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41.发明申请公开(公告)号：US20160153310A1
公开(公告)日：2016-06-02
申请号：US14905994
申请日：2014-07-02
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Frederick M. Schwarz , Bruce L. Morin
CPC classification number: F01D25/04 , F01D5/06 , F01D5/142 , F01D9/06 , F01D15/12 , F01D17/105 , F02C3/107 , F02K3/06 , F04D29/321 , F04D29/325 , F05D2220/32 , F05D2220/36 , F05D2240/24 , F05D2260/40311 , F05D2260/96 , F05D2270/02 , F05D2270/333 , Y02T50/673
Abstract: A turbine section including a high pressure turbine, an intermediate pressure turbine and a fan drive turbine, the fan drive turbine driving a gear reduction to in turn drive a fan, and effecting a reduction in the speed of the fan relative to an input speed from the fan drive turbine and said high pressure turbine driving a high pressure compressor, and the intermediate pressure turbine driving a low pressure compressor, with the intermediate pressure turbine having a number of turbine blades in at least one row, and the turbine blades operating at least some of the time at a rotational speed, and the number of turbine blades in the at least one row, and the rotational speed being such that the following formula holds true for the at least one row of the intermediate pressure turbine: (number of blades×speed)/60?5500 Hz.
Abstract translation: 包括高压涡轮机，中压涡轮机和风扇驱动涡轮机的涡轮机部分，风扇驱动涡轮驱动齿轮减速，从而驱动风扇，并且相对于输入速度降低风扇的速度， 所述风扇驱动涡轮机和所述高压涡轮机驱动高压压缩机，并且所述中压涡轮机驱动低压压缩机，所述中压涡轮机具有至少一排中的多个涡轮叶片，并且所述涡轮叶片至少工作 一些转速的时间，以及至少一行中的涡轮叶片的数量，并且转速使得下列公式对于至少一排中压涡轮机成立：（叶片数量 ×速度）/ 60〜5500Hz。
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公开(公告)号：US20140318147A1
公开(公告)日：2014-10-30
申请号：US14144710
申请日：2013-12-31
Applicant: United Technologies Corporation
Inventor： David A. Topol , Bruce L. Morin
CPC classification number: F02C3/10 , F01D5/02 , F01D5/12 , F02C7/24 , F02C7/36 , F05D2260/96 , Y02T50/673 , Y10T29/49236 , Y10T29/4932
Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: (number of blades×rotational speed)/60 s≧5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.
Abstract translation: 燃气涡轮发动机具有风扇和具有风扇驱动涡轮机转子的涡轮机。 风扇驱动涡轮转子驱动压缩机转子。 齿轮减速实现了风扇相对于来自驱动压缩机转子的风扇驱动涡轮转子的输入速度的降低。 压缩机转子在压缩机转子的多排中的至少一排中具有多个压缩机叶片。 叶片在旋转速度下至少运行一些时间。 至少一行中的压缩机叶片的数量和转速使得至少一排压缩机转子涡轮机的以下公式成立：（叶片数×转速）/60s≥5500Hz， 转速为每分钟转。 还公开了一种设计燃气涡轮发动机和压缩机组件的方法。
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公开(公告)号：US20140271112A1
公开(公告)日：2014-09-18
申请号：US13970670
申请日：2013-08-20
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol , Detlef Korte
IPC: F02C7/24
CPC classification number: F02C7/24 , F02K3/06 , F05D2200/36 , F05D2210/31 , F05D2240/301 , F05D2240/307 , Y02T50/673
Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a compressor section driven by the turbine section, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.
Abstract translation: 根据本公开的示例性方面的燃气涡轮发动机包括涡轮部分，其包括风扇驱动涡轮机，由涡轮机部分驱动的压缩机部分，由风扇驱动涡轮机驱动的齿轮架构，以及风扇 由风扇驱动器通过齿轮架构驱动。 涡轮部分的至少一个级包括可旋转刀片阵列和一组叶片。 叶片数与叶片数之比大于或等于约1.55。 叶片的机械末端旋转马赫数被配置为在接近速度下大于或等于约0.5。
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公开(公告)号：US11168614B2
公开(公告)日：2021-11-09
申请号：US16833782
申请日：2020-03-30
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol , Detlef Korte
IPC: F02C7/24 , F02K3/06 , F01D5/02 , F01D5/12 , F01D9/04 , F01D15/12 , F01D17/10 , F01D5/16 , F01D25/04 , F02C3/107 , F02C7/36
Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a compressor section driven by the turbine section, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.
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公开(公告)号：US20200347785A1
公开(公告)日：2020-11-05
申请号：US16411728
申请日：2019-05-14
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , Anthony R. Bifulco
IPC: F02C7/24
Abstract: An acoustic liner for a gas turbine engine includes an acoustic panel that is curved about a central axis. The acoustic panel includes a support backing, a face sheet, and a cellular structure disposed between the support backing and the face sheet. The face sheet has elongated slots that extend along respective slot centerlines in the plane of the face sheet. The slot centerlines are sloped at oblique angles to the central axis.
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Patent Agency Ranking
公开(公告)号：US20200291859A1
公开(公告)日：2020-09-17
申请号：US16740611
申请日：2020-01-13
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol , Detlef Korte
IPC: F02C7/24 , F02K3/06 , F01D5/02 , F01D5/12 , F01D9/04 , F01D15/12 , F01D17/10 , F01D5/16 , F01D25/04 , F02C3/107 , F02C7/36
Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.
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公开(公告)号：US10704415B2
公开(公告)日：2020-07-07
申请号：US14761689
申请日：2014-03-11
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Bruce L. Morin , Wesley K. Lord
Abstract: In accordance with one aspect of the disclosure, a gas turbine engine, method of using and designing such is disclosed. The gas turbine engine may comprise a fan including a plurality of blades, and a variable area fan nozzle. The fan may be configured to have a design point fan tip leading edge relative flow angle βADP, and may be further configured to have an off-design point fan tip leading edge relative flow angle β at an off-design fan operating point. The variable area fan nozzle may be configured to manipulate the amount of air flowing through the fan so that the absolute value of a difference between the design point fan tip leading edge relative flow angle βADP and the off-design point fan tip leading edge relative flow angle β is in a specified range.
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公开(公告)号：US10519859B2
公开(公告)日：2019-12-31
申请号：US15452604
申请日：2017-03-07
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Ian T. Marchaj , Bruce L. Morin , Mani Sadeghi , Daniel L. Gysling
Abstract: Disclosed is a flutter damper, including an acoustic liner having a perforated radial inner face sheet and a radial outer back sheet, the acoustic liner being configured for peak acoustical energy absorption at a frequency range that is greater than a frequency range associated with fan flutter, a chamber secured to the radial outer back sheet, the chamber being in fluid communication with the acoustic liner, and the chamber being configured for peak acoustical energy absorption at a frequency range that is associated with one or more fan flutter modes, and at least one stiffening structure connected to a top surface of the chamber that tunes the top surface out of the frequency range associated with one or more fan flutter modes.
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公开(公告)号：US10337407B2
公开(公告)日：2019-07-02
申请号：US14774273
申请日：2014-02-19
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol
Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan, a compressor section having a low pressure compressor and a high pressure compressor, a combustor section, and a turbine section having a low pressure turbine, the low pressure turbine for driving the low pressure compressor and the fan; a gear reduction effecting a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor; and at least one stage of the compressor section having a ratio of vanes to blades that is greater than or equal to 1.8. The corrected tip speed of the blades is greater than or equal to 480 ft/sec at an approach speed.
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公开(公告)号：US20180299477A1
公开(公告)日：2018-10-18
申请号：US16018754
申请日：2018-06-26
Applicant: United Technologies Corporation , MTU Aero Engines AG
Inventor： David A. Topol , Bruce L. Morin , Detlef Korte
IPC: G01P3/48 , F02C7/32 , F02K3/06 , F02C7/24 , F01D5/12 , F02C3/04 , F02C7/36 , F01D5/02 , F02C3/10
Abstract: A gas turbine engine has a fan section including a fan, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, and a gear reduction. The low pressure compressor and the low pressure turbine have a number of blades in each of at least one of a plurality of blade rows. The blades are rotatable at least some of the time at a rotational speed in operation. The number of blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: 5500≤(number of blades×rotational speed)/60≤10000, the rotational speed being in revolutions per minute.
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