194 results about "Afterburner" patented technology

According to the dual-fuel afterburner without the inner duct stabilizer, an oil injection rod and an air injection rod are arranged in a supporting plate with a cavity structure, an oil injection nozzle and an air injection nozzle which are communicated are formed in the wall face of the supporting plate, and cooperative combustion of gas fuel and liquid fuel is achieved. According to the afterburner, a stable ignition source is provided through the characteristic that gas fuel is easy to burn, efficient and stable combustion is achieved in cooperation with fuel injection, and therefore a traditional inner duct stabilizer is omitted, the weight of the afterburner is remarkably reduced, and weight reduction and performance optimization of the afterburner are achieved. Evaporative flame stabilizers distributed in the circumferential direction are arranged in the outer duct area to achieve stable combustion of airflow in the outer duct, and flame stability and cross flame performance are guaranteed in cooperation with the supporting plates of the inner duct and fuel gas sprayed out of the supporting plates.

The invention discloses an engine configuration using liquid hydrogen as a coolant and fuel and an overall design method, and relates to the field of aero-engines, and the engine configuration comprises an air inlet channel, a fan, a precooler, an air compressor, a main combustion chamber, a high-pressure turbine, a low-pressure turbine, a mixing chamber, an afterburner and an exhaust nozzle. Only the inner culvert airflow is cooled, on the basis of solving the difficulty of rotation limitation under the high Mach number working condition of the core engine, the problem of cooling capacity waste generated by precooling the outer culvert airflow is avoided, and the liquid hydrogen amount depends on the combustion amount needed by the main combustion chamber and the afterburner; therefore, the waste problem caused by the fact that the amount of liquid hydrogen needed by pre-cooling is larger than the amount of liquid hydrogen needed by combustion is avoided. According to the overall design of the engine, from the angle of flight / intake / launch integration, the size and performance parameters of an engine body are determined in a ground state based on the take-off task requirement of an aircraft, the size of an intake and exhaust system is determined in a cruise state based on the cruise task requirement of the aircraft, and a control rule is determined by the engine demand flow and the intake passage capture flow.

This invention discloses a high-disturbance spray boom structure based on plasma-excited coupled aerodynamically assisted atomization, relating to the field of afterburner technology for aero-engines. It includes: a shell with an air inlet; a fuel pipe with a fuel delivery channel; the fuel pipe extending into the shell and forming an air channel with it; a first insulating platform with a collision atomization zone and an atomization development zone; the first insulating platform is connected to the output end of the fuel pipe, and its neck has several fuel-gas atomization channels; a second insulating platform with a plasma discharge region; the second insulating platform is snapped into the first insulating platform; a high-voltage electrode located between the first and second insulating platforms; and a grounding electrode located between the second insulating platform and the shell. This invention employs the coupling of plasma excitation with aerodynamically assisted atomization and cross-collision to perform three consecutive actions on the fuel spray, thereby improving the fuel atomization performance of the high-disturbance spray boom.

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 discloses a main combustion chamber and afterburner integrated test platform and method, and belongs to the technical field of aero-engines. The integrated test platform comprises a main combustion chamber test section, a heat exchanger section, a back pressure valve section and an afterburner test section which are connected in sequence; the heat exchanger section comprises a heat exchange flange and a straight and straight mixing section which are coaxially connected; the heat exchange flange is provided with an air inlet pipe, an air collecting ring and a jet hole, external air is introduced into the straight mixing section through the jet hole, and the tail gas of the main combustion chamber and the external air are mixed and cooled and then input into the afterburner test section; the main combustion chamber test section and the afterburner test section are connected with the gas supply system, the oil supply system and the measurement and control system, the heat exchanger section is connected with the gas supply system and the measurement and control system, and the back pressure valve section is connected with the water supply system and used for cooperatively controlling test parameters. Decoupling between working conditions of the main combustion chamber and the afterburner is easier to realize, so that the working condition range of an integrated test of the main combustion chamber and the afterburner is widened.

To provide a cooking device capable of achieving improvement in combustibility of a grill burner and uniform grilling of an object to be cooked while improving a lamp black burning effect by an after burner.SOLUTION: An after burner is attached to the upper side of an exhaust duct in a posture in which a burning surface is directed downward. A control plate 61 is provided in a communication port between a grill chamber and the exhaust duct. The control plate 61 includes an opening part 63, a right closing part 64, a left closing part 65, a right half opening part 66, and a left half opening part 67. The opening part 63 is provided in the center of the control part 61 and opens the communication port. The right closing part 64 and the left closing part 65 are provided on both of the right and left end sides of the control plate 61, and close the communication port. The right half opening part 66 is provided between the right closing part 64 and the opening part 63, and the left half opening part 67 is provided between the left closing part 65 and the opening part 63. The right half opening part 66 and the left half opening part 67 close the lower half of the communication port and open the upper half.SELECTED DRAWING: Figure 9

The invention relates to the field of aero-engine afterburner tests, and discloses a fan-shaped afterburner inlet flow field simulation device and a control method thereof.The fan-shaped afterburner inlet flow field simulation device comprises a linkage ring, a knuckle bearing, a transmission cylinder and a servo electric cylinder, and the telescopic end of the servo electric cylinder stretches out and draws back in the axial direction of an afterburner and is used for driving the transmission cylinder to rotate through stretching out and drawing back; and the corresponding guide blades are driven by the linkage ring and the rocker arms of the other guide blades to rotate along the rotating shafts of the corresponding guide blades. Through the structural design of the servo electric cylinder and the transmission mechanism, the actual inlet flow field state of the fan-shaped afterburner is simulated, and particularly, the angle of the air inlet guide blade can be synchronously adjusted through continuous change, so that the test environment simulation of the fan-shaped afterburner with different air inlet cosine angles is realized; the fan-shaped afterburner inlet airflow cosine angle distribution is obtained, so that the test environment is closer to the real working environment, and the efficient development of the aero-engine afterburner performance test can be supported.

The invention provides a dual-fuel composite pre-cooling engine system based on jet flow / heat exchanger pre-cooling and a using method of the dual-fuel composite pre-cooling engine system, and relates to the field of aircraft engine systems. An air inlet channel, a fuel pump, a precooler, a jet device, an air compressor, a main combustion chamber, a turbine, an afterburner and an exhaust nozzle sequentially form an engine main body in the axial direction. According to the system, the characteristics of jet flow precooling and heat exchanger precooling are combined, liquid ammonia / kerosene dual-fuel layout is adopted, and a cooling working medium can realize conversion of precooling working places through a control valve according to precooling requirements, so that incoming flow air in different flight states is efficiently precooled; and the problems that a heat exchanger is large in precooling weight and flow resistance and a jet device is poor in precooling heat exchange effect are effectively solved. And finally, the system gives consideration to the simple jet flow pre-cooling structure and the high efficiency of heat exchanger pre-cooling, and the pre-cooling effect and the propelling performance of the pre-cooling engine are effectively improved.

The invention discloses a hydrogen and kerosene dual-fuel integrated support plate flame stabilizer, an afterburner and an aero-engine, and belongs to the technical field of aero-engine afterburners, the support plate flame stabilizer comprises a straight support plate body, the straight support plate body is arranged in the radial direction and is perpendicular to a flow path, and the straight support plate body is arranged in the radial direction; a plurality of hydrogen spraying holes and kerosene spraying holes are evenly distributed in the side face of the straight plate supporting plate body in the longitudinal direction, and a kerosene header pipe and a hydrogen header pipe are arranged in the straight plate supporting plate body. According to the hydrogen-kerosene dual-fuel integrated support plate flame stabilizer, the afterburner and the aero-engine, hydrogen-kerosene dual-fuel combustion can be coordinated, the combustion efficiency is improved, the temperature distribution is uniform, the structure is simple, and the weight is light.

The invention discloses a rotating detonation afterburner based on centrifugal nozzle discrete injection, and belongs to the technical field of detonation propulsion. The combustion chamber comprises a gas collecting cavity, an injection unit, a centrifugal nozzle, an outer ring and an inner column, wherein an annular detonation combustion chamber is formed between the inner wall surface of the outer ring and the outer wall surface of the inner column. The centrifugal nozzles are uniformly arranged at the intersection of the gas collection cavity and the injection unit in the circumferential direction to realize primary atomization of the fuel oil; the air is accelerated by the injection unit from the air collection cavity and then is mixed with the atomized fuel oil, secondary atomization is achieved, the particle size of liquid drops is remarkably reduced, and the fuel activity and mixing uniformity are improved. The injection unit adopts a Laval nozzle profile, air is accelerated to supersonic speed in the injection unit, and pressure of a combustion chamber is prevented from being transmitted forwards. The rotary detonation engine is suitable for the high-temperature low-oxygen environment, supports the annular combustion chamber configuration and the empty cylinder combustion chamber configuration, can achieve stable rotary detonation detonation and propagation of liquid fuel and air, and provides technical support for engineering application of the rotary detonation engine.

The invention discloses a thrust augmentation type continuous detonation turbine engine. An afterburner is formed on the rear side of a rectification exhaust tail cone of the engine; the continuous detonation main combustion chamber consists of a shell and a central inner column which are coaxially arranged; a pre-explosion pipe, an air film cooling hole and a cold air mixing hole are formed in the shell; a fuel channel and a secondary air channel are arranged in the central inner column; a fuel swirler and a secondary air hole are arranged on the central inner column; the supporting plates are evenly distributed in the afterburner in the circumferential direction. A secondary inner column, a plane air ring and an annular protruding oil injection ring are arranged at the tail of the supporting plate. The annular protruding oil injection ring injects secondary fuel at high pressure in the axial direction, and the secondary fuel is mixed with tail gas of a combustion chamber for combustion. The plane air ring injects air in the axial direction, so that efficient mixing and combustion of fuel in the inner ring are achieved. According to the engine, a continuous detonation turbine engine and an afterburner are integrally designed, and the purpose of releasing a large amount of energy in a short time is achieved.

The present application belongs to the field of detonation turbine engine design, and is a device for separating detonation and afterburner airflow exhaust, comprising an inner nozzle, an outer nozzle, an inner regulating device, and an outer regulating device; when the detonation turbine engine is working, the high-temperature combustion gas in the afterburner combustion chamber enters the inner nozzle, the high-temperature combustion gas in the detonation combustion chamber enters the high-temperature combustion gas channel, and the outer cooling gas is divided into two streams and enters the first cooling channel and the second cooling channel respectively, cooling the wall surfaces of the outer nozzle and the inner nozzle. At the same time, the outer cooling gas in the first cooling channel and the second cooling channel enters the heat shield, so that the inner regulating device will not come into contact with the high-temperature combustion gas generated by the detonation combustion chamber, and at the same time, it is continuously cooled by the outer cooling gas, thereby minimizing the ambient temperature of the inner regulating device located in the outer high-temperature area, and effectively reducing the risk of ablation of the cylinder wall and the inner regulating device caused by the outer high-temperature area.

The invention belongs to the technical field of aero-engines, and particularly relates to a dual-mode afterburner outer culvert pneumatic area calculation method, which comprises the following steps: aiming at a conventional mode and an outer culvert throttling mode, respectively carrying out circulation section division on an outer culvert gas path of an afterburner; based on three-dimensional numerical simulation, resistance proportions of all sections of the outer culvert channel in the two modes are calculated respectively, and according to the resistance proportions, total pressure distribution of the outer culvert on-way channel is obtained; calculating the flow of each section based on the total pressure distribution and the given initial geometric area of each section, and iterating the final geometric area of each section meeting a flow rationality condition and an outlet static pressure balance condition according to the flow of each section; and according to the final geometric area and the flow of each section, calculating and determining the corresponding total pneumatic area of the external culvert in the two modes.

FIELD: rocket science.SUBSTANCE: invention concerns the design of a ramjet engine. The ramjet engine body is a structure consisting of several layers, including: an internal heat-protection layer made of a composite material based on a thread frame of heat-resistant fibers impregnated with an elastic polymer binder, a power body based on a high-temperature polymer composite material and an external heat-protection layer made of a composite material based on an elastic polymer binder and a finely dispersed mineral filler. Thanks to the use of layers of external and internal thermal protection, it becomes possible to use a polymer composite material for the power casing.EFFECT: reducing the labor intensity of manufacturing and weight of the product and increased energy-mass perfection of the propulsion system.1 cl, 2 dwg

A kind of marine low-pressure selective catalytic reduction denitration system regeneration method utilizes marine engine supercharger front high-temperature exhaust waste heat, to heat up to ammonium bisulfate decomposition temperature point to SCR system, to realize the regeneration of SCR system.Supercharger front high-temperature exhaust gas is introduced into SCR reactor by exhaust gas bypass valve, and the flow of high-temperature exhaust gas can be changed by adjusting the opening of exhaust gas bypass valve according to the volume of catalyst in SCR reactor and the set regeneration temperature.The invention can save the traditional engine external afterburner, significantly reduce the floor area of SCR system, simplify the operation of SCR system, and realize the full utilization of engine exhaust waste heat utilization.

This invention discloses a drag reduction and thrust reversal device, relating to the field of fluid dynamics technology. The invention includes a front drag reduction air intake assembly, an automatic rotary cutting assembly, two fluid delivery pipes, and the drag reduction and thrust reversal assembly. The automatic rotary cutting assembly is mounted on the front drag reduction air intake assembly. In this invention, the drag reduction and thrust reversal device, composed of the front drag reduction air intake assembly, the automatic rotary cutting assembly, the fluid delivery pipes, and the drag reduction and thrust reversal assembly, efficiently captures oncoming fluid through an inclined inlet at the front. A filter baffle at the inlet blocks large debris, which is then shredded and mixed by fluid-driven rotary cutting blades. The mixture is then transported through the fluid pipes to the end delivery pipes. Under the action of the tail jet pressurization device, the fluid forms a vortex and is ejected at high speed, generating reverse thrust. This converts oncoming resistance into propulsive power, directly propelling the ship forward, achieving the dual effects of drag reduction and thrust reversal, and converting drag energy into forward propulsion.

The invention belongs to the technical field of aero-engines, and particularly relates to a dual-mode afterburner outer culvert channel system and a working method, the dual-mode afterburner outer culvert channel system comprises an outer culvert channel outer casing, an outer culvert channel inner casing and an outer culvert channel located between the outer culvert channel outer casing and the outer culvert channel inner casing, and the dual-mode afterburner outer culvert channel system further comprises a throttling device arranged at an inlet of the outer culvert channel and used for adjusting the area of the inlet of the outer culvert channel; the cold air flow dividing device is arranged on a flow dividing window of a casing in the outer culvert channel behind the throttling device and is used for adjusting the area of the flow dividing window; the driving control unit is connected with the throttling device and the cold air flow dividing device and is configured to drive the throttling device and the cold air flow dividing device to be switched between the opening state and the closing state so as to achieve different outer culvert channel circulation areas.

The invention belongs to the field of afterburners of aero-engines, and particularly relates to a cyclone nozzle of an afterburner. The cyclone nozzle comprises a nozzle body, a fuel oil channel and a gas channel are formed in the nozzle body, the fuel oil channel is provided with a fuel oil channel outlet, and a containing cavity located at the fuel oil channel outlet is formed in the nozzle body; the pressure difference floating seat is accommodated in the accommodating cavity; the fuel quantity control valve is fixedly installed on the nozzle body, located in the containing cavity and used for fuel flow control. The nozzle opening is formed in the nozzle main body, and the nozzle opening is communicated with the containing cavity through a nozzle opening channel; the cyclone channel is formed in the nozzle body, an air inlet of the cyclone channel is communicated with the air channel, and an air outlet of the cyclone channel is communicated with the nozzle channel.

The invention relates to the field of aero-engines, and discloses a method for machining a combustion header pipe mounting seat of an afterburner of a certain type of aero-engine, which comprises the following steps of: changing three 8.2 through holes in a triangular base of a part into M8 threaded holes, fixing a workpiece on a tooling plate by using three inside hexagonal bolts, and expanding the M8 threaded holes to 8.2 through holes after milling, so as to meet the clamping of the workpiece; a numerical control spark machine tool is used for machining hexagonal holes with the opposite distance being 4 mm in the end face of the inside hexagonal bolt, two pressing plates are pressed above a tool plate after a workpiece is milled, the inside hexagonal bolt can be taken off without detaching the tool plate from a three-jaw self-centering chuck, and positioning errors caused by secondary clamping are avoided. The process plate adopts a circular structure to facilitate positioning and clamping, a short cylindrical boss is designed above the process plate to avoid over-positioning, the process plate limits the degree of freedom of rotation around the X-axis and the Y-axis and the degree of freedom of movement along the Z-axis, the short cylindrical boss limits the degree of freedom of movement in the X-axis and the Y-axis, over-positioning does not exist, and rotation around the Z-axis is determined through alignment of a dial indicator.

The present invention discloses a turbine afterburner matching system for a small turbojet engine supercharged afterburner combustion chamber. The system comprises a mixing channel and a supercharged annular cavity. The mixing channel communicates with the supercharged annular cavity, and an anti-backflow unit is provided between the mixing channel and the supercharged annular cavity. The anti-backflow unit is used to control the directional flow of fluid through the mixing channel into the supercharged annular cavity. This ensures a stable front-to-rear pressure drop ratio for the turbine in the supercharged afterburner combustion chamber of the small turbojet engine, prevents premature fuel autoignition, and ensures the stability of the turbine's operating state.

The present application belongs to the field of afterburner design, and is an asymmetric afterburner internal step ignition structure, comprising a cone, a support plate and an ignition nozzle; an annular step ignition zone is provided in the middle of the cone, and the step ignition zone comprises a first step zone, a transition zone and a second step zone. During the ignition process of the afterburner, the fuel passes through the fuel nozzle in the first step zone through the casing and the support plate into the interior of the step, and is ignited by the ignition nozzle located at the center of gravity of the first step zone. The oil-gas mixture is first ignited and burned in the first step zone. Since the step angle of the first step zone is 90°, it can produce an obvious sudden expansion effect on the airflow, thereby constructing a stable recirculation zone, and at the same time has excellent flame propagation characteristics. The flame quickly propagates to the transition zone and the second step zone. The setting of the transition zone can make the vortex formed in the first step zone during the flame transfer more stable and will not fall off, thereby realizing fast and reliable ignition of the afterburner.

The invention belongs to the technical field of aero-engines, and particularly relates to a central pre-combustion chamber with an afterburning nozzle, and the central pre-combustion chamber is arranged in a central cone of an afterburner. The pre-combustion chamber forms a stable backflow area through a head swirler, and fuel oil is ignited by an ignition plug to generate high-temperature fuel gas. An innovative fuel gas shunting structure comprises a center shunting cone and a plurality of radial flame transfer pipes which are distributed in the circumferential direction, high-temperature fuel gas which is combusted in a centralized mode in the center is effectively shunted and guided to a backflow area of each radial supporting plate stabilizer of the afterburner, and fuel oil is added through an afterburning nozzle. The high-temperature fuel gas jet flow ejected by the radial flame transfer tube 5 of the central pre-combustion chamber is high in temperature and large in flow, the ignition energy of the high-temperature fuel gas jet flow is far larger than the spark energy generated by a conventional sparking plug, and the problems that the afterburner 1 is insufficient in ignition reliability and narrow in ignition boundary under high-temperature gas inlet can be solved.

The present invention discloses a combined atomization device with a single oil circuit and dual fuel supply modes, which includes a main combustion stage oil pipe, a pre-combustion stage oil pipe, a centrifugal pressure nozzle, a main combustion stage oil circuit switch conversion mechanism, a pre-combustion stage oil circuit switch conversion mechanism, and a direct injection nozzle. The top end of the main combustion stage oil pipe is integrally connected to the fuel supply pipe, and the bottom end of the main combustion stage oil pipe is connected to the wall surface of the centrifugal pressure nozzle. The pre-combustion stage oil pipe is sleeved inside the main combustion stage oil pipe, and the bottom end of the pre-combustion stage oil pipe is connected to the inlet of the centrifugal pressure nozzle. An annular main combustion stage oil circuit is formed between the main combustion stage oil pipe and the pre-combustion stage oil pipe, and a pre-combustion stage oil circuit is formed inside the pre-combustion stage oil pipe. The main combustion stage oil circuit switch conversion mechanism is installed at the top end of the main combustion stage oil circuit; the pre-combustion stage oil circuit switch conversion mechanism is installed at the top end of the pre-combustion stage oil circuit; the direct injection nozzle is installed on the pipe wall of the main combustion stage oil pipe and is communicated with the main combustion stage oil circuit. The present invention can realize the conversion of two fuel atomization modes in the afterburner, and achieve ignition and stable combustion in the afterburner.

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.

The invention relates to an aero-engine afterburner self-adaptive adjustable blade which is used for being arranged between a turbine of an aero-engine and an afterburner, the aero-engine afterburner self-adaptive adjustable blade comprises a plurality of blades, each blade comprises a main body and a stabilizer assembly, and the stabilizer assembly is connected to the tail of the main body; the multiple blades are circumferentially arranged around the central axis of the aero-engine, and in every two adjacent blades in the circumferential direction, a main body of the blade located on the upstream and a main body of the blade located on the downstream are at least partially overlapped; the stabilizer assembly is configured to operably perform unfolding movement or folding movement, so that the blades are in an expanded state or a contracted state; the multiple blades are configured to be capable of synchronously switching between the expansion state and the contraction state in response to the change of the working condition of the engine.

The application provides a two-stage annular mixer of a booster combustion chamber and a bypass ratio self-adaptive matching method, and relates to the technical field of an aero-engine.The two-stage annular mixer of the booster combustion chamber comprises a casing, a first-stage annular mixer arranged in the casing, an outer bypass formed between the first-stage annular mixer and the casing, a second-stage annular mixer arranged in the casing and located inside the first-stage annular mixer, a split flow channel formed between the first-stage annular mixer and the second-stage annular mixer, a mixing channel formed by the casing and the second-stage annular mixer, and the outer bypass and the split flow channel are communicated with the mixing channel.In the embodiment of the application, a small mixer mixing effect is realized under a small overflow loss and a flow separation risk, a low cooling gas temperature is ensured, and a low flow resistance and high-efficiency stable working requirement of a wide outer / inner bypass range can be met.

To provide a cooling appliance capable of suppressing the waste of gas in an afterburner.SOLUTION: A gas stove comprises a grill device and a distribution device 80. The grill device comprises a grill chamber, an exhaust duct, and an afterburner 7. The grill chamber comprises an upper burner 21 and lower burners 23, 24. The exhaust duct communicates with the rear part of the grill chamber, and discharges combustion exhaust gas flowing from inside of the grill chamber to the outside. The afterburner 7 is attached into the exhaust duct. The distribution device 80 distributes gas to the upper burner 21, the lower burners 23, 24, and the afterburner 7, and links firepower of the afterburner 7 to firepower of the upper burner 21.SELECTED DRAWING: Figure 5

The application relates to a radar stealth control structure of a radial flame channel of an aero-engine afterburner, which comprises the following: a plurality of radial flame channels, which are arranged in a circumferential direction in a V-shaped cross section in an outer wall of the afterburner and are connected to trailing edges of an outer annular flame stabilizer and an inner annular flame stabilizer; a plurality of pairs of baffles, which are hinged to two side walls of each radial flame channel; a plurality of driving rods, one end of each of which extends into each radial flame channel from an outer end; a plurality of pairs of hinged rods, which are arranged in each radial flame channel and are hinged between each pair of baffles and each driving rod; a linkage ring, which is sleeved around an outer periphery of the outer annular flame stabilizer, is connected to the other end of each driving rod, can move in an axial direction of the outer wall of the afterburner, and then drives each pair of baffles to close or open through each driving rod and each pair of hinged rods; when each pair of baffles is closed, an included angle exists therebetween; and when each pair of baffles is opened, the two side walls of each radial flame channel are flattened.

This application belongs to the field of aero-engine technology, and specifically relates to a double-layer structure afterburner heat insulation plate. The heat insulation plate is installed between the inner cone and the rear bearing cavity, isolating the rear bearing cavity from the inner cone's internal cavity. The heat insulation plate includes an inner layer, an outer layer, a skin, and a mounting edge. The mounting edge is located at the circumferential outer edge of the inner layer and has multiple circumferentially distributed through holes. The mounting edge is installed on the mounting edge of the bottom surface of the inner cone through these through holes. The outer layer is attached to the surface of the inner layer near the inner cone. The inner layer includes a spherical surface protruding into the inner cone's internal cavity. The afterburner heat insulation plate has a double-layer structure, and the cavity height is maintained by fixed supports and rivets. Heat insulation cotton is added to the double-layer cavity to enhance the heat insulation effect of the heat insulation plate.

The invention belongs to the field of afterburners of aero-engines, and particularly relates to an afterburner oil supply device. The oil supply device comprises an oil supply main body, a fuel oil flow channel is formed in the oil supply main body and comprises two or more fuel oil flow channel sections which are sequentially communicated, an inlet of one fuel oil flow channel section is communicated with an oil supply pipeline, a plurality of oil spraying holes are formed in the oil supply main body, and the oil spraying holes are communicated with the fuel oil flow channel sections. The fuel injection holes are communicated with the corresponding fuel runner sections; the fuel oil supply assemblies are in one-to-one correspondence with the fuel oil flow channel sections, and all the fuel oil supply assemblies are arranged in the corresponding fuel oil flow channel sections respectively. The radial partition oil supply and combustion requirements of the afterburner can be met, and better afterburner oil-gas mixing and combustion effects are provided; and the structure design is more flexible, and finer fuel injection control and a more accurate oil-gas matching effect can be achieved.