Abstract: The mainstream jet engine now is the turbo-jet engine. Here in this invention the turbine is replaced by a piston motor. This will increase fuel efficiency and separating the turning parts from the burning parts will benefit maintenance effort. Unlike traditional turbo-jet engine to use turbines to drive fans and compressors, this invention uses a rotary motor driven by high pressure fluid generated by separate working chambers piston engine to drive fans and compressors Sealing is a major issue in engine design. Without proper sealing between moving parts, the working media will loose pressure and the power output will be low. So tradition rotary engine has sealing problems due to its complex movement of pistons. Here a linear movement piston is used when the working media is gas and a rotary piston when the working media is fluid, thus solving the problems with sealing, cooling and lubricating difficulties. This invention uses some techniques developed in the previous inventions by me, application No.

Abstract: An aircraft engine is provided with at least one pulse detonation device connected to an engine exhaust nozzle portion with a plurality of supersonic injectors. The flow from the pulse detonation device is passed through the supersonic injectors into the primary engine flow path. The injector flow is injected at a velocity such that the injected flow penetrates into the primary flow path. This injected flow creates a virtual obstacle for the primary flow, and their interaction creates a virtual surface or an aero-lobe, thus controlling the nozzle exit area to reduce engine noise.

Abstract: A continuous detonation wave engine (CDWE) is disclosed. An example embodiment includes a generally annular combustion chamber dimensioned to allow a fuel mixture to detonate, a mixing chamber, and a fuel mixture channel that provides for fluid communication between the mixing chamber and the combustion chamber. At least part of the fuel mixture channel features a quenching structure dimensioned to substantially prevent detonation from spreading from the combustion chamber via the fuel mixture channel to the mixing chamber. Other embodiments are described and claimed.

Abstract: A detonation chamber for a pulse detonation combustor including: a plurality of dimples disposed on at least a portion of an inner surface of the detonation chamber wherein the plurality of dimples enhance a turbulence of a fluid flow through the detonation chamber.

Abstract: A supersonic propulsion system is provided. The supersonic propulsion system includes a plurality of systems for efficiently creating cyclic detonations and at least one rocket booster device. Each of the systems include at least a first initiator chamber configured to generate an initial wave, at least one main chamber coupled to the first initiator chamber. The main chamber is configured to generate a main wave and to output products of supersonic combustion. The products are generated within the main chamber. The main chamber is configured to enable the main wave to travel upstream and downstream within the main chamber when the first initiator chamber is located outside the main chamber. The system further includes an initial connection section located between the first initiator chamber and the main chamber that enhances a combustion process via shock focusing and shock reflection.

Abstract: Pulse detonation/deflagration apparatus for providing enhanced pressure wave operating frequency and/or magnitude, and methods of increasing the frequency or the magnitude of deflagration to detonation waves, are provided. A pulse detonation/deflagration apparatus can include a main/outer pulse detonation/deflagration actuator/engine (PDA/E) with multiple smaller internal combustion chambers or tubes positioned inside the cavity of the main/outer PDA/E with each performing the function of individual PDA/Es. The output pressure waves created by the internal PDA/Es can be utilized for propulsion or for controlling large scale flows, where needed.

Abstract: A device and method for improving the performance of a pulse detonation engine. The device includes at least one of an exhaust structure and an ejector. The exhaust structure can be configured as a straight, converging or diverging nozzle device, and connected to the engine to control the flow of a primary fluid produced during a detonation reaction. The ejector is fluidly coupled to the engine, using the movement of the primary fluid to promote entrainment of a secondary fluid that can be mixed with the primary fluid. The secondary fluid can be used to increase the mass flow of the primary fluid to increase thrust, as well as be used to cool engine components. Device positioning, sizing, shaping and integration with other engine operating parameters, such as fill fraction, choice of fuel and equivalence ratio, can be used to improve engine performance. In addition to thrust augmentation and enhanced cooling, the disclosed device can be used for engine noise reduction.

Abstract: A pressure rise combustor is provided with fuel provided at intermittent periods. The fuel is pulsed at timings such that the phase lag between the addition of the fuel and a resultant pressure rise is minimized. The fuel is pulsed such that the unsteady addition of heat reinforces the amplitude of an unsteady pressure fluctuation.

Abstract: A turbine disk assembly including a rotatable cylindrical member rotatably coupled to a shaft and a plurality of turbine blades extend circumferentially outward from said cylindrical member. The turbine blades include at least two different geometrical shapes, a first of the geometrical shapes is configured to facilitate extracting power from a first pulsed detonation combustor product stream. A second of said geometrical shapes is configured to facilitate extracting power from a second pulsed detonation combustor product stream that is different from the first pulsed detonation combustor product stream.

Abstract: The present invention is a constant volume rocket motor that uses a non-detonating constant-volume, bipropellant combustion process in pulse-mode operation. Opening and closing of the combustion chamber exhaust outlet is controlled by an actuated reciprocating thrust valve (RTV). Fuel enters the combustion chamber at low pressure with the RTV closed. The valve opens after or during combustion when combustion chamber pressure is at or near maximum. The motor has applications in reaction control systems and attitude control systems in spacecraft.

Abstract: A constant volume combustor device includes, in one form, a detonative combustion. In one form the wave rotor of the constant volume combustor is supported by magnetic bearings. The constant volume combustor device includes a rotor having a number of fluid passageways that rotate about an axis. End plates having at least one inlet port and at least one outlet port are located on either end of the rotor. Relatively compressed air enters the rotor through the at least one inlet port, is burned with fuel in a pulsed combustion process, and exits at least one exit port. The pulsed combustion process can be a pulsed detonation combustion process or a pulsed deflagration combustion process.

Abstract: A swirler having cross-sectional area comparable to the area of a detonation chamber is placed upstream of the detonation chamber to enhance the fuel-air mixing. The swirler has a first region and a second region, each of which induces swirl in the flow through the swirler. Each region induces a different direction of swirl in the flow. The residual net swirl present in the flow downstream of the swirler is controlled by the relative properties of each region of the swirler. The swirler also provides high optical blockage to inhibit the upstream propagation of flow due to the detonation shockwave.

Abstract: The present invention provides a reactor utilizing high-voltage discharge for processing exhausted hydrogen gas emitted during membrane plating, etching, or washing of semiconductors, where higher than 95% of destruction and removal efficiency (DRE) of hydrogen gas is obtained.

Abstract: A rotary piston engine having a housing defining an un-obstructed circular toroidal chamber, a toroidal segment piston in said chamber, a detonation chamber having an outlet substantially tangential to the outer diameter of said toroidal chamber wherein said piston is driven in a continuous circular orbit by energy derived from pulse-detonation generated shock wave and pulse-jet gas flow.

Abstract: A pulsejet is provided that includes an inlet, an exhaust nozzle, and a combustion chamber between the inlet and the exhaust nozzle. The pulsejet additionally includes a flow-turning device positioned over an end of the inlet. The flow turning device forms an air flow pathway (AFP) having a substantially 180° turn. The 180° turn aligns a direction of exhaust exiting the combustion chamber through the inlet along a positive axial thrust line and substantially parallel with exhaust exiting the combustion chamber through the nozzle.

Abstract: A continuous detonation system, including: a rotatable member including a forward end, an aft end, a circumferential wall and a longitudinal centerline axis extending therethrough; an outer circumferential wall, wherein the rotatable member is positioned therein so that the circumferential wall of the rotatable member is spaced radially inwardly from the outer circumferential wall; at least one helical channel formed by a plurality of helical sidewalls extending between the circumferential wall of the rotatable member and the outer circumferential wall, each helical channel being open at the forward end and the aft end of the rotatable member so as to provide flow communication therethrough; an air supply for providing air to each helical channel; and, a fuel supply for providing fuel to each helical channel. In this way, a mixture of the fuel and air is continuously detonated within each helical channel in a manner such that combustion gases exit therefrom with an increased pressure and temperature.

Abstract: The invention provides a liquid fueled pulse detonation air breathing engine. The invention also provides an embodiment of an engine which has an axial flow compressor final stage configured for momentary closure of air flow at the final stage of the compressor followed by pulse detonation combustion that powers compressor, appliances, propeller, and produce thrust. Fuel enters the configured final stage of the axial flow air compressor to allow a mixture of air and fuel before entering the engine's pulse detonation combustion chamber, also referred to as detonation canister, having an inlet opening for receiving the fuel and air mixture charge and an open end down stream to discharge the resulting combustion products. Once the fuel and air mixture enters the detonation chamber there is closure for an instant of the compressor final stage's novel stators and blades before detonation of the fuel and air mixture.

Abstract: A pulse detonation combustor assembly contains at least one PDC tube, a mechanical air flow valve which directs an air flow into the PDC tube, where the mechanical air flow assembly changes a rate of the air flow into the PDC tube during a fill stage of the PDC tube. The assembly also contains a fuel flow control valve which directs fuel to the PDC tube and changes the rate of the fuel flow into PDC tube. By controlling the flow of the fuel and air into the PDC tube the equivalence ratio profile of the PDC tube can be tailored and controlled.

Abstract: A method of rotating a crank shaft and in internal detonation engine are provided. The internal detonation engine comprises a deflagration to detonation transition section. The deflagration to detonation transition section is connected to a main cylinder, which houses a piston. Inducing a detonation wave from the deflagration wave and passing the detonation wave through a fluid, gives rise to high pressure and temperature in a cylinder and pushes a piston towards bottom dead center. An internal detonation reciprocating engine may be a single cylinder and may be either a two or four stroke engine. A two-stroke internal detonation reciprocating engine is similar to a four-stroke internal detonation reciprocating engine but has different valve placements. Detonations produce a more thorough combustion of the fuel and may, thereby, yield reduced emissions of carbon monoxide as compared to internal combustion engines.

Abstract: A system of controlling airflow into a pulse detonation engine includes a rotary airflow controller valve receiving air from a high-speed inlet. An engine frame includes a plurality of detonation chambers. A sealing mechanism is between the rotary airflow controller valve and the engine frame. The sealing mechanism is associated with the engine frame and limits leakage of a gas from a first of the detonation chambers to a second of the detonation chambers.

Abstract: A pulse detonation engine comprises a primary air inlet; a primary air plenum located in fluid communication with the primary air inlet; a secondary air inlet; a secondary air plenum located in fluid communication with the secondary air inlet, wherein the secondary air plenum is substantially isolated from the primary air plenum; a pulse detonation combustor comprising a pulse detonation chamber, wherein the pulse detonation chamber is located downstream of and in fluid communication with the primary air plenum; a coaxial liner surrounding the pulse detonation combustor defining a cooling plenum, wherein the cooling plenum is in fluid communication with the secondary air plenum; an axial turbine assembly located downstream of and in fluid communication with the pulse detonation combustor and the cooling plenum; and a housing encasing the primary air plenum, the secondary air plenum, the pulse detonation combustor, the coaxial liner, and the axial turbine assembly.

Abstract: The present invention provides a pulse detonation insert that induces flow obstructions within the pulse detonation chambers, wherein the flow obstructions are operable to induce turbulence within a primary fluid flow passing over the obstructions. This turbulence may take the form of vortices that enhance the mixing of the oxidizer and fuel within the primary flow. Additionally, supports couple to the pulse detonation chamber walls and flow obstructions to hold the flow obstructions in place within the pulse detonation chamber. The combustion/Detonation of the mixed oxidizing fuel results in an increased velocity of the primary flow exiting the pulse detonation chamber and reduced the amount of unburnt fuel within the exhaust.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: In one embodiment, a combustion apparatus comprises: an oxidant inlet, a fuel inlet, a mixing zone configured to mix fuel and oxidant to form a mixture, a combustion zone in fluid communication with the mixing zone and configured to combust the mixture, and an outlet in fluid communication with the combustion zone. The mixing zone, the combustion zone, and/or the outlet comprises an electro-dynamic swirler configured to swirl ionized gas. The electro-dynamic swirler comprises a plurality of electrodes in electrical communication with a power source. In one embodiment, a method of creating thrust comprises: mixing a fuel and an oxidant to form a mixture, igniting the mixture to form an ignited mixture, combusting the ignited mixture to form a flame, and using the flame to create thrust, wherein the flame and/or the fuel and oxidant are electro-dynamically swirled.

Abstract: A pulsed detonation engine may include a flame tube with a lateral wall and a transverse base defining a combustion chamber and a supply device cyclically feeding said combustion chamber with a combustible charge. The transverse base of the flame tube may be movable for reciprocating in translation inside said flame tube in order to be able to occupy two boundary positions, a first position corresponding to the detonation phase of the combustible charge in the combustion chamber of said flame tube and a second position corresponding to the phase wherein the combustible charge is supplied to said combustion chamber.

Abstract: A pulse combustion device has a number of combustors with upstream bodies and downstream nozzles. Coupling conduits provide communication between the combustors. For each given combustor this includes a first communication between a first location upstream of the nozzle thereof and a first location along the nozzle of another. There is second communication between a second location upstream of the nozzle and a second communication between a second location upstream of the nozzle of a second other combustor and a second nozzle location along the nozzle of the given combustor.

Abstract: A linear acoustic pulsejet (LAP) including a body having a first side panel and an opposing substantially parallel second side panel. A plurality of substantially parallel intercostals are orthogonally connected to each of the first and second side panels to create a plurality of pulsejet cells within an interior of the body. The LAP additionally includes a linear inlet cap over a top of the body that forms an air flow pathway (AFP) between the linear inlet cap and an exterior of an inlet section of the body. The linear inlet cap forms the AFP to have a substantially 180° turn between an exterior of the body and an interior of the pulsejet cells.

Abstract: The present specification generally relates to improvements to combustors such as burners and engines. In one aspect, the specification presents an acoustically enhanced ejector system which can be used as part of an intake system for a combustor. In another aspect, the specification teaches the use of a combustor combustion chamber as an oscillator to magnify a harmonic frequency of a pulsating frequency of the combustor. In still other aspects, the specification presents a combustion chamber having an inlet with a plurality of tangentially spaced apertures, and an in-line intake system connected to the apertures.

Abstract: A pulse jet engine comprising a quarter wave gas resonator (1) which is arranged to cycle at an ultrasonic frequency, wherein the resonator (1) is closed, or substantially closed, to gas flow at the pressure antinode (4) thereof. The shape and dimensions of the quarter wave gas resonator (1) at least partially determine the ultrasonic frequency at which the resonator (1) cycles.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: An engine contains at least one pulse detonation combustor which is positioned upstream of a turbine section, a stage of which is a Curtiss type turbine stage. Following the initial Curtiss type turbine stage, is either of a Rateau type turbine stage or a high efficiency turbine stage, or a combination thereof.

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of the fan section, the booster compressor including a first compressor blade row and a second compressor blade row connected to a second drive shaft and interdigitated with the first compressor blade row; and, a pulse detonation system for powering the first and second drive shafts. The pulse detonation system powers only the second drive shaft during a first designated condition of the gas turbine engine and both the first drive shaft and the second drive shaft during a second designated condition of the gas turbine engine. The first and second drive shafts are powered independently of each other by the pulse detonation system.

Abstract: A steady-state detonation combustor and a steady-state detonation wave generating method, in which a stabilized detonation wave can be generated by generating a hypersonic and unburned premixed gas. An rich premixed gas whose gas fuel is rich is combusted in a rich premixed gas combustion chamber (11) to generate a fist high-temperature and high-pressure burned gas, while at the same time, a lean premixed gas whose oxygen is rich is combusted in a lean premixed gas combustion chamber (12) to generate a second high-temperature and high-pressure burned gas, and subsequently, after each high-temperature and high-pressure burned gas is accelerated to hypersonic speed and at the same time mixed together through an interpenetrating nozzle (40), a premixed gas obtained by the mixture and containing the gas fuel and the oxygen which are unreacted is impinged on a steady-state detonation stabilizer (60), so that a stabilized detonation wave is generated.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: Pressure and density of a gaseous mixture are increased in the process of introducing the gaseous mixture into the combustor of an air-breathing pulse detonation engine employing atmospheric oxygen as an oxidizer. The exit valve 20 able to be opened and closed is provided at the outlet of the combustor 15, an air cooler 12 is provided in the exit of the intake, and density is increased by exchange of heat of the air received at the intake with a coolant in the air cooler 12. Furthermore, by closing the exit valve 20 provided in the outlet of the combustor during the process of loading the gaseous mixture, transition to the detonation process is possible without expansion of the high-pressure high density air obtained by ram-compression at the intake.

Abstract: A propulsion system for creating a propulsive force has a combustion chamber, a pair of electrodes within the combustion chamber and a power supply attached to the electrodes to create a high voltage field within an initiation zone of the combustion chamber, and an injector for introducing a propellant, preferably in atomized form, into the high voltage field for creating the propulsive force. In a preferred embodiment of the present invention, the propellant is hydrogen peroxide. In another embodiment of the present invention, a second propellant is introduced into the combustion chamber for increasing the propulsive force.

Abstract: The present invention is a gas turbine engine system containing a compressor stage, a pulse detonation stage, and a turbine stage. During operation of the engine system, compressed flow from the compressor stage is directed to at least one off-axis pulse detonation tube in the pulse detonation stage. The off-axis pulse detonation tube detonates a mixture containing the compressed flow and a fuel to create a detonation wave, which is routed to an inlet portion of a reverse flow turbine. The exhaust flow from the reverse flow turbine is then directed away from the engine through ducting.

Abstract: A pressure wave apparatus utilizing the principles of pulsed detonation and wave rotor technologies. The apparatus includes inlet and outlet ports that interface with a plurality of fluid flow passageways on a rotor. A buffer gas is routed through some of the inlet and outlet ports and into and out of the plurality of fluid flow passageways. One of the inlet ports is a buffer gas inlet port that when placed in registry with a fluid flow passageway allows the flow of buffer gas into the respective passageway. Fuel is delivered into the buffer gas proximate the buffer gas inlet port so that only a portion of the buffer gas inlet port receives any fuel.

Abstract: The present invention is a pulse detonation combustion system, having a plurality of detonation initiation devices coupled to a main combustion chamber, where each of the detonation initiation devices is operating out-of-phase with each other. Each of the detonation initiation devices assists in the initiation of a detonation in the main combustion chamber, out-of-phase from each other such that the operational frequency of the pulse detonation combustion system is related to the number of detonation initiation devices multiplied by the operational frequency of a single detonation initiation device.

Abstract: A pulse detonation system for a gas turbine engine having a longitudinal centerline axis extending therethrough. The pulse detonation system includes a rotatable cylindrical member having a forward surface, an aft surface, and an outer circumferential surface, where at least one stage of circumferentially spaced detonation passages are disposed therethrough.

Abstract: A pulsed combustion device includes a support structure and a combustor carousel supported by the support structure and rotating relative thereto about an axis. The carousel has a number of combustion conduits in a circumferential array. Each conduit cyclically receives a charge and discharges combustion products of the charge.

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a drive shaft; a booster compressor positioned downstream of the fan section including a plurality of stages, where each stage includes a stationary compressor blade row and a rotating compressor blade row connected to the drive shaft and interdigitated with the stationary compressor blade row; and, a combustion system for producing pulses of gas having increased pressure and temperature of a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet.

Abstract: The invention is a turbojet propulsion system which includes a compressor section, a turbine section coupled by a shaft to the compressor section, a combustion section mounted between the compressor section and the turbine section, and an exhaust duct coupled to the aft end of said turbine section. A fuel delivery system is incorporated for supplying fuel to the combustion section. A laser assembly provides electromagnetic radiation to the combustion section. An electrical generator coupled to the turbine driven shaft powers the laser assembly.

Abstract: The disclosed device is directed toward a periodic combustion propulsion system. The periodic combustion propulsion system comprises at least one periodic combustion chamber configured to contain periodic combustion and at least one supersonic exhaust nozzle coupled to the at least one periodic combustion chamber. The supersonic exhaust nozzle is configured with a variable exhaust expansion ratio.

Abstract: An engine includes at least one pulse detonation chamber configured to receive and detonate a fuel and an oxidizer. The pulse detonation chamber has an outlet end and includes a porous liner adapted to fit within an inner surface of the pulse detonation chamber within a vicinity of the outlet end. The engine also includes a casing housing the pulse detonation chamber.

Abstract: Distributed initiation (e.g., multipoint or continuous) is utilized to obtain constant volume-like combustion performance in a pulse combustion device in the absence of detonation. A number of such devices may be utilized as turbine engine combustors.

Abstract: A method for operating a pulse detonation system. The method includes providing a pulse detonation chamber including a plurality of detonation tubes extending therein, and detonating a mixture of fuel and air within each detonation tube such that at least a first tube is detonated at a different time than at least a second detonation tube.

Abstract: A two-stage pulse detonation system includes a pre-combustor and a geometric resonator connected via a converging-diverging nozzle to the pre-combustor to create a high temperature and high pressure conditions in the resonator in order to create optimal conditions for detonation initiation. A mixture of a fuel and a gas is burned in the pre-combustor and is passed through the nozzle into the geometric resonator, where the burned mixture is detonated. The detonation propagates through the resonator exit nozzle thus generating thrust.

Abstract: The disclosed device is directed toward a periodic combustion propulsion system. The periodic combustion propulsion system comprises at least one periodic combustion chamber configured to contain periodic combustion and at least one supersonic exhaust nozzle coupled to the at least one periodic combustion chamber. The supersonic exhaust nozzle is configured with a variable exhaust expansion ratio.

Abstract: This invention pertains to apparatus and method for initiating detonation in a chamber of tubular or other shapes, which can be a combustor for a propulsion engine, such as a pulse detonation engine. This invention is characterized by detonation initiation at high pressure and temperature that is generated by imploding shocks induced by jets in different directions around the chamber. Further, the detonation initiation is achieved without energy depositing devices, such as electric spark plugs and lasers and without any fuel or other chemical additives.