Abstract: In one embodiment, a turbine system includes a number of sensors, each of the number of sensors disposed in a respective location of the turbine system, and a controller including a memory storing one or more processor-executable routines and a processor. The processor configured to access and execute the one or more routines encoded by the memory wherein the one or more routines, when executed cause the processor to receive one or more signals from the number of sensors during any stage of operation of the turbine system, and convert the one or more signals to audio output.

Abstract: A dual engine generator has two engines and two inverters and two alternators configured within an inner enclosure. An outer enclosure extends around the inner enclosure to create plenums for airflow. Cooling airflow enters through the outer cover and flows through the plenums and into the inner enclosure to cool the components therein. Airflows out of the inner enclosure and into the plenums and finally exhausts from the outer enclosure. The outer enclosure may have rounded sides to form a pill shaped enclosure. The outer covers are configured to deflect or flex in high wind condition and protect the inner enclosure. The engines are configured above the alternator and inverter and a flow of air reduces heat from the engines from overheating the other components. The mufflers are within the plenums and a flow of exhaust air from within the inner enclosure is directed over the mufflers.

Abstract: A gas turbine engine exhaust nozzle comprises a housing having an aft end that terminates in a row of chevrons. At least one surface of the housing has scalloped root regions proximate bases of adjacent chevrons. The scalloped root regions have a reduced thickness relative to the rest of the aft end.

Abstract: The present invention relates to acoustical panels. As to provide a cabin panel to be made as simple as possible and with optimized acoustical properties, an acoustical cabin panel (10) is provided that comprises a first (12) and a second layer (14) which are spaced apart from each other by an intermediate layer (16) and altogether with a plurality of reinforcement elements (22) form a core composite component (24). Furthermore, the panel comprises at least one intermediate layer (18) made of a core material (20) which is arranged in the intermediate layer. The reinforcement elements extend from the first to the second layer through the at least one intermediate film and are tension and pressure resilient in a finished state of the panel. The reinforcement elements can be at least partially infiltrated with a curable matrix material and at least one prepreg (26) soaked with matrix material is provided in the first and/or second layer.

Abstract: A method and system for a compartment baffle system are provided. The compartment baffle system includes a first baffle portion extending along an axial length of a casing, the first baffle portion extending radially outwardly from an outer surface of the casing and a second baffle portion extending at least partially radially outwardly from the first baffle portion. The second baffle portion further includes a radially inner edge supported by the first baffle portion and a radially outer edge extending proximate an inner surface of a radially outer cowl. The compartment baffle system also includes a gap extending between the radially outer edge and the inner surface.

Abstract: A propulsion system for a supersonic aircraft includes an engine including an engine core and an engine bypass, a compression surface upstream of the engine, a shroud surrounding the engine configured to direct airflow passing over the compression surface towards the engine, and a plurality of vortex generators positioned upstream of the engine. The vortex generators have a height such that when the supersonic aircraft is flown at a predetermined altitude and predetermined speed, the plurality of vortex generators create vortices that propagate partially outside of a boundary layer formed proximate a surface of a supersonic inlet. The vortices cause a high-velocity portion of the airflow to move towards the engine core and a low-velocity portion of the airflow to move towards the engine bypass. The plurality of vortex generators are disposed aft of a terminal shock and have a height greater than the thickness of the boundary layer.

Abstract: In a damping device according to the present invention, a damping device 63 is mounted on a bypass pipe 61 that supplies an amount of high-pressure air to a combustor transition piece 33. The damping device 63 includes a fluid introducing unit 71 that forms a fluid introduction space B by covering an outer peripheral portion of the bypass pipe 61, a plurality of acoustic boxes 73a and 73b that forms resonance spaces Da and Db with the base portions connected to the fluid introducing unit 71 and the end portions extending along the outer peripheral portion of the bypass pipe 61 in the circumferential direction, and partition plates 74a and 74b that form resonance ducts Ea and Eb of a predetermined length by partitioning the resonance spaces Da and Db.

Abstract: An apparatus for dampening of acoustic noise generated by air-cooling of at least one turbine component provided with the nacelle of a wind turbine is provided. Apparatus for dampening of acoustic noise generated by air-cooling of at least one wind turbine component provided with the nacelle of a wind turbine, comprising at least one acoustic dampener, with the acoustic dampener including at least one acoustic dampening channel structure having at least one acoustic dampening channel connected with at least one inlet opening of the acoustic dampening channel structure and at least one outlet opening of the acoustic dampening channel structure.

Abstract: An acoustic panel structure includes a perforated front sheet, a back sheet and a porous core between the front sheet and the back sheet. The acoustic panel structure also includes a non-perforated back cover which overlaps a portion of the back sheet. A portion of the back sheet is perforated and another portion of the back sheet is not perforated. An acoustic chamber is formed at least in part by the space between the back sheet and the back cover. The acoustic chamber may also be formed in part by the space between the front sheet and back cover. The space between the back sheet and the back cover is generally elongated in the x-y direction of the acoustic panel structure. This relatively thin acoustic panel structure is configured to attenuate long wavelength and low frequency noise.

Abstract: An engine nacelle includes an exterior wall and an interior wall. A forward bulkhead is coupled between the interior wall and the exterior wall, wherein the forward bulkhead comprises a curvilinear body portion that extends between the interior wall and the exterior wall. The engine nacelle also includes an aft bulkhead coupled between the interior wall and the exterior wall, wherein the aft bulkhead comprises a curvilinear body portion that extends between the interior wall and the exterior wall.

Abstract: A nacelle inlet is provided for a turbo-fan engine for an aircraft. The nacelle inlet includes a lip skin having first and second pad-ups, an inner barrel including a forward flange and a bulkhead. The bulkhead, the lip skin and the inner barrel are attached together at mating surfaces. The bulkhead and the lip skin are attached together at mating surfaces. Each mating surface defines a surface geometry which approximates a developable surface having parallel ruling lines. The developable surfaces are formed by machining.

Abstract: An aircraft includes a fuselage including a propulsion system supported within an aft portion. A thrust reverser is mounted proximate to the propulsion system for directing thrust in a direction to slow the aircraft. The thrust reverser directs thrust at an angle relative to a vertical plane to reduce interference on control surfaces and reduce generation of underbody lift.

Abstract: A method for mounting an acoustic panel on the inner surface of an annular housing of an axial turbomachine, the cylindrical acoustic panel comprising at least one longitudinal slot so as to define two opposing ends, the method comprising: bringing the two opposing ends closer together so as to reduce the cylinder radius of the acoustic panel; inserting the acoustic panel into the fan casing when the opposing ends are close together; deploying the opposing ends so as to increase the cylinder radius of the acoustic panel; and securing the acoustic panel to the housing by gluing.

Abstract: A method and apparatus for attenuating sound. Air, through which acoustic waves are traveling, is received within a core comprised of a plurality of cells. The sound created by the acoustic waves is attenuated using a set of channels through a number of cell interfaces between cells of the plurality of cells by allowing the air to flow between the cells of the plurality of cells through the set of channels.

Abstract: An acoustic protection device for a fan casing of an aircraft turbomachine, including an acoustic protection panel forming a ferrule sector centered on a center axis, and one or more vibration damping strips pressed on a first side to an external surface of the panel and on a second side to an internal surface of the casing, each damping strip including two opposite edges each including an upstream end and a downstream end spaced from each other along a direction of the center axis. At least one of both opposite edges of at least one of the damping strips is shaped such that liquid present on this edge can flow by gravity towards either or both of its upstream and downstream ends.

Abstract: A thermally-controlled component and thermal control process are disclosed. The thermally-controlled component includes thermally-responsive features. The thermally-responsive features are configured to modify a flow path to control temperature variation of the thermally-controlled component. The thermally-responsive features deploy from or retract toward a surface of the thermally-controlled component in response to a predetermined temperature change. The thermal control process includes modifying the flow path in the thermally-controlled component to control temperature variation of the thermally-controlled component and/or cooling a region of the thermally-controlled component through the thermally-responsive features deploying from or retracting toward a surface of the thermally-controlled component.

Abstract: An exhaust system for an APU in a gas turbine engine aircraft having a tail cone in which exhaust noise is reduced. An APU exhaust liner is attached to the APU and extending to an oval cutout at the tail cone The liner has a first constant cross section from the APU to a point proximate the tail cone, and a second, increasing cross section from the point proximate the tail cone to the liner cutout. The liner is turned toward the side of the tail. The liner cutout is further enlarged with at least one larger cutout.

Abstract: An interactive liner design and impedance modeling tool comprises software utilized to design acoustic liners for use in constrained spaces, both regularly and irregularly shaped. A graphical user interface allows the acoustic channel geometry to be drawn in a liner volume while the surface impedance calculations are updated and displayed in real-time. A one-dimensional transmission line model may be used as the basis for the impedance calculations.

Abstract: An acoustic panel liner includes a face sheet, back plate, and liner core positioned therebetween, which may be used in an engine nacelle. Elongated chambers contain variable amounts of septa at a calibrated depth or depths. The septa may have varying DC flow resistance. The chambers may have a hexagonal or other polygonal cross sections. The septa, such as mesh caps, may be bonded to an inner wall of a corresponding chamber. The insertion depths may be the same or different. If different, the pattern of distribution of the depths may be randomized.

Abstract: A turbine exhaust case for a gas turbine engine has an outer shroud and an inner shroud concentrically defining therebetween an annular gaspath for channelling hot gases. A plurality of circumferentially spaced-apart turbine exhaust struts extends radially across the hot gaspath. The inner shroud may be provided in the form of an acoustic panel. The acoustic panel has a radially outwardly facing surface defining the radially inner flow boundary surface of the gaspath. The acoustic panel is mechanically fastened to the radially inner end of the struts.

Abstract: A duct and a method for damping pressure waves caused by thermoacoustic instability, the duct comprising: a locking Helmholtz resonator positioned on the circumference of the duct, the locking Helmholtz resonator tuned to a first frequency; and one or more damping Helmholtz resonators positioned on the circumference of the duct at or adjacent to 90 and/or 270 degrees from the locking Helmholtz resonator, wherein the damping Helmholtz resonators are tuned to a second frequency which is different from the first frequency of the locking Helmholtz resonator.

Abstract: A fan ramp for use in a thrust reverser portion of a nacelle is disclosed. The nacelle is included in a propulsion system. The fan ramp extends circumferentially about an axial fan ramp centerline. The fan ramp includes a forward edge, an aft edge, and a noise suppression feature. The forward edge is disposed proximate an aft end of a fan case. The fan case at least partially surrounds a fan section of a gas turbine engine. The aft edge is disposed proximate a forward end of an array of cascades. The array of cascades is operable to permit a bypass airstream to pass there through during a thrust reversing operation. The noise suppression feature is operable to suppress noise within the propulsion system.

Abstract: The invention relates to a device for reducing the noise emitted by a conduit belonging to an aircraft system that comprises an acoustic grid positioned in a segment of the conduit. Said grid has a length L1, along the axis (X-X) of said segment, and includes multiple meshes, each of which fits into a square with side m, the following relationships being satisfied: L1>1.8?fmin, m<?fmax/2, where fmax is the maximum frequency of the noise to be attenuated, fmin is the minimum frequency of the nice to be attenuated, ?fmax is the wavelength of the noise at frequency fmax, and ?fmin is the wavelength of the noise at frequency fmin. Application to air intake and outlet conduits of aircraft systems.

Abstract: A mixer for mixing flows in a turbofan engine is provided. The mixer includes a plurality of chevron lobes, each of the plurality of lobes comprising a crown, a keel, a first trailing edge, a second trailing edge, and a first transverse edge extending between first trailing edge and second trailing edge, said mixer configured to receive two separate incoming exhaust flows and mix the two flows into at least one rotational exhaust flow that is ejected out at least one of said first trailing edge and said second trailing edge.

Abstract: In accordance with at least one embodiment: a muffler with a case (comprised of at least one inlet, at least one outlet, and a body), and elongated members comprised of material capable of a predetermined resonance. The elongated members have sufficient length after their final point of attachment to vibrate when exposed to flowing exhaust gasses. This vibration results in resonance that is noise canceling and/or sound enhancing. The muffler may also contain any combination of sound baffles, sound absorbent material, and other sound altering devices.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: Apparatus and methods for attenuating noise, vibration, and/or acoustic energy in a turbomachine. The apparatus includes a first base plate having first base plate perforations defined therethrough. The apparatus also includes first and second spacers each extending from the first base plate, being spaced apart from each other, and defining a first channel therebetween. The apparatus further includes a first insert disposed in the first channel and having insert perforations defined therethrough, the first base plate, the first and second spacers, and the first insert at least partially defining a first acoustic chamber therebetween, wherein the base plate perforations and the insert perforations are in communication with the first acoustic chamber.

Abstract: A gas turbine engine (10) comprising a fan (14) and an acoustic absorption intake and/or fan case liner (40). The liner (40) comprises at least four liner sections (42) demarcated by circumferentially spaced inhomogeneities (50). The inhomogeneities (50) cause scattering of a fan tone acoustic field when the gas turbine engine (10) is in use. To be accompanied when published by FIG. 3.

Abstract: A duct and a method for damping pressure waves caused b thermoacoustic instability, the duct comprising: a locking Helmholtz resonator positioned on the circumference of the duct, the locking Helmholtz resonator tuned to a first frequency, and one or more damping Helmholtz resonators positioned on the circumference of the duct at or adjacent to 90 and/or 270 degrees from the locking Helmholtz resonator, wherein the damping Helmholtz resonators are tuned to a second frequency which is different from the first frequency of the locking Helmholtz resonator.

Abstract: Honeycomb sections are bonded together with seams made up of an adhesive that is carried by a linked-segment seam support. The seams are particularly useful for splicing together curved honeycomb sections that contain acoustic septum. The curved acoustic honeycomb sections are spliced or seamed together to form engine nacelles and other acoustic damping structures.

Abstract: A gas turbine including a combustion oscillation suppressing device disposed to a transition piece of a combustor to define a gas space S, and having a plurality of vent holes through which a gas space L and the inside of the transition piece communicate. The combustion oscillation suppressing device is formed of a first member and a second member which define the gas space, the second member having the plurality of vent holes, and a portion of a transition piece wall includes the second member. A distance from the second member to the first member in the radial direction of the transition piece is a height of the gas space, and when the height of the gas space varies along an axial direction of the transition piece, an opening ratio ?p of each of the plurality of vent holes is adjusted depending on the height of the gas space.

Abstract: A device controlling vortex structures in a turbulent air jet flowing out of an air exhaust channel, which is constituted by at least one pair of needles connected to an AC voltage source to create a pulsating corona discharge between the needles. One needle is located along and the other across the flow to produce a resonant effect of the pulsating corona discharge on the vortex structures. In a method for controlling vortex structures in a turbulent air jet, at least one pair of needles connected to an AC voltage source create a pulsating corona discharge between the needles. One needle is located along and the other across the flow to produce a resonant effect of the pulsating corona discharge on the vortex structures.

Abstract: An inlet for an aircraft engine nacelle having an acoustic barrel panel, a septum buried or sandwiched within the acoustic barrel panel, and a perforated face sheet. The inlet may be located forward of an engine and/or an engine fan housed within the aircraft engine nacelle. Both the septum and the perforated face sheet may have perforations or holes therein of varying sizes and varying distances apart from each other. The smaller the area of the perforations or holes and the further apart they are from each other, the higher the impedance of a given area of the septum or the perforated face sheet. The impedance of the septum and/or the perforated face sheet may progressively increase in a direction away from the engine fan.

Abstract: A core nacelle for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a core cowl positioned adjacent to an inner duct boundary of a fan bypass passage having an associated cross-sectional area that radially extends between a fan exhaust nozzle and the inner duct boundary. The core cowl includes at least one groove that is selectively exposed to change the cross-sectional area at an axial location of the fan exhaust nozzle.

Abstract: The present invention relates to a method for manufacturing a body provided with a honeycomb structure, which body is designed preferably axis-symmetrical, with sheet-metal parts being provided with a wave-like structuring by means of a forming process and formed to ring elements, with the sheet-metal parts being shaped and designed in a substantially identical way as regards their structuring, with adjacent sheet-metal parts being arranged offset to one another and then joined together to form the honeycomb structure.

Abstract: A turbine exhaust case for a turbofan engine comprises a plurality of arcuate acoustic panels assembled into a circumferentially extending inner shroud with circumferential gaps between adjacent acoustic panels. An outer shroud extends circumferentially about the inner shroud. The inner shroud and the outer shroud define an annular gaspath therebetween. A plurality of circumferentially spaced-apart exhaust struts extends radially across the annular gaspath and structurally connects the individual acoustic panels forming the inner shroud to the outer shroud.

Abstract: A noise reduction system with a chamber includes a chamber (17) which is provided at a portion of a supply path connecting a flow path in an upstream side of a combustor in a jet engine to a plurality of microjet nozzles which is provided at an exhaust side peripheral edge of a main nozzle of the jet engine, wherein the supply path is configured to supply part of compressed air from the flow path into the chamber (17), and the chamber (17) is configured to inject the compressed air through the plurality of the microjet nozzles (63) to a jet flow.

Abstract: A protocol for assembling a fuel shut off pin valve assembly for a gas turbine engine. The protocol outlines a specific sequence of events in order to ensure failsafe incorporation of a fuel shut off valve assembly within the low pressure turbine area and specifically within the engine casing of the gas turbine.

Abstract: Method and apparatus for converting acoustic energy in an unshifted acoustic signal into an upshifted acoustic signal, and emitting the upshifted acoustic signal. Using a frequency converter, acoustic energy in a first frequency range in the unshifted acoustic signal is converted into acoustic energy in the upshifted acoustic signal in a second frequency range. The upshifted acoustic signal is emitted into an ambient acoustic environment. The second frequency range is higher than the first; the unshifted acoustic signal comprises an unwanted acoustic noise signal. The effect of the acoustic noise signal can be reduced by: enhanced attenuation of upshifted signal during atmospheric propagation; enhanced attenuation of upshifted signal by a non-gaseous attenuating medium; or decreased coupling of upshifted signal into resonances of a human body or mechanical structure.

Abstract: A method for manufacturing a sound absorber, the outer wall of which is provided with a plurality of recesses, with funnel-like cone elements each being assigned to the recesses inside the sound absorber, said cone elements having a larger opening facing radially outwards and a smaller opening facing radially inwards, with adjacent cone elements each being provided on a strip-shaped first carrier band, with cup elements being provided radially on the inside relative to the cone elements, and each cup element receiving one cone element, with adjacent cup elements each being provided on a strip-shaped first carrier band, with the first carrier bands being arranged adjacently in a first direction, and the second carrier bands being arranged adjacently in a second direction, with the directions crossing each other and the carrier bands being joined to one another and to the outer wall to form a rigid body.

Abstract: The present invention relates to a gas-turbine exhaust cone having an outer wall, which is provided with a plurality of recesses, a honeycomb-structured layer, which is arranged on the inside of the outer wall and extends along the inside of the outer wall, an inner wall, which extends substantially parallel to the outer wall and is connected to the honeycomb structure, and at least one annular chamber, which adjoins the inner wall and is centered relative to a central axis, with the inner wall being provided with passage recesses connecting the area of the honeycomb structure to the annular chamber.

Abstract: A silencer for an auxiliary power unit of an aircraft includes an inlet, an outlet, a housing, and a flow channel of a porous material, arranged in the housing. Between the housing and the flow channel an intermediate space is created, which space is divided by at least one outer partition into outer cells arranged around the flow channel. A central body includes a casing that is permeable to gas at least in some regions and that is arranged in the interior of the silencer. The central body includes inner partitions dividing the central body into inner cells. The resonance frequency of the inner cells corresponds to the natural frequency of the first radial mode in a channel formed between the central body and the inner surface of the inlet. In this manner particularly good sound attenuation is achieved with a compact installation space of a silencer.

Abstract: A silencer for an auxiliary power unit of an aircraft comprises an inlet, an outlet, a housing and a flow channel with a porous wall material that is arranged in the housing. An intermediate space is formed between the housing and the flow channel and divided into outer cells that are arranged around the flow channel by means of one or more partitions. This makes it possible to realize a very compact silencer with very good sound insulation properties. The sound dampening can be additionally improved by dividing the outer cells into outer regions and inner regions, by adapting the shape of the silencer to a tail section of an aircraft and by arranging an annular channel on the outlet.

Abstract: A disclosed embodiment provides an acoustic panel having a composite laminate panel having a back sheet layer, a face sheet layer, and a core layer disposed therebetween. The core layer has one or more depressions at an interface between the core layer and the face sheet layer, with the face sheet having a generally uniform thickness across the composite laminate panel. A hole extends through the composite laminate panel at the depressions, and a bolt assembly extends through the hole such that it is countersunk.

Abstract: A turbojet engine nacelle includes at least one nacelle element having a wall delimiting a main flow channel. The flow channel has a variable flow passage section. The cross-sectional area of at least one first area of the channel is greater than that of at least one second area of the channel. The turbojet engine nacelle further includes at least one secondary flow conduit establishing a fluidic connection between the first area and the second area of the channel and is arranged parallel to the channel. The wall of the nacelle element is at least partly covered by an acoustic damping panel extending in said areas of the channel. The conduit passes through the acoustic damping panel in order to establish the secondary flow via the acoustic damping panel. The present invention can be used in the avionics field.

Abstract: An acoustic damper is fixed to a vibration source to be provided along an outer surface of the vibration source, forms a passage that takes in air vibration generated by the vibration source, and is provided with an acoustic portion having a partition member closing the downstream side of the passage with respect to the propagation of the air vibration to serve as resistance against the air vibration, in which the acoustic portion allows air vibration to propagate in a reverse direction, and has partition members arranged so as to serve as resistance against the air vibration in respective directions. This allows to prevent from generating stress due to pressure fluctuation caused when reducing vibration, and to achieve a relatively compact size.

Abstract: An airborne mobile platform that has at least one turbofan engine assembly having a fan driven by a core engine, a short nacelle around the fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is disposed substantially coaxial with and extending forwardly into the fan exhaust duct to provide a mixer duct between the mixer duct shell and a core engine shroud of the core engine. At least a portion of the mixer duct shell has a honeycomb core structure having an inner surface and an outer surface, with an acoustic lining on one of the inner or outer surfaces. The acoustic lining attenuates sound emanating from the turbofan engine assembly.

Abstract: A combustor liner has a plurality of resonators formed thereon. Each resonator has a radially outer wall and at least one side wall. The outer wall can be free of holes. Each resonator has an inner cavity defined between the outer wall, the at least one side wall and the outer peripheral surface of the liner. The at least one side wall of each resonator surrounds a subset of a plurality of holes that extend substantially radially through the liner. A plurality of cooling passages extends generally longitudinally within the liner. Each cooling passage has an inlet in fluid communication with the exterior of the liner and an outlet in fluid communication with the inner cavity of a respective one of the resonators. A coolant, such as compressor air, can enter and flow along the cooling passages to thereby cool the liner and purge the inner cavity of the resonator.

Abstract: A nozzle has a nozzle surface area and a nozzle rim, on which first and second guiding devices are alternatingly provided in the circumferential direction, where the first guiding devices are of the nozzle-type design and the second guiding devices are of the diffuser-type design. The first guiding devices each have a first azimuthal guide wall and two wall elements. The second guiding devices each have a second azimuthal guide wall and two wall elements. A wall element connects a first guiding device and a second guiding device. At least some of the first azimuthal guide walls of the first guiding device and at least some of the second azimuthal guide walls of the second guiding device have differing axial lengths, so that first and second trailing edges thereof have differing axial positions.

Abstract: A silencer (1) for an exhaust gas installation of an internal combustion engine, particularly of a motor vehicle includes a housing (2), with a circumferential shell (3) closed in the peripheral direction (5) and an end base (6) at each of two longitudinal ends, spaced apart in an axial direction (4). A silencer insert (15) is arranged in the housing (2) and has an inlet pipe (8) for exhaust gas and an outlet pipe (10) for exhaust gas. The shell (3) has, in the peripheral direction (5) has at least one inlet shell-segment (11) and one outlet shell-segment (12). The inlet shell-segment has at least one inlet opening (7), into which such an inlet pipe (8) is inserted from the inside. The outlet shell-segment (12) has at least one outlet opening (9) into which such an outlet pipe (10) is inserted from the inside.