公开(公告)号：US20190382121A1

公开(公告)日：2019-12-19

申请号：US16012047

申请日：2018-06-19

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  William G. Sheridan

IPC: B64D27/02 ,  B64D27/12 ,  B64D27/24 ,  B64D31/14 ,  B64D33/04 ,  F02C6/20 ,  F02C6/14 ,  B64D31/02 ,  H02K7/18 ,  H02K7/20 ,  H02K11/00

Abstract: A propulsion system for an aircraft includes at least two gas turbine engines and at least one auxiliary propulsion fan. The at least one auxiliary propulsion fan is configured to selectively receive a motive force from either or both of the at least two gas turbine engines through at least one shaft operatively coupled to the at least one auxiliary propulsion fan.

公开(公告)号：US20190353098A1

公开(公告)日：2019-11-21

申请号：US16531704

申请日：2019-08-05

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz

IPC: F02C3/107 ,  F01D5/14 ,  F02K3/06

Abstract: A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

公开(公告)号：US20190323380A1

公开(公告)日：2019-10-24

申请号：US16404217

申请日：2019-05-06

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Frederick M. Schwarz ,  Lisa I. Brilliant

IPC: F01D25/14 ,  F01D5/12 ,  F02C7/24 ,  F02C7/36 ,  F02C3/04 ,  F01D25/26 ,  F01D5/08 ,  F02C7/14 ,  F01D5/02 ,  F01D9/04 ,  F01D25/24

Abstract: A gas turbine engine having an engine axis and method of manufacturing the same is disclosed. The gas turbine engine may comprise a fan configured to drive air, a low pressure compressor section having a core flow path and configured to draw in and compress air flowing along the core flow path, a spool configured to drive the fan, and geared architecture configured to adjust the fan speed. The gas turbine engine may also include a housing defining a compartment that encloses the geared architecture. The housing is disposed between the core flow path and the axis, and includes a shielded mid-section that is in thermal communication with the core flow path of the low pressure compressor section. The shielded mid-section includes an outer layer and an insulator adjacent to the outer layer.

公开(公告)号：US10451004B2

公开(公告)日：2019-10-22

申请号：US15173288

申请日：2016-06-03

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Brian D. Merry ,  Christopher M. Dye ,  Steven B. Johnson ,  Frederick M. Schwarz

IPC: F01D5/06 ,  F02C7/20 ,  F02C7/36 ,  F02K3/06 ,  B64D27/26 ,  F01D15/12 ,  F01D25/24 ,  F01D25/28 ,  F02C3/107 ,  F02C9/20 ,  F02C9/18 ,  F01D9/02 ,  F02K1/15

Abstract: A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a first nacelle which at least partially surrounds a second nacelle and the fan rotor, the fan section configured to communicate airflow into the first nacelle and the second nacelle, a first turbine, and a second turbine followed by the first turbine. The first turbine is configured to drive the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.

公开(公告)号：US10436121B2

公开(公告)日：2019-10-08

申请号：US14038886

申请日：2013-09-27

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Frederick M. Schwarz ,  Robert E. Malecki

IPC: F02K3/06 ,  F02C7/36 ,  F02C3/10 ,  F02K3/068 ,  B64D33/02 ,  F02C3/107

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

公开(公告)号：US20190292985A1

公开(公告)日：2019-09-26

申请号：US15928506

申请日：2018-03-22

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  Nathan Snape

IPC: F02C7/18

Abstract: A gas turbine engine includes a plurality of rotating components housed within a main compressor section and a turbine section. A first tap is connected to the main compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A cooling air valve is configured to selectively block flow of cooling air across the heat exchanger. A cooling compressor is connected downstream of the heat exchanger. A shut off valve stops flow between the heat exchanger and the cooling compressor. A second tap is configured to deliver air at a second pressure which is higher than the first pressure. A mixing chamber is connected downstream of the cooling compressor and the second tap. The mixing chamber is configured to deliver air to at least one of the plurality of rotating components. A system stops flow between the cooling compressor and the plurality of rotating components. A controller is configured to modulate flow between the heat exchanger and the plurality of rotating components under certain power conditions of the gas turbine engine. The controller is programmed to control the cooling air valve, the shut off valve and the system such that flow is stopped between the heat exchanger and the cooling compressor only after the cooling compressor has been stopped.

公开(公告)号：US10358932B2

公开(公告)日：2019-07-23

申请号：US14754161

申请日：2015-06-29

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  William K. Ackermann

IPC: F01D11/00 ,  F01D11/02 ,  F16J15/44

Abstract: An assembly is provided for rotational equipment. The assembly includes a circumferentially segmented stator and a rotor radially within the stator. The assembly also includes a seal assembly configured for substantially sealing a gap radially between the stator and the rotor. The seal assembly includes a carrier and a non-contact seal seated with the carrier. The carrier includes a plurality of discrete carrier segments circumferentially arranged around the non-contact seal.

公开(公告)号：US20190219484A1

公开(公告)日：2019-07-18

申请号：US16264760

申请日：2019-02-01

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  Marnie A. Rizo ,  David P. Houston ,  David M. Nissley ,  Paul J. Hiester ,  Timothy Dale ,  Timothy B. Winfield ,  Madeline Campbell ,  James R. Midgley

IPC: G01M15/14 ,  G05B23/02 ,  F02C9/00 ,  B64D27/10

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

公开(公告)号：US10316758B2

公开(公告)日：2019-06-11

申请号：US14888154

申请日：2014-05-02

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  William G. Sheridan

IPC: F02C7/36 ,  F02K3/06

Abstract: A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between about 219 and 328.

公开(公告)号：US10316757B2

公开(公告)日：2019-06-11

申请号：US14875762

申请日：2015-10-06

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  Robert E. Malecki

IPC: F02K3/06 ,  F02C7/36 ,  F02C3/10 ,  F02K3/068 ,  B64D33/02 ,  F02C3/107

Abstract: A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine configured to drive the geared architecture. The turbine has an exit point, and a diameter (Dt) defined as the radially outer diameter of a last blade airfoil stage in the turbine at the exit point. A nacelle at least partially surrounds a core engine housing. The fan configured to deliver air into a bypass duct is defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point, with a downstream most point of the core engine exhaust nozzle being defined at a distance (Lc or Ln) from the exit point.