151results about "Ram jet engines" patented technology

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.

An air-breathing turbojet engine (301) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (302) configured to compress inlet air by one or more shocks, a cooler (303) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (304) to compress the air further. A precooler (311) between the inlet (302) and the cooler (303) cools compressed inlet air using compressed air from the turbo-compressor (304). A combustor (305) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (306) expands and is driven by combustion products, and a second turbine (307) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (309) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

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 invention provides a solid powder pre-combustion assembly and a heater assembly, comprising: a fluid interconnection device having a first interface for receiving fluidized powder fuel, a second interface for receiving fluidized powder fuel, and a third interface for receiving fluidized powder fuel; the first fuel gas generator is used for receiving auxiliary gas and generating high-temperature fuel gas, and a gas outlet of the fuel gas generator is communicated with the second connector of the fluid interconnection device; and an inlet of the pre-combustion chamber communicates with the third connector, and the pre-combustion chamber is used for receiving high-temperature fuel gas and powder fuel, conducting pre-combustion and removing a surface oxidation layer of the powder fuel. Compared with the prior art, the solid powder pre-combustion assembly and the heater assembly are simple in structure, and the solid powder combustion completeness and combustion efficiency can be improved.

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.

This invention relates to an inlet structure and determination method for a subsonic and scramjet engine. The outer compression section of the inlet consists of an outer compression section horizontal plane, a first panel, a second panel, and a third panel. An angled lip is provided at the inlet of the lower end face of the inner compression section. When determining the inlet structure parameters, the design of the subsonic and scramjet engine inlet is combined. Only the inlet Mach number needs to be determined to automatically identify and select the appropriate ramjet engine inlet geometry determination process, ultimately outputting the optimal inlet geometry that meets the design requirements. This invention enables rapid two-dimensional inlet structure design for subsonic and scramjet engines and provides optimized shape and performance parameters, laying the foundation for subsequent overall propulsion performance analysis of the ramjet engine.

An intentional fluid manipulation apparatus (IFMA) assembly that includes an upstream intentional momentum shedding apparatus (IMSA) configured to impart a first induced velocity to a local free stream flow during a nominal operation requirement. The upstream IMSA creates a streamtube. The IFMA includes a downstream IMSA, with some or all of the downstream IMSA being located in a downstream portion of the streamtube. The downstream IMSA imparts a second induced velocity to the local free stream flow within the streamtube. The second induced velocity at the location of the downstream IMSA has a component in a direction opposite to the direction of the first induced velocity at the location of the downstream IMSA.

The application discloses a combined engine and a control method thereof. The combined engine comprises a control assembly, an air inlet channel, a turbine core engine, a nuclear reactor system, a propellant supply system and a nozzle assembly which are connected through pipelines. The control assembly is used for controlling the air inlet end of the nuclear reactor system to be selectively connected with the air inlet channel, the turbine core engine or the propellant supply system through the pipelines. The air outlet end of the nuclear reactor system is connected with the nozzle assembly. The air inlet end of the nuclear reactor system is connected with the air inlet channel, the turbine core engine and the propellant supply system through the pipelines respectively by the control assembly, so that the combined engine enters the nuclear ram mode, the nuclear turbine mode and the non-suction mode respectively. The combined engine has the characteristics of high fuel economy, working in the atmosphere-free environment, realizing different speed domain flights by using different engine modes and meeting the needs of space transportation of the super-speed aircraft.

A combined power intake duct with a translational linkage flow regulation geometry device and its operating method are disclosed. This invention features a wide-range intake flow regulation range. Based on the principles of changing the capture flow and increasing the throat flow capacity for stable intake operation, this invention proposes a linkage flow regulation geometry device involving the translational lip and throat. The geometry device consists of a translational lip, a translational throat, and a four-bar linkage. It can achieve multi-dimensional optimization of intake performance by coordinating and controlling the actual capture flow and throat compression through single-degree-of-freedom adjustment. The geometry device operates in two states during movement: 1) When the four-bar linkage moves backward, the actual capture flow decreases, the ram throat height increases, and the turbine throat height decreases; 2) When the four-bar linkage reaches a certain position and continues to move backward, the actual capture flow decreases, the ram throat height decreases, and the turbine throat height increases.

The invention discloses a double-flow-channel combustion chamber and a rocket-based combined power engine. The problem that a combustion chamber cannot stably work in a wide range in a high-performance mode is solved. The first flow channel is provided with a first isolation section and a first combustion chamber, and a first flame stabilizer is arranged on the side, close to the first isolation section, of the first combustion chamber; the second flow channel is provided with a second isolation section and a second combustion chamber, a second flame stabilizer is arranged on the side, close to the second isolation section, of the second combustion chamber, the first flame stabilizer and the second flame stabilizer are both used for injecting fuel, the second flow channel is parallel to the flowing direction of the first flow channel, and the expansion ratio of the second flow channel is larger than or equal to that of the first flow channel; the first combustion chamber and the second combustion chamber communicate with the second-stage combustion chamber, and the expansion ratio of the first flow channel and the expansion ratio of the second flow channel are both smaller than the expansion ratio of the second combustion chamber. The power device communicates with the second-stage combustion chamber, and the jetting direction of the power device inclines towards the flowing direction of the second flow channel and is used for jetting high-temperature wake flow into the second-stage combustion chamber.

The invention discloses an end face combustion solid-liquid fuel gas generator which comprises a combustion chamber and a solid grain arranged in the combustion chamber, an exhaust port communicated with a combustion cavity is formed in one end of the combustion chamber, and the end, facing the exhaust port, of the solid grain is a combustion end. A space formed by the end face of the combustion end and the cavity wall of the end, provided with the exhaust port, of the combustion cavity is a combustion space, the spring is arranged in the combustion cavity, and the cavity wall, located between the end of the non-combustion end of the solid grain and the cavity wall of the combustion cavity, of the combustion cavity is provided with an annular groove capable of abutting against the combustion end. The movement of the solid grain towards the exhaust port is limited; the annular groove enables the two ends of the spring to abut against the solid grain and the combustion cavity respectively and to be in a compressed state, so that elastic force for moving towards the exhaust port is continuously applied to the solid grain, and the combustion space is kept unchanged in the combustion process of the solid grain.

The application discloses a cross-medium scramjet engine and relates to the field of engines.The cross-medium scramjet engine comprises an engine body, a guide vane is fixedly installed on the circumferential outer side of a tail-end injection port of the engine body, an air inlet channel is formed in the inner side of the end of the engine body away from the guide vane, a butt joint seat is fixedly arranged on the outer side of the end air inlet channel of the engine body, an air inlet cylinder is fixedly arranged at the end of the butt joint seat, air inlet holes are formed in the circumferential outer wall of the air inlet cylinder, and a flow-through vane is fixedly arranged on the circumferential outer side of the air inlet cylinder; a turbine is arranged at a position close to the injection port in the inner side of the engine body; and a combustion assembly is arranged between the turbine and the air inlet channel.The cross-medium scramjet engine is helpful to improve the mixing uniformity of fuel and air in the combustion chamber, realize more stable and efficient combustion, reduce problems such as local overheating or insufficient combustion, and improve the reliability and thrust output stability of the 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 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.

An environmental control system includes a ram air circuit having a ram air duct and at least one ram heat exchanger arranged within the ram air duct, a compression device, and an expansion device. The compression device is configured to receive at least one of a first medium, a second medium, and a third medium. The environmental control system is operable in a plurality of modes including at least one normal mode and a failure mode. In the at least one normal mode, the first medium output from the compression device is exhausted into the ram air duct upstream from the at least one ram heat exchanger, and in the failure mode, only the third medium output from the compression device is exhausted into the ram air duct upstream from the at least one ram heat exchanger.

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 invention discloses a water-air dual-purpose engine. The engine takes air or water as a combustion medium; the engine includes: a medium introduction part for introducing a combustion medium into the engine; the trapped vortex combustion part is used for accommodating the fuel and the combustion medium under the condition that the combustion medium is air, so that the fuel is combusted; the metal powder combustion part is used for accommodating fuel and a combustion medium under the condition that the combustion medium is water, so that the fuel is combusted; and the substance spraying part is used for spraying substances generated by combustion out of the engine so as to generate traction force. According to the water-air dual-purpose engine, the two different propelling modes of turbojet and metal powder water stamping are integrated, the two propelling modes are compatible at the same time according to different environments and working conditions, and the working states are synchronously switched during environment switching so as to adapt to combustion media corresponding to the different environments; therefore, the engine can provide power for the aircraft in different propulsion modes in the air environment and the water environment respectively, and the stability of fuel combustion is kept.

The application discloses a kind of throttling-based auxiliary ignition device, including circular cross-section combustion chamber, kerosene injection support plate and injection hole installation point, combustion chamber includes isolation section, groove section and expansion section, kerosene injection support plate is installed in isolation section export, injection hole installation point and the installation position of kerosene injection support plate are in the same section, injection hole installation point is spaced distribution with kerosene injection support plate, and a group of air injection holes are installed in injection hole installation point.The application can promote kerosene reliable ignition to start stable combustion under the condition of not using flammable gas, high-frequency spark plug and igniter, and can be reused, simplify auxiliary ignition device, reduce system complexity, improve reliability.

The present application relates to a ram air turbine emergency power drive device, comprising: a first set of turbine blades and a second set of turbine blades coaxially arranged in an axial direction, wherein the second set of turbine blades can be rotated relative to the first set of turbine blades around the axial direction from a first position to a second position and kept in the second position, wherein in the first position, the first set of turbine blades is superposed with the second set of turbine blades, and in the second position, the first set of turbine blades is at a predetermined non-zero angle with the second set of turbine blades. In this way, the number of blades is increased without affecting the space in the ram air turbine cabin, so that the turbine starting speed is faster, and the RAT access to the power grid time is reduced. In addition, increasing the number of blades can also improve the efficiency of the turbine in converting wind energy, thereby improving the efficiency of the emergency power generation. The present application also relates to an aircraft comprising the ram air turbine emergency power drive device.

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 provides a subsonic combustion ramjet engine air inlet channel based on an adjustable partition plate and a back pressure resisting method of the subsonic combustion ramjet engine air inlet channel. The subsonic combustion ramjet engine air inlet channel comprises a lower molded surface, an upper molded surface, a side wall and a partition plate assembly. The side wall is connected with the upper molded surface and the lower molded surface, an inner channel is defined by the upper molded surface, the lower molded surface and the side wall, and the inner channel comprises an inner contraction section, a throat section, an expansion section and an equal straight section which are sequentially arranged in the fluid flowing direction; the partition plate assembly is arranged in the expansion section in the fluid flowing direction and comprises a front partition plate, a rear partition plate and a rotating shaft, the front partition plate and the rear partition plate are hinged through the rotating shaft, the end, away from the front partition plate, of the rear partition plate is fixedly connected with the side wall, the rotating shaft is rotationally installed on the side wall, and the expansion section is divided into an upper channel and a lower channel through the partition plate assembly. The front partition plate rotates around the rotating shaft, the expansion ratio and flow distribution of the upper channel and the lower channel are adjusted, and the back pressure resistance of the upper channel and the back pressure resistance of the lower channel are improved, so that the overall back pressure resistance of the air inlet channel is improved, and flow loss is avoided.

This invention provides a solid powder combustion power system, including a combustion power unit and an atomizing device. The combustion power unit consists of a powder inlet flange, a first-stage combustion chamber, an igniter assembly, and a tail nozzle connected coaxially in sequence. The atomizing device delivers solid powder into the combustion chamber via a pipeline through the powder inlet flange. The first-stage combustion chamber includes a water spray flange, a water inlet flange, and a connecting flange. Inner and outer pipes are installed between the water inlet flange and the connecting flange to form an annular sealed cooling water supply chamber, with the inner cavity of the inner pipe serving as the combustion chamber. The water inlet flange has a water inlet channel, and the water spray flange has a water spray channel with a water mist nozzle. The two are connected via a water supply channel with a switch valve, allowing water to be sprayed into the combustion chamber as needed. The igniter ignites the solid powder within the chamber, and the tail nozzle has a flame outlet for energy output. This system employs an integrated water-cooled structure on the combustion chamber wall, ensuring uniform cooling and facilitating installation and maintenance. The cooling water achieves uniform wall cooling and can also be sprayed into the combustion chamber to generate combustible gas, improving combustion efficiency and energy release.

A system and method for ram air intake for pulse combustion systems is disclosed that improves the ability of pulse combustions to ingest air into the inlet pipe when the pulse combustion system is moving in a direction opposite the direction the open end of the inlet pipe is facing and the system and method includes the ability to increase the thrust output from the pulse combustion system.

The invention relates to the technical field of high-temperature thermal protection of scramjet engines, in particular to a jet flow impact type active and passive composite cooling structure of a scramjet engine. Wherein a jet flow and regeneration channel combined structure is adopted for active cooling, supercritical hydrocarbon fuel serves as a medium for active cooling, and a closed-loop design of flow dividing, flow converging, jet flow and regeneration is adopted; passive cooling is achieved by laying an ultra-high-temperature ceramic matrix composite coating on the wall face of the combustion chamber, heat in the combustion chamber can be effectively prevented from being conducted to a wall face matrix, the combination performance with a metal or composite matrix is good, and the passive cooling device is suitable for the thermal environment under the high Mach number working condition. And through the synergistic effect of active and passive cooling, safe and stable operation of the scramjet engine under the high Mach number working condition is guaranteed.

This invention discloses a high-efficiency propellant-type water ramming test engine, comprising a front end cap, an outer cylinder, a rear end cap, a combustion chamber, and a nozzle connected in sequence, as well as an inner cylinder, an injection element, a piston, and a propellant grain. The front end cap serves as an air intake channel; the outer cylinder has a water inlet, forming a sandwich between the outer and inner cylinders; the rear end of the inner cylinder has a circular hole that forms a water flow channel with the injection element, and the interior of the inner cylinder forms a combustion chamber; the front end of the piston contacts the filling gas, and the rear end is in close contact with the propellant grain; the injection element sandwich forms a water flow channel, and a small circular hole on the injection element allows water to be injected into the combustion chamber; the propellant grain is a metal-rich fuel propellant, and during combustion, water in the propellant reacts with the metal to release energy through a chemical reaction. During operation, the propellant grain gradually burns and recedes, while the filling gas pushes the piston through the air intake channel to maintain an optimal constant distance between the propellant grain combustion surface and the water nozzle, ensuring that the combustion efficiency remains at its highest level.

The application relates to the technical field of detonation aero-engine, in particular to a detonation wave efficient initiation device which comprises a detonation mechanism, a combustion cabin, a fuel pipe arranged on the outer wall of the combustion cabin and a pre-detonation pipe penetrating through the outer wall of the combustion cabin; an adjusting mechanism which comprises an elastic piece arranged on the inner wall of the pre-detonation pipe, a fixing frame arranged on the end surface of the outer wall of the pre-detonation pipe and a moving assembly; a connecting mechanism which comprises a connecting ring and a connecting column; and a limiting mechanism which comprises a limiting column and a positioning assembly; the adjusting mechanism is arranged to control the displacement adjustment of the elastic piece by the connecting ring, so that the roughness of the inner wall of the pre-detonation pipe is changed, thereby adapting to different application scenarios; after the surface roughness is adjusted, the elastic piece is fixed by the limiting mechanism, the constant roughness is maintained, meanwhile, the pre-detonation pipe is clamped, fixed and limited, and the shaking of the pre-detonation pipe is prevented from affecting the process of the accurate detonation wave initiation.

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.