Abstract: A butterfly valve for a conduit defining a passage for a flow of a fluid therethrough in a flow direction. The butterfly valve includes a shaft rotatably mounted to the conduit and defining a longitudinal axis along its length. The butterfly valve includes a valve body coupled to the shaft, such that the valve body is rotatable along with the shaft about the longitudinal axis between a closed position and a fully open position. The valve body includes a first major surface, a second major surface opposite to the first major surface, a perimeter surface, a central plane, a first lobe, a second lobe, and a third lobe.

Abstract: A butterfly valve for a conduit defining a passage for a flow of a fluid therethrough in a flow direction. The butterfly valve includes a shaft rotatably mounted to the conduit and defining a longitudinal axis along its length. The butterfly valve includes a valve body coupled to the shaft, such that the valve body is rotatable along with the shaft about the longitudinal axis between a closed position and a fully open position. The valve body includes a first major surface, a second major surface opposite to the first major surface, a perimeter surface, a central plane, a first lobe, a second lobe, and a third lobe.

Abstract: A thermal management system for a gas turbine engine includes a plurality of inlet and outlet vanes extending from a bypass inner wall towards an outer cowl, a plurality of first removable shims, a plurality of second removable shims, and a third removable shim. Each first removable shim extends between and engages the outer cowl and a corresponding inlet vane. Each second removable shim extends between and engages the outer cowl and a corresponding outlet vane. The third removable shim extends between and engages a heat exchanger and the outer cowl. Each of the plurality of first removable shims, each of the plurality of second removable shims, and the third removable shim is removable for positionally adjusting the outer cowl relative to the bypass inner wall.

Abstract: A pressurised fluid duct failure event characterising method including: monitoring acoustic signals; detecting a failure event occurrence based on an acoustic signal exceeding a first failure threshold, and identifying a failure marker; estimating, for each acoustic signal, a failure detection time, based on the failure marker and a rate of change in sound pressure level for each acoustic signal; determining an acoustic weighted position for acoustic sensors; determining an estimate of a region containing the failure event as a first sub-volume of the duct based on the acoustic weighted position and a predetermined relationship between sound pressure level and regions of the duct; calculating, for each acoustic signal, one or more spectral characteristics; and determining at least one failure characteristic based on a comparison between the one or more spectral characteristics, the estimate of the region containing the failure event, and one or more predefined failure event parameters.

Abstract: A pressurised fluid duct failure event location determination method including: monitoring acoustic signals; detecting a failure event occurrence based on an acoustic signal of those exceeding a first failure threshold, and identifying a failure marker; estimating, for each acoustic signal, a failure detection time; determining an acoustic weighted position for acoustic sensors based on the respective failure detection time for each acoustic signal and the position of each microphone; determining an estimate of a region containing the failure event as a first sub-volume of the duct based on the acoustic weighted position and a predetermined relationship between sound pressure level and regions of the duct; discretising the first sub-volume of the duct into a first three-dimensional grid having voxels according to a first grid mesh size; and determining a first estimate of the failure event location using a time-domain beamforming algorithm applied to the acoustic signals and the first sub-volume.

Abstract: A thermal management system for a gas turbine engine includes a plurality of inlet and outlet vanes extending from a bypass inner wall towards an outer cowl, a plurality of first removable shims, a plurality of second removable shims, and a third removable shim. Each first removable shim extends between and engages the outer cowl and a corresponding inlet vane. Each second removable shim extends between and engages the outer cowl and a corresponding outlet vane. The third removable shim extends between and engages a heat exchanger and the outer cowl. Each of the plurality of first removable shims, each of the plurality of second removable shims, and the third removable shim is removable for positionally adjusting the outer cowl relative to the bypass inner wall.

Abstract: An inlet assembly includes a web extending between a first end and a second end along a first axis. The web includes first and second major surfaces. The inlet assembly includes a load transfer flange at least partially disposed around the web and configured to be fixedly coupled to a gas turbine engine. The inlet assembly includes first vanes spaced apart from each other at least along the first axis. Each first vane extends between the first major surface of the web and the load transfer flange. The inlet assembly includes second vanes spaced apart from each other at least along the first axis. Each second vane extends between the second major surface of the web and the load transfer flange.

Abstract: Disclosed herein is a method of automatically determining an operating condition of at least part of a gas turbine engine 10 for an aircraft, the method comprising: measuring one or more gas pressure waves by a gas pressure detector 401, wherein the gas pressure detector 401 is located in the gas turbine engine 10; and automatically determining, by a computing system, an operating condition of at least part of a gas turbine engine 10 in dependence on an output signal of the gas pressure detector 401.

Abstract: A bleed valve system comprises a duct, featuring a central longitudinal axis and allowing a main flow of fluid to pass from a first environment at a first static pressure to a second environment at a second static pressure along a bleed direction, a valve comprising a valve member arranged within the duct between the first and second environments and movable to partially obstruct the duct and deviate the main flow of fluid to direct at least a part of it towards a portion of an internal wall of the duct; and an ejector, arranged within the duct, downstream of the valve member and offset from the central longitudinal axis in correspondence of said portion of the internal wall, adapted to supply an additional flow of fluid within the duct to accelerate the main flow of fluid and reduce the second static pressure.

Abstract: A louvre offtake arrangement for a gas turbine engine includes a first duct, where a primary flow flows; a second duct, defining an offtake, connected to the first duct at an inlet; and a plurality of louvres arranged at the inlet. At least one louvre protrudes from the second duct into the first duct to divert part of the primary flow towards the second duct.

Abstract: A tip clearance control (TCC) system 100 is provided to control the gap 176 between the tips of turbine blades 172 of a gas turbine engine 10 and the casing 174 within which they rotate. The inlet 120 to the TCC system is provided in a bifurcation panel 110 that extends across a bypass duct 22 of the gas turbine engine 10. The inlet is provided on a first major surface 112 of the bifurcation panel 110 that is defined such that the direction (R) that points through the bifurcation panel from the first major surface 112 to a second major surface 114 corresponds to the fan rotation direction. This provides a particularly effective and/or efficient TCC system 100.

Abstract: A pressure relief arrangement for a gas turbine engine comprises a panel and a plurality of pressure relief mechanisms provided in the panel. The mechanisms have a first configuration and a second configuration. In the first configuration the panel is sealed to prevent fluid flow through the panel in a thickness direction and in the second configuration the mechanisms are arranged so that a plurality of holes are provided in the panel so fluid can flow through the panel in a thickness direction.

Abstract: A gas turbine engine for an aircraft is provided. The engine includes an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The engine further includes core casings surrounding the engine core. The engine further includes an aerodynamic cowl which surrounds the core casings. The engine further includes a propulsive fan located upstream of the engine core, the fan generating a core airflow which enters the core engine and a bypass airflow which enters a bypass duct surrounding the aerodynamic cowl. The engine further includes one or more engine accessories mounted in a space between the core casings and the aerodynamic cowl.

Abstract: A pressure relief arrangement for a gas turbine engine comprises a hinged door and a mount spaced from the door. An expandable structure is connected to the door and to the mount such that when the door hinges open the expandable structure expands.

Abstract: Disclosed herein is a method of automatically determining an operating condition of at least part of a gas turbine engine 10 for an aircraft, the method comprising: measuring one or more gas pressure waves by a gas pressure detector 401, wherein the gas pressure detector 401 is located in the gas turbine engine 10; and automatically determining, by a computing system, an operating condition of at least part of a gas turbine engine 10 in dependence on an output signal of the gas pressure detector 401.

Abstract: There is disclosed a gas turbine engine having a thermal management system, the gas turbine engine comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan located upstream of the engine core. The thermal management system comprises an oil tank; a heat exchanger; an oil coolant circuit that connects the oil tank and the heat exchanger; and an oil pump that pumps oil around the oil coolant circuit. The oil tank is located within the engine core, and the oil pump is electrically driven such that it is operable independently of the core shaft.

Abstract: A modular fluid-driven linear actuator comprising an actuator support comprising a first part and a second part configured for relative linear movement, wherein the actuator support is configured to hold a plurality of removable actuator modules in parallel and in respective module-receiving locations defined between the first part and the second part. At least one actuator module is removably installed in the actuator support and held in a respective module-receiving location of the plurality. A first fluid manifold is configured to provide a drive fluid to each actuator module when received in a respective module-receiving location, to act on a piston of the actuator module having a respective piston area to drive the piston in a first direction. A total piston area of the modular actuator is variable by installing and uninstalling actuator modules in the actuator support.

Abstract: A modular fluid-drive rotary actuator comprising a housing defining a chamber, a rotary driver disposed within the chamber. The rotary driver may comprise a modular driver assembly of at least two drive wheels, and/or the housing may comprise at least a first end part, a second end part and an interchangeable intermediate part disposed between them. There is also disclosed a method of reconfiguring a fluid-driven rotary actuator by replacement, addition or removal of one or more drive wheels or housing parts.

Abstract: A gas turbine engine for an aircraft is provided. The engine includes an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The engine further includes core casings surrounding the engine core. The engine further includes an aerodynamic cowl which surrounds the core casings. The engine further includes a propulsive fan located upstream of the engine core, the fan generating a core airflow which enters the core engine and a bypass airflow which enters a bypass duct surrounding the aerodynamic cowl. The engine further includes one or more engine accessories mounted in a space between the core casings and the aerodynamic cowl.

Abstract: A louvre offtake arrangement for a gas turbine engine includes a first duct, where a primary flow flows; a second duct, defining an offtake, connected to the first duct at an inlet; and a plurality of louvres arranged at the inlet. At least one louvre protrudes from the second duct into the first duct to divert part of the primary flow towards the second duct.