Abstract: Increased efficiency and reliability is achieved in an engine with a combustion chamber structured as detonation resonator with an outlet to an exhaust nozzle. The resonator is formed as an aspherical reflector symmetrical with respect to the engine axis. The engine uses gaseous fuels and a gaseous oxidizer in a single stage combustion process. A pyrolyzing chamber for hydrocarbon fuel. Pyrolizing is achieved by contact of fuel flow with heated back side of the reflector. A mixture of fuel and oxidizer is supplied into the combustion chamber through an annular supersonic injection system. To initiate detonation, these engines may have a detonation initiator formed as a tube plugged at the distal end and open at the end inserted into the combustion chamber and located along the axis of the engine. Detonation products ejected through the exhaust nozzle create thrust that pushes the engine in the opposite direction.

Abstract: A pulse detonation combustion system includes: an inlet pipe; an intake cone disposed in the inlet pipe, having an end provided with a pneumatic valve, and including an atomizing air transfer tube, a fuel transfer tube, and a conical swirl nozzle connected to the atomizing air transfer tube and the fuel transfer tube; an atomizing air intake tube connected with the atomizing air transfer tube; a fuel supply tube connected with the fuel transfer tube; a pulse detonation combustion chamber located downstream of and communicated to the inlet pipe, and provided with a spark plug mounting seat for mounting a spark plug; a gas energy distribution adjustment device located downstream of and communicated to the pulse detonation combustion chamber; and a transition section located downstream of and communicated to the gas energy distribution adjustment device.

Abstract: A rotary detonation rocket engine generator system can include an axial drive shaft operably coupleable to an electrical generator. At least one support arm is radially coupled to the axial drive shaft and has corresponding rotary detonation rocket engines. An air-fuel mixing chamber receives ambient air and fuel to form an air-fuel mixture and deliver the air-fuel mixture to an annular combustion chamber. At least one pulse detonation combustion chamber is in fluid communication with the annular combustion chamber to receive an oxidizer and fuel to form an oxidizer-fuel mixture. The at least one pulse detonation combustion chamber creates a detonation wave that travels along the at least one pulse detonation chamber to the annular combustion chamber and ignites the air-fuel mixture as the detonation wave travels around the annular combustion chamber to generate thrust force that causes rotation of the axial drive shaft to drive the electrical generator.

Abstract: A system and method is disclosed for the start-up and control of pulsejet engines and this system includes an Electronic Fuel Injection (“EFI”) system that further includes one or more electrically controlled fuel injectors that can be selectively operated for start-up and control of such pulsejet engines. According to the system and method, the rate and/or pattern of fuel delivery to pulsejet engines can be varied not only by controlling the amount of time the fuel injectors are open versus closed to define a “duty cycle,” but also with the capability to selectively disable one or more fuel injectors in the programmed manner for start-up and control of such pulsejet engines.

Abstract: The present invention discloses an isolation section suppressing shock wave forward transmission structure for a wave rotor combustor and a wave rotor combustor, and belongs to the new concept field of unsteady combustion. The isolation section suppressing shock wave forward transmission structure for a wave rotor combustor includes a wave rotor and a gas inlet port, and the wave rotor is provided with several wave rotor channels. When the wave rotor rotates, the several wave rotor channels communicate with the isolation section sleeve sequentially through the fan-shaped hole. The present invention suppresses reflected shock waves by changing a flow blockage ratio and a shape of the pneumatic valve to consume back transmission pressure, which is beneficial to a fuel intake process, so that steady working of the wave rotor combustor in a state of deviating from a design point can be implemented.

Abstract: Provided is a gas turbine combustor that can achieve improvement of the combustion stability. A gas turbine combustor includes a pilot burner of the diffusion combustion type, a pilot flow control valve that regulates a flow rate of fuel to be supplied to the pilot burner, a main burner of the premix combustion type arranged on an outer circumference side of the pilot burner, main flow control valves that regulate flow rates of fuel to be individually supplied to burner sectors into which the main burner is sectioned in a circumferential direction, and a controller configured to control the pilot flow control valve and the main flow control valves. The controller controls the main flow control valves such that, when fuel is to be supplied to all the burner sectors, a difference in fuel flow rate occurs between at least one burner sector and the other burner sectors among the burner sectors.

Abstract: A flow control actuator includes at least one side wall, an upstream wall coupled to an upstream end of the side wall, a downstream cap coupled to a downstream end of the side wall, the downstream cap comprising at least one orifice disposed therein, at least one fuel injector disposed in at least one of the upstream wall, and the sidewall, the fuel injector dispersing fuel into the interior of the flow control actuator, and at least one oxidizer inlet disposed in at least one of the upstream wall and the sidewall, the at least one oxidizer inlet introducing an oxidizer into the interior of the flow control actuator. The flow control actuator includes at least one external fuel injector disposed adjacent to the side wall. The fuel from the fuel injector and oxidizer from the oxidizer inlet ignite in the interior of the flow control actuator.

Abstract: A pulse combustor system for efficiently operating a pulse combustor. The pulse combustor system includes the pulse combustor and a duct. The pulse combustor has a combustion chamber defining an internal space, a conduit having a first end in fluid communication with the internal space and a second end in fluid communication with an environment outside of the pulse combustor system, and a fuel injector configured to inject fuel into the internal space of the combustion chamber. The duct has two openings, with one opening disposed adjacent to the second end of the conduit. The pulse combustor system has an average operating frequency, and the duct has a length that is about one quarter of a wavelength corresponding to the average operating frequency. The pulse combustor and the duct each has a central longitudinal axis, and the two axes are substantially aligned.

Abstract: An engine assembly includes an engine compartment containing an internal combustion engine and a cooler compartment adjacent the engine compartment containing a heat exchanger. The engine and cooler compartments have an opening defined therebetween. A forced air system is operable to drive an airflow. A method for cooling the engine and its compartment is disclosed.

Abstract: A wave rotor includes an inlet plate, an outlet plate, and a rotor drum positioned therebetween. The inlet plate is formed to include an inlet port arranged to receive gasses. The outlet plate is formed to include an outlet port arranged to receive the gasses flowing out of the wave rotor. The rotor drum is arranged to rotate relative to the inlet and outlet plates. A piston assembly is used to counteract forces from pressure built up in the rotor drum.

Abstract: A combustor is configured to operate in a rotating detonation mode and a deflagration mode. The combustor includes a housing and at least one initiator. The housing defines at least one combustion chamber and is configured for a deflagration process to occur within the at least one combustion chamber during operation in the deflagration mode and a rotating detonation process to occur within the at least one combustion chamber during operation in the rotating detonation mode. The at least one initiator is configured to initiate the rotating detonation process within the at least one combustion chamber during operation in the rotating detonation mode and to initiate the deflagration process within the at least one combustion chamber during operation in the deflagration mode.

Abstract: An engine without a compressor or a turbine comprises a combustion chamber for burning a fuel-air mixture formed by mixing a fuel with outside air; and an outside air introduction part for introducing outside air into the combustion chamber. The outside air introduction part comprises an intake main port for introducing outside air into the combustion chamber from the direction along the central axis of the combustion chamber and a plurality of intake sub-ports for introducing outside air into the combustion chamber from the direction toward the central axis. The intake sub-ports comprise ejection openings capable of ejecting outside air toward a collision point inside the combustion chamber. Streams of outside air ejected from the ejection openings of the intake sub-ports mutually collide at the collision point and are thereby compressed.

Abstract: A system of injecting fuel into the combustion chamber of an engine, including at least two fuel circuits, one permanent flow circuit and one intermittent flow circuit, fuel proportioning and distribution devices for proportioning fuel and distributing fuel between the two circuits and a controller. When an order to fill circuits with fuel after the circuit with intermittent flow has been drained is received, the controller is adapted to control the proportioning and distribution devices to obtain a predetermined fuel flow higher than the flow corresponding to the filling order and to supply the resulting surplus of fuel to the intermittent flow circuit for a predetermined duration.

Abstract: The injection speed of the injection valves in an internal combustion engine is increased by using a single injection valve configured to carry out multiple fuel injections and combustions per rotation cycle. The single-valve propulsion thermal reactor has a casing with upper and lower walls consecutively defining a sleeve for taking in a pressurized air flow, a combustion chamber, and a gas discharge nozzle. The thermal reactor has a single injection valve to inject fresh gas into the combustion chamber, and at least one valve to exhaust burnt gases, which extends about transverse axes. The valves are cylindrical and have multiple surfaces which have a circular cross-section and are separated by facets that define, by a rotation of the valves, the intake and discharge ports for the gases. Preferably, a thermal ignition tank is built into the combustion chamber.

Abstract: A fuel nozzle system for enabling a gas turbine to start and operate on low-Btu fuel includes a primary tip having primary fuel orifices and a primary fuel passage in fluid communication with the primary fuel orifices, and a fuel circuit capable of controlling flow rates of a first and second low-Btu fuel gases flowing into the fuel nozzle. The system is capable of operating at an ignition status, in which at least the first low-Btu fuel gas is fed to the primary fuel orifices and ignited to start the gas turbine, and a baseload status, in which at least the second low-Btu fuel gas is fired at baseload. The low-Btu fuel gas ignited at the ignition status has a content of the first low-Btu fuel gas higher than that of the low-Btu fuel gas fired at the baseload status. Methods for using the system are also provided.

Abstract: Reducing jet noise by weakening Mach cones in a jet exhaust gas streamtube. The Mach cones are weakened by modifying exhaust gas flow in a longitudinal axial core of the exhaust gas streamtube.

Abstract: Methods and systems for operating a gas turbine engine including a fuel delivery system and a plurality of combustor assemblies are provided. The fuel delivery system comprises a primary fuel circuit configured to continuously supply fuel to each of the plurality of combustor assemblies during a first mode of operation and a second mode of operation. At least one secondary fuel circuit of the fuel delivery system is configured to supply fuel to each of the plurality of combustor assemblies during the second mode of operation. The secondary fuel circuit includes at least one isolation valve coupled in flow communication with each of the plurality of combustor assemblies. The at least one isolation valve facilitates preventing fluid flow upstream into the secondary fuel circuit during the first mode of operation. The fuel delivery system, using the isolation valve, replaces a purging system in the gas turbine engine.

Abstract: A pulse detonation combustor (PDC) includes a combustion tube, an inlet located on an upstream end of the combustion tube which receives a flow of a fuel/air mixture, an enhanced DDT region located within the tube downstream of the inlet, a nozzle disposed on a downstream end of the tube and a fortified region disposed downstream of the enhanced DDT region and upstream of the nozzle. A combustion initiation system that provides multiple initiation locations at different axial stations along the length of the tube are positioned downstream of the inlet and upstream of the fortified region. The initiator system is operable to initiate combustion of a fuel-air mixture within the tube at a selected one of the initiation locations.

Abstract: One embodiment of the present invention is an engine. Another embodiment is a unique combustion system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for engines and combustion systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: The engine (10) includes at least one firing tube (12) wherein an exhaust stream (32) from the firing tube (12) drives a turbine (30). A scroll ejector attenuator (40) is secured between and in fluid communication with an outlet end (28) of the firing tube (12) and an inlet (76) of the turbine (30). The attenuator (40) defines a turning, narrowing passageway (72) that extends a distance the exhaust stream (32) travels before entering the turbine (30) to attenuate shockwaves and mix the pulsed exhaust stream (32) into an even stream with minimal temperature differences to thereby enhance efficient operation of the turbine (30) without any significant pressure decline of exhaust stream (32) pressure and without any backpressure from the attenuator (40) on the firing tube (12).

Abstract: The flow through the core of a hybrid pulse detonation combustion system is passed through a compressor and then separated into a primary flow, that passes directly to the combustor, and a bypass flow, which is routed to a portion of the system to be used to cool components of the system. The bypass flow is routed to a nozzle of the pulse detonation combustor. The flow is then passed back into the primary flow through the core downstream of where it was extracted.

Abstract: In one embodiment, a pulse detonation system includes a pulse detonation tube including a base tube and a thermally protective layer disposed adjacent to an inner surface of the base tube. The thermally protective layer is configured to limit temperature fluctuations at the inner surface of the base tube to less than approximately 20 degrees Celsius during operation of the pulse detonation system, and the thermally protective layer does not comprise a ceramic coating.

Abstract: The invention relates to a jet engine (1), in particular an aircraft jet engine, including at least one combustion chamber (3). The combustion chamber (3) is connected to at least one compressed gas intake (4) and to at least one burnt gas outlet (5). Said burnt gas outlet(s) (5) include an exhaust valve. The exhaust valve includes two rotary parts (7), referred to as rotary exhaust parts, the rotary exhaust parts (7) including curved walls (8) and intermediate walls (9) connecting the curved walls (8), and rotating in a coordinated, continuous fashion such that said valve is in a closed position in order to block the gas during at least one combustion step, and in an open position in order to define a space (10) through which the gas flows out from the combustion chamber (3) during at least one expansion step.

Abstract: A pulse detonation combustor including a plurality of nozzles engaged with one another via mating surfaces to support a gas discharge annulus in a circumferential direction. The pulse detonation combustor also including multiple pulse detonation tubes extending for the nozzles and a plurality of thermal expansion control joints coupled to the plurality of pulse detonation tubes. Each of the plurality of thermal expansion control joints is configured to facilitate independent thermal growth of each of the plurality of pulse detonation tubes. The thermal expansion control joints may be configured as a bellows expansion joint or a sliding expansion joint.

Abstract: Apparatus and methods for combustion of fuel includes, in some embodiments, a fuel nozzle which injects fuel into a combustion channel of a wave rotor combustor or a pulse detonation combustor. In some embodiments the combustion process includes a backward-propagating detonation wave within a substantially closed channel which compresses discrete quantities of combustible and noncombustible mixture. Yet other embodiments include a precombustion chamber integrated into the wave rotor, the outlet stator or both.

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

Abstract: The engine (10) includes at least one firing tube (12) wherein an exhaust stream (32) from the firing tube (12) drives a turbine (30). A scroll ejector attenuator (40) is secured between and in fluid communication with an outlet end (28) of the firing tube (12) and an inlet (76) of the turbine (30). The attenuator (40) defines a turning, narrowing passageway (72) that extends a distance the exhaust stream (32) travels before entering the turbine (30) to attenuate shockwaves and mix the pulsed exhaust stream (32) into an even stream with minimal temperature differences to thereby enhance efficient operation of the turbine (30) without any significant pressure decline of exhaust stream (32) pressure and without any backpressure from the attenuator (40) on the firing tube (12).

Abstract: According to the invention, the said engine (I), which comprises at least one flame tube (2) with a mobile transverse end wall (18), comprises an external envelope (3) around the said flame tube (2), which defines a peripheral annular space (4) in which fixed flow guides (11, 12, 13, 14) are positioned, these flow guides forming flow channels (10) for the air, and at least one mobile plug (25), connected to the said mobile end wall (18), to close off and open one of the flow channels (10).

Abstract: An engine including a means for supplying fuel to the combustion chamber of the fire tube, that includes a variable-volume transfer chamber for receiving fuel, a fuel transfer means from the tank of the engine towards the transfer chamber, and injection means for injecting fuel into the combustion chamber from the transfer chamber. The engine further includes an elastic return at least partially defined by the fuel contained in the transfer chamber.

Abstract: A pulse detonation device contains a pulse detonation combustor which detonates a mixture of oxidizer and fuel. The fuel is supplied through fuel ducts and the fuel flow is controlled by fuel flow control devices. Oxidizer flow is provided through a main inlet portion and a flow control device directs the oxidizer flow to either the combustor or to a bypass duct, or both. The combustor further contains an ignition source. Each of the flow control devices, fuel flow control devices and ignition source are controlled by a control system to optimize performance at different thrust/power settings for the device.

Abstract: In one embodiment, a pulse detonation tube includes a continuous base tube having a substantially uniform wall thickness. The pulse detonation tube also includes a local flexural wave modifying feature configured to locally vary a flexural wave speed such that the flexural wave speed through the pulse detonation tube is different than an expected detonation wave speed, and/or to locally dissipate flexural wave energy.

Abstract: One embodiment of the present invention is a unique pulse detonation combustion system. Another embodiment is a unique gas turbine engine including a unique pulse detonation combustion system. In some embodiments, the pulse detonation combustion system includes an inlet section, a vortex generator and at least one flame accelerator. The inlet section, vortex generator and the at least one flame accelerator may be operative to initiate a deflagration to detonation transition. In some embodiments, the pulse detonation combustion system may include a flame accelerator configured with a directionally-dependent drag coefficient. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for pulse detonation combustion systems, gas turbine engines, and other machines and engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: A rotating valve assembly includes an inner cup having at least one inlet port; an outer cup having at least one inlet port, the outer cup rotatably mounted concentric with the inner cup by a bearing arrangement; and a cooling system located between the inner cup and the bearing arrangement for providing a thermal barrier between the inner cup and the bearing arrangement. The valve assembly also includes a labyrinth sealing arrangement located around the at least one inlet port of one of the inner and outer cups for preventing leakage of pressure waves generated by detonations or quasi-detonations within a combustion chamber of the inner cup.

Abstract: A detonation chamber and a pulse detonation combustor including a detonation chamber, wherein the detonation chamber includes a plurality of aerodynamic jets disposed adjacent an exterior of a sidewall of the detonation chamber. The detonation chamber further includes a plurality of openings formed in the sidewall of the detonation chamber, wherein each of the plurality of openings is in fluidic communication with one of the plurality of aerodynamic jets. The plurality of aerodynamic jets are adapted to create a plurality of jet flows of a fluid within the detonation chamber during a combustion cycle defining a plurality of initiation obstacles within the detonation chamber to enhance a turbulence of a fluid flow and flame acceleration through the detonation chamber.

Abstract: A constant volume combustor device includes, in one form, a detonative combustion. 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. In one form the wave rotor of the constant volume combustor is supported by magnetic bearings.

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 and a pulse detonation combustor including a detonation chamber, wherein the detonation chamber includes a plurality of initiation obstacles and at least one injector in fluid flow communication with each of the plurality of initiation obstacles. The plurality of initiation obstacles are disposed on at least a portion of an inner surface of the detonation chamber with each of the plurality of initiation obstacles defining a low pressure region at a trailing edge. The plurality of initiation obstacles are configured to enhance a turbulence of a fluid flow and flame acceleration through the detonation chamber. The at least one injector in provides a cooling fluid flow to each of the plurality of initiation obstacles, wherein the cooling fluid flow is one of a fuel, a combination of fuels, air, or a fuel/air mixture.

Abstract: The invention relates to a pulse detonation engine operating with an air-fuel mixture. According to the invention, the engine (1) includes at least two predetonation tubes (4, 5) which operate under conditions close to thermal cutoff conditions and the shock waves from which are focused in the combustion chamber (19).

Abstract: A vortex cannon based on pulse detonation engine comprises a combustion chamber, a fuel source, an oxidizer source, a purge gas source, a valve allowing delivery of fuel from the fuel source to the combustion chamber, a valve allowing deliver of either oxidizer or purge gas from the oxidizer and purge gas sources to the combustion chamber, an ignition source for the combustion chamber for initiating detonation of fuel and oxidizer, and a conical barrel outlet from the combustion chamber. The combustion chamber is configured for control over the detonation front. A control system provides for varying the rate and quantity of fuel and oxidizer injected to the combustion chamber for varying the frequency and strength of pulse generation. Ring vortices may be generated either in single pulses or at high rates of fire which maintain a consistent track.

Abstract: Such apparatuses are used for discharging active substances along with various carrier media, water often being used as a carrier medium. In order to ensure that the water perfectly nebulizes together with the active substance, the mist pipe (10) comprises at least three additional pipes (3 to 6) which partly surround each other to form annular chambers (31 to 33). Such a mist pipe (10) allows the size distribution of the drops to be kept within narrow limits even when water is used as a carrier medium. The nebulizer and the mist pipe (10) are mainly used in the health sector, in agriculture, plantations and greenhouses, for protecting supplies and for disinfection purposes on humans, in animal husbandry, and in food production.

Abstract: Apparatus and methods for combustion of fuel. Some embodiments of the inventions include a fuel nozzle which injects fuel into a combustion channel of a wave rotor combustor or a pulse detonation combustor In some embodiments the combustion process includes a backward-propagating detonation wave within a substantially closed channel which compresses discrete quantities of combustible and noncombustible mixture. Yet other embodiments include a precombustion chamber integrated into the wave rotor, the outlet stator or both.

Abstract: A helical cross flow pulse detonation engine.

Abstract: An engine contains at least one pulse detonation combustor having a combustion chamber and an exit nozzle coupled to and downstream of the combustion chamber. During operation of the at least one pulse detonation combustor a detonation occurs within the combustion chamber and at least one of a fuel fill fraction and purge fraction of the at least one pulse detonation combustor are utilized to offset a temperature peak of said detonation from a pressure peak of said detonation. The fuel fill fraction is defined as 1?purge fraction, and the purge fraction is the ratio of the purge time of the at least one pulse detonation combustor to a sum of the purge time of the at least one pulse detonation combustor and a fuel fill time of the at least one pulse detonation combustor.

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 gas turbine engine (10) having a pressure-rise combustor (30) and the pressure-rise combustor (30) is positioned upstream of a stage of turbine nozzle guide vanes (32). The vanes (33) of the stage of turbine nozzle guide vanes (32) forms an ejector and each of the vanes (33) has an upstream portion (34) and a downstream portion (36). The upstream portions (34) of the vanes (33) have leading edges (38) and the upstream portions (34) are arranged substantially straight and parallel to define constant area mixing passages (40) for a flow of gases there-through. The downstream portions (36) of the vanes (33) are arranged at an angle to the upstream portions (34) of the vanes (33) to turn the flow of gases there-through.

Abstract: A pulse detonation engine including one or more fuel injectors comprising one or more piezoelectric driving stacks wherein a flow control member of each injector is driven directly by the one or more piezoelectric stacks without additional amplification means or interposing elements while a flow area of the nozzle is variably adjustable to deliver controlled flow rates in a desired flow profile to improve engine performance and reduce emissions. The pulse detonation engine configured to support variable mission and operational requirements including delivery of required thrust using specific fuel types and with power and performance of the pulse detonation engine variably adaptable. The fuel injectors associated with the pulse detonation engine configure to deliver specified flow rates with minimal linear movement of the flow control member. The injector and drive electronics configured to deliver higher frequency operation and response with increased operational stability.

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: A pulse detonation combustor includes at least one plenum located along the length of the pulse detonation combustor. The plenum can be located: 1) proximate an air valve; 2) between a fuel injection port and an ignition source; 3) downstream of both the fuel injection port and the ignition source; and 4) proximate an exit nozzle of the pulse detonation combustor. In addition, the pulse detonation combustor can have multiple plenums, for example, proximate the air valve and proximate the exit nozzle. The location and dimensions of the plenum can be selectively adjusted to control mechanical loading on the wall, the velocity of fluid flowing within the combustor, and the pressure generated by the pulse detonation combustor.

Abstract: A pulsed detonation engine may include a detonation tube for receiving fuel and an oxidizer to be detonated therein, one or more fuel-oxidizer injectors for injecting the fuel and oxidizer into the detonation tube, one or more purge air injectors for injecting purge air into the detonation tube for purging the detonation tube, and an ignition for igniting the fuel and oxidizer in the detonation tube so as to initiate detonation thereof. The detonation tube has an upstream end, a downstream end, and an axially extended portion extending from the upstream end to the downstream end and having a perimeter. The fuel-oxidizer injectors and purge air injectors may be disposed at least along the axially extended portion. The ignition may include a plurality of igniters disposed at or near the perimeter of the axially extended portion, spaced about the perimeter, at or near the upstream end of the detonation tube.

Abstract: An engine contains a compressor stage, a plurality of pulse detonation combustors and a rotary inlet valve structure having a plurality of inlet ports through which at least air flows to enter the pulse detonation combustors during operation of the engine. Downstream of the pulse detonation combustors is a turbine stage. Further, the ratio of the pulse detonation combustors to the inlet ports is a non-integer.