United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Yuan Dong

Lisa I Brilliant

Gabriel L Suciu

Frederick M Schwarz

Stanley J Balamucki

Becky E Rose

Sean Nolan

Sanjay S Hingorani

Simon W Evans

Jody Kirchner

Sue-Li Chuang

Robert W Fessenden

Li Xing Pan

William G Sheridan

Brian D Merry

Matthew E Bintz

Dilip Prasad

Steven D Roberts

Sean P McCutchan

Larry W Spires

J Axel Glahn

Andrew J Murphy

Steven J Sirica

Jordan T Wall

Taryn Narrow

Daniel Carminati

Daniel J Monahan

Abhir A Adhate

James T Thoresen

Lucas D Sorensen

Sue-Li Kingsley Chuang

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor that delivers flow to a core flowpath and a compressor section including first and second compressors. The core flowpath passes through the first compressor. The core flowpath in the first compressor has an outer diameter relative to the engine longitudinal axis. The outer diameter has a slope angle relative to the axis.

Compressor flowpath

A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

Repeating airfoil tip strong pressure profile

A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

A rotor blade for a gas turbine engine includes a blade extending from a root and a contoured tip portion at a first end of the blade. The first end is opposite the root. The contoured tip portion includes a first sloped region and a second sloped region. The second sloped region is steeper than the first sloped region, relative to a platform.

Rotor blade for a gas turbine engine

Rotor of a gas turbine engines having a rotor hub and a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at an airfoil root of the blade at the hub and extends to a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade.

Full-span forward swept airfoils for gas turbine engines

A compressor may include an airfoil configured to rotate about an axis. The airfoil may have a span measured radially from a root of the airfoil to a tip of the airfoil. A compressor case may be radially adjacent to the airfoil. The airfoil and the compressor case may define a clearance between the tip of the airfoil and a radially inner surface of the compressor case. A ratio of the clearance to the span may be between 1.5 to 2.5 percent.

Gas turbine engine airfoil

A compressor tip clearance management system includes a compressor section having a low pressure compressor and a high pressure compressor arranged downstream from the low pressure compressor. A variable stator vane is arranged upstream from the high pressure compressor. The variable stator vane is connected to an actuator, and a controller is in communication with the actuator. The controller is configured to provide a command to the actuator to move the variable stator vane in response to a high pressure compressor clearance condition. A high pressure compressor rotor speed is altered, and the high pressure compressor clearance condition is changed to a desired clearance.

Compressor tip clearance management for a gas turbine engine

A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

LPC flowpath shape with gas turbine engine shaft bearing configuration

A fan blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by pressure and suction sides to provide an exterior airfoil surface that extends in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Gas turbine engine fan blade airfoil profile

A blade for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a platform and an airfoil that extends from the platform. The airfoil extends in span between a root and a tip region and in chord between a leading edge and a trailing edge. A sweep angle is defined at the leading edge and a dihedral angle is defined relative to the chord of the airfoil. The sweep angle and the dihedral angle are localized at the tip region and extend over a distance of the airfoil equivalent to about 10% to about 40% of the span.

A method of designing an airfoil for a gas turbine engine according to one embodiment of this disclosure can include localizing a sweep angle at a leading edge of a tip region of the airfoil, and localizing a dihedral angle at the tip region of the airfoil. The dihedral angle can be applied by translating the airfoil in direction normal to a chord of the airfoil.

A turbine engine includes a fan section that rotates about a longitudinal axis and a compressor section. The turbine engine also includes a fan exit stator located between the fan section and the compressor section, the fan exit stator including an airfoil. The airfoil defines an entrance angle with respect to a leading edge of the airfoil and a line parallel to the longitudinal axis, and the airfoil defines an exit angle with respect to a trailing edge of the airfoil and a line parallel to the longitudinal axis. A difference between the entrance angle and the exit angle is between about 45&#xb0; and about 65&#xb0;. The turbine engine also includes a turbine section, and a portion of the compressor section is driven by a portion of the turbine section.

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Glastonbury, CT, United States of America

Yuan Dong