171results about "Explosive combustion chamber" patented technology

The invention discloses a wide-range rotary detonation combustion chamber runner structure and an operation method thereof, and belongs to the technical field of rotary detonation ramjet engines. Comprising an equipment throat, an expansion section and an outer ring switching section, the end, away from the equipment throat, of the expansion section is connected with an outer ring switching device, the outer ring switching device is connected with an equal direct oil injection section, the equal direct oil injection section is connected with an expansion concave cavity section, and the expansion concave cavity section is connected with an exhaust nozzle shell; an inner column is arranged in the outer ring switching device and the equal-straight oil injection section and movably connected with a plug type spray pipe. According to the flow channel structure of the wide-range rotary detonation combustion chamber, a plurality of static and dynamic pressure measuring points are arranged on the surface of the outer ring switching section, so that the combustion state is monitored in real time; the plug type spray pipe moves back and forth in the inner column guide rail through the lead screw, and the throat area is flexibly adjusted so as to adapt to the working range from the low Mach number to the high Mach number. The equal direct oil injection section ensures different combustion requirements through a multi-group annular injection design; the expansion concave cavity section is provided with a pre-explosion pipe connector and a plasma igniter installation panel, and the multi-work ignition requirement is met.

The invention discloses a rocket-based rotary detonation engine combustion chamber structure and application thereof, and belongs to the technical field of rocket engine research. The combustion chamber structure comprises an isolation section, an equal-direction oil injection section and an expansion-concave cavity section which are connected in sequence. A central rocket is arranged in the isolation section and the equal direct oil injection section; and four groups of first ring-jet injection points are arranged on the equal straight oil injection section, two groups of second ring-jet injection points are arranged on the central rocket, and two groups of third ring-jet injection points are arranged on the expansion-concave cavity section. Oil is injected into the expansion-concave cavity section firstly, a pre-detonation pipe is adopted for igniting an oil-gas mixture, rotating detonation waves are formed, the built-in center rocket can provide thrust for the aircraft at the ground zero speed and form an injection effect for air suction, extra thrust is provided when the cruising thrust of the aircraft is insufficient, and when the rotating detonation combustion flameout boundary exists, the detonation effect is achieved. A flame stabilizing mechanism is provided for rotary detonation combustion, and stable work of wide-speed-range and wide-height-range rotary detonation combustion is achieved.

The present disclosure is directed to a rotating detonation combustor that includes a forward wall, a radially inner wall, and a radially outer wall. The forward wall is disposed at an inlet end of the rotating detonation combustor. The radially inner wall surrounds a longitudinal axis and extends downstream from the forward wall to an outlet end of the rotating detonation combustor. The radially outer wall extends downstream from the forward wall to the outlet end and surrounds the radially inner wall to define at least one annular plenum between the radially inner wall and the radially outer wall. At least one partition is proximate to 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 design method of a bending detonation combustion chamber with a V-shaped passivated front edge, and belongs to the field of hypersonic velocity detonation engines of aircrafts. The oblique detonation engine aims at solving the problems of an existing oblique detonation engine in the aspects of detonation, standing, mixing and the like. According to the method, by determining the temperature increasing amplitude of the V-shaped front edge, the distance and flow parameters of a detached shock wave are determined by using a two-dimensional detached shock wave theory, the geometrical shape and interference type of a bent shock wave are determined based on a bent shock wave theory, and the crotch shape of the passivated V-shaped front edge is determined. And finally, constructing a geometric model and a combustion chamber geometric model according to the passivation angle and curvature of the V-shaped front edge. According to the method, the length of the engine can be effectively shortened, the formation and stability of detonation waves are improved, effective standing of the detonation waves is achieved, and normal work of the engine is guaranteed. The combustion chamber has the advantages of saving space, being easy to detonate, compact in structure and the like, and is suitable for a propelling system of a hypersonic aircraft.

A system and method are disclosed for improving the operation of pulse combustors that are used in pulsejets by disposing a turbulator in the inlet pipe near where it is connected to the combustion chamber to cause increased turbulence in the combustion chamber to enhance fuel / air mixing, increased combustion chamber pressure, and flame-holding.

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 pulse detonation engine oil supply, in particular to a pulse detonation combustion chamber oil supply system which comprises a fuel oil filter, a flow regulating valve, a fuel oil flow meter, a fuel oil electric heater, an electromagnetic valve, a high-pressure direct-drive electronic fuel injector, a detonation chamber wall surface regeneration cooling mechanism and a fuel oil control system. Fuel oil is electrically heated in the starting process, an electric heater and the high-temperature wall face of the detonation chamber are used for heating the fuel oil within the full working condition range, the initial temperature of the fuel oil entering the pulse detonation combustion chamber is increased, the atomization mixing effect of the fuel oil in the pulse detonation combustion chamber is improved, and the combustion heat efficiency in the pulse detonation combustion chamber is improved; the regenerative cooling structure plays a role in heat exchange and detonation wave formation acceleration on the outer wall and the inner wall of the pulse detonation combustion chamber, the overall structure is simplified while the function is guaranteed, and the reliability of system operation is improved. And the requirements of the pulse detonation combustion chamber on stable oil supply flow under different working conditions and special conditions are met.

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.

The invention belongs to the technical field of engines, and discloses a rotary detonation combustion chamber which comprises a combustion chamber body, the combustion chamber body comprises a combustion chamber shell composed of an inner sleeve structure and an outer sleeve structure, a fluid channel is formed in the combustion chamber shell, and the inlet end of the fluid channel is a combustion chamber air inlet; the fluid channel comprises a combustion chamber throat section, a first expansion section, a U-shaped groove section, a second expansion section and a combustion chamber downstream straight section which are sequentially arranged. One end of the combustion chamber throat section communicates with the combustion chamber air inlet. And the U-shaped groove section comprises a U-shaped groove fillet section and a U-shaped groove straightway section which are arranged in sequence. The whole U-shaped groove fillet section is of a topological fluid diode structure, and one end, close to the first expansion section, of the topological fluid diode structure is of an arc-shaped structure. The topological fluid diode structure provides a new view angle and a new solution for optimization of the combustion chamber, and design and development of the more efficient and more stable combustion chamber are promoted.

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.

A system and method is disclosed for improving fuel injection systems for pulse combustors that reduce protrusions that can detrimentally impact the performance of pulsejet aircraft to which such pulse combustors are associated. The fuel injection system disclosed herein can be used with a pulse combustor to minimize protrusions, with the pulse combustor including at least an inlet pipe, a combustion chamber, an exhaust pipe, an ignition device, a fuel supply system, with the fuel injection system.

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.

A system and method are disclosed for improving the operation of pulse combustors that are used in pulsejets by disposing a turbulator in the inlet pipe near where it is connected to the combustion chamber to cause increased turbulence in the combustion chamber to enhance fuel / air mixing, increased combustion chamber pressure, and flame-holding.

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.

A combustor system is provided that includes rotating detonation combustors (RDCs) arranged in series. In some embodiments, a combustor system includes a volume that receives a core oxidant-fuel mixture. The combustor system includes a first RDC having a first detonation chamber. The first detonation chamber receives a first pilot oxidant-fuel mixture and is defined by a first channel formed in the peripheral wall. The combustor system also includes at least one additional RDC having a second detonation chamber. A second detonation chamber receives a second pilot oxidant-fuel mixture and is defined by a second channel formed in the peripheral wall. The first pilot oxidant-fuel mixture reacts in the first detonation chamber and the second pilot oxidant-fuel mixture reacts in the second detonation chamber to generate a rotating detonation combustion wave directed by the first channel and the second channel to support the reaction consuming the core oxidant-fuel mixture.

The invention provides a valve-free liquid fuel pulse detonation burner which can continuously feed and automatically generate countercurrent detonation pulse waves to generate high-temperature, high-speed and high-pressure oxidation reaction. The pulse detonation burner comprises a group of pulse detonation burner pipes containing detonation promoters, and the pulse detonation burner pipes are composed of a plurality of reducing pipe sections, expanding pipe sections and straight pipe sections which are connected in series. A heat transfer strengthening structure is arranged on the outer surface of the pulse detonation burner tube; a group of liquid fuel atomizing nozzles; a plurality of groups of ignition devices are arranged on the straight pipe section of the pulse detonation burner pipe; the detonation wave discharge pipe is arranged at the downstream of the straight pipe section of the pulse detonation burner pipe; and a secondary vortex air jacket.

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.

The invention relates to the field of aero-engines, and provides a rotary detonation afterburner system capable of regulating and controlling the injection temperature of aviation kerosene, which comprises a rotary detonation afterburner, a fuel tank, an oxygen bottle, a first fuel supply pipeline, a second fuel supply pipeline and a first gas supply pipeline, a first channel is arranged in the side wall of the rotary detonation afterburner, the fuel tank is communicated with the first channel through a first oil supply pipeline, and the first channel is communicated with the interior of the rotary detonation afterburner through a second oil supply pipeline; the oxygen bottle communicates with the interior of the rotary detonation afterburner through a first gas supply pipeline, and the oxygen bottle is used for supplying oxygen to the interior of the rotary detonation afterburner in the detonation stage of the rotary detonation afterburner. The rotary detonation afterburner solves the problem that the rotary detonation afterburner is difficult to detonate and maintain.

Combustor systems are provided that include rotating detonation combustors (RDCs) arranged in series. In some embodiments, a combustor system includes a volume that receives a core oxidizer-fuel mixture. The combustor system includes a first RDC having a first detonation chamber. The first detonation chamber receives a first pilot oxidizer-fuel mixture and is bounded by a first channel formed in a peripheral wall. The combustor system further includes at least one additional RDC having a second detonation chamber. The second detonation chamber receives a second pilot oxidizer-fuel mixture and is bounded by a second channel formed in the peripheral wall. The first pilot oxidizer-fuel mixture reacts in the first detonation chamber and the second pilot oxidizer-fuel mixture reacts in the second detonation chamber to generate rotating detonation combustion waves that are guided by the first channel and the second channel to support a reaction that consumes the core oxidizer-fuel mixture.

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 present invention provides a detonation combustion chamber and a gas turbine, the detonation combustion chamber comprising: a detonation combustion chamber shell, the inner cavity of the detonation combustion chamber shell is coaxially sleeved with a detonation outer ring and a detonation inner ring from the outside to the inside, the detonation outer ring and the detonation inner ring are both tubular structures, a main flow channel annular cavity for detonation combustion is formed between the annular walls of the detonation outer ring and the detonation inner ring, the air inlet end of the main flow channel annular cavity is connected to air, and the air is mixed with the fuel in the cavity for detonation combustion; a first annular cavity is arranged between the detonation combustion chamber shell and the annular wall of the detonation outer ring, and the first annular cavity is used as an outer cooling channel to pass circulating coolant; the inner cavity of the detonation inner ring is provided with an inner cooling channel, and circulating coolant is passed into the inner cooling channel. The structure of the present invention can effectively improve the cooling effect of the detonation combustion chamber, ensure the long-term stable operation of the gas engine, and improve the power adjustment range of the gas turbine, improve the combustion efficiency, and reduce the difficulty of processing.

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.

The invention discloses an oblique detonation engine suitable for supersonic wide-range operation conditions, the lower surface of a combustion chamber of the oblique detonation engine is arranged to be a wedge surface, the wedge surface of the combustion chamber adopts a curved surface structure design in the direction perpendicular to the flowing direction, so that the angle of the wedge surface continuously changes in the direction perpendicular to the incoming flow direction of an inlet of the combustion chamber. According to the oblique detonation engine provided by the invention, through the structural design of the three-dimensional curved wedge surface combustion chamber, wedge surfaces with continuously changed angles are formed at different positions perpendicular to the incoming flow direction in the combustion chamber, and different positions have different detonation angles, so that quick triggering of local detonation points is realized; and formation and development of oblique detonation waves in the whole combustion chamber are accelerated. Through the design of the combustion chamber of the three-dimensional curved wedge surface structure, the stable operation working condition range of the combustion chamber of the supersonic propulsion system based on oblique detonation combustion under the wide-area flight condition is widened, the through-flow area of the combustion chamber is increased while rapid and stable ignition is guaranteed, and the flow pressure loss in the combustion chamber is reduced.