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：US11280271B2

公开(公告)日：2022-03-22

申请号：US14433444

申请日：2013-10-03

Applicant: United Technologies Corporation

Inventor： Glenn Levasseur ,  Gabriel L. Suciu ,  Gary M. Stetson

IPC: F02C7/36 ,  F02C9/20 ,  F02C3/06 ,  F01D17/16 ,  F02C7/32

Abstract: A gas turbine engine has a fan which includes a plurality of fan blades that are rotatable about an axis, and a compressor section, where the combustor section includes a first compressor and a second compressor aft of the first compressor. At least one first variable guide vane controls operation of the first compressor and at least one second variable guide vane controls operation of the second compressor. A combustor is in fluid communication with the compressor section and a turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

公开(公告)号：US11130580B2

公开(公告)日：2021-09-28

申请号：US15809244

申请日：2017-11-10

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Brian Merry ,  Stephen H. Taylor ,  Charles E. Lents

IPC: F02C6/08 ,  B64D13/08 ,  F02C9/18 ,  F02K3/115 ,  B64D27/10 ,  F01D15/10 ,  F02C3/04 ,  F02K3/06 ,  H02K7/18 ,  F02C7/18 ,  F02K3/04 ,  B64D13/06

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

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：US10808608B2

公开(公告)日：2020-10-20

申请号：US16027496

申请日：2018-07-05

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Nathan Snape ,  Christopher M. Dye

IPC: F02C7/12 ,  F02C3/04 ,  F01D25/12 ,  F02C7/14 ,  F02C7/06 ,  F01D25/18 ,  F02C7/32

Abstract: A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment.

公开(公告)号：US10794293B2

公开(公告)日：2020-10-06

申请号：US16025094

申请日：2018-07-02

Applicant: United Technologies Corporation

Inventor： Paul R. Adams ,  Shankar S. Magge ,  Joseph B. Staubach ,  Wesley K. Lord ,  Frederick M. Schwarz ,  Gabriel L. Suciu

IPC: F02C7/36 ,  F02C3/107 ,  F02C9/18 ,  F02K3/06 ,  F02K3/075 ,  F01D5/06 ,  F01D25/24 ,  F02C3/04 ,  F02C7/20 ,  F04D19/02 ,  F01D11/12

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

公开(公告)号：US10794273B2

公开(公告)日：2020-10-06

申请号：US14789036

申请日：2015-07-01

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Jeffery F. Perlak ,  Joseph B. Staubach ,  Gabriel L. Suciu ,  James D. Hill ,  Frederick M. Schwarz

IPC: F02C3/107 ,  F02C3/14 ,  F02C3/10

Abstract: A gas turbine engine according to the present disclosure includes a first compressor and a first turbine for driving the first compressor. A core section includes a second compressor and a second turbine for driving the second compressor. A third turbine is arranged fluidly downstream of the first turbine and the second turbine and configured to drive a power take-off. A first duct system is arranged fluidly between the low-pressure compressor and the core section. The first duct system is arranged to reverse fluid flow before entry into the core section.

公开(公告)号：US10794214B2

公开(公告)日：2020-10-06

申请号：US15589009

申请日：2017-05-08

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Mark F. Zelesky ,  Ioannis Alvanos ,  Brian Merry

IPC: F01D11/24 ,  F01D25/24

Abstract: A clearance control system for a gas turbine engine comprises at least one case support associated with an engine case defining an engine center axis. A clearance control ring is positioned adjacent the at least one case support to form an internal cavity between the engine case and the clearance control ring. The clearance control ring includes a first mount feature. An outer air seal has a second mount feature cooperating with the first mount feature such that the clearance control ring can move independently of the engine case in response to changes in temperature. An injection source inject flow into the internal cavity to control a temperature of the clearance control ring to allow the outer air seal to move in a desired direction to maintain a desired clearance between the outer air seal and an engine component. A gas turbine engine and a method of controlling tip clearance in a gas turbine engine are 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.