151 results about "Ramjet" patented technology

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

The invention provides a high-Mach-number boron powder ramjet engine based on air fluidization. The high-Mach-number boron powder ramjet engine comprises a boron powder conveying module and a kerosene conveying module. The boron powder conveying module can convey boron powder to the combustion part; the kerosene conveying module can convey kerosene to the combustion part; and the combustion part is provided with a stable combustion support plate. The problems of boron powder conveying and ignition stable combustion of the high-Mach-number boron powder ramjet engine based on air fluidization are solved, and incoming flow air is directly used as driving gas and fluidizing gas to fluidize boron powder fuel; kerosene is combusted, and injected fluidized air generates a high-temperature low-speed oxygen-enriched area behind the stable combustion support plate, so that removal of an oxide layer on the surface of the boron powder fuel is accelerated, and the combustion effect of the boron powder fuel is improved. The problem that in the prior art, a boron powder fuel conveying system is too complex to be integrated with an engine system is solved, the defects that reliable ignition and stable combustion are difficult to achieve due to the poor ignition combustion characteristic of boron powder fuel are overcome, and high integration of the boron powder fuel conveying system and the engine and ignition stable combustion of the boron powder fuel are achieved.

The application discloses a strut plate embedded plasma igniter suitable for a scramjet engine and relates to the technical field of the scramjet engine.The strut plate embedded plasma igniter comprises a strut plate main body, and fuel main flow channels, oxygen main flow channels and an embedded channel are arranged on the strut plate main body; a plurality of fuel injection holes are arranged on the side wall of the fuel main flow channel; the oxygen main flow channel is in communication with the embedded channel, and an anode core is arranged in the embedded channel; a detachable head is arranged at one end of the embedded channel facing a combustion chamber, and a plurality of communication holes are arranged on the detachable head. A circular axis-symmetrical hypersonic combustion chamber is connected to the outside of the strut plate main body. The strut plate embedded plasma igniter suitable for the scramjet engine has the advantages of a stable combustion strut plate and a plasma igniter, has the characteristics of good insulation and partial stable combustion, and can ensure stable ignition of the axis-symmetrical supersonic combustion chamber.

The invention discloses a two-stage supporting plate injector applied to a wide-working-condition scramjet engine combustion chamber, belongs to the technical field of scramjet engines, solves the problem that an existing alternating supporting plate is low in fuel and air mixing efficiency under the conditions of a high Mach number and a high equivalence ratio, and comprises a supporting plate body. One end of the supporting plate body is a wedge-shaped tail edge formed by arranging a plurality of wedge-shaped slopes in a staggered mode, and first-stage fuel oil spraying holes are formed in the staggered points of the wedge-shaped slopes in the wedge-shaped tail edge. The two side faces of the supporting plate body are symmetrically provided with a plurality of second-stage fuel oil spraying holes inclined in the wedge-shaped tail edge direction. The first-stage fuel injection hole and the second-stage fuel injection hole form a two-stage fuel injection system, the fuel input mode is changed fundamentally, and the fuel flow ratio of the two-stage fuel injection system is dynamically adjusted according to different inlet Mach numbers and equivalence ratios, so that the two-stage support plate injector can be used in various working conditions, and the service life of the two-stage support plate injector is prolonged. And the optimal mixing efficiency can be realized through the optimal fuel distribution strategy.

The invention relates to a device and a method for measuring and correcting thrust of a powder fuel ramjet engine. The device comprises a fixed rack, a heater, an engine, a slide rail and a bearing pier, the heater works, the cold-state thrust of the heater is measured through the working force sensor, after the chamber pressure fluctuation of the heater does not exceed + / -5%, the engine starts to work, and the hot-state thrust Fhot is measured through the working force sensor; when the heater and the engine are disconnected, the heater works independently; the thrust F'hot is measured through the working force sensor, and the value of the engine thrust absolute value F. According to the method, the thrust when the heater and the engine work at the same time and the thrust when only the heater works are measured respectively, and the on-way pressure data are combined, so that the engine thrust absolute value F is obtained. Measurement of absolute values of thrust and friction force of the powder ramjet engine based on the direct connection type ramjet experiment table can be realized.

The invention discloses a distributed discharge-bypass pressurization coupled wide-area ramjet engine and a working method thereof, and belongs to the field of wide-area ramjet engines, the distributed discharge-bypass pressurization coupled wide-area ramjet engine comprises a bypass pressurization flow channel and a dual-mode ramjet combustion chamber flow channel, and the dual-mode ramjet combustion chamber flow channel comprises a fixed geometric air inlet channel; the bypass supercharging flow channel comprises a distributed drainage hole communicating with the compression face side of the fixed geometric air inlet channel, and a drainage cavity, a first Laval nozzle, a bypass supercharging chamber, a second Laval nozzle and a bypass exhaust nozzle which communicate with the distributed drainage hole in sequence. According to the distributed discharge-bypass supercharging coupled wide-area ramjet engine and the working method thereof, by arranging the bypass supercharging flow channel, the problem of structure and control complexity and difficulty caused by adjustable parts in integrated adjustment of the engine is solved, meanwhile, the starting capacity and thrust gain under the low Mach number are increased, and the starting efficiency is improved. And the engine has the wide-speed-range working capacity.

The application discloses a combustion chamber thrust control method in the acceleration process of a scramjet engine, and adopts a gradually expanding series connection cavity combustion chamber, fuel injection holes are arranged on the upstream of two cavities, the first cavity is a pre-combustion zone, and the second cavity is a main combustion zone; the thrust control method is as follows: when the combustion chamber inlet Mach number is low, the combustion mode of the pre-combustion zone is controlled to be kept in a weak subsonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode; when the combustion chamber inlet Mach number is high, the combustion mode of the pre-combustion zone is controlled to be kept in a supersonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode. The application is applied to the field of scramjet engines, can guarantee that the combustion chamber can generate higher thrust in a wide flight Mach number range, simultaneously avoids thermal choking, and the thrust mutation caused by combustion mode conversion is smaller, thrust control is more stable, and smooth transition of thrust during combustion mode conversion can be realized.

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

This application relates to a fuel injection device and method for matching non-uniform incoming flow, and a ramjet engine. The device includes injectors disposed on the combustion chamber wall, comprising an adjustable injector group and a non-adjustable injector group; the adjustable and non-adjustable injector groups are arranged intersectingly along the circumference of the combustion chamber wall; by using a fixed pressure and a fixed nozzle area, the non-adjustable injector group continuously injects a fixed proportion of fuel into the combustion chamber; by adjusting the injection pressure and nozzle area in real time, the adjustable injector group injects a calculated proportion of fuel into the combustion chamber. This solution can match non-uniform incoming flow at the combustion chamber inlet within a wider flight envelope, improving fuel combustion efficiency and enhancing the performance of the ramjet engine.

This invention discloses a method and system for gas-thermal coupling control of multiple fuel cell stacks in a rotary-wing unmanned aerial vehicle (UAV), relating to the fields of UAV power systems and fuel cell thermal management technology. The system captures the rotor downwash airflow as a source by setting an integrated annular ramjet inlet on each arm, and supplies it to distributed fuel cell stacks after adjustment by adaptive valves and ejectors. Exhaust gases from each stack converge into a central thermal energy aggregation chamber and can be drawn back by ejectors via branch lines. The method intelligently switches operating modes based on stack temperature: at low temperatures, a thermal circulation mode is activated, utilizing waste heat from the exhaust gas to preheat and humidify the intake air; at normal / high temperatures, it switches to a forced exhaust mode for rapid heat dissipation. This invention utilizes rotor airflow to achieve heat dissipation and intake with zero parasitic power consumption, solving the problems of high power consumption, difficult low-temperature start-up, and large temperature differences among multiple stacks in traditional systems, significantly improving the environmental adaptability and endurance of the UAV.

The present application belongs to the technical field of ramjet engine, and discloses a composite supporting nacelle for air intake and exhaust and a design method thereof. The load bearing base of the multifunctional strong damping supporting body of the composite supporting nacelle is located on the central axis of the reverse thrust fairing. The load bearing base is in a herringbone structure, the front section is fixed through a front elastic support, and the rear section is fixed through a rear support. The front end of the load bearing base is provided with a ring-shaped high-temperature exhaust collector, and four composite curved exhaust pipes are arranged in sequence along the circumference of the middle section, and each composite curved exhaust pipe is provided with a corresponding outlet on the reverse thrust fairing. The design method comprises the following steps: designing the overall technical scheme; designing the multifunctional strong damping supporting body; designing the composite curved exhaust pipe; and designing the reverse thrust fairing. The composite supporting nacelle for air intake and exhaust and the design method thereof can meet the urgent requirement of integrating high-temperature exhaust and normal-temperature air intake in the design of small gas turbine engines, and have engineering practical value.

The invention discloses a water ramjet injector, which is characterized in that a core component is a centrifugal two-stage atomizing nozzle, and a two-stage duct design is adopted in the centrifugal two-stage atomizing nozzle: the first stage is a communicating cavity, high-pressure water flow enters the first atomizing cavity and then is accelerated by the communicating cavity, a strong shearing effect is generated due to sudden change of the aperture, and high-Reynolds-number turbulent flow, namely first-stage atomization, is formed; and then the fluid enters the second atomization cavity and interacts with the gas medium at an outlet, and the liquid film is further crushed into micron-sized fog drops, namely second-stage atomization. The nozzles are restrained to the inner wall of the combustion chamber through threads in a circumferential axial symmetry mode, and the jet flow axes and angles of the multiple nozzles are reasonably arranged, so that atomized water media achieve self-impact collision in the center area of the combustion chamber. In the collision process, through momentum accumulation and turbulence disturbance, liquid drops are promoted to be subjected to secondary crushing, the atomization uniformity and the mixing efficiency are remarkably improved, and the method is suitable for large-scale industrial use and popularization.

The application discloses a single-double-flow switching type binary hypersonic variable-geometry inlet and a switching method thereof. The inlet comprises an outer compression section, an inner compression section, a splitter plate and a divergent section. The splitter plate is arranged between a lip cover and a lower surface of the divergent section, is parallel to the lower end surface of the lip cover, has a length equal to that of the divergent section, and divides the divergent section into two splitter flow channels. The lower surface of the inner compression section is rotationally connected to one end of the surface of the outer compression section located downstream. A locking device is arranged on a rotating shaft. During operation, the lower surface of the inner compression section and the surface of the outer compression section are locked when the lower surface of the inner compression section is smoothly connected to the lower surface of the divergent section or the splitter plate downstream, so that the two surfaces cannot rotate relative to each other. The single-double-flow switching type binary hypersonic variable-geometry inlet is matched with two different working modes of a downstream ramjet and a scramjet, the height of a throat of the inlet can be controlled, and the inlet designed at a high Mach number can be smoothly started at a low Mach number.

The application discloses a shaft symmetry double-S-bend stamping engine air inlet channel, which comprises an air inlet channel center body and an air inlet channel lip cover installed outside the air inlet channel center body; an air inlet channel inner passage is arranged between the air inlet channel center body and the air inlet channel lip cover, the air inlet channel inner passage is in a double-S-bend shape and comprises an S1 section and an S2 section which are smoothly and transitionally connected, the starting point and the ending point of the center line of the S1 section and the S2 section are both tangentially horizontal, the S1 section and the S2 section have different areas and different center line change rules, and the air inlet channel throat is always located at the S2 section when the air inlet channel lip cover moves forward and backward. The air inlet channel lip cover is moved to improve the flow capture capacity of the air inlet channel under the low Mach number condition on the premise of ensuring the high-efficiency compression and stable work of the air inlet channel, the required adjusting mechanism is simple, and the air inlet channel has high realizability and operability. The air inlet channel can effectively enhance the anti-back pressure capacity, reduce the self-starting Mach number of the air inlet channel and meet the aerodynamic performance requirements of the engine on the air inlet channel under the wide-speed-range flight working condition.

The present application relates to a kind of composite cavity combustion chamber and hypersonic ramjet engine, wherein, composite cavity combustion chamber, comprising: combustion chamber main body (1) and the cavity structure (2) installed on the combustion chamber main body (1);The cavity structure (2) includes: first cavity portion (21) and second cavity portion (22);The first cavity portion (21) first cavity less than the second cavity that the second cavity portion (22) is surrounded;Along the width direction perpendicular to the direction of incoming flow of combustion chamber, the second cavity portion (22) with the first cavity portion (21) is connected and is arranged in the opposite side of the first cavity portion (21), and the first cavity is communicated with the second cavity;The opening side of the first cavity portion (21) is connected with the opening side of the second cavity portion (22) on the combustion chamber main body (1).

The application belongs to the technical field of ramjet engine, and discloses a kind of ignition enhanced ramjet engine combustion chamber and its working method.The ignition enhanced ramjet engine combustion chamber includes combustion chamber front expansion section, ignition flame stabilizing section and combustion chamber rear expansion section connected in sequence from front to back along the direction of incoming flow, and the center cavities of each section are through from front to back.The annular flame stabilizing groove is arranged on the inner wall of ignition flame stabilizing section, the circumferential direction of flame stabilizing groove is arranged with uniformly distributed secondary amplification groove, the primary amplification groove is arranged on the secondary amplification groove, and the top end of primary amplification groove is arranged with igniter interface.The working method uses the characteristics that the combustible mixture gas flow in each level of backflow area formed by each level of groove is relatively stable, gradually slows down the combustible mixture gas, obtains the combustible mixture gas with lower speed, and then ignites with igniter, which improves the ignition reliability and enhances the flame stability, and expands the ignition Mach number range of each type of igniter on the ramjet engine.

The application provides a rotary ram shock wave supercharged gas turbine engine and belongs to the technical field of turbine engines, which comprises an outer casing and an inner casing, an outer channel is formed between the inner casing and the outer casing, a guide vane is installed in the outer casing and is used for guiding airflow, a first mechanism is installed in the inner casing and is used for supercharging and heating the airflow and doing work, a second mechanism is arranged on one side of the inner casing and is used for supercharging and heating the airflow for the second time. The application solves the problems of large windward area of the centrifugal compressor, many parts of the multistage axial flow compressor and low efficiency of small flow by using the rotary ram compressor, higher flow and pressure ratio are obtained in compact size, greater thrust is obtained under the condition of smaller outer size of the original engine, the outer size is reduced by about 32%, and the thrust is increased by 17.3% under the condition of the same fuel consumption rate when the engine is at rest.

The present application belongs to the technical field of guidance, and particularly relates to a ramjet composite range increasing projectile with hidden controllable folding wings and a design method. The ramjet composite range increasing projectile comprises a central cone and a composite range increasing projectile body located outside the central cone, the central cone is coaxially fixedly connected with the composite range increasing projectile body, and characterized in that the composite range increasing projectile body is provided with hidden controllable folding wings, and the hidden controllable folding wings are detachably fixedly connected with the central cone through an ejection device. The present application aims to provide a ramjet composite range increasing projectile with hidden controllable folding wings and a design method. While the advantages of the ramjet composite range increasing projectile are retained, the hidden controllable folding wings are designed and applied to the ramjet composite range increasing projectile, and the design of the ramjet composite range increasing projectile with hidden controllable folding wings is completed.

The utility model discloses a scramjet engine combustion chamber structure, a cooling channel structure and a combustion system.The combustion chamber structure comprises a combustion chamber wall, a combustion chamber is defined by the combustion chamber wall, and a plurality of cooling channels are arranged in the combustion chamber wall at intervals in the circumferential direction of the combustion chamber wall; the extending direction of the cooling channel is consistent with the extending direction of the combustion chamber, the inner wall of the cooling channel is in a bamboo joint shape, the combustion system comprises a fuel chamber and the combustion chamber structure, and an outlet of the fuel chamber is connected with an inlet of the combustion chamber and an inlet of the cooling channel. An outlet of the cooling channel is connected with an inlet of the combustion chamber. According to the scramjet engine combustion chamber structure, the cooling channel structure and the combustion system, heat exchange between the cooling medium and the combustion chamber wall can be enhanced, the cooling effect is good, aviation fuel serves as the cooling medium of the combustion system, heat absorbed by the cooling medium can be further utilized, and the combustion efficiency in the combustion chamber is improved.

The present application relates to the technical field of cross-medium ramjet engine and aircraft, in particular to high-mobility cross-medium ramjet engine and aircraft, comprising an engine structure, the engine structure comprises an engine shell, a communication control valve pipe, a control atomizer, a vent valve seat and an airflow passage pipe, the rear end of the engine shell is provided with the communication control valve pipe, the rear end of the communication control valve pipe is provided with the control atomizer in communication, the rear end of the control atomizer is provided with the vent valve seat, the rear end of the vent valve seat is provided with the airflow passage pipe in communication, the engine structure further comprises a powder fuel storage, a first communication control pipe, a second communication control pipe, a combustion chamber, a combined tail nozzle, a closed adjusting block and a protection ring block, the control atomizer and the airflow passage pipe are provided in communication with the combustion chamber, and the rear end of the combustion chamber is provided with the combined tail nozzle, through the structure, cross-medium driving work is facilitated.

The invention relates to the technical field of rocket engines, in particular to a solid-liquid mixed combined type ramjet engine which comprises a rocket body, supporting parts are fixedly installed at the front end of the rocket body and distributed annularly, and a spray pipe system is fixedly installed at the front end of each supporting part and connected with the rocket body. The supporting part comprises a positioning device, a transmission device and a displacement device, the displacement device is slidably inserted into the inner center of the positioning device, and the transmission device is fixedly installed in the positioning device and located above the displacement device; the positioning device comprises a positioning cavity, a combustion chamber, a heat insulation sleeve layer, a solid fuel column, a guide sleeve, a connecting base frame, a driving motor, a transmission worm, a movable hole and a limiting groove, the combustion chamber is fixedly installed at the front end of the positioning cavity, the heat insulation sleeve layer is fixedly installed on the inner wall of the combustion chamber, and the solid fuel column is fixedly installed on the inner wall of the heat insulation sleeve layer. And through the arrangement of the supporting component, the purpose that an oxidizing agent can be injected into different areas in the combustion column is achieved.

The application discloses a hyperplane climbing track group intelligent optimization method, which comprises the following steps: constructing a hyperplane climbing section dynamics equation, and obtaining to-be-optimized parameters in the hyperplane climbing section dynamics equation; designing a hierarchical cost function considering multiple constraints and turbojet, subsonic ramjet and supersonic ramjet three power mode switching Mach numbers; and optimizing the cost function through a chicken swarm algorithm, thereby completing the hyperplane climbing track group optimization. The hyperplane climbing track parameter optimization is performed through the improved chicken swarm algorithm, the suitable power mode switching Mach number is selected while the complex constraints of the hyperplane climbing section are met, the climbing scheme capable of significantly shortening the climbing time and total fuel consumption is obtained, and the problem that the track is difficult to optimize under the hyperplane cross-power mode and long-time window climbing condition is solved.

The invention belongs to the field of engine propulsion systems, and particularly relates to a fuel stamping propulsion system and method. Wherein the fuel stamping propulsion system comprises a fuel storage tank and a self-pressurization control device; the fuel storage tank is arranged outside a high-temperature fuel gas flow channel of the engine in a winding mode, n partition plates are arranged in the fuel storage tank, the n partition plates divide the fuel storage tank into n + 1 sub-cabin storage tanks in the axial direction of the engine, and each partition plate is provided with a communication control valve communicated with the adjacent sub-cabin storage tanks. The self-pressurization control device comprises a pressurization part and a supply part. The advantages that ammonia gas is prone to gasification and phase change and large in heat sink are fully utilized, the technical method of ammonia gas self-pressurization is utilized, the complexity degree of an engine supply pipeline is greatly reduced, extra devices such as a fuel pump and a power battery are omitted, a stamping propulsion system is greatly simplified through simple self-pressurization control, and the cost is reduced. And structural thermal protection and system use requirements with higher Mach number are met.

The invention relates to the technical field of engines, in particular to a magnetofluid-water ramjet mixed cross-medium engine which comprises a mounting protective cover, a magnetofluid power generation propelling system, a conveying spray pipe, a water ramjet engine system, a first adjusting nozzle and a first motor. The conveying spray pipe and the water ramjet engine system are arranged in the installation protective cover, and the water ramjet engine system is connected to the upper end of the conveying spray pipe. Metal fuel powder arranged in the water ramjet engine system can be timely and sufficiently supplemented, so that the water ramjet engine system can stably run, the fuel powder in the water ramjet engine system can be fully contacted and reacted with clear water in the running process, the contact mixing reaction efficiency is high, and the service life of the water ramjet engine system is prolonged. The aircraft is simple in structure, rapid and thorough in response, capable of providing sufficient power in water and capable of assisting in providing power when flying in the air, so that the aircraft flies stably and has a longer flight stroke.

The invention belongs to the technical field of prediction and control of a supersonic air inlet channel global flow field of a ramjet engine, and particularly relates to a prediction method and a prediction system for a wide-speed-domain variable-geometry air inlet channel global flow field. The method comprises the steps that 1, wide-speed-domain variable-geometry air inlet channel flow field information is obtained; 2, based on the information obtained in the step 1, a variable geometry air inlet channel global flow field prediction data set is constructed; 3, a variable geometry air inlet channel wall surface sensor configuration scheme is constructed, and the pressure value of the variable geometry air inlet channel is collected; and 4, designing a global flow field prediction model structure of the air inlet channel based on the variable geometry air inlet channel acquired in the step 3, and realizing prediction of the global flow field of the wide-speed-domain variable geometry air inlet channel. The method is used for solving the problem that the accurate perception of the hypersonic flow field is in the future development of hypersonic velocity.

The invention relates to the technical field of thermal protection of high-temperature parts of hypersonic flight vehicles, in particular to a periodic swing type jet impingement cooling structure of a scramjet engine based on the coanda effect. Comprising a fluid oscillation jet flow hole structure, a combustion chamber wall face cooling channel and a fuel distribution and recovery device. The combustion chamber wall face cooling channel comprises a cooling channel body, a confluence cavity is formed above the cooling channel body, and a plurality of fluid oscillation jet hole structures are formed in the position, corresponding to the cooling channel body, of the bottom of the confluence cavity. The fuel flow dividing and recycling device is composed of a plurality of flow adjusting devices on a fuel pipeline, and a fuel inlet and a confluence cavity inlet are jointly connected with a fuel pump and a fuel tank through the fuel pipeline. Unpowered periodic swing type jet cooling is adopted, a jet channel is combined with a fluid oscillator, periodic swing of jet is driven through the fluid pressure difference and the coanda effect in the engine, use of extra power equipment is avoided, system complexity is reduced, and cooling efficiency is improved.

The invention belongs to the technical field of aero-engine control, and discloses an anti-interference switching control method for a rocket ramjet combined engine. The working mode and the control requirement of the rocket ramjet combined engine are analyzed, and the control loop switching rate based on the mode and protection is designed. For different working modes and flight conditions, feedforward adjustment is introduced on the basis of a linear active disturbance rejection algorithm, and the rapidity and disturbance rejection performance of adjustment of the rocket ramjet in all states are guaranteed. Furthermore, the invention provides a mixed residence time method depending on a system state, so that the conservative property of switching limitation is reduced, and the effects of ensuring the action sensitivity when the engine is close to the switching demand and ensuring the switching stability of the engine are achieved. The smooth switching strategy is introduced at the switching moment, a control system is built, and finally the self-adaptive switching control function of the rocket ramjet combined engine is achieved.

The invention relates to the field of ramjet kerosene shunting and regenerative cooling, and discloses a ramjet regenerative cooling kerosene shunting device which comprises an annular piece, a plurality of kerosene flow channel inlets, a plurality of kerosene flow channel outlets formed in the other side of the annular piece, an annular confluence groove and a plurality of flow channels formed in the annular piece. The annular confluence groove is communicated with an outlet of each kerosene flow channel through the shunting channel; wherein the shunting channel comprises a convergence section, a throat channel and an expansion section, the convergence section comprises a rapid convergence section and a saturation pressure section, and the expansion section comprises a primary expansion section, an equal-diameter channel and a secondary expansion section. According to the invention, through the arrangement of the shunting channel, when kerosene passes through the convergence section of the shunting channel, the fluid speed is accelerated, the pressure is reduced, and the static pressure at the throat part is reduced to the saturated vapor pressure, the liquid is sharply vaporized, so that the flow of the kerosene after passing through the expansion section is kept stable, and the kerosene flow in each cooling channel connected with the downstream is uniform and kept stable; the regeneration cooling effect of each area is ensured.

The invention discloses a trapped vortex cavity subsonic combustion ramjet combustion chamber coping with distorted incoming flow. The trapped vortex cavity subsonic combustion ramjet combustion chamber comprises a combustion chamber case, a diffuser, a center cone, a rectifying grating, a splitter plate, a trapped vortex cavity stabilizer, a tail cone, a heat shield, n first radial supporting plates and n second radial supporting plates. The integrated design of the center cone and the rectification grating is adopted, efficient rectification of distorted incoming flow is achieved, uniform flow field distribution is formed in front of a combustion area, meanwhile, the standing vortex cavity stabilizer serves as an on-duty level, and the problems of ignition flame stabilization and efficient combustion under the distorted incoming flow are effectively solved.

The invention belongs to the technical field of ramjet engines, and particularly relates to a throat stepless adjustable spray pipe with fixed inlet and outlet areas, the spray pipe adopts a contraction and expansion structure and comprises a round-to-square section, a contraction section and an expansion section, and the inner molded surface of the spray pipe can be actively adjusted according to the flow change of an upstream engine to adapt to different incoming flow conditions. The motor rod is driven by the motor, the single-degree-of-freedom linear motion of the motor rod is converted into synchronous rotation motion of the contraction section and the expansion section, and meanwhile the throat area change meets the expected design target. Stepless adjustment of the throat is achieved under the condition that the inlet and outlet area and the length of the spray pipe are not changed, and the performance of the spray pipe under different incoming flow conditions can be guaranteed while the requirement for matching of different flows of an upstream engine can be met. In addition, according to the design of the wall surface type surface at the lap joint, the attachment degree of the contraction section and the expansion section of the spray pipe under the driving of a single power source is good through the design of the wall surface type surface at the lap joint according to the steps.