Abstract: A cowl for a turbomachine nozzle, the cowl having a plurality of repeated patterns disposed circumferentially at a trailing edge, each pattern being asymmetrical about a midplane of the pattern containing a longitudinal axis of the cowl, and each pattern having a first portion inclined radially towards the inside of the cowl and a second portion inclined radially towards the outside of the cowl.

Abstract: An axisymmetric nozzle for a gas turbine engine has divergent flaps and seals disposed about a central longitudinal axis thereof. The divergent flaps and seals each have an inner surface defining cooling air inlet holes at an upstream portion, cooling air exit holes at a downstream portion, and cooling air channels disposed within the divergent flap and seal, and communicating at a first end with the inlet holes and at a second end with the exit holes for conducting cooling air therethrough. Some of the exit holes of the divergent flap are disposed along lateral edge regions thereof. The divergent seals are interposed between adjacent divergent flaps, and each include lateral edge regions extending laterally beyond the exit holes. The lateral edge regions of each divergent seal have an outer surface overlying at least a portion of the exit holes of an adjacent divergent flap.

Abstract: A device for reducing the jet noise of a turbomachine, the turbomachine having a longitudinal axis and a substantially cylindrical nozzle extending along the longitudinal axis of the turbomachine, having a downstream end for mixing the flows of gas inside and outside the nozzle, the device comprising a plurality of corrugations disposed to extend the downstream end of the nozzle, and a plurality of notches disposed between pairs of successive corrugations in the plurality of corrugations, and the plurality of corrugations and the plurality of notches are symmetrical relative to at least one axis perpendicular to the longitudinal axis of the turbomachine so as to generate a dual counter-rotating effect between the gas flows.

Abstract: An acoustic liner 20 includes a remote panel 26, a proximate panel 28 transversely spaced from the remote panel and a resonator chamber 34b residing between the panels. Perforations 38 penetrate the proximate panel in registration with the resonator chamber 34b. A neck 56 with an inlet 58 recessed from the proximate panel establishes communication between the chamber and a fluid stream G flowing past the proximate panel. A bypass coolant passage 66 guides coolant through the perforations without guiding it through the resonator chamber.

Abstract: The invention relates to a ventilation system for a convergent divergent exhaust nozzle in a bypass turbojet comprising an afterburn chamber surrounded by an annular passage through which circulates a stream of cooling air, a convergent divergent axisymmetric nozzle arranged downstream of said afterburn chamber, each circle of flaps comprising alternately a plurality of controlled flaps, and a plurality of follower flaps, a circle of cold flaps arranged radially outside said nozzle and hinged at their upstream end to a conical shell linked to the downstream part of the casing.

Abstract: An embodiment of the invention is a technique to suppress noise in a jet engine. A noise suppressor includes an exhaust duct to exhaust a first stream from an air stream taken from an inlet. The exhaust duct has an exit end. The air stream is divided into the first stream and a second stream. The second stream has an axis. A stream director attached to vicinity of the exit end directs the first stream to a direction that is away from or sideward relative to the axis.

Abstract: A method for operating a gas turbine engine includes channeling compressed air from the gas turbine engine to a noise suppression system, and selectively operating the noise suppression system such that air discharged from the noise suppression system generates a streamwise vortex downstream from each respective chevron.

Abstract: An external jet nozzle mixer includes identically formed lobes. The external mixer works with the internal mixer further to mix the engine internal bypass flow with the internal jet engine core flow to level the disparate flow velocities, to reduce the peak velocities from the jet engine core and increase the lower bypass velocities of the engine internal bypass flow, and thereby reduce noise. The internal lobe contours act as lifting flutes, causing mixing of the primary hot and cold flows to mix before exiting the nozzle. The external lobe contours act as venturi chutes, accelerating the cooler ambient secondary air flow. The lobes thus act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. Also obtained is an increased thrust efficiency and a consequently decreased fuel consumption and engine emissions.

Abstract: A method enables a gas turbine engine to be operated. The method comprises channeling exhaust gases from a core engine through an exhaust assembly and past at least one flow boundary surface, and selectively operating a noise suppression system extending from the at least one flow boundary surface to facilitate attenuating noise generated during engine operation.

Abstract: A gas turbine engine including a gas exhaust duct assembly which comprises an annular shroud forming at least a section of an inner wall of a gas exhaust duct, a plurality of strut members radially projecting from the shroud, and a plurality of insert members. The strut members are disposed in a circumferentially, substantially equally spaced-apart relationship, and extend in a substantially axial direction. The insert members each define a circumferential section of an outer wall of the gas exhaust duct and are positioned in a space between adjacent strut members, and are secured to same.

Abstract: There is disclosed a lobe mixer comprising: a cylindrical partition wall 22 which partitions a high-speed inner flow from a low-speed outer flow; a lobe portion 24 which is disposed on a downstream side of the partition wall and whose transverse sectional shape is a waveform shape and is enlarged on the downstream side; and a connecting reinforcement member 26 which is disposed between lobe side walls 24a extending in a radial direction of the lobe portion to connect the adjacent lobe side walls to each other. The connecting reinforcement member 26 is a thin member disposed along a flow line of a fluid which flows between the lobe side walls.

Abstract: The invention relates to a rocket engine nozzle comprising a system for controlling jet separation of the flow in the nozzle, wherein said control system exhibits a plurality of separation triggering elements (5, 10) arranged in such a way as to generate, from mutually spaced initiation points (9), distinct zones (6) of jet separation, so as to form a three-dimensional separation of the flow. The flow control system can exhibit at least two triggering elements (5, 10).

Abstract: The present invention reveals a method and apparatus for controlling the effective area and thrust vector angle of a fluid flow. In one embodiment, the fluid flow is controlled in an advanced, high aspect ratio, complex aperture geometry nozzle using asymmetric injection into the subsonic portion of the fluid flow. The present invention vectors the primary flow by partially blocking the flow with an opposed flow across the flow field. A fluidic flow field is defined in a flow container that directs a pressurized, primary fluidic flow from the container towards an exit of the container. A nozzle may cooperate with the exit of the flow container to control the fluidic flow as it exits the flow container. One or more injectors associated with the container are proximate to the effect throat of the primary flow while other are located downstream of to introduce an opposing fluidic flow that interacts with the primary fluidic flow.

Abstract: An external jet nozzle mixer (20) includes identically formed lobes (48), equal in number those of the internal mixer (42). External mixer (20) works with internal mixer (42), further to mix the engine internal bypass flow with the internal jet engine core flow, to level the disparate exhaust flow velocities, to reduce the peak velocities from the jet engine core and increase the lower bypass velocities of the engine internal bypass flow, and thereby to reduce noise. The internal lobe contours act as lifting flutes, causing mixing of the primary hot and cold flows before exiting the nozzle. The external lobe contours act as venturi chutes, accelerating the cooler ambient air secondary flow. The lobes thus act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. Also obtained is an increased thrust efficiency and a consequently decreased fuel consumption and engine emissions.

Abstract: The present invention relates to a variable cycle boost propulsor system for use on an aircraft. The variable cycle boost propulsor system includes an engine, a turbine, a fan connected to the turbine, and a valve system for delivering the fluid output from the engine to the turbine for driving the turbine and the fan and thereby generating additional thrust for the aircraft. The engine is also used to provide power to one or more systems onboard the aircraft.

Abstract: The invention relates to a dual flow turbojet comprising an engine of axis X housed completely inside a tubular pod whose inside wall co-operates with the case of said engine to define an annular passage for a secondary flow delivered by a fan, said pod having an air inlet upstream of the engine, thrust-reverser means in its middle section, and a common nozzle for ejecting both the primary flow and the secondary flow, the outlet of the nozzle being situated downstream from the engine, said thrust-reverser means being suitable for taking up an active position in which the secondary flow is diverted outwards and forwards from said pod, said nozzle having an enclosure defined by the inside and outside walls of said nozzle.

Abstract: A lobe mixer (22) used in jet engines, is provided with diameter expansion portions (22a) in which the diameters gradually increase in the downstream direction, in which adjacent diameter expansion portions have different diameters from each other, and which guide the core gasses that flow along the inner sides of the adjacent diameter expansion portions, peripherally outwards at different angles. Thus, the gasses are mixed without the thrust of the jet engine being reduced, and the noise of the jet is reduced.

Abstract: The invention concerns a cooling system for a post combustion jet nozzle, The jet nozzle includes a primary gas duct allowing a primary flow of gas, and a secondary air duct allowing a secondary flow of air. The secondary air duct surrounds the primary gas duct and is separated from it by a protective thermal shroud. The secondary air duct has a downstream end, and dampers surrounding an output section of the primary gas duct. The system includes a protective thermal shell in the secondary air duct, at the downstream end of said duct. The protective thermal shell bears an annular diaphragm extending out in front of the dampers and being provided with support sectors and inter-sector zones equipped with slots. The inter-sector zones define spaces between the diaphragm and the protective thermal shroud.

Abstract: A pneumatic inhaler that is able to deliver a controlled burst or dose of aerosol from a reservoir of liquid medication. The inhaler is suitable for the aerosolization of liquid medication that is in solution or suspension form. The inhaler is also ideal for the delivery of unique and specialty liquid medications in short aerosol bursts because no additional formulation development is needed and has the further advantage of being able to deliver multiple medications, as mixed by the patient, doctor, or pharmacist, with a single burst at a repeatable output. Because medication and propellant are not mixed until aerosolization occurs, the inhaler is appropriate for more pharmaceutical agents than the current inhalers available and at a substantial cost savings.

Abstract: A gas turbine engine of a fan bypass type includes an exhaust shroud having a perforated tubular wall extending between a forward end and an aft end. The exhaust shroud is adapted to be affixed to a gas turbine engine exhaust for discharging engine exhaust gases without substantial blockage thereto. The perforations formed in the shroud wall communicate the regions at both sides of the shroud wall to pass fluid flow across the perforated shroud wall so that the mixing of the engine exhaust gases with a surrounding fluid flow is enhanced. The mixing of the engine exhaust gases with the surrounding fluid flow is further enhanced by a trailing edge of the exhaust shroud. The trailing edge is deviated from a straight line in a circumferential direction of the tubular wall to effectively increase the peripheral length of the mixing boundary. Thus, the jet noise contribution volume of the engine exhaust gases is effectively reduced, thereby significantly reducing the aero-engine jet noise.

Abstract: An aerospike nozzle includes a nozzle body and a spike, the nozzle body and the spike defining a generally annular gas flow channel therebetween; the spike comprising an upper, generally conical section having a decreasing diameter in a downstream direction; a middle, generally cylindrical section having a generally constant diameter; and a lower, generally conical section having a decreasing diameter in a downstream direction; a shroud disposed radially outward of the spike and longitudinally downstream of the nozzle body, the shroud comprising a first generally cylindrical section and, downstream of the first section, a second generally concave section; a plurality of struts connected between the spike and the shroud; an ambient air inlet defined between the nozzle body and the shroud; a mixing and combustion chamber defined between the spike and the first section of the shroud; and an expansion chamber defined between the spike and the second section of the shroud.

Abstract: Disclosed is a means for reducing jet engine exhaust noise wherein mixing is enhanced between adjacent exhaust flows and between exhaust flow and free-stream flow. The device does so with a very small degradation in aircraft performance. The device is a segmented, triangular or trapezoidal shaped, curved extension to a nozzle's sleeve that results in a serrated trailing edge. The nozzle extensions enhance the natural free mixing of the jet's exhaust flows and therefore reduce the acoustic energy associated with the velocity differences between the streams in which they are imbedded. The novel structure forces adjacent flows to penetrate into one another to a greater depth than that achievable with free mixing and results in a more uniform flow in a shorter stream wise distance.

Abstract: The invention relates to lobe-type mixer tube for a bypass jet engine in which a bypass flow delivered by a fan is mixed with a hot gas stream issuing from the core engine, said lobe-type mixer tube having essentially circumferentially equally spaced gutter-shaped, radially moderately alternately inward and outward directed ducts to guide both the bypass flow and the hot gas stream, said ducts each having an end formed by an essentially plane extreme section and the respective extreme sections of some ducts when viewed in a longitudinal section extending in the direction of flow being inclined vis-a-vis the extreme sections of the other ducts. In this arrangement the radially outward directed ducts, when viewed circumferentially, alternately issue at points offset one from the other.

Abstract: Augmentation of the thrust achieved by a supersonic nozzle of continuous curvature is achieved by causing secondary combustion to occur in an annular region of the interior of the divergent section of the nozzle. The secondary combustion forms a secondary combustion gas that complements the primary combustion gas passing through the nozzle and maintains a wall pressure that is equal to or greater than ambient pressure at low altitudes, eliminating the negative component of the thrust in an overexpanded nozzle at takeoff.

Abstract: An engine nozzle arrangement for an aircraft comprises a nozzle (18) for hot exhaust gases, and a shroud (20) for the stream of hot exhaust gases emerging from the exhaust nozzle. The shroud includes an intake for a boundary layer of air for cooling the stream of exhaust gases.

Abstract: A gas turbine engine exhaust nozzle includes an exhaust duct for channeling a gas jet. A plurality of adjoining chevrons are disposed at an aft end of the duct to define an exhaust outlet. Each of the chevrons is triangular in configuration, with a base, apex, side trailing edges converging therebetween, and radially opposite first and second surfaces bounded thereby. The trailing edges of adjacent chevrons are spaced laterally apart to define respective diverging slots disposed in flow communication with the duct. The chevrons have a concave contour axially between the bases and apexes which promotes jet mixing through the slots.

Abstract: An ejector mountable on a planar wall of a housing for gathering and conveying particulate material to within such housing generally consisting of a flat base plate segment mountable on such planar wall member, a first set of flat plates joined together and with the base plate segment defining a converging fluid passageway, a second set of flat plate segments joined together and with the first set of plate segments and the base plate segment defining a restricted fluid passageway communicating with the converging fluid passageway and a third set of flat plate segments joined together and with the second set of plate segments and the base plate segment defining a diverging fluid passageway communicating with the restricted fluid passageway, one of the plate segments having a first inlet opening through which a stream of pressurized fluid may be injected into the converging and restricted fluid passageways, one of the plate segments having a second inlet opening through which fluid may be suctioned by a negativ

Abstract: A control system for selectively adjusting effective flow areas of an aircraft engine exhaust nozzle to change operational characteristics of the nozzle. The control system includes a chamber having a hollow interior, a plurality of outlet passages extending from the hollow interior of the chamber to sites within the exhaust nozzle, and an adjustable inlet extending from a pressurized air source to the hollow interior of the chamber for delivering a jet of pressurized air to the chamber. The inlet is adjustable to direct pressurized air to selected one or more outlet passages for the delivery of air via the one or more outlets to corresponding one or more sites within the exhaust nozzle, thereby to change the operational characteristics of the nozzle.

Abstract: Mixing of a subsonic exhaust plume from a nozzle is induced and enhanced, and shaping of the plume is controlled, by periodic pulsed jets of fluid injected into the plume from two or more locations spaced circumferentially about the plume. The jets are pulsed out of phase relative to one another. An active flow control apparatus includes at least first and second pulsed jet nozzles circumferentially spaced apart about the exhaust plume near an exit plane of the exhaust nozzle, and a pulse generator operable to supply periodic pulses of fluid to each of the pulsed jet nozzles such that the pulses to each nozzle are out of phase relative to the pulses to the other nozzle or nozzles.

Abstract: A stream of primary gas flowing through a bi-stable thrust vectoring nozzle becomes attached to a first or second surface extending downstream of the nozzle, each surface incorporating one or more control ports for controlling to which surface the stream is attached, wherein relative to the longitudinal axis of the nozzle, the angle of discharge from the first surface is substantially different from the angle of discharge from the second surface, preferably with the first surface substantially aligned with the longitudinal axis of the nozzle. In one embodiment, a plurality of nozzles are arranged with the respective first surfaces substantially aligned with the longitudinal axis of the nozzle combination and each of the respective second surfaces arranged to laterally deflect a respective portion of the stream of primary gas in a respective direction along each of two orthogonal lateral axes.

Abstract: An exhaust induced ejector nozzle (20) includes a convergent portion (30) and a divergent portion (34) disposed downstream from the convergent portion (30). The ejector nozzle (20) also includes an ejector (40) having an outlet (46) disposed within the divergent portion (34). The ejector (40) is operable to provide entrainment of a nacelle airflow (18) from a nacelle airflow area (44) into the divergent portion (34). The ejector nozzle (20) further includes a convergent ejector (60) extending between the convergent portion (30) and the nacelle airflow area (44). The convergent ejector (60) is operable to draw a portion of an exhaust airflow (42) into the nacelle airflow area (44) to increase a flow rate of the nacelle airflow (18).

Abstract: An infrared suppressor system for a gas turbine engine is provided for suppressing infrared radiation associated with the hot metal parts of the engine and with the hot exhaust gases exhausted from the engine. The system features a unique arrangement of multiple baffles that are connected together as a single baffle module. The baffle arrangement permits mixing of hot and cool gas flows while eliminating line-of-sight infrared radiation in an axially compact suppressor. The baffles are additionally linked together into a single insertable baffle module that permits removal of the baffle module when infrared suppression is unnecessary.

Abstract: An infrared radiation Coanda suppressor of an exhaust system encloses a Coanda surface for guiding exhaust gases into the atmosphere. The infrared radiation suppressor further uses cooling films for reducing the temperature of the Coanda surface and other exposed surfaces. A cool air entrainment flow of the exhaust gases prevents the buildup of recirculating hot gases in the exhaust chamber. As a result, IR radiation is suppressed from areas of the exhaust system which are exposed to IR detectors.

Abstract: A rear mixer ejector of a turbomachine for a supersonic aircraft has an ejector casing of rectangular cross-section fitted with movable flaps defining a variable geometry exhaust nozzle, and cold air mixing devices mounted on the casing side walls upstream of the nozzle. The mixing devices comprise a plurality of U-section chutes extending into the casing from apertures in the casing side walls, each chute being formed by a pair of lateral walls and a transverse wall disposed between them. The chutes are adjustable between a deployed position and an inoperative position, either by pivoting the chutes about axes at the downstream edges of the apertures or by pivoting only the transverse walls about the upstream edges of the apertures. In each case the lateral walls of the chutes remain in the gas flow path through the casing, parallel to the casing axis, in both the deployed and the inoperative positions, and the transverse walls are parallel to the casing axis in the inoperative position.

Abstract: An axially symmetric turbojet engine exhaust-nozzle includes a system for masking the exhaust jet, such system including inner hot flaps (7), hinging on a conical shell (4) and driven by a first control ring (10) powered by linear actuators (11), and outer cold flaps (20) hinging on an annular assembly (12) and driven by a second control ring (22) powered by linear actuators (23) to vary the angle of the cold flaps. The annular assembly (12) is axially displaceable by a third set of linear actuators (30). Each cold flap (20) comprises an inner skin (40) fitted with orifices (43) and hinging on the annular assembly(12) and an outer skin (41) spaced radially away of the inner skin (40) to define an air flow channel between the skins.