14 results about "Scramjet engine" patented technology

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.

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.

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.

This invention discloses an experimental apparatus and method for studying the ignition and combustion mechanism of condensed-phase particles under supersonic airflow, relating to the field of solid scramjet engine technology. The apparatus includes an air heater, a nozzle, an isolation section, and a combustion chamber connected in sequence. The airflow, after being rectified by the isolation section, enters the combustion chamber and reacts with gaseous fuel injected through the gaseous fuel inlet, providing a high-temperature, high-pressure combustion environment for the condensed-phase particles. A high-speed camera records the ignition and combustion process of the condensed-phase particles in detail through a high-speed camera lens.

The present disclosure relates to the technical field of engines, and particularly provides a flow passage variable-width speed range ramjet engine, a combined power device and an aircraft. The engine body is provided with a first flow passage and an adjusting assembly. The bottom wall of the first flow passage is fixed, and the upper wall is adjustable. The upper wall has two states. In the first state, the first flow passage is a double-mode scramjet engine flow passage. In the second state, the first flow passage is a common subsonic ramjet engine flow passage. The upper wall comprises a variable-length sleeve plate, a front adjusting plate, a middle adjusting plate and a rear adjusting plate which are sequentially hinged. One end of the variable-length sleeve plate away from the front adjusting plate is hinged to the engine body. One end of the rear adjusting plate away from the middle adjusting plate is hinged to a connecting piece. The connecting piece is slidingly installed on a first sliding rail. The first sliding rail is fixedly arranged on the engine body. The present disclosure greatly widens the lower limit of the effective working speed range of the ramjet engine and reduces the technical difficulty of combination with a turbine engine.

The application discloses a plasma intelligent control system and method for shock wave of a separation section, and relates to the technical field of a scramjet engine.The flat plate is erected on a support seat through a single truss, and a double truss is arranged on each side of the single truss; one end of the double truss is connected to the support seat, and the other end is connected to an upper sharp wedge mechanism; a first pressure sensor module, a distributed plasma exciter module and a second pressure sensor module are sequentially arranged on the flat plate in the direction from the upstream to the downstream; the upper surface of the upper sharp wedge mechanism is provided with an inclination angle; the lower surface of the upper sharp wedge mechanism is parallel to the upper surface of the flat plate; and the upper sharp wedge mechanism is located above the first pressure sensor module and the second pressure sensor module.The application captures the precursor of instability in advance and drives the excitation action, thereby achieving active and advanced stability maintenance in structure and replacing the traditional passive response type control.

The application discloses a kind of dual-mode scramjet internal flow field automation CFD solving method, comprising: using Python script to build AnsysFluent automation solving environment, load the grid file of dual-mode scramjet model, using Python script to configure the physical model and boundary condition of dual-mode scramjet, using Python script to complete the setting of numerical discrete format, solving method and the flow field initial value construction of calculation domain;Flow field initial value is used as the starting point of iteration, using Python script to execute the dynamic climbing calculation under first-order accuracy, obtain stable initial flow field distribution;Stable initial flow field distribution is used as the initial value of calculation, using Python script to switch second-order high-precision calculation and carry out convergence determination.This method effectively solves the problems of artificial intervention tedious, shock capturing difficulty and convergence determination subject in high Mach number strong compressible flow calculation.

This invention discloses a high-frequency guided method for reconstructing flow field schlieren images of a lightweight scramjet engine, belonging to the technical field of image reconstruction. The method includes acquiring one-dimensional sparse pressure sequence data of the scramjet engine wall; generating an initial two-dimensional feature tensor by projecting the one-dimensional sparse pressure sequence data; inputting the initial two-dimensional feature tensor into a multi-scale feature encoder to obtain intermediate features and deep global semantic features at each scale; inputting the deep global semantic features into a spatial decoder for multi-stage cascaded decoding; and then completing cross-scale feature aggregation. After multi-stage cascaded decoding by the spatial decoder, a high-resolution scramjet engine flow field schlieren image is reconstructed. This invention eliminates redundant feature extraction branches and can directly and accurately map extremely sparse pressure sequence data into a high-resolution flow field schlieren image, achieving accurate capture and dynamic compensation of high-frequency detail features with low computational power consumption.

The application provides a real-time active oxygen supplementing regulation method and device for combustion of a scramjet engine, comprising the following steps: acquiring wall surface pressure data and flame self-luminous radiation data in a combustion chamber in real time; determining a current flame combustion state based on a flame combustion state discrimination criterion and the current acquired wall surface pressure data and flame self-luminous radiation data in the combustion chamber; and generating a corresponding control signal for the combustion chamber according to the current flame combustion state to realize active oxygen supplementing and combustion active regulation of the oxygen injection valve of the combustion chamber. The application can strengthen the combustion effect, improve the combustion efficiency, and improve the precision and stability of combustion control.

The present application relates to the field of hypersonic aerodynamics and intelligent control technology, especially to a bulge assisted oblique detonation wave initiation intelligent control method based on a feedforward neural network, which is used for the design and control of high-performance propulsion systems such as scramjet engines and oblique detonation wave thrusters. Wave structure data under different bulge geometric parameters and inflow conditions are obtained through numerical simulation technology. Image processing technology is used to extract feature parameters, label wave structure types, and construct sample data sets. A multi-layer feedforward neural network is used to train the samples, and the optimal network structure is selected through cross-validation to establish a nonlinear mapping model between the bulge geometric parameters and the wave structure categories. The neural network is used to quickly identify and predict the wave structure under unknown conditions, and the probability distribution is output to evaluate the confidence. According to the wave structure category and the control rule library, the height, length, position or inclination angle of the bulge is adjusted to realize real-time intelligent control of the oblique detonation wave system.