109 results about "Afterburner" patented technology

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 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 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 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.

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 relates to an aero-engine afterburner test technology, in particular to a method for measuring a 3D3C flow field of an afterburner stabilizer, and aims to solve the problem that a three-dimensional flow field of a complex flow channel of the stabilizer cannot be accurately represented by traditional two-dimensional velocity field (2D2C) measurement. An adopted measuring device comprises PIV equipment, and a front-section cylinder body, an optical measuring-section cylinder body and a rear-section cylinder body which are connected in sequence; an afterburner stabilizer is mounted in the optical measurement section cylinder to serve as a tested piece; airflow sequentially flows through the front-section cylinder body, the optical measurement section cylinder body and the rear-section cylinder body, and flow field characteristics formed when the airflow flows through a tested piece arranged in the optical measurement section cylinder body are measured by the externally arranged PIV equipment. According to the invention, 3D3C flow field measurement is realized, flow field characteristics can be comprehensively captured, the operation difficulty is reduced, data processing is accelerated, the test time is shortened, and the method has practical application value.

The invention belongs to the technical field of aero-engines, and particularly relates to an outlet gas sampling device and method for afterburner performance test, the working condition of an engine is set, and a gas analysis test section is arranged at the downstream of a tail nozzle (2) of an afterburner (1); setting the angular position of the rotary displacement mechanism (5); only the mixed sampling pipeline (8) is opened to complete the combustion efficiency test of the afterburner; only the single-point sampling pipeline (10) is opened to complete an outlet temperature test; the circumferential angle of the rotary displacement mechanism (5) is replaced, and the step 3 is executed till all the angles are completed; and adjusting the working condition of the engine, and skipping to the step 1 until all working conditions are completed.

This invention discloses a porous rectifying structure for the afterburner of a variable cycle engine, comprising a rectifying orifice plate installed in an annular channel between the turbine outlet and the flame stabilizer. The orifice plate is symmetrically arranged along its radial central axis, with the clogging area facing the turbine inflow and axially aligned with the heat shield inlet. The orifice plate height is 2 / 3 of the channel height between the central cone and the heat shield. The plate has uniformly distributed pores with an equivalent diameter of 10-20 mm along the radial direction, with a clogging ratio of 5%-25%. This invention forms a multi-stage mixing interface through the radially gradient porosity pore arrangement, promoting shear mixing of the high-temperature internal combustion gas and the external bypass gas. Simultaneously, the static pressure difference across the orifice plate drives the external bypass gas flow into the heat shield cooling channel, increasing the cooling flow rate, improving the uniformity of the upstream velocity field of the flame stabilizer, eliminating the recirculation zone inside the heat shield, reducing the thermal load on the casing wall, and controlling the total pressure loss to within 3%, achieving high-stability combustion and reliable thermal protection under a wide bypass ratio.

The invention belongs to the technical field of aero-engines, and particularly relates to a method for obtaining the influence of the nozzle area and the stress application fuel flow on the drop pressure ratio, and the method comprises the steps: carrying out the test of an engine in a training state, locking stress application fuel, independently controlling the vertical step of the nozzle area, and obtaining a first model of the nozzle area on the drop pressure ratio; locking the area of the nozzle, independently controlling the up-down step of the stress application oil, and obtaining a second model of the stress application oil to the drop pressure ratio; identifying the first model and the second model to obtain a first transfer function taking a nozzle area as an input variable and taking a drop pressure ratio as an output variable, and a second transfer function taking afterburner fuel as an interference variable and taking the drop pressure ratio as the output variable; and comparing the first transfer function with the second transfer function to obtain a first influence coefficient of the thrust augmentation fuel oil or the nozzle area.

This invention discloses a pneumatic concave cavity flame stabilization device for an afterburner, comprising a central cylinder, an annular combustion chamber shell, a diffuser, an afterburner shell, a pneumatic concave cavity pipe, and a central cylinder connecting rod. The diffuser fixes the annular combustion chamber shell and the afterburner shell together. The central cylinder is fixed to the center of the annular combustion chamber shell via the central cylinder connecting rod. The pneumatic concave cavity pipe extends through the central cylinder into the afterburner shell. The vertical jet from the pneumatic concave cavity exhaust port of this invention can enhance the recirculation zone and stably maintain a low-speed recirculation vortex under different operating conditions, stabilizing the combustion operating range and reducing the risk of flameout. The vertical jet from the pneumatic concave cavity exhaust port of this invention can entrain fuel and air into the cavity, prolonging the fuel residence time and fundamentally improving combustion efficiency. The vertical jet from the pneumatic concave cavity exhaust port of this invention can fill the low-pressure zone at the rear edge of the cavity, reducing resistance, significantly reducing total pressure loss, and improving engine thrust efficiency.

The application discloses a nozzle afterburner regulator timing abnormality exit fault detection system and disposal method, comprising an afterburner fuel test closed-loop control system and an oil pressure measuring mechanism, the oil pressure measuring mechanism is used for collecting the front and rear oil pressures of a metering valve, the rear oil pressure of a constant pressure valve and the rear fuel oil pressure of an afterburner locking valve of the nozzle afterburner regulator in a current operation cycle. The application has the beneficial effect that: under the condition of no test driving, the emergency afterburner three-zone timing abnormality fault of an engine caused by the nozzle afterburner regulator can be reproduced quickly and effectively, so compared with the traditional test platform, the system does not need test driving and does not need to build the huge test system required for test driving, not only can effectively monitor and evaluate the speed regulation quality of the fuel regulator, but also improves the work efficiency, reduces the fuel consumption, thereby reduces the detection cost, saves a large amount of manpower and material resources, and has significant economic benefits.

This invention discloses an afterburning fuel control method for multi-electric aero-engines, belonging to the field of aerospace engine control technology. Addressing the engine control instability caused by the dead-zone characteristics of the afterburning fuel actuator during experiments with multi-electric aero-engines using electric pumps for fuel metering, this invention designs an afterburning fuel closed-loop control loop based on a neural network inverse model and an improved μ-correction adaptive control method. This invention also discloses an afterburning fuel control device for multi-electric aero-engines. Simulation results show that the control method of this invention can effectively avoid the dead-zone region of the actuator and achieve a smooth transition of engine state throughout the entire operating range of the afterburning fuel actuator, effectively improving the stability of the control system. Furthermore, the improved μ-correction adaptive control algorithm proposed in this control method has better dynamic characteristics than the traditional μ-correction adaptive control algorithm.

The invention discloses a possibility-based optimization design method for an aero-engine thrust augmentation cylinder, which comprises the following steps of: firstly, analyzing and identifying an important region of interest of the thrust augmentation cylinder, cutting the region by utilizing a sub-model technology, and realizing parametric modeling and finite element modeling; calculating the low-cycle fatigue life of the position with the maximum stress by adopting a local stress-strain method; the method comprises the following steps: representing geometric configuration, material attributes and load environment parameters of a stress application cylinder as fuzzy variables, and carrying out reliability analysis through a method of combining fuzzy simulation and adaptive Kriging to obtain a failure possibility degree; and finally, by taking the minimum volume of the stress application cylinder as a target and the design value of the failure possibility degree as a constraint, solving an optimization design model based on the failure possibility degree by adopting a method based on an enhanced expectation improvement learning function and combining an active constraint criterion and an adaptive Kriging model, so as to obtain design parameters of the stress application cylinder. According to the method, the problem that a traditional optimization algorithm is poor in adaptability under multi-constraint and fuzzy uncertainty is solved.

The invention relates to a turbofan engine afterburner organization combustion oil injection stability integrated test device which comprises a tester and a test piece. The tester is rectangular; rectangular observation windows are formed in the upper portion and the two side walls of the tester, visual glass is arranged on the outer sides of the observation windows, the visual glass is pressed on the side walls of the tester through rectangular glass frames, and the glass frames are connected to the side walls of the tester through screws. An ignition electric nozzle mounting opening is formed in the side wall of one side of the tester, an ignition electric nozzle mounting plate is connected to the side wall of the side of the tester in a sliding mode, a mounting hole used for mounting an ignition electric nozzle is formed in the ignition electric nozzle mounting plate, and the ignition electric nozzle mounting plate can drive the ignition electric nozzle to slide in the runner of the tester in the axial direction; the test piece comprises a positioning mounting plate, a ventilation and spray rod assembly, a stabilizer assembly and a flame transfer groove assembly; the test piece is connected above the tester through the positioning mounting plate, and the ventilation and spray rod assembly, the stabilizer assembly and the flame transfer groove assembly extend into a flow channel of the tester.

The invention relates to the technical field of aero-engines, in particular to a fixing structure of an afterburner stabilizer which comprises a base body and a plurality of stabilizer bodies arranged on the base body. The fixing structure comprises a first hinge seat, a first connecting rod and a cartridge receiver cylinder; the first hinge seat is arranged at the top end of the stabilizer body; the lower end of the first connecting rod is hinged to the first hinge seat, the upper end of the first connecting rod is arranged to be a disc-shaped body, and a first threaded rod is integrally formed in the center of the top face of the disc-shaped body. The first threaded rod is sequentially provided with a first gasket and a first nut after penetrating through the rear cartridge receiver cylinder; the disc-shaped body is used for being supported on the inner wall face of a cartridge receiver barrel, and the first gasket is used for being pressed on the outer wall face of the cartridge receiver barrel. The stabilizer body is detachably connected with a cartridge receiver cylinder through the first connecting rod, the disc-shaped body on the top of the first connecting rod, the first threaded rod, the first gasket and the first nut.

The invention belongs to the technical field of aero-engines, and particularly relates to a central pre-combustion chamber with a fuel gas shunting structure, 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 backflow areas of all radial supporting plate stabilizers of the afterburner, and therefore comprehensive and reliable ignition is achieved.

Engine (1) for aircraft, comprising several shafts (2) and a gas turbine (3) with a compressor (4) for compressing gas and a turbine (5), wherein the compressor (4) and the turbine (5) are rotationally fixed to the at least one shaft (2), and a combustion chamber (6), further comprising at least one fuel cell (7), wherein the gas turbine (3) and the at least one fuel cell (7) are connected to each other in such a way that, in the operating state, compressed gas (15) from the gas turbine (3) can be supplied to the at least one fuel cell (7), wherein it has at least one electric motor, wherein the fuel cell (7) is arranged for its energy supply and wherein the electric motor is arranged for at least indirectly driving the shafts (2), wherein each shaft (2) can be driven at least indirectly exclusively by an electric motor. wherein the fuel cell (7) has a reformer (9) for supplying the fuel cell (7) with fuel, and wherein the engine has an afterburner which is designed separately from the combustion chamber (6) and is configured to burn unburned residual fuel (16) of the fuel cell (7) as well as a portion of a fuel not converted by the reformer (9) and to supply the exhaust gases of the afterburner to the gas turbine (3) after the combustion chamber (6).

This invention discloses a concave cavity flame stabilizer structure, its fuel injection and cooling method, and an afterburner. The structure includes a symmetrical Tesla valve, a fluid oscillator, and shape memory alloy ribs. The Tesla valve is located at the front end of the evaporation chamber of the afterburner, utilizing its unidirectional flow resistance characteristics to suppress fuel backflow caused by pressure fluctuations and ensure stable fuel supply. The fluid oscillator replaces the traditional injection hole, converting the fuel-air mixture into a high-frequency, large-scale periodic swept jet injected into the concave cavity, improving atomization, evaporation, and mixing efficiency. The shape memory alloy ribs are located on the cold side of the bottom and rear wall of the concave cavity, automatically adjusting their shape according to changes in wall temperature to achieve dynamic matching between cooling intensity and heat load. This invention, through the system integration of fuel supply, combustion, and cooling, solves the problems of fuel backflow, low atomization efficiency, and poor adaptability of the cooling system under wide operating conditions in traditional concave cavity structures, significantly improving the combustion efficiency and thermal protection reliability of the afterburner.

The present application belongs to the technical field of afterburner fuel flow regulation, and relates to a self-adaptive fuel flow regulation device for a multi-hole self-excited sweeping fuel injection rod, comprising a nozzle connecting part and a flow sensing and regulating part. The nozzle connecting part comprises a fixed fuel baffle, a top connecting strip, a bottom connecting strip and a side connecting strip. The fixed fuel baffle is an arc-shaped long strip plate body, and a concave elastic groove is arranged on one side of the long strip plate body. The flow sensing and regulating part is rotatably installed in the concave elastic groove. The flow sensing and regulating part comprises a long strip-shaped movable fuel baffle, and an inside flow sensing tab and an outside flow sensing tab are welded on the inside and outside of the movable fuel baffle. When fuel flows through the self-adaptive fuel flow regulation device, a pressure difference is generated on both sides, which drives the movable fuel baffle to move inward, and the fuel flow size of the corresponding fuel injection hole is self-adaptively regulated, so that the fuel flow uniformity of each fuel injection hole is ensured, the fuel atomization degree is ensured, and the combustion efficiency of the afterburner is finally improved.

The invention provides a variable-cycle afterburner and a variable-cycle engine thereof. An inner culvert channel is located in the center area of the afterburner and used for guiding low-pressure turbine inner culvert airflow; the outer culvert channel surrounds the outer side of the afterburner and is used for guiding fan outer culvert airflow, and the outer culvert channel and the afterburner converge in a mixing area located in the rear middle section of the afterburner. The supporting plate assembly is arranged in an inner channel of the mixing area, and a concave cavity-step combined structure is arranged at the joint of the lower portion of the supporting plate assembly and the center cone and used for preliminarily inducing airflow to change a track to generate a vortex system and breaking the layering phenomenon. The mixing reinforcing structure is a lobe array, is arranged at the shunting ring of the outer culvert and is positioned at the downstream of the supporting plate, the inner side edge of a lobe is connected with the outer wall surface of the inner culvert, and the outer side edge of the lobe is connected with the inner wall surface of the outer culvert to form a transition structure for communicating the inner culvert and the outer culvert; and single / double external culvert mode switching is realized, so that the mixing efficiency is dynamically optimized, and the overall thermal mixing performance is improved.

The invention relates to the technical field of aero-engines, and discloses an adjustable mixing stabilizer structure, a control method and an applied afterburner. The adjustable mixing stabilizer structure comprises a stabilizer main body, an air entraining cover plate and a first driving mechanism; when the duct is large, the first driving mechanism is controlled to drive the air entraining cover plate to rotate along the hinge shaft, so that the throttling area of the mixing channel is increased, the flow of cooling air passing through the mixing channel is increased, the flow of the cooling air of the heat insulation screen is reduced, the content of oxygen in the duct is increased, and the combustion efficiency of the afterburner is improved; when the duct is small, the first driving mechanism is controlled to drive the air entraining cover plate to rotate along the hinge shaft so as to reduce the throttling area of the mixing channel, so that the flow of cooling air passing through the mixing channel is reduced, the flow of cooling air of the heat screen is increased, the flow of cooling air is guaranteed, and it is guaranteed that the relative cold air amount entering the heat screen is within a stable and controllable range; and safe operation requirements of the afterburner are met.

The application discloses an integrated afterburning scheme based on a high-ejection-speed fan-shaped nozzle and belongs to the field of aero-engine combustion.The high-ejection-speed fan-shaped nozzle is used to spray fuel into an afterburning chamber in a fan-shaped oil mist surface shape along a branch plate in a radial direction, so that the radial space uniformity of fuel behind the branch plate is improved.A large-area flame surface profile is formed behind the branch plate, a larger physical space is provided for chemical reaction, and the fuel particles are rapidly combusted, so that higher combustion efficiency is obtained.The high-ejection-speed fan-shaped nozzle is applied to the combustion scheme on both sides of an integrated branch plate flame stabilizer.Under the premise that the branch plate blockage ratio is not changed, new flow loss is not introduced, and the overall performance in a non-afterburning state is not reduced, the combustion efficiency of the integrated afterburning chamber is improved by improving a fuel supply strategy from the perspective of organizing combustion, so the application has great application prospect.

This invention relates to the field of aero-engine technology, and discloses an inlet swirl adjustment device and its design method for an afterburner test component. A flow divider ring is used to segment the test flow channel into an inner flow channel and an outer bypass flow channel for simulating the airflow in the inner flow channel. The remaining linkage control structures are installed in the outer bypass casing or outer bypass flow channel to reduce interference with the inner flow channel. The linkage adjustment assembly performs overall synchronous adjustment of all adjustable guide vanes, ultimately achieving adjustment of the inner airflow angle, thereby simulating the inner airflow of the afterburner and reducing testing costs and overall engine verification risks. The inlet airflow swirl adjustment device of this invention comprehensively considers the relevant requirements of total pressure recovery coefficient and airflow uniformity index, ensuring high-precision and high-stability simulation of the inner airflow of the afterburner. Furthermore, its simple structure allows for easy assembly and replacement, providing flexible support for afterburner performance testing under different operating conditions.

The application provides a rotating detonation afterburner of an aero-engine, and relates to the technical field of gas turbine engines, and specifically comprises a rotating detonation afterburner connected to an exhaust end of a casing of the engine, an air inlet of the rotating detonation afterburner is connected to an exhaust port of the casing, and an airflow injection channel is arranged on the exhaust end of the casing and / or the air inlet of the rotating detonation afterburner, the airflow injection channel is connected to an internal space and an external space of the air inlet of the rotating detonation afterburner, and when exhaust gas of the exhaust end of the engine passes through the air inlet of the rotating detonation afterburner, the air in the external space of the connection part between the rotating detonation afterburner and the casing is introduced into the rotating detonation afterburner through the airflow injection channel. Through the processing scheme, the stability of the detonation wave is stabilized, and the stability of the rotating detonation afterburner is improved.