Abstract: The present invention includes a gas turbine engine exhaust centerbody comprising a forward portion having at least one flap at a downstream end thereof to induce a local low pressure zone in the exhaust gas flow without significantly impeding it, a tailcone partially inserted into the forward portion so that at least part of the at least one flap overlaps a portion of the tailcone, a centerbody cavity containing gas at a pressure higher than the local low pressure, and at least one opening between the forward portion and the tailcone adjacent to and upstream of one of the local low pressure zones and providing fluid communication between the cavity and the exhaust gas flow. A ventilation flow out of the centerbody cavity through each opening is induced by a positive difference between the pressure of the gas in the centerbody cavity and the low pressure of the adjacent zone.

Abstract: A wall structure (2) configured to be exposed to a thermal load. The wall structure having at least two layers including a first layer (5) and a second layer (6). The second layer (6) is located closer to a source of the thermal load than the first layer (5), and the layers (5, 6) are arranged so that heat is allowed to be conducted from the second layer (6) to the first layer (5). Each of the first and second layers (5, 6) are adapted to carry a significant portion of a structural load, and the second layer (6) exhibits a higher thermal conductivity and/or a lower thermal expansion than the first layer (5). The invention reduces the thermal strain in the wall structure (2).

Abstract: A system and method for cooling at least a portion of an engine are provided. The engine is cooled using a fuel, such as a high heat sink fuel, that is subsequently used for combustion in the engine. The fuel can be used to cool one or more of the gases and/or components in the engine, thereby cooling the engine including an exhaust nozzle. For example, the fuel can be circulated through one or more heat exchanging devices that are disposed inside or outside a passage of the engine, and the fuel can absorb thermal energy from the engine or air that flows in the engine passage. In any case, the cooling of the engine can result in a reduction to the infrared signature of the engine.

Abstract: An exhaust nozzle includes an exhaust duct with an outlet and a row of radial apertures upstream therefrom. A radial frame surrounds the duct upstream from the apertures. A row of flaps are hinged to the frame to selectively cover and uncover the apertures for controlling exhaust flow discharged therethrough. An arcuate unison bar surrounds the duct adjacent to the frame and includes circumferentially spaced apart cams engaging corresponding cam followers affixed to the flaps. An actuator is joined to the bar for selective rotation thereof between opposite first and second directions to pivot open and closed the flaps atop the apertures.

Abstract: The present invention provides a plate-like body connecting method in which functions such that cooling capability is not reduced. There is provided a plate-like body connecting method for connecting at least two plate-like bodies each having passages that extend along at least one of opposite surfaces thereof and between the opposite surfaces, comprising the steps of disposing the plate-like bodies so that edges of the plate-like bodies are abutted with each other; connecting the plate-like bodies along the edges; forming a groove in the one surfaces of the plate-like bodies along the edges; and attaching a closure to and along the groove.

Abstract: A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

Abstract: A nozzle comprising an inner barrel, a centerbody disposed inside the inner barrel so as to define a first channel for a primary flow, and an outer barrel surrounding the inner barrel so as to define a second annular channel for a secondary flow, the free end of the inner barrel presenting a surface having a first series of undulations made up of negative undulations alternating with positive undulations, the centerbody presenting a second series of undulations made up of negative undulations alternating with positive undulations, the negative and positive undulations of the series of undulations being of radial height that varies angularly and being disposed in such a manner that the negative undulations of the first series are disposed facing negative undulations of the second series, and vice versa, so as to generate turbulence and radial shear between the primary and secondary flows while conserving a constant flow section around the entire circumference of the nozzle.

Abstract: Methods for manufacturing outlet nozzles for use in rocket motors are disclosed. The methods include providing an inner wall, positioning an outer wall around the inner wall, positioning spacers in the space between the inner and outer walls, and laser welding between the spacers and at least one of the inner and outer walls in order to form cooling ducts between the spacers and to provide T-shaped joints between the spacers and at least one of the inner and outer walls, the T-shaped joints including a rounded shape within the cooling ducts.

Abstract: The rocket motor nozzle assembly of this invention includes a throat insert and a carbon or silica protective eyelid. The throat insert has a carbon throat support and a refractory metal shell. The shell is positioned radially inside the throat support to cover the inner surface of the throat support. The protective eyelid covers a sufficient portion of the forward surface region of the shell and the underlying converging portion of the throat support to protect these components against particle impingement. The protective eyelid extends sufficiently far forward along the converging/diverging pathway to cover and protect the forward face or edge of the throat insert and prevent the combustion gases from passing under the throat insert and reaching the radially outer surfaces of the throat insert. However, the protective eyelid leaves the throat surface region of the shell exposed to the converging/diverging pathway.

Abstract: The pointed end of a central spike, used in an engine of a space launcher, may have a deformable structure that can adopt either a folded position or a deployed position.

Abstract: A nozzle extension including flange elements of C/SiC is able to be manufactured relatively simply and provides sufficient physical properties. The carbon-fiber structure of the nozzle extension is wound, and a flange for fastening the nozzle extension to the nozzle outlet of the combustion chamber is integrated in the nozzle extension by a continuous fiber flow of the wound carbon-fiber structure. The foregoing may be intended for use in rocket propulsion units.

Abstract: A turbofan exhaust nozzle includes a fan duct defined between a fan nacelle and a core engine cowling. The fan duct includes a longitudinal endwall and a partition spaced therefrom to define a secondary flow duct. The fan duct also includes a primary outlet, and the secondary duct includes a secondary outlet. The partition between the two ducts includes an aperture covered by a selectively movable flap.

Abstract: Disclosed is a segmented mixing device 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 while limiting high frequency noise through utilization of root radii in the order of 15 to 20% of the segment's width at the intersection of a parent nozzle.

Abstract: A convergent-divergent rocket nozzle is formed by joining two coaxially aligned conical sections at a throat plane, each diverging outward from the throat plane. Coolant channels formed in the wall of the nozzle are arranged in spirals around the nozzle axis. Preferably, the conical sections are formed from platelet laminates rolled into conical form with a single spiral seam, and at least one of the conical sections is a composite of two or more component conical sections separately formed and then joined in a nested arrangement with the seams not superimposed. A further preferred construction is one in which the convergent end of one conical section is split radially into strips that are then spread apart to serve as bonding surfaces to bond to the other conical section.

Abstract: A high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the gas flow path of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector the is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, such as on a centerbody.

Abstract: A gas turbine engine exhaust nozzle includes coaxial inner and outer conduits. The inner conduit has a main outlet at an aft end thereof, and a row of radial apertures spaced upstream from the outlet. The outer conduit has an auxiliary outlet at an aft end thereof, and surrounds the inner conduit over the apertures to form a bypass channel terminating at the auxiliary outlet. A plurality of flaps are hinged at upstream ends thereof to selectively cover and uncover corresponding ones of the apertures and selectively bypass a portion of exhaust flow from the inner conduit through the outer conduit in confluent streams from both the main and auxiliary outlets.

Abstract: Method and arrangement for providing an outlet nozzle (10) for use in a liquid fuel rocket engine. The nozzle forms a body of revolution having an axis (11) of revolution and a cross section that varies in diameter along said axis. The nozzle has a wall structure having a plurality of mutually adjacent cooling channels, helically extending substantially in parallel from the inlet end (13) of the nozzle to its outlet end (14). The nozzle includes at least two longitudinally arranged sections (10a,10b,10c). A shift between a positive and a negative channel angle in the transition from one section to an adjacent section balances any roll momentum generated by friction of rocket exhausts against the nozzle wall.

Abstract: Method and arrangement for providing a liquid fuel rocket engine member having a load bearing wall structure (11,14) including a plurality of cooling channels (11) for handling a coolant flow. Each cooling channel (11) is provided with a flow guiding surface (15) extending at an angle to the cooling channel axis, for providing the axial coolant flow with an added flow component in the radial direction.

Abstract: Method and arrangement for providing a liquid fuel rocket engine member (10). The member has a load bearing wall structure (14) having a plurality of cooling channels (11). For improving the heat transfer, a material with a higher thermal conductivity than the load bearing wall structure (14) is applied to the wall structure.

Abstract: In a first assembly step a coolant liner is formed having an outside surface. A plurality of coolant channels are formed on the outside surface. At least two inner conformal throat support sections and at least two outer conformal throat support sections are formed. The two inner conformal throat support sections are formed around the outside surface of the coolant liner to form a closed inner conformal throat outside surface. The inner conformal throat support sections are joined at inner throat seam lines. At least two outer conformal throat support sections are assembled around the closed inner conformal throat outside surface so as to cover the inner throat seam lines, to form a liner throat support assembly. An inlet manifold and an outlet manifold are assembled around the liner throat support assembly to form a hot isostatic press (HIP) assembly. A HIP'ed assembly is formed by HIP'ing the HIP assembly.

Abstract: A method and a system (10) for exhausting gas via a nozzle (19, 21) of a gas turbine engine, for example. The system (10) comprises a nozzle (19, 21) comprising a nozzle body portion (32, 33, 74) defining a nozzle exit (42, 43), characterised in that the nozzle body portion (32, 33, 74) comprises fluid injection means (60), positioned upstream of the exit (42, 43) relative to a fluid flow (F1, F2) created by the operation of the system, for injecting fluid (68) upstream of the exit (42, 43).

Abstract: A method and apparatus for augmenting thrust in a rocket traveling through atmospheric gas. Rocket motor designs are provided where a throat(s) from one or more rocket motors eject high-speed primary exhaust gas in a configuration which peripherally surrounds an outlet for induced, secondary gas. The secondary gas is mixed with the jet of primary exhaust gas to add momentum, and therefore thrust. Either expansion deflection or plug type rocket discharge nozzles can be utilized. In one embodiment, a thrust augmentation of over one hundred percent is achieved. In another embodiment, a plurality of rocket motor assemblies each containing a thrust augmenting rocket motor design is affixed to a rocket body. Such rocket motors enhance rocket thrust performance, and enables more efficient payload to rocket motor selection, or, alternatively, allows higher loads to be carried with the same amount of thrust.

Abstract: A segmented exhaust nozzle for attenuating noise from a turbofan jet engine without adversely impacting the operability or operability limit related performance of the engine. The exhaust nozzle includes spaced apart fan nozzle inner and outer walls which form an annular exhaust gas flow path therebetween. The fan nozzle outer wall is segmented at the downstream end. The outer wall curves inwardly towards the inner wall and then turns back away from the inner wall to form an arcuate protrusion that extends into the exhaust gas flow path forming an aerodynamic throat. Through the segmented portion of the nozzle, the outer wall then continues to curve away from the inner wall before again curving back towards the inner wall at a nozzle exit station. The nozzle exit effective area is approximately equal in cross sectional area to a conventional exhaust nozzle exit area.

Abstract: The rocket motor nozzle assembly of this invention includes a throat insert and a carbon or silica protective eyelid. The throat insert has a carbon throat support and a refractory metal shell. The shell is positioned radially inside the throat support to cover the inner surface of the throat support. The protective eyelid covers a sufficient portion of the forward surface region of the shell and the underlying converging portion of the throat support to protect these components against particle impingement. The protective eyelid extends sufficiently far forward along the converging/diverging pathway to cover and protect the forward face or edge of the throat insert and prevent the combustion gases from passing under the throat insert and reaching the radially outer surface of the throat insert. However, the protective eyelid leaves the throat surface region of the shell exposed to the converging/diverging pathway.

Abstract: The present invention features a fluid flow regulator that functions to significantly influence fluid flow across the surface of an object, as well as to significantly effect the performance of the object subjected to the fluid. The fluid flow regulator comprises a pressure recovery drop that induces a sudden drop in pressure at an optimal pressure recovery point on said surface, such that a sub-atmospheric barrier is created that serves as a cushion between the molecules in the fluid and the molecules at the object's surface. More specifically, the present invention fluid flow regulator functions to significantly regulate the pressure gradients that exist along the surface of an object subject to fluid flow. Regulation of pressure gradients is accomplished by selectively reducing the pressure drag at various locations along the surface, as well as the pressure drag induced forward and aft of the object, via the pressure recovery drop.

Abstract: Thrust chamber housing for a driving mechanism for space travel applications which is fastened to an injection head with its first end and which includes a combustion chamber housing of a highly heat-resistant steel, a nozzle element of a platinum-iridium alloy, and an expansion nozzle housing of a highly heat-resistant steel, whereby the combustion chamber housing is welded through a first intermediate ring with the nozzle element and the nozzle element is welded through a second intermediate ring with the expansion nozzle housing by means of welded connections, whereby the first and the second intermediate ring are made of a platinum-rhodium alloy.

Abstract: The present invention comprises a rocket motor nozzle that replaces the fixed-wall throat with a “wall” created by injecting gas radially into the nozzle. By injecting gas into the flow of combustion products that are going through the nozzle, this will deflect the combustion product flow, restricting and accelerating such flow, just as a fixed-wall does in a standard nozzle. However, by using the “gas-wall” described herein, no erosion will result at the restriction area.

Abstract: A high temperature sealing system is operative to seal an interface between adjacent hot structures and to minimize parasitic flow between such structures that move relative to one another in-plane or out-of-plane. The sealing system may be used to seal thrust-directing ramp structures of a reusable launch vehicle and includes a channel and a plurality of movable segmented sealing elements. Adjacent ramp structures include edge walls which extend within the channel. The sealing elements are positioned along the sides of the channel and are biased to engage with the inner surfaces of the ramp structures. The segmented sealing elements are movable to correspond to the contour of the thrust-directing ramp structures. The sealing system is operative to prevent high temperature thrust gases that flow along the ramp structures from infiltrating into the interior of the vehicle.

Abstract: A valve assembly for supplying cooling air to a nozzle of a gas turbine engine. The nozzle can move between a first configuration and a second configuration. The valve assembly includes a first member having openings that receive cooling air. The valve assembly also includes a second member having openings. The second member resides adjacent the first member to receive the cooling air from the first member. Finally, the valve assembly includes actuator for moving the first member from a first position, which allows a first flow rate of cooling air to pass through the openings, to a second position, which allows a second flow rate of cooling air to pass through the openings. The second flow rate is less than the first flow rate.

Abstract: Insulation for a rocket motor is provided, as is a method for insulating a rocket motor. The insulation includes a cured elastomer and vapor-grown carbon fibers dispersed in the cured elastomer. The cured elastomer is preferably formed from a precursor composition comprising an EPDM terpolymer. Generally, the vapor-grown carbon fibers have an internal graphitized tube surrounded by a sheath of vapor-deposited amorphous carbon.

Abstract: Methods and arrangements for tailoring rocket exhaust plume signatures of rocket exhaust systems are disclosed. The tailoring of the rocket exhaust plume signature is accomplished by providing and locating at least one structure having materials and/or additives that modify the radiant intensity pattern of the rocket exhaust plume upon ablation or melting into the rocket exhaust plume. The materials and/or additives can be manufactured or incorporated into preexisting ablative materials of the rocket exhaust system or be provided as stand-alone structures.

Abstract: An outlet device for a jet engine (2) includes an outlet channel (3), which has an upstream end (3′) for being connected to the jet engine and a downstream end (3″) and which defines a main flow direction (a) for a jet from the jet engine. The outlet channel has in the proximity of the downstream end an elongated shape, seen in a section across the flow direction, and includes at least two outlet portions (4, 5, 6), which are separated from each other, for a respective outlet flow of said jet. A first outlet portion (5, 6) includes means for controlling the outlet direction of the outlet flow in a first plane and a second outlet portion (4) includes means for controlling the outlet direction of the outlet flow in a second plane, which forms an angle to the first plane. An aircraft may include such an outlet device.

Abstract: A combustion apparatus including a structural jacket, a forward liner, and an aft liner, and an associated method of constructing the same are provided. The structural jacket defines a passage including a first end, a second end, and a neck. The neck is positioned between and separates the first and second ends of the passage. The forward liner is positioned within the first end of the passage and has a throat fitted within the neck of the passage. The aft liner is positioned within the second end of the passage and has an upstream potion abutting the throat of the forward liner. The forward and aft liners are brazed together, and are bonded to the passage of the structural jacket, to form a longitudinal combustion chamber so that combustion of the propellant is contained and directed through the combustion chamber from the forward liner to the aft liner.

Abstract: The invention relates to an apparatus for increasing the heat transfer to the coolant at the inside of a nozzle wall provided with coolant channels of expander cycle rocket engines. To achieve this the invention proposes that, to disturb the boundary layer at the nozzle wall and thereby increasing the heat transfer, the inside of the nozzle wall facing the flame has a particularly chosen increased surface roughness of such a magnitude that it penetrates the viscous sub-layer of the boundary layer at the nozzle wall.

Abstract: A pulsed detonation engine having improved efficiency has a detonation chamber for receiving a detonable mixture, an igniter for igniting the detonable mixture, and an outlet for discharging detonation products. A diverging-converging nozzle is provided at the outlet of the detonation chamber. The geometry of the diverging-converging nozzle is selected to enable a relatively short nozzle to significantly improve efficiency of the pulsed detonation engine.

Abstract: Systems and methods are disclosed for tailoring rocket exhaust plume signatures of fuel-propelled vehicles. An uncombusted fuel is injected or sprayed into the exhaust plume of an ignited rocket. The injected or sprayed fuel modifies the rocket exhaust plume to simulate an exhaust plume of a different vehicle. One or more parameters associated with the injected or sprayed fuel can be modified, so that appropriate tailoring can be achieved.

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: The invention concerns a cooling system for a post combustion jet nozzle (1), said jet nozzle comprising a primary gas duct (2) allowing a primary flow of gas, a secondary air duct (3) allowing a secondary flow of air, said secondary air duct surrounding the primary gas duct and separated from it by a protective thermal shroud (4), said secondary air duct having a downstream end (3a), dampers (5) surrounding an output section of the primary gas duct, the system comprising a protective thermal shell (7) in the secondary air duct, at the downstream end of said duct. According to the invention, the protective thermal shell bears an annular diaphragm (8) extending out in front of the dampers and being provided with support sectors (11) and inter-sector zones equipped with slots, said inter-sector zones defining spaces between the diaphragm and the protective thermal 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: According to one embodiment of the invention, a system for altering a fluid flow includes a nozzle having a fluid flow and including a converging portion, a diverging portion downstream of the converging portion, and a throat coupling the converging portion to the diverging portion, at least one port located in a wall of the nozzle and angled with respect to the fluid flow, and at least one pulse detonation device operable to inject a plurality of detonation waves in a pulsed manner through the port and into the fluid flow. The pulsed detonation waves operate to alter the fluid flow.

Abstract: A central turbo ducted fan integrated power plant unit comprises a central turbine having turbine blades and retaining ring, fan duct stator, internal combustion engine, fuel tank. There are provided central turbine blades on the retaining ring. The fan duct stator includes an outer ring, an inner ring, stator guide blades, turbine guide blades and the fixing supports on which the internal combustion engine and fuel tank are fixed. The central turbine is mounted on the fan duct stator and the inner hole or guide hole containing the central turbine blades is communicated with the inner hole of the inner ring. On the front end of the fan duct stator is provided a guide ring with diameter or inner diameter being reduced gradually from the front toward the rear, thereby the volume of air flow to be expelled can be increased and the air stream can be guided to be expelled linearly reward and the disordered air flow can be prevented from occurring.

Abstract: A gas turbine engine exhaust nozzle includes coaxial inner and outer conduits. The inner conduit has a main outlet at an aft end thereof, and a row of radial apertures spaced upstream from the outlet. The outer conduit has an auxiliary outlet at an aft end thereof, and surrounds the inner conduit over the apertures to form a bypass channel terminating at the auxiliary outlet. A plurality of flaps are hinged at upstream ends thereof to selectively cover and uncover corresponding ones of the apertures and selectively bypass a portion of exhaust flow from the inner conduit through the outer conduit in confluent streams from both the main and auxiliary outlets.

Abstract: A method of making heat engine components, such as rocket nozzle throats, includes forming a body having at least one surface, wherein the body is made of a refractory metal, and compressing a ceramic powder between a die and the at least one surface of the body, with sufficient heat and pressure to densify and thus form a solid ceramic coating on the at least one surface of the body. Nozzle throats made according to the invention have bodies made of refractory metals, such as tantalum, and ceramic coatings on the inner surfaces of the annular bodies, wherein the ceramic coatings are made of ceramic materials such as tantalum carbide.

Abstract: A gas turbine engine component includes a perforate metal wall having pores extending therethrough. The wall has a first surface covered by a thermal barrier coating. The pores have first ends which are covered by the thermal barrier coating, and the pores are ventilated from an opposite second surface of the wall for cooling the thermal barrier coating.

Abstract: The invention refers to a full-flow rocket nozzle having a longitudinal contour essentially corresponding to a parabola. For obtaining a flow separation control and also a side load reduction the invention suggests that the parabola contour shape from the point of 50% expansion ratio onwards or more distant from the throat of the nozzle is changing over into (i) strictly conical shape having an angle to the center line of between 15° and 25°, (ii) a slight outward curvature, i.e. implying a contour shape with positive 2nd derivative of the radius r, or (iii) a slight inward curvature, i.e. with a negative 2nd derivative of r but lying externally of the parabola contour, in the last case the 3rd derivative of r being constantly equal to zero (0).

Abstract: To control the separation of a hot gas ejection jet within a rocket engine nozzle, the device includes a jettisonable annular structure for placing around the outside wall of the nozzle at the gas cutlet section thereof, which structure defines a radial extension around the nozzle so that in the presence of an outlet jet coming from the nozzle, it creates a low pressure zone in the vicinity of the bottom face of the device, thereby displacing in the gas outlet direction jet separation from the interior wall of the nozzle.

Abstract: A rocket nozzle assembly is provided for operatively coupling to a rocket motor. The assembly includes a nozzle insert structure having a converging-diverging passageway with a throat region. First and second thrust control cylinders are parallel to each other and rotatable about respective axes transverse to the passageway central axis. The first and second thrust control cylinders intersect the throat region and include respective grooves. A thrust control cylinder-rotating subassembly is operatively associated with the first and second thrust control cylinders to rotate the cylinders about their respective axes and move the grooves relative to the throat region. Movement of the grooves changes the effective cross-sectional throat area of the passageway at the throat region for controlling the thrust when the rocket nozzle assembly is operatively engaged with an operating rocket motor.

Abstract: A method of making heat engine components, such as rocket nozzle throats, includes forming a body having at least one surface, wherein the body is made of a refractory metal, and compressing a ceramic powder between a die and the at least one surface of the body, with sufficient heat and pressure to densify and thus form a solid ceramic coating on the at least one surface of the body. Nozzle throats made according to the invention have bodies made of refractory metals, such as tantalum, and ceramic coatings on the inner surfaces of the annular bodies, wherein the ceramic coatings are made of ceramic materials such as tantalum carbides (TaC and Ta2C).

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 outlet device for a jet engine (2) includes an outlet channel (3), which has an upstream end (3′) for being connected to the jet engine and a downstream end (3″) and which defines a main flow direction (a) for a jet from the jet engine. The outlet channel has in the proximity of the downstream end an elongated shape, seen in a section across the flow direction, and includes at least two outlet portions (4, 5, 6), which are separated from each other, for a respective outlet flow of said jet. A first outlet portion (5, 6) includes means for controlling the outlet direction of the outlet flow in a first plane and a second outlet portion (4) includes means for controlling the outlet direction of the outlet flow in a second plane, which forms an angle to the first plane. An aircraft may include such an outlet device.