-
公开(公告)号：US09970452B2
公开(公告)日：2018-05-15
申请号：US14623761
申请日：2015-02-17
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Richard David Conner , Michael Todd Barton , Yoseph Gebre-Giorgis , Harry Lester Kington , John Repp
CPC classification number: F04D29/284 , F01D5/225 , F01D9/04 , F02C3/09 , F05D2250/38 , F05D2260/941 , Y02T50/671
Abstract: Embodiments of a forward-swept impeller are provide, as are embodiments of a gas turbine engine containing a forward-swept impeller. In one embodiment, the gas turbine engine includes a shaft and a forward-swept impeller mounted to the shaft. The forward-swept impeller includes, in turn, an inboard impeller section, an outboard impeller section circumscribing the inboard impeller section, and a plurality of hub flow paths extending over the forward-swept impeller from the inboard impeller section to the outboard impeller section. The plurality of hub flow paths each have a flow path exit that is tilted in a forward direction, as taken along a line tangent to the flow path exit.
-
公开(公告)号：US12286896B2
公开(公告)日：2025-04-29
申请号：US18408260
申请日：2024-01-09
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce D. Reynolds , John Repp
Abstract: A turbomachine includes a shroud and a hub spaced apart from the shroud to channel an airflow along a direction. The turbomachine includes a plurality of airfoils coupled between the shroud and the hub. At least one airfoil of the plurality of airfoils includes a leading edge spaced apart from a trailing edge in the direction of the airflow and a pressure side opposite a suction side. The at least one airfoil has a chord that is 0% chord at the leading edge and 100% chord at the trailing edge, and a span defined as 0% at the hub and 100% at the shroud. The turbomachine includes at least one local vortex generator array defined onto the hub between 0% chord to −10% chord at 0% span so as to extend forward of the leading edge.
-
23.发明申请公开(公告)号：US20220010682A1
公开(公告)日：2022-01-13
申请号：US16927161
申请日：2020-07-13
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Jeffrey Gluck , Bruce Reynolds , John Repp
Abstract: Systems and methods for air injection passageway integration and optimization in turbomachinery using surface vortex generation. An airfoil including a leading edge, a trailing edge, a pressure side, and a suction side, and is configured to influence an airflow as it passes from the leading edge to the trailing edge. The airfoil defines an aerodynamic passageway having an inlet on the pressure side and an outlet on the suction side to deliver air from the airflow through the airfoil to the suction side. The outlets are configured to inject the air at areas on either airfoil side targeted due to their propensity to generate undesirable boundary layer growth and associated flow losses. Outlet may also be included in the hub and the shroud of the turbomachine.
-
公开(公告)号：US20210381380A1
公开(公告)日：2021-12-09
申请号：US16892145
申请日：2020-06-03
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Nick Nolcheff , John Repp , Bruce Reynolds , John Gunaraj
Abstract: A rotor for a turbofan booster section associated with a fan section of a gas turbine engine includes a rotor blade having an airfoil extending from a root to a tip and having a leading edge and a trailing edge. The airfoil has a plurality of chord lines spaced apart in a spanwise direction. Each chord line of the plurality of chords lines is defined between the leading edge and the trailing edge and has a normalized chord value. From the hub, the normalized chord value decreases to a minimum value between about 20% to about 90% span and increases from the minimum value to the tip. The rotor includes a rotor disk coupled to the rotor blade configured to be coupled to the shaft or the fan to rotate with the shaft or the fan, respectively, at the same speed as the shaft and fan.
-
25.发明申请公开(公告)号：US20200158128A1
公开(公告)日：2020-05-21
申请号：US16198279
申请日：2018-11-21
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： David Richard Hanson , John A. Gunaraj , Jeffrey Hayes , Nick Nolcheff , John Repp
IPC: F04D29/38
Abstract: A rotor for a compressor includes a hub and a plurality of airfoils having a root, a tip opposite the root and a span that extends from 0% at the root to 100% at the tip. Each of the airfoils is coupled to the hub at the root and is spaced apart from adjacent ones of the airfoils over the span by a throat dimension. The throat dimension has a maximum value at a spanwise location between 60% of the span and 90% of the span of the adjacent ones of the airfoils, and at 10% of the span of the adjacent ones of the airfoils above or below the spanwise location of the maximum value, the throat dimension is less than 97% of the maximum value. The throat dimension at 5% of the span of the adjacent ones of the airfoils is less than 70% of the maximum value.
-
Patent Agency Ranking
公开(公告)号：US10330121B2
公开(公告)日：2019-06-25
申请号：US14632024
申请日：2015-02-26
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce David Reynolds , David Richard Hanson , John Repp , Michael Todd Barton , John A. Gunaraj
Abstract: Methods and apparatuses are provided for a compressor. The compressor includes a first stage having a first rotor and a first stator, and a second stage downstream from the first stage in a direction of a fluid flow. The compressor also includes a secondary flow system that directs fluid from the second stage into the first stator to improve at least one of a performance and a stability of the compressor.
-
公开(公告)号：US20180355791A1
公开(公告)日：2018-12-13
申请号：US16056101
申请日：2018-08-06
Applicant: Honeywell International Inc.
Inventor： Nick Nolcheff , John Repp , Bruce David Reynolds , David Richard Hanson
CPC classification number: F02C3/08 , F04D17/025 , F04D19/02 , F04D29/286 , F04D29/321 , F04D29/542 , F05D2220/32 , F05D2220/3216 , F05D2240/35
Abstract: Methods and apparatus are provided for an axi-centrifugal compressor in a gas turbine engine for a business aviation or rotorcraft propulsion unit. The compressor includes an axial compressor section operable to affect a first pressure ratio along the flow path between a compressor inlet and a first section exit, and a centrifugal compressor section operable to affect a second pressure ratio along the flow path between a second section inlet and the compressor exit. The pressure rise across the axial and centrifugal compressor section is configured to have a tuning factor is in a range between 2.8 and 4.5 and a loading factor in a range between 0.6 and 0.8.
-
公开(公告)号：US20180238341A1
公开(公告)日：2018-08-23
申请号：US15439430
申请日：2017-02-22
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Nick Nolcheff , John Repp , Jeffrey Hayes , John A Gunaraj , Yoseph Gebre-Giorgis
CPC classification number: F04D29/325 , B64C11/14 , B64D33/02 , B64D2033/0233 , F01D5/021 , F01D5/141 , F01D5/34 , F02C7/05 , F02K3/06 , F04D29/34 , F04D29/541 , F05D2220/323 , F05D2220/36 , F05D2240/12 , F05D2240/20 , Y02T50/671 , Y02T50/673
Abstract: Embodiments of a core-protecting fan module are provided, as are embodiments of a turbofan engine containing such a fan module. In an embodiment, the core-protecting fan module contains a nose member, a fan rotor downstream of the nose member, a full span stator downstream of the fan rotor, and a splitter structure downstream of the fan rotor. The fan rotor includes a plurality of fan blades, which extends from a rotor hub and which is angularly spaced about a rotational axis. Certain fundamental angular relationships are observed between the angles formed by rotational axis, the nose member, the fan rotor, and a leading edge of the splitter structure to reduce contaminant ingestion by the core flow path and to promote moisture shedding to reduce susceptibility to icing within the fan module, while further avoiding or minimizing negative impacts to other structural and functional aspects of the turbofan engine.
-
公开(公告)号：US20170276070A1
公开(公告)日：2017-09-28
申请号：US15079538
申请日：2016-03-24
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Nick Nolcheff , John Repp , Bruce David Reynolds , David Richard Hanson
CPC classification number: F02C3/08 , F04D17/025 , F04D19/02 , F04D29/286 , F04D29/321 , F04D29/542 , F05D2220/32 , F05D2220/3216 , F05D2240/35
Abstract: Methods and apparatus are provided for an axi-centrifugal compressor in a gas turbine engine for a business aviation or rotorcraft propulsion unit. The compressor includes an axial compressor section operable to affect a first pressure ratio along the flow path between a compressor inlet and a first section exit, and a centrifugal compressor section operable to affect a second pressure ratio along the flow path between a second section inlet and the compressor exit. The pressure rise across the axial and centrifugal compressor section is configured to have a tuning factor is in a range between 2.8 and 4.5 and a loading factor in a range between 0.6 and 0.8.
-
公开(公告)号：US20160252098A1
公开(公告)日：2016-09-01
申请号：US14632024
申请日：2015-02-26
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce David Reynolds , David Richard Hanson , John Repp , Michael Todd Barton , John A. Gunaraj
CPC classification number: F04D29/684 , F01D9/065 , F04D27/0215 , F04D27/0238 , F04D29/522 , F04D29/682 , F05D2270/10 , F05D2270/101
Abstract: Methods and apparatuses are provided for a compressor. The compressor includes a first stage having a first rotor and a first stator, and a second stage downstream from the first stage in a direction of a fluid flow. The compressor also includes a secondary flow system that directs fluid from the second stage into the first stator to improve at least one of a performance and a stability of the compressor.
Abstract translation: 为压缩机提供了方法和装置。 压缩机包括具有第一转子和第一定子的第一级，以及在流体流动方向上从第一级下游的第二级。 压缩机还包括二次流系统，其将来自第二级的流体引导到第一定子中以改进压缩机的性能和稳定性中的至少一个。
-
-
-
-
-
-
-
-
-
Search Results
30
records
(took 0.024 seconds)
