ,  Daniel SUMMERS-LEPINE

IPC: F02K1/82 ,  F02K3/06

Abstract: An exhaust casing for a gas turbine engine comprises a shroud configured to surround an exhaust cone, and a heat shield attached to the shroud. The heat shield has a first end and a second end axially spaced apart from each other. A gap is defined radially between the shroud and the heat shield. The gap is configured to enclose at least a portion of an interface defined between at least one strut and the shroud. The exhaust casing is configured to surround the exhaust cone and configured to be connected to the exhaust cone via the at least one strut.

公开(公告)号：US20180258790A1

公开(公告)日：2018-09-13

申请号：US15454589

申请日：2017-03-09

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Kapila JAIN ,  Patricia PHUTTHAVONG ,  Marc TARDIF ,  Ho-Wing LEUNG

IPC: F01D25/12 ,  F01D9/04 ,  F01D25/24 ,  F02C3/04

CPC classification number: F01D25/14 ,  F02C3/107 ,  F02C6/206 ,  F05D2240/35 ,  F05D2260/201

Abstract: The turbine housing assembly includes a housing surrounding a plurality of turbine shroud segments mounted to support members of the housing. Impingement holes extend through the housing and have outlet openings communicating with a cavity between the shroud segments and the support. A deflector rail protrudes axially away from the support members into the cavity. The deflector rail defines a flow-redirecting surface to redirect the cooling air flow from the impingement holes radially outwardly, away from the turbine shroud segments.

公开(公告)号：US20240117748A1

公开(公告)日：2024-04-11

申请号：US17938736

申请日：2022-10-07

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Marc TARDIF ,  Alexandre SEGUIN ,  Sylvain VIGNOLA

IPC: F01D11/00

CPC classification number: F01D11/008 ,  F05D2220/323 ,  F05D2240/55 ,  F05D2240/80

Abstract: A rotor assembly has: blades having airfoils and roots protruding from platform segments; a rotor disc having a peripheral face defining recesses, and slots, a recess located between two adjacent ones of the slots and bounded by a step; feather seals located radially between the peripheral face and the platform segments, a feather seal having a core extending from a trailing end to a leading end and overlapping a gap defined between two platform segments and tabs protruding from the core, the tabs including: trailing tabs positioned axially outside the recess; and leading tabs, a leading tab extending from a root to a tip and having one or more of: the tip axially positioned outside of the recess; and a fillet at an intersection between the tip and an edge of the leading tab, the edge extending between the tip and the core, and facing the step.

公开(公告)号：US20200332669A1

公开(公告)日：2020-10-22

申请号：US16385428

申请日：2019-04-16

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Marc TARDIF ,  Michael PAPPLE ,  Jason GILLESPIE

IPC: F01D11/08 ,  F01D25/12

Abstract: A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.

公开(公告)号：US20190211694A1

公开(公告)日：2019-07-11

申请号：US16352019

申请日：2019-03-13

Applicant: PRATT & WHITNEY CANADA CORP.

Inventor： Roger HUPPE ,  Marc TARDIF ,  Herve TURCOTTE

IPC: F01D5/18 ,  F01D25/12 ,  F01D5/08 ,  F02C7/18 ,  F01D9/06 ,  F02C3/04

CPC classification number: F01D5/187 ,  F01D5/081 ,  F01D9/065 ,  F01D25/12 ,  F02C3/04 ,  F02C7/18 ,  F05D2220/32 ,  F05D2260/231

Abstract: An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.