Raytheon Technologies Corporation

United Technologies Corporation

Rtx Corporation

Raytheon Technologies Corporatino

Thomas E Clark

William M Barker

Ken Francis Blaney

Daniel J Whitney

Dean W Johnson

Kathryn S Read

Gary L Grogg

Danielle Mahoney

Brian C McLaughlin

Eric Charles Mundell

Winston Gregory Smiddy

Andrew Joseph Lazur

William Wangard, Iii

Jeremy Drake

Murat Yazici

Michael G McCaffrey

Paul M Lutjen

Marc J Muldoon

Jose R Paulino

Richard K Hayford

Jeffrey T Morton

Carson A Roy Thill

Daniel S Rogers

Craig R McGarrah

Anthony B Swift

Andrew E Breault

Andrew J Lazur

Terence P Tyler

Jeffrey Vincent Anastas

Andrew D Keene

Peter J Milligan

Jung Muk Choe

Jaimie Schweitzer

Dairus Deylan McDowell

Kevin D Tracy

Sudarshan Thirumalai

Bryan Kupferman

Karhryn S Read

William Wangard Iii

Winston G Smiddy

Jr Terence P Tyler

An assembly is provided for an aircraft propulsion system. This assembly includes an open propulsor rotor and an open guide vane structure. The open propulsor rotor is configured to rotate about a rotational axis. The open guide vane structure is axially next to the open propulsor rotor. The open guide vane structure includes a plurality of open guide vanes arranged circumferentially about the rotational axis. The open guide vanes include a first open guide vane. The first open guide vane includes a first vane leading edge, a first vane trailing edge and a fixed first vane chord length measured from the first vane leading edge to the first vane trailing edge. The first open guide vane is configured to translate axially along the rotational axis.

Translatable guide vanes for aircraft propulsion system

A fan is disposed at an intake of a nacelle, and a gas turbine is contained within a first portion of the nacelle. The gas turbine includes compressor, combustor, and turbine stages through which a core stream of gas passes past the fan and past the compressor, combustor, and turbine stages. A first heat exchanger is disposed in a second portion of the nacelle, where a fan stream passes through the second portion of the nacelle. A first switching valve receives gas from the fan stream and directs the received gas to a second switching valve or the turbine stage. The second switching valve directs gas to a boost compressor and receives gas from the first switching valve or a high pressure bleed from the compressor stage. A third switching valve receives gas from the boost compressor and directs compressed gas to the first heat exchanger or the turbine stage.

Boost compressor augmentation for hybrid electric thermal management systems

An aircraft powerplant assembly includes a gearbox and a power coupler. The gearbox includes a first power transfer apparatus and a second power transfer apparatus independent of the first power transfer apparatus. The power coupler includes a clutch. The first power transfer apparatus is configured to operatively couple and transfer mechanical power through the gearbox between a first engine apparatus and an engine accessory. The power coupler is configured to operatively couple a power coupled accessory to the second power transfer apparatus when the clutch is engaged. The second power transfer apparatus and the power coupler are collectively configured to operatively couple and transfer mechanical power through the gearbox between a second engine apparatus and the power coupled accessory when the clutch is engaged. The power coupler is configured to operatively decouple the second power transfer apparatus from the power coupled accessory when the clutch is disengaged.

Aircraft powerplant gearbox with selectable power coupler

A mandrel for a molding process includes a first portion and a second portion that form a mandrel having a channel between the first and second portions. At least one locking pin is configured to urge the first and second portions away from each other to establish a gap in response to inserting the locking pin into the channel along a respective end of the mandrel. A molding apparatus and method of forming a matrix composite component are also disclosed.

Constant cross section mandrel for CMC components

A method includes providing a ceramic matrix composite (CMC) seal arc segment that has radially inner and outer sides, attaching an abradable layer on the radially outer side of the CMC seal arc segment, providing a carrier to support the CMC seal arc segment, the carrier including a ridge, and sliding the CMC seal arc segment relative to the carrier such that during the sliding the ridge cuts a groove into the abradable layer, the ridge remaining disposed in the groove to thereby provide a labyrinth seal that partitions the cooling cavity between the carrier and the CMC seal arc segment into sub-cavities.

Air seal system with backside abradable layer

A fiber-reinforced component for use in a gas turbine engine includes a first braided fiber sleeve forming a cooling channel and a plurality of fiber plies enclosing the first braided fiber sleeve, with the plurality of fiber plies forming first and second walls separated by the first braided fiber sleeve. The fiber-reinforced component further includes a matrix material between fibers of the braided fiber sleeve and the plurality of fiber plies.

CMC component with integral cooling channels and method of manufacture

A gas turbine engine includes a primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is disposed radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features defining a purge airflow. A heat shield is positioned immediately downstream of the first seal and is configured in relation to the first seal such that the purge airflow enters a mixing plenum defined between the heat shield and the first seal.

Gas turbine engine including flow path flex seal and heat shield

A gas turbine engine includes a primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is positioned radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features. A bifurcation feature is positioned downstream of the first seal. The bifurcation feature is configured to direct a portion of an airflow through the first seal into a cooling air passage. A second seal contacts the bifurcation and the first seal, such that fluid radially inward of the bifurcation feature.

Gas turbine engine including flow path flex seal with cooling air bifurcation

A seal system for a gas turbine engine includes a ceramic matrix composite (CMC) seal arc segment and a carrier supporting the CMC seal arc segment. The CMC seal arc segment defines radially inner and outer sides and has an abradable layer disposed on the radially outer side. There is a cooling cavity radially between the carrier and the abradable layer. The carrier includes a ridge that projects into a groove in the abradable layer and provides a labyrinth seal that partitions the cooling cavity into sub-cavities.

A blade outer air seal assembly includes a support structure. A blade outer air seal has a plurality of seal segments arranged circumferentially about an axis and mounted in the support structure by a carrier. The carrier has first and second hooks that extend radially outward from a platform along an axial length of the carrier. The first and second hooks are in engagement with the support structure.

BOAS carrier with dovetail attachments

A method of assembling a blade outer air seal assembly includes engaging a first blade outer air seal with a first attachment surface on a first attachment body. The first blade outer air seal includes a first attachment body passage for accepting the first attachment body. A second blade outer air seal is engaged with a second attachment surface on the first attachment body. The second blade outer air seal includes a second attachment body passage for accepting the first attachment body. Rotation is prevented of the first attachment body relative to the first blade outer air seal with a first post engaging the first blade outer air seal. Rotation is prevented of the first attachment body relative to the second blade outer air seal with a second post engaging the second blade outer air seal.

Gas turbine engine component

A method for forming a ceramic matrix composite (CMC) component with an internal cooling channel includes partially densifying a first fiber preform to form a portion of a final ceramic matrix volume, machining a first channel into a surface of the partially densified first fiber preform, covering the first channel with a fibrous member to form a near net shape fiber preform with an internal passage formed by the first channel and the fibrous member, and densifying the near net shape fiber preform.

Internal cooling circuits for CMC and method of manufacture

A blade outer air seal assembly includes a support structure and a blade outer air seal having a plurality of segments extending circumferentially about an axis and mounted to the support structure. At least two segments have a base portion extending from a first circumferential side to a second circumferential side, a first protrusion extending from the first circumferential side and having a first slot, and a second protrusion extending from the second circumferential side and having a second slot. A feather seal is arranged in the first slot and the second slot between the at least two segments such that the feather seal is pivotable between first and second positions in response to relative movement between the first and second protrusions.

Feather seal for CMC BOAS

Artist's sink liner with flexible drain retention locking ring

Growing community of inventors

Sanford, ME, United States of America

Thomas E Clark