Raytheon Technologies Corporation

United Technologies Corporation

Rtx Corporation

United Technologies Coproration

Carey J Clum

Dominic J Mongillo, Jr

Timothy J Jennings

Christopher Whitfield

Dennis M Moura

Thomas N Slavens

Tracy A Propheter-Hinckley

Evan P Molony

San Quach

Nicholas M LoRicco

Sasha M Moore

Charles Thistle

John McBrien

Clifford J Musto

Brandon W Spangler

Mark F Zelesky

Christopher W Robak

Michael J Birnkrant

James Thompson Beals

Matthew A Devore

Timothy S Snyder

Joel H Wagner

Robin Prenter

Ricardo Trindade

Adam P Generale

Weina Li

Jonathan Perry Sandoval

Mark Borja

Timothy Charles Nash

Wolfgang Balzer

Shayan Ahmadian

Jose L Santana

Tracy A Propheter-hinckley

Jr Dominic J Mongillo

Nicholas M Loricco

Dominic J Mongillo

Nicholas M Lo Ricco

Thomas N Stevens

According to one embodiment, a gas turbine engine component assembly is provided. The gas turbine engine component assembly comprising: a first component having a first surface and a second surface opposite the first surface, the first component including a cooling hole extending from the second surface to the first surface through the first component, wherein the second surface of the first component is oriented relative to a first airflow path such that airflow in the first airflow path separates from the second surface of the first component; and a first fairing secured to the first component proximate the second surface of the first component, the first fairing being configured to redirect airflow in the first airflow path such that the airflow exits the first fairing oriented parallel with the second surface of the first component.

Apparatus and method for mitigating airflow separation around engine combustor

Core assemblies for manufacturing airfoils and airfoils made therefrom are described. The core assemblies having a leading edge hybrid skin core positioned relative to a plurality of core bodies and configured to define a leading edge cavity at a leading edge of the manufactured airfoil. The leading edge hybrid skin core extends from a root region toward a tip region in a radial direction, the leading edge hybrid skin core extends above at least one of the plurality of core bodies to define an exit in a tip region of the manufactured airfoil, and the leading edge hybrid skin core has a height-to-width ratio of about 0.8 or less.

Leading edge hybrid cavities for airfoils of gas turbine engine

A dual wall liner for a gas turbine engine may comprise a shell having a first side and a second side, a panel contacting the shell, the panel at least partially defining a hot gas path through which a hot gas flows, wherein the first side of the shell faces the panel, wherein the shell includes a thermal barrier coating (TBC) disposed on the first side of the shell. The TBC may thermally protect the shell from heat from a hot gas path.

Panel burn through tolerant shell design

Airfoils bodies having a first core cavity and a second core cavity located within the airfoil body that is adjacent the first core cavity. The second core cavity is defined by a first cavity wall, a second cavity wall, a first exterior wall, and a second exterior wall, wherein the first cavity wall is located between the second core cavity and the first core cavity and the first and second exterior walls are exterior walls of the airfoil body. The first cavity wall includes a first surface angled toward the first exterior wall and a second surface angled toward the second exterior wall. At least one first cavity impingement hole is formed within the first surface and a central ridge extends into the second core cavity from at least one of the first cavity wall and the second wall and divides the second core cavity into a two-vortex chamber.

Airfoil having internal hybrid cooling cavities

A gas turbine engine component comprises a body having a leading edge, a trailing edge, and a radial span. One internal channel in the body provides an upstream supply pressure. Another internal channel in body receives the upstream supply pressure and provides a downstream supply pressure. At least one axial rib separates an internal area adjacent to the trailing edge into a plurality of individual cavities. At least one pressure regulating feature is located at an entrance to at least one individual cavity entrance to control downstream supply pressure to the trailing edge. Exits formed in the trailing edge communicate with an exit pressure. The rib and pressure regulating features cooperate such that the downstream supply pressure mimics the exit pressure along the radial span.

Trailing edge pressure and flow regulator

An airfoil may include an airfoil body that defines a skin chamber, a skin chamber outlet opening, and a tip flag chamber. In various embodiments, the skin chamber is in fluidic communication with the tip flag chamber via the skin chamber outlet opening. In various embodiments, the airfoil body further defines an outlet hole disposed on at least one of a trailing edge and a pressure side of the airfoil. The tip flag chamber may be in fluidic communication with the outlet hole (e.g., cooling circuit air in the tip flag chamber may exit the airfoil via the outlet hole). In various embodiments, the tip flag chamber extends parallel and directly adjacent to the outermost tip of the airfoil.

Airfoil cooling circuit

An airfoil has a pressure side wall and a suction side wall circumferentially spaced apart from the pressure side wall. Each radially extends between a root region and a tip region, and each axially extends between a leading edge and a trailing edge. The airfoil also has a gusset, a squealer pocket, and a tip pocket. The gusset extends between the pressure side wall and the suction side wall. The squealer pocket is formed in the tip region and is disposed between the leading edge, the gusset, the pressure side wall, and the suction side wall. The tip pocket is formed in the tip region and is spaced apart from the squealer pocket. The tip pocket is disposed between the trailing edge, the gusset, the pressure side wall, and the suction side wall.

Tip cooling design

A gas turbine engine component assembly is provided. The gas turbine engine component assembly, comprising: a first component having a first surface, a second surface opposite the first surface, and a cooling hole extending from the second surface to the first surface through the first component; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween in fluid communication with the cooling hole for cooling the second surface of the second component; and a particulate capture device attached to at least one of the first component and the second component, the particulate capture device configured to aerodynamically separate the airflow from the particulate.

Apparatus and method for mitigating particulate accumulation on a component of a gas turbine

A component for a gas turbine engine includes an outer surface bounding a hot gas path of the gas turbine engine, and a cooling passage configured to deliver a cooling airflow therethrough. The cooling passage includes a passage wall located opposite the outer surface to define a component thickness and a plurality of protrusions located along the passage wall. Each protrusion has a protrusion height extending from the passage wall and a protrusion streamwise width extending along the passage wall in a flow direction of the cooling airflow through the cooling passage. One or more cooling holes extend from the passage wall to the outer surface. A cooling hole inlet of a cooling hole is located at the passage wall, in a protrusion wake region downstream of a protrusion of the plurality of protrusions.

Film cooling hole arrangement for gas turbine engine component

A gas turbine engine component assembly comprising: a first component having a first surface and a second surface opposite the first surface, wherein the first component includes a cooling hole extending from the second surface to the first surface; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween; and a lateral flow injection feature integrally formed in the first component and fluidly connecting a flow path located proximate to the second surface of first component to the cooling channel, the lateral flow injection feature being configured to direct airflow from the airflow path through a passageway and into the cooling channel at least partially in a lateral direction parallel to the second surface of the second component such that a cross flow is generated in the cooling channel.

A gas turbine engine component assembly is provided. The gas turbine engine component assembly comprising: a first component having a first surface and a second surface; a threaded stud including a first end and a second end opposite the first end, the threaded stud extending from the second surface of the first component; and a faired body operably secured to the threaded stud, wherein the faired body is shaped to redirect the airflow in a lateral direction parallel to the second surface of the first component such that a cross flow is generated.

A component includes a component body. The component further includes a first passage disposed in the component body. The first passage includes a first end and a second end opposite the first end. The component further includes a second passage. The second passage extends from the second end of the first passage. The second passage includes a turn. The component further includes a third passage. The third passage extends from the second end of the first passage. The component further includes a first projection extending from a passage surface of the component body within the first passage. The first projection is disposed between the first and the second end of the first passage and is configured to direct debris transiting the first passage away from the second passage and into the third passage.

Component having a dirt tolerant passage turn

A gaspath component for a gas turbine engine includes a platform. A body extends outward from the platform and includes at least one internal cooling passage. The internal cooling passage includes at least one particulate redirection feature defined at an end of the body opposite the platform. The at least one particulate redirection feature includes a first face oblique to and facing an expected flow of fluid through the internal cooling passage and defines at least one corresponding opening.

Turbine blade tip dirt removal feature

A combustor may comprise a combustion chamber encapsulated by a combustor casing, wherein the combustor casing comprises a casing forward portion, a casing radially inward portion, a casing radially outward portion, and a casing aft portion. The casing radially inward portion may comprise a dirt collector system comprising a hot panel adjacent to the combustion chamber; a cold shell panel coupled to the hot panel such that the hot panel is between the combustion chamber and the cold shell panel; and a dirt collector panel coupled to the cold shell panel such that the cold shell panel is between the hot panel and the dirt collector panel. The cold shell panel and the dirt collector panel at least partially define a dirt collecting space.

Dirt collector system

A gas turbine engine component comprising: a first component having a receiving aperture extending; a second component having a first surface and a second surface; and a passageway portion including a first end, a second end opposite the first end, and an outer surface extending from the second end to the first end, the passageway portion extending from the second surface of the second component through the receiving aperture of the first component, wherein the outer surface of the passageway portion and the first component define a gap therebetween, the gap fluidly connecting airflow in an airflow path proximate the second surface of the first component to the cooling channel, wherein the gap is configured to direct the airflow along the outer surface and the outer surface is shaped to redirect the airflow in a lateral direction parallel to the second surface such that a lateral airflow is generated.

Growing community of inventors

West Hartford, CT, United States of America

Carey J Clum