-
公开(公告)号：US11300141B2
公开(公告)日：2022-04-12
申请号：US16532554
申请日：2019-08-06
Applicant: United Technologies Corporation
Inventor： David A. Topol
Abstract: An exemplary section of a gas turbine engine according to this disclosure includes, among other things, a first array of airfoils including a first number of airfoils, and a second array of airfoils downstream of the first array of airfoils. The second array includes a second number of airfoils. The second number of airfoils is at least 1.19 times the first number of airfoils thereby providing a predetermined modal.
-
公开(公告)号：US10655538B2
公开(公告)日：2020-05-19
申请号：US16143662
申请日：2018-09-27
Applicant: United Technologies Corporation
Inventor： Jonathan Gilson , Bruce L. Morin , Ramons A. Reba , David A. Topol , Wesley K. Lord
Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.
-
公开(公告)号：US10605172B2
公开(公告)日：2020-03-31
申请号：US13970670
申请日：2013-08-20
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol , Detlef Korte
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.
-
公开(公告)号：US10233948B2
公开(公告)日：2019-03-19
申请号：US14768839
申请日：2013-12-18
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： David A. Topol , Dilip Prasad
Abstract: In accordance with one aspect of the disclosure, a rotor for a gas turbine engine is disclosed. The rotor may include a rotor disk and a plurality of blade extending radially outward from the rotor disk. At least one of the blades may have a physical nonuniformity. The blades may be distributed about the rotor disk based on any physical nonuniformities of the blades to generate at least one decay-resistant harmonic.
-
公开(公告)号：US10113559B2
公开(公告)日：2018-10-30
申请号：US14850438
申请日：2015-09-10
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Matthew A. Turner , Colin J. Kling , David A. Topol
Abstract: A gas turbine engine impact liner is disclosed. The impact liner may include a base sheet, a plurality of stanchions extending from the base sheet, and a plurality of supports, each being operatively associated with one of the plurality of stanchions.
-
Patent Agency Ranking
公开(公告)号：US20170191424A1
公开(公告)日：2017-07-06
申请号：US15259232
申请日：2016-09-08
Applicant: United Technologies Corporation
Inventor： David A. Topol , Bruce L. Morin
CPC classification number: G01P3/48 , F01D5/02 , F01D5/12 , F02C3/04 , F02C3/10 , F02C7/24 , F02C7/32 , F02C7/36 , F02K3/06 , F05D2200/13 , F05D2200/14 , F05D2220/32 , F05D2220/36 , F05D2260/40311 , F05D2260/96 , Y02T50/673 , Y10T29/49236 , Y10T29/4932
Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section that has a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of blade rows of the compressor rotor, and the blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in at least one of the blade rows and the rotational speed are such that the following formula holds true for the at least one of the plurality of blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 sec≧about 5500 Hz. A method of designing a gas turbine engine is also disclosed.
-
公开(公告)号：US20160195021A1
公开(公告)日：2016-07-07
申请号：US15014363
申请日：2016-02-03
Applicant: United Technologies Corporation
Inventor： David A. Topol , Bruce L. Morin
CPC classification number: G01P3/48 , F01D5/02 , F01D5/12 , F02C3/04 , F02C3/10 , F02C7/24 , F02C7/32 , F02C7/36 , F02K3/06 , F05D2200/13 , F05D2200/14 , F05D2220/32 , F05D2220/36 , F05D2260/40311 , F05D2260/96 , Y02T50/673 , Y10T29/49236 , Y10T29/4932
Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction is configured to effect a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. Each of the compressor rotor and the fan drive turbine rotor includes a number of blades in each of a plurality of blade rows. The number of blades are configured to operate at least some of the time at a rotational speed.
Abstract translation: 根据本公开的示例的燃气涡轮发动机尤其包括风扇，具有风扇驱动涡轮转子的涡轮部分和压缩机转子。 齿轮减速构造成相对于来自风扇驱动涡轮转子的输入速度降低风扇的速度。 压缩机转子和风扇驱动涡轮机转子中的每一个在多个叶片行中的每一个中包括多个叶片。 叶片的数量被配置为在转速下的至少一些时间运行。
-
公开(公告)号：US10533447B2
公开(公告)日：2020-01-14
申请号：US14606087
申请日：2015-01-27
Applicant: United Technologies Corporation
Inventor： Bruce L. Morin , David A. Topol , Detlef Korte
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.
-
公开(公告)号：US10247018B2
公开(公告)日：2019-04-02
申请号：US15246752
申请日：2016-08-25
Applicant: United Technologies Corporation
Inventor： David A. Topol , Glen E. Potter , Flavien L. Thomas
Abstract: A disclosed fan section of a gas turbine engine includes a fan rotor having a plurality of fan blades and a duct defining a passageway aft of the fan rotor. A fan exit guide vane is disposed within the duct downstream of the fan blades. The fan exit guide vane includes a plurality of exit guide vanes positioned downstream of the fan rotor with at least two of the plurality of exit guide vanes including different aft geometries for guiding airflow through the passage to reduce pressure distortions at the fan blades.
-
公开(公告)号：US20170343574A1
公开(公告)日：2017-11-30
申请号：US15662528
申请日：2017-07-28
Applicant: United Technologies Corporation , MITU Aero Engines AG
Inventor： David A. Topol , Bruce L. Morin , Detlef Korte
CPC classification number: G01P3/48 , F01D5/02 , F01D5/12 , F02C3/04 , F02C3/10 , F02C7/24 , F02C7/32 , F02C7/36 , F02K3/06 , F05D2200/13 , F05D2200/14 , F05D2220/32 , F05D2220/36 , F05D2260/40311 , F05D2260/96 , Y02T50/673 , Y10T29/49236 , Y10T29/4932
Abstract: A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine rotor, a compressor rotor, and a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor. Each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least a plurality of the blade rows of the first turbine rotor: (number of blades×rotational speed)/60≧5500, and the rotational speed being an approach speed in revolutions per minute, and the following formula holds true for at least a plurality of the blade rows of the compressor rotor: (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute.
-
-
-
-
-
-
-
-
-
Search Results
50
records
(took 0.018 seconds)
