111results 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 invention relates to the technical field of aero-engines, and discloses an efficient thermal protection hybrid thermodynamic cycle combined engine and a control method thereof.The combined engine composed of an external rotary detonation engine and an internal turbine engine is adopted, and modal switching of the combined engine is achieved through an inlet adjusting valve and an adjustable exhaust nozzle; the working upper limit of a turbine engine can be effectively joined, power relay can be completed, so that a thrust gap is spanned, the structural design complexity is relatively low, the system stability is relatively high, and in a stress application rotating detonation mode, the reliability is high. Combustible mixtures sprayed by the rotary detonation engine are guided into the afterburner of the turbine engine through the afterburner diverter valve for detonation combustion, an original cooling device of the afterburner is used for thermal protection, and an outer casing is prevented from being damaged due to high temperature.

A rotating detonation combustion system includes a detonation channel including an inner wall and an outer wall and extend in a longitudinal direction from an inlet of the detonation channel to an outlet of the detonation channel. A first mixing element and a second mixing element are disposed in the rotating detonation combustion system. The first mixing element forming a ring on the inner wall and the second mixing element forming a ring on the outer wall of the detonation channel adjacent the inlet. Each of the first mixing element and the second mixing element comprise a plurality of protrusions disposed circumferentially along the inner wall and the outer wall and extend into the detonation channel such that the plurality of protrusions affects vectors of at least a portion of the fuel and at least a portion of the fluid passing through recesses between the plurality of protrusions.

A detachable strut structure in a combustion chamber of a scramjet-oblique detonation engine includes two strut components and two hydraulic telescopic mechanisms. The strut components are disposed in the combustion chamber, the combustion chamber defines an inlet and an outlet, and a connection between the inlet and the outlet is a centerline, and the two strut components are mirrored and arranged on two sides of the centerline respectively. Each strut component includes a first surface, a second surface, a third surface, the first surface is arranged parallel to the centerline, the second surface is arranged on a side of the first surface close to the inlet. An angle α is defined between the second and first surfaces. The third surface is arranged on a side of the first surface close to the outlet, an angle β is defined between the third surface and an extension line of the first surface.

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.

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 the fuel and the flow of the air in the detonation chamber to generate a fuel-air mixture, detonating the fuel-air mixture in the detonation chamber to generate rotating detonation waves within the detonation chamber, and controlling, during operation of the rotating detonation combustor from a first power operating state to a second power operating state, different from the first power operating state, the air inlet wall to control the flow of the air through the air inlet into the detonation chamber to control a discharge coefficient and an operating mode within the detonation chamber.

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 relates to the technical field of aero-engines, and discloses a center detonation type oblique detonation engine which comprises a fuel injection supporting plate and a detonation supporting plate, the fuel injection supporting plate is erected and fixed in a combustion chamber, and a plurality of first reverse jet holes are formed in the end face, directly facing an air inlet channel, of the fuel injection supporting plate; a plurality of fuel injection holes are formed in the end face, back to the gas inlet channel, of the fuel injection supporting plate, and a first gas conveying flow channel and a fuel conveying flow channel are formed in the fuel injection supporting plate; the detonating supporting plate is erected and fixed in the combustion chamber and is parallel to the fuel injection supporting plate, and a plurality of second reverse jet holes are formed in the end face, directly facing the fuel injection holes, of the detonating supporting plate; the detonation supporting plate is arranged to replace an original wedge face, the detonation supporting plate and the fuel injection supporting plate both adopt reverse jet flow for thermal protection, arch shock waves formed through induction of the reverse jet flow play a role in intensifying detonation, meanwhile, oblique detonation waves are located in the center for detonation, and separation of a boundary layer of the lower wall face cannot be caused.

The invention discloses an inner-column-free rotary detonation combustion chamber based on center rotational flow, and belongs to the technical field of detonation engines. Comprising a combustion chamber front wall face, an outer wall face, a center rotational flow air inlet system and an oil supply system. The combustion chamber is divided into an outer ring rotary detonation combustion area and a center rotational flow subsonic velocity combustion area. The outer ring area forms supersonic airflow through a Laval nozzle to be mixed with fuel, and detonation waves propagating in a self-sustaining mode are formed through ignition of a pre-detonation pipe. The central area generates a backflow area through a swirler, and fuel and air are fully mixed and then ignited by a sparking plug. The double combustion modes work cooperatively to adapt to different working conditions, and it is ensured that the combustion chamber operates stably under the wide-range condition. The design does not need inner column support, the structure is simplified, and the combustion efficiency and the working stability are improved.

A fuel injector, a rotary detonation engine, and an aircraft are disclosed. The fuel injector includes: a main body having an annular receiving cavity and an annular combustion chamber, the receiving cavity being located radially outside the combustion chamber, a common wall between the receiving cavity and the combustion chamber, the common wall having leakage holes distributed thereon, the leakage holes connecting the receiving cavity and the combustion chamber; fuel and an elastic pressing device, both disposed within the combustion chamber, the elastic pressing device pressing fuel against the common wall; a pressure detection device configured to detect pressure data from the elastic pressing device; a vibration generating device disposed on the common wall and configured to drive the common wall to vibrate; and a control device electrically connected to the pressure detection device and the vibration generating device. The fuel is a solid powder fuel. The control device controls the vibration generating device to drive the common wall to vibrate at a set amplitude and a set frequency based on the pressure data and the fuel demand data of the combustion chamber, so that the leakage amount from the leakage holes meets the fuel demand data, thereby achieving precise control of solid powder fuel combustion.

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 present disclosure refers to a eating system (1) for producing and applying heat, comprising a heat generator (2) operable to produce a heated gas; and a heating applicator (3) operable to extract heat from the heated gas and apply the heat for heating. The heat generator (2) comprises a rotating detonation combustor (4) operable to produce and exhaust the heated gas. Further, a method for producing and applying heat in a heating system (1) is provided.

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.

The invention discloses a bimodal afterburner structure and a modal adjusting method thereof.The bimodal afterburner structure comprises an inner duct structure, an outer duct structure, a combustion chamber and a spray pipe, a partition plate is arranged between the inner duct structure and the outer duct structure, and a change-over valve is connected to the partition plate; one end of the combustion chamber is connected with the inner duct structure and the outer duct structure, and the other end of the combustion chamber is connected with the spray pipe. Through the design of a slow combustion / deflagration dual-mode inner and outer ring parallel structure, autonomous dual-mode switching and a dual-mode shared combustion chamber, efficient energy conversion and dynamic adjustment of the power device in a wide speed domain range are achieved, the thrust coefficient of a slow combustion mode is 0.3-0.5, the detonation mode is improved to 0.6-0.8, the flight requirement of Mach number 1-Mach number 5 can be met, and the flight speed of the power device is improved. And meanwhile, the detonation combustion heat efficiency is improved by 15% compared with a slow combustion mode, and the fuel power ratio is reduced by 15%-20%.

This group of inventions describes a group of devices based on a detonation engine. 1. Detonation engine as propulsion for aircraft, rocket system, space missile, space system 2. Detonation engine as the basis for energy systems 3. Detonation engine as the basis for a gas laser 4. Detonation engine as the basis for a free electron laser 5. Detonation engine as the basis for a cannon

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.