Abstract: A propulsion system includes a main gas turbine engine adapted for generating propulsive thrust during subsonic and supersonic flight operations and a supplementary propulsion unit adapted for generating additional thrust. The supplementary propulsion unit has an air intake and an exhaust for gas accelerated by the supplementary propulsion unit to provide the additional thrust and is adapted to generate the additional thrust during a limited range of subsonic flight operations, and to be dormant during other flight operations. The propulsion system has housing for the supplementary propulsion unit, including intake and exhaust covers which are moveable between deployed and stowed configurations. During the limited range of subsonic flight operations the intake and exhaust cover are moved to the deployed configuration to open the intake and the exhaust. During other flight operations the intake and exhaust cover are moved to the stowed configuration to close the intake and the exhaust.

Abstract: A hybrid propulsion system and methods for cooling the same are provided. The system may comprise a gas turbine and a secondary engine. The gas turbine engine may have a core passage and an engine compartment. The secondary engine may be a supersonic and/or hypersonic engine. The system may comprise a thermal barrier, an inlet and an exhaust. The thermal barrier may longitudinally envelope the gas turbine engine. The thermal barrier may comprise an inner envelope, an outer envelope, an upstream opening, and a downstream opening. The inlet may be in fluid communication with the ambient environment and the gas turbine engine via the upstream opening. The exhaust may be in fluid communication with the ambient environment and the gas turbine engine via the downstream opening. The engine compartment may be located between a boundary of the core passage and the inner envelope.

Abstract: A flow control system (22) includes a fuel passage network (34) that has first (36) and second (38) network portions that are in a parallel flow arrangement with each other. A fueldraulic device (40) is located in the first network portion. Operation of the fueldraulic device varies flow through the first network portion. A flow restriction orifice (42) is located in the fuel passage network and is arranged in series with, and upstream of, the fueldraulic device. The flow restriction orifice is operable to generate a pressure differential that varies responsive to the flow through the first network portion. A flow control valve (44) is located in the second network portion. The flow control valve is operable responsive to the pressure differential across the flow restriction orifice to control flow through the second network portion.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a translating inlet assembly are provided. In one embodiment, a core engine of a gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include an inner flowpath surface. A core casing can enclose the core engine. A forward end of the core casing can include a translating inlet assembly moveable between a first position and a second position. The translating inlet assembly and the inner flowpath surface can together define an inlet to an engine airflow path. A translating inlet assembly can define a first inlet area in the first position and a second inlet area in the second position, the first inlet area being greater than the second inlet area.

Abstract: An engine system that produces all required thrust for an aerospace vehicle from takeoff through space operation utilizing a turbo ram rocket exhaust nozzle and M-Spike rocket consisting of airbreathing and non-airbreathing propulsion apparatuses. The airbreathing system consists of a turbine engine, a ramjet or scramjet, and the non-airbreathing system is a rocket motor. The turbine engine consists of a turbojet or turbofan configuration. The air breathing turbine, ramjet or scramjet feature an air inlet mechanism, and combustion fuel. The non-airbreathing rocket system includes separate oxidizer system, and either a separate or same source of combustion fuel as the turbine. Airflow velocities in the turbine bypass duct, and burner system, include subsonic and supersonic velocities for ramjet or scramjet operation. The rocket engine utilize either cryogenic or a non-cryogenic fuel and oxidizer system.

Abstract: A cooling system for a turbine engine including a heat exchanger in fluid communication with a first fluid inlet stream and disposed upstream and in fluid communication with a core engine. The heat exchanger operative to cool the first fluid inlet stream. The heat exchanger including a heat exchanger inlet for input of a heat exchanging medium for exchange of heat from the first fluid inlet stream to the heat exchanging medium. The heat exchanger further including a heat exchanger outlet for discharge of a heated output stream into one of a turbine of a downstream engine, an augmentor or a combustor of the core engine. The heated output stream provides an additional flow to the downstream engine. A turbine engine including the cooling system is disclosed.

Abstract: An aircraft includes a fuselage extending between a forward end and an aft end. The aircraft additionally includes a propulsor mounted to the fuselage at the aft end of the fuselage, the propulsor including an outer nacelle and the outer nacelle defining an inlet. Additionally, the aircraft includes a deployable assembly attached to at least one of the fuselage or the outer nacelle, the deployable assembly movable between a stowed position and an engaged position. The deployable assembly alters an airflow towards the propulsor or into the propulsor through the inlet defined by the outer nacelle when in the engaged position to increase an efficiency of the aft fan and/or of the aircraft.

Abstract: The system (10) provides a modulator ring (54) secured adjacent an exhaust end (56) of an inner wall of a by-pass fan air duct (40) and secured adjacent a convergent nozzle section (38) of the engine wherein flow of by-pass fan air (42) and core engine air (18) converge within an exhaust nozzle (34), The modulator ring (54) may be actuated to extend beyond the exhaust end (56) of the duct (40) to decrease air passage (62) forcing more air into the core engine (14) for greater thrust from the core engine (14) with minimal loss of thrust from the by-pass fan air (42). The modulator ring (54) stay alternatively rotate to uncover trailing-edge slots in the exhaust end (56) of the duct to also modulate flow rates of by-pass fan air (42) into the convergent nozzle section (38).

Abstract: A nozzle is provided that is capable providing flowpaths for a combined cycle aircraft propulsion system that in one form includes a gas turbine engine and a ramjet. The gas turbine engine produces an exhaust flow that is offset from an exhaust flow from the ramjet. The two streams can be flowed independent of each other or together depending on the application and relevant portion of a flight envelope. The nozzle includes a movable portion that can selectively open and close an exhaust flowpath for the gas turbine engine. The nozzle includes a surface that provides expansion for both low speed (gas turbine engine) flow and high speed (ramjet) flow.

Abstract: An engine accessory system for a gas turbine engine includes a first accessory component defined along an accessory axis and a second accessory component mounted to the first accessory component along the accessory axis.

Abstract: A variable cross-section nozzle of an aircraft has at least two movable parts enabling modification of the cross-section of the nozzle. A control device for the nozzle includes a system for regulating the engine, the system being connected to a control element which drives at least one electric motor which creates the movement of the movable parts, including an electric motor for each movable part and two mechanical transmission systems for each movable part connected only to the same electric motor.

Abstract: A heat-exchange architecture for exhaust of a turbine engine, including heat-exchange elements partially obstructing a hot exhaust gases stream. In an architecture built into an exhaust line for gas streams of a turbine engine, the turbine engine is a turboshaft engine including a gas generator and a free turbine supplying power to a shaft via a through-shaft and an upstream reduction gear. An annular plate-shaped exchanger is installed in the axisymmetric portion of the ejector. The exchanger includes inlet and outlet channels, connected to the inlet and outlet of a central channel wound into a helix shape or sine curve in the annular plate. The channels are connected at the other ends to mechanical or electromechanical mechanisms for recovering and recycling energy. The energy is recovered via a cold fluid heated in the central channel by transfer of heat from residual gases.

Abstract: A building energy recovery system having at least one wind drive mounted to the building structure proximate to the roof junction, a generator and an electrical output. The wind drive includes a radial flow wind turbine having a rotor, a rotatable shaft of the rotor, an open windward facing air intake duct, and a side facing air discharge opening disposed perpendicular to the rotatable shaft of the rotor. The intake duct is in communication with a wind blowing toward the front wall and the discharge opening is in communication with a wake region adjacent to the discharge opening and induced by a wind flowing behind the roof junction.

Abstract: The invention relates to multimode aircraft operated at supersonic and subsonic flight speeds in a wide range of flight altitudes. The invention is particularly applicable in multimode super-maneuverable aircraft capable of cruising at supersonic speed and having low radar signature. The technical effect to be achieved by the invention is to provide an aircraft which has low radar signature, is super-maneuverable at high angles of attack, has high aerodynamic efficiency at supersonic speeds while keeping high aerodynamic efficiency at subsonic modes, and can accommodate bulky cargo in its internal compartments. The integral aerodynamic configuration aircraft comprises fuselage (1) with strake (2), wing with panels (3) smoothly blended into fuselage (1), all-moving horizontal tail (AMHT) (4), and all-moving vertical tail (AMVT) (5). Mid-fuselage is flattened and formed longitudinally by a set of aerodynamic profiles.

Abstract: An aircraft powerplant is disclosed that can be operated in at least three modes including as a gas turbine engine, as an engine having a ramburner, and as an engine having a forward compression combustor engine such as a ramjet and/or scramjet. An airflow valve is provided to direct air to a downstream portion of the aircraft engine and can be positioned as a function of the aircraft operating modes. The valve can be used to cocoon the gas turbine engine.

Abstract: The jet (40) has an air-fuel combustion chamber (42) and a plurality of rocket engines (11) arranged upstream from the combustion chamber (42), each rocket engine having its own combustion chamber with the wall thereof being cooled by lateral injection of fuel through said wall.

Abstract: Some embodiments provide a dual use hydrazine propulsion thruster system. The dual use hydrazine propulsion thruster system may include a storage vessel configured to store hydrazine and a chemical thruster configured to provide thrust using the hydrazine. The dual use hydrazine propulsion thruster system may also include a convertor configured to catalytically convert hydrazine to a product comprising ammonia, and an electric thruster configured to provide thrust by ionizing ammonia and accelerating the ionized ammonia out of the electric thruster.

Abstract: The present invention relates to a mechanical system and method that modulates airflow in an aircraft inlet diffuser that is used in conjunction with an aircraft engine that integrates both a center turbine engine and a high Mach engine such as a constant volume combustor (CVC) arrangement or ramjet arrangement with intakes formed co-centrically about the turbine. The modulation system uses two articulating components, a movable air flow duct and an articulating cone. The air flow duct, in a first position, is in exclusive air flow communication with the circular intake face of the turbine in a first position to receive air from the intake diffuser. In this configuration the expandable cone is in a retracted position and does not redirect airflow and allows the aircraft to operates in low speed mode as only the turbo jet receives airflow. In a transition speed mode, the air flow duct is retracted to allow air flow to both the turbo jet and the CVC.

Abstract: A cyrogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid.

Abstract: The invention relates to a turbomachine thermomechanical part forming a body of revolution about a longitudinal axis, and including at least one abradable ring for a labyrinth seal. In characteristic manner, the abradable ring is made up of angular sectors that present different stiffnesses between adjacent pairs of sectors. The invention is applicable to a compressor, a turbine, or to a rotor and stator assembly.

Abstract: The present invention provides a plasma arc torch that can be used for lean combustion. The plasma arc torch includes a cylindrical vessel, an electrode housing connected to the first end of the cylindrical vessel such that a first electrode is (a) aligned with a longitudinal axis of the cylindrical vessel, and (b) extends into the cylindrical vessel, a linear actuator connected to the first electrode to adjust a position of the first electrode, a hollow electrode nozzle connected to the second end of the cylindrical vessel such that the center line of the hollow electrode nozzle is aligned with the longitudinal axis of the cylindrical vessel, and wherein the tangential inlet and the tangential outlet create a vortex within the cylindrical vessel, and the first electrode and the hollow electrode nozzle create a plasma that discharges through the hollow electrode nozzle.

Abstract: A stable and play-free pivot connection of a bicycle derailleur between a parallelogram link and either a base member or a chain cage. The pivot points are formed with first and second bushings that are arranged as far from one another as possible. The parts are connected by pivot pins having a anti-rotation contour for preventing rotation of the pin within an inner pivot part. The axial displacement of the pivot pin is prevented by a securing element or by a deforming element engaging a holding contour on the pivot pin. One of the first and second bushings is protected and arranged in a blind hole and the other of the first and second bushings is covered on the outside by a pin head of the pivot pin.

Abstract: An insensitive combined cycle missile propulsion system includes a solid fuel contained within a first section of the missile, a liquid oxidizer contained within a second section of the missile and a solid oxidizer contained within a third section of said missile. A first conduit has a first valve communicating the fuel and the oxidizer and a second conduit, spatially removed from the first conduit, has a second valve communicating the fuel and the oxidizer. An inlet system for delivering atmospheric oxygen for combustion with the fuel rich gases generated within the missile and a nozzle exhausts combustion products that result from combustion of the fuel, the liquid and solid oxidizers, and air.

Abstract: A gas delivery system includes a gas booster module for delivering natural gas from a utility gas service to power generation equipment installed in or around a building in a manner that meets the minimum volume and pressure requirements of the power generation equipment. The gas delivery system advantageously uses pipe of a relatively small size for delivering gas to the power generation equipment, thereby substantially reducing installation costs and eliminating the need for a welded gas line. The gas delivery system also provides a control system that facilitates close control over the gas flow and ensures compliance with local building codes and safety regulations and requirements.

Abstract: A combined ramjet and rocket engine, operates as a ramjet in atmosphere, and a rocket engine in space is provided with a rectilinear duct. The engine has interior of the rectilinear duct paired and hinged rectilinear baffles which move from one of the duct walls. These hinged rectilinear baffles move into and out of the planar duct wall and define both the constriction of the Venturi for the ramjet engine configuration or the closure of the duct for the rocket engine configuration. The hinged rectilinear baffles are given leading and trailing wall surfaces so that during movement within the duct for the ramjet configuration, alteration of the slope of the collector and diffuser of the Venturi can occur. The hinged rectilinear baffle bordering the diffuser is manifolded to supply fuel when the engine is ramjet configured and fuel and oxidant when the engine is rocket engine configured.

Abstract: An integral turbo jet/ram jet propulsion system has co-axial forward and rear portions and includes a converging supersonic compression chamber and an inlet in the forward portion for directing air into the supersonic compression chamber. A diverging subsonic compression chamber is disposed rearwardly of the supersonic compression chamber. An inlet sonic throat connects the supersonic compression chamber and the subsonic compression chamber and directs a flow of air from the supersonic compression chamber through the throat and into the subsonic compression chamber. A high by-pass turbo jet engine is disposed adjacent to the subsonic compression chamber for heating and expanding the air which passes into the engine. A subsonic expansion exhaust chamber and supersonic expansion exhaust chamber are disposed rearwardly of the engine with an exhaust sonic throat between the two.

Abstract: A jet engine includes a gas turbine, peripheral ramjet channels located oide the gas turbine and including burners wherein a gas circulation stream in the engine is subdivided over part of its length into a central turbojet stream which flows through the gas turbine, the peripheral ramjet channels and peripheral discharge channels terminating by flowing into a part of the gas circulation stream downstream of the turbine. A plurality of flaps are provided for opening and closing the central stream. A ring surrounds the central stream and extends in front of part of the peripheral discharge channels and ramjet channels. A motor is also provided for rotating the ring wherein the ring has a plurality of slots formed therein, the channels being arranged such that the slots successfully pass, during rotation of the ring in a continuous direction, in front of the discharge channels only, then in front of none of the channels and then in front of the ramjet channels only.

Abstract: A combined turbofan and ramjet aircraft engine includes a forward bypass duct which allows the engine to operate more efficiently during the turbofan mode of operation. By mounting a ramjet preburner in the forward duct and isolating this duct from the turbofan bypass air, a transition from turbofan operation to ramjet operation can take place at lower flight Mach numbers without incurring pressure losses or blockage in the turbofan bypass air.

Abstract: A variable-cycle storable fuels engine having three cycles is provided. The engine operates as an air-augmented rocket from launch to approximately Mach 3. From approximately Mach 3 to Mach 8, the engine operates as a dual mode ramjet/scramjet to altitudes exceeding 100,000 feet. Thereafter, the engine operates as a conventional pure rocket for final sub-orbital boost. A fixed geometry air inlet provides improved ram recovery at altitudes below 100,000 feet.

Abstract: Turboramjet engine having a variable air intake, a ram air duct annularly enveloping the basic turbojet engine, and shutoff devices with which the ram air duct is energized with air when the basic turbojet engine is shut off and with which the basic turbojet engine is energized with air when the ram air duct is shut off. An air tract communicates with the air intake on a tubular duct carried axially on the basic turbojet engine and expands to an annular inlet section to the ram air duct. The shutoff devices are associated at least on the inlet side with quadrilateral sections of the tubular duct. As shutoff means for the inlet flow, use is made of an axially adjustable and internally swept sleeve valve and at least one flap that can be pivoted against a partial inlet area of the inlet section, which flap seals off the sleeve valve in lid-fashion when the inlet section is in a jointly exposed position.

Abstract: The combined propulsion engine comprises an air-breathing jet engine and a non-air-breathing rocket engine, the jet engine comprising an air inlet device and an external nozzle which, in association with a central body, defines an air-breathing combustion chamber, while the rocket engine comprises a non-air-breathing annular combustion chamber and at least one turbo pump for feeding it with propellant. The combustion chamber of the rocket engine is disposed inside a central body towards the rear thereof and is situated at the downstream end of the air-breathing combustion chamber, the rocket engine combustion chamber being itself delimited by a streamlined portion extending the central body and constituting a spike which penetrates into the throat of the external nozzle while ensuring aerodynamic continuity for the stream of combustion gases leaving the air-breathing combustion chamber.

Abstract: A stationary centerbody includes vaulted areas displaced substantially by 180.degree. axially as well as along the circumference with respect to one another on its largest relative diameter. The vaulted areas, while enclosing a shut-off surface which is diagonal on the circumference side with respect to the end of an engine shroud forming a casing lip, encloses an axially spaced circumferential surface which is shut off in the manner of an annular slide valve from the end of the shroud.

Abstract: An integrated turboramjet engine has an axially slidable inlet cone, a rigid guide ring arranged concentrically outside the inlet cone, and a lock ring which can be slid axially against the inlet cone, whereby the air volume flowing into the air inlet can be introduced alternately into a turbo-inlet duct or a ram-air inlet duct in a low-loss and flow-promoting manner.

Abstract: A combined turborocket and ramjet propulsion unit comprises an air intake duct, and an air compressor with two counter-rotating rotor stages disposed downstream of the duct in an annular path formed between an outer casing and a central body which is maintained in position relative to the casing by structural arms. The rotor stages of the compressor are driven by a power turbine driven by combustion gases from a gas generator which burns liquid propellants. The power turbine includes two interleaved counter-rotating modules arraanged within the central body in a position longitudinally between the two compressor stages which they drive, and the gas generator is disposed axially within the central body to the rear of the turbine modules, the generator supplying the combustion gases to the turbine in an upstream direction through a central ejection chamber and an annular reversing chamber which redirects the gases rearwards through the turbine.

Abstract: A turboramjet engine is provided having a ram air duct which is constructed between an outer duct wall and a housing-type shrouding of a basic turbo-engine, and having shut-off devices by means of which the ram air duct, when the basic engine is switched on and the ramjet engine is switched off, is to be shut-off on the air supply side, with the simultaneous release of intake air into the basic engine. In this case, among other things, the shrouding is to form an air shaft which is widened in a polygonal manner in the direction of the inlet-side end of the ram air duct and is to have breakthroughs there, the shut-off devices being flaps which are pivotally arranged at the breakthroughs and which, when the ramjet engine is switched on, expose the breakthroughs and in the process form a shut-off body of the air shaft which is symmetrically acutely folded together with respect to the ram air flow.

Abstract: An aerospace propulsor is powered by liquid hydrogen. The hydrogen is passed through a turbine which drives a liquid oxygen pump. During non-atmospheric flight, the liquid oxygen is directed to the propulsor's combustion chamber where it is mixed with the liquid hydrogen and combustion takes place. However during atmospheric flight, the liquid oxygen is pumped to drive a turbine which powers a fan adapted to provide additional pressurization of air entering the engine for combustion.

Abstract: First and second axially adjustable rings, each ring having a front and rear end, function together in a telescopic manner for blocking and exposing an annular inflow surface in a combined turboramjet engine. The annular inflow surface is blocked by the device when a basic engine and fan are inoperative to direct ram air through a ring duct to a ramjet combustion chamber. When in an inoperative position, the first and second rings are moved above one another into an intermediate housing with their front ends and a flow divider forming a lug edge and, when in a blocking position of the annular inflow surface, each ring is moved out of the intermediate housing to a different degree. The first ring is moved out farthest and has an end surface mating with the local oblique contour of the front wall.

Abstract: A cold-air bypass duct in a hypersonic engine having a turbojet engine and a ramjet engine in the same housing. A common air inlet and exhaust nozzle are provided for the turbojet and ramjet engines. The air inlet is designed to provide the necessary air for the ramjet engine at maximum air requirements. Less air is needed by the turbojet engine when it is in operation. The excess air entering the inlet during turbojet engine operation enters a bypass duct normally used for ramjet operation. The air entering the bypass duct exits the housing behind the ram burner and turbojet nozzle throat. When the ramjet operates, a vent in the bypass duct opens to direct a portion of the air into the ram burner. When the ramjet engine is operating at maximum air requirements, all the air entering the bypass duct enters the ram burner. The bypass duct is a single duct with a vent and an associated barrier that are openable to multiple positions.

Abstract: An aircraft converting aerodynamic compression ram thermal stream energy into thrust power generation comprising:(a) an aerodynamic compression ram thermal stream generating multiple vane diffuser as the air-inlet to a pressure plenum;(b) a ram thermal-pressure stream induction double throttle duct consisting of a main ramflow inducing nozzle and a fuel injection ramflow inducing nozzle wihch coalesce to form the air-oulet of a pressure plenum-engine pod;(c) a compressed air shooting annular slotted ignition chamber downstream of the fuel injecting ramflow inducing nozzle with flame bed-walled combustion chamber forming a ramjet engine;(d) a ram thermal-pressure stream induction annular slotted thrust nozzle extending rearward from the combustion chamber of the ramjets and exit nozzle of a turbojet engine thereby creating a tailpipe for the turbo-ram induction jet engine; and(e) an aerodynamic compression ram thermal stream sink double-walled shockcone airframe, the double walled shockcone housing the ram the

Abstract: A compact, multimode propulsion system for aerial vehicles having multiple thermal cycles for selective independent or integrated operation in multiple modes to enable the propelled vehicle to operate in a full range of velocities under a variety of atmospheric conditions and efficiencies, the system including a turbo-jet unit with an air compressor, a gas turbine and an intermediate fuel injected combustor in association with an auxiliary combustor unit that includes a supplemental fuel and selective oxidizer supply, and a conversion system for selectively converting the auxiliary combustion unit into a secondary burner mode for the turbo-jet unit, an independent rocket mode operable with or without the turbo-jet unit or in independent ram-jet mode operable with or without the turbo-jet unit, the turbo-jet unit and auxiliary combustor system having separate exhaust systems for directional mobility of the vehicle to which the propulsion system is applied.

Abstract: Composite changeover-type reaction power plant for propelling aircraft and spacecraft operating in subsonsic, supersonic and hypersonic regions, embracing a turbofan bypass engine with front fan, and a ramjet engine with a cycle corresponding to that of the fan, where the fan has individually independently supported, operationally counterrotating statorless rotors and variable fan blades which are feathered and immobilized during ramjet operation for low drag during said ramjet operation.

Abstract: An improved power plant incorporating a gas turbine and a magnetohydrodynamic (MHD) generating means is provided wherein a plurality of coolant conducting fins are disposed in the diffuser region of the duct between the MHD generator and the turbine. The cooling fins are provided with means to supply compressed coolant to the diffuser duct for mixing with the combustion gases to provide optimum gas pressure, temperature and velocity for driving the turbine.

Abstract: A turbofan-ramjet engine is provided with optional duct burning in the turbofan mode. A core engine is provided with a first annular passage therearound for directing airflow around the inlet of the core engine. A second annular passage having a fan directs air flow into the inlet of the core engine and said first annular passage. A third annular passage extends around the first and second annular passage with the first and third annular passage each directing their flow into a common annular combustion chamber. Variable vanes are provided in the first and third annular passages to provide for blocking off the passage or acting as swirl vanes depending on mode of operation. Blocking vanes are also provided in the annular exhaust passage of the core engine. Coannular nozzles are provided with one located at the exhaust of the core engine and an outer one located aft of the common annular combustion chamber.

Abstract: A valve for controlling the fluid flow between two generally coannular ducts. The valve is provided with a pair of back-to-back, mutually facing flow mixers of the daisy variety, each associated with a section of the coannular ducts. Each mixer transists the two coannular duct passages into a plurality of circumferentially separated chute passages communicating alternatively with the inner and outer ducts, the mixers being indexed with respect to each other by half a chute width. A stage of radially extending, pivotable vanes cooperates with the chutes of both mixers and allows both the inner and outer upstream duct flows to be directed to either the inner or outer downstream duct, depending upon the orientation of the vanes. Alternatively, either or both of the fluid streams can be blocked off entirely. Complete flexibility for flow inversion is obtained without turning or indexing the flow through the vanes by more than half a chute width, thereby maintaining high efficiency.

Abstract: A variable ratio bypass turbojet engine comprising a front fan core engine and a forward fan connected thereto by an inverter valve arrangement which comprises a fixed forward duct system and a fixed after duct system selectively shaped and joined by flapper valve mechanism. The flapper valve mechanism may be actuated so that the bypass ratio of the engine can be varied to produce a low bypass ratio engine during cruise operation mode wherein the front fan air is conducted to the core engine fan and the auxiliary inlet air is discharged overboard, or to a high bypass ratio engine for take-off condition mode wherein the front fan air is discharged to atmosphere through a thrust nozzle and the auxiliary inlet air is directed to the core engine fan.

Abstract: Air breathing gas turbine engine design concepts, and methods and apparatus for providing variable engine cycles are disclosed. The preferred engine embodiments involve two flow control valve systems for internal management of air, each of which have the capacity to cross-over or invert adjacent annular flow patterns, and at least one of which is capable of switching from inverted to straight through flow.