72 results about "Scramjet engine" patented technology

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 relates to the technical field of scramjet engines, in particular to a supersonic combustion chamber and a scramjet engine. The supersonic combustion chamber comprises a combustion chamber body, a support plate combustion stabilizer, a first wall surface fuel injection device and a second wall surface fuel injection device. The combustion chamber body comprises an isolation section, an expansion section and an equal straight section which are sequentially connected. The supporting plate combustion stabilizer is arranged on the inner wall of the expansion section, the first wall face fuel injection device is arranged on the upstream of the supporting plate combustion stabilizer, and the second wall face fuel injection device is arranged on the downstream of the supporting plate combustion stabilizer. An oil supply layout of one-way support plate oil supply and two-way wall surface oil supply is formed, one-way wall surface oil supply is located at the upstream of the support plate oil supply, multi-stage energy release in the combustion chamber body can be effectively improved, the other-way wall surface oil supply is located at the downstream of the support plate oil supply, the height of the combustion chamber body at the downstream of the support plate combustion stabilizer can be effectively covered, and the combustion chamber body is more stable. Therefore, stable work of the supersonic combustion chamber is ensured.

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 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.

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 invention discloses an intelligent solution method for a supersonic combustion flow field based on hierarchical physical constraints, and relates to the crossing field of combustion simulation calculation and deep learning of a supersonic combustion chamber, and the method comprises the steps: S1, verifying the reliability of simulation data obtained through supersonic combustion flow numerical simulation calculation; s2, constructing a three-dimensional combustion flow field data set based on the final turbulence model parameters obtained in S1; s3, constructing a physical mechanism model according to the generation principle of the combustion flow of the scramjet engine; s4, constructing a neural network model according to the three-dimensional combustion flow field data set; and S5, performing staged constraint on the prediction result of each layer of the neural network model by adopting a physical mechanism model to obtain a three-dimensional flow field intelligent solution model. According to the method, high-precision generation of the three-dimensional supersonic combustion flow evolution process is guaranteed, meanwhile, the generation speed of the model is further increased, and technical support is provided for accelerating the numerical simulation process of the scramjet engine and promoting the optimization process of the scramjet engine.

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

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

This invention relates to a turbulent combustion simulation method for temperature and pressure adaptation in scramjet engines, comprising: constructing a flame front database; estimating the base temperature and pressure range of the target combustion flow field, wherein the base temperature includes: estimating the minimum and maximum temperatures at the oxidizer end and the minimum and maximum temperatures at the fuel end; determining four sets of estimated operating conditions; solving the flame front equations to obtain four sets of laminar flame front solutions; using a predefined PDF function to perform ensemble averaging on the four sets of laminar flame front solutions to obtain four sets of ensemble average solutions for the laminar flame front solutions; selecting a reference flame front database based on each set of ensemble average solutions, and obtaining the scaling factor of each database within the group relative to each reference flame front database; constructing a flame front database using the reference flame front database and the scaling factor; performing CFD calculations on the flow field of the scramjet engine, and updating the local flame front database with the flame front database after each calculation.

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.

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 discloses a scramjet engine inlet-engine coupling rapid calculation model and a high-safety control method of the scramjet engine inlet-engine coupling rapid calculation model. Establishing an air inlet flow characteristic and total pressure recovery coefficient characteristic database under different flight attack angles (-5 degrees to + 15 degrees) and Mach numbers (the cruise Mach number is more than 6), and constructing an air inlet real-time performance prediction model based on characteristic data; based on air inlet channel-isolation section flow field coupling analysis, the non-starting boundary of the air inlet channel is accurately calibrated in combination with the shock wave string position, the wall surface pressure gradient and the separation area evolution characteristics; a Min-Max layered limiting protection control strategy is adopted, the flow capture rate, the total pressure recovery coefficient and the air inlet supercharge ratio are monitored in real time, the back pressure adjusting amount of a combustion chamber is dynamically restrained, and it is ensured that key parameters are within a safety threshold value. According to the method, the problem that a traditional method is insufficient in adaptability to wide-area flight conditions is solved, and the anti-disturbance capacity of the air inlet channel and the working stability of the engine are remarkably improved.

This invention discloses an active cooling structure and application method for optimizing kerosene heat sink utilization, relating to the design of scramjet engines. The structure includes variable cross-section passage zones I, II, and III, a flow restrictor, a local manifold, and a spacer. By providing different heat exchange efficiencies through the variable cross-section passage zones I, II, and III, the invention adjusts the kerosene heat sink utilization rate according to the actual thermal load on the cooling panel, improving the temperature uniformity of the cooling panel structure and effectively guiding the engine's thermal structure design.

The application discloses an engine test heterogeneous nozzle determination method and device and electronic equipment, and relates to the technical field of liquid rocket engines. The method comprises the following steps: determining the inner surface of the device nozzle by adopting the streamline tracking mode; determining the inner wall surface structure of the device nozzle by extending the preset size outward based on the inner surface of the device nozzle; setting the reinforcing rib structure at the first preset position of the nozzle aerodynamic surface and setting the liquid collecting cavity at the second preset position of the nozzle aerodynamic surface; and setting the flanges at the inlet of the device nozzle aerodynamic surface and the outlet of the device nozzle aerodynamic surface respectively according to the test bench inlet size and the combustion chamber inlet structure size, so as to obtain the engine test heterogeneous nozzle. The inner surface of the device nozzle obtained by the streamline tracking mode guarantees the smooth outlet and the good matching performance of the heterogeneous combustion chamber, can effectively simulate the high Mach number coming flow condition, and can solve the problem of difficult simulation of the coming flow of the heterogeneous scramjet engine combustion chamber.

A health monitoring system and adaptive adjustment method for a scramjet engine cooling channel, relating to the field of aero-engine management and condition monitoring technology, is disclosed. The system includes: a condition perception and monitoring module, which uses fiber optic measuring points deployed within the cooling channel to perceive multi-physical field information such as temperature, pressure, and vibration in real time, achieving in-situ, passive, and interference-resistant measurement; a data transmission and conversion module, which converts optical signals into digital signals; an intelligent analysis and decision-making module, which analyzes the cooling channel condition based on real-time data and historical models, identifies abnormal conditions such as overheating, coking, thermal stress concentration, and flow instability, and generates corresponding pump and valve adjustment commands; and a strategy execution and adjustment module, which dynamically adjusts the cooling medium flow rate by regulating pump speed or valve opening, achieving closed-loop adaptive adjustment of the cooling channel condition and improving engine thermal protection efficiency and operational reliability. This invention is applicable to the integrated and intelligent development needs of future hypersonic aircraft.

The application provides a hypersonic scramjet cooling channel optical fiber sensor fixing and integrated installation structure and belongs to the technical field of sensors; wall embedded measurement array is used to realize in-situ distributed measurement of internal temperature, pressure and strain parameters of the cooling channel; the fourth group of optical fibers and the fifth group of optical fibers are arranged on four edges of the rectangular channel outlet of the outlet integrated measurement array, eight optical fibers are used to realize ultra-high density monitoring of the outlet area, the optical fiber lead is provided with a redundant length and is protected by a heat insulation sleeve, and the reliability in an extreme environment is ensured. The application comprises a wall embedded measurement array, micro grooves or micro through holes are formed on the inner wall of the rectangular cooling channel and the outer wall of the rectangular cooling channel, and the packaged optical fiber sensor is fixed in a flush or slightly convex manner; an outlet integrated measurement array is arranged in the outlet area of the rectangular cooling channel; and the wall embedded measurement array and the outlet integrated measurement array are connected through optical fiber leads.

The present application relates to the performance evaluation method and equipment of the super-combustion ramjet engine based on direct connection test, wherein the method comprises: obtaining the first test bench thrust-time curve, the first flow parameter and the first pressure parameter of the super-combustion ramjet engine under the free jet test working condition; setting the inflow condition of the super-combustion ramjet engine during the ground direct connection test based on the first airflow parameter, taking the first airflow parameter as the airflow parameter of the selected second engine characteristic section, and taking the second airflow parameter as the airflow parameter of the fuel inlet section of the combustion chamber of the super-combustion ramjet engine; obtaining the second test bench thrust-time curve, the second flow parameter and the second pressure parameter of the super-combustion ramjet engine after the second engine characteristic section under the ground direct connection test working condition; and evaluating the performance of the super-combustion ramjet engine based on the second test bench thrust-time curve, the second flow parameter and the second pressure parameter.

The invention discloses an intelligent control method and system for multi-point fuel injection of a scramjet engine. The method comprises the steps that a multi-core parallel computing architecture is adopted to build a simulation model, and a multi-core parallel computing one-dimensional model is built; the security constraint is fused into a learning process, and a Safe SAC intelligent agent considering the security constraint is constructed; reinforcement learning multi-point intelligent injection pre-training based on multi-core parallel computing one-dimensional model interaction is carried out, and a pre-trained Safe SAC reinforcement learning agent is obtained; designing a thrust indirect control law based on the wall surface pressure; calculating the thrust of the engine according to the wall surface pressure; deploying a pressure sensor on the wall surface of the isolation section, and calculating the front edge position of a shock train in the isolation section by adopting a pressure ratio method; and the engine thrust and the shock train leading edge position are input into a pre-trained Safe SAC reinforcement learning agent, and the fuel flow of multiple oil injection points is controlled. According to the invention, intelligent regulation and control of multi-point fuel injection of the combustion chamber are realized under the safety constraint condition of the aircraft.

An experimental model based on plasma synthetic jet array excitation is provided to promote the enhancement of cavity mixing. A wall shear stress sensor is arranged upstream of the leading edge of the cavity, a wall shear stress sensor is arranged on the rear vertical surface of the cavity, and a dynamic pressure sensor is arranged on the bottom surface of the cavity. By applying amplitude and phase modulation to different rows or columns of plasma synthetic jet exciters, a traveling wave standing wave excitation method is realized to intelligently enhance the mixing of the cavity. First, a plurality of groups of excitation parameters are randomly initialized and the flow control benefits are evaluated; second, a proxy model is constructed to predict the best combination of excitation parameters by means of an intelligent optimization algorithm; finally, the discharge trigger signal sequence of the power supply is solved in reverse, and the benefits are improved through iterative testing. The plasma synthetic jet traveling wave standing wave excitation method has higher penetration depth, stronger mixing ability, better adaptability, and is more intelligent, and can better meet the needs of flame stable combustion under the wide working boundary of a scramjet engine.

The present application relates to a kind of strong flame combustion environment under scramjet speed field measurement system and method, belong to aerospace engine testing and flow field measurement technical field, comprising: first camera, second camera, particle generator, engine combustion chamber and sheet light beam;Through system construction, regulate timing to realize the extremely short exposure of nanosecond level of double camera wrong time, and the light of laser double cavity is respectively matched with the exposure period of double camera;Again, calibration image is collected to solve perspective conversion matrix and error matrix, and flow field image is synchronously collected and preprocessed is completed;Subsequently, through cross-correlation algorithm, original velocity field is solved, and after outlier rejection, bilinear interpolation and error matrix correction, actual velocity field is obtained by combining calibration coefficient conversion.The measurement method of the present application can accurately capture and analyze the dynamic change of the flow field inside the combustion chamber under the strong flame combustion environment, which provides important support for the research and development of scramjet.

The application provides a super-combustion ramjet engine spanwise pulse injection device, which comprises a spanwise injection cylinder and at least two pulse injection assemblies in parallel and communicated with the spanwise injection cylinder, the pulse injection assembly comprises a first valve body, a second valve body, a movable valve core arranged at an outlet end of the second valve body and a nozzle seat which are sequentially arranged along a fuel flow direction; at least two buffer cavities which are symmetrical along a second valve body axis are arranged at a connection position of a third valve seat and a second valve seat, a buffer elastic member is arranged in the buffer cavities, and a pressure relief pipeline which is communicated with the second cavities is arranged at one end of the buffer spring close to the buffer cavities. Advantageous effects: through the design of four buffer cavities, a pressure relief pipeline and a buffer elastic member, when the pressure in the pipeline changes, the buffer elastic member is compressed or rebounds, the water hammer effect is effectively reduced, the rate of fuel entering the second cavities from the pipeline is accelerated, the injection rate is further accelerated, the pulse lag phenomenon is reduced, and the pulse frequency is improved.

The invention belongs to the technical field of engine parameter calculation, and particularly relates to an air suction cruise hypersonic aircraft nozzle parameter equivalent estimation method, which comprises the following steps of: enabling an air inlet channel, an isolation section and a combustion chamber of a scramjet engine to be equivalent to a combustion chamber of a rocket engine; the contraction ratio of the jet pipe of the scramjet engine is equivalent to the expansion ratio of the jet pipe of the rocket engine; the method comprises the following steps: enabling air flow and fuel injection flow entering a combustion chamber of the scramjet engine to be equivalent to premixed gas which is uniformly mixed according to a preset proportion, and determining formula parameters of an equivalent propellant based on the premixed gas; converting the flow field parameters at the air inlet of the scramjet engine into the total temperature and the total pressure of a combustion chamber of an equivalent rocket engine; calculating nozzle parameters of the scramjet engine by adopting a thermodynamic calculation method based on the parameters obtained in the above steps; the problems that an existing scramjet engine nozzle parameter is large in solving difficulty and high in calculation cost are solved.

The invention discloses a combustion mode judgment method for a scramjet engine combustion chamber, which comprises the following steps of: firstly, acquiring combustion field information in the combustion chamber by utilizing a flow field sensing technology or a CFD (computational fluid dynamics) numerical simulation means, including the average temperature of a central axis at the rear end of the combustion chamber and the average flow speed in the combustion chamber; therefore, dimensionless parameters capable of reflecting the macroscopic kinetic energy and heat energy characteristics of the combustion chamber are constructed. Then, structural characteristic parameters such as the flame perimeter and the flame area are extracted and used for quantifying the diffusion degree and the morphological characteristics of the flame; by integrating the parameters to construct composite dimensionless parameters, accurate identification of different combustion modes can be realized. According to the method, key leading factors in the combustion process are fully reflected in parameter selection, and the combustion mode of the scramjet engine can be discriminated with high reliability and high precision.

The application relates to the field of aerospace power design and manufacturing, and discloses a multi-combustion mode combined hydrogen fuel aviation power system for a super-speed aircraft, which is based on a single-rotor turbojet engine and is additionally provided with an outer channel detonation combustion chamber and a scramjet engine after the turbine, is compact and simple in structure, is in series combination of various parts, can minimize the volume and weight of the power system, is more suitable for small and medium-sized aircrafts for super-speed, can meet the wide-speed-domain flight of the aircraft, can effectively utilize the high-speed and high-pressure airflow of the outer channel by adjusting the opening and closing mode of the nozzle with the suction hole to adapt to different working modes of the combined engine, and can meet the technical requirements of low carbon, high energy efficiency, high speed and long-time flight when the turbine engine works independently, the turbine engine and the detonation combustion chamber work together, the scramjet engine works independently and the scramjet and the turbine engine work together.

The invention discloses an automatic CFD solving method for an internal flow field of a bimodal scramjet engine, and the method comprises the steps: constructing an AnsysFluent automatic solving environment through employing a Python script, loading a grid file of a bimodal scramjet engine model, configuring a physical model and boundary conditions of the bimodal scramjet engine through employing the Python script, and carrying out the automatic solving of the internal flow field of the bimodal scramjet engine. A Python script is utilized to complete setting of a numerical discrete format and a solving method and construction of a flow field initial value of a computational domain; taking the flow field initial value as an iteration starting point, and executing dynamic climbing calculation under first-order precision by utilizing a Python script to obtain stable initial flow field distribution; the stable initial flow field distribution is used as an initial calculation value, and a Python script is used for switching second-order high-precision calculation and carrying out convergence judgment. According to the method, the problems of tedious manual intervention, difficulty in shock wave capture and subjective convergence judgment in high-Mach-number strong compressible flow calculation are effectively solved.

The invention relates to the technical field of supersonic air inlet channel starting and active flow control, aims to solve the problems of air inlet channel flow separation, shock wave boundary layer interference and starting delay of an existing scramjet engine, and provides a miniature supersonic spray pipe array device for enhancing air inlet channel starting and an engine. The miniature supersonic spray pipe array device comprises a rotating mechanism, and a plurality of spray pipes are mounted on the rotating mechanism and arranged at intervals; an air flow channel in each spray pipe is an acceleration flow channel, and an air flow channel between every two adjacent spray pipes is a speed reduction flow channel; the acceleration flow channel is sequentially divided into an acceleration flow channel inlet section, a middle contraction section and an acceleration flow channel outlet section in the gas flowing direction, and the section size of the middle contraction section is smaller than the section size of the acceleration flow channel inlet section and the section size of the acceleration flow channel outlet section; according to the invention, the problems of shock wave boundary layer interference, flow separation and starting delay of the lip of the air inlet channel under the condition of high Mach number are effectively solved.

A coreless non-circular rotating detonation scramjet engine, provided by the present invention, comprises: a combustion-type air heater; a shape transition nozzle connected to the combustion-type air heater; an isolator connected to the shape transition nozzle; a combustor connected to the isolator and comprising a cavity having an obround cross section and an expansion nozzle; multiple fuel injection ports formed in a ring shape along the outer circumferential surfaces of the isolator, the cavity, and the expansion nozzle; and an igniter for igniting a compound located in the cavity.

This invention relates to the field of supersonic inlet start-up and active flow control technology, aiming to solve the problems of inlet flow separation, shock wave boundary layer interference, and start-up lag in existing scramjet engines. It provides a miniature supersonic nozzle array device and engine for enhancing inlet start-up. The miniature supersonic nozzle array device includes a rotating mechanism on which multiple nozzles are mounted, arranged at intervals. The airflow channel inside each nozzle is an acceleration channel, and the airflow channel between two adjacent nozzles is a deceleration channel. The acceleration channel is sequentially divided into an acceleration channel inlet section, a central contraction section, and an acceleration channel outlet section along the gas flow direction. The cross-sectional size of the central contraction section is smaller than that of the acceleration channel inlet section and the acceleration channel outlet section. This invention effectively solves the problems of shock wave boundary layer interference, flow separation, and start-up lag at the inlet lip under high Mach number conditions.