241 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 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 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 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.

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 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 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.

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 present invention discloses a method for predicting the ignition of a scramjet combustion chamber, belonging to the technical field of ignition prediction. The method comprises the following steps: numerically simulating the combustion chamber using a three-dimensional density-based steady-state injection; determining whether convergence conditions have been met; extracting flow field data of the combustion chamber after convergence; and calculating the residence time t of aviation kerosene microclusters in the ignition region based on the flow field data. z ; Calculate the ignition delay time t of the ignition area based on the flow field data c Calculate Da in the ignition area. Predict ignition conditions based on Da. The scramjet engine combustion chamber ignition prediction method of the present invention calculates Da based on injection field data and predicts ignition success based on the relationship between Da and 1. This helps adjust actual ignition conditions and improve ignition success rates. It can meet the ignition prediction needs of combustion chambers of various configurations.

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 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.

The present invention relates to a rocket-based combined cycle (RBCC) engine with a parallel compressor flow path, which includes a compressor flow path and an RBCC engine flow path. The compressor flow path and the RBCC engine flow path are arranged in parallel, and the two flow paths share an inlet duct. After the inlet duct flow path, a rocket thrust chamber, a primary combustion chamber, a secondary combustion chamber and a nozzle are provided; the airflow in the compressor flow path converges into the secondary combustion chamber and burns in the secondary combustion chamber.

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 invention discloses a central water injection type water ramjet engine which comprises a front end socket, a combustion chamber, a rear end socket, an afterburning chamber and a spray pipe which are connected in sequence. The front explosive blocking ring is arranged at the position of the front sealing head, the rear explosive blocking ring is assembled on a boss of the rear sealing head, the combustion chamber is filled with the propellant grain, the jetting assembly is arranged on the central axis in the combustion chamber, and the ignition assembly is arranged at the position of an ignition connector of the rear sealing head. The engine is integrally assembled in the underwater navigation body to serve as a propelling system. The central injection assembly is adopted, the structure is simple, installation is easy, and the manufacturing cost is reduced; water is directly injected into the center of the combustion chamber through central water injection, so that the water and fuel can be rapidly mixed, the reaction is more sufficient, and the combustion efficiency can be greatly improved.

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).