Abstract: A center plug for attenuating noise in a gas turbine engine includes an inner skin, the inner skin having a substantially cylindrical shape and extending along an axial centerline; an outer skin positioned radially outside the inner skin; a forward bulkhead disposed proximate a forward end of the inner skin, the forward bulkhead connected to and extending radially outward from the inner skin; an aft bulkhead disposed proximate an aft end of the inner skin, the aft bulkhead connected to and extending radially outward from the inner skin; and a first resonator cavity disposed within a volume defined by the inner skin, the outer skin and the forward bulkhead and the aft bulkhead, the first resonator cavity including a first perforated disk extending between the inner skin and the outer skin and forming a first sub-cavity and a second sub-cavity.

Abstract: An acoustic liner includes a core including a plurality of cells extending between a first side of the core and a second side of the core opposite the first side of the core. The acoustic liner further includes a back skin including a light-reflecting layer. The light-reflecting layer includes a first side attached to the first side of the core. The acoustic liner further includes a face skin. The face skin includes a first side attached to the second side of the core.

Abstract: Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The ceramic matrix composite shell comprises a fibrous preform. The fibrous core has a greater porosity than the fibrous preform. A method of making the ceramic matrix component is also disclosed.

Abstract: An aerospace component, for example, used in a gas turbine engine, includes the following structurally-integrated layers: a metallic layer and a composite layer having reinforcing fibers embedded in a matrix material. The aerospace component can also include an insulating layer disposed between the metallic layer and the composite layer where the insulating layer has a thermal conductivity that is lower than a thermal conductivity of the composite layer.

Abstract: A mixing system for a power generation system. The power generation system includes a rotary machine, an exhaust processing system, and a duct system. The rotary machine is configured to produce an exhaust stream. The exhaust processing system is positioned to receive and process the exhaust stream. The duct system is oriented to channel an air stream to the exhaust processing system and to channel the exhaust stream from the rotary machine to the exhaust processing system. The mixing system is within the duct system. The mixing system includes a plurality of supports, a plurality of links extending between at least two of the supports, and at least one wrap circumscribing at least two of the links. The at least one wrap is oriented to change an effective direction of momentum of the exhaust stream and the air stream.

Abstract: A gas turbine engine including: a tail cone; a low pressure compressor; a low pressure turbine; a low speed spool interconnecting the low pressure compressor and the low pressure turbine; and an electric generator located within the tail cone, the electric generator being operably connected to the low speed spool; and one or more acoustic panel assemblies located around the electric generator, the one or more acoustic panel assemblies including: an acoustic backing sheet, an outer porous liner, and an acoustic liner interposed between the outer porous liner and the acoustic backing sheet.

Abstract: An ejector comprises a primary nozzle having an annular wall forming part of an outer boundary of an exhaust portion of a primary flow path of a gas turbine engine. The annular wall has a downstream end defining a plurality of circumferentially distributed lobes. The ejector further comprises a secondary nozzle having an annular wall disposed about the primary nozzle, the primary nozzle and the secondary nozzle defining a secondary flow passage therebetween for channeling a secondary flow. The secondary nozzle defines a mixing zone downstream of an exit of the primary nozzle. A flow guide ring is mounted to the primary nozzle lobes. The ring has an aerodynamic surface extending from a leading edge to a trailing edge respectively disposed upstream and downstream of the exit of the primary nozzle. The aerodynamic surface of the ring is oriented to guide the high velocity primary flow into the mixing zone.

Abstract: An acoustic liner includes a first face sheet, a second face sheet spaced from the first face sheet, and a plurality of sidewalls extending between the first face sheet and the second face sheet. The plurality of sidewalls defines a plurality of cells. Each cell of the plurality of cells defines a cavity between the first face sheet and the second face sheet. A bulk absorber is disposed within at least one cell of the plurality of cells. The bulk absorber further defines the cavity of the at least one cell of the plurality of cells. The first face sheet defines a plurality of apertures extending through a thickness of the first face sheet. Each aperture of the plurality of apertures is aligned with a respective cell of the plurality of cells.

Abstract: An aircraft nacelle including a conduit rigidly connected to an air inlet and a conduit rigidly connected to an engine, a link between the conduit rigidly connected to the air inlet and the conduit rigidly connected to the engine. The link includes a sleeve that receives one end of a first conduit and is rigidly connected to a second conduit, and linking elements that link the sleeve and the first conduit and that are distributed about the entire circumference of the nacelle, each linking element having a threaded stem oriented in a radial direction and being configured to enable tightening from inside the first conduit.

Abstract: A gas turbine engine including: a tail cone; a low pressure compressor; a low pressure turbine; a low speed spool interconnecting the low pressure compressor and the low pressure turbine; an electric generator located within the tail cone, the electric generator being operably connected to the low speed spool; a structural support housing at least partially enclosing the electric generator, the structural support housing being located within the tail cone; and a mounting system located within the tail cone between the structural support housing and the tail cone, wherein the mounting system attaches the tail cone to the structural support housing.

Abstract: Systems and methods are presented for acoustic dampening in a rotating machine. The rotating machine has a rotatable shaft defining an axis of rotation and a gas flowpath. A system comprises an acoustic panel affixed to an annular casing. The acoustic panel comprises an acoustic treatment member extending between a radially inner skin and radially outer skin. The radially inner skin extends the full axial and circumferential dimensions of the acoustic panel, and the acoustic treatment member overlays the entirety of the radially inner skin. The acoustic panel is positioned so that the radially outer skin abuts the casing, the abutting surfaces being configured to effect relative axial movement between the surfaces while maintaining contact between the surfaces. The acoustic panel is affixed in position to the casing by one or more fasteners passing through the casing and a portion of the radially outer skin.

Abstract: An exhaust system for an engine includes an exhaust nozzle located adjacent an outlet end of the engine to receive a primary flow of exhaust gasses expelled from the engine, an inlet opening formed between the exhaust nozzle and the outlet end of the engine through which a secondary flow is provided to the exhaust nozzle, and a vortex generator arranged within the exhaust system at a position where both the primary flow and the secondary flow are present. The vortex generator interrupts at least one of the primary flow and the secondary flow.

Abstract: Embodiments of the present invention are directed to devices adapted to enhance propulsion (i.e., propulsion enhancing devices) of watercraft such as, for example, personal watercraft and the like. To this end, a propulsion enhancing device in accordance with the disclosures made herein may be attached to or integral with (e.g., formed unitarily with a housing thereof) a propulsion unit of a watercraft. The propulsion unit generates a stream of water that provides for propulsion of the watercraft. A propulsion enhancing device in accordance with the disclosures made herein includes internal structures that enhance velocity and/or volumetric attributes of a stream of water generated by the propulsion unit by transforming non rotational flow at the inlet of the propulsion enhancing device to rotational flow at the outlet of the propulsion enhancing device.

Abstract: An acoustic treatment panel intended to be disposed on at least one wall of a turbojet engine in contact with a fluid flow, the panel including a first acoustically reflective plate, a second plate and a plurality of cavities mounted between the first plate and the second plate and including a plurality of cells. The second plate is a one-piece plate through which a plurality of channels pass, each opening out on the one hand onto a first orifice formed on a first face of the second plate and on the other hand onto a second orifice formed on a second face of the second plate, the length of each channel extending between its first orifice and its second orifice being greater than the thickness of the second plate.

Abstract: An assembly method for an air intake of a jet engine of an aircraft, comprising providing an acoustic panel, an outside panel, flanges and a collar, arranging a rear edge of the acoustic panel on an assembly chassis. The flanges and the collar are attached at the rear edge of the acoustic panel. A rear edge of the outside panel is attached to the attached collar. A first set of geometric data of a front edge of the acoustic panel is measured and a second set of geometric data of a front edge of the outside panel is measured. A U-shaped lip as a single element is produced based on the two sets of geometric data. An inside leg of the lip is attached to the front edge of the acoustic panel and an outside leg of the lip is attached to the front edge of the outside panel.

Abstract: An assembly for the rear of a bypass turbomachine (10) comprises a primary nozzle (11) comprising a trailing edge defining a primary flow path portion and a secondary nozzle (110) defining a secondary flow path portion, defined about a longitudinal axis (X), said secondary nozzle being configured to eject a mixture of the flows coming from a secondary flow path (Vs) and from a primary flow path (Vp) of the turbomachine (10), the secondary nozzle being of convergent-divergent shape with a throat (112) corresponding to a minimum cross section of the nozzle (110), the secondary nozzle (110) comprising, at the throat (112), a periodic succession of lobes (116, 118) which are situated along the internal circumference of the secondary nozzle (110). The assembly also comprises a lobed mixer (130) at the downstream end of the primary nozzle (11), this having an alternation of hot lobes (134) extending inside the secondary flow path and of cold lobes (132) extending inside the primary flow path.

Abstract: An exhaust device for a combined cycle power plant, includes a diffuser and a plenum. The diffuser includes a first wall, a second wall, a diffuser inlet, a diffuser outlet, and a diffuser flow path. The first and second walls extend circumferentially about a centerline axis of the exhaust device. The second wall is spaced from the first wall. The diffuser flow path is defined between the first and second walls, and extends from the diffuser inlet to outlet. The plenum includes an inlet wall portion and a non-circular plenum outlet, where the inlet wall portion is coupled to the diffuser outlet. The non-circular plenum outlet is spaced from the diffuser outlet along an axial direction of the exhaust device.

Abstract: The application discloses a noise attenuating system (100, 200, 300, 400, 700, 800) for use with a positive airway pressure system providing a flow of gas, comprising: an expansion chamber (140, 240, 340, 434, 740) having a volume; an intake tube (115, 215, 315, 415, 715, 815) having an inlet (117, 217, 317, 410, 717, 812) and outlet port or portion separated by a length, wherein a portion of the inlet port or portion extends outside of the expansion chamber; and either: a noise attenuator (110, 120, 170, 210, 220, 223, 224, 226, 228, 232, 270, 320, 411, 413, 422, 423, 424, 425, 426, 428, 431, 720, 722) having a bottom and protruding sidewall forming a cavity, wherein the noise attenuator is positioned near the inlet portion of the intake tube such that a portion of the intake tube extends into the cavity of the noise attenuator; or: an acoustic deflector (110, 120, 170, 210, 220, 223, 224, 226, 228, 270, 232, 320, 411, 413, 422, 423, 424, 425, 426, 428, 431, 720, 722) positioned near the outlet port or porti

Abstract: Aspects of the disclosure are directed to a sound-absorbing exhaust center plug for an aircraft gas turbine engine comprising a center plug adapted for attachment within an exhaust nozzle of the engine. The center plug comprises an outer skin, and at least one cavity within the center plug and the at least one cavity extending between a forward bulkhead and an aft bulkhead, where the aft bulkhead is canted inwardly. The outer skin includes a plurality of outer skin openings providing an acoustic pathway through the outer skin and into the at least one cavity.

Abstract: An exhaust mixer for a gas turbine engine where each outer lobe has at the downstream end a circumferential offset in a direction corresponding to that of the swirl component of the flow entering the mixer. The mixer has a crest line having at least a downstream portion curved with respect with respect to a circumferential direction of the mixer and/or a center line at the downstream end tilted with respect to a radial line extending to the tip of the outer lobe to define the circumferential offset. A method of mixing a core flow and a bypass flow surrounding the core flow with an annular mixer is also provided.

Abstract: A nacelle inlet structure for an aircraft propulsion system. This nacelle inlet structure includes an inlet lip, a bulkhead and a reinforcement structure. The inlet lip extends circumferentially about an axial centerline. The bulkhead extends circumferentially about the axial centerline. The bulkhead is configured with the inlet lip to form a cavity axially between the inlet lip and the bulkhead. The reinforcement structure extends circumferentially about the axial centerline. The reinforcement structure is connected to and extends axially between the inlet lip and the bulkhead thereby radially dividing the cavity into an inner sub-cavity and an outer sub-cavity.

Abstract: A transition piece can include: an inner transition piece providing a gas channel; a plurality of cooling rings disposed on the inner transition piece; a plurality of film holes formed on the inner transition piece; and a plurality of impingement holes formed on the plurality of cooling rings, wherein the plurality of film holes are arranged to correspond to the plurality of cooling rings. Each of the plurality of cooling rings includes a cavity such that the plurality of film holes connect the cavity to the gas channel and the plurality of impingement holes connect the cavity to an outside of the inner transition piece.

Abstract: A combustor is provided with a combustion liner (12) in which a combustion gas flows through the inner peripheral side and a plurality of through-holes (14) are formed, and a sound attenuator (20) having a space-forming member (21) configured to form a space (S) connected to the plurality of through-holes (14) at the outer peripheral side of the combustion liner (12). The space-forming member has a main body having an opening configured to bring an outside and the space in communication with each other, and a lid joined to an outer surface of the main body via a welded portion and configured to protrude outward and close the opening.

Abstract: Systems and methods for an exhaust system for a gas turbine engine are provided. The method includes: receiving a primary exhaust fluid in a primary manifold; receiving a secondary cooling fluid in a secondary plenum; directing the primary exhaust fluid through a plurality of eductor primary flow paths associated with a respective plurality of ducts spaced circumferentially about a primary manifold, the plurality of eductor primary flow paths extending through the secondary plenum; and mixing the primary exhaust fluid and the secondary cooling fluid in a mixing chamber associated with a respective one of the plurality of ducts to create a mixed fluid flow.

Abstract: A cooling system for an aircraft has at least one moveable member configured to cover an opening formed within an aircraft outer skin. An actuator moves the at least one moveable member between a fully open position where external atmosphere air can be directed through the opening to an internal passage enclosed by the aircraft outer skin and a fully closed position where the opening is covered. A controller selectively controls the actuator to move the at least one moveable member between the fully open and fully closed positions. An aircraft engine and a method of cooling an aircraft engine in an aircraft are also disclosed.

Abstract: An acoustic liner can comprise a support layer including a set of partitioned cavities defining a set of cells with open faces, as well as a first and second facing sheet operably coupled to the support layer that overlie and close the open faces. A set of aerogel fillings can be provided within at least some of the set of cells.

Abstract: An aircraft nacelle air intake structure including an air intake lip with a U-shaped section to ensure air is shared between an inner face and an outer face. An acoustic panel is on the intake lip rear on the side of the inner face level with a first assembly point. An outer panel is on the intake lip rear on the side of the outer face level with a second assembly point. A front reinforcing frame on the intake lip inside is fixed to the air intake lip on the side of the inner face level with a first fixing point and on the side of the outer face level with a second fixing point. The length of the chord between the stop point and each fixing point is between ? and ? of the total chord between the stop point and the assembly point.

Abstract: A blade containment system includes a plurality of circumferentially-arranged rotatable blades. Each blade has a blade compliance. An annular containment structure is arranged around the rotatable blades. The containment structure includes a liner that has a liner compliance. The blade compliance and the liner compliance are configured such that a strain induced on a respective one of the blades upon impact with the liner is less than a threshold critical strain beyond which the rotatable blades fracture.

Abstract: A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.

Abstract: A fan nacelle for an aircraft gas turbine engine. The nacelle includes an aft nacelle portion including a radially outer surface and a radially inner surface, the radially outer and inner surfaces defining an internal cavity therebetween. The nacelle further includes an aft nacelle segment translatable along a translation vector having an axial component, wherein the aft nacelle segment is configured to translate between a forward deployed position in which the nacelle defines a first primary fan nozzle exit area (A1) and a clean position in which the nacelle defines a second primary fan nozzle exit area (A2) less than the first primary fan nozzle exit area A1, wherein in the forward deployed position, the aft nacelle segment is at least partly located within the internal cavity.

Abstract: An exhaust stack assembly includes an exhaust stack having an internal surface that defines an interior of the exhaust stack. The exhaust stack is configured to receive exhaust gas from at least one turbomachine component and exhaust the exhaust gas to atmosphere. The exhaust gas assembly further includes a plurality of attenuation assemblies disposed in the interior, each of the plurality of attenuation assemblies including a base substrate generally oriented in the direction of flow of the exhaust gas through the interior, each of the plurality of attenuation assemblies further including a plurality of attenuation modules mounted to the base substrate. Each of the plurality of attenuation modules includes a fiber mesh. The fiber mesh is exposed to the exhaust gas in the interior.

Abstract: A sealing apparatus includes a sealing portion configured to be positioned between a platform of a turbine blade and an inner shroud of a turbine vane of a turbine, and an acoustic damper having a closed space portion which is configured to communicate, via an opening, with a disc cavity configured to be positioned between the turbine blade and the turbine vane.

Abstract: Acoustic structures in which acoustic septa are located in the cells of a honeycomb for reducing the noise generated from a source. The honeycomb used to form the acoustic structure has walls that contain convex and concave contours which make the honeycomb flexible. The acoustic septa are formed by inserting planar acoustic inserts into the honeycomb cells to form a septum cap which is friction-locked within the cell and then permanently bonded in place. The planar acoustic septum is configured to match the unique shape of the cell contours to provide desired friction-locking when inserted into the cells and desired acoustic properties after being permanently bonded in place.

Abstract: A sound attenuation apparatus includes a body and a stacked cavity structure. The structure includes a structure inlet formed in the top surface of the body; a passage extending from the structure inlet into the interior of the body to a base surface within the body; a first cavity with a first cavity inlet fluidly coupled to the passage and being formed by a first arm, a first side wall within the interior of the body, and the base surface; and a second cavity with a second cavity inlet fluidly coupled to the passage and being formed by a second arm, the first side wall, and the first arm. Each of the first and second arms extends from the first side wall and includes at least two arm segments oriented at different angles relative to the first side wall longitudinal axis.

Abstract: An ejector/mixer for a gas turbine engine includes an annular wall having upstream end adapted to be fastened to an engine case and a downstream end forming a plurality of lobes. A support member interconnects the lobes, and includes an annular blade located radially inwardly of the bight of the lobes. The lobes extend radially inwardly downstream relative to the annular wall and the support member includes an annular blade and has spaced apart joint surfaces spaced apart to coincide with the joint surfaces of a respective lobes. The spaced-apart joint surfaces of the support member being profiled to mate with the corresponding joint surface of the lobes.

Abstract: A composite component comprising fibres embedded in a matrix material. A plurality of pins extend through the matrix material. Each pin of the plurality of pins comprises a hollow bore such that the bore of each pin defines a plurality of holes in the composite component.

Abstract: An exhaust section of a gas turbine engine includes an exhaust plug defining a plurality of exhaust plug apertures circumferentially spaced from each other. Also included is an exhaust nozzle radially offset from the exhaust plug defining an exhaust pathway between the exhaust plug and the exhaust nozzle. Further included is an exhaust plug liner having a non-uniform outer surface axially aligned with the exhaust plug apertures. The exhaust plug liner is rotatable relative to the exhaust plug between a first position and a second position to selectively change a cross-sectional area of the exhaust pathway during thrust reversal operation to reduce an amount of reverse thrust necessary.

Abstract: Methods and systems for restricting movement in a flow mixer including a plurality of lobes that each includes a trough, using a stiffener mechanism, are provided. The stiffener mechanism includes a first end portion configured to be coupled to a first lobe of the plurality of lobes and a second end portion configured to be coupled to a second lobe of the plurality of lobes. The stiffener mechanism also includes a body portion coupled between the first end portion and the second end portion such that the body portion is configured to restrict movement of the first and second lobes.

Abstract: An acoustically dampened gas turbine engine (10) having a gas turbine engine combustor (12) with an acoustic damping resonator system (14) is disclosed. The acoustic damping resonator system (14) may be formed from one or more resonators (16) formed from a resonator housing (18) positioned within the gas turbine engine combustor (12) at an outer housing (20) forming a combustor basket (22) and extending circumferentially within the combustor (12). In at least one embodiment, the resonator housing (18) may include resonator chambers (26) that may be welded in place within resonator chamber (26) receivers (24) but easily replaceable without exposing the resonator housing (18) to damage. In another embodiment, an inner surface (32) of the resonator chamber (26) may be offset radially outward from an inner surface (34) of the resonator housing (18), thereby creating a flow-path discontinuity and reducing heating of the resonator chamber (26).

Abstract: A system includes a silencer baffle which mounts in a fluid conduit along a fluid flow path, where the silencer baffle includes a plurality of baffle sections. At least two baffle sections of the plurality of baffle sections couple together via mating interlock structures, and at least one baffle section of the plurality of baffle sections has first and second baffle portions which couple together via mating joints.

Abstract: A gas turbine engine is provided having a fan case and a translating sleeve positioned downstream from the fan case. A flow channel extends between the fan case and the translating sleeve. The flow channel includes an inner diameter and an outer diameter. A structural guide vane is positioned within the flow channel and extends from the inner diameter to the outer diameter. A liner is positioned between an aft end of the fan case and an aft end of the translating sleeve to reduce vibratory stress on the structural guide vane.

Abstract: A gas turbine engine includes a fan, a nacelle arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow from the fan. A nozzle associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position.

Abstract: A method of making a light weight component is provided. The method including the steps of: forming a metallic foam core into a desired configuration; applying an external metallic shell to an exterior surface of the metallic foam core after it has been formed into the desired configuration; and attenuating the component to a desired frequency by forming a plurality of openings in the external metallic shell.

Abstract: A damping system and apparatus are disclosed for dampening acoustic pressure oscillations of a gas flow in a combustor of a gas turbine engine having at least one combustor with a combustor liner. A second inner cap portion is disposed on the at least one combustor inner liner. The second inner cap portion can have a hot surface, a cold surface, at least one burner opening protruding from the cold surface, and at least one neck ring having an internal opening and protruding from the cold surface. At least one resonating tube having a resonating tube neck is integrated with and protruding from the at least one neck ring. The at least one resonating tube is disposed between adjacent burner openings, and is configured such that the radial dimension is greater than or equal to the axial dimension.

Abstract: A damping system and apparatus are disclosed for dampening acoustic pressure oscillations of a gas flow in a combustor of a gas turbine engine having at least one combustor with a combustor liner. A second inner cap portion is disposed on the at least one combustor. The second inner cap portion can have a hot surface, a cold surface, at least one burner opening protruding from the cold surface, and at least one neck ring having an internal opening and protruding from the cold surface. At least one extended resonating tube having a resonating tube neck is integrated with and protruding from the at least one neck ring. The at least one extended resonating tube is disposed between adjacent burner openings, and is configured such that the radial dimension is less than the axial dimension.

Abstract: A mixer of a bypass turbine aeroengine according to one embodiment, includes circumferential inner and outer flow surfaces in a wavy configuration to form a plurality of lobes of the mixer. The mixer has an upstream end portion of sheet metal with a first thickness and a downstream end portion of sheet metal with a second thickness less than the first thickness.

Abstract: A gas turbine engine has a surface cooler having a cooler member defining a fluid passage and at least partially contained in an airfoil shaped flow guide member. The cooler member is disposed within a bypass duct and supported on one of the duct walls. The flow guide member provides a smooth outer surface to guide a main portion of a bypass air stream passing over the surface cooler, and defines an inner air channel between the flow guide member and that supporting one of the duct walls for a secondary portion of the bypass air stream to pass through the inner air channel.

Abstract: A turbocharger turbine having housing walls defining a diffuser. Within the diffuser, a center body within the diffuser is supported by de-swirl vanes extending from the diffuser wall. The center body forms a de-swirl passageway having an increasing mean diameter of flow from an upstream end of the center body to a leading edge of the de-swirl vanes. A trailing edge of the de-swirl vanes is near the downstream end of the center body. Annular-type guide vanes surround the center body within the de-swirl passageway. A wastegate system is configured to vent wastegate flow into the diffuser through injection ports on the de-swirl vanes, annular guide vanes and/or center body.

Abstract: The invention relates to a damper for reducing the pulsations in a chamber of a gas turbine. The damper includes a resonator cavity and a neck in flow communication with the resonator cavity and the chamber. The neck includes a mouth to communicate with the chamber. The air flow inside the chamber flows across the mouth of the neck. The neck is so configured that the longitudinal axis of its mouth is angled 0-90° relative to the direction of the air flow inside the chamber. The damper of this invention may effectively alleviate the detrimental effect of the grazing flow and thus it enables the placement of the damper also at locations where strong grazing flows are present.

Abstract: A Helmholtz damper for a combustor of a gas turbine includes an enclosure defining a damping volume from which a neck portion extends and which has a flow path (F) for cooling and purging air with an inlet opening and an outlet opening to the enclosure. The outlet opening is formed in the neck portion. A seal is arranged at the neck portion adjacent to the outlet opening for cooling and purging air such that a cooling effect of the seal is provided.