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公开(公告)号：US20130192199A1
公开(公告)日：2013-08-01
申请号：US13762970
申请日：2013-02-08
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu
CPC classification number: F01D25/16 , F01D25/162 , F02C7/06 , F02C7/36 , F02K3/04 , F05D2250/19 , F05D2260/40311
Abstract: A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft.
Abstract translation: 燃气涡轮发动机包括具有入口壳体和中间壳体的芯壳体，其分别提供入口壳体流动路径和中间壳体流动路径。 轴支撑轴向设置在入口壳体流动路径和中间壳体流动路径之间的压缩机部分。 轴包括主轴和具有波纹管的挠性轴。 柔性轴在第一端处固定到主轴，并且包括与第一端相对的第二端。 齿轮架构耦合到轴，并且风扇耦合到由齿轮架构旋转地驱动。 齿轮架构包括在第二端支撑的太阳齿轮。 第一轴承相对于中间壳体支撑轴，第二轴承相对于入口壳体支撑轴。 第二轴承从柔性轴径向向外设置。
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公开(公告)号：US10830152B2
公开(公告)日：2020-11-10
申请号：US15184253
申请日：2016-06-16
Applicant: United Technologies Corporation
Inventor： Karl L. Hasel , Joseph B. Staubach , Brian D. Merry , Gabriel L. Suciu , Christopher M. Dye
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 or equal to about 8 at cruise power. A gear arrangement is configured to drive the fan section. A compressor section includes both a first compressor section and a second compressor section. A turbine section is configured to drive the gear arrangement, and may have a low pressure turbine with four stages and a low pressure turbine pressure ratio greater than about 5:1, and a high pressure turbine with two stages. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor section and a pressure ratio across the second compressor section, and greater than about 40, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the second compressor section is greater than about 7.
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公开(公告)号：US20200025105A1
公开(公告)日：2020-01-23
申请号：US16374995
申请日：2019-04-04
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu , Michael G. McCaffrey
Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.
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公开(公告)号：US10480419B2
公开(公告)日：2019-11-19
申请号：US14745567
申请日：2015-06-22
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, passing the tapped air through a first heat exchanger and then to a cooling compressor. A second tap taps air from a location closer to the downstream most end than the location of the first tap, and the first and second taps mix together and are delivered into the high pressure turbine. The cooling compressor is positioned downstream of the first heat exchanger, and upstream of a location where air from the first and second taps mix together.
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公开(公告)号：US10352241B2
公开(公告)日：2019-07-16
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
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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Patent Agency Ranking
公开(公告)号：US10287905B2
公开(公告)日：2019-05-14
申请号：US15036019
申请日：2014-11-11
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： James D. Hill , Gabriel L Suciu , Brian D. Merry , Ioannis Alvanos
Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine. The engine includes a first rotor disk, a second rotor disk, and a circumferentially segmented seal. The segmented seal engages the first rotor disk and the second rotor disk. The segmented seal further includes a fore surface contacting the first disk, an aft surface contacting the second disk, and a radially outer surface. Further, (1) the aft surface and (2) one of the fore surface and the radially outer surface include perforations to allow fluid to flow through the interior of the segmented seal.
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公开(公告)号：US20190107055A1
公开(公告)日：2019-04-11
申请号：US16158581
申请日：2018-10-12
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu , William K. Ackermann
Abstract: A gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row.
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公开(公告)号：US20190107051A1
公开(公告)日：2019-04-11
申请号：US16152502
申请日：2018-10-05
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu , Karl L. Hasel
Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.
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公开(公告)号：US20180258860A1
公开(公告)日：2018-09-13
申请号：US15978517
申请日：2018-05-14
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Jesse M. Chandler , Brian D. Merry , Nathan Snape
CPC classification number: F02C7/185 , F02C3/04 , F02C7/32 , F02C9/18 , F05D2220/323 , F05D2240/90 , F05D2260/213 , F05D2260/4031
Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A tap connects to the compressor section. A heat exchanger connects downstream of the tap. A cooling compressor connects downstream of the heat exchanger, and the cooling compressor connects to deliver air to at least one of the rotating components. A core housing has an outer peripheral surface and a fan housing defines an inner peripheral surface. At least one bifurcation duct extends between the outer peripheral surface to the inner peripheral surface. The heat exchanger is disposed within the at least one bifurcation duct.
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公开(公告)号：US09915204B2
公开(公告)日：2018-03-13
申请号：US14676513
申请日：2015-04-01
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Brian D. Merry , James D. Hill
CPC classification number: F02C7/18 , F01D5/088 , F01D11/001 , F01D25/24
Abstract: Systems and methods are disclosed herein for distributing cooling air in gas turbine engines. A tangential on board injector (“TOBI”) may supply cooling air to a turbine section of a gas turbine engine. The cooling air may be split into a first cooling air path and a second cooling air path. The first cooling air path may fluidly connect the TOBI and the interior of a first stage rotor blade. The second cooling air path may fluidly connect the TOBI and a cavity. The cavity may be located between a first disk and a second disk. The cooling air paths from a single cooling air source may thermally isolate portions of the turbine section.
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