公开(公告)号：US10781749B2

公开(公告)日：2020-09-22

申请号：US15371589

申请日：2016-12-07

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler

IPC: F02C6/08 ,  F02K3/06 ,  F02C7/36 ,  F02C3/20 ,  B64D33/02 ,  F02C3/10 ,  B64D27/20 ,  F02K3/062 ,  B64C1/16 ,  B64C21/00

Abstract: A boundary layer ingestion engine includes a gas generator and a turbine fluidly connected to the gas generator. A fan is mechanically linked to the turbine via a shaft such that rotation of the turbine is translated to the fan. A boundary layer ingestion inlet is aligned with an expected boundary layer, such that the boundary layer ingestion inlet is configured to ingest fluid from a boundary layer during operation of the boundary layer ingestion engine.

公开(公告)号：US10760491B2

公开(公告)日：2020-09-01

申请号：US14771055

申请日：2014-02-28

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  William K. Ackermann

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

Abstract: A gas turbine engine is provided 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 in communication with said mid-span pre-diffuser inlet and said combustor section.

公开(公告)号：US10760483B2

公开(公告)日：2020-09-01

申请号：US15784629

申请日：2017-10-16

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Brian Merry ,  Gabriel L. Suciu ,  James W. Norris

IPC: F02C1/00 ,  F02C3/073 ,  F02K1/04 ,  F02K3/068

Abstract: A non-metallic tailcone in a tip turbine engine includes a tapered wall structure disposed about a central axis. The non-metallic tailcone is fastened to a structural frame in the aft portion of the tip turbine engine. The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The second temperature gas stream is a lower temperature than the first temperature gas stream. The second temperature gas stream is discharged at an inner diameter of the tip turbine engine over an outer surface of the tailcone. Discharging the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone.

公开(公告)号：US20200256430A1

公开(公告)日：2020-08-13

申请号：US16274360

申请日：2019-02-13

Applicant: United Technologies Corporation

Inventor： Hung Duong ,  Marc J. Muldoon ,  Gabriel L. Suciu ,  Joseph H. Polly ,  Nicholas D. Leque

IPC: F16H1/22 ,  F02C7/32

Abstract: A gas turbine engine includes a low speed input shaft connected to a low speed tower shaft, in turn connected to a lower speed spool and a high speed input connected to a high speed tower shaft, in turn connected to a higher speed spool. The low speed input shaft is connected to drive a first plurality of accessories. The high speed input shaft is connected to drive a second plurality of accessories. The high speed tower shaft and the low speed tower shaft are coaxial and rotate about a common axis. An angle drive is provided into the low speed input shaft and the high speed input shaft, respectively.

公开(公告)号：US20200256258A1

公开(公告)日：2020-08-13

申请号：US16274340

申请日：2019-02-13

Applicant: United Technologies Corporation

Inventor： Hung Duong ,  Marc J. Muldoon ,  Gabriel L. Suciu

IPC: F02C7/32 ,  F02C3/113 ,  F01D15/10 ,  F02K3/06 ,  F01D15/12

Abstract: A gas turbine engine has a low speed input shaft drives a first plurality of accessories. A high speed input shaft drives a second plurality of accessories. The first plurality of accessories rotating about a first set of rotational axes perpendicular to a first plane. The second plurality of accessories rotating about a second set rotational axes perpendicular to a second plane. The first and second planes extending in opposed directions away from a drive input axis. Compressed air is tapped and passes through a heat exchanger, then to a boost compressor, and then to at least one rotatable components in a main compressor section and a main turbine section. The boost compressor driven on a boost axis, which is non-parallel to the first set of rotational axes and the second set of rotational axes.

公开(公告)号：US10718272B2

公开(公告)日：2020-07-21

申请号：US15490155

申请日：2017-04-18

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Wesley K. Lord ,  Matthew R Feulner ,  Gabriel L. Suciu ,  Jesse M. Chandler

IPC: B64D27/14 ,  F02C9/18 ,  F02K3/06 ,  B64D33/02 ,  B64C21/06 ,  B64C21/02

Abstract: A gas turbine engine comprises a housing having an inlet leading to a fan rotor. A bypass door is mounted upstream of the inlet to the fan rotor, and is moveable away from a non-bypass position to a bypass position to selectively bypass boundary layer air vertically beneath the engine. An aircraft is also disclosed.

公开(公告)号：US10718268B2

公开(公告)日：2020-07-21

申请号：US15807911

申请日：2017-11-09

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler ,  Joseph Brent Staubach ,  Brian D. Merry ,  Wesley K. Lord

IPC: F01D25/12 ,  F02C7/18 ,  F02C7/143 ,  F01D17/10 ,  F02C3/13

Abstract: A gas turbine engine comprises a main compressor section having a downstream most end, 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 cooling air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.

公开(公告)号：US10590852B2

公开(公告)日：2020-03-17

申请号：US15409889

申请日：2017-01-19

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Hung Duong ,  Jonathan F. Zimmitti ,  William G. Sheridan ,  Michael E. McCune ,  Brian Merry

IPC: F02C7/32 ,  F02C7/06 ,  F02C7/36

Abstract: A gas turbine engine assembly includes a turbine section having first and second turbines mounted for rotation about a common rotational axis within an engine static structure, first and second turbine shafts coaxial with one another and to which the first and second turbines are respectively operatively mounted, first and second towershafts respectively coupled to the first and second turbine shafts, an accessory drive gearbox mounted to the engine static structure, and a transmission transitionable between a first mode where an electric machine is driven at a first speed relative to the first towershaft, and a second mode where the electric machine is driven at a different, second speed relative to the first towershaft.

公开(公告)号：US10563585B2

公开(公告)日：2020-02-18

申请号：US15058229

申请日：2016-03-02

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Wesley K. Lord ,  Nathan Snape ,  Gabriel L. Suciu

IPC: F02C7/18 ,  F02C6/08 ,  F02K3/115

Abstract: A gas turbine engine component includes a heat exchange structure having an upstream end and a downstream end. A diffusing duct is associated with the upstream end. A throttle member controls air flow through the heat exchange structure, wherein the throttle member: (a) has a non-circular cross section; and (b) is mounted to the downstream end of the heat exchange structure or is mounted between the upstream end and the diffusing duct.

公开(公告)号：US20200049077A1

公开(公告)日：2020-02-13

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.