Abstract: Systems and methods for passively directing aircraft engine nozzle flow are disclosed. One system includes an aircraft nozzle attachable to an aircraft turbofan engine, with the nozzle including a first flow path wall bounding a first flow path and being positioned to receive engine exhaust products, and a second flow path wall bounding a second flow path and being positioned to receive engine bypass air. The first flow path wall is positioned between the first and second flow paths, and the second flow path wall is positioned between the second flow path and an ambient air flow path. Multiple flow passages can be positioned in at least one of the first and second flow path walls to passively direct gas from a corresponding flow path within the flow path wall through the flow path wall to a corresponding flow path external to the flow path wall.

Abstract: The present disclosure relates to an improved and more efficient turbine deployed in the generation of thermal energy. The improved turbine comprises aerodynamic blades and supersonic nozzles to generate impulse and reaction forces which results into higher efficiency.

Abstract: The invention concerns a combustion chamber (10) comprising a neck (15) downstream of the injection (11) of gases, and downstream of this neck a divergent section (20) whereof the outer face of the wall (30), when in operation, is cooled by a cooling system using a cryogenic product and surrounding this outer face. This divergent section (20), on the inner face (32) of its wall (30), comprises a coating (40) acting as temperature compensator so that the temperature of the inner face (42) of the coating (40) is higher than the condensation temperature of the combustion gases on this inner face (42) under operating conditions, such that no condensation is formed on this inner face (42).

Abstract: An aircraft gas turbine engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments. Each of the basesheet segments includes a panel body between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. Leading and trailing edge ribs supporting the overlapping flanges are clipped together with longitudinally spaced apart and transversely extending forward and aft slits in retainers.

Abstract: An apparatus for reducing heating effects of an exhaust plume of a jet engine on an impinged surface includes fluid injectors disposed adjacent and aimed into an exhaust plume zone that's to be occupied by an exhaust plume when the engine is running. A flow generator transmits fluid flow into such an exhaust plume through the injectors. Each injector emits fluid in at least two divergent directions to increase the cross-sectional area of the exhaust plume by forming fluidic lobes in the exhaust plume.

Abstract: An exhaust eductor system includes a primary exhaust nozzle configured to transport an active flow stream; and a mixing duct at least partially surrounding the primary exhaust nozzle and configured to transport a passive flow stream that is entrained by mixing with the active flow stream from the primary exhaust nozzle. The mixing duct has an interior, and a baffle on the interior of the mixing duct is configured to prevent the mixed flow streams from exiting the mixing duct.

Abstract: A nozzle section of a gas turbine engine includes a regulator system in fluid communication with a secondary flow duct and a tertiary flow duct to selectively regulate communication of secondary airflow into the tertiary flow duct.

Abstract: An exhaust mixer (16) is provided for use in an engine exhaust system (10) downstream from an additive injector (14). The mixer (16) includes a first wall structure (20), and a second wall structure (22) carrying a set of mixer vanes (30), the second wall structure (22) being cone shaped and extending radially outwardly from the first wall structure (20).

Abstract: Supply of a liquid component in a combustion chamber of a rocket engine is controlled by a feed valve provided with an obturator mobile between a pen position and a closed position of at least one supply pipe, which has an inlet that communicates with a tank for containing the liquid component and an outlet that communicates with the combustion chamber; the displacement of the obturator from its closed position to its open position being triggered by a pressurized fluid supplied to the outlet of the supply pipe.

Abstract: An aircraft engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments having panel bodies between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges include first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. A basesheet assembly includes a frame includes first and second side rails supporting the first and second basesheet side edges. An aircraft gas turbine engine exhaust nozzle includes divergent seals with the basesheet assemblies in sealing engagement against divergent flaps along the first and second side rails.

Abstract: A propulsion system, such as for use in a missile, includes a multinozzle grid having a pair of plates that are separably mechanically coupled together. When coupled together in a first configuration, the plates provide multiple nozzles in a first nozzle configuration (geometry). Separation of the plates, such as by separating an aft plate from a forward plate that remains with the missile, reconfigures the multinozzle grid to a second configuration that has nozzles in a second nozzle configuration (geometry). The nozzle configurations may be suitable for different types of propulsion mechanisms. The hybrid propulsion system utilizing the multinozzle grid may include a pair of pressurized gas sources, for example a solid rocket fuel and a combustion chamber for a ramjet.

Abstract: A propulsion system may include a cylindrical support member and a tubular rotatable member rotatably mounted within the support member that may be adapted to permit fluid flow therethrough. The tubular rotatable member may extend past a down stream end of the support member. An exemplary embodiment of a propulsion system may also disclose a vane attached on an interior surface of the tubular member and may include a blade which extends in a direction toward a rotational axis of the rotatable member such that rotation of the tubular member and the vane attached thereon draws fluid into the tubular member to accelerate the fluid flow through the tubular member. Additionally, a nozzle may be attached to the down stream end of the support member and include a primary nozzle and a secondary nozzle within the primary nozzle. The secondary nozzle may be engaged with the primary nozzle by a stator.

Abstract: A plurality of hatches are distributed on the rear periphery of an engine cowl. The hatches are opened only when turbojet engine power is greater than cruising power. In an open position, the hatches draw hot air jets passing through an intermediate chamber before ejecting the jets through longitudinally-distributed communication components.

Abstract: A swept fan ramp for a pivot door thrust reverser includes a cylindrical unit with an elliptical flared portion, a side portion, and a rectangular flared portion. The elliptical flared portion is operable to reduce drag on airflow in a reverse direction and is coupled to the top center and bottom center of the circumference of the cylindrical unit. The side portion is operable to reduce side spillage airflow and is coupled to the central left side and central right side of the aft circumference of the cylindrical unit. The rectangular flared portion is operable to promote separation of airflow into an upper airflow path and a lower airflow path, is coupled with the aft circumference of the cylindrical unit, and is connected to the elliptical flared portion and the side portion.

Abstract: An exhaust nozzle includes a conical duct terminating in an annular outlet. A row of vortex generating duplex tabs are mounted in the outlet. The tabs have compound radial and circumferential aft inclination inside the outlet for generating streamwise vortices for attenuating exhaust noise while reducing performance loss.

Abstract: Eyeball seals for a gimbaled rocket engines, and associated systems and methods are disclosed. A system in accordance with a particular embodiment includes a rocket body, an engine carried by and movable relative to the rocket body, and a seal assembly. The seal assembly can include a sealing surface carried by one of the rocket body and the engine, and a seal element carried by the other of the rocket body and the engine. The seal element is in contact with the sealing surface. The seal assembly can further include a cylinder and a piston slideably received in the cylinder, with one of the piston and the cylinder carrying the seal element. The cylinder includes ports that are in fluid communication with a region external to the rocket body. Accordingly, pressures external to the rocket body can force the seal element and/or the sealing surface into contact with each other.

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, consequently, decreased fuel consumption and engine emissions.

Abstract: The invention relates to an arrangement for propelling an aircraft comprising a jet engine and an outlet nozzle arranged downstream of the jet engine. The jet engine is of a type which generates an internal core flow and an external fan flow. A part of the outlet nozzle has an internal surface which defines a gas duct for the fan flow. The aforementioned part of the outlet nozzle includes a wall structure that is arranged at a distance from the internal surface such that it separates the gas duct for the fan flow from an internal gas duct for the core flow.

Abstract: An engine exhaust nozzle comprises a plurality of CMC staves attached to one or more support rings arranged axially. The support rings provide a circumferential load path between the staves and for attaching the exhaust nozzle to the metallic engine components. The staves are fixed to the support rings with a spacing intended to accommodate for relative movement due to the difference in CTE for the CMC and metallic components and due to thermal gradients through the wall thickness. The resulting apparatus is lightweight, relieves the nozzle of thermal stresses, and is easier to manufacture and repair.

Abstract: A composite structure and an associated exhaust washed structure are provided which may be formed of ceramic matrix composite (CMC) materials. A method of fabricating a composite structure which may include the CMC material is also provided. A composite structure may include a corrugated septum extending in a lengthwise direction. The composite structure may also include a flute within which the corrugated septum is disposed to form, for example, a partitioned flute assembly.

Abstract: An acoustically treated exhaust centerbody comprising a body including a body fore portion and a body aft portion. An internal passageway extends through the body in the axial direction. A resonator in the body fore portion includes a plurality of acoustic chambers. A plurality of ribs in the resonator form fore/aft walls of the acoustic chambers. Each rib is shaped substantially as a section of an annulus. A plurality of radial fins extend between adjacent ribs. The fins form sidewalls of the acoustic chambers. A skin overlies the acoustic chambers and forms an outer surface of the resonator.

Abstract: A nozzle allowing the reduction of noise generated by aircraft jet engines having at least one output ring which disturbs the propulsion jet, and having at least one pair of conduits from which the propulsion of jet air is ejected at the output of the nozzle. The conduits of each pair are disposed convergent with respect to each other in order to generate a triangle of interaction of the air jets at the output ring.

Abstract: A supersonic combustor as a component of a rocket nozzle offers improved utilization of available chemical energy that may be released from combustion gasses flowing through the rocket nozzle. A subsonic combustor sub-sonically accelerates an exothermically reacting combustion gas up to a nozzle throat. The supersonic combustor expands and super-sonically accelerates the exothermically reacting combustion gas beyond the nozzle throat. The dimensions of the supersonic combustor may be selected such that the supersonic combustor achieves a slow rate of cooling of the combustion gasses without creating shockwaves within the supersonic combustor. A supersonic discharge expands and super-sonically accelerates the now substantially non-reacting combustion gas through a supersonic discharge of the rocket nozzle. The momentum of the combustion gas leaving the supersonic discharge propels the rocket nozzle in the opposite direction due to the principle of conservation of momentum.

Abstract: Disposed at or toward the trailing edge of one or more nozzles associated with a jet engine are injection ports which can selectively be made to discharge a water stream into a nozzle flow stream for the purpose of increasing turbulence in somewhat of a similar fashion as mechanically disposed chevrons have done in the known art. Unlike mechanically disposed chevrons of the known art, the fluid flow may be secured thereby increasing the engine efficiency. Various flow patterns, water pressures, orifice designs or other factors can be made operative to provide desired performance characteristics.

Abstract: A gas turbine exhaust component, such as an exhaust nozzle, is provided that includes circumferential ribs arranged along a longitudinal axis. Longitudinal strakes are supported by the circumferential ribs and arranged about the longitudinal axis. Fastening elements are used to secure the longitudinal strakes to one another to provide a desired exhaust component contour. In one example, fastening elements are used to secure the longitudinal strakes to the circumferential ribs using a bracket. The longitudinal strakes are preformed into the desired exhaust component contour, using low cost forming methods. The longitudinal strakes are formed to fit the circumferential ribs and provide the desired exhaust component contour. The longitudinal strakes are secured to one another to maintain the desired exhaust component contour. The longitudinal strakes may be provided by a low cost steel, in one example.

Abstract: A method and device for adding special substances to the gaseous outflow of the jet engine in order to create a “virtual wall” of increased pressure zone behind the jet engine nozzle which can serve as a support for the jet engine gas outflow to push against, thus increasing the thrust power of the jet engine. This increase in acceleration power results in the accelerated movement of the jet engine equipped vehicle or higher fuel efficiency. A device that adds the special substances to the gaseous outflow is proposed. Characteristics of the special substances, which if added into the jet engine gaseous outflow may create a “virtual wall” of increased pressure zone behind the jet engine, are proposed.

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: An integral composite rocket motor nozzle. The novel nozzle includes a first layer of a first reinforcement material, a second layer of a second reinforcement material, and a common matrix material surrounding the first and second reinforcement materials such that the reinforcement materials and matrix material form a single integral composite structure. In an illustrative embodiment, the first reinforcement material includes graphite fibers for providing structural support, and the second reinforcement material includes glass or quartz fibers for providing thermal insulation on a first side of the first layer. The nozzle may also include a third layer of a third reinforcement material for providing thermal insulation on a second side of the first layer. In a preferred embodiment, the first layer is shaped to form an integrated dome and nozzle structure.

Abstract: An exhaust duct assembly includes a cooling liner spaced apart from an exhaust duct case that articulate for use in a short take off vertical landing (STOVL) type of aircraft. The cooling liner assembly is attached to the exhaust duct case through a foldable attachment hanger system. The foldable attachment hanger system provides a low profile (foldable up/down) for a limited access installation envelope typical of a three bearing swivel duct (3BSD) which rotates about three bearing planes to permit transition between a cruise configuration and a hover configuration. In this way, each cooling liner segment may be formed as a complete cylindrical member requiring joints only between the swivelable duct sections.

Abstract: Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter.

Abstract: A variable area fan nozzle for use with a gas turbine engine system includes a nozzle section that is attachable to a gas turbine engine for influencing flow through the fan bypass passage of the gas turbine engine. The nozzle section is movable between a first length and a second length that is larger than the first length to influence the flow. For example, the nozzle section includes members that are woven together to form collapsible openings that are open when the nozzle section is moved to the first length and that are closed when the nozzle section is moved to the second length.

Abstract: A fastener for a turbine engine includes a bolt and a ceramic matrix composite (CMC) cap mechanically secured to the bolt. CMC components are attached to an exhaust nozzle case with the fasteners. The fastener includes a bolt made of metal and having a head. A CMC cap is located on the head of the bolt and secured to the bolt by a mechanical device. The mechanical device allows the CMC cap to have a different thermal expansion than the bolt. The CMC cap also acts as a thermal shield for the bolt, aiding in the reduction of a heat signature of the turbine engine.

Abstract: Jet engine nozzles with projections (e.g., chevrons) and injected flow, and associated systems and methods are disclosed. A method in accordance with one embodiment includes generating a first flow of gas with a jet engine, delivering the first flow through a nozzle having a trailing edge perimeter that includes multiple projections extending in an aft direction, and injecting a pressurized second flow of fluid into the first flow proximate to the projections. In other embodiments, other mixing enhancement devices (e.g., vortex generators) are carried by the projections. It is expected that the combination of the projections and the mixing enhancement devices will reduce engine exhaust noise levels below the levels achievable with either projections or injected flow individually.

Abstract: An exhaust educator system includes a primary exhaust nozzle configured to transport an active flow stream; and a mixing duct at least partially surrounding the primary exhaust nozzle and configured to transport a passive flow stream that is entrained by mixing with the active flow stream from the primary exhaust nozzle. The mixing duct has an interior, and a baffle on the interior of the mixing duct is configured to prevent the mixed flow streams from exiting the mixing duct.

Abstract: A turbofan engine includes a fan variable area nozzle having a multiple of vents through a fan nacelle and rotatable elements rotatable within the vents by an actuator system. Rotation of the rotatable element within each vent changes the effective area of the fan nozzle exit area to permit efficient operation at a multiple of flight conditions.

Abstract: A gas turbine engine exhaust nozzle includes a fluted shell terminating in a row of chevrons. The nozzle is radially serpentine circumferentially around the shell and has a circumferentially serpentine trailing edge.

Abstract: A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

Abstract: A mixer for a turbofan engine includes a centerbody and a mixer nozzle. The mixer nozzle surrounds at least a portion of the centerbody and is spaced apart to define a core flow path between the mixer nozzle and the centerbody. The mixer nozzle is configured, when bypass air flows through the turbofan engine, to direct at least a portion of the bypass air to impinge on the centerbody. The mixer nozzle includes a plurality of circumferentially spaced mixer lobes that extend axially in a rearward direction and have a cross-section shape defined by a set of equations.

Abstract: An annular exhaust duct of a gas turbine engine includes inner and outer cases co-axially disposed, and radially spaced apart by a plurality of airfoils radially extending therebetween. The outer case includes a connection apparatus for supportably connecting the exhaust duct to the gas turbine engine and the inner case includes an axial end portion at a rear end thereof for connection with a tail cone. The axial end portion bends radially and inwardly at a rear extremity thereof to stiffen the rear end of the inner ease.

Abstract: An exhaust nozzle for a gas turbine engine includes a modular chevron. The chevron includes dual skins fixedly joined together at a base flange, and at a rim extending along the trailing edge of the chevron between the base and apex thereof.

Abstract: An exhaust nozzle assembly includes an inner liner exposed to hot combustion gases and an outer liner spaced a radial distance from the inner liner to form an annular chamber. The inner liner includes a hot side that is directly exposed to the hot combustion gas flow and a cold side is exposed to cooling air within the chamber. The outer liner includes an outer surface exposed to cooling air flow up to a restriction preventing communication of cooling air flow. A plenum is attached to the outer liner to define a plenum chamber that extends into an end portion. The plenum chamber receives cooling air flow from a supply opening and communicates that air to an end portion through a plurality of impingement openings that provide impingement flow of cooling air outboard of the restriction.

Abstract: An outlet device for a jet engine includes an outlet nozzle and a plurality of guide vanes arranged movably in the outlet nozzle for guiding a gas from the jet engine for the purpose of steering a craft equipped with the jet engine. At least one of the guide vanes can be adjusted into such a position that hot parts of the engine located inside the outlet nozzle are at least substantially concealed seen from the outlet side of the outlet nozzle.

Abstract: The cross sectional flow area of a fan discharge nozzle on one side of a central plane of an associated gas turbine engine power plant is greater than the corresponding flow area of the fan discharge nozzle on an opposite side of the central plane to compensate for the blockage of fan airflow by a pylon.

Abstract: Dielectric barrier discharge plasma actuators are used to manipulate exhaust flow within and behind a jet engine nozzle. The dielectric barrier discharge plasma actuators may be used to direct cooling airflow near the surface of the nozzle to reduce heating of the nozzle, create thrust vectoring, and reduce noise associated with the exhaust flow exiting the nozzle.

Abstract: A device for attenuating acoustical noise from an exhaust nozzle of a turbojet is disclosed. The turbojet exhaust nozzle includes a central body defining the inner surface of the primary gas flow path. The central body includes an outer wall defining one, and only one, resonator cavity. The single resonator cavity is in fluid communication with the primary gas flow path via orifices through at least one upstream portion of the outer wall. The single resonator cavity and plurality of orifices compose a Helmholtz resonator.

Abstract: A device for attenuating acoustic noise from an exhaust nozzle of a turbojet is disclosed. The turbojet exhaust nozzle includes a central body spaced apart from an external casing, defining a gas exhaust channel therebetween. The central body includes a flexible sleeve within an external wall, defining at least one resonator cavity therebetween. The at least one resonator cavity is in fluid communication with the gas exhaust channel via a plurality of orifices through at least one upstream portion of the external wall. The at least one resonator cavity and plurality of orifices compose at least one Helmholtz resonator.

Abstract: Within a turbine engine a heat exchanger may be provided to cool compressor air flows to be utilized for cabin ventilation or other functions. A fluid flow acting as a coolant for the heat exchanger is generally drawn from the by-pass duct of the engine and a dedicated outlet duct is provided such that the exhausted fluid flow from the heat exchanger is delivered along the conduit duct to a low pressure region. In such circumstances, an appropriate pressure differential across the heat exchanger is maintained for operational efficiency when required, whilst the exhausted fluid flow is isolated and does not compromise the usual engine ventilation vent exit area sizing. Over-sized ventilation vents would create an aerodynamic step and therefore drag upon the thrust of the engine.

Abstract: A support and air mixer includes an inner annular member, an outer annular member shaped and positioned such that an annular air gap is formed between the inner annular member and the outer annular member, and a plurality of air mixing members, each air mixing member having a mixing passageway therethrough and each air mixing member having an inner portion connected to the inner annular member, an outer portion connected to the outer annular member, and first and second opposing side portions interconnecting the inner and outer portions and the inner and outer annular members such that the mixing passageway is in fluid communication with the annular air gap.

Abstract: Systems and methods for passively directing aircraft engine nozzle flow are disclosed. One system includes an aircraft nozzle attachable to an aircraft turbofan engine, with the nozzle including a first flow path wall bounding a first flow path and being positioned to receive engine exhaust products, and a second flow path wall bounding a second flow path and being positioned to receive engine bypass air. The first flow path wall is positioned between the first and second flow paths, and the second flow path wall is positioned between the second flow path and an ambient air flow path. Multiple flow passages can be positioned in at least one of the first and second flow path walls to passively direct gas from a corresponding flow path within the flow path wall through the flow path wall to a corresponding flow path external to the flow path wall.

Abstract: A method of assembling a crank assembly for a turbine engine. The turbine engine includes a valve assembly including an outer fairing and an inner fairing coupled to the outer fairing with a strut. The valve assembly further includes an annular slide valve coupled between the inner and outer fairings. The valve assembly is positioned within a duct having a radially outer duct wall and a radially inner duct wall. The method includes coupling a first arm of a crank assembly to the annular slide valve, and coupling a second arm of the crank assembly to the outer fairing such that the crank assembly controls movement of the annular slide valve and for moving the outer and inner fairings between a first operational position and a second operational position to facilitate regulating an amount of airflow channeled through the turbine engine.