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公开(公告)号：US11661894B2
公开(公告)日：2023-05-30
申请号：US17555606
申请日：2021-12-20
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Frederick M. Schwarz , Gabriel L. Suciu
CPC classification number: F02C7/36 , F02C3/10 , F02C3/107 , F02C7/20 , F02C7/32 , F05D2220/36 , F05D2260/40311 , Y02T50/60 , Y10T29/49321
Abstract: An example gas turbine engine includes a propulsor assembly consisting of a fan module and a fan drive turbine module, an epicyclic gear train, a high spool and a low spool. A weight of the propulsor assembly is less than 40% of a total weight of a gas turbine engine. The high spool includes an outer shaft, a high pressure turbine and a high pressure compressor. The low spool includes an inner shaft, a low pressure turbine and a low pressure compressor. The inner shaft drives the propulsor through the gear train to drive the propulsor. A weight of the propulsor is greater than a weight of the low pressure turbine.
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公开(公告)号：US20230160338A1
公开(公告)日：2023-05-25
申请号：US18096872
申请日：2023-01-13
Applicant: Raytheon Technologies Corporation
Inventor： Frederick M. Schwarz
CPC classification number: F02C3/107 , F01D5/142 , F02K3/06 , F05D2260/40311 , F05D2220/327 , F05D2220/36 , Y02T50/60
Abstract: A turbine engine includes a first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.
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公开(公告)号：US20230122852A1
公开(公告)日：2023-04-20
申请号：US18082888
申请日：2022-12-16
Applicant: Raytheon Technologies Corporation
Inventor： Frederick M. Schwarz , Michael Winter , Charles E. Lents , Nathan Snape , Alan H. Epstein
IPC: B64D13/08 , F02C7/141 , F02C7/16 , F02C9/18 , F25B9/14 , F25J3/04 , B01D45/08 , B01D53/00 , B03C1/28 , B64D27/12 , B64D33/08 , F01D15/00 , F01D25/12 , F02C6/08 , F02C7/143 , F02C7/18 , F02K3/06 , B64D13/06 , F17C7/02 , F17C7/04
Abstract: An engine-driven cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for condensing a chilled air stream into liquid air for an aircraft use. The first air cycle machine includes a plurality of components operably coupled to a gearbox of a gas turbine engine and configured to produce a cooling air stream based on a first engine bleed source of the gas turbine engine. The second air cycle machine is operable to output the chilled air stream at a cryogenic temperature based on a second engine bleed source cooled by the cooling air stream of the first air cycle machine.
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公开(公告)号：US11614036B2
公开(公告)日：2023-03-28
申请号：US17530544
申请日：2021-11-19
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Paul R. Adams , Shankar S. Magge , Joseph B. Staubach , Wesley K. Lord , Frederick M. Schwarz , Gabriel L. Suciu
IPC: F02C7/36 , F02C3/107 , F02C9/18 , F02K3/06 , F02K3/075 , F01D5/06 , F01D25/24 , F02C3/04 , F02C7/20 , F04D19/02 , F01D11/12
Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.
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公开(公告)号：US11608786B2
公开(公告)日：2023-03-21
申请号：US17730782
申请日：2022-04-27
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Frederick M. Schwarz , Daniel Bernard Kupratis
Abstract: A gas turbine engine includes a propulsor section including a propulsor, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture, a turbine section including a low pressure turbine and a high pressure turbine, and a power density of greater than or equal to 4.75 and less than or equal to 5.5 lbf/in3, wherein the power density is a ratio of a thrust provided by the engine to a volume of the turbine section.
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Patent Agency Ranking
公开(公告)号：US11598286B2
公开(公告)日：2023-03-07
申请号：US17538048
申请日：2021-11-30
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Daniel Bernard Kupratis , Frederick M. Schwarz
Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section, a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a low pressure turbine and a high pressure turbine. The low pressure turbine drives the low pressure compressor and the gear arrangement to drive the propulsor. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.
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公开(公告)号：US11585268B2
公开(公告)日：2023-02-21
申请号：US17737179
申请日：2022-05-05
Applicant: Raytheon Technologies Corporation
Inventor： Frederick M. Schwarz
Abstract: A turbine engine includes a first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.
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公开(公告)号：US11560851B2
公开(公告)日：2023-01-24
申请号：US17086715
申请日：2020-11-02
Applicant: Raytheon Technologies Corporation
Inventor： Frederick M. Schwarz , Karl L. Hasel
Abstract: A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.
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公开(公告)号：US11480108B2
公开(公告)日：2022-10-25
申请号：US17230271
申请日：2021-04-14
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Paul R. Adams , Shankar S. Magge , Joseph B. Staubach , Wesley K. Lord , Frederick M. Schwarz , Gabriel L. Suciu
IPC: F02C7/36 , F02C3/107 , F02C9/18 , F01D11/12 , F01D5/06 , F01D25/24 , F02C3/04 , F02C7/20 , F04D19/02 , F02K3/075 , F02K3/06 , F04D29/32 , F04D29/54
Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is between 0.50 and 0.55, or is greater than 0.55 and less than or equal to 0.65.
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公开(公告)号：US20220268203A1
公开(公告)日：2022-08-25
申请号：US17737179
申请日：2022-05-05
Applicant: Raytheon Technologies Corporation
Inventor： Frederick M. Schwarz
Abstract: A turbine engine includes a first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.
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