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公开(公告)号：US10221862B2
公开(公告)日：2019-03-05
申请号：US14695534
申请日：2015-04-24
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Jesse M. Chandler , Joseph Brent Staubach , Brian D. Merry
Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main 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. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.
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公开(公告)号：US10156206B2
公开(公告)日：2018-12-18
申请号：US14512520
申请日：2014-10-13
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Brian D. Merry
Abstract: A gas turbine engine that has a fan case radially surrounding a fan hub, and plural fan exit guide vanes rotatably connected between the fan hub and the fan case. The fan section also has a thrust reverser opening within the case, axially upstream of the guide vanes. The gas turbine engine also has a core case containing a core engine.
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公开(公告)号：US20180283277A1
公开(公告)日：2018-10-04
申请号：US15943825
申请日：2018-04-03
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu
CPC classification number: F02C7/06 , F01D25/16 , F05D2240/50
Abstract: A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing.
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公开(公告)号：US20180252162A1
公开(公告)日：2018-09-06
申请号：US15970516
申请日：2018-05-03
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Nathan Snape , Gabriel L. Suciu , Brian D. Merry , James D. Hill , William K. Ackermann
IPC: F02C7/12
CPC classification number: F02C7/12 , F05D2260/213
Abstract: The present disclosure provides systems and methods related to thermal management systems for gas turbine engines. For example, a thermal management system comprises a thermally neutral heat transfer fluid circuit, a first heat exchanger disposed on the fluid circuit, and a second heat exchanger disposed on the fluid circuit.
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公开(公告)号：US20180223668A1
公开(公告)日：2018-08-09
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
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.
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Patent Agency Ranking
公开(公告)号：US09790792B2
公开(公告)日：2017-10-17
申请号：US14302884
申请日：2014-06-12
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Gabriel L. Suciu , Stephen P. Muron , Christopher M. Dye , Ioannis Alvanos , Brian D. Merry
CPC classification number: F01D5/02 , F01D5/022 , F01D5/025 , F01D5/027 , F01D5/066 , F01D5/14 , F01D5/147 , F01D11/006 , F05D2240/20 , F05D2260/15 , Y10T29/49316 , Y10T29/4932
Abstract: A spool for a gas turbine engine includes at least one rotor disk defined along an axis of rotation and at least one rotor ring defined along the axis of rotation, with the rotor ring being in contact with the rotor disk. The rotor disk and rotor ring are contoured to define a smooth rotor stack load path.
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公开(公告)号：US20170254219A1
公开(公告)日：2017-09-07
申请号：US15062442
申请日：2016-03-07
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Jesse M. Chandler , Brian D. Merry
Abstract: A gas turbine engine comprises an engine having a compressor section, and a turbine section. A firewall and accessory pumps are mounted on a downstream side of the firewall. The accessory pumps are driven by electric motors mounted on the firewall on an upstream side of the firewall.
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公开(公告)号：US09726112B2
公开(公告)日：2017-08-08
申请号：US14190144
申请日：2014-02-26
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Brian D. Merry
IPC: F02K3/06
CPC classification number: F02K3/06 , F05D2210/40 , F05D2250/314 , Y02T50/671
Abstract: A gas turbine engine has a propulsor including a fan and a power turbine, an engine core aerodynamically connected to the propulsor by a transition duct, and a bypass valve in the transition duct that allows for air from the engine core to bypass the power turbine.
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公开(公告)号：US20170167384A1
公开(公告)日：2017-06-15
申请号：US14967518
申请日：2015-12-14
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Brian D. Merry , Jesse M. Chandler , William K. Ackermann , Matthew R. Feulner , Om P. Sharma
CPC classification number: F02C7/18 , F01D5/08 , F02C3/04 , F02C7/143 , F02C7/185 , F02C9/18 , F02K3/06 , F04D27/0215 , F04D29/582 , F05D2220/32 , F05D2220/3219 , F05D2240/24 , F05D2260/205 , F05D2260/211 , F05D2260/213 , Y02T50/676
Abstract: A compressor section for use in a gas turbine engine comprises a compressor rotor having a hub and a plurality of blades extending radially outwardly from the hub and an outer housing surrounding an outer periphery of the blades. A tap taps air at a radially outer first location, passing the tapped air through a heat exchanger, and returning the tapped air to an outlet at a second location which is radially inward of the first location, to provide cooling air adjacent to the hub. A gas turbine engine is also disclosed.
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公开(公告)号：US20170122220A1
公开(公告)日：2017-05-04
申请号：US15411173
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 , F02C9/18 , F01D25/24 , F01D5/06 , F01D9/02 , F01D21/00 , F02C3/06 , F02K3/06
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, 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 at cruise power, and a low fan pressure ratio of less than 1.45 measured across a fan blade alone. A gear arrangement drives the fan section. A compressor section includes both a low pressure compressor and a high pressure compressor. A guide vane includes a forward attachment, the forward attachment positioned aft of a plumbing connection area. A turbine section drives the gear arrangement, and may have a low pressure turbine with a low pressure turbine pressure ratio greater than 5:1, and a two stage high pressure turbine. 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.
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