46 results about "Ablation" patented technology

The invention discloses a wall surface heat flow prediction method considering material ablation response under a semi-supervised learning framework, and relates to the field of aerospace, and the method comprises the following steps: constructing a data set; building a neural network model, and introducing the ablation physical information loss item into a loss function of the neural network model; pre-training the neural network model; predicting the label-free data through a pre-trained neural network model, and generating a pseudo label; mixing the labeled data and the unlabeled data, and performing semi-supervised self-training on the pre-trained neural network model to obtain a final neural network model; and wall surface heat flow prediction of material ablation response is carried out through the final neural network model. According to the method, the requirement of the model for a labeled data set is reduced, the model generalization ability under the small sample size training condition is enhanced, the physical interpretability of the network model is improved, and the wall surface heat flow prediction precision is improved.

The invention belongs to the technical field of high-temperature refractory composite materials, and discloses an ablation-resistant connecting piece and a preparation method thereof. The ablation-resistant connecting piece comprises the main body connecting part made of the high-temperature alloy material and the ablation-resistant fixing part, the internal main body connecting part can be effectively protected by the fixing part through ablation resistance and heat resistance, the main body connecting part is made of the high-temperature alloy, the high-temperature strength is high, and the service life of the main body connecting part is prolonged. The fixing part and the high-temperature-resistant binder are co-cured and formed on the main body connecting part, and the bonding strength of the fixing part and the main body connecting part is further improved through the embedding structure, so that the ablation-resistant connecting piece has the heat-proof performance of an ablation material and the high-temperature strength performance of high-temperature metal, is suitable for connection of short-time high-speed aircrafts, and is suitable for being used in a large-scale production line. The preparation method of the ablation-resistant connecting piece is relatively simple in required equipment, convenient to operate, stable in process, high in production efficiency and suitable for large-scale industrial production.

The application belongs to the field of flame detector structure design, and particularly relates to a variable-diameter probe structure of an aviation high-temperature-resistant ion flame detector. The structure comprises a cooling air inlet pipe (1) which comprises a large-diameter section (11), a small-diameter section (12) and a variable-diameter section (13) between the large-diameter section (11) and the small-diameter section (12), wherein the variable-diameter section (13) is provided with cooling air inlet holes (14); a probe head (2) is arranged at the end of the small-diameter section (12) of the cooling air inlet pipe (1), the probe head (2) is a hollow cone structure, the apex of the cone is provided with a spherical surface, and a plurality of small holes (21) are arranged on the probe head (2). The application improves the cooling effect of the cooling air flow on the probe of the ion flame detector, reduces the ablation influence of the hot air flow on the probe of the ion flame detector, and further improves the high-temperature resistance of the ion flame detector.

The invention discloses a dynamic measurement system for the ablation rate of a thermal protection material of an aircraft under high-temperature fuel gas. The dynamic measurement system comprises a sample placement unit, a fuel gas generation unit, laser displacement equipment and a data acquisition and processing system, the sample placement unit is used for placing a thermal protection material sample piece, the fuel gas generation unit comprises a fuel gas generator, and a flame nozzle of the fuel gas generator is used for being aligned with the surface of the thermal protection material sample piece; the laser displacement equipment dynamically captures displacement change signals of surface retreating of the thermal protection material sample piece under the action of the fuel gas generator; and the data acquisition and processing system calculates the dynamic ablation rate of the thermal protection material under the high-temperature fuel gas according to the displacement change signal. The laser displacement equipment is matched with a data acquisition system to finish dynamic acquisition of the retreat amount and calculation of the dynamic ablation rate, a dynamic measurement system is established, and the problems that the ablation rate of the thermal protection material cannot dynamically respond to changes and the accuracy of an ablation model prediction result is low when traditional static measurement is adopted at present are solved.

The invention relates to a thermal management-ablation resistance integrated ceramic matrix composite based on a shell silicide gradient structure and a preparation method thereof. According to the material, a carbon fiber preform is used as a reinforcing framework, and an interface layer is deposited on the surface; zrC-SiC ceramic is used as a matrix; zrSi2, MoSi2 and WSi2 high-melting-point silicides coated by the HfC shell layer are used as phase change heat absorption functional phases in gradient distribution, and when different silicides absorb instantaneous high heat flow step by step, the HfC shell layer can form an HfO2 high-melting-point anti-oxidation diffusion barrier after high-temperature oxidation to inhibit functional phase oxidation. According to the invention, a gradient structure of which the melting point of a functional phase is gradually increased along the thickness direction is realized through a gradient impregnation-curing-cracking combined reactive infiltration process, so that the problems of phase change medium volatilization, insufficient heat absorption capacity and interface mismatch failure of a traditional passive heat-proof structure at the temperature of 2000 DEG C or above are solved; and the service requirements of light weight, extreme thermal shock and long-time oxidation ablation resistance of high-speed aircrafts are met.

The invention relates to a method for measuring an ablated surface of a reentry capsule based on three-dimensional scanning, solves the problem of measuring the ablated surface of the reentry capsule, and belongs to the technical field of three-dimensional scanning. According to the method, a three-dimensional scanning measurement unit composed of a handheld laser tracker and a visual target is adopted, 360-degree surface ablation morphology data of the reentry capsule are spliced in a unified space reference based on a structured light scanning global positioning splicing technology, and a digital stereoscopic model of the reentry capsule is constructed; and extracting a theoretical three-dimensional model of the reentry capsule, positioning the scanning point cloud of the reentry capsule structure to the theoretical three-dimensional model of the reentry capsule structure through profile fitting, and realizing measurement of the ablation surface of the reentry capsule structure based on point cloud comparison.

The application discloses a ceramic-based composite structure integrating heat-proof and electromagnetic wave absorbing functions based on a cross-scale porous superstructure, which comprises, from bottom to top, a metal bottom plate, a micro-macro cross-scale porous superstructure electromagnetic wave absorbing layer, a micro-nano scale porous structure ceramic heat insulation layer and a high-temperature ceramic ablation-resistant layer. The application is innovative in that different scale porous structures are designed into a dense gradient multi-functional integrated superstructure according to electromagnetic wave absorbing and heat insulation requirements by a cross-scale artificial configuration design method, has the structural and functional integration characteristics, and is convenient for realizing ceramic 3D printing integrated forming, and can provide a new method for realizing electromagnetic protection / thermal protection integration of future high-speed aircraft components.

The present application relates to the field of electrophysiology, and particularly relates to an ablation state evaluation system. The system comprises an impedance acquisition device and a control unit. The impedance acquisition device is configured to acquire impedance information of tissue at an ablation site in real time before and after energy delivery at different frequencies. The control unit is configured to calculate a frequency response index based on the received impedance information, obtain an energy delivery difference index at a current position, and obtain an energy delivery difference index of an ablation path in combination with a spatial position of an energy delivery electrode. The impedance information of tissue at the ablation site before and after energy delivery is acquired in real time at different frequencies. The frequency response index of tissue at the ablation site before and after energy delivery is calculated based on the impedance information, an energy delivery difference index at a current position is obtained, and an energy delivery difference index of an ablation path is obtained in combination with a spatial position of an energy delivery electrode. The present application realizes rapid evaluation of the effectiveness of energy delivery on an ablation path during an ablation process.

The present application belongs to the technical field of high-temperature refractory composite material, and discloses a kind of ablation-resistant connecting piece and preparation method thereof.The ablation-resistant connecting piece includes a main connecting part of high-temperature alloy material and an ablation-resistant fixing part.The fixing part effectively protects the internal main connecting part by having ablation resistance and heat resistance.The main connecting part is processed by high-temperature alloy, which has high high-temperature strength.The fixing part is co-solidified with a high-temperature resistant adhesive on the main connecting part, and the bonding strength of the fixing part and the main connecting part is further improved by the embedded structure, so that the ablation-resistant connecting piece has both the heat resistance of ablation-resistant material and the high-temperature strength of high-temperature metal, and is suitable for connecting short-time high-speed aircraft.The preparation method of the ablation-resistant connecting piece requires relatively simple equipment, is convenient to operate, has stable process, high production efficiency, and is suitable for large-scale industrial production.

The invention belongs to the technical field of high-temperature fire-resistant composite materials, and discloses a screw for connection of an aerobic long-time hypersonic aircraft and a manufacturing method of the screw. The screw comprises a core body made of a high-temperature alloy material, a heat insulation layer reinforced by a phase-change material and an ablation-resistant shell reinforced by ablation-resistant powder, and the core body and the heat insulation layer are integrally wrapped by the ablation-resistant shell. The screw is ablated through the ablation-resistant shell to prevent heat, and the core body in the screw is protected through phase-change heat absorption and low-thermal-conductivity heat insulation of the heat insulation layer. The screw has long-time anti-ablation capacity and heat insulation capacity and also has the high-temperature strength performance of high-temperature alloy metal, meanwhile, the ablation-resistant shell of the screw integrally wraps the core body and the heat insulation layer, the risk of structural failure caused by high-temperature falling does not exist, the overall reliability of the screw is high, and the screw is suitable for connection of aerobic long-time hypersonic aircrafts. The screw manufacturing method is relatively simple in required equipment, convenient to operate, stable in process, high in production efficiency and suitable for large-scale industrial production.

The invention discloses a heat insulation-phase change composite thermal protection device which comprises a heat insulation unit and a heat absorption and removal unit. The heat insulation unit comprises an anti-ablation layer, a buffer heat insulation layer and a main body heat insulation layer; the heat absorbing and discharging unit comprises a high-temperature phase change layer, a medium-temperature phase change layer and a sensible heat recovery layer, and the high-temperature phase change layer and the medium-temperature phase change layer communicate with the sensible heat recovery layer through exhaust pipelines; the anti-ablation layer, the sensible heat recovery layer, the buffer heat insulation layer, the high-temperature phase change layer, the main body heat insulation layer and the medium-temperature phase change layer are sequentially arranged from outside to inside; the heat insulation performance of the materials selected by the buffer heat insulation layer is superior to that of the main heat insulation layer; and after being gasified, the phase-change materials of the high-temperature phase-change layer and the medium-temperature phase-change layer enter the sensible heat recovery layer through the exhaust pipeline to absorb and take away part of heat penetrating through the anti-ablation layer. According to the composite thermal protection device, external heat flow can be effectively blocked, internal heat can be timely dissipated, and therefore efficient heat insulation and temperature control of the nose cone part of the aircraft are achieved.

The application relates to the technical field of aircrafts, and discloses a calculation method, equipment and medium for the multi-physical effect of ablation particles, the method comprising the following steps: acquiring equivalent inertia items induced by respective multi-physical effects; calculating a multi-physical particle dynamics process according to a particle sliding resistance item and the equivalent inertia items induced by the respective multi-physical effects; performing dimensionless processing on a radiation heat transfer equation of spherical particles in the volume of an arbitrary sampling unit to obtain a dimensionless radiation heat transfer equation; and calculating a thermal radiation particle thermodynamic process by using the dimensionless radiation heat transfer equation. The method can reflect complex physical processes under extreme mechanical environments, ensure that numerical simulation can reproduce the original appearance of ablation physics, has engineering interaction of ablation problems, automatically performs dimensionless processing, avoids numerical rigidity caused by dimensional differences of different physical problems, and fills the gap that the prior art cannot calculate complex processes of multi-physical effects of ablation particles under extreme mechanical conditions.

In some examples, a medical system includes processing circuitry configured to determine the periodicity of the flow rate of fluid delivered by the medical system to a balloon for medical ablation. The balloon is configured to inflate to facilitate heat transfer between the vessel wall and the fluid flowing through the balloon. The medical system is configured to vary the flow rate of the fluid delivered to the balloon between an upper and lower limit flow rate within this period. When the medical system delivers fluid at the upper limit flow rate, the lower limit flow rate, and this period, the upper limit flow rate, the lower limit flow rate, and the period maintain tissue at a tissue depth within a narrow temperature range around a target temperature.

The invention discloses a digital intelligent design method for an aerodynamic thermal protection system of a wide-area aircraft, and belongs to the technical field of aerospace thermal protection. Firstly, a flight path model is established and an aerodynamic thermal environment is calculated; constructing a structured database containing the thermal physical attributes of the materials; sample data is generated through high-fidelity heat conduction-ablation coupling physical simulation, and a proxy model is trained to achieve rapid performance prediction; on the basis, a multi-objective optimization model with the weight and the cost as objectives and the back wall temperature, the ablation depth, the interface temperature and the total thickness as constraints is established, and a global optimization algorithm is adopted for scheme automatic optimization; and finally, through high-fidelity simulation verification, the design is ensured to meet the precision indexes that the highest surface temperature deviation is less than or equal to 100 DEG C, the back temperature deviation is less than or equal to 20 DEG C and the ablation prediction error is less than or equal to 2 times. The wide-range adaptive capacity that the coverage Mach number is 2-8 and the flight height is smaller than or equal to 60 km is achieved, the design efficiency and precision are remarkably improved, and the system weight and cost are reduced.

The utility model provides a hanging support and an ablation system. The hanging support comprises a supporting part, a telescopic hanging rod and a clamping piece. The supporting part can be fixed to corresponding medical equipment, and a first clamping groove is formed in the first side face of the supporting part; a rotating part rotatably matched with the supporting part is arranged at the end part of the telescopic hanging rod, a first end surface of the rotating part corresponds to the first side surface, and a second clamping groove is formed in the first end surface; the clamping piece is matched with the two clamping grooves and is constructed to be capable of switching the matching state with the two clamping grooves, and the matching state comprises that the clamping piece is located in the first clamping groove and the second clamping groove at the same time, and the clamping piece is integrally located in the first clamping groove or the second clamping groove. Based on the technical scheme of the utility model, the equipment is used as a bracket base, and a conventional base with large volume and heavy weight, a telescopic structure of the telescopic hanging rod and a rotary connecting structure between the telescopic hanging rod and the supporting part are omitted, so that the adjustment of the hanging height and the upright or laying down of the telescopic hanging rod are realized, and the requirements of use and storage are met at the same time.

The invention discloses a steam generation device based on LC induction heating, ablation equipment and a control method. The device comprises a power control circuit, an LC resonance induction circuit and an induction heating pipe. The output end of the power control circuit is electrically connected with the input end of the LC resonance induction circuit, and the output end of the LC resonance induction circuit is in coupling connection with the induction heating pipe in an electromagnetic induction mode. The problems that an existing ablation equipment heating technology is low in heating efficiency and low in heating speed are effectively solved, sufficient steam can be generated within hundreds of milliseconds based on LC induction heating, and compared with traditional resistance type heating, the device has the advantages of being high in response speed and energy density, achieving non-contact type heating, being easy to control accurately and the like; the performance requirement of a clinical practical application scene for ablation equipment is met, and powerful support is provided for further application of the steam ablation technology in the medical field.

The invention relates to the technical field of ablation protective sheaths, and discloses an ablation protective sheath and a control method thereof.The ablation protective sheath comprises a sheath body and a plugging device, the sheath body is provided with a main channel, a temperature measuring channel, a cooling liquid channel and a negative pressure channel, the temperature measuring channel, the cooling liquid channel and the negative pressure channel are arranged around the main channel, the plugging device is arranged in the main channel in a penetrating mode, and the end, away from the near end, of the plugging device forms an expansion part and is driven to expand so that a heat preservation area can be formed between the expansion part and the far end. The heat preservation area is formed by the plugging device and the sheath body, medical staff can measure the temperature of the heat preservation area in real time through the temperature measuring channel, and heat radiation damage is avoided; meanwhile, normal saline in the heat preservation area can be updated through the cooling liquid channel and the negative pressure channel, the pressure cannot be increased while the temperature of the heat preservation area is reduced, and safety is higher.

The application discloses a method for calculating aerodynamic heat and thermal response coupling considering ablation pyrolysis product injection, relates to the field of aerodynamic heat and thermal protection coupling numerical calculation, and establishes the following methods: a numerical calculation method for aerodynamic heat environment under the condition of mass injection; a numerical calculation method for thermal response considering surface ablation reaction, internal pyrolysis reaction, pyrolysis gas flow and solid heat conduction; and a method for coupling data transmission between the two methods. The method can simulate the generation, transportation and injection process of ablation pyrolysis thermal protection system surface ablation reaction products and internal pyrolysis reaction products of high-speed aircraft, and can calculate the coupling aerodynamic heat / thermal response result of the ablation pyrolysis thermal protection system of high-speed aircraft considering mass injection.

The application provides a simulation method based on material ablation effect under cooperation of atomic oxygen and ultraviolet, and relates to the technical field of spacecraft simulation calculation, and the simulation method comprises the following steps: obtaining atomic oxygen environment influence parameters, wherein the atomic oxygen environment influence parameters comprise spacecraft speed, incoming flow direction and atomic oxygen density; establishing a three-dimensional model of a spacecraft, inputting the atomic oxygen environment influence parameters into the three-dimensional model, and obtaining atomic oxygen flux on the surface of the spacecraft; determining a material ablation reaction coefficient on the surface of the spacecraft according to the cooperation of atomic oxygen and ultraviolet radiation on the surface of the spacecraft; and obtaining the ablation thickness of the material on the surface of the spacecraft through the atomic oxygen flux on the surface of the spacecraft and the material ablation reaction coefficient on the surface of the spacecraft. The simulation method is simple in steps, easy to operate, and accurate in simulation.

The invention discloses an aerodynamic heat and thermal response coupling calculation method considering ablation pyrolysis product injection, relates to the field of aerodynamic heat and thermal protection coupling numerical calculation, and establishes an aerodynamic heat environment numerical calculation method under a mass injection condition. A thermal response numerical calculation method considering a surface ablation reaction, an internal pyrolysis reaction, pyrolysis gas flow and solid heat conduction; the invention relates to an aerodynamic heat / thermal response coupling calculation method considering ablation pyrolysis product injection, and a coupling data transmission method strategy between the two. According to the method, the generation, transportation and injection processes of surface ablation reaction products and internal pyrolysis reaction products of the high-speed aircraft ablation pyrolysis thermal protection system can be simulated, and a coupled aerodynamic thermal / thermal response result of the high-speed aircraft ablation pyrolysis thermal protection system considering mass injection can be calculated.

The invention provides a step heating design method and device for simulating ballistic ablation response and a medium, and the method comprises the steps: obtaining a target response parameter in a flight ballistic environment, determining a heating time length, debugging a test state, and carrying out ablation thermal response trial calculation, and preparing an initial simulation ablation retreat amount and a simulation key position highest temperature for iterative calculation. Step heating state parameters of the ground simulation test are determined through iterative calculation on the basis of the target response parameters, and the first iteration step is used for determining the total enthalpy coefficient of the total enthalpy of the ground incoming flow with the matched target ablation retreat amount as the target; and the second iteration step takes the highest temperature of the matched target key position as a target and is used for determining a heat flux density coefficient of the ground heat flux density. According to the technical scheme, the total heating amount, the ablation amount and the highest temperature response of the key position in the step heating state can be completely matched with the response parameters in the flight trajectory state.

The invention discloses a method for designing a multilayer anti-ablation-radiation regulation and control integrated thermal protection structure in an interference area. The method comprises the following steps: determining flight parameters of a hypersonic aircraft under a thermal protection structure design scheme; setting parameter constraint conditions of the thermal protection structure; establishing an aerodynamic thermal numerical simulation calculation model for the local interference area; thermal environment variables of the interference area under different groups of flight parameters are calculated, so that an aerodynamic thermal environment database is constructed; determining interference area thermal environment parameters corresponding to the flight parameters; calculating the ablation retreating amount and the back surface temperature corresponding to different thermal protection materials; optimizing by using an optimization algorithm to obtain a thermal protection material with the minimum ablation amount and the lightest mass, taking the thermal protection material as an ablation layer, and determining the design thickness of the ablation layer; reserving an air layer at the bottom of the ablation layer; based on the parameter constraint conditions and a thermal protection material database, determining a thermal insulation layer and parameters thereof; and sequentially combining the ablation layer, the air layer and the thermal insulation layer from top to bottom to form a thermal protection structure, thereby completing the design scheme.

The invention discloses a method for acquiring ablation morphology of a C / C composite material based on a physical information neural network, and relates to the field of aerospace. Comprising the following steps: taking coordinate data, ablation time, a dimensionless parameter Sh and a dimensionless parameter A of a C / C composite material three-dimensional space model under a Cartesian coordinate system as inputs of a pre-trained multilayer perceptron network, and outputting an oxygen molar concentration and a solid volume fraction through the pre-trained multilayer perceptron network; and solving a gas phase diffusion equation and a solid phase evolution equation through the oxygen molar concentration and the solid volume fraction output by the multi-layer perceptron network to obtain the ablation morphology of the C / C composite material. According to the method, the gas-solid coupling equation can be efficiently and accurately solved, the limitation of a traditional numerical method is overcome, the ablation morphology of the C / C composite material under the gas-solid coupling effect is accurately obtained with low calculation cost, and powerful theoretical support and technical guarantee are provided for application of the C / C composite material in the fields of aerospace and the like.

The invention provides an aircraft aerodynamic force / thermal environment-material ablation response coupling calculation method, and relates to the technical field of aerospace, and the method comprises the steps: dispersing an aircraft reentry trajectory into a plurality of trajectory points, obtaining the trajectory parameter of each trajectory point, determining the temperature and wall surface temperature of an aircraft heat-proof material at the trajectory starting moment, and calculating the temperature of the aircraft heat-proof material at the trajectory starting moment; calculating an initial external flow field of each ballistic point according to the ballistic parameters of each ballistic point, forming initial thermal environment data along the ballistic, calculating the ablation response of the material, obtaining an initial material thermal response calculation result, and calculating the thermal response of the material through multiple coupling iterative calculations. Thermal environment data at each ballistic point moment and a thermal response calculation result at each ballistic point moment are obtained, and when the heat flow relative deviation of the current coupling iteration M and the coupling iteration M-1 does not exceed a threshold value, a convergence standard is reached, and a material ablation response coupling calculation result is obtained. The problem that the material ablation response of the trajectory of the aircraft is difficult to calculate accurately in real time is solved.

The invention discloses a nozzle ablation calculation method and device and a storage medium, and belongs to the technical field of aerospace. Comprising the following steps: 1, acquiring spectral data of main components of high-temperature fuel gas in a spray pipe, and processing the spectral data to obtain radiation data of the high-temperature fuel gas; 2, according to the high-temperature fuel gas radiation data obtained in the step 1, obtaining actual temperature offline intensity, broadening of Doppler and Lorentz effects and monochromatic absorption coefficients, and calculating a Planck average absorption coefficient; 3, inputting the Planck average coefficient into a two-way coupling model, wherein the two-way coupling model introduces a DO model and simulates a wall surface retreating process by adopting a dynamic grid technology; and iteratively updating the flow field, and outputting distribution data of the ablation rate and the ablation amount of the spray pipe. When numerical simulation is carried out on the nozzle ablation problem, a reasonable algorithm is introduced, a proper radiation heat exchange mathematical model is combined, radiation heat is considered in the calculation process, and the method has extremely high practical application value.

This invention discloses an integrated composite structure based on a porous superstructure for active cooling, heat protection, and electromagnetic wave absorption. From bottom to top, it comprises: a metal base plate, a ceramic-based high-temperature absorbing superstructure layer, a fiber-reinforced aerogel insulation layer, and an ultra-high temperature ablation-resistant coating. This integrated composite structure utilizes a three-dimensional superstructure, formed by a periodic array of porous cells using a lightweight ceramic-based absorbing agent, as its core component. It comprehensively leverages the synergistic effects of multiple mechanisms: the active thermal protection through phase change and sweating of the porous superstructure, the efficient heat insulation of the aerogel nanoporous network structure, the high-temperature resistance and ablation resistance of the ultra-high temperature ceramic coating, and the high-loss absorption of electromagnetic waves in the wide frequency and angular domains of the three-dimensional superstructure. This results in an integrated structural and functional structure that easily integrates thermal and electromagnetic protection, providing a new approach for achieving integrated high-temperature resistance, ablation resistance, heat insulation, and electromagnetic wave absorption protection for future high-speed aircraft shell components.

The invention relates to a nozzle throat liner ablation prediction system based on data mining, which belongs to the technical field of aerospace and comprises a data import and loading module, a full model training module, a selective verification module, a prediction module and a result export module. The data importing and loading module is used for importing historical data and inputting the historical data into the model training module; the full model training module processes the data through a plurality of data mining algorithms to generate a prediction model; the selective verification module is used for selecting different types of models for training and verification according to requirements; the prediction model performs ablation prediction through real-time data input; the result export module stores the prediction result; the method has the beneficial effects that the ablation trend of the nozzle throat liner can be predicted in advance, replacement is carried out in time, nozzle faults caused by serious ablation are avoided, assistance is provided for designers to accurately estimate an inner trajectory and provide accurate ablation data, and the safety and reliability of a system are improved.

This invention provides a motor-unlocked launch device ablation test fixture, comprising: a base plate (1), a front main arm (2), a worktable (3), an electric push rod (4), a mechanical safety device (5), a guide groove (6), a mounting base (7), an auxiliary arm (8), and a rear main arm (9). The launch device is mounted on the worktable, and a movable push arm is installed on the lower part of the worktable, allowing the launch device to move with the worktable during the test, moving away from the engine exhaust flame, simulating the relative positional relationship between the launch device and the engine during the launch beam separation process. By controlling the locking and movement of the push arm through the electric push rod and the mechanical safety device, the ablation time of the launch device can be accurately controlled according to the command, simulating the ablation conditions during the actual engine launch process.

The application provides a kind of aircraft aerodynamic force / thermal environment-material ablation response coupling calculation method, it is related to aerospace technical field, the method is to reentry trajectory of aircraft is dispersed into multiple trajectory points, obtains the trajectory parameter of each trajectory point, simultaneously determines the temperature and wall surface temperature of heat protection material of aircraft at trajectory start time, according to the trajectory parameter of each trajectory point, calculates the initial external flow field of each trajectory point, forms initial thermal environment data along trajectory, calculates the ablation response of material, obtains initial material thermal response calculation result, through multiple coupling iteration calculation, obtains the thermal environment data of each trajectory point time and the thermal response calculation result of each trajectory point time, when the heat flow relative deviation of current coupling iteration M and coupling iteration M-1 is not more than threshold value, reach convergence standard, obtain material ablation response coupling calculation result.The application solves the problem that the material ablation response of aircraft trajectory is difficult to calculate accurately in real time.

The invention provides an X-ray-based in-situ nondestructive testing system and method for an ablation process of a heat-proof material, and the system is characterized in that a micro-focus X-ray source and a digital flat panel detector are externally arranged at the two sides of a test cabin of an arc heating test system; the incident window and the emergent window which are made of the low-density alloy form an X-ray penetration path, so that in-situ, real-time and nondestructive testing of the internal structure of the test model in a high-temperature airflow ablation state is realized. The external design effectively isolates the influence of severe environments such as high temperature, vibration and the like on the precise imaging equipment, and ensures the equipment safety and the imaging quality. According to the invention, the dynamic initiation and evolution process of internal cracks, debonding, layer uncovering and other defects of the heat-proof material in the ablation process can be observed and analyzed in real time, key technical support is provided for quantitative research on the damage mechanism of the material under a force and heat load, and the evaluation capability of a ground aerodynamic heat test is remarkably improved.