97 results about "Scramjet engine" patented technology

The invention relates to the technical field of engine thermal management, and particularly provides an active cooling method for coupling rotational flow cooling and regenerative cooling, which comprises the following steps: arranging a rectangular auxiliary inlet channel on the side wall of a rectangular main inlet channel of a regenerative cooling channel, and constructing a rotational flow regenerative cooling channel; when the rectangular main inlet channel performs regenerative cooling, coolant induced rotational flow is injected into the rectangular auxiliary inlet channel, and the rotational flow strength is regulated and controlled by adjusting the ratio of the flow of the coolant induced rotational flow to the total flow; and a cooling process simulation model is constructed, and the cooling effect of the rotational flow regeneration cooling channel is analyzed based on simulation calculation data. By means of the innovative rotational flow regeneration cooling channel design, controllable rotational flow is introduced into a traditional regeneration cooling channel, and the problems that a scramjet engine combustion chamber is insufficient in heat transfer in the extremely high-temperature environment and the fuel chemical heat sink utilization rate is low are solved. And fluid mixing is enhanced through rotational flow, a thermal boundary layer is destroyed, the fuel pyrolysis reaction is optimized, and efficient heat exchange is achieved.

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 center sudden expansion step stable combustion supersonic combustion chamber and a combustion method, and belongs to the technical field of supersonic combustion chambers. The center sudden expansion step stable combustion supersonic combustion chamber comprises an expansion section and an equal straight section, one end of the expansion section is connected with an isolation section of a scramjet engine, and the other end of the expansion section is connected with the equal straight section; a wedge-shaped supporting plate is arranged in the expansion section, penetrates through the expansion section and forms a center sudden expansion step structure in the expansion section, a mounting hole for mounting an igniter is formed in the center of the end, close to the equal straight section, of the expansion section, and the mounting hole is located right behind the supporting plate; the expansion section is provided with a plurality of wall face fuel oil injection hole arrays, and the wall face fuel oil injection hole arrays are located above the supporting plate. By the adoption of the supersonic combustion chamber with the center sudden expansion step stable combustion and the combustion method, the problems that an existing supersonic combustion chamber is difficult in stable combustion and needs to carry oxygen, and application is limited can be solved.

The invention discloses a method and a system for measuring a time average velocity field of a combustion chamber of a scramjet engine, and belongs to the technical field of average velocity field measurement. The method comprises the following steps: establishing a measurement system, matching the number of exposure frames of a high-speed camera with the number of cavities of a laser, collecting a time sequence image for pre-processing, obtaining the pre-processed time sequence image, dividing the time sequence image into grid slices for cross-correlation calculation, and obtaining a time sequence speed according to a space and pixel conversion coefficient; constructing a velocity model according to a Gaussian mixture model algorithm, and obtaining a probability matrix of a data point belonging to a correct velocity vector according to an expectation maximization algorithm; and calculating to obtain a time average velocity field according to the correct velocity vector and the error velocity vector judgment condition. The method effectively improves the efficiency and accuracy of measuring the distribution of the time average velocity field, and can be used for measuring the time average velocity field.

The invention provides an ammonia fuel cell power generation dual-fuel scramjet engine capable of achieving energy gradient utilization, belongs to the technical field of aircraft thermal protection and power generation, and solves the problems that a traditional airborne power generation system is difficult to use due to the high incoming flow Mach number in hypersonic flight, and coking is caused in a regeneration cooling channel of a fuel decomposition gas turbine. Ammonia and aviation kerosene generate high-temperature fuel gas in a combustion chamber of the scramjet engine to push the engine to generate thrust; high-temperature waste heat of the combustion chamber is recycled through the regeneration cooling channel, fuel is heated, a part of ammonia is decomposed, and decomposed gas drives a turbine generator to generate electric energy. According to the power generation system for gradient utilization of energy, factors such as actual use temperature and pressure are fully considered, the energy utilization efficiency is effectively improved, energy waste is reduced, and reasonable adaptation of energy is achieved, so that the power generation efficiency of the whole system is improved, and the electric energy requirement of a hypersonic aircraft is greatly met.

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.

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 invention discloses a scramjet combustion chamber without a concave cavity and a supporting plate and a combustion stability enhancing method, the combustion chamber comprises a combustion chamber main body and a main injection mechanism arranged on the wall surface of the combustion chamber main body, the main injection mechanism is used for injecting main fuel jet flow; the auxiliary injection mechanism is located on the upstream of the auxiliary injection mechanism and used for injecting auxiliary fuel jet flow, and plume of the auxiliary fuel jet flow forms a pneumatic type backflow area under the retarding effect of the main fuel jet flow; the ignition device is located between the main fuel jet flow and the auxiliary injection mechanism and used for igniting the auxiliary fuel in the pneumatic type backflow area, and the main flow fuel is continuously preheated and ignited while stable combustion of the auxiliary fuel is achieved. The combined jet flow combustor is applied to the field of scramjet engines, supersonic stable combustion is achieved on the basis of the pneumatic type backflow area formed through combined jet, meanwhile, the overall combustion stability is greatly enhanced through the continuous ignition and preheating effect of the pneumatic type backflow area on downstream main fuel jet flow, and combustion oscillation in the flow direction is greatly weakened.

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 provides a scramjet engine natural vibration co-frequency fuel pulse injection system and method, a fuel injection device and a self-adaptive opening and closing air inflow device.The fuel injection device comprises a driving device and a fuel conveying rotor shaft, and the output end of the driving device is fixedly connected with the fuel conveying rotor shaft; and the self-adaptive opening and closing air inlet device is communicated with an air inlet of the driving device. The method comprises the steps that 1, the air inlet channel starts to capture air; 2, the natural vibration frequency of an engine is obtained, and the natural vibration frequency information of the engine is fed back to a driving motor of the self-adaptive opening and closing air inlet device; 3, the driving motor drives the sealing piece set to act, and the opening degree of the corresponding rotating speed of the pneumatic motor is adjusted; 4, pulse frequency information of fuel injection is obtained; and 5, the preset program starts phase matching work, and finally the phase is adjusted to the phase where the fuel pulse injection wave crest and the natural vibration pressure signal wave crest are overlapped.

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.

The present invention discloses an injection configuration for a fuel injection module of a scramjet engine, comprising a fuel injection assembly, wherein the fuel injection assembly is provided with a row of fuel injection holes, the direction of the center line of the fuel injection holes is perpendicular to the flow direction of the airflow in the combustion chamber, and the aperture of the fuel injection holes gradually decreases from the center of the fuel injection assembly to both sides; beneficial effect: the present invention makes the jet penetration depth of each fuel injection hole consistent by changing the aperture of the fuel injection holes, the incoming flow in the central area of ​​the combustion chamber can be fully utilized, the fuel is evenly distributed in the combustion chamber, the mixing effect of the fuel and the high-speed incoming flow is enhanced, and the components are more uniform when entering the downstream concave cavity of the combustion chamber for ignition and combustion, thereby improving the combustion efficiency.

System and methods for shaping a waverider inlet manifold are provided. In implementations, a method of controlling the shape of a shapeable waverider inlet manifold for a hypersonic scramjet includes: measuring a flight Mach number of the hypersonic scramjet; comparing the measured Mach number to a current Mach-number-specified-by-position; adjusting an inlet stream surface of the shapeable waverider inlet manifold towards a desired disposition based on the comparing; iteratively ensuring that the flight Mach number matches the Mach-number-specified-by-position by correcting the disposition of the inlet stream surface until a difference between the Mach numbers is within a predetermined tolerance.

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 present invention discloses a controllable scramjet engine and its operating method, belonging to the field of scramjet engines. The engine comprises a supersonic combustion chamber, a controllable gas generator disposed on the outer wall of the supersonic combustion chamber, the output end of the controllable gas generator being connected to the input end of the supersonic combustion chamber via a regenerative cooling pipe disposed on the wall of the tail nozzle of the supersonic combustion chamber; the controllable gas generator comprises a main body, fuel filled within the main body, and an electric pyrolysis mechanism for pyrolyzing the fuel. Using the above-mentioned controllable scramjet engine and its operating method, pyrolysis gas is used as fluidizing gas to carry powdered fuel for injection, effectively enhancing the engine's specific impulse performance. The gas fuel can also serve as a pilot fuel, first burning to increase pressure and temperature, and then carrying a flow to eject high-energy powder particles, thereby solving the problem of short residence time of powdered fuel in the combustion chamber and low combustion efficiency.

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.

The present invention discloses a method and system for measuring the time-averaged velocity field of a scramjet combustion chamber, belonging to the technical field of average velocity field measurement. The present invention constructs a measurement system, matches the exposure frame number of a high-speed camera with the cavity number of a laser, collects time-series images for pre-processing, obtains the pre-processed time-series images, divides them into grid slices for cross-correlation calculation, and obtains the time-series velocity based on the spatial and pixel conversion coefficients; constructs a velocity model based on a Gaussian mixture model algorithm, and obtains a probability matrix of a data point belonging to a correct velocity vector based on an expectation maximization algorithm; and calculates the time-averaged velocity field based on the correct velocity vector and the error velocity vector judgment conditions. The present invention effectively improves the efficiency and accuracy of measuring the time-averaged velocity field distribution and can be applied to measuring the time-averaged velocity field.

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.