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    GAS TURBINE ENGINE AIRFOIL
    82.
    发明申请
    GAS TURBINE ENGINE AIRFOIL 审中-公开

    公开(公告)号：US20200149545A1

    公开(公告)日：2020-05-14

    申请号：US16741918

    申请日：2020-01-14

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Edward J. Gallagher ,  Lisa I. Brilliant ,  Joseph C. Straccia ,  Stanley J. Balamucki ,  Mark A. Stephens ,  Kate Hudon

    IPC: F04D29/38 ,  F04D29/32 ,  F02K3/06 ,  F02C7/36 ,  F02C3/04 ,  F01D25/24 ,  F01D17/16 ,  F01D5/14 ,  F04D29/54

    Abstract: A gas turbine engine includes a combustor section arranged between a compressor section and a turbine section. The compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged upstream of the combustor section. A fan section has an array of twenty-six or fewer fan blades. The low pressure compressor is downstream from the fan section. An airfoil is provided in the compressor section outside the high pressure compressor section and includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a stagger angle and span position that defines a curve with the stagger angle greater than 35° from 0% span to at least 50% span. The stagger angle has a positive slope from 20% span to 100% span.

    Efficient, low pressure ratio propulsor for gas turbine engines
    83.
    发明授权
    Efficient, low pressure ratio propulsor for gas turbine engines 有权

    公开(公告)号：US10605202B2

    公开(公告)日：2020-03-31

    申请号：US15936552

    申请日：2018-03-27

    Applicant: United Technologies Corporation

    Inventor： Edward J. Gallagher ,  Byron R. Monzon

    IPC: F02K3/075 ,  F02K3/06 ,  F01D17/14 ,  F02K1/06

    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.

    GAS TURBINE ENGINE AIRFOIL
    84.
    发明申请
    GAS TURBINE ENGINE AIRFOIL 审中-公开

    公开(公告)号：US20200096005A1

    公开(公告)日：2020-03-26

    申请号：US16681187

    申请日：2019-11-12

    Applicant: United Technologies Corporation

    Inventor： Edward J. Gallagher ,  Lisa I. Brilliant ,  Joseph C. Straccia ,  Stanley J. Balamucki ,  Mark A. Stephens ,  Kate Hudon

    IPC: F04D29/38 ,  F01D5/14 ,  F01D9/04 ,  F02C3/04 ,  F04D29/32 ,  F04D29/54 ,  F04D29/56

    Abstract: A gas turbine engine includes a combustor section, a compressor section, and a turbine section. A high pressure compressor is arranged upstream of the combustor section. A fan section has an array of twenty-six or fewer fan blades. A low pressure compressor is counter-rotating relative to the fan blades and downstream from the fan section. The airfoil outside of the compressor section is rotatable relative to an engine static structure. The airfoil includes pressure and suction sides extending from a 0% span position to a 100% span position. A relationship between a leading edge angle and span position that defines a curve with at least one of a decreasingly negative slope or a positive slope from 80% span to 100% span. The leading edge angle is defined by a tangential plane normal to an engine axis and a line tangent to a camber mean line at the leading edge.

    Gas turbine engine airfoil
    85.
    发明授权
    Gas turbine engine airfoil 有权

    公开(公告)号：US10550852B2

    公开(公告)日：2020-02-04

    申请号：US15460333

    申请日：2017-03-16

    Applicant: United Technologies Corporation

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

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

    Abstract: A gas turbine engine includes a combustion section arranged between a compressor section and a turbine section that extend in an axial direction. A fan section is arranged upstream from the compressor section. An airfoil is arranged in one of the fan section, the compressor section and the turbine section. The airfoil includes pressure and suction sides 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. The airfoil has a leading edge that is projected onto a plane from various views and the plane is perpendicular to a viewing direction which corresponds to the various views. The plane is parallel with the axial direction in a 0° view. The various views include the 0° view which projects into an axial plane in the axial direction. A 90° view projects into a tangential plane in a tangential direction normal to the axial direction and views between the 0° and 90° views. The airfoil has a maximum leading edge projection in a 20° to 40° view. The radial direction is normal to the axial and tangential directions.

    Gas turbine engine airfoil
    87.
    发明授权
    Gas turbine engine airfoil 有权

    公开(公告)号：US10358925B2

    公开(公告)日：2019-07-23

    申请号：US15294939

    申请日：2016-10-17

    Applicant: United Technologies Corporation

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

    IPC: F01D5/14 ,  F02K3/00 ,  F02C3/107 ,  F02K3/06 ,  F04D29/38

    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 has a relationship between a stacking offset and a span position that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X1 is provided from the zero-crossing position to a negative-most value on the curve. A second axial stacking offset X2 is provided from the zero-crossing position to a positive-most value on the curve. A ratio of the second to first axial stacking offset X2/X1 is less than 2.0.

    Efficient, low pressure ratio propulsor for gas turbine engines
    88.
    发明授权
    Efficient, low pressure ratio propulsor for gas turbine engines 有权

    公开(公告)号：US10233868B2

    公开(公告)日：2019-03-19

    申请号：US15252811

    申请日：2016-08-31

    Applicant: United Technologies Corporation

    Inventor： Edward J. Gallagher ,  Byron R. Monzon

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

    Abstract: A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A first turbine is coupled with the first shaft such that rotation of the first turbine is configured to drive the fan, through the first shaft and gear assembly, at a lower speed than the first shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 (N) of the fan blades, a solidity value (R) that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0.

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