Abstract: The invention relates to a turbomachine (1) for an aircraft comprising a gas generator (4) designed such that hot gases are ejected from a combustion chamber (28) towards the upstream side of the turbomachine, and also comprising a plurality of hollow mixer struts (40) connected to an output (36) from the gas generator through which hot gases can pass, each mixer strut comprising an output mixer portion (40c) located inside the annular fan duct (8) so as to mix the secondary air and the hot gases ejected by these output mixer portions towards the downstream side of the turbomachine.

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. Also disclosed is a means for improving the thrust of jet engines and for reducing the noise and drag of aircraft. The device is a segmented, triangular or trapezoidal shaped, curved extension to a nozzle's sleeve or surface trailing edge that results in a serrated trailing edge. The extensions enhance the natural free mixing of the adjacent 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 achieved with free mixing and results in a more uniform flow in a shorter stream wise distance. As a result of the enhanced mixing, drag on flow surfaces that are downstream of the devices and are scrubbed by the flows adjacent to the devices can be reduced.

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: A second stage external jet nozzle mixer (20) includes identically formed lobes which equal in number the lobes of the first stage internal mixer. The external mixer works with the internal mixer, and furthers the mixing of the jet engine internal bypass flow with the internal jet engine core flow. This mixing levels the disparate flow velocities attendant with the jet engine exhaust, reduces the peak velocities from the jet engine core and increases the lower bypass velocities of the jet engine internal bypass flow. The lobes include complex curvatures that greatly enhance mixing of the gases and ambient cooling air, and thereby reduce noise. At the lobe terminus, the lobe dimensional characteristics may be adjusted to thereby adjust the total terminus area to achieve a match to a jet engine to cause that jet engine to run at a determined RPM and noise level. Noise attenuation may also be adjusted by changing lobe dimensions.

Abstract: A gas turbine engine exhaust nozzle includes a nacelle having an inlet and main outlet at opposite ends, and a main duct extending therebetween. A secondary bypass duct extends radially through the nacelle upstream of the main outlet and includes an unobstructed secondary inlet joined to the main duct, and a secondary outlet surrounding the main outlet for collectively discharging exhaust flow in confluent streams.

Abstract: A method facilitates assembling a gas turbine engine. The method comprises mounting a core engine to a vehicle, coupling a fuselage radially outward and around the core engine, and coupling an exhaust nozzle to the core engine to channel exhaust gases discharged from the core engine. In addition, the method also comprises coupling an infrared suppression system in flow communication with the engine exhaust nozzle for channeling exhaust gases discharged from said exhaust nozzle to facilitate suppressing an exhaust infrared signature of the core engine during operation, wherein the infrared suppression system includes an access door and a flow channel that is coupled to the access door such that the flow channel is movable with the access door from a closed position to an open position.

Abstract: A wet scrubber having a series of chambers for capturing and cooling exhaust gases generated during static test firing of rocket motors. Exhaust gas enters an inlet to a first chamber and is cooled and slowed by a spray solution. HCL gas is condensed and absorbed by the spray solution and precipitates to a liquid slurry at the bottom of the device. The remaining exhaust products enter a demister chamber where nozzles continue to spray the gasses as they pass upward and through a mesh-style demister at the top of the vessel. The demister filters liquid and solid waste particles from the gas stream, and the clean, dry gases are accelerated through a centrifugal fan into the atmosphere. A deflector is positioned within the inlet to the first chamber for containing parts in the event of a motor mal-function.

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: 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: An exhaust flow nozzle for a jet engine having a plurality of flow altering components extending from a lip portion of a secondary exhaust nozzle that are movable between first and second positions. In the first position the flow altering components are disposed substantially parallel to an exhaust gas flow path and thereby do not produce drag or a reduction of thrust from the engine. In the second position the flow altering components bend or are deformed to project into the exhaust gas flow path exiting from the secondary exhaust nozzle. The flow altering components are comprised of a shape-memory alloy material which deforms in response to heat. One or more additional layers of material are bonded or otherwise coupled to the shape-memory alloy layer of each flow altering component to assist in returning the shape-memory alloy layer to its unheated shape.

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: A segmented exhaust nozzle for attenuating noise from a 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: 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: 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: A tortuous path quiet exhaust eductor system may be mounted to a gas turbine engine, such as commercial aircraft APU. The system includes an oil cooler, eductor primary nozzle, oil cooler air nozzle, and surge air dump nozzle. The primary nozzle, oil cooler air nozzle, and surge air dump nozzle direct exhaust flow to entrain APU compartment cooling air, oil cooling air and surge air in a direction having both radial and axial components with respect to the APU centerline axis. The exhaust flow is directed into an eductor mixing duct angled away from the centerline axis and then is turned to enter an exit duct angled toward the centerline axis so that direct line of sight acoustic paths from the tail pipe exit to the turbine exit are blocked, suppressing core noise. The tail pipe ducts may be acoustically treated, further enhancing noise suppression.

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: 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: Disclosed is an exhaust mixer (50) including a passage (54) extending from an inlet (56) to an outlet (58) that is coincident with a centerline axis of mixer (50). Several ridges (68) are circumferentially disposed about the axis and each flare away from the centerline axis relative to a direction along the centerline axis from inlet (56) toward outlet (58). Ridges (68) each define a corresponding one of several inner channels (74) radially disposed about passage (54) that each intersect passage (54) between inlet (56) and outlet (58). Several outer channels (84) are also radially disposed about passage(54) and are each positioned between a corresponding pair of inner channels (74). Ridges are each shaped to turn inner channels (74and outer channels (84) about the axis as ridges (68) extend along the indicated direction. Inner channels (74) diverge away from the axis and one another in this direction while outer channels (84) converge toward the axis and one another in this direction.

Abstract: A centerbody for a gas turbine engine includes a thermal control system which minimizes the thermal stresses between the centerbody and at least one stiffener. The centerbody stiffener extends radially inward from a centerbody shell. A cavity is defined within each centerbody stiffener. The thermal control system includes a plurality of openings circumferentially disposed around the centerbody. Each opening extends through the centerbody shell into each cavity. The openings include pairs of entrance openings and exit openings which permit circumferential flow to develop within the centerbody cavity.

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: A jet noise suppressor is disclosed. The jet noise suppressor includes first and second high frequency pulse injection units located exteriorly of the turbine engine exhaust nozzle. The pulse injection units are located oppositely, diametrically across the exhaust nozzle. The pulse injection units include a resonance tube in outlet fluid communication with an injector nozzle. High pressure gas is supplied to the resonance tubes. A second source of high pressure gas includes a switch for supplying an alternating flow of high pressure gas to the injector nozzles. The combination of the high frequency pulses emitted from the injection units with the alternating low frequency, high amplitude mode of injection across the exhaust nozzle, provides a dramatic reduction of noise as well as exhaust temperature.

Abstract: A gas turbine exhaust passage is realized which may suppress the radiation of strong ultra low frequency noise to the outside without amplifying the turbulence of an high speed exhaust gas flow rate or a pressure pulsation generated in a gas turbine. At least a portion of a wall of a gas turbine exhaust passage (3, 5) is formed of an acoustically transmissive material (36, 56) for allowing a low frequency noise of several tens of Hz or less to pass therethrough sufficiently. Also, the acoustically transmissive material is made of one or more of a porous material, a porous heat insulating material, a mesh having a large flow resistance, cloth or film material. Further, the acoustically transmissive material is supported by a porous plate or a frame. In a case where an acoustically transmissive material is used only for an exhaust chimney (5), the exhaust chimney (5) is supported by a rack (11). Further, a soundproof panel (12) may be attached to the rack (11).

Abstract: An exhaust nozzle includes an annular exhaust duct for discharging exhaust from an aft end thereof. Circumferentially adjoining chevrons extend from the duct aft end around only an arcuate portion thereof leaving a plain arcuate shelf between terminal ones of the chevrons. Each of the chevrons has a triangular configuration and a compound arcuate contour both circumferentially and axially.

Abstract: A gas turbine engine exhaust nozzle includes an outlet for discharging exhaust, and a plurality of circumferentially spaced apart stub airfoils disposed adjacent the outlet. The airfoils shed vortices for mixing exhaust flow and attenuating noise and reducing infrared signature with minimum performance loss.

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: The present invention relates to a gas turbine engine exhaust nozzle for suppressing jet noise. The nozzle has an arrangement of tabs that extend in radially inward and outward directions. The tabs cause the formation of vortices which pull into mixing engagement the exhaust gas streams from the core and fan ducts and the ambient air. In addition, various construction details are developed for the tab arrangement including tabs that are directed radially inwardly and outwardly, and tabs that are continuous with the nozzle duct therebetween.

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: 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 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.

Abstract: An IR Suppressor (20) operative to suppress the infrared signature radiated from the high-temperature exhaust of an engine (12) and comprising a primary exhaust manifold (22), first and second mixing ducts (24a, 24b), and a secondary flow shroud (26). More specifically, the primary exhaust manifold (22) includes an elongate duct (28) adapted for receiving a primary flow of the high-temperature engine exhaust from the engine and at least two high aspect ratio nozzles (30a, 30b) integrated in combination with the elongate duct (28). The high aspect ratio nozzles (30a, 30b) are positioned so as to define at least two outlets at the same axial station along the length of the elongate duct (28). The first and second mixing ducts (24a, 24b) are disposed over the high aspect ratio nozzles (30a, 30b) so as to form a mixer/ejector defining at least two ejector inlets (44).

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 internal flow mixer is attached to the rearward end of a turbofan engine for inverting the engine fan and core streams. Rearwardly of the internal flow mixer is a multi-lobed exhaust nozzle which mixes ambient air with the gases flowing from the flow mixer. An ejector housing lined with acoustic absorption material receives the flow from the exhaust nozzle and discharges the thrust producing gases in noise abatement condition.

Abstract: An engine exhaust noise suppression system comprising a plurality of bypass ports surrounding the exhaust tailpipe and opening to the atmosphere to create streams of cool air which generate vortices effective to dampen the exhaust noise vibrations and to contain the same as they dissipate into the atmosphere. Fluid is injected into the air streams to enhance the vorticular action thereof by applying a helical swirling motion to the streams.

Abstract: A system is provided for reducing the emissions of hydrochloric acid (HCl) from solid fuel rockets, especially during ground disposal. An aqueous solution (40, FIG. 2 ) of an alkali metal hydroxide is injected as a mist (50) into the rocket chamber (32) as the rocket fuel (30) is burned. The reaction of the alkali metal with hydrogen chloride (HCl) produces a salt and thereby minimizes the presence of hydrochloric acid in the rocket exhaust. An injected neutralizing material which reduces hydrochloric acid, but which produces less thrust than an equal weight of rocket fuel, can be injected into an operating rocket which carries a payload high above the earth, with the injected material being injected only while the rocket is at a lower altitude when hydrochloric acid is most undesirable. The injected material can be produced by a small auxiliary rocket device (82, FIG.

Abstract: A propulsion system assembly for a vehicle such as an aircraft having a nacelle and shroud is disclosed. Various construction details are developed which permit the flow of cooling air into the full circumference of the shroud despite the blockage of the shroud by upstream structure such as a pylon on an aircraft engine. In one embodiment, the nozzle is disposed downstream of an engine in the nacelle and inwardly of the shroud. A passage 44 extends rearwardly in the pylon and underneath the shroud to duct cooling air into regions of the shroud blocked by the upstream structure.

Abstract: A method and apparatus for suppressing jet noise in a gas turbine engine powerplant 10 is disclosed. Various construction details are developed for providing sound suppression at sea level take-off operative conditions and not providing sound suppression at cruise operative conditions. In one embodiment, the powerplant 10 has a lobed mixer 152 between a primary flowpath 44 and a second flowpath 46, a diffusion region downstream of the lobed mixer region (first mixing region 76), and a deployable ejector/mixer 176 in the diffusion region which forms a second mixing region 78 having a diffusion flowpath 72 downstream of the ejector/mixer and sound absorbing structure 18 bounding the flowpath throughout the diffusion region. The method includes deploying the ejector/mixer 176 at take-off and stowing the ejector/mixer at cruise.

Abstract: A jet engine for use in a hypersonic aircraft. A fuel rich injectant is injected into the mixing zone of a duct at a substantial downstream angle to form an ejector. The temperature of the injectant is sufficiently low that spontaneous combustion of the injectant is delayed until the injectant reaches the combustion zone of the duct. In order to prevent combustion from propagating upstream from the combustion zone, the equivalence ratio of the mixture of injectant and ambient air flowing along the wall of the duct is maintained at a relatively high value, and the velocity profile of the flow along the wall of the duct is maintained at a relatively high value.

Abstract: An interturbine duct for interconnecting a gas turbine engine with a power turbine. The interturbine duct is of generally frusto-conical form and double skinned so that an annular cross-section volume is defined between the skins. Cooling air introduced in a generally tangential direction at the larger diameter extent of the duct follows a generally spiral path to the smaller diameter extent where it is exhausted into the efflux of the gas turbine engine. The spiral path of the cooling air ensures good heat transfer coefficients over both of the skins and obviates the need for baffles to direct the cooling air flow.

Abstract: A method of suppressing the formation of contrails from the exhaust of an engine operating in cold temperatures including the steps of providing a combined nucleating agent and freeze-point depressant selected from the group of water soluble monohydric, dihydric, trihydric or other polyhydric alcohols, or mixtures thereof, forming the solution into a vapor, and injecting the solution into the exhaust of the engine. The solution may include a non-corrosive surfactant. Another solution may include an organic or an inorganic nucleating agent, or mixtures thereof, in monohydric, dihydric or polyhydric alcohols, or mixtures thereof, and in addition may contain one or more surfactants.

Abstract: The diverging portion of discontinuous curvature is fitted with a controlled injection device for controlled injection of an additional jet of gas into the diverging portion in the same direction as the main flow of gas ejected by the nozzle, and said controlled injection device is disposed outside the diverging portion in the vicinity of the join zone between the first and second lengths in order to inject a determined flow of said additional gas into the diverging portion along a direction which is not tangential to the wall of the diverging portion.

Abstract: This invention relates to a double-flow turboshaft engine with confluent nozzle, i.e. of the type having a central generator emitting via its outlet orifice a flow of hot gas and an annular bypass duct which surrounds the central generator and through which passes a bypass flow of relatively cold gas, the gas flows being freely confluent within the nozzle which extends the annular duct in convergent manner beyond the outlet orifice of the central generator, wherein it includes controlled means for momentarily varying on the periphery the passage of the flux of relatively cold gas in the annular bypass duct, with a view to compensating, at least partly, the degradation of the performances of the turboshaft engine when they are limited either by the maximum operational temperature of the central generator or by the maximum power of the engine, but not by these two conditions simultaneously.

Abstract: A coolant pressure regulating appartus having a predetermined number of coolant pressure regulating valves which rotate against a spring bias to divert coolant flow and excess pressure from the coolant air fan duct into the engine core exhaust flow. The pressure regulator valve regulates the pressure within a closed cavity suppressor to the pressure level within the core without any affect upon the upstream fan duct pressure.

Abstract: A propulsion system having an afterburning turbojet with turbine bypass ducts in which the afterburner fuel is added. The hot gas products and fuel mixture enters the afterburner at sonic velocity. The turbojet turbine bypass has a flow control connected to the exit of the combustor to control the amount of turbine bypass flow.

Abstract: A technique for producing thrust by generating a hybrid plume plasma exhaust is disclosed. A plasma flow is generated and introduced into a nozzle which features one or more inlets positioned to direct a flow of neutral gas about the interior of the nozzle. When such a neutral gas flow is combined with the plasma flow within the nozzle, a hybrid plume is constructed including a flow of hot plasma along the center of the nozzle surrounded by a generally annular flow of neutral gas, with an annular transition region between the pure plasma and the neutral gas. The temperature of the outer gas layer is below that of the pure plasma and generally separates the pure plasma from the interior surfaces of the nozzle. The neutral gas flow both insulates the nozzle walls from the high temperatures of the plasma flow and adds to the mass flow rate of the hybrid exhaust. The rate of flow of neutral gas into the interior of the nozzle may be selectively adjusted to control the thrust and specific impulse of the device.

Abstract: A jet pump is provided for pumping a secondary fluid through a duct by injecting a primary fluid into the duct in the form of sheet-like jets lying in a plane parallel to the direction of flow of secondary fluid through the duct.

Abstract: An infrared radiation screening arrangement includes a special core structure and an outer gas guiding ring which together serve to block the direct view of a hot engine exhaust gas outlet and associated hot engine parts in order to hinder IR-detection of the engine. The gas guiding ring and the core structure define a cooling air guide channel with cooling air inlets for cooling the components of the arrangement and for cooling the exhaust gas stream which is ejected in an essentially unaltered direction.

Abstract: An exhaust system having a protective arrangement includes a plurality of spaced overlapping inclined armor plate blockers providing a ballistic shield over the entire exhaust area. A louvered grate is provided extending between the tops of the blockers to permit egress of gases from the engine but to block destructive elements from being thrown into the interior of an engine. Cooling air is supplied to the blockers and is caused to move along a path which carries it across the surfaces of the blockers so as to materially reduce the temperature of these surfaces. Further, this cooling air is directed not only over surfaces of the blockers which are directly visible externally of the vehicle, but also over a surface which, while not directly visible from the exterior of the vehicle, is a potential source of reflected heat which could be sensed exteriorly of the vehicle.

Abstract: A fluid dynamic pump, such as an ejector, has a common wall which separates a primary high energy fluid flow passage and a secondary low-energy fluid flow passage. The common wall includes a plurality of adjoining lobes extending in the direction of flow to the outlets of the passages at the downstream end of the common wall. A lobe on one side of the wall forms a corresponding trough on the other side. The common wall therefore has a wave-like shape at its downstream end. Primary and secondary fluid flow from the trough outlets on their respective sides of the common wall and are mixed together rapidly in a mixing region downstream of the outlets. A diffuser is located at the downstream end of the mixing region and enhances the pumping ability of the ejector. Secondary fluid is thereby drawn or pumped into the mixing region as a result of this transfer of energy from the primary stream to the secondary stream. Viscous losses are low during the process.

Abstract: The temperature of heated exhaust gases from a gas turbine is increased to value sufficient to vaporize fog oil to form smoke. The turbine itself is unmodified and requires no bleed air control valve at the air inlet to the compressor to regulate exhaust gas temperature.