58results about "Explosive combustion chamber" patented technology

A method of controlling a deflagration combustion process in a pistonless combustor includes scavenging combustion products of the previous cycle, introducing air into the combustor thereby initiating a flow pattern having a first flow component within the combustor, introducing air into the pistonless combustor in a nonparallel angle in relation to the previous air input and thereby creating a second flow component to the flow pattern for increasing speed of combustion propagation, introducing fuel mixed into the air creating a fuel-air mixture flowing within the flow pattern, igniting the fuel-air mixture within the pistonless combustor and introducing fuel by direct injection after the ignition thereby increasing pressure within the pistonless combustor.

The present disclosure is directed to a supersonic oblique rotating detonation wave engine (SORDE) and systems and methods for generating a supersonic oblique rotating detonation wave. The SORDE is configured to produce and sustain a supersonic oblique rotating detonation wave through the injection of fuel at supersonic speeds into an inlet air flow between Mach 1 and Mach 7. The SORDE and method include injecting fuel into the inlet air in an amount to generate an equivalence ratio of 0.2 to 2.5. Some embodiments include a plurality of fuel injector ports each having a diameter of about 0.010 inches to about 0.040 inches; an annular wedge disposed in or upstream of the detonation chamber with an angle of about 5 degrees to about 40 degrees relative to a longitudinal axis of the engine; and / or a cylindrical center body disposed in or upstream of the detonation chamber.

The invention relates to the technical field of aero-engine combustion chamber design, in particular to a multi-layer mixing section structure of a pulse detonation combustion chamber. The multi-layer mixing structure is characterized in that the multi-layer mixing structure comprises a multi-layer mixing section arranged in a detonation pipe, an oil supply pipe is arranged on the detonation pipe, the multi-layer mixing section comprises an inner mixing ring and an outer mixing ring, and the inner mixing ring and the outer mixing ring form three Venturi type circulation channels through connecting structures; and a fuel nozzle connected with an oil supply pipe is arranged at the axis of the inner mixing ring. By means of the multi-layer mixing section, incoming flow air is divided into multiple jet flows, the mixing efficiency of fuel oil and air is improved, the mixing uniformity of the fuel oil and air is remarkably improved, and the mixing distance is shortened. Meanwhile, according to the regenerative cooling structure, a fuel oil channel is designed in a multi-layer mixing section, recycling of the wall face of the pulse detonation combustion chamber is achieved, meanwhile, the oil injection temperature is increased, the fuel oil and air mixing effect is improved, and the combustion efficiency is improved.

An accessory power unit assembly includes a combustor assembly that includes a pulse detonation tube that defines a curved passage between an inlet that receives a compressed airflow and a discharge end. The detonation tube operates as a resonant cavity to sustain a standing pressure wave. Fuel injected into the pulse detonation tube is timed to coincide with movement of a standing wave toward a discharge end. The standing wave provides an increase in pressure of an output exhaust gas driving the turbine without a corresponding increase in load.

The application relates to the field of aeroengines, in particular to a rotating detonation combustion chamber with a pressure regulating structure and an injection channel, comprising a duct shell, a combustion chamber body arranged in the duct shell and a Tesla valve arranged at the inlet of the combustion chamber body, the combustion chamber body being provided with an annular channel, the Tesla valve comprising a shell and a flow channel, one end of the shell being fixedly connected with one end of the duct shell, the other end of the shell being coaxially connected with the combustion chamber body, the flow channel being coaxially arranged in the shell, the inlet end of the flow channel being used for air inlet, and the outlet end of the flow channel being communicated with the annular channel in the combustion chamber body; the injection channel is formed between the duct shell, the shell and the combustion chamber body, the inlet end and the outlet end of the flow channel are both provided with a pressure regulating structure, the pressure regulating structure comprises a gas collection cavity and an air inlet hole used for connecting the gas collection cavity with the flow channel, and the injection channel is communicated with the gas collection cavity. The application has the effect of improving the problem that the pressure backflow causes the total pressure loss of the combustion chamber to be large.

The application relates to the technical field of aero-engines, and provides a rotating detonation combustion chamber detonation wave propagation direction induced air intake pre-rotation structure which comprises an air intake shell, an air intake flow channel is arranged in the air intake shell, the air intake flow channel comprises an air intake passage, a connecting passage, an expansion passage and a backflow passage, one end of the air intake passage is used for introducing air and fuel, the other end is communicated with the connecting passage, the small end of the expansion passage is communicated with one end of the connecting passage away from the air intake passage, and the large end of the expansion passage is communicated with a combustion chamber; the inlet end of the backflow passage is communicated with one end of the connecting passage away from the expansion passage, and the outlet end of the backflow passage is communicated with the air intake passage; a plurality of guide vanes are arranged in the expansion passage and are arranged in a circumferential interval, and a pre-rotation passage for pre-rotating gas is formed between two adjacent guide vanes. The rotating detonation combustion chamber detonation wave propagation direction induced air intake pre-rotation structure can improve the control accuracy of the detonation wave mode in the combustion chamber.

The invention discloses an air film cooling structure and a rotary detonation engine, and relates to the technical field of rotary detonation engine thermal protection, the air film cooling structure comprises a cold air cavity and a Tesla valve type air film pipe, the Tesla valve type air film pipe comprises a main flow pipe, a flow dividing pipe and a collecting pipe, and an inlet of the flow dividing pipe is connected and communicated with the side portion of the main flow pipe; an outlet of the flow dividing pipe is connected and communicated with an inlet of the collecting pipe, an outlet of the main flow pipe is connected and communicated with the side portion of the collecting pipe, the main flow pipe and the collecting pipe are linear, the flow dividing pipe is an arc-shaped pipe and protrudes upwards relative to the collecting pipe, and an inlet of the main flow pipe is connected and communicated with the cold air cavity. And an outlet of the collecting pipe is connected and communicated with a combustion chamber of the rotary detonation engine. The rotary detonation engine comprises a combustion chamber and the air film cooling structure, and the combustion chamber is provided with a combustion inlet and a combustion outlet. According to the invention, high-temperature fuel gas anti-invasion can be inhibited when detonation waves pass through, the influence of high-temperature fuel gas on the cold air cavity is reduced, and the cooling efficiency is improved.

A rotary knock combustion system includes a knock channel including an inner wall and an outer wall and extending in a longitudinal direction from an inlet of the knock channel to an outlet of the knock channel. The first mixing element and the second mixing element are disposed in the rotary knock combustion system. The first mixing element forms a ring on the inner wall and the second mixing element forms a ring on the outer wall of the knock channel adjacent the inlet. Each of the first and second mixing elements includes a plurality of protrusions disposed circumferentially along the inner and outer walls and extending into the knock channel such that the plurality of protrusions affect a vector of at least a portion of the fuel and at least a portion of the fluid passing through the recesses between the plurality of protrusions.

Systems and methods are provided herein useful to thrust augmentation in a gas turbine engine. In some embodiments, the systems include augmentors that incorporate a rotating detonation architecture. An exhaust system of a gas turbine engine includes an augmentor and a peripheral wall surrounding an exhaust system core. The augmentor comprises a detonation chamber disposed within the exhaust system core. The detonation chamber includes a channel formed in the peripheral wall. A core inlet path delivers a core air-fuel mixture and a pilot inlet path delivers a pilot air-fuel mixture to the detonation chamber. The core air-fuel mixture combusts in the detonation chamber along the midline the detonation chamber. The pilot air-fuel mixture detonates in the detonation chamber adjacent the peripheral wall to create a rotating detonation wave that supports the combustion reaction occurring along the midline of the detonation chamber.

The invention belongs to the technical field of continuous detonation engines, and particularly relates to a continuous detonation turbine engine based on coaxial shear injection, a fuel confluence cavity and a plurality of rows of array type coaxial shear injection units are arranged at one end, facing a diffuser, of a flame tube of the engine; each coaxial shearing injection unit is composed of a central air channel and an annular fuel jet gap; the fuel oil pipeline comprises a main fuel oil supply pipeline attached to the outer wall of the flame tube, a fuel oil collecting pipe and a plurality of fuel oil flow dividing thin pipes. The fuel oil collecting pipe is arranged on the peripheral side of the outlet end of the flame tube and communicates with the main fuel oil supply way and the fuel oil flow dividing thin pipes. The multiple fuel oil flow dividing thin pipes are distributed in the circumferential direction of the flame tube. And the annular fuel jet flow gap is communicated with the fuel oil confluence cavity, so that air flow jetted from the central air channel and gasified fuel jetted from the annular fuel jet flow gap are coaxially sheared and mixed. The engine is favorable for meeting the requirements of continuous detonation combustion on mixing uniformity, reaction induction time and combustion stability.

The application relates to a double-channel rotating detonation combustion chamber structure with an inlet total pressure adjusting function, which comprises a Tesla valve arranged at the inlet of the rotating detonation combustion chamber, the Tesla valve comprises a shell and a flow channel, the inlet end and the outlet end of the flow channel are respectively provided with pressure adjusting structures, the radial outer side of the flow channel is provided with an injection channel, the injection channel comprises an air outlet channel and a pressurizing channel communicated with the air outlet channel, and the side, away from the pressurizing channel, of the air outlet channel is provided with an opening and closing assembly for opening or closing the air outlet channel, the opening and closing assembly comprises a supporting plate, a plurality of opening and closing doors and a self-locking driving part, and the self-locking driving part is used for synchronously driving the plurality of opening and closing doors to be close to or away from the outlet side of the air outlet channel. The application has the effects that when the rotating detonation combustion chamber is in an air thin belt, the air intake of the air inlet channel is automatically enhanced, the rotating detonation combustion chamber preferentially intakes the air in the air inlet channel, the air supply of the combustion chamber is ensured, and the adaptability and stability of the combustion chamber in different environments are improved.

A system and method for ram air intake for pulse combustion systems is disclosed that improves the ability of pulse combustions to ingest air into the inlet pipe when the pulse combustion system is moving in a direction opposite the direction the open end of the inlet pipe is facing and the system and method includes the ability to increase the thrust output from the pulse combustion system.

The present disclosure relates to a rotary detonation combustor including a front wall, a radially inner wall, and a radially outer wall. The front wall is disposed at an inlet end of the rotary detonation combustor. The radially inner wall surrounds a longitudinal axis and extends downstream from the front wall to an outlet end of the rotary detonation combustor. The radially outer wall extends downstream from the front wall to the outlet end and surrounds the radially inner wall to define at least one annular cell between the radially inner wall and the radially outer wall. At least one partition is proximate the inlet end and defines at least two mixing zones. A plurality of oxidizer inlets and a plurality of fuel inlets are disposed at the inlet end in fluid communication with the at least two mixing zones.

The invention discloses a device and a method for realizing efficient extraction of graded detonation energy, and the method creatively changes the existing pulse detonation technology, and adopts two-stage detonation on the premise of not changing the initial filling pressure instead of adopting single detonation; the method specifically comprises the steps that initial detonation waves are formed at low filling pressure in two straight-pipe-shaped detonation combustion chambers through first-stage detonation, and overfilled fuel is pre-compressed; in the second-stage detonation, direct detonation is conducted through shock wave focusing at high filling pressure in an expansion cavity, the pressure of a combustion device is improved through two-stage detonation combustion, efficient extraction of fuel chemical energy is achieved, and the thermodynamic cycle performance of a power system adopting the combustion mode is improved; and meanwhile, the problem of fuel waste caused by over-filling is solved, and the structure is simpler and more efficient.

The invention relates to the technical field of aero-engines, and discloses a pulse detonation turbine engine which comprises an air inlet assembly, a supercharging assembly, a combustion chamber casing, a pulse detonation combustion chamber, a nozzle assembly, a transition assembly and an exhaust assembly. The first end of the pulse detonation combustion chamber is provided with a first channel and a second channel, and the second end is provided with a first exhaust channel and extends outwards from the combustion chamber casing. A two-way channel is formed between the combustion chamber casing and the pulse detonation combustion chamber, and the two-way channel is communicated with the first channel; the transition assembly is arranged at the end, close to the exhaust direction, of the pressurization assembly and communicates with the second channel. The two channels are formed between the pulse detonation combustion chamber and the combustion chamber casing, the intensity of return detonation can be effectively weakened through the cavity effect of the two channels, and then the return pressure faced by the pressurization assembly is reduced; meanwhile, after air enters the two channels through the pressurizing assembly, the wall face of the pulse detonation combustion chamber can be cooled.

A rocket engine system having a coolant source including a coolant fuel, the coolant source for supplying the coolant fuel to a first surface of a wall that partially defines a combustion chamber of the rocket engine, the coolant fuel film cooling the first surface of the wall, and the oxidizer source for supplying the oxidizer to a second surface of the wall. The monitor is configured to control the flow of the coolant fuel and the oxidizer and to ensure that a stoichiometric ratio of the coolant fuel and oxidizer mixture is suitable for causing combustion of the coolant fuel and oxidizer mixture after the coolant fuel has film cooled the first surface of the wall.

An explosive combustion system includes an explosive combustor. The explosive combustor includes an explosive manifold and one or more explosive chamber walls defining an explosive chamber. The explosive manifold includes a plurality of explosive fluid paths defined by an overall structure of the explosive manifold, and a plurality of explosive port sets each including a plurality of explosive ports disposed about a surface of the explosive manifold. A respective one of the plurality of explosive port sets provides fluid communication from a corresponding one of the plurality of explosive fluid paths to the explosive chamber through the plurality of explosive ports corresponding to the respective one of the plurality of explosive port sets. The plurality of explosive ports can be symmetrically oriented about a reference element of the explosive combustor.

The present application belongs to the technical field of continuous detonation engine, and particularly relates to a continuous detonation turbine engine based on coaxial shearing injection, wherein a flame tube of the engine is provided with a fuel converging cavity and a plurality of arrayed coaxial shearing injection units at an end facing a diffuser; each coaxial shearing injection unit is composed of a central air passage and an annular fuel jet slit; a fuel pipeline comprises a fuel total oil supply path attached to the outer wall of the flame tube, a fuel converging pipe and a plurality of fuel shunt thin pipes; the fuel converging pipe is arranged at the outer circumferential side of the outlet end of the flame tube and is in communication with the fuel total oil supply path and the fuel shunt thin pipes; the plurality of fuel shunt thin pipes are distributed along the circumference of the flame tube; the annular fuel jet slit is in communication with the fuel converging cavity, so that the airflow injected from the central air passage and the gasified fuel injected from the annular fuel jet slit are coaxially sheared and mixed. The above engine helps to meet the requirements of continuous detonation combustion on mixing uniformity, reaction induction time and combustion stability.

The rotary detonation combustion device comprises a combustion chamber outer ring, a combustion chamber head, a combustion chamber center body, a high-temperature-resistant sealing ring, a circular seam width adjusting shaft sleeve, an axial positioning ring and a fixed cylinder barrel. Wherein the combustion chamber center body comprises a combustion chamber center body front section, a combustion chamber center body rear section, a fuel cavity and a fuel spray hole; and the circular seam width adjusting shaft sleeve comprises a shaft sleeve main body, a shaft sleeve axial positioning boss and a shaft sleeve external telescopic arm. When the device works, the cylinder barrel drives the shaft sleeve to be externally connected with the telescopic arm to enable the circular seam width adjusting shaft sleeve to move in the axial direction, the width of an air injection circular seam jointly formed by the shaft sleeve body tail end molded surface and the front end of the combustion chamber center body rear section is adjusted to the optimal value, the detonation wave propagation stability can be enhanced, and active regulation and control over the working mode of the combustion chamber are achieved; and the stable working range of the rotary detonation combustion chamber is widened. The method can be used in the fields of detonation propulsion and the like.

The application discloses a reinforced mixing structure of an oblique detonation engine, which comprises a mixing cavity arranged on the inner bottom surface of a combustion chamber of the oblique detonation engine, a reinforced mixing step structure arranged on one side of the mixing cavity close to the inlet end of the combustion chamber, and an open structure with a flat bottom surface arranged on the other side of the mixing cavity away from the inlet end of the combustion chamber, wherein the open structure extends horizontally to the detonation wedge surface, the reinforced mixing step structure comprises a plurality of first steps and second steps arranged alternately along the width direction of the combustion chamber of the oblique detonation engine, the second steps are arranged protruding from the first steps, the steps on the side of the first steps and the second steps away from the outlet end of the combustion chamber are all inclined surfaces arranged in a downward inclination, and fuel injection holes are arranged on the inclined surfaces of the first steps and the second steps, and the fuel injection holes are communicated with a fuel delivery pipe arranged below the combustion chamber through connecting branch pipes. The reinforced mixing step structure is arranged to form a flow vortex to reinforce the mixing.

Disclosed is an aircraft having a rocket-based combined cycle (RBCC) propulsion system including a rotating detonation rocket engine (RDRE) and an airbreathing jet engine. The RDRE is located within a throat area of the jet engine and is configured to be cooled by air bypassed from upstream of the jet engine.

This invention discloses a turbine power generation device based on pulse detonation, comprising a compressor, a pulse detonation combustion chamber, a gas turbine, a coupling, a generator, and a controller. During operation, air enters from the compressor and is split into inner and outer bypass air. The inner bypass air passes through a cyclone separator and enters the mixing chamber. Fuel, controlled by a solenoid valve, enters the mixing chamber through a nozzle at the center of the cyclone separator, mixing with the air to form a premixed gas. After passing through a central conical blunt body, the premixed gas is ignited by a high-energy electric nozzle in a concave cavity and enters the detonation section. The flame accelerates in this section, ultimately forming a detonation wave. This detonation wave ejects the outer bypass air and enters the tailpipe together, further increasing the pressure. The ejected airflow enters the gas turbine and drives the generator to generate electricity. Compared to traditional slow-burning power generation devices, this invention has higher power generation efficiency, a wider range of power output variation, and can achieve rapid changes in power output.

The invention discloses a rotary detonation combustion chamber which comprises a gas and oil supply assembly, an injection assembly and a combustion chamber body. Wherein an annular channel is formed in the combustion chamber body, and the gas supply and oil supply assembly communicates with the gas supply system, can form high-speed airflow to enter the annular channel and communicates with the energy supply system and the injection assembly through the gas supply and oil supply system, so that fuel can enter the injection assembly through the gas supply and oil supply assembly; the high-speed airflow and the fuel jet flow can collide in the annular channel to be subjected to secondary atomization to form a combustible mixture, the combustion chamber body is provided with an igniter used for detonating the combustible mixture so as to form rotating detonation waves, and through the combustion mode of rotating detonation, the combustion efficiency is improved. By means of the self-pressurization structure, airflow expands and accelerates at the downstream of detonation waves and is exhausted from the tail of the combustion chamber in a high-temperature fuel gas mode, the heat efficiency of combustion can be improved through the self-pressurization characteristic of the self-pressurization structure, emission of pollutants such as NOx is reduced, and the self-pressurization structure is more compact.

A method of operating a rotating detonation combustor (104, 104', 108, 108', 114, 114') includes providing (S702) a flow of air (80, 98) through an air inlet (154, 154') to flow into a detonation chamber (128), providing (S703) a flow of fuel (68, 102) from at least one fuel injector (160, 160') into the detonation chamber (128), mixing (S704) the flow of the fuel (68, 102) and the flow of the air (80, 98) in the detonation chamber (128) to generate a fuel-air mixture (111), detonating (S705) the fuel-air mixture (111) in the detonation chamber (128) to generate rotating detonation waves (106, 112) within the detonation chamber (128), and controlling (S706-S715), during operation of the rotating detonation combustor (104, 104', 108, 108', 114, 114') from a first power operating state to a second power operating state, different from the first power operating state, the air inlet wall (164, 164') to control the flow of the air (80, 98) through the air inlet (154, 154') into the detonation chamber (128) to control a discharge coefficient and an operating mode within the detonation chamber (128).

This invention provides a rotary detonation combustion device based on wall microstructures, comprising: a combustion chamber and an ignition mechanism, an inner wall cooling mechanism, an outer wall cooling mechanism, and a fuel mechanism connected to the combustion chamber; the fuel mechanism for supplying combustible fuel to the combustion chamber; the ignition mechanism for controlling the ignition of the combustible fuel in the combustion chamber; the inner wall cooling mechanism for cooling the inner wall of the combustion chamber after the combustible fuel in the combustion chamber ignites and generates detonation; and the outer wall cooling mechanism for cooling the outer wall of the combustion chamber after the combustible fuel in the combustion chamber ignites and generates detonation. In this embodiment of the invention, the channel effectively increases the residence time of the coolant on the inner and outer walls, prevents the coolant from being blown away by the detonation shock wave, extends the cooling time for the coolant to form a cooling layer, and induces localized high-temperature and high-pressure electricity in the combustion chamber after the detonation wave sweeps, thus mitigating the impact of the inert coolant on the propagation stability of the detonation wave.

A method of operating a rotating detonation combustor includes providing a flow of air through an air inlet to flow into a detonation chamber, providing a flow of fuel from at least one fuel injector into the detonation chamber, mixing the flow of fuel and the flow of air in the detonation chamber to produce a fuel-air mixture, during operation of the rotary detonation combustor from a first power operating state to a second power operating state different from the first power operating state, the fuel-air mixture is detonated in the detonation chamber to generate a rotary detonation wave in the detonation chamber, and during operation of the rotary detonation combustor from the first power operating state to a second power operating state different from the first power operating state, the air inlet wall is controlled to control air flow into the detonation chamber through the air inlet, therefore, the emission coefficient and the operation mode in the detonation chamber are controlled.

We provide a rotary detonation engine that can compress an oxidizer without using exhaust gases. [Solution] A rotary detonation engine according to one aspect of the present disclosure includes a combustor having a cooling mechanism and generating a rotary detonation wave, a compressor for compressing an oxidant, a turbine for rotationally driving the compressor, a combustion oxidant passage for supplying the oxidant compressed by the compressor into the combustor, a cooling oxidant passage for supplying the oxidant compressed by the compressor to the cooling mechanism as a refrigerant, and a turbine oxidant passage for supplying the oxidant that has passed through the cooling mechanism to the turbine as a working fluid.

The embodiments disclosed herein disclose a fluidization device and a rotary detonation engine, and relate to but are not limited to the rotary detonation engine technology. The fluidization device comprises a first container having a space for accommodating solid powder fuel, a second container having a space for accommodating solid powder oxidizer, a driving mechanism and a gas supply mechanism. Thus, when the solid powder fuel and the solid powder oxidizer are fluidized, the gas supply mechanism can control the feeding amount of the solid powder fuel and the solid powder oxidizer moving to the fluidization position by the driving mechanism, so as to control the usage amount of the solid powder fuel and the solid powder oxidizer. Meanwhile, by controlling the gas supply amount supplied to the first container and the second container by the gas supply mechanism, the fluidization gas amount entering the first container and the second container can also be controlled, so that by controlling the feeding speed of the driving mechanism and the gas supply amount supplied to the first container and the second container, different fluidization effects of the solid powder fuel and the solid powder oxidizer can be achieved to adapt to different fluidization processing requirements.

The application discloses a fuel injection structure of a rotary detonation combustion chamber, and belongs to the technical field of fuel injection. The fuel injection structure comprises a containing cavity arranged on an air inlet device, the containing cavity is used for containing fuel, the air inlet device comprises an inner cylinder and an outer cylinder, a gap between the inner cylinder and the outer cylinder forms an air inlet channel, the air inlet channel comprises a first air inlet flow channel, a contraction flow channel, a second air inlet flow channel and an expansion flow channel, the containing cavity is sequentially provided with a first injection structure, a second injection structure and a third injection structure, the first injection structure is used for injecting fuel into the first air inlet flow channel, the second injection structure is used for injecting fuel into the contraction flow channel, and the third injection structure is used for injecting fuel into the second air inlet flow channel. The application has the effect of facilitating efficient and sufficient mixing of fuel added in the air inlet channel and air entering the rotary detonation combustion chamber, and improving the combustion efficiency in the rotary detonation combustion chamber.