Abstract: An assembly is provided for a gas turbine engine. This assembly includes an integrally bladed rotor and a damper. The integrally bladed rotor is rotatable about an axis. The integrally bladed rotor includes a plurality of rotor blades and a rotor disk. The rotor blades are arranged circumferentially around and project radially out from the rotor disk. The rotor disk includes a flange, a groove and a plurality of slots. The groove extends circumferentially around the axis within the flange. The groove projects radially into the flange from an inner side of the flange. The slots are arranged circumferentially about the axis along the groove. Each of the slots projects radially into the flange from the inner side of the flange. The damper is mounted to the rotor disk and seated within the groove.

Abstract: A gas turbine engine includes a compressor section, a combustor and a turbine section. At least one casing surrounds at least one of the compressor and turbine section. The at least one casing is formed of sheet metal. The at least one casing has at least one potential peak displacement point due to vibration across a speed range of the one of the compressor and turbine section. A damper is placed on the casing at the at least one potential peak displacement point. The damper has one end fixed to the casing, and a second end not fixed to the casing and the second end provides an interference fit such that the second end cannot move to a relaxed position of the second end due to the wall of the casing. A method is also disclosed.

Abstract: A gas turbine engine includes a compressor section, a combustor and a turbine section. At least one case surrounds at least one of the compressor section and the turbine section. The case is formed from sheet metal and includes a case wall and at least one potential displacement location due to vibration. At least one retaining ring is assembled to the case and biased against the case wall at the at least one potential displacement location to dampen vibration.

Abstract: A gas turbine engine includes a compressor section, a combustor and a turbine section. At least one casing surrounds at least one of the compressor and turbine section. The at least one casing is formed of sheet metal. The at least one casing has at least one potential peak displacement point due to vibration across a speed range of the one of the compressor and turbine section. A damper is placed on the casing at the at least one potential peak displacement point. The damper has one end fixed to the casing, and a second end not fixed to the casing and the second end provides an interference fit such that the second end cannot move to a relaxed position of the second end due to the wall of the casing. A method is also disclosed.

Abstract: A gas turbine engine under this disclosure could be said to include a compressor section, a combustor and a turbine section. At least one casing surrounds at least one of the compressor and turbine section. The at least one casing is formed of sheet metal, and has at least one potential peak displacement point due to vibration across a speed range of the at least one of the compressor and turbine section. A stiffening element is fixed to a wall of the casing at the at least one potential peak displacement point. A method is also disclosed.

Abstract: A gas turbine engine under this disclosure could be said to include a compressor section, a combustor and a turbine section. At least one casing surrounds at least one of the compressor and turbine section. The at least one casing is formed of sheet metal, and has at least one potential peak displacement point due to vibration across a speed range of the at least one of the compressor and turbine section. A stiffening element is fixed to a wall of the casing at the at least one potential peak displacement point. A method is also disclosed.

Abstract: An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a case, a housing, a component and a spring element. The case includes an aperture that extends axially along an axis through the case. The housing is attached to the case with a cavity formed by and axially between the housing and the case. The component includes a base and a projection. The base is disposed within the cavity and axially engages the case. The projection projects out from the base and axially through the aperture. The spring element is disposed within the cavity. The spring element is compressed axially between and engages the base and the housing.

Abstract: An aircraft engine including a core casing defining a core passage, a nacelle located radially outward from and circumferentially around the casing, and a bypass passage radially defined between nacelle and casing. A mixer mounted to the casing, including a wall having a leading edge attached to casing and defining a plurality of lobes that are circumferentially spaced apart. A damper ring surrounds the mixer at an axial location between leading edge and trailing edge of the wall, and extends circumferentially uninterrupted, and extends axially from first side to second side. The first side being spaced axially away from the leading edge, and the damper ring having an annular ring surface extending between the first and second sides and second side. The damper ring is radially mounted to the mixer wall via an annular ring surface to dampen the wall.

Abstract: An aircraft engine including a core casing extending circumferentially about an axis and defining a core flow passage, a nacelle located radially outward from and around the casing, and a bypass passage defined between the nacelle and the casing. A mixer includes a peripheral wall having a leading edge and a trailing edge, the leading edge attached to the casing. A first mixer portion and a second mixer portion are circumferentially spaced apart and respectively extend away from the leading edge from a first location to a second location. The first and second mixer portions respectively have a first and second stiffness, the first stiffness greater than the second stiffness and greater than a stiffness of an upstream portion of the wall extending from the first mixer portion at the first location toward the leading edge.

Abstract: A stator vane assembly includes a support structure a support structure and a plurality of cantilever stator vanes Each cantilever stator vane of the plurality of cantilever stator vanes extends from the support structure to a distal end of that cantilever stator vane. At least one notch is formed in each of the distal ends. At least one damper is received within the at least one notch formed in each of the distal ends. A method is also disclosed.

Abstract: A stator vane for a gas turbine stator vane stage is provided that includes an airfoil having leading and trailing edges, a vane tip, suction and pressure side surfaces, and at least one aero passage. The leading and trailing edges are chordwise spaced apart. The vane tip is spanwise spaced apart from a radial base end. The suction side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The pressure side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The at least one aero passage extends through the airfoil between the suction and pressure side surfaces, and is disposed proximate and spanwise separated from the vane tip. The stator vane is configured to be cantilevered with the vane tip being unsupported.

Abstract: An orifice pack is provided for delivering pressurized air to a compressor bleed valve of a gas turbine engine. The orifice pack has a diffusion chamber in serial flow communication with a tapering passage and a first outlet passage for venting a first portion of the pressurized air from the diffusion chamber. A second outlet passage branches off from the diffusion chamber at an axial location between the inlet and the tapering passage. The second outlet passage is fluidly connected to the compressor bleed valve for directing a second portion of the pressurized air from the diffusion chamber to the compressor bleed valve.

Abstract: An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a case, a housing, a component and a spring element. The case includes an aperture that extends axially along an axis through the case. The housing is attached to the case with a cavity formed by and axially between the housing and the case. The component includes a base and a projection. The base is disposed within the cavity and axially engages the case. The projection projects out from the base and axially through the aperture. The spring element is disposed within the cavity. The spring element is compressed axially between and engages the base and the housing.

Abstract: Methods and systems for filtering pressurized air used to control a compressor bleed valve of a gas turbine engine are provided. One method comprises receiving the pressurized air in a conduit via an inlet of the conduit, releasing a first portion of the pressurized air out of the conduit via an outlet of the conduit, releasing a second portion of the pressurized air from the conduit via a port disposed between the inlet and the outlet of the conduit, directing the second portion of the pressurized air from the port to a filter along an upward flow path, filtering the second portion of the pressurized air using the filter, and directing the second portion of the pressurized air from the filter toward the compressor bleed valve.

Abstract: The gas turbine engine can have a pneumatic actuator; an intake device secured to a gas path wall delimiting the gas path, the intake device having a tubular body protruding from the gas path wall into the gas path and an inlet aperture formed in the tubular body, the inlet aperture spaced-apart from the gas path wall and facing downstream relative a flow orientation of the gas path, the intake device having an internal conduit extending from the inlet aperture, along the tubular body, to an outlet across the gas path wall; and a fluid line fluidly connecting the outlet of the intake device to the pneumatic actuator.

Abstract: The gas turbine engine can have a pneumatic actuator; an intake device secured to a gas path wall delimiting the gas path, the intake device having a tubular body protruding from the gas path wall into the gas path and an inlet aperture formed in the tubular body, the inlet aperture spaced-apart from the gas path wall and facing downstream relative a flow orientation of the gas path, the intake device having an internal conduit extending from the inlet aperture, along the tubular body, to an outlet across the gas path wall; and a fluid line fluidly connecting the outlet of the intake device to the pneumatic actuator.

Abstract: An orifice pack is provided for delivering pressurized air to a compressor bleed valve of a gas turbine engine. The orifice pack has a diffusion chamber in serial flow communication with a tapering passage and a first outlet passage for venting a first portion of the pressurized air from the diffusion chamber. A second outlet passage branches off from the diffusion chamber at an axial location between the inlet and the tapering passage. The second outlet passage is fluidly connected to the compressor bleed valve for directing a second portion of the pressurized air from the diffusion chamber to the compressor bleed valve.

Abstract: Methods and systems for filtering pressurized air used to control a compressor bleed valve of a gas turbine engine are provided. One method comprises receiving the pressurized air in a conduit via an inlet of the conduit, releasing a first portion of the pressurized air out of the conduit via an outlet of the conduit, releasing a second portion of the pressurized air from the conduit via a port disposed between the inlet and the outlet of the conduit, directing the second portion of the pressurized air from the port to a filter along an upward flow path, filtering the second portion of the pressurized air using the filter, and directing the second portion of the pressurized air from the filter toward the compressor bleed valve.

Abstract: A turbomachine airfoil element comprises an airfoil having: an inboard end; an outboard end; a leading edge; a trailing edge; a pressure side; and a suction side. A span between the inboard end and the outboard end is 1.75-2.20 inches. A chord length at 50% span is 1.05-1.35 inches. At least two of: a first mode resonance frequency is 2400±10% Hz; a second mode resonance frequency is 4950±10% Hz; a third mode resonance frequency is 7800±10% Hz; a fourth mode resonance frequency is 8700±10% Hz; and a fifth mode resonance frequency is 12500±10% Hz.

Abstract: A system comprises a pipe defining a fluid passage, the pipe having a first rate of thermal expansion. A housing defines an opening for receiving an end of the pipe for fluid circulation between the fluid passage and an interior of the housing, the housing having a second rate of thermal expansion lesser than the first rate of thermal expansion, at least one annular gap defined between a periphery of the opening and the end of the pipe when the system is below a safety condition threshold temperature. Seal(s) seal the annular gap, wherein the pipe and the housing are configured and the first and second rates of thermal expansion are selected so that, when the system exceeds the safety condition threshold temperature, the end of the pipe contacts the periphery of the opening by thermal expansion to seal the annular gap independent of the at least one seal.