384 results about "Ramjet" patented technology

The invention discloses a ramjet numerical simulation method, device and equipment and a storage medium, and relates to the technical field of aerodynamics, and the method comprises the steps: determining a dimensionless temperature corresponding to a gas component of an initial flow field; based on the dimensionless temperature, determining a molecular viscosity collision integral function of the gas component in a polynomial fraction fitting mode; determining the viscosity coefficient of the gas component according to the molecular viscosity collision integral function; and generating a computational grid of the ramjet based on the viscosity coefficient so as to perform numerical simulation operation on the ramjet. Therefore, in the numerical simulation process of the ramjet, the polynomial fraction is adopted to fit the viscosity collision integral function, the calculation time of the gas viscosity coefficient can be effectively shortened, and the calculation precision is kept to meet the use requirement; in addition, the method can adapt to the combustion flow field temperature requirement and the component requirement of the ramjet engine, and the numerical simulation efficiency of the ramjet engine can be improved.

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

The invention provides a multi-cavity scramjet engine combustion chamber and a control method thereof. The combustion chamber is simple in structure, capable of adapting to different equivalence ratio ranges and different flight heights, low in resistance, high in combustion performance, low in thermal load of the wall face of the combustion chamber, convenient to maintain and capable of facilitating aircraft-engine integrated design. The combustion chamber comprises an isolation section, a first concave cavity group, a second concave cavity group, a third concave cavity group, an injection module and a combustion chamber expansion section, each concave cavity group is provided with two concave cavities which are oppositely arranged up and down; the isolation section is a first square tube and is connected with the first concave cavity group; the first concave cavity group is connected with the second group of concave cavities through a second square tube; the second group of concave cavities is connected with a third concave cavity group through a third square pipe; a combustion chamber expansion section is connected behind the third concave cavity group; the injection unit comprises a front injection device and a rear injection device; front injection devices are oppositely arranged on the upper wall surface and the lower wall surface of the second square pipe; rear injection devices are oppositely arranged on the upper and lower wall surfaces of the third square pipe.

The invention relates to the technical field of ramjet engines, in particular to a composite ramjet engine and a combined power device. According to the composite ramjet engine, a scramjet engine runner serves as a basic runner, and the runner can be switched into an equivalent subsonic combustion ramjet engine runner through profile adjustment. The basic flow channel comprises an air inlet channel, an isolation section, a combustion chamber front section, a combustion chamber rear section and an expanding exhaust nozzle. A flow channel molded surface adjusting mechanism is arranged on the inner wall of the expanding exhaust nozzle; the mechanism has two states: in the first state, the adjusting mechanism is attached to the inner wall of the expansion type exhaust nozzle, and the molded surface of a flow channel of the exhaust nozzle is in an expansion type and corresponds to a dual-mode scramjet mode; and in the second state, the adjusting mechanism protrudes towards the inner cavity of the exhaust nozzle, and the exhaust nozzle runner comprises a first expansion section, a first contraction section and a second expansion section in the airflow direction and corresponds to the equivalent subsonic combustion stamping mode. The working range of a conventional speed range scramjet engine can be widened to Ma2.5-7.0, and the conventional speed range scramjet engine can be directly combined with a common turbine engine.

The invention belongs to the technical field of design of turbo-ramjet combination engines, and particularly relates to a turbo-ramjet combination engine and a control method thereof. According to the design, a complex outer turbine channel is separated from a smooth ramjet channel, and an afterburner and a ramjet combustion chamber are designed to share the space; the designed center stamping channel has multiple functions of a turbine engine stator force bearing frame, a center deflation channel, a subsonic combustion stamping rectification channel, a scramjet channel and the like. Self-adaptive switching of four working modes including a turbine engine conventional working mode, a turbine engine center deflation working mode, a subsonic combustion stamping working mode and a scramjet working mode is achieved, the space of a center stamping channel is utilized to the maximum extent, the working speed range of the turbine modes is widened, the cross sectional area is greatly reduced, the flight resistance is effectively reduced, and the service life of the turbine engine is prolonged. And wide-speed-range work with the Mach number of 0-7 is realized.

The invention provides a powder fuel-based cross-medium ramjet engine which comprises a powder fuel cabin, a combustion chamber and a spray pipe are fixedly connected to the axial front end of the powder fuel cabin, a high-pressure gas tank is fixedly connected to the axial rear end of the powder fuel cabin, and a plurality of gas generators are fixedly arranged on the inner wall of the combustion chamber; a plurality of air inlet channels are arranged on the outer wall of the combustion chamber; the high-pressure gas tank is communicated with the powder fuel cabin through a high-pressure gas flow channel and a fluidization pipeline, and is communicated with the gas inlet channel through a gas inlet channel pressurizing pipeline; a liquid medium cavity is formed in the air inlet channel and communicates with the combustion chamber through a liquid medium main supply pipeline, an upstream liquid supply pipeline and a downstream liquid supply pipeline. Combustion energy release of the powder fuel in the air mode and the underwater mode of the cross-medium ramjet engine is effectively promoted, and the method is suitable for the powder fuel cross-medium ramjet engine.

The invention discloses an integrated closed circulation system for an RBCC engine, and solves the problems that a constant-pressure extrusion mode of an existing RBCC engine propellant supply system is not suitable for variable working condition adjustment of an RBCC rocket thrust chamber, and a high-pressure storage tank is huge in structure, large in weight and not beneficial to application and expansion of the system on the day. According to the system, high-temperature oxygen is generated through catalytic decomposition of hydrogen peroxide and acts on the turbine, then the hydrogen peroxide pump and the coal oil pump are driven to work, integrated pressurization, supply and adjustment of propellants can be carried out on the multiple rocket thrust chambers and the ramjet combustion chamber at the same time, and therefore wide-range fuel oil adjustment is met, and a widely-applied electric pump system can be replaced; and the structural envelope and weight are greatly reduced, so that the power requirement of a longer voyage is met.

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 discloses a multistage self-adaptive enhanced mixing device for a solid fuel ramjet engine, which belongs to the technical field of solid fuel ramjet engines and comprises a front flange, a rear flange, an afterburning chamber barrel, a mixing partition plate, a heat insulating layer, a pressure measuring seat and a temperature measuring seat. The front flange and the rear flange are welded with the afterburning chamber barrel, a water pressure test is carried out after welding is completed, the mixing partition plates are bonded to the inner wall of the afterburning chamber barrel through high-temperature-resistant glue, and the mixing partition plates are distributed in a multi-stage mode, are separated through a heat insulation layer and are bonded through the high-temperature-resistant glue. In order to verify the effect of the multi-stage self-adaptive enhanced mixing device for the solid fuel ramjet engine, a pressure measuring point and a temperature measuring point are arranged on an afterburning chamber barrel and a mixing partition plate; according to the multi-stage self-adaptive enhanced mixing device, the combustion efficiency of a solid fuel ramjet can be remarkably improved, and the efficiency is further improved along with the increase of the number of the mixing partition plates.

The invention particularly relates to a parallel combination engine starting and fuel oil supply system cross-linking structure and a control method, and the parallel combination engine starting and fuel oil supply system cross-linking structure comprises an air turbine, an aircraft accessory cartridge receiver, an engine transmission system and a stamping fuel oil pump; an air turbine shaft is connected with an aircraft accessory casing, a starting clutch is arranged between the air turbine shaft and the aircraft accessory casing, the aircraft accessory casing is connected with an engine transmission system, and the engine transmission system is connected with a turbine engine; an air turbine shaft is connected with a ramjet fuel pump, an oil supply clutch is arranged between the air turbine shaft and the ramjet fuel pump, and the ramjet fuel pump is used for The air turbine is connected with an airplane auxiliary power system through an air entraining pipeline; an auxiliary air entraining valve is arranged on the air entraining pipeline; the air turbine is connected with the turbine engine fan rear flow channel through an air entraining pipeline, and a fan air entraining valve is arranged on the air entraining pipeline. The air turbine is connected with the stamping air inlet channel through an air entraining pipeline, and a stamping air entraining valve is arranged on the air entraining pipeline.

The invention belongs to the technical field of scramjet engine control, and particularly relates to a semi-physical simulation test platform of a scramjet engine control system, which comprises a fuel supply and control system, an engine simulator system, a flight control simulation system, a sensor simulation system, an upper computer management system and the like. The test platform can provide test equipment support for development (including control law research, control parameter optimization, performance prediction, fault diagnosis and the like) of a scramjet engine control system, and comprehensively examine system matching performance and control performance of key components such as a controller, a fuel pump, a fuel valve and the like. And stable and reliable operation of the scramjet engine control system is guaranteed.

The invention relates to the technical field of ramjet engines, in particular to a wide-speed-range scramjet combustion chamber, a scramjet engine and a combined power device. The wide-speed-range scramjet combustion chamber comprises a front combustion section and a rear combustion section; the front combustion section comprises a front-section runner wall surface, a front-section concave cavity flame stabilizer, a front-section wall surface fuel injector and a front-section multi-mode flame stabilizer; the rear combustion section comprises a rear section runner wall surface, a rear section concave cavity flame stabilizer, a rear section wall surface fuel injector and a rear section multi-mode flame stabilizer; the front-section wall surface fuel injector and the rear-section wall surface fuel injector are fixedly mounted on the wall surfaces in front of the front-section concave cavity flame stabilizer and the rear-section concave cavity flame stabilizer respectively, and the front-section multi-mode flame stabilizer and the rear-section multi-mode flame stabilizer are located in the middle of a front-section flow channel and the middle of a rear-section flow channel respectively and located at the same axial position as the front-section concave cavity flame stabilizer and the rear-section concave cavity flame stabilizer; the lower limit of the working speed domain of the scramjet combustion chamber in the conventional speed domain of Ma 4.0-7.0 can be widened to Ma 3.0-3.5.

The invention belongs to the technical field of scramjet engine control, and particularly relates to a regenerative cooling ramjet engine temperature control system which comprises a novel cooling oil supply system with an oil drain valve as a core and an engine thrust and temperature decoupling control method. The working boundary of the engine is widened, and it is guaranteed that in the temperature control process, the temperature is controllable, and the thrust of the engine is kept constant.

The invention relates to the field of scramjet engines, in particular to a method for inhibiting combustion asymmetry of a scramjet engine, which comprises the following steps of: releasing high-temperature airflow at the upstream of shock wave string head shock wave in an isolation section of the engine, or releasing and igniting auxiliary fuel in the isolation section of the engine, therefore, high-temperature airflow is formed at the upstream of shock waves of the shock wave string head. According to the embodiment of the invention, the flow field in the isolation section is reconstructed through combustion of high-temperature airflow or auxiliary fuel, so that the quality of the flow field at the inlet of the combustion chamber is improved, and the purpose of inhibiting combustion asymmetry of the combustion chamber is finally achieved.

A propulsion system includes a turbine engine and a ramjet engine. A turbine engine inlet section restriction device permits an inlet airflow to a turbine engine inlet in a first (low speed) operating state of the propulsion system, and restricts the inlet airflow to the turbine engine inlet in a second (high speed) operating state of the propulsion system. A combustor inlet flow control device, in the first operating state, provides an airflow from a compressor to the turbine engine combustor, and restricts an airflow from a ramjet bleed air passage, and, in the second operating state, provides a ramjet inlet bleed airflow to the turbine engine combustor, while restricting airflow from the compressor. In the second operating state, the turbine engine drives at least one accessory gearbox to drive at least one accessory component.

The turbine-based combined ramjet engine is characterized in that the turbine-based combined ramjet engine comprises an inner culvert channel, and the inner culvert channel is provided with a core diffuser, a core engine, a low-pressure turbine, an afterburner and an injection spray pipe; the core diffuser, the core machine, the low-pressure turbine, the afterburner and the injection spray pipe are sequentially arranged in the airflow direction. The outer culvert channel surrounds the radial outer side of the inner culvert channel, the outer culvert channel is provided with a mode selection valve, an outer culvert diffuser, a ramjet combustion chamber and a single-side expansion spray pipe, and the mode selection valve, the outer culvert diffuser, the ramjet combustion chamber and the single-side expansion spray pipe are sequentially arranged in the airflow direction; the mode selection valve is used for controlling opening and closing of the outer culvert channel; and the supersonic through-flow fan is arranged on the upstream sides of the inner culvert channel and the outer culvert channel.

The invention discloses a three-dimensional S-shaped bent inner concave runner design method and system. The design method comprises the following steps that a runner inlet molded line and a runner outlet molded line of a supersonic speed stamping runner are obtained; n discrete points are extracted in the circumferential directions of the inlet molded line and the outlet molded line to form N groups of inlet and outlet discrete point pairs; the change rule of S-shaped curve parameters is set in the circumferential direction of the runner inlet molded line / runner outlet molded line, and N S curves are obtained; acquiring a coordinate set of N * M discrete points on the curve set by M parallel interception planes; enveloping to generate an S-shaped curved surface, calculating the area of the two-dimensional cross section graph, and obtaining a flow channel sectional area change rule in the flow direction and a slope change rule of a center curve of the S-shaped curved surface; and judging whether the design is reasonable. The method can be suitable for computer-aided design of flow channels transited from various patterns including irregular patterns to circles or ellipses, and can meet the design of flow channels with offset inlet and outlet positions.

The invention relates to the technical field of engines, and discloses a turbine-based combined cycle engine operation method and a turbine-based combined cycle engine, and the turbine-based combined cycle engine operation method comprises the steps that a pulse detonation combustion mode is adopted in a pure turbine mode; a rotary detonation combustion mode is adopted in a pure stamping mode; a pulse detonation combustion mode is adopted in the transition mode, and meanwhile, a rotary detonation combustion mode is adopted. By utilizing the advantages of self-pressurization and high unit thrust of pulse detonation combustion, the high-speed performance is improved, and the working Mach number upper limit of a pure turbine mode is widened; by utilizing the advantages of high combustion speed, high combustion efficiency, simple structure and short length of rotary detonation combustion, the lower limit of the working Mach number of a pure stamping mode is widened. And in the transition mode, the characteristics of a pure turbine mode and a pure stamping mode are achieved at the same time, and enough thrust is provided for the aircraft.

The invention relates to a combustion chamber of a scramjet engine, which adopts kerosene as a cooling medium to realize phase change sweating cooling of the combustion chamber (7). The kerosene comes from a propellant of a scramjet engine and is stored in the hypersonic aircraft; a supporting plate (10) is arranged in the combustion chamber (7), a fixed wall surface (1) and a porous wall plate (2) are arranged on the supporting plate (10), the porous wall plate (2) is embedded and fixed in the fixed wall surface (1), and an increasing bag cavity (4) is formed in the inner side of the porous wall plate (2); according to the combustion chamber, phase change sweating cooling of the combustion chamber is achieved, meanwhile, the temperature of kerosene serving as fuel is increased, and the afterburning effect is achieved.

The invention belongs to the technical field of aero-engine design and calculation, and relates to a calculation method for stress of inlet, engine and exhaust parts of a subsonic combustion ramjet, which can obtain axial thrust of an air inlet, a combustion chamber and an exhaust nozzle of the ramjet according to working states of the engine at different speeds and heights and related parameters of the ramjet provided by a power unit. The decomposition of engine part force is realized; after the axial force of the air inlet channel, the engine combustion chamber and the exhaust nozzle is obtained, on one hand, the axial force is provided for the structure specialty, and the structure specialty can conveniently design an aircraft structure frame beam according to the ramjet force transmission requirement; and on the other hand, the aircraft longitudinal moment caused by axial stress of the combustion chamber of the ramjet engine is calculated by the operation and stability specialty in combination with the gravity center position of the aircraft, operation and stability design and check are carried out, and the method can finally meet the requirements of the aircraft for structural design and flight quality calculation related to the ramjet engine.

The invention relates to the technical field of rocket engines, in particular to a rocket support plate and concave cavity combined oblique detonation engine which comprises a mixing section, a combustion chamber and an exhaust nozzle which are sequentially connected, the mixing section is of a hollow structure, a fuel-rocket support plate assembly is arranged in the mixing section, and a detonation slope-concave cavity combined support plate assembly is arranged in the combustion chamber; a middle flow channel for constant-pressure combustion is formed in the middle of the detonation slope-concave cavity combined supporting plate assembly, and an upper flow channel and a lower flow channel for constant-volume combustion are formed by the upper surface and the lower surface of the detonation slope-concave cavity combined supporting plate assembly and the inner wall of the combustion chamber. According to the device, an oblique detonation combustion mode, a rocket mode, a subsonic combustion ramjet mode and a scramjet mode are integrated into the same engine, a split flow channel structure is formed, full-speed domain coverage from low Mach to ultrahigh Mach (more than 8 Mach) is achieved, and combustion organizations in different flight stages are optimized through a mode switching mechanism.

The invention provides a ramjet air turbine pump shell assembly welding forming process which comprises a welding method combining turbine stator non-blade end TIG welding and turbine stator blade end brazing and an aging heat treatment procedure in the welding process. The defects that in the turbine shell assembly production process, the turbine stator blade end laser welding or TIG welding difficulty is large, deformation of the two ends of a product is large due to large residual stress after welding, welding seams crack, and the service life is short are effectively overcome, and welding quality control over the turbine shell assembly is achieved; and the product consistency and reliability of the turbine shell assembly are greatly improved.

The invention belongs to the technical field of design of turbine-based ramjet combination engines, and particularly relates to a turbine-based ramjet combination engine and a design method thereof. A turbine engine and a subsonic combustion ramjet engine are designed to be connected in parallel to intake air and share a thrust augmentation / ramjet combustion chamber; main advantages of a series-connection type turbine-based ramjet combination engine and a parallel-connection type turbine-based ramjet combination engine are integrated, and compared with the series-connection type turbine-based ramjet combination engine, mode selection valves can be arranged at inlets of an air inlet channel low-speed channel and an air inlet channel high-speed channel when the series-connection type turbine-based ramjet combination engine works in a ramjet engine mode; the high Mach number working reliability is good, the efficiency is high, compared with a parallel type turbine-based ramjet combined engine, a set of ramjet combustion chamber and a combined spray pipe channel are reduced, the weight can be reduced by 12%-15%, the mission load and transonic speed performance of an aircraft can be improved, and the control difficulty is reduced.

The invention discloses an afterburner adopting a gas-liquid bimodal-cross hole spray rod for oil supply, and relates to the technical field of aero-engines. Comprising a turbine located at the axis position of an inner duct in the air incoming flow direction; the mixer is arranged between the turbine and the inner duct wall body; the rectification support plates are provided with air inlet channels, and the multiple rectification support plates penetrate through the outer duct and are vertically and uniformly arranged in the circumferential direction of the inner duct; the oil injection rod is provided with an air flow channel and a fuel oil channel, and the air inlet channel is communicated with the air flow channel; and the oil injection rod is radially arranged in the rectification support plate. On the basis that the structural complexity of an afterburner of a turbine engine and a combustion chamber of a subsonic combustion ramjet engine is not increased, the atomization performance and the oil-gas mixing uniformity of fuel oil in the afterburner / ramjet combustion chamber are remarkably improved, the combustion efficiency of the engine is improved, and the combustion instability is reduced.

The invention discloses a nozzle matching method for cross-medium stamping, and relates to the field of cross-medium stamping engine nozzles. According to the matching method, on the basis of actual performance parameters of an existing air ramjet and water ramjet engine, the detailed reaction mechanism of combustion is considered in design calculation, and an extensible spray pipe scheme is used for solving the achieving problem of change of the expansion ratio of the air ramjet and water ramjet engine. According to the method, the design process of an air inlet channel and a combustion chamber of a conventional cross-medium ramjet engine is supplemented, and the integrated design of the cross-medium ramjet engine is perfected.

The present invention discloses a strut-type supersonic combustor coupled with wall-swept nozzles, relating to the technical field of scramjet engines. The combustor comprises a central strut and multiple wall-swept nozzles. One end of the combustor body serves as an air inlet, and the other end serves as an air outlet. The central strut is fixedly mounted within the combustor body and is provided with multiple fuel injection holes for injecting fuel into the combustor body. Multiple sweep nozzles are circumferentially arranged on the inner wall of the combustor body, each sweep nozzle being configured to sweep fuel into the combustor body. The strut-type supersonic combustor coupled with wall-swept nozzles provided by the present invention enhances fuel diffusion and mixing, thereby improving combustion efficiency.

This invention discloses a one-dimensional modeling method for scramjet engine volumetric dynamics, comprising the following steps: 1) establishing a mathematical model of the scramjet engine's inlet, combustion chamber, and tail nozzle; 2) establishing mathematical models of the scramjet engine's scramjet shockless mode, scramjet oblique shock mode, and subcombustion mode; and 3) establishing a dynamic model of the scramjet engine based on volumetric dynamics. This invention utilizes volumetric dynamics to establish a one-dimensional scramjet engine model, enabling rapid calculation of changes in combustion chamber axial flow field parameters, meeting the requirements of engine control system design.

This invention belongs to the field of ramjet technology and provides a wide-speed range ramjet fuel supply system and method driven by a pyrolysis oil-gas turbine. The system comprises a fuel tank, a centrifugal pump, a motor, an oil-gas turbine, a battery, an electromagnetic control module, fuel lines, a pyrolysis fuel vapor line, a combustion chamber, and a regenerative cooling channel on the outer wall of the combustion chamber. Through the connection between the centrifugal pump, the regenerative cooling channel, and the oil-gas turbine, the system utilizes the cracking of hydrocarbon fuel into small-molecule pyrolysis kerosene vapor to drive the turbine rotor to produce work. The motor then assists in starting the ramjet, thereby resolving the power generation challenges of the ramjet mode and alleviating thermal protection issues.

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.