Abstract: An aircraft propulsion unit is described. The unit may include a gas generator with a fan surrounded by a casing. A nacelle may extend around the casing and define an annular compartment with the casing wherein some equipment may be housed. An air inlet may be configured so a ventilation air flow penetrates inside the compartment. An air outlet may be configured so a ventilation air flow is evacuated from the compartment. The propulsion unit may also include an air flow adjustment regulator. The air flow adjustment regulator may be configured to maintain a nominal value of the ventilation air flow circulating through at least one of the air inlet and of the air outlet under nominal operating conditions, and to reduce the value of this ventilation air flow when a fire is detected inside the compartment.

Abstract: A two-stage water-lift and water separation muffler for use with marine electrical generators installed onboard marine vessels. A first stage water-lift muffler has internal tunable sound suppression structure to provide a first stage of exhaust silencing, and a second stage water separator having internal tunable sound suppression structure to provide a second stage of exhaust silencing while separating entrained cooling water from exhaust gases. The combined first stage water-lift muffler and second stage water separator include internal baffling and silencing structure which is easily adaptable to different generator configurations thereby allowing for structural adjustments to the muffler to optimize sound suppression and minimize backpressure for particular generator and/or exhaust conditions.

Abstract: A nacelle assembly includes a fixed inlet segment and a translating inlet segment. The translating inlet segment includes a slider beam laterally between a pair of tracks of the fixed inlet segment. The slider beam is mated with and slidable longitudinally along the pair of tracks. The translating inlet segment is configured to translate longitudinally between a retracted position and an extended position. An aft end of the translating inlet segment is abutted against a forward end of the fixed inlet segment when the translating inlet segment is in the retracted position. An airflow inlet into an inlet passage of the nacelle assembly is opened longitudinally between the aft end of the translating inlet segment and the forward end of the fixed inlet segment when the translating inlet segment is in the extended position.

Abstract: A gas turbine engine includes an engine core having a compressor, a combustor fluidly connected to the compressor, and a turbine fluidly connected to the combustor. A core nacelle is disposed radially outward of the engine core. A cavity is disposed between an inner surface of the core nacelle and an outer surface of the engine core. The cavity includes a vent disposed at an aft end. The vent includes at least one flap configured to be maintained in an unrestricted position and in a restricted position. An actuator is configured to control the position of the at least one flap.

Abstract: An assembly for a turbine engine comprises a circumferential band defining an air flow path, a component having a leading edge and extending from the band into the flow path, and a trough located in a surface of the band along the leading edge.

Abstract: An assembly is provided for a turbofan engine. This turbofan engine assembly includes a cowling, a door and an actuation mechanism configured to actuate movement of the door in response to receiving a control signal. The cowling is configured to form a compartment at least partially around a case of the turbofan engine. The cowling includes an exhaust port that is fluidly coupled with the compartment. The door is configured to at least partially open and close the exhaust port. This variable exhaust port may be opened in case increased airflow is needed through the compartment, such as when increased cooling airflow is needed through an environmental air precooler that cools compressed air for the aircraft cabin and the precooler exhausts its cooling air into the compartment.

Abstract: An exhaust mixer for a gas turbine engine has a lobe cross-over offset. The proposed geometric feature leads to improved exhaust performance and potential weight reduction.

Abstract: A long-duct mixed-flow nozzle system for a turbofan engine, the nozzle system may include an inner housing configured to enclose a core and form a core flow duct, the inner housing terminating in a core nozzle having a core exit aperture, an outer housing forming a fan flow duct and terminating in a fan exit aperture at a location downstream of the core exit aperture, the fan exit aperture having a plurality of chevrons, and the core exit aperture having a plurality of chutes separated by radially extending lobes configured to mix exhaust gas from the core flow duct with bypass gas flow in the fan flow duct, the radially extending lobes varying in profile from each other.

Abstract: Systems and methods for an exhaust system for a gas turbine engine are provided. The exhaust system includes an eductor system adapted to receive a primary exhaust fluid. The eductor system includes a body that extends along a first axis, with a plurality of ducts spaced apart about a circumference of the body. Each of the plurality of ducts extend from the body along a second axis transverse to the first axis to define a plurality of eductor primary flow paths that terminate in a mixing chamber. The mixing chamber is adapted to receive a secondary cooling fluid and to mix the primary exhaust fluid with the secondary cooling fluid to create a mixed fluid flow.

Abstract: A turbine exhaust system including a bypass fluid duct that includes: an outer exit shroud and an inner mold line (IML) forming a channel from an inlet to an outlet of the bypass fluid duct, a moveable diffuser configured to move between a forward and aft locations, and a door configured to pivot between minimum and maximum door angles, wherein the moveable diffuser and the door together form the outer exit shroud. The bypass fluid duct is configured to: receive an incoming fluid stream in a first direction through the inlet; bypass the incoming fluid stream around a turbine generator; and direct a flow path of the incoming fluid stream from the inlet through the outlet to join an ambient fluid flow. A movement of the door and the diffuser, from an unextended position to a fully or partially extended position varies a shape and volume of the channel.

Abstract: A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Abstract: A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to merge the third stream with the primary stream. The fixed geometry nozzle is configured to operate between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode the total pressure of the primary stream is substantially the same as the total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

Abstract: A variable area fan nozzle comprises an actuator flap and a follower flap. The actuator flap has a portion in contact with a portion of the follower flap. A bias member biases the follower flap outwardly. An actuator actuates the actuator flap inwardly and outwardly to, in turn, move the follower flap against the bias member and to vary an area of an exhaust nozzle. The flap actuator is operable to drive the actuator flap out of contact with the follower flap into a thrust reverser position.

Abstract: A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Abstract: One embodiment of the present invention is a unique gas turbine engine system. Another embodiment is a unique exhaust nozzle system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine systems and exhaust nozzle systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: Nacelle air management systems for a high bypass ratio engine are provided. The nacelles air management systems may include an outer cowl and an inner cowl that are configured to provide dual bypass flow channels around an engine core. These systems may be employed to accommodate larger engine fans. The nacelle air management system may also include one or more blocker doors are configured to at least partial obstruct and/or direct air from the dual bypass flow channels to create reverse thrust.

Abstract: A midframe portion (213) of a gas turbine engine (210) is presented, and includes a compressor section (212) configured to discharge an air flow (211) directed in a radial direction from an outlet of the compressor section (212). Additionally, the midframe portion (213) includes a manifold (214) to directly couple the air flow (211) from the compressor section (212) outlet to an inlet of a respective combustor head (218) of the midframe portion (213).

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

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

Abstract: A turbofan engine includes an engine core and a ducted fan assembly, with the ducted fan assembly including an annular cowling and a dilating fan duct nozzle. The nozzle includes continuous nozzle sections with intermeshing tiles. The tiles include drive tiles that are pivotal between nominal and dilated positions and driven tiles that intermesh with the drive tiles and shift with the drive tiles between the positions. The intermeshing tiles cooperatively adjust an orifice size of the nozzle by shifting between the positions and thereby affect the thrust and noise developed by the ducted fan assembly.

Abstract: An example auxiliary power unit (APU) exhaust silencer includes cooling features to protect the outer skin and other components from heat generated by gases passing through an exhaust duct. Cooling air flow through a cooling air passage in thermal contact with the exhaust silencer carries heat away from other nearby components and the aircraft skin.

Abstract: A trans cowl for a jet engine includes a chevron coupled with a linear actuators. The chevron is movable by the linear actuator forward or aft to change a gas flow path formed by an core cowl and thrust reverser translating cowl. In a first position, the chevrons are disposed substantially parallel the gas flow path to attenuate drag and/or loss of engine thrust. In a second position, the chevrons are moved aft to project, or further project, into the gas flow path. In one embodiment, the linear actuator comprises a first component that is coupled with the outer cowl. A second component of the linear actuator is coupled with the chevron. When installed, the linear actuator can be coupled with a controller and an electrical power source. A position sensor coupled with the controller senses a position of the linear actuator and/or the chevron.

Abstract: A nacelle assembly for a bypass gas turbine engine includes a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section. The variable area fan nozzle is in communication with a fan bypass flow path, the first fan nacelle section defines an intermittent trailing edge which defines a multiple of ports and the second fan nacelle section defines a multiple of doors, each of the multiple of doors match each of the multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position relative to the first fan nacelle section.

Abstract: A jet engine adapted for reducing far-field noise levels is disclosed. The jet engine has a plurality of flap assemblies that are disposed around a perimeter of an exhaust port of the jet engine. Each flap assembly is movable between a first position and a second position. The jet engine also includes a plurality of actuators respectively coupled to the plurality of flap assemblies. Each actuator is configured to selectively move the coupled flap assembly between the first and second positions. The jet engine also includes a controller that is coupled to each actuator. The controller is configured to cause at least one actuator to move the coupled flap assembly between the first and second positions at a determined frequency.

Abstract: The invention relates to a turbojet nacelle (1) and to an aeroacoustic noise reduction device for said nacelle, the nacelle comprising a nozzle (10) on the downstream end thereof, said nozzle (10) comprising an inner wall inside which a first flow from the turbojet circulates and an outer wall (102) outside which a second flow corresponding to the surrounding outside air circulates. The invention also relates to an aeroacoustic noise reduction device (20) of the turbojet, comprising a plurality of chevrons (201) arranged on the circumference of the nozzle (10). The invention is characterized in that the aeroacoustic noise reduction device also comprises a slide (202) arranged on the circumference of the nozzle in such a way that it can rotate about the axis of the nozzle, each chevron (201) being connected to the slide (202) by means of a guiding element (204) that can move along the slide (202) during the rotation thereof in order to ensure the displacement of the chevron (201).

Abstract: This trailing edge is of the type with moving chevrons (9a, 9b, 9c). It comprises actuating means able to cause these chevrons (9a, 9b, 9c) to move from a passive position in which they are directed substantially in the direction of the airflow leaving the said engine, into an active position in which at least part of each chevron is inclined with respect to this direction. This trailing edge is notable in that the said active position is derived from the said passive position by pivoting, possibly combined with a translational movement, of at least part of each chevron about an axis (Aa, Ab, Ac) substantially parallel to the said direction.

Abstract: A jet engine adapted for reducing far-field noise levels is disclosed. The jet engine has a plurality of flap assemblies that are disposed around a perimeter of an exhaust port of the jet engine. Each flap assembly is movable between a first position and a second position. The jet engine also includes a plurality of actuators respectively coupled to the plurality of flap assemblies. Each actuator is configured to selectively move the coupled flap assembly between the first and second positions. The jet engine also includes a controller that is coupled to each actuator. The controller is configured to cause at least one actuator to move the coupled flap assembly between the first and second positions at a determined frequency.

Abstract: According to the invention, the shape of each of said chevrons is defined: by two lateral sides (21, 22) having front ends (21A, 22A) that are connected to said nozzle (4, 10) and that, when moving away from the latter, converge towards each other without joining so that the rear ends (21R, 22R) of said lateral walls (21, 22) are spaced from each other; and by a curvilinear transverse line (23.1) connecting the rear ends (21R, 22R) of said lateral walls (21, 22) by forming two rounded side protrusions (24.1, 25.1) separated by an intermediate rounded recess (26.1).

Abstract: The invention relates to a turbojet nacelle (1) and to an aeroacoustic noise reduction device for said nacelle, the nacelle comprising a nozzle (10) on the downstream end thereof, said nozzle (10) comprising an inner wall inside which a first flow from the turbojet circulates and an outer wall (102) outside which a second flow corresponding to the surrounding outside air circulates. The invention also relates to an aeroacoustic noise reduction device (20) of the turbojet, comprising a plurality of chevrons (201) arranged on the circumference of the nozzle (10). The invention is characterised in that the aeroacoustic noise reduction device also comprises a slide (202) arranged on the circumference of the nozzle in such a way that it can rotate about the axis of the nozzle, each chevron (201) being connected to the slide (202) by means of a guiding element (204) that can move along the slide (202) during the rotation thereof in order to ensure the displacement of the chevron (201).

Abstract: An arcuate or semi-circumferential shield in the shape of a visor that is adapted to be secured to an external surface of a nozzle housing of a jet nozzle to attenuate jet noise generated by the jet nozzle. The shield may be fixedly secured to the nozzle housing or movably supported so that it can be moved between retracted and deployed positions. When used as a fixedly mounted component, or when positioned in its deployed position, a downstream edge of the shield extends past a downstream edge of the nozzle housing and the shield is spaced apart from an outer surface of the nozzle housing to form a channel therebetween. The shield is preferably orientated at approximately a bottom dead center of the nozzle housing. The shield operates to attenuate jet installation noise, and particularly jet installation noise experienced during the takeoff phase of flight of a jet aircraft, as well as to reduce jet installation noise that would otherwise propagate forwardly toward the cockpit of the aircraft.

Abstract: This trailing edge is of the type with moving chevrons (9a, 9b, 9c). It comprises actuating means able to cause these chevrons (9a, 9b, 9c) to move from a passive position in which they are directed substantially in the direction of the airflow leaving the said engine, into an active position in which at least part of each chevron is inclined with respect to this direction. This trailing edge is notable in that the said active position is derived from the said passive position by pivoting, possibly combined with a translational movement, of at least part of each chevron about an axis (Aa, Ab, Ac) substantially parallel to the said direction.

Abstract: In order to provide noise suppression, bumps or undulations are provided on a nozzle surface in order to vary the convergent-divergent ratio between that surface and an opposed nozzle surface. By such an approach, a circumferential variation in the shock cell pattern is created and the flow is deflected so as to enhance turbulent mixing thereby suppressing noise.

Abstract: An example turbofan engine sound control system includes a core nacelle (12) housing a compressor and a turbine. The core nacelle is disposed within a fan nacelle (34). The fan nacelle includes a turbofan. A bypass flow path downstream from the turbofan is arranged between the two nacelles. A controller (50) is programmed to manipulate the nozzle exit area to control sound propagating from the engine. In one example, the controller manipulates the nozzle exit area using hinged flaps (42) to control engine sound. The hinged flaps open and close to adjust the nozzle exit area and the associated bypass flow rate.

Abstract: The invention uses tubes for the reduction of radiated duct noise. The tubes may be used to attenuate narrow band noise by diverting a portion of the acoustic energy through the tubes back into the duct. Upon reintroducing the diverted flow out of phase with the primary flow, acoustic cancellation is achieved. The frequency at which this occurs is dependent on the difference in the path lengths through the tubes and the separation distance—the distance separating the inlet and outlet of the tube. This invention can be used over varying frequency ranges by using an arrangement of flexible, constant length tubes. The effective frequency range is tailored by varying the separation distance and also allows the varying the angular arrangement of the tubes for more effective attenuation of spinning modes having varying propagation angle.

Abstract: The invention uses tubes for the reduction of radiated duct noise. The tubes may be used to attenuate narrow band noise by diverting a portion of the acoustic energy through the tubes back into the duct. Upon reintroducing the diverted flow out of phase with the primary flow, acoustic cancellation is achieved. The frequency at which this occurs is dependent on the difference in the path lengths through the tubes and the separation distance—the distance separating the inlet and outlet of the tube. This invention can be used over varying frequency ranges by using an arrangement of flexible, constant length tubes. The effective frequency range is tailored by varying the separation distance and also allows the varying the angular arrangement of the tubes for more effective attenuation of spinning modes having varying propagation angle.

Abstract: An embodiment of the technology described herein is an auxiliary power unit assembly. The auxiliary power unit assembly includes an auxiliary power unit being installable in an aircraft having a cabin, a duct connecting the cabin and the auxiliary power unit, and a noise reduction feature in the duct.

Abstract: An arcuate or semi-circumferential shield in the shape of a visor that is adapted to be secured to an external surface of a nozzle housing of a jet nozzle to attenuate jet noise generated by the jet nozzle. The shield may be fixedly secured to the nozzle housing or movably supported so that it can be moved between retracted and deployed positions. When used as a fixedly mounted component, or when positioned in its deployed position, a downstream edge of the shield extends past a downstream edge of the nozzle housing and the shield is spaced apart from an outer surface of the nozzle housing to form a channel therebetween. The shield is preferably orientated at approximately a bottom dead center of the nozzle housing. The shield operates to attenuate jet installation noise, and particularly jet installation noise experienced during the takeoff phase of flight of a jet aircraft, as well as to reduce jet installation noise that would otherwise propagate forwardly toward the cockpit of the aircraft.

Abstract: A device for reducing engine exhaust noise in an aircraft engine includes a number of oscillating jets which direct an oscillating flow into the engine exhaust. The oscillating jets are connected via channels to a turbomachinery source of high pressure gas, where the source is located upstream from the engine exhaust portion. The high pressure gas passes through the oscillating jets, which have a nozzle with a triangular shaped orifice and an exhaust pipe, and exits the oscillating jets having a oscillating flow. The oscillating flow mixes with the engine exhaust thus reducing the noise created by the engine exhaust.

Abstract: Exhaust system assemblies employ wire bushings for thermal compensation to minimize noise due to “stick-slip” while the system cools. In a preferred embodiment, the wire bushings are formed of at least one strand of wire coiled about male end portions of pipe sections used in the system, each of which male end portions are inserted into a female end portion and held therein by an interference fit. The strand of wire is spot welded at spaced locations to the male end portions to provide bearing surfaces between the spot welds that move slightly as the exhaust system heats and cools to relieve thermal stress in small, substantially inaudible increments.

Abstract: An exhaust processor has an exhaust tube and a cover mounted for rotation at least partially around the exhaust tube to adjust the tuning frequency of the exhaust processor. A method of operating the exhaust processor is also disclosed.

Abstract: In order to reduce noise effectively and to provide a light-weight noise reducing apparatus and an exhaust nozzle for a jet engine, a mixer for mixing an exhaust gas and an external air is disposed behind an engine, a flow channel for accelerating an exhaust gas toward a mixer and a flow channel for introducing an external air to a mixer are formed by a flow channel forming section in a noise reducing mode.

Abstract: The invention is an adjustable exhaust system and an adjustable insert for an exhaust system. The exhaust system includes a header flange, at least one conduit coupled to the header flange, a collector coupled to the conduit, and an adjustable insert for tuning engine performance.

Abstract: A nozzle where the sector into which noise reduction is sought (the direction of populated areas) has a moving lip, driven by actuators in an oscillatory, periodic or sinusoidal motion, around the mid position, to and from the axis, so as to cause the shear layer to become more irregular and turbulent, thus enhancing scattering effects on the sound of internal sources, including back reflection into the jet, widening the directivity pattern and spreading the acoustic energy into a broader spectrum, so as to reduce the acoustic energy transmitted in these sensitive directions; the nozzle lip may also be corrugated, undulated or have vortex generators, in the sector over which noise reduction is sought. In all other sectors, for which noise radiation is of no concern, even if enhanced, the nozzle lip is smooth and fixed, leading to a thin shear layer, across which sound is more easily, so that acoustic energy exits from the jet in these harmless directions.

Abstract: Device for connecting two tubular pieces of an aircraft turbine engine.

Abstract: A gas turbine engine includes an exhaust flow guide attached to the engine exhaust end. The exhaust flow guide is made of a piece of sheet metal curved about the central longitudinal axis of the engine to form a partial portion of a nozzle having curved leading and trailing edges, and side edges. The exhaust flow guide asymmetrically affects the downstream jet noise contribution volume of the exhaust gases discharged from the engine exhaust end to change the jet noise directivity and reduce its power level. The exhaust flow guide can further include additional features such as perforations, corrugated surfaces, reinforcing strips and irregularly shaped trailing edges to further improve the mixing of exhaust gas flow with surrounding fluid flow to further reduce the downstream jet noise contribution volume, thereby resulting in jet noise reduction. The exhaust flow guide according to the present invention also improves attenuation of engine noise generated in the engine and emitted from the engine exhaust end.

Abstract: A apparatus and method for augmenting the thrust of a jet engine while reducing the noise generated thereby, including extending the thrust reverser sleeve, exposing the cascades while retaining the thrust reverser doors or buckets in their stowed position.

Abstract: An ejector nozzle (10) including an first cowling (12), a second cowling (14), and opposed upright sidewalls (16) that together form an internal nozzle exhaust path is provided. A reconfigurable plug assembly (18) having separable first and second diverters (20), (22) is located in the exhaust path to direct engine exhaust (24) in either dual paths formed around diverter outer surfaces (76), or dual paths formed between the diverter inner surfaces (74) and centerbody exterior surfaces. When the diverters direct exhaust airflow between themselves, first and second ejectors (26), (28) formed in the first and second cowlings (12), (14), respectively, are available to entrain ambient air (30) into the exhaust stream (24). The preferred ejectors include translatable aft flaps (32), (34). Mixing devices, such as a lobed mixer (90) are incorporated into the diverters (20), (22) to improve engine noise suppression.

Abstract: A method of reducing noise from a jet engine at a distance D from the engine comprising the step of up-shifting frequencies of noise from the jet engine to higher frequencies to cause increased noise fall-off in the distance D. Where the jet engine has an exhaust nozzle for exhaust gases carrying the noise of minimum cross-sectional area A, the method comprises, disposing a frequency shifting plate containing a plurality of small openings therethrough of total cross-sectional area A across the exhaust nozzle and passing all exhaust gases from the jet engine through the frequency shifting plate. Each of the plurality of small openings has an individual cross-sectional area which up-shifts frequencies of noise from the jet engine passing therethrough to higher frequencies. Various embodiments of apparatus to perform the method are disclosed.

Abstract: A noise suppression system for a jet engine of an aircraft having a device for blocking a flow of exhaust gases down the exhaust nozzle of the engine, and a device for separating the exhaust gases into a plurality of exhaust streams spaced a predetermined distance apart and having a smaller cross-sectional area than that of the original flow of exhaust gases. The separating device also directs the exhaust streams into the ambient air from points outside the periphery of a nacelle covering the engine. The aforementioned predetermined distance separating the exhaust streams, and the cross-sectional area of the exhaust streams, are chosen so that engine noise is shifted to a higher frequency spectrum than that of the original flow of exhaust gases, and then substantially dissipated within a short distance of the aircraft.

Abstract: A sound suppressor exhaust structure for a jet engine. The exhaust structure has a nozzle body having a generally cylindrical forward intake opening to accommodate an aft end of a jet engine, and a generally quadrilateral aft exit opening. An openable sound suppression chamber is situated aft of the exit opening and is formed by two opposing enclosure structures each having a major surface and each hingedly attached in a clamshell manner to the nozzle body. The opposing enclosure structures are movable between an open and a closed configuration such that, when the enclosure structures are closed, a chamber having a generally quadrilateral cross section is defined and situated aft of the exit opening. Each enclosure structure has a major surface through which are substantially uniformly disposed a plurality of small apertures such that from about 15% to about 25% per square foot of each major surface is void and from about 45 to about 65 apertures are provided per square foot.