公开(公告)号：US10415466B2

公开(公告)日：2019-09-17

申请号：US14921575

申请日：2015-10-23

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  James D. Hill

IPC: F02C3/14 ,  F02C3/04 ,  F02K3/00 ,  F01D25/24 ,  F01D25/30 ,  F02K3/04 ,  F02C3/10

Abstract: A gas turbine engine includes a propulsor with a power turbine, a power turbine shaft extending forward therefrom defining a centerline axis, and a fan driven by the power turbine shaft. The fan is aligned with the centerline axis forward of the power turbine and is operatively connected to be driven by the power turbine through the power turbine shaft. A gas generator operatively connected to the propulsor is included downstream from the fan and forward of the power turbine, wherein the gas generator defines a generator axis offset from the centerline axis. The gas generator is operatively connected to the power turbine to supply combustion products for driving the power turbine.

公开(公告)号：US10371055B2

公开(公告)日：2019-08-06

申请号：US14837009

申请日：2015-08-27

Applicant: United Technologies Corporation

Inventor： Nathan Snape ,  Gabriel L. Suciu ,  Brian D. Merry

IPC: F02C7/18 ,  F02K3/075 ,  F02C7/277 ,  F02C6/08 ,  F02C7/32 ,  F02C3/04 ,  F02K3/02 ,  F02K3/065

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.

公开(公告)号：US10337406B2

公开(公告)日：2019-07-02

申请号：US14771047

申请日：2014-02-28

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  William K. Ackermann

IPC: F01D5/02 ,  F01D9/04 ,  F01D9/06 ,  F02C3/04 ,  F02C6/08 ,  F02C7/04 ,  F02C7/06 ,  F02C7/14 ,  F02C7/18 ,  F01D17/10 ,  F01D25/12 ,  F01D25/24 ,  F04D29/54

Abstract: A gas turbine engine that includes a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and a bearing compartment.

公开(公告)号：US20190186381A1

公开(公告)日：2019-06-20

申请号：US16284671

申请日：2019-02-25

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler

IPC: F02C9/18 ,  F01D25/16 ,  F04D29/52 ,  F04D29/54 ,  F04D29/70 ,  F02C7/25 ,  F02C6/08 ,  F01D25/24 ,  F04D29/58

CPC classification number: F02C9/18 ,  F01D25/162 ,  F01D25/24 ,  F02C6/08 ,  F02C7/052 ,  F02C7/25 ,  F04D29/522 ,  F04D29/542 ,  F04D29/5853 ,  F04D29/701 ,  F05D2220/32 ,  F05D2240/15

Abstract: A compressor intermediate case for a gas turbine engine includes a plurality of intermediate case struts joining the compressor intermediate case to an inner engine structure. Each strut of the plurality of intermediate case struts includes a leading edge. A turning scoop is disposed at the leading edge of each strut of the plurality of intermediate case struts. A plurality of diffusers extends radially outwardly from the compressor intermediate case so that each diffuser of the plurality of diffusers engages with a corresponding turning scoop. A substantially annular structural fire wall extends radially outwardly from the compressor intermediate case. An environmental control system manifold is disposed on the compressor intermediate case. The environmental control system manifold includes an exit port.

公开(公告)号：US20190153866A1

公开(公告)日：2019-05-23

申请号：US15819047

申请日：2017-11-21

Applicant: United Technologies Corporation

Inventor： Dmytro Mykolayovych Voytovych ,  Alexander Staroselsky ,  Om P. Sharma ,  Gabriel L. Suciu ,  Brian Merry ,  Ioannis Alvanos

IPC: F01D5/02 ,  F01D5/22 ,  F04D29/32 ,  F01D5/34

Abstract: An exemplary gas turbine engine includes a turbine section and a fan mechanically connected to the turbine section such that rotation of the turbine drives rotation of the fan. The fan includes a hub, a plurality of blade bodies extending radially outward from the hub to a first partial shroud, and a plurality of blade tips extending radially outward from the partial shroud.

公开(公告)号：US10240526B2

公开(公告)日：2019-03-26

申请号：US15682848

申请日：2017-08-22

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Frederick M. Schwarz ,  William K. Ackermann ,  Daniel Bernard Kupratis

IPC: F02C3/113 ,  F02C7/36 ,  F02C9/16 ,  F02K3/06 ,  F02C3/06 ,  F02K3/072

Abstract: A compressor section is in fluid communication with a fan, which includes a first compressor section and a second compressor section. A turbine section includes a first turbine section driving the fan and the first compressor section and a second turbine section driving the second compressor section and the second compressor rotor. A first performance quantity is defined as a product of the first speed squared and the first area. A second performance quantity is defined as a product of the second speed squared and the second exit area. A performance ratio of the first performance quantity to the second performance quantity is between about 0.2 and about 0.8. A gear reduction is included between the first turbine section and the first compressor section.

公开(公告)号：US20190063325A1

公开(公告)日：2019-02-28

申请号：US16059415

申请日：2018-08-09

Applicant: United Technologies Corporation

Inventor： Nathan Snape ,  Gabriel L. Suciu ,  Brian D. Merry

IPC: F02C7/18 ,  F02K3/075 ,  F02K3/065 ,  F02K3/02 ,  F02C6/08 ,  F02C7/32 ,  F02C7/277 ,  F02C3/04

CPC classification number: F02C7/185 ,  F02C3/04 ,  F02C6/08 ,  F02C7/277 ,  F02C7/32 ,  F02K3/025 ,  F02K3/065 ,  F02K3/075 ,  F05D2260/213 ,  Y02T50/671 ,  Y02T50/676

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.

公开(公告)号：US10215094B2

公开(公告)日：2019-02-26

申请号：US14821409

申请日：2015-08-07

Applicant: United Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Karl L. Hasel

IPC: F02C1/00 ,  F02C7/06 ,  F02C7/36 ,  F02C3/04

Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.

公开(公告)号：US20190032518A1

公开(公告)日：2019-01-31

申请号：US15662318

申请日：2017-07-28

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler

IPC: F01D25/28 ,  F02K3/06 ,  B64D27/12 ,  B64D27/26

Abstract: A gas turbine engine according to an example of the present disclosure includes a mount arrangement including a gas turbine engine, an adapter ring extending circumferentially around the gas turbine engine and mounted to both the gas turbine engine and to an aircraft beneath the gas turbine engine, and at least one thrust link including a first end mounted to the gas turbine engine fore of the adapter ring, and an opposite second end mounted to the adapter ring above the gas turbine engine.

公开(公告)号：US20190003391A1

公开(公告)日：2019-01-03

申请号：US16111870

申请日：2018-08-24

Applicant: United Technologies Corporation

Inventor： Nathan Snape ,  James D. Hill ,  Gabriel L. Suciu ,  Brian Merry

IPC: F02C7/14

Abstract: An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.