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    Turbofan arrangement with blade channel variations
    61.
    发明授权
    Turbofan arrangement with blade channel variations 有权

    公开(公告)号：US11118459B2

    公开(公告)日：2021-09-14

    申请号：US16519475

    申请日：2019-07-23

    Applicant: United Technologies Corporation

    Inventor： Edward J. Gallagher ,  Sue-Li Chuang

    IPC: F01D5/14 ,  F04D29/32 ,  F02C3/10 ,  F04D29/38

    Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section having a fan including a rotor hub rotatable about an engine longitudinal axis and an array of fan blades, a geared architecture, 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. The low pressure turbine drives the fan through the geared architecture. Each fan blade of the array of fan blades includes pressure and suction sides and extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, and facing pressure and suction sides of adjacent fan blades define a channel in a chordwise direction having a width between the facing pressure and suction sides at a given span position of the adjacent fan blades. The width at each pressure side location along the channel is defined as a minimum distance to a location along the suction side, and the width converges along the channel for at least some span positions.

    Gas turbine engine airfoil
    63.
    发明授权
    Gas turbine engine airfoil 有权

    公开(公告)号：US10519971B2

    公开(公告)日：2019-12-31

    申请号：US15113868

    申请日：2014-08-22

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Edward J. Gallagher ,  Byron R. Monzon ,  Ling Liu ,  Linda S. Li ,  Darryl Whitlow ,  Barry M. Ford

    IPC: F04D29/38 ,  F01D5/14 ,  F04D29/32 ,  F02C3/04 ,  F02C7/36 ,  F02K3/06

    Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil geometry corresponds to axial leading and trailing edge curves and an axial stacking offset curve. The airfoil extends from a root and a zero axial reference point corresponds to axial center of the root. XLE corresponds to an axial distance from a leading edge to the reference point at a given span position. XTE corresponds to a axial distance from a trailing edge to the reference point at a given span position. Xd corresponds to an axial stacking offset at a given span position. (XLE−Xd)/(Xd−XTE) at 100% span position is about 1 and (XLE−Xd)/(Xd−XTE) at 90% span position is about 1.

    EFFICIENT, LOW PRESSURE RATIO PROPULSOR FOR GAS TURBINE ENGINES
    66.
    发明申请
    EFFICIENT, LOW PRESSURE RATIO PROPULSOR FOR GAS TURBINE ENGINES 审中-公开

    公开(公告)号：US20180230943A1

    公开(公告)日：2018-08-16

    申请号：US15936552

    申请日：2018-03-27

    Applicant: United Technologies Corporation

    Inventor： Edward J. Gallagher ,  Byron R. Monzon

    IPC: F02K3/06 ,  F02K1/06 ,  F01D17/14

    CPC classification number: F02K3/06 ,  F01D17/14 ,  F02K1/06 ,  F05D2220/327 ,  F05D2260/4031

    Abstract: A gas turbine engine includes a bypass flow passage and a core flow passage. The bypass flow passage defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is located upstream of the bypass flow passage. The bypass flow passage includes an inlet and an outlet that define a design fan pressure ratio of approximately 1.3 to 1.55. A first, inner shaft and a second, outer shaft are concentric. A first turbine is coupled with the first shaft, and the first shaft is coupled with the fan. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number of the fan blades, the number (N) being 18, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.1, and a ratio of N/R that is from 16.4 to 18.0.

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