Abstract: A screen for an aeronautical gas turbine engine includes a plurality of first filaments and a plurality of second filaments. The plurality of second filaments define a plurality of openings with the plurality of first filaments. The plurality of first filaments and the plurality of second filaments define a first arrangement in which the plurality of first filaments are perpendicular to the plurality of second filaments. Upon application of a force to the screen, at least some of the plurality of first filaments and at least some of the plurality of second filaments are configured to translate from the first arrangement to a second arrangement in which a size of at least one of the openings is configured to change from a first size to a second size.

Abstract: A nacelle for a gas turbine engine includes an interior portion along a radial direction of the gas turbine engine and a heat exchanger. The interior portion defines a heat exchanger duct including an inlet and an outlet. The heat exchanger is disposed in the heat exchanger duct between the inlet and the outlet. The inlet is disposed aft of the outlet along a central axis of the nacelle.

Abstract: A screen for an aeronautical gas turbine engine includes a plurality of first filaments and a plurality of second filaments. The plurality of second filaments define a plurality of openings with the plurality of first filaments. The plurality of first filaments and the plurality of second filaments define a first arrangement in which the plurality of first filaments are perpendicular to the plurality of second filaments. Upon application of a force to the screen, at least some of the plurality of first filaments and at least some of the plurality of second filaments are configured to translate from the first arrangement to a second arrangement in which a size of at least one of the openings is configured to change from a first size to a second size.

Abstract: A fan assembly is provided for a gas turbine engine including: a plurality of fan blades, the plurality of fan blades including a first fan blade; and an actuation assembly including: a first linkage connected to the first fan blade; a first pivot point rotatable with the first fan blade, the first linkage further connected to the first pivot point; a control point moveable relative to the first pivot point and connected to the first linkage for changing a relative position of the first fan blade within the plurality of fan blades.

Abstract: Methods, apparatus, and articles of manufacture to measure pressure of a fluid are disclosed. An example apparatus to measure a total pressure of a fluid includes a probe head including a plurality of ports, the plurality of ports fluidly coupled to a measurement chamber, and a pressure sensor operatively coupled to the measurement chamber to measure the total pressure inside the measurement chamber, the total pressure corresponding to fluid flow through one of the plurality of ports.

Abstract: A fan assembly is provided for a gas turbine engine including: a plurality of fan blades, the plurality of fan blades including a first fan blade; and an actuation assembly including: a first linkage connected to the first fan blade; a first pivot point rotatable with the first fan blade, the first linkage further connected to the first pivot point; a control point moveable relative to the first pivot point and connected to the first linkage for changing a relative position of the first fan blade within the plurality of fan blades.

Abstract: A screen for an air intake portion of a machine is provided. The screen includes an assembly of screen members that form a plurality of screen cells, wherein at least a portion of the screen cells define an irregular configuration.

Abstract: An apparatus for providing foreign object debris protection an air intake of an aircraft engine. The apparatus includes a frame and a plurality of cross-members. The cross-members are positioned in the frame to define a plurality of screen openings. At least one of the cross-members has an aerodynamically efficient cross section.

Abstract: A bell-mouth scoop assembly includes an actuator comprising a plurality of hinge members configured to rotate in unison about a respective hinge axis of rotation from a first stowed position to a second deployed position and at least one linkage arm extending outwardly from at least one of the plurality of hinge members. The bell-mouth scoop assembly further comprises a bell-mouth panel comprising a panel longitudinal centerline and pivotably coupled to each linkage arm, in the first stowed position the bell-mouth panel (configured to conform to an outer surface of the with the panel longitudinal centerline aligned about a circumference of the flow discharge nozzle, in the second deployed position the bell-mouth panel configured to extend away from the outer surface of the flow discharge nozzle with the longitudinal centerline aligned parallelly with the nozzle centerline.

Abstract: A gas turbine engine that includes a nozzle assembly that has features that facilitate airflow into and through a bypass passage of the gas turbine engine during a reverse thrust operation is provided. The nozzle assembly of the gas turbine engine also includes features that increase the effectiveness of the thrust reverse system of the gas turbine engine. Methods for reversing the thrust of a gas turbine engine are also provided.

Abstract: A compression inner core cowl for a jet engine can have a concave annular section of the core cowl. The concave section can begin within the expanse of an outer fan cowl and extend aft of the fan cowl toward the tail cone, defining an annular bypass section between the fan cowling and the core cowling. The concave geometry of the core cowling reduces the strength of supersonic shock waves generated at the corners of the cowlings adjacent a supersonic airflow stream to increase overall efficiency of the engine.

Abstract: An aspect of the present disclosure is directed to a housing and rotatable shaft apparatus. The apparatus includes a static structure, a rotor assembly, and a bearing assembly. The bearing assembly includes a first static race coupled to the static structure, a first rotatable race coupled to the rotor assembly, a first rolling bearing element disposed in contact between the first static structure and the first rotatable race, a support member coupled to the first static race, a second static race disposed within and coupled to the support member, and a second rotatable race disposed within the support member. The second rolling bearing element is disposed between the second static race and the second rotatable race.

Abstract: A bell-mouth scoop assembly includes an actuator comprising a plurality of hinge members configured to rotate in unison about a respective hinge axis of rotation from a first stowed position to a second deployed position and at least one linkage arm extending outwardly from at least one of the plurality of hinge members. The bell-mouth scoop assembly further comprises a bell-mouth panel comprising a panel longitudinal centerline and pivotably coupled to each linkage arm, in the first stowed position the bell-mouth panel (configured to conform to an outer surface of the with the panel longitudinal centerline aligned about a circumference of the flow discharge nozzle, in the second deployed position the bell-mouth panel configured to extend away from the outer surface of the flow discharge nozzle with the longitudinal centerline aligned parallelly with the nozzle centerline.

Abstract: A gas turbine engine that includes a nozzle assembly that has features that facilitate airflow into and through a bypass passage of the gas turbine engine during a reverse thrust operation is provided. The nozzle assembly of the gas turbine engine also includes features that increase the effectiveness of the thrust reverse system of the gas turbine engine. Methods for reversing the thrust of a gas turbine engine are also provided.

Abstract: An aircraft engine assembly can include an engine having an exhaust turbine with drive shaft and exhaust collector, a propeller rotationally coupled to the drive shaft and defining a rotational axis, and an exhaust stub to direct exhaust gases away from the aircraft.

Abstract: An apparatus and method for an exhaust assembly for a turbine engine including a reverse flow portion. The exhaust assembly can include an exhaust stub extending between an exhaust collector and an exhaust outlet to define an exhaust conduit having the reverse flow portion. A flow of combustion gases can exit the engine and travel through the exhaust assembly. The exhaust assembly can be applied to a turboprop engine.

Abstract: An apparatus for providing foreign object debris protection an air intake of an aircraft engine. The apparatus includes a frame and a plurality of cross-members. The cross-members are positioned in the frame to define a plurality of screen openings. At least one of the cross-members has an aerodynamically efficient cross section.

Abstract: A vent nozzle configured to minimize adverse flow interactions between merging gas flows that have two different speeds. The nozzle includes a first wall that defines an outer surface and an inner surface. A first flowpath is positioned on one side of the first wall such that the first flowpath is adjacent to the outer surface. A second flowpath is positioned on another side of the first wall such that the second flowpath is adjacent to the inner surface. A second wall is spaced-apart from the first wall and that defines a portion of the second flowpath and an extension of the second wall extends beyond the first wall. The extension of the second wall approaches an imaginary line that is defined by an extension of the outer surface.

Abstract: A turbofan exhaust nozzle apparatus includes: an inner shell having a centerline axis, the inner shell disposed inside a surrounding outer shell to define an annular flow duct therebetween terminating in an outlet at a trailing edge of the outer shell, the inner shell terminating in a trailing edge aft of the outlet; and a pylon interrupting circumferentially the duct; wherein an aft portion of the outer shell is non-axisymmetric around a common centerline axis of the nozzle, and the trailing edge of the outer shell is vertically non-coplanar.

Abstract: A compression inner core cowl for a jet engine can have a concave annular section of the core cowl. The concave section can begin within the expanse of an outer fan cowl and extend aft of the fan cowl toward the tail cone, defining an annular bypass section between the fan cowling and the core cowling. The concave geometry of the core cowling reduces the strength of supersonic shock waves generated at the corners of the cowlings adjacent a supersonic airflow stream to increase overall efficiency of the engine.