United Technologies Corporation

Rtx Corporation

Raytheon Technologies Corporation

Jason Elliott

Byron R Monzon

Xuetao Li

William R Graves

Michael M Joly

Christopher S McKaveney

Myron Klein

Ethan C Drew

Daniel C Falk

Bronwyn Power

Christopher Beaudry Miller

Gabriel L Suciu

Christopher Marc Dye

Tahany Ibrahim El-Wardany

Edward J Gallagher

Anthony R Bifulco

Thomas J Robertson, Jr

Darin S Lussier

Darryl Whitlow

David P Houston

Becky E Rose

Agnieszka M Wusatowska-Sarnek

Allan R Penda

James R Murdock

Shaoluo L Butler

George Alan Salisbury

Scott C Billings

James D Campbell, Jr

James H Moffitt

Royce E Tatton

Barclay Bingham Young, Jr

Jay Thomas Abraham

Po-Ching Yeh

Jun Jiang

Agnieszka M Wusatowska-Samek

Adam Cade Wilcox

Ryan Clifford Cox

A gas turbine engine is provided and includes a first fan blade including a suction surface, a second fan blade comprising a pressure surface and neighboring the first fan blade and a throat region interposed between the suction surface of the first fan blade and the pressure surface of the second fan blade. The throat region includes a passage throat located at a minimum distance between the pressure and suction surfaces. The first and second fan blades are configured such that a pre-compression region is defined in the throat region ahead of the passage throat. Each of the first and second fan blades includes a mean camber line defining a flattened suction surface.

Fan blade or vane with improved bird impact capability

A blade includes an airfoil section and a root section that extends in a chord-wise direction between leading and trailing root faces and in a radial direction that is perpendicular to the chord-wise direction from an inner root face to the second end of the airfoil section. The root section defines a dovetail cross-section that is perpendicular to the chord-wise direction. The airfoil section and the root section are comprised of a composite material architecture that includes inter-section fiber plies that each extend from the leading to the trailing edge and from the airfoil section into the root section. Wedges extend in the root section and are interleaved with the inter-section fiber plies. Each of the wedges tapers in the chord-wise direction and the radial direction.

Composite blade with tapered root section

A turbomachine airfoil element includes an airfoil that has pressure and suction sides spaced apart from one another in a thickness direction and joined to one another at leading and trailing edges. The airfoil extends in a radial direction a span that is in a range of 27.5-28.5 inches (699-724 mm). A chord length extends in a chordwise direction from the leading edge to the trailing edge at 50% span and is in a range of 13.0-14.0 inches (330-356 mm). The airfoil element includes at least two of a first mode with a frequency of 34&#xb1;10% Hz, a second mode with a frequency of 90&#xb1;10% Hz, a third mode with a frequency of 178&#xb1;10% Hz, a fourth mode with a frequency of 222&#xb1;10% Hz, a fifth mode with a frequency of 307&#xb1;10% Hz and a sixth mode with a frequency of 389&#xb1;10% Hz.

Gas turbine engine airfoil frequency design

In one exemplary embodiment, a turbofan engine comprises a fan section. A core section includes a turbine section arranged fluidly downstream from the compressor section. A combustor is arranged fluidly between the compressor and turbine sections. The fan and core sections are configured to produce a thrust in a range 27,000-35,000 pounds-f (120,102-156,688 N). An airfoil is arranged in the fan section. The airfoil has first and second modes each having a frequency. The first mode has the lowest frequency, and the second mode has the second lowest frequency wherein the second mode frequency is 140 Hz or less at a redline engine speed.

A dovetail root of a fan blade configured for operation within the fan assembly of a gas turbine engine. The dovetail root is imparted with a compressive residual stress layer along the outer faces of the dovetail root preventing crack formation within the dovetail root when the gas turbine engine is in operation. To further protect and structurally enhance the dovetail root, a wear covering is disposed on the dovetail root. The wear covering may consist of a composite laminate layer that is bonded to the metallic core of the dovetail root providing additional cracking protection of the dovetail root when the gas turbine engine is in operation.

Fan blade root

A turbojet fan blade adapted for securement to an engine rotor includes a titanium sheath that extends continuously over its leading edge from blade tip to blade root. The sheath provides impact protection over the leading edge of the blade, and also incorporates enhanced radial securement of the sheath to the rotor. As such, a radial lock system for the sheath is disclosed that includes a fully integrated sheath portion that extends completely over the frontal face of a fan blade dovetail, providing a radial lock of the sheath to the rotor in the event of failure of any adhesive used to attach the sheath to the leading edge. Because the sheath extends continuously from blade tip to the blade root, lightning strike protection is enhanced by avoidance of any discontinuities otherwise associated with related art sheaths that typically extend over only a flow path portion of the leading edge.

Radial lock for fan blade sheath

A fan blade includes a root including a front surface, a rear surface, a first side surface connected to the front surface and the rear surface, and a second side surface connected to the front surface and the rear surface. The front surface engages the first side surface and the second side surface by one or more blunted surfaces, and the rear surface engages the first side surface and the second side surface by one or more blunted surfaces. A blade extends from the root.

Fan blade attachment of gas turbine engine

A gas turbine engine includes a spool, a turbine coupled to drive the spool and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the spool drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 16 of the propulsor blades.

Efficient, low pressure ratio propulsor for gas turbine engines

A fan case of a gas turbine engine includes a fan blade containment section defined about an engine axis and a plurality of helical ribs adjacent to the fan blade containment section.

Energy-absorbing fan case for a gas turbine engine

This disclosure relates to a design methodology used in manufacturing a broaching tool for cutting slots in aerospace disk applications. The method includes modeling geometry of the slot and the broach tool, which is based upon an initial minimum tooth rise that is determined empirically for the particular disk material. The number of broach inserts and teeth per insert is calculated, and the broach inserts are modeled. The stresses and deformation of the lug are calculated in a finite elements environment based upon simulated incremental broach tool movements. If the lug is not within specifications based upon design constraints, then the broach tool is revised and the simulations are repeated until the slot is within the desired specifications.

Broach tool design methodology and systems

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Huntington, IN, United States of America

Jason Elliott