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公开(公告)号：US20180334961A1
公开(公告)日：2018-11-22
申请号：US15599149
申请日：2017-05-18
发明人： Paul R. Hanrahan , Jonathan Ortiz
CPC分类号： F02C7/18 , F01D5/081 , F01D9/02 , F02C3/04 , F02C6/08 , F02C7/185 , F02C9/18 , F02K3/06 , F05D2220/32 , F05D2240/24 , F05D2240/35 , F05D2260/213
摘要： A cooling system for a component of a gas turbine engine includes a first airflow passage configured to direct a first airflow to a mixing chamber and a second airflow passage to configured direct a second airflow to the mixing chamber, the second airflow having a higher temperature than the first airflow, and a cooling airflow passage to direct a cooling airflow from the mixing chamber to the component, the cooling airflow comprising the first airflow and the second airflow. The airflow passages are configured and sized to allow an amount of cooling airflow for unrestricted engine operation. When the first airflow passage is disabled, the second airflow passage and cooling airflow passage are configured and sized to allow an amount of cooling airflow which is adequate to permit continued safe engine operation restricted to within only a portion of its normal parameters and operating envelope.
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公开(公告)号：US11773742B2
公开(公告)日：2023-10-03
申请号：US16806168
申请日：2020-03-02
CPC分类号： F01D11/20 , F01D5/02 , F01D9/065 , F01D11/24 , F02C3/04 , F02C9/18 , F04D29/542 , F05D2220/32 , F05D2240/35 , F05D2260/213 , F05D2300/50212 , Y02T50/60
摘要： A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.
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公开(公告)号：US11739697B2
公开(公告)日：2023-08-29
申请号：US15601524
申请日：2017-05-22
CPC分类号： F02C9/18 , F01D9/065 , F01D17/02 , F02C7/18 , F05D2220/32 , F05D2260/20 , F05D2270/30
摘要： A bleed air cooling system for a gas turbine engine includes one or more bleed ports located at one or more axial locations of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. One or more safety sensors are configured to sense operational characteristics of the gas turbine engine, and a controller is operably connected to the one or more safety sensors and configured to evaluate the sensed operational characteristics for anomalies in operation of the bleed air cooling system.
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公开(公告)号：US11459906B2
公开(公告)日：2022-10-04
申请号：US15601544
申请日：2017-05-22
摘要： A bleed air cooling system for a gas turbine engine includes one or more bleed flowpaths operably connected to a bleed outlet to divert a bleed airflow from a gas turbine engine flowpath. Each bleed flowpath includes two or more bleed ports to divert a bleed airflow from a gas turbine engine flowpath, and a bleed duct in fluid communication with the bleed ports and configured to convey the bleed airflow from the two or more bleed ports to the bleed outlet. A valve is located at each bleed port of and is configured to move between an opened position and a closed position, and one or more sensors are located along the bleed flowpath to sense one or more conditions of the bleed air cooling system. The valve at a particular bleed port is moved to the opened position based on the sensed one or more conditions.
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公开(公告)号：US10934845B2
公开(公告)日：2021-03-02
申请号：US16794900
申请日：2020-02-19
发明人： Matthew A Devore , Jonathan Ortiz
摘要： An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber. A tangential onboard injector (TOBI) may be disposed in the first airflow path. A radial onboard injector (ROBI) may be disposed in the second airflow path.
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公开(公告)号：US20200347741A1
公开(公告)日：2020-11-05
申请号：US16806168
申请日：2020-03-02
摘要： A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.
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公开(公告)号：US10711640B2
公开(公告)日：2020-07-14
申请号：US15484168
申请日：2017-04-11
IPC分类号： F01D17/16 , F01D25/12 , F02C3/04 , F01D17/10 , F01D9/04 , F01D11/14 , F02C6/08 , F01D5/08 , F01D9/06
摘要： A gas turbine engine comprises a compressor section, a combustor, and a turbine section. The turbine section includes a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap taps air having been compressed by the compressor, the tapped air being passed through a heat exchanger. A vane section has vanes downstream of the combustor, but upstream of the first stage blade, and the air downstream of the heat exchanger passes radially inwardly of the combustor, along an axial length of the combustor, and then radially outwardly through a hollow chamber in the vanes, and then across the blade outer air seal, to cool the blade outer air seal.
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公开(公告)号：US10577964B2
公开(公告)日：2020-03-03
申请号：US15475656
申请日：2017-03-31
摘要： A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.
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公开(公告)号：US20180334918A1
公开(公告)日：2018-11-22
申请号：US15601544
申请日：2017-05-22
CPC分类号： F01D17/105 , F01D5/02 , F01D17/145 , F01D21/003 , F01D25/12 , F02C7/12 , F02C9/18 , F04D27/009 , F05D2220/32 , F05D2260/606
摘要： A bleed air cooling system for a gas turbine engine includes one or more bleed flowpaths operably connected to a bleed outlet to divert a bleed airflow from a gas turbine engine flowpath. Each bleed flowpath includes two or more bleed ports to divert a bleed airflow from a gas turbine engine flowpath, and a bleed duct in fluid communication with the bleed ports and configured to convey the bleed airflow from the two or more bleed ports to the bleed outlet. A valve is located at each bleed port of and is configured to move between an opened position and a closed position, and one or more sensors are located along the bleed flowpath to sense one or more conditions of the bleed air cooling system. The valve at a particular bleed port is moved to the opened position based on the sensed one or more conditions.
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公开(公告)号：US20180298774A1
公开(公告)日：2018-10-18
申请号：US15490304
申请日：2017-04-18
CPC分类号： F01D11/04 , F01D5/06 , F01D5/082 , F01D9/041 , F01D11/001 , F01D11/02 , F01D25/12 , F05D2220/32 , F05D2240/128 , F05D2240/56 , F05D2260/14 , F05D2260/601
摘要： Gas turbine engines and turbines thereof including a stator section having a plurality of vanes, a rotating section having a plurality of blades, the rotating section being axially adjacent the stator section along an axis of the turbine, the stator section being aftward of the rotating section along the axis of the turbine, and a primary tangential onboard injector located radially inward from the stator section and configured to direct an airflow from the stator section in a forward direction toward the rotating section, the primary tangential onboard injector turning the airflow in a direction of rotation of the rotating section.
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