公开(公告)号：US11585276B2

公开(公告)日：2023-02-21

申请号：US16598048

申请日：2019-10-10

Applicant: United Technologies Corporation

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

IPC: F02C7/36 ,  F02K3/06 ,  F02C3/107 ,  F01D25/16 ,  F01D5/06 ,  F02C3/06 ,  F02C7/06 ,  F04D29/32

Abstract: A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section.

公开(公告)号：US11486269B2

公开(公告)日：2022-11-01

申请号：US16286970

申请日：2019-02-27

Applicant: United Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Todd A. Davis ,  Gregory E. Reinhardt ,  Enzo DiBenedetto

IPC: F01D25/16 ,  F02C7/36 ,  F02K3/04 ,  F02C7/06 ,  F01D15/12 ,  F01D5/02

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing.

公开(公告)号：US10830149B2

公开(公告)日：2020-11-10

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 ,  F02C6/08 ,  F02C7/277 ,  F02C7/32 ,  F02C3/04 ,  F02K3/02 ,  F02K3/075

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：US10760423B2

公开(公告)日：2020-09-01

申请号：US15947119

申请日：2018-04-06

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Stephen P. Muron ,  Ioannis Alvanos ,  Christopher M. Dye ,  Brian D. Merry ,  Arthur M. Salve, Jr. ,  James W. Norris

IPC: F01D5/02 ,  F01D5/06

Abstract: A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk at an interface, where the interface is defined along a spoke. A spool for a gas turbine engine includes the rotor disk, the plurality of blades with the interface defined along the spoke radially inboard of a blade platform, a rotor ring axially adjacent to the rotor disk, and a plurality of core gas path seals which extend from the rotor ring. Each of the plurality of core gas path seals extends from the rotor ring at a seal interface, with the seal interface defined along a spoke and the plurality of core gas path seals being axially adjacent to the blade platform.

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：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：US10082034B2

公开(公告)日：2018-09-25

申请号：US14755952

申请日：2015-06-30

Applicant: United Technologies Corporation

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

IPC: F01D5/34 ,  F01D5/22 ,  F01D5/02

CPC classification number: F01D5/225 ,  B23P15/006 ,  F01D5/02 ,  F01D5/10 ,  F01D5/34 ,  F05D2220/32 ,  F05D2240/24 ,  F05D2260/941

Abstract: A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk, adjacent ones of the plurality of blades are joined by a flexible web.

公开(公告)号：US20180230909A1

公开(公告)日：2018-08-16

申请号：US15947890

申请日：2018-04-09

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Ioannis Alvanos ,  Brian D. Merry

IPC: F02C7/20 ,  F02K1/04 ,  F01D9/04 ,  F01D25/30

CPC classification number: F02C7/20 ,  F01D9/041 ,  F01D25/30 ,  F02K1/04 ,  F05D2300/6033 ,  Y02T50/672 ,  Y02T50/673 ,  Y10T29/4932

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a turbine case extending along a turbine axis and a CMC turbine exhaust case mounted to the turbine case. The CMC turbine exhaust case includes a CMC core nacelle aft portion, a CMC tail cone connected to the CMC core nacelle aft portion, and a multiple of CMC turbine exhaust case struts extending between the CMC core nacelle aft portion and the CMC tail cone. The CMC core nacelle aft portion, the CMC tail cone and the turbine case are arranged along the turbine axis. The CMC turbine exhaust case is mounted to the turbine case at a flange such that the CMC tail cone is axially spaced apart from the turbine case relative to the turbine axis.

公开(公告)号：US20180187605A1

公开(公告)日：2018-07-05

申请号：US15908475

申请日：2018-02-28

Applicant: United Technologies Corporation

Inventor： Michael E. McCune ,  Brian D. Merry ,  Gabriel L. Suciu

IPC: F02C7/36 ,  F02C3/107 ,  F01D25/16 ,  F01D21/04 ,  F02K3/075 ,  F02K3/02 ,  F02C3/06 ,  F01D21/02

CPC classification number: F02C7/36 ,  F01D21/02 ,  F01D21/04 ,  F01D25/16 ,  F02C3/06 ,  F02C3/107 ,  F02K3/025 ,  F02K3/075 ,  F05D2270/091 ,  Y02T50/671

Abstract: A turbofan engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path, and a speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan.