公开(公告)号：US20210222619A1

公开(公告)日：2021-07-22

申请号：US16746299

申请日：2020-01-17

Applicant: United Technologies Corporation

Inventor： Amanda J. L. Boucher ,  Joseph B. Staubach

IPC: F02C6/18 ,  F02C1/05

Abstract: Gas turbine engines are described. The gas turbine engines include a compressor section, a combustor section, a turbine section, a nozzle section, wherein the compressor section, the combustor section, the turbine section, and the nozzle section define a core flow path that expels through the nozzle section, and a waste heat recovery system. The waste heat recovery system includes a heat recovery heat exchanger arranged at the nozzle section, wherein the heat recovery heat exchanger is arranged within the nozzle section such that the heat recovery heat exchanger occupies less than an entire area of an exhaust area of the nozzle section and a heat rejection heat exchanger arranged to reduce a temperature of a working fluid of the waste heat recovery system.

公开(公告)号：US20210222618A1

公开(公告)日：2021-07-22

申请号：US16746284

申请日：2020-01-17

Applicant: United Technologies Corporation

Inventor： Joseph B. Staubach ,  Amanda J. L. Boucher

IPC: F02C6/18 ,  F02C7/141

Abstract: Gas turbine engines are described. The gas turbine engines includes a compressor section, a combustor section, a turbine section, and a nozzle section. The compressor section, the combustor section, the turbine section, and the nozzle section define a core flow path that expels through the nozzle section. A cooling duct is provided that is separate from the core flow path. A waste heat recovery system is arranged with a heat rejection heat exchanger arranged within the cooling duct and a blower is arranged within the cooling duct and configured to generate a pressure drop across the heat rejection heat exchanger.

公开(公告)号：US20210207530A1

公开(公告)日：2021-07-08

申请号：US16733504

申请日：2020-01-03

Applicant: United Technologies Corporation

Inventor： Marc J. Muldoon ,  Joseph B. Staubach ,  Jesse M. Chandler ,  Neil Terwilliger ,  Gabriel L. Suciu

IPC: F02C6/20

Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

公开(公告)号：US20200224589A1

公开(公告)日：2020-07-16

申请号：US16248884

申请日：2019-01-16

Applicant: United Technologies Corporation

Inventor： Brendan T. McAuliffe ,  Joseph B. Staubach ,  Nagendra Somanath

IPC: F02C6/02 ,  F02K3/06 ,  F02C6/18 ,  F01D13/02

Abstract: A gas turbine engine includes a primary flowpath fluidly connecting a compressor section, a combustor section, and a turbine section. A heat exchanger is disposed in the primary flowpath downstream of the turbine section. The heat exchanger includes a first inlet for receiving fluid from the primary flowpath and a first outlet for expelling fluid received at the first inlet. The heat exchanger further includes a second inlet fluidly connected to a supercharged CO2 (sCO2) bottoming cycle and a second outlet connected to the sCO2 bottoming cycle. The sCO2 bottoming cycle is an overexpanded, recuperated Brayton cycle.

公开(公告)号：US10550768B2

公开(公告)日：2020-02-04

申请号：US15346206

申请日：2016-11-08

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Joseph B. Staubach ,  Nathan Snape

IPC: F02C7/143 ,  F02C7/18 ,  B64D13/06 ,  F02C6/08

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage is fluidly coupled to a bleed port of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system includes a pump fluidly coupled to the first cooling stage to receive and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air. The intercooled cooling system includes an air cycle machine in fluid communication to outputs of the cooling stages to selectively receive the cool bleed air or the intercooled cooling air.

公开(公告)号：US20190017445A1

公开(公告)日：2019-01-17

申请号：US16025022

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

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

公开(公告)号：US20180230912A1

公开(公告)日：2018-08-16

申请号：US15941240

申请日：2018-03-30

Applicant: United Technologies Corporation

Inventor： Karl L. Hasel ,  Joseph B. Staubach ,  Brian D. Merry ,  Gabriel L. Suciu ,  Christopher M. Dye

IPC: F02C7/36 ,  F02K3/06 ,  F02C3/107 ,  F01D5/12 ,  F01D9/02 ,  F01D25/24 ,  F01D25/28 ,  F02C3/04 ,  F02C7/28 ,  F02K3/02 ,  F02C7/06

CPC classification number: F02C7/36 ,  F01D5/12 ,  F01D9/02 ,  F01D25/24 ,  F01D25/28 ,  F02C3/04 ,  F02C3/107 ,  F02C7/06 ,  F02C7/28 ,  F02K3/025 ,  F02K3/06 ,  F05D2220/32 ,  F05D2240/12 ,  F05D2240/55 ,  F05D2260/40 ,  F05D2260/4031 ,  F05D2260/40311

Abstract: A gas turbine engine includes, among other things, a fan, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10. A gear arrangement drives the fan. A compressor section includes both a low pressure compressor and a high pressure compressor. A turbine section drives the gear arrangement. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor and a pressure ratio across the high pressure compressor, and greater than 50, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the high pressure compressor is greater than 7.

公开(公告)号：US20180087398A1

公开(公告)日：2018-03-29

申请号：US15278612

申请日：2016-09-28

Applicant: United Technologies Corporation

Inventor： Matthew P. Forcier ,  Joseph B. Staubach

IPC: F01D25/12 ,  F28D15/02 ,  F01D5/08

CPC classification number: F01D25/12 ,  F01D5/08 ,  F02C7/08 ,  F05D2220/36 ,  F05D2260/208 ,  F28D15/02 ,  F28D15/046 ,  F28D2021/0026 ,  F28F21/02 ,  Y02T50/671 ,  Y02T50/676

Abstract: A graphene heat pipe for a gas turbine engine includes a body of graphene. The body has a hot side to accept heat from the gas turbine engine, a cold side to reject heat from the body, and an adiabatic portion to flow heat within the body between the hot side and the cold side.

公开(公告)号：US09650954B2

公开(公告)日：2017-05-16

申请号：US14597510

申请日：2015-01-15

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Michael E. McCune ,  Jesse M. Chandler ,  Alan H. Epstein ,  Steven M. O'Flarity ,  Christopher J. Hanlon ,  William F. Schneider ,  Joseph B. Staubach ,  James A. Kenyon

IPC: F02C3/107 ,  B64D35/04 ,  F01D13/00 ,  F02K3/077 ,  F02C7/36 ,  F02C3/06 ,  F02C3/067 ,  F02C9/18 ,  B64D27/18

CPC classification number: F02C3/107 ,  B64D27/18 ,  B64D35/04 ,  F01D13/003 ,  F02C3/06 ,  F02C3/067 ,  F02C7/36 ,  F02C9/18 ,  F02K3/077 ,  F05D2220/323 ,  F05D2220/36 ,  F05D2240/40 ,  F05D2250/312

Abstract: A gas generator has at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine is positioned downstream of a path of the products of combustion having passed over the at least one gas generator turbine rotor. The fan drive turbine drives a shaft and the shaft engages gears to drive at least three fan rotors.

公开(公告)号：US20170122219A1

公开(公告)日：2017-05-04

申请号：US15411147

申请日：2017-01-20

Applicant: United Technologies Corporation

Inventor： Karl L. Hasel ,  Joseph B. Staubach ,  Brian D. Merry ,  Gabriel L. Suciu ,  Christopher M. Dye

IPC: F02C7/36 ,  F02C7/06 ,  F02K3/06 ,  F02C9/18 ,  F01D21/00 ,  F01D9/02 ,  F01D25/24 ,  F02C3/06 ,  F01D5/06

CPC classification number: F02C7/36 ,  F01D5/06 ,  F01D5/12 ,  F01D9/02 ,  F01D21/003 ,  F01D25/24 ,  F01D25/28 ,  F02C3/04 ,  F02C3/06 ,  F02C3/107 ,  F02C7/06 ,  F02C7/28 ,  F02C9/18 ,  F02K3/025 ,  F02K3/06 ,  F05D2220/32 ,  F05D2220/323 ,  F05D2240/12 ,  F05D2240/35 ,  F05D2240/55 ,  F05D2260/40 ,  F05D2260/4031 ,  F05D2260/40311

Abstract: A gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than 8 at cruise power. A gear arrangement drives the fan section. A guide vane includes a forward attachment, the forward attachment positioned aft of a plumbing connection area. A compressor section includes both a first compressor and a second compressor. A lubrication system and a compressed air system are in fluid communication with the gear arrangement. A turbine section drives the gear arrangement, and includes a low pressure turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, and greater than about 50, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the second compressor is greater than about 7.