公开(公告)号：US20190383157A1

公开(公告)日：2019-12-19

申请号：US16012059

申请日：2018-06-19

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Frederick M. Schwarz ,  William G. Sheridan

IPC: F01D15/10 ,  F02C7/36 ,  F02K3/06 ,  H02K11/00 ,  H02K7/18

Abstract: A gas turbine engine according to an exemplary embodiment of this disclosure includes, among other things, a fan driven by a fan shaft rotatable about an engine axis; a fan drive electric motor providing a supplemental drive input to the fan; a turbine section driving an input shaft; a geared architecture driven by the input shaft of the turbine section and coupled to the fan shaft to provide a main drive input for driving the fan; and a generator including a rotor supported on the input shaft and a stator disposed on a static structure relative to the rotor, wherein the generator communicates electric power to power the fan drive electric motor.

公开(公告)号：US10364749B2

公开(公告)日：2019-07-30

申请号：US14974779

申请日：2015-12-18

Applicant: United Technologies Corporation

Inventor： Paul R. Hanrahan ,  Daniel Bernard Kupratis

IPC: F02C7/18 ,  F02C7/141 ,  F02C3/10 ,  F02K3/06 ,  F01D25/16 ,  F02C6/08 ,  F02C7/12

Abstract: A heat exchanger for a gas turbine engine includes a cooling air scoop that has a leading edge nose that is positioned adjacent to a downstream-most stator cascade of a fan section of the gas turbine engine. The cooling air scoop subtends only a sector of a circumference of the gas turbine engine. The heat exchanger also includes a cold side that is connected to the cooling air scoop and a hot side that is connected to a compressor section of the gas turbine engine.

公开(公告)号：US10364041B2

公开(公告)日：2019-07-30

申请号：US15218915

申请日：2016-07-25

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Arthur W. Utay ,  Tania Bhatia Kashyap ,  Thomas N. Slavens ,  Kevin L. Rugg ,  Mark F. Zelesky ,  Brian D. Merry ,  Gabriel L. Suciu

IPC: B64D41/00 ,  F04D17/10 ,  F02C3/08 ,  F02C3/14

Abstract: An auxiliary power unit may comprise a twin centrifugal compressor including a first blade. A turbine may be disposed aft of the twin centrifugal compressor. The turbine may include a second blade. The first blade comprises a first material and the second blade comprises a second material. The first material may the same as the second material. The twin centrifugal compressor may include forward centrifugal compressor and an aft centrifugal compressor disposed aft of the forward centrifugal compressor.

公开(公告)号：US10288010B2

公开(公告)日：2019-05-14

申请号：US15405498

申请日：2017-01-13

Applicant: United Technologies Corporation

Inventor： David P. Houston ,  Daniel Bernard Kupratis ,  Frederick M. Schwarz

IPC: F02K3/06 ,  F02K3/072 ,  F02C7/36 ,  F01D25/16 ,  F01D5/02 ,  F02C7/06 ,  F01D5/06 ,  F01D9/02 ,  F02C3/04 ,  F02C3/107 ,  F02C9/18 ,  F02K1/78 ,  F04D27/00 ,  F04D29/32 ,  F01D9/04

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.

公开(公告)号：US10253694B2

公开(公告)日：2019-04-09

申请号：US14635366

申请日：2015-03-02

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Anthony R. Bifulco

IPC: F02C7/143 ,  F01D9/04 ,  F02C7/12 ,  F02C7/18 ,  F02C9/18 ,  F02C3/13 ,  F02C3/32 ,  F02C7/10

Abstract: A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages.

公开(公告)号：US20190093553A1

公开(公告)日：2019-03-28

申请号：US16142337

申请日：2018-09-26

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Daniel Bernard Kupratis ,  Francis R. Moon

IPC: F02C3/14 ,  F23G7/06 ,  F23R7/00 ,  F02C3/107 ,  F02K3/11 ,  F02C7/36 ,  F02C7/06 ,  F02C5/02 ,  F02K3/06

Abstract: The engine (10) includes a low spool (16) disposed aft of an air inlet (12) and a high spool (34) disposed aft of the low spool (16). An intake reverse-duct (44) is disposed radially outward of the high spool (34) and reverses direction of low pressure compressed air from the low spool (16) into a forward-flow high pressure compressor (40) of the high spool (34). A discharge reverse-manifold (48) directs flow of an exhaust gas stream (50) from a forward-flow low pressure turbine (20) into a rearward-flow direction and into at least one pulse detonation firing tube (54). An annular bypass air duct (72) directs cooling air along the engine (10)—The at least, one firing tube is positioned radially outward of the high spool (34), overlies the high spool (34) and is also positioned within the bypass air duct (72).

公开(公告)号：US20190024584A1

公开(公告)日：2019-01-24

申请号：US16143684

申请日：2018-09-27

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Francis R. Moon

IPC: F02C6/02 ,  F01D17/14 ,  F02C3/05 ,  F02C3/10 ,  F02C3/107 ,  F02C3/13 ,  F02C3/14 ,  F02C7/36 ,  F02C9/18 ,  F02K3/06 ,  F02K3/065 ,  F02K3/075

Abstract: A gas turbine engine comprises a fan at an axially outer location, the fan rotating about an axis of rotation, delivering air into an outer bypass duct, a radially middle duct, and a radially inner core duct. Air from the inner core duct is directed into a compressor, and then flows axially in a direction back toward the fan through a combustor section, and across a core turbine section, and is then directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method of operating a gas turbine engine is also disclosed.

公开(公告)号：US10174680B2

公开(公告)日：2019-01-08

申请号：US14687995

申请日：2015-04-16

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis

IPC: F02C7/36 ,  F02C3/10

Abstract: A gas turbine engine comprises a plurality of distributed fan rotors. A gas generator has a core fan rotor. at least one compressor rotor. at least one gas generator turbine rotor, and a combustion section. A fan drive turbine is downstream of the at least one gas generator turbine rotor. A shaft is configured to be driven by the fan drive turbine, the shaft engaging gears to drive the plurality of distributed fan rotors. The core fan rotor delivers a portion of air into the at least one compressor rotor, and a portion of bypass air into a bypass duct which bypasses the gas generator. The bypass air mixes with products of combustion downstream of the at least one gas generator turbine rotor.

公开(公告)号：US20180320591A1

公开(公告)日：2018-11-08

申请号：US16030893

申请日：2018-07-10

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis

IPC: F02C6/02 ,  B64C27/14 ,  F02C7/32 ,  F02C3/14 ,  F02C6/18 ,  F02C7/36 ,  F01D15/12 ,  F01D13/00 ,  B64D35/08 ,  F02C7/10

CPC classification number: F02C6/02 ,  B64C27/14 ,  B64D35/08 ,  F01D13/003 ,  F01D15/12 ,  F02C3/145 ,  F02C6/18 ,  F02C7/10 ,  F02C7/32 ,  F02C7/36 ,  F05D2220/329 ,  F05D2260/40311 ,  F05D2260/532

Abstract: A drive architecture comprises a rotor and a gearbox for driving the rotor. A pair of input gears provides rotational drive to the gearbox. A first recuperative cycle engine drives one of the pair of gears and a second engine drives the other of the pair of gears. The first recuperative cycle engine and the second engine are both gas turbine engines. A power takeoff from a drive shaft of the second engine supplies rotational drive to drive at least one component in the first recuperative cycle drive.

公开(公告)号：US20180306116A1

公开(公告)日：2018-10-25

申请号：US16017042

申请日：2018-06-25

Applicant: United Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Paul R. Hanrahan

IPC: F02C7/18 ,  F02C9/18 ,  F02C7/143 ,  F02C6/08

CPC classification number: F02C7/185 ,  F02C6/08 ,  F02C7/143 ,  F02C9/18 ,  F05D2240/40 ,  F05D2260/20 ,  F05D2260/211 ,  Y02T50/675

Abstract: A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section via a primary flowpath and a turbine section fluidly connected to the combustor via the primary flowpath. Also included is a cascading cooling system having a first inlet connected to a first compressor bleed, a second inlet connected to a second compressor bleed downstream of the first compressor bleed, and a third inlet connected to a third compressor bleed downstream of the second compressor bleed. The cascading cooling system includes at least one heat exchanger configured to incrementally generate cooling air for at least one of an aft compressor stage and a foremost turbine stage relative to fluid flow through the turbine section.