United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

John T Ols

Steven D Porter

Steven Wayne Burd

Stephen Karl Kramer

Paul K Sanchez

Richard N Allen

Albert K Cheung

Caroline Karanian

Sandra S Pinero

James B Hoke

Jorge I Farah

William Arthur Sowa

John S Tu

Jonathan Lemoine

Meggan Harris

Wangen Lin

Kyle C Lana

Fumitaka Ichihashi

Jeffrey Jay Lienau

Renee J Jurek

Karl A Mentz

Robert Newman

Giovanni Whitman

William G Askey

Reid Dyer Curtis Smith

Irving Segalman

Rajendra Prasad Uppara Allabanda

Sheree R Swenson-Dodge

Dae K Dempsey

Reid D Smith

Venu Sambarapu

Prabaharan Manoharan

Carolina Karanian

Reid DC Smith

A brush seal system includes a component including a first mating surface. The brush seal system further includes a brush seal including a brush seal backing plate, a retaining ring, and a plurality of bristles retained between the brush seal backing plate and the retaining ring. The brush seal backing plate includes a second mating surface mounted to the first mating surface. The brush seal system further includes at least one cooling channel extending from an exterior side of the component to an interior side of the component so as to bypass the brush seal.

Combustor to vane sealing assembly and method of forming same

Flow diverters for installation in mid-turbine frame systems at a conduit outlet of gas turbine engines are described. The flow diverters include a diverter body having a connector portion defining a diverter inlet, a diverter extension at least partially defining a diverter outlet, and a curved portion arranged between the connector portion and the diverter extension, the curved portion configured to change a direction of flow from a first direction to a second direction that is about 90&#xb0; from the first direction as the flow passes from the diverter inlet to the diverter outlet.

Flow diverter for mid-turbine frame cooling air delivery

A vane has an airfoil extending between a radially outer platform and a radially inner platform. At least one of the platforms has nominally radially thinner portions, and a pad defining a radially thicker portion. The pad has a radial thickness that is greater than a thickness of the nominal radially thinner portions. The pad surrounds an outer periphery of the airfoil on a side of the radially outer platform. The pad has a varying radial thickness. A mid-turbine frame and a gas turbine engine are also disclosed.

Turbine vane with platform pad

A vane pack for a gas turbine engine includes an annular arrangement of vanes. A ring is secured around the vanes and extends proud of an axial end of the vanes.

Gas turbine engine turbine vane ring arrangement

A gas-turbine engine is provided. The gas-turbine engine comprises a high pressure turbine with an aft blade platform. A static structure may be disposed aft of the high pressure turbine and proximate a cavity defined by the aft blade platform. A vane of the static structure may have a vane platform with a shaped tip extending into the cavity.

Hardware geometry for increasing part overlap and maintaining clearance

A heat shield panel for a gas turbine engine combustor may comprise a perimetrical rail extending at least partially around a perimeter of the heat shield panel and defining a main backside cavity, a plurality of air admittance holes, each air admittance hole extending along a hole axis, and a midrail dividing the main backside cavity into a first backside cavity and a second backside cavity, wherein at least a portion of the midrail is axially offset from at least one of the hole axes.

Interrupted midrail for combustor panel

A vane for use in a gas turbine engine has an airfoil extending between a leading edge and a trailing edge, and is generally hollow. The vane has a radially outer platform and a radially inner platform. At least one of the radially inner and radially outer platforms has thicker portions circumferentially located beyond one of the leading edge and the trailing edge, and thinner portions circumferentially beyond circumferential edges of the airfoil. A mid-turbine frame and a gas turbine engine are also disclosed.

Turbine vane with non-uniform wall thickness

A turbine vane has an airfoil extending between an inner platform and an outer platform. The airfoil is hollow. A hollow within the airfoil extends between an inner leading edge to an inner trailing edge. An inner radius is defined at the inner trailing edge in a circumferential direction measured between walls which extend from the inner leading edge to the inner trailing edge. The inner radius varies along a radial length defined between the inner and outer platforms. A mid-turbine frame and a gas turbine engine are also disclosed.

Turbine vane with variable trailing edge inner radius

A gas-turbine engine is provided comprising a high-pressure turbine, a low-pressure turbine aft of the high-pressure turbine, and a housing around the low-pressure turbine and the high-pressure turbine. A stator may be disposed between the low-pressure turbine and the high-pressure turbine with a boss formed on the stator. A bushing may be retained in the boss with a pin coupled to the housing and partially within the bushing. A method of mounting a full-ring stator is also provided, the method including the steps of inserting a bushing into a boss of the full-ring stator by press fitting, and inserting a pin into the bushing of the full-ring stator.

Mid-turbine frame stator with repairable bushing and retention pin

A method is provided for reworking a component. The method includes at least partially filling a cavity in a non-fusion weldable base alloy with a multiple of layers of a multiple of laser powder deposition spots formed of a filler alloy. Each of the multiple of laser powder deposition spots at least partially overlaps at least one of another of the multiple of laser powder deposition spots. The filler alloy may be different than the non-fusion weldable base alloy.

Laser powder deposition weld rework for gas turbine engine non-fusion weldable nickel castings

An interface within a gas turbine engine includes a sealing surface defined by a portion of a vane platform. A seal is in contact with said sealing surface. A barrier is transverse to the sealing surface.

Flow discourager for vane sealing area of a gas turbine engine

An airfoil extends between radially inner and radially outer platforms. The airfoil extends between a leading edge and a trailing edge, and merges into facing surfaces of the radially inner and outer platforms. A variable fillet merges a facing surface of one of the radially inner and outer platforms into a wall of the airfoil. The variable fillet has a length extending away from a surface of the airfoil and a height extending away from the facing surface of at least one of the radially inner and outer platforms outwardly on to the airfoil. The variable fillet has a greater length at one of the leading edge and the trailing edge. A spaced portion has a shorter length at locations spaced from at least one of the leading and trailing edges. A mid-turbine frame and a gas turbine engine are also disclosed.

Turbine vanes with variable fillets

A unitary one-piece hub has first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members. The windows include a lip configured to surround the cross members. A gas turbine engine and a method of providing a hub for a gas turbine engine are also disclosed.

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Northborough, MA, United States of America

John T Ols