Abstract: A gas turbine engine includes a bypass duct extending about a longitudinal centerline of the gas turbine engine. The bypass duct includes at least one bypass duct wall defining at least a portion of a bypass flow path through the bypass duct. The at least one bypass duct wall includes a scoop extending into the bypass flow path. The gas turbine engine further includes a bleed conduit including an inlet connected to the bypass duct within the scoop of the at least one bypass duct wall and at least one louver mounted to the bleed conduit within the inlet. The at least one louver extends between a leading edge and a trailing edge opposite the leading edge. The leading edge is located within the bypass flow path and the trailing edge is located within the bleed conduit.

Abstract: A system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The takeoff conduit projects out from the flowpath wall. The porous cover for the internal conduit passage is disposed at the takeoff port.

Abstract: A multi-engine aircraft includes first and second gas turbine engines for providing motive power to the aircraft, each of the first and second engines comprising an air system, an air pressure line extending between and fluidly connecting the air systems of the first and second engines, and a pressure wave damper communicating with the air pressure line extending between the first and second engines.

Abstract: A gas turbine engine includes a bypass duct extending about a longitudinal centerline of the gas turbine engine. The bypass duct includes at least one bypass duct wall defining at least a portion of a bypass flow path through the bypass duct. The at least one bypass duct wall includes a scoop extending into the bypass flow path. The gas turbine engine further includes a bleed conduit including an inlet connected to the bypass duct within the scoop of the at least one bypass duct wall and at least one louver mounted to the bleed conduit within the inlet. The at least one louver extends between a leading edge and a trailing edge opposite the leading edge. The leading edge is located within the bypass flow path and the trailing edge is located within the bleed conduit.

Abstract: A method for controlling noise emitted by a hybrid-electric gas turbine engine for an aircraft during a takeoff flight condition includes applying a first total rotational force to a shaft with a turbine and an electric motor. The first total rotational force includes a first electric rotational force applied by the electric motor and a first thermal rotational force applied by the turbine. The first total rotational force has a first rotational force ratio of the first electric rotational force to the first thermal rotational force. The method further includes controlling the noise emitted by the gas turbine engine by reducing the first rotational force ratio from an initial rotational force ratio of the rotational force ratio as an altitude of the aircraft increases and maintaining the first total rotational force substantially constant while reducing the rotational force ratio.

Abstract: There is provided a turbofan engine having: a fan rotatable about an axis within a case; a first acoustic treatment lining a first portion of the case upstream of the fan; and a second acoustic treatment lining a second portion of the case downstream of the fan, the first acoustic treatment having different noise-attenuating characteristics relative to the second acoustic treatment.

Abstract: A system has: a combustor; a plenum surrounding the combustor; a transfer tube having an inlet fluidly connected to the plenum and at least two outlets, a first flow passageway defined between the inlet and a first outlet, a second flow passageway defined between the inlet and a second outlet, the second flow passageway connected to a discharge region outside of the plenum; a flow valve disposed within the second flow passageway and operable between an open position and a closed position, in the open position the flow valve fluidly connects the plenum with the discharge region, in the closed position the flow valve blocking fluid communication between the plenum and the discharge region; and a controller communicatively coupled to the flow valve to control operation thereof by: causing the flow valve to open for a time period; and subsequent to the time period, causing the flow valve to close.

Abstract: Sandwich-structured noise-attenuating and/or structural panels and methods of manufacturing such panels using ultrasonic welding are described. The method includes: receiving a backing member, a sheet and a cellular structure; assembling the cellular structure between the backing member and the sheet; and ultrasonically welding the backing member and the sheet together.

Abstract: A system has: a combustor; a plenum surrounding the combustor; a transfer tube having an inlet fluidly connected to the plenum and at least two outlets, a first flow passageway defined between the inlet and a first outlet, a second flow passageway defined between the inlet and a second outlet, the second flow passageway connected to a discharge region outside of the plenum; a flow valve disposed within the second flow passageway and operable between an open position and a closed position, in the open position the flow valve fluidly connects the plenum with the discharge region, in the closed position the flow valve blocking fluid communication between the plenum and the discharge region; and a controller communicatively coupled to the flow valve to control operation thereof by: causing the flow valve to open for a time period; and subsequent to the time period, causing the flow valve to close.

Abstract: A method for operating a gas turbine engine comprises providing fuel flow and compressed airflow to a combustor with a fuel-to-air ratio, the compressed airflow being from a compressed air source; detecting at least one parameter indicative of the fuel-to-air ratio being below a predetermined value; and bleeding compressed air from the compressed air source when the at least one parameter indicative of the fuel-to-air ratio is below the predetermined value to increase the fuel-to-air ratio to at least the predetermined value.

Abstract: An aircraft component and an associated method of manufacturing such aircraft component are disclosed. The aircraft component comprises a backing member, a porous sheet spaced apart from the backing member, and a cellular structure disposed between the backing member and the porous sheet. The cellular structure is attached to at least one of the backing member and the porous sheet by one or more non-adhesive connections.

Abstract: The gas turbine engine includes a fluid system fluidly connecting at least two components of the gas turbine engine, and a tunable resonator in fluid flow communication with the fluid system. The tunable resonator has a resonating volume that varies as a function of a volume of an inflatable member located inside the tunable resonator. The inflatable member having a means for varying the volume of the inflatable member, to thereby tune the resonating volume to a selected frequency of pressure fluctuations or acoustic waves within the fluid system.

Abstract: There is provided an acoustic treatment for an aircraft engine. The acoustic treatment has: a facing sheet configured to line a gas path portion of the aircraft engine having in use an airflow passing over the facing sheet, the facing sheet having a thickness and having perforations extending through the thickness; and a backing member spaced from the facing sheet to define a cavity between the facing sheet and the backing member, a ratio of the thickness of the facing sheet to an effective diameter of the perforations is less than or equal to 0.22.

Abstract: There is provided a turbofan engine having: a fan rotatable about an axis within a case; a first acoustic treatment lining a first portion of the case upstream of the fan; and a second acoustic treatment lining a second portion of the case downstream of the fan, the first acoustic treatment having different noise-attenuating characteristics relative to the second acoustic treatment.

Abstract: A multi-engine aircraft includes first and second gas turbine engines for providing motive power to the aircraft, each of the first and second engines comprising an air system, an air pressure line extending between and fluidly connecting the air systems of the first and second engines, and a pressure wave damper communicating with the air pressure line extending between the first and second engines.

Abstract: A bleed air system is disclosed for a turbofan engine having a bypass duct. The bleed air system includes a bleed air conduit having an inlet end connected to the bypass duct and a bleed air flow control valve disposed downstream of the inlet end, and a louver assembly disposed within the opening at inlet end of the bleed air conduit. The louver assembly includes an airfoil-shaped louver blade which extends across the opening in a direction transverse to the flow direction within the bypass duct. A reinforcing rib is attached to the louver blade and extends across the louver assembly in a direction generally parallel to the flow direction.

Abstract: Herein provided are methods and systems for reducing an acoustic signature of a gas turbine engine. An acceleration command for the engine is received. In response to receiving the acceleration command: a fuel flow to the engine is increased for a first predetermined time period; subsequent to the first predetermined time period, the fuel flow to the engine is reduced for a second predetermined time period; and subsequent to the second predetermined time period, the fuel flow to the engine is increased for a third time period.

Abstract: The gas turbine has a plurality of inlet guide vanes each having a blade having a leading edge, a trailing edge, a span extending along the leading edge, and pivot members at opposite ends of the leading edge, and being pivotally mounted across the radial-to-axial intake via the pivot members, the pivot axis extending axially across a radial portion of the radial-to-axial intake. The plurality of inlet guide vanes include a plurality of first inlet guide vanes, and a plurality of second inlet guide vanes, the second inlet guide vanes having a trailing edge recess differing from the corresponding portion of the first inlet guide vanes. During operation, when the inlet guide vanes are pivoted past a given angle toward the tangential orientation, a radial flow of gas is allowed through the trailing edge recesses to avoid or impede vortex whistle.

Abstract: An aircraft component and an associated method of manufacturing such aircraft component are disclosed. The aircraft component comprises a backing member, a porous sheet spaced apart from the backing member, and a cellular structure disposed between the backing member and the porous sheet. The cellular structure is attached to at least one of the backing member and the porous sheet by one or more non-adhesive connections.

Abstract: A method for operating a gas turbine engine comprises providing fuel flow and compressed airflow to a combustor with a fuel-to-air ratio, the compressed airflow being from a compressed air source; detecting at least one parameter indicative of the fuel-to-air ratio being below a predetermined value; and bleeding compressed air from the compressed air source when the at least one parameter indicative of the fuel-to-air ratio is below the predetermined value to increase the fuel-to-air ratio to at least the predetermined value.