Abstract: A Brayton cycle engine including a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines a gas flowpath of the engine. An inner wall assembly is extended from the longitudinal wall into the gas flowpath. The inner wall assembly defines a detonation combustion region in the gas flowpath upstream of the inner wall assembly.

Abstract: The invention relates to a method for manufacturing a fibre-reinforced pressure vessel having fibre-reinforced polar caps as well as a corresponding pressure vessel having these polar caps. Therein, the method comprises the steps of applying fibre composite material onto a provided winding body having the shape of the polar caps at at least one of the ends, using a winding process; of intermediately curing the fibre composite material for dimensional stabilisation, said fibre composite material, however, subsequently still remaining chemically active for later cross-linking; of severing the fibre composite material for producing a polar cap reinforcing layer which is detached from the winding body and placed onto a liner underlay of the pressure vessel. Subsequently, the polar cap reinforcing layer is cross-linked with the fibre composite material of the pressure vessel for producing the pressure vessel reinforcing layer.

Abstract: A stator temperature control system for a gas turbine engine is provided. The stator temperature control system includes a casing circumferentially surrounding a stator assembly, the casing having a top portion and a bottom portion; an air source having an inlet and an outlet; and a supply line in fluid communication with the outlet of the air source and the bottom portion of the casing, wherein the bottom portion of the casing receives a flow of air from the air source via the supply line to increase a temperature of the bottom portion of the casing.

Abstract: A gas turbine engine, in which a compressed gas from a compressor of an axial-flow type is burned in a combustor and an obtained combustion gas drives a turbine, includes: a diffuser of an annular shape connected to an outlet of the compressor, the diffuser including a diffuser inner tube and a diffuser outer tube that are tubular members disposed concentrically with each other; and a plurality of partition members that are disposed in a diffuser flow path, which is an annular space formed between the diffuser inner tube and the diffuser outer tube, and divide the diffuser flow path in a circumferential direction.

Abstract: A gas turbine engine has a fan inlet and a fan configured to deliver air to an exhaust nozzle. A core gas turbine engine. including in serial order extending further into the engine, a core turbine section, a combustor and a core compressor section. A core engine inlet duct is spaced from the fan inlet. A method is also described.

Abstract: This disclosure generally relates to lightweight thermionic microengines for aerial vehicles. The aerial vehicles include a propulsion system. The propulsion system includes a combustor. The propulsion system further includes a thermionic generator that receives heat from the combustor and generates electricity. The propulsion system further includes one or more propulsion motors that receive the electricity generated by the thermionic generator. The propulsion motors may provide power to one or more propellers to generate lift and thrust for a UAV.

Abstract: Plenums for bifurcated ducts and bifurcated ducts are provided for stabilizing flow therethrough. The plenum comprises an outer cylindrical body intersected by a pair of exhaust duct stubs that are configured to be coupled to a corresponding pair of exhaust ducts and an inner body. The outer cylindrical body includes an axial rear end portion. The inner body is disposed in the axial rear end portion and increases in diameter in the aft direction. The inner body comprises one of a generally axi-symmetrical inner body or a non-axi-symmetrical inner body. The bifurcated duct comprises the plenum and the pair of exhaust ducts. Exhaust systems are also provided.

Abstract: The invention relates to a mixing arrangement for mixing a fuel with a stream of oxygen containing gas flowing along an axis in an axial channel, especially in the second combustor of a gas turbine with sequential combustion. The mixing is improved and the mixing length reduced by said mixing arrangement comprising an injector with at least one injector ring, which is passed by said stream of gas inside and outside.

Abstract: A gas generator for a reverse core engine propulsion system has a variable cycle intake for the gas generator, which variable cycle intake includes a duct system. The duct system is configured for being selectively disposed in a first position and a second position, wherein free stream air is fed to the gas generator when in the first position, and fan stream air is fed to the gas generator when in the second position.

Abstract: A separate propulsion unit incorporating a free turbine and a fan receives gases from a plurality of core engines. The core engines each include a compressor, a turbine and a combustion section. The core engines in combination pass gases across the free turbine. A method is also disclosed.

Abstract: A gas turbine engine comprises a high spool, a low spool and an intermediate spool. The high spool comprises a high pressure turbine coupled to a high pressure compressor. The intermediate spool comprises an intermediate pressure turbine coupled to a ducted fan. The low spool comprises a low pressure turbine coupled to an open-rotor propeller. A variable area turbine section positioned between the intermediate pressure turbine and the low pressure turbine variable turbine section is configured to vary an expansion ratio across the intermediate pressure turbine to control rotational speeds of the low spool and the intermediate spool.

Abstract: A noise reducing device in a jet engine that has a cylindrical casing, a cylindrical partition wall that is inserted in the casing while protruding partially from a trailing edge of the casing, and a compressor that compresses air that is taken into the cylindrical partition wall, with the inside of the cylindrical partition wall serving as a duct in which a core stream of high-speed air flows, and the space between the cylindrical partition wall and the casing serving as a duct in which a bypass stream of low-speed air flows, and being connected with a wing of an airplane by a pylon that has a projection portion that extends beyond the casing to the downstream of the core stream and the bypass stream, the noise reducing device includes a nozzle, disposed at the pylon on the downstream of the core stream, that injects a fluid toward a noise generation source that is produced from the mutual approach of an ambient air stream that is produced outside of the bypass stream and the core stream.

Abstract: The present invention discloses a dual-vortical-flow hybrid rocket engine, including a main body and a nozzle communicating with an end of the main body. The main body includes a plurality of disk-like combustion chambers arranged longitudinally, and a central combustion chamber formed along the axial portion and communicating the disk-like combustion chambers. Each of the disk-like combustion chambers is provided with a plurality of oxidizer injection nozzles at its inner circumference surface. Inside the disk-like combustion chambers, the oxidizer is injected in nearly the tangent directions of the circumference, and the injection directions are opposite for the neighboring disk-like combustion chambers, which creates vortical flows with opposite rotating directions so as to increase the total residence time of the combustion reactions of the oxidizer and the solid-state fuel in the disk-like combustion chambers of the present invention.

Abstract: A high efficiency power plant for a UAV with a high pressure ratio gas turbine engine used for low power operation such as loiter speed and a low pressure ratio gas turbine engine used for high power operation. A power turbine receives hot gas flows from the two engines to drive an output shaft. At low power operation, only the high pressure ratio engine is operated. At high power operation, both engines are operated where the exhaust from the high pressure ratio engine is used to drive a turbine of the low pressure ratio engine. A compressor of the low pressure ratio engine supplies compressed air to a combustor that produces a hot gas stream that is passed through the power turbine.

Abstract: Specific impulse and rocket engine efficiency can be improved by injecting reactants, e.g., a propellant combination or a monopropellant and a catalyst, into a plasma flow of a rocket engine. In some aspects, a catalyst or a propellant is carried by plasma formed by passing a flow of a feed gas through an electrical arc. In some aspects, reactants are combusted in supersonic plasma flow to generate combustion ionization in the plasma flow.

Abstract: A supplemental thrust system for an airplane comprising a submerged inlet duct, a diffuser section in communication with the inlet duct, and a duct outlet in communication with the diffuser section, an outlet nozzle in communication with the duct outlet, a turbocharger unit comprising one or more turbocharges in communication with each other to provide compressed air to the engine and the cabin, a heat exchanger unit comprising a plurality of heat exchangers in the diffuser section to provide cooling liquid to cool pressurized air in the cabin, engine intake air and the engine jacket, where each turbocharger is coupled to the intake manifold of an engine to increase the power of the engine by introducing compressed air into the manifold.

Abstract: A gas turbine engine includes a propulsion unit mounted to rotate about a first axis, and a core engine mounted to rotate about a second axis, and wherein the first and second axes are non-parallel. A gas turbine engine includes a propulsion unit driven by a free turbine which is adjacent to the propulsion unit and an associated fan, and having a gas generator core engine including a compressor, combustor and turbine section. A method is also disclosed.

Abstract: An apparatus includes a slotted multi-nozzle grid with a plate having multiple elongated slotlettes through the plate. Each of at least some of the slotlettes has a convergent input, a divergent output, and a narrower throat portion separating the convergent input and the divergent output. At least some of the slotlettes are arranged in multiple rows. The plate further includes multiple cooling channels through the plate. At least some of the cooling channels are located between the rows of slotlettes. Each cooling channel is configured to transport coolant through the plate in order to cool the plate, such as to cool the plate as hot combustion gases pass through the plate. Each of at least some of the rows may include at least two slotlettes, and two adjacent slotlettes in one row may be separated by a structural ligament (which may have a tear-drop cross-sectional shape).

Abstract: An air vehicle propulsion system incorporates an engine core with a power shaft to drive an outer blade row. The power shaft extends through and is supported by a counter rotation transmission unit which drives an inner blade row in counter rotational motion to the outer blade row. The counter rotation transmission unit exchanges power from the engine core with the shaft. An actuator engages the shaft for translation from a first retracted position to a second extended position.

Abstract: A rocket multi-nozzle grid assembly includes an insulator grid, a plurality of refractory nozzle fittings and a grid support plate. The grid support plate braces the plurality of refractory nozzle fittings and the insulator grid. The insulator grid includes insulator orifices, each of the nozzle fittings includes a fitting orifice and the grid support plate includes a plurality of plate orifices. The rocket multi-nozzle grid assembly includes a plurality nozzles. Each nozzle includes a plate orifice aligned with one insulator orifice and one fitting orifice. Exhaust gases are directed through the plurality of nozzles. The multi-nozzle grid assembly substantially maintains its surface area throughout operation of a rocket motor.

Abstract: A liquid propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, and a plurality of holder plates arranged inside the tank body, around an axis L of the tank body. The holder plates cause the liquid propellant A to adhere to them by surface tension, thereby establishing, inside the tank body, a liquid propellant holding area LA in which the liquid propellant A is held and a gas accumulation area GA in which vapor produced by evaporation of part of the liquid propellant A accumulates. The tank body has a propellant inlet open into the liquid propellant holding area LA and a gas outlet open into the gas accumulation area GA.

Abstract: Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.

Abstract: A rocket multi-nozzle grid assembly includes an insulator grid, a plurality of refractory nozzle fittings and a grid support plate. The grid support plate braces the plurality of refractory nozzle fittings and the insulator grid. The insulator grid includes insulator orifices, each of the nozzle fittings includes a fitting orifice and the grid support plate includes a plurality of plate orifices. The rocket multi-nozzle grid assembly includes a plurality nozzles. Each nozzle includes a plate orifice aligned with one insulator orifice and one fitting orifice. Exhaust gases are directed through the plurality of nozzles. The multi-nozzle grid assembly substantially maintains its surface area throughout operation of a rocket motor. In one example, the surface area is maintained through ablation of the insulator grid. In another example, the surface area is maintained through cooling of the plurality of nozzle fittings with the ablated fragments passing through the nozzles.

Abstract: A joint including a first component and a second component. The second component including a shape memory material component. The first component has a half dovetail shaped slot and the shape memory material second component has a half dovetail shaped attachment feature disposed in the half dovetail shaped slot in the first component.

Abstract: A core engine of a variable cycle gas turbine engine includes a low pressure spool for generating streams of bypass air and pressurized air, and a high pressure spool for further pressurizing the stream of pressurized air to generate streams of combustion air and supercharged auxiliary air. A peripheral case surrounds the engine case to form a peripheral duct. An auxiliary combustor and propulsor are positioned within the peripheral duct. A bleed duct extends from the high pressure spool to the auxiliary combustor. Variable ductwork directs airflow through the bleed duct and peripheral duct in two modes. A first mode comprises directing the stream of auxiliary air to the auxiliary combustor, and directing stream of inlet air through the peripheral duct. A second mode comprises directing the stream of auxiliary air into the stream of bypass air, and preventing inlet air from entering the peripheral duct.

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: A gas turbine propulsion system comprises a turbofan engine, a peripheral duct, an annular frame, an auxiliary turbine and an auxiliary fan. The turbofan engine is configured to produce bypass air and combustion air. The bypass air flows through a bypass duct and the combustion air flows through a core engine. The peripheral duct surrounds the turbofan engine and is configured to selectively receive peripheral inlet air. The annular frame is disposed aft of the bypass duct and the peripheral duct, and is rotatable to alternately guide the bypass air out the bypass duct or the peripheral duct. The auxiliary turbine is connected to an aft end of the core engine and is configured to receive the combustion air. The auxiliary fan is connected to the auxiliary turbine and is configured to receive airflow from the peripheral duct.

Abstract: A spindle device driven by driving fluid, having accurate rotation and superior quietness. The device includes a stator having at least one inlet for introducing the driving fluid, a rotor having at least one flange arranged outside of the stator and provided with nozzles for jetting the driving fluid, and a static-pressure fluid bearing for rotatably supporting the rotor with respect to the stator. The stator has a first inner path for conducting the driving fluid introduced from the inlet to the rotor and the rotor has a second inner path communicating with the first inner path of the stator for conducting the driving fluid to the nozzles. The driving fluid introduced from the inlet of the stator is jetted from the nozzles of the rotor through the first and second inner paths to rotate the rotor.

Abstract: Systems and methods of controlling solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and propellant gas flow rate, rely on pulse width modulation of a control valve duty cycle. A control valve that is movable between a closed position and a full-open position is disposed downstream of, and in fluid communication with, a solid propellant gas generator. The solid propellant in the solid propellant gas generator is ignited, to thereby generate propellant gas. The control valve is moved between the closed position and the full-open position at an operating frequency and with a valve duty cycle. The valve duty cycle is the ratio of a time the control valve is in the full-open position to a time it takes the valve to complete one movement cycle at the operating frequency. The valve duty cycle is controlled to attain a desired solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and/or propellant gas flow rate.

Abstract: The present invention provides a mixer assembly including a mixer and a plurality of scoops attached around a circumference of the mixer element via a combination of sliding brackets and fixed brackets. The sliding brackets provide attachment and allow relative movement between the plurality of scoops and the mixer element while the fixed brackets provide only attachment between the mixer element and the plurality of scoops. The first sliding bracket includes a first flange that is fixedly mounted to the mixer element and a second flange that is biased against a surface of the scoop. A second sliding bracket is fixedly mounted to the mixer element and adjustably mounted to the scoop. The second sliding bracket includes an attachment hole and the scoop includes an adjustment slot. A fastener assembly including a spacer is received within the adjustment hole and the adjustment slot to allow relative movement between the mixer element and the scoop.

Abstract: A gas turbine engine includes a turbomachinery core operable to generating a first flow of pressurized combustion gases, the core having an exit plane; a fan disposed upstream of the core adapted to extract energy from the core and generate a first flow of pressurized air; a bypass duct surrounding the core which receives a portion of the flow of pressurized air from the fan; a duct burner disposed in the bypass duct, upstream of the exit plane, for receiving the first flow of pressurized air and generating a second flow of pressurized combustion gases; and an exhaust duct disposed downstream of the core and operable to receive and the first and second flows of pressurized combustion gases and to discharge the combined flows downstream.

Abstract: A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel.

Abstract: A spindle device driven by driving fluid, having accurate rotation and superior quietness. The device includes a stator having at least one inlet for introducing the driving fluid, a rotor having at least one flange arranged outside of the stator and provided with nozzles for jetting the driving fluid, and a static-pressure fluid bearing for rotatably supporting the rotor with respect to the stator. The stator has a first inner path for conducting the driving fluid introduced from the inlet to the rotor and the rotor has a second inner path communicating with the first inner path of the stator for conducting the driving fluid to the nozzles. The driving fluid introduced from the inlet of the stator is jetted from the nozzles of the rotor through the first and second inner paths to rotate the rotor.

Abstract: A gas turbine powerplant for a V/STOL aircraft in which a vectorable exhaust nozzle at a forward part (eg. nose) of the aircraft selectably receives a supply of cold fan air through a duct leading from the fan through a passageway (28) coaxial with the engine axis and located within the hub of the fan (12), with the aid of controllable flaps and diverters (26, 30).

Abstract: Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.

Abstract: An exhaust nozzle assembly for a gas turbine engine, the assembly comprises a main axis, an inner nozzle, an outer nozzle, a translatable centre-body and an actuator for translating the centre-body between a forward position and a rearward position. The centre-body is disposed radially inwardly of the inner nozzle thereby partly defining an inner duct for a core engine gas flow, an outer duct for a bypass gas flow is defined by the inner nozzle and the outer nozzle. The outer nozzle extends downstream of the inner nozzle and with the centre-body defines a final mixing duct having a final exhaust exit area. When the centre-body is in the forward position the final exhaust exit area is at a maximum area, and when the centre-body is in a rearward position the final exhaust exit area is at a minimum area.

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: A center plug is configured for variable area use in an exhaust nozzle. A forward portion of the plug includes an inlet having a closing flap therein, and an aft portion of the plug includes an outlet. The flap is selectively open during aircraft takeoff operation and up to a preset climb altitude for bypassing a portion of core exhaust flow through the plug, with the flap being closed from preset altitude and cruise operation. Opening the flap correspondingly increases the available flow area for discharging the exhaust flow for reducing velocity thereof, and attenuating noise.

Abstract: A variable cycle propulsion system for a supersonic airplane, the system comprising at least one engine having means for producing exhaust gas and a gas exhaust nozzle for generating thrust for supersonic flight speeds, and at least one separate auxiliary propulsion assembly dissociated from said engine, having no gas generator, and capable of generating thrust for takeoff, landing, and subsonic flight speeds. The system further comprises gas flow tapping means movable between a position in which they tap off at least a fraction of the exhaust gas produced by said engine and feed it to said propulsion assembly to enable it to generate thrust for takeoff, landing, and subsonic cruising flight, and a position in which the gas produced by the engine is directed solely to the engine nozzle for supersonic cruising flight.

Abstract: A microjet based control system helps control the feedback loop that is created by an impinging jet of a STOVL aircraft. A plurality of microjets circumferentially encompass the outer periphery of the impinging jet and each issues a jet flow toward the jet flow created by the impinging jet. A control system is connected to the microjets for adaptively controlling the microjets based on various operating parameters collected by various input devices.

Abstract: A turbo shaft engine for amplifying an air stream flow rate includes a turbine fan assembly and gas generator. The gas generator includes a primary air duct defining intake and outlet ports. A combustion chamber is connected to the primary air duct for igniting an admixture of fuel and a portion of the intake flow to form an energized motive flow. The motive flow is discharged from the combustion chamber back into the primary air duct over a Coanda-profiled guide member so as to amplify the flow rate of incoming intake flow by momentum transfer. A portion of the motive flow is returned directly to the fan assembly for amplifying incoming intake flow. The remaining motive flow is again combusted and used to rotate turbine blades. A resonance chamber with volume adjustment is included for tuning a pulse of intake flow into the primary combustion chamber.

Abstract: A micro air vehicle is propelled by a bipropellant micro engine supplied with hydrogen peroxide and a hydrocarbon fuel. The hydrogen peroxide is decomposed to produce steam and oxygen at high temperature and the hydrocarbon fuel is burnt within a combustion chamber with oxygen produced from such decomposition. Products of such decomposition and combustion are exited through a nozzle to produce thrust. In a preferred embodiment there is also a ducted turbofan located downstream of the nozzle exit.

Abstract: A gas turbine engine turbine assembly includes a high pressure spool having a high pressure turbine drivingly connected to a high pressure compressor by a high pressure shaft which is rotatable about an engine centerline. A low pressure turbine has counter rotatable low pressure inner and outer shaft turbines drivingly connected to coaxial low pressure inner and outer shafts respectively which are at least in part rotatably disposed co-axial with and radially inwardly of the high pressure spool. The low pressure inner shaft turbine is drivingly connected to a forward fan blade row by the low pressure inner shaft. The low pressure outer shaft turbine is drivingly connected to a aft fan blade row by the low pressure outer shaft. A single direction of rotation booster is drivenly connected to the low pressure outer shaft and axially located aft and downstream of the aft fan blade row.

Abstract: A turbo shaft engine for amplifying an air stream flow rate includes a plurality of blades positioned within a turbine housing for rotation by an intake air flow. A gas generator having a primary air duct defines intake and outlet ports, the intake port receiving the intake flow from the housing. A combustion chamber is connected to the primary air duct for igniting an admixture of fuel and a portion of the intake flow to form an energized motive flow. The motive flow is discharged from the combustion chamber back into the air intake of the primary air duct so as to amplify the flow rate of incoming intake flow by momentum transfer. A portion of the motive flow is returned directly to the housing inlet port for amplifying incoming intake flow. The remaining motive flow is again combusted and used to rotate the turbine blades.

Abstract: A rocket motor is disclosed that is suitable for use with an ejection seat. The rocket motor includes a combustion chamber, and has a plurality of spaced apart outlets arranged in fluid communication with the combustion chamber, to allow the passage of exhaust gases upon ignition of a propellant within the combustion chamber. Each exhaust outlet is arranged to direct the exhaust gases along respective fixed lines of thrust. The arrangement is such that the resultant thrust generated by the rocket motor acts along the line non-parallel with at least one of the fixed lines of thrust so as to direct the rocket motor along the curved trajectory. The secondary outlets may be smaller than the primary outlets and are selectively closable by valves.

Abstract: This jet engine comprises a complete low bypass turbofan unit 1 including secondary fan unit 2 and engine core, ahead of which sits a larger diameter main fan unit 3 driven from the front of the secondary fan via a long drive shaft 4. A center portion of the main fan feeds the secondary fan via long central duct 5. The outer portion of the main fan is ducted into a split exhaust plenum chamber 6. This chamber has thrust vectoring exhaust mechanisms 7 immediately attached to each side, and then each side then feeds a thrust splitter 10, and then feeds to variable area outer ducting 8, using double-hinged inner and outer doors 9. Both outer ducts rejoin the central duct downstream in a second plenum chamber 11 just ahead of the secondary fan. Between this chamber and secondary fan is an intercooler 12.

Abstract: A propulsion pod for a twin tube, rotary inlet valve, pulse detonation engine includes a shared, two-dimensional, low aspect ratio wedge nozzle in which each side of the wedge is transitioned into the discharge end of the detonation tubes of the engine. The nozzle design results in a quasi-separate exhaust flow path with two separate nozzle throat areas, one for each detonation tube. Actuation of the center wedge of the nozzle provides pitch vectoring of the exhaust. Actuation of flaps integrated into the side walls of the nozzle provides yaw vectoring. The detonation tubes are cooled by flowing bypass air over cooling means coupled to the detonation tubes.

Abstract: A modular rocket motor having a cylindrical housing with one end having an inwardly curved rim and the other end having internal threads. A heat-resistant plastic liner and a threaded nozzle cap are slidingly insertable with the cylindrical housing, with the threaded nozzle cap mating with the internal threads of the cylindrical housing. The motor is threaded into a rocket frame. A nozzle cap with liner attached thereto for a modular rocket motor that is colour coded to indicate the performance classification of the rocket motor.

Abstract: A method and device for reducing the amount of noise generated by a supersonic jet engine is provided. The method creates an envelope of air around the supersonic jet exhaust. The temperature and velocity of this envelope is controlled to eliminate or reduce the formation of noise making Mach waves.

Abstract: A method and device for reducing the amount of noise generated by a supersonic jet engine is provided. The method creates an envelope of air around the supersonic jet exhaust. The temperature and velocity of this envelope is controlled to eliminate or reduce the formation of noise making Mach waves.