Abstract: Acoustic structures for a gas turbine engine are disclosed. In various embodiments, the acoustic structures comprise a panel in the form of a cellular structure having a plurality of cells; and a housing configured to house the cellular structure, the housing including a first rail configured for attachment to a static structure within the gas turbine engine and a second rail configured for attachment to the static structure. In various embodiments, the acoustic structures comprise a liner in the form of a non-segmented or two-piece structure configured to extend about the inner barrel of a nacelle structure. The disclosed panels and liners are configured to maximize the operational area of the acoustic structures by eliminating obstructions from adjacent flow paths.

Abstract: A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A fan case surrounds the fan rotor. A core engine has a compressor section and includes a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of aft connecting members rigidly connected to the fan case, and to the core engine. A plurality of fan exit guide vanes are rigidly connected to the fan case, with the fan exit guide vanes including structural fan exit guide vanes which are rigidly connected to the core engine, and non-structural fan exit guide vanes, and the non-structural fan exit guide vanes being provided with an acoustic feature to reduce noise.

Abstract: Acoustic structures for a gas turbine engine are disclosed. In various embodiments, the acoustic structures comprise a panel in the form of a cellular structure having a plurality of cells; and a housing configured to house the cellular structure, the housing including a first rail configured for attachment to a static structure within the gas turbine engine and a second rail configured for attachment to the static structure. In various embodiments, the acoustic structures comprise a liner in the form of a non-segmented or two-piece structure configured to extend about the inner barrel of a nacelle structure. The disclosed panels and liners are configured to maximize the operational area of the acoustic structures by eliminating obstructions from adjacent flow paths.

Abstract: A face sheet of an acoustic liner for a gas turbine engine, having: a first surface, a second surface, and a sheet body extending from the first surface to the second surface; a forward end, an aft end, and the sheet body extending from the forward end to the aft end; and a hole formed in the sheet body, extending from the first surface to the second surface, the hole defining a hole opening at the first surface of the face sheet, and an elongated edge that extends aft at the hole opening, wherein the elongated edge defines a flow ramp surface that tapers toward the first surface and is configured to direct flow toward the first surface, wherein the flow ramp surface is a streamlining feature.

Abstract: A face sheet of an acoustic liner for a gas turbine engine, having: a first surface, a second surface, and a sheet body extending from the first surface to the second surface; a forward end, an aft end, and the sheet body extending from the forward end to the aft end; and a hole formed in the sheet body, extending from the first surface to the second surface, the hole defining a hole opening at the first surface of the face sheet, and an elongated edge that extends aft at the hole opening, wherein the elongated edge defines a flow ramp surface that tapers toward the first surface and is configured to direct flow toward the first surface, wherein the flow ramp surface is a streamlining feature.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of H-frame connecting members rigidly connected to the fan case, and to the core engine. The H-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions which are generally parallel to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine including a plurality of radial struts rigidly connected to the fan case, and to the core engine. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of H-frame connecting members rigidly connected to the fan case, and to the core engine. The H-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions which are generally parallel to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

Abstract: A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A fan case surrounds the fan rotor. A core engine has a compressor section and includes a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of aft connecting members rigidly connected to the fan case, and to the core engine. A plurality of fan exit guide vanes are rigidly connected to the fan case, with the fan exit guide vanes including structural fan exit guide vanes which are rigidly connected to the core engine, and non-structural fan exit guide vanes, and the non-structural fan exit guide vanes being provided with an acoustic feature to reduce noise.

Abstract: An assembly for a gas turbine engine has: a fan case with a forward end, an aft end, and an inner surface extending from the forward end of the fan case to the aft end of the fan case; an acoustic liner having a forward end, an aft end, and an acoustic liner body that extends continuously from the forward end of the acoustic liner to the aft end of the acoustic liner; an inlet case located forward of the forward end of the fan case, the forward end of the acoustic liner being secured to the inlet case, and wherein the acoustic liner body is positioned against an inner surface of the inlet case and an inner surface of the fan case, and the aft end of the acoustic liner is secured to the fan case between the forward end and the aft end of the fan case.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine including a plurality of radial struts rigidly connected to the fan case, and to the core engine. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

Abstract: A fluid wash system for a gas turbine engine is disclosed, the gas turbine engine defining an axial direction and comprising a borescope port that provides access to a component within a core flow path of the gas turbine engine. In various embodiments, the fluid wash system includes a wash line fluidly connected to a pump configured to provide a pressurized flow of wash liquid; a spray nozzle connected to the wash line and configured for extending into the borescope port to provide the pressurized flow of wash liquid to the component within the core flow path; and an attachment mechanism configured to releasable mount the spray nozzle to the borescope port, the attachment mechanism including an alignment mechanism configured to orient the spray nozzle and direct the pressurized flow of wash liquid in a predetermined direction toward the component.

Abstract: A system for dampening noise generated by a gas turbine engine is disclosed. In various embodiments, the system includes a fan exit guide vane having a leading edge, a trailing edge and a pocket that extends in a chordwise direction between the leading edge and the trailing edge; and an acoustic panel configured to be received within the pocket, the acoustic panel including a back sheet, a face sheet and a core disposed between the back sheet and the face sheet, the core having a first cavity extending between a first wall and a second wall and a second cavity disposed within the first cavity.

Abstract: A system for retaining stators and reducing air leakage in a gas turbine engine having an axis includes a stator having an inner platform, an outer platform, a low pressure side, a high pressure side, and at least one foot, and designed to turn air. The system also includes a case positioned radially outward from the stator and having at least one recess designed to interface with the at least one foot to resist movement of the stator relative to the case. The system also includes a bladder positioned between the outer platform of the stator and the case and designed to receive pressurized fluid having a greater pressure than ambient pressures experienced at the low pressure side of the stator and to further resist movement of the stator relative to the case in response to receiving the pressurized fluid.

Abstract: A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A fan case surrounds the fan rotor. A core engine has a compressor section and includes a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of aft connecting members rigidly connected to the fan case, and to the core engine. A plurality of fan exit guide vanes are rigidly connected to the fan case, with the fan exit guide vanes including structural fan exit guide vanes which are rigidly connected to the core engine, and non-structural fan exit guide vanes, and the non-structural fan exit guide vanes being provided with an acoustic feature to reduce noise.

Abstract: A passive transpirational flow acoustic liner assembly for a gas turbine engine includes a guide vane assembly and a conduit configured to deliver airflow received from the guide vane. The guide vane assembly includes an airfoil having a transpirational flow acoustic liner. The acoustic liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet and having a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures, a backing sheet having a plurality of apertures and being coupled to the segmented member such that the segmented member is positioned between the face sheet and the backing sheet, and a plenum coupled to the backing sheet opposite the segmented member and fluidly connected to the conduit.

Abstract: An acoustic liner for a gas turbine engine includes an acoustic panel that is curved about a central axis. The acoustic panel includes a support backing, a face sheet, and a cellular structure disposed between the support backing and the face sheet. The face sheet has elongated slots that extend along respective slot centerlines in the plane of the face sheet. The slot centerlines are sloped at oblique angles to the central axis.

Abstract: A non-metallic engine case inlet compression seal for a gas turbine engine includes a non-metallic longitudinal leg section that extends from the non-metallic arcuate interface section and a non-metallic mount flange section that extends from the longitudinal leg section.