66 results about "Ablation" patented technology

The invention discloses an aircraft end ablation simulation algorithm structure and method based on a meshless method. The structure comprises a pneumatic calculation module, an ablation calculation module and an SPH boundary movement / structure heat transfer module. The method comprises the following steps: dividing a flight trajectory into a plurality of discrete flight state points according to the flight trajectory of reentry flight of the aircraft; ballistic data, a geometric shape and an atmosphere model of the aircraft end are obtained; according to the incoming flow parameter and the geometric shape, calling a pneumatic calculation module to calculate a wall surface pneumatic flow field parameter of the end; calculating an effective heat flow and a linear ablation rate; calling a meshless SPH structure heat transfer / boundary movement module program; according to the method, characterization through particles replaces characterization through grids, so that characterization of the end appearance after ablation does not need grid movement, the length-width ratio of the grids is changed to cause singularity of the grids, calculation of a temperature field in the ablation process cannot be continuously carried out, and calculation can be stably carried out in the large-deformation ablation simulation process.

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 relates to the technical field of thermal protection of hypersonic aircrafts, in particular to a multi-scale prediction method for an ablation behavior of a thermal protection structure based on data driving, which comprises the following steps: performing molecular dynamics simulation on a gas-solid interface of an ablation type thermal protection material to obtain simulation calculation activation energy; obtaining the experimental activation energy of the material; performing Bayesian optimization on the simulation calculation activation energy and the experiment measured activation energy to obtain optimized activation energy; performing macroscopic wall surface ablation simulation by taking the optimized activation energy as a boundary condition to obtain an ablation parameter database recording ablation wall surface temperatures, ablation rates and ablation depths corresponding to different flow rates, altitudes and time; training a radial basis function neural network based on the ablation parameter database, and establishing an ablation prediction model; obtaining a prediction result; the calculation efficiency and accuracy of the thermal protection structure ablation behavior multi-scale prediction method can be improved.

The invention relates to a multi-scale prediction method for the oxidative ablation behavior of a carbon / carbon porous heat-proof material, and belongs to the technical field of aerospace science, in particular to a multi-scale prediction model for the oxidative ablation behavior of the carbon / carbon porous heat-proof material. A chemical reaction standard enthalpy value and reaction activation energy are analyzed based on a gas-solid interface micro-model of a carbon / carbon porous heat-proof material, and then pore scale fluid mechanics calculation is performed in combination with a porous medium model; explaining an evolution mechanism between gas-solid interfaces by using a micro-scale and pore-scale simulation method, and predicting an ablation process surface evolution process; according to the invention, fluid flow, thermal reaction and heterogeneous chemical reaction in the porous medium in the ablation process can be coupled and simulated at the same time; according to the method, the action mechanism of the heterogeneous chemical reaction at the gas-solid interface in the oxidation ablation process of the carbon / carbon porous heat-proof material can be disclosed, and the distance change after ablation and the ablation surface roughness in the ablation process can be predicted more accurately.

The invention relates to the technical field of medical instruments, and provides a steam gun and a steam ablation system. The steam gun comprises a steam gun body and a steam needle assembly, the fixed end of the steam needle assembly is connected to the steam gun body, and the steam needle assembly can move relative to the steam gun body so that the steam needle assembly can stretch out or retract from the steam gun body; the steam needle assembly comprises an outer sheath tube and a steam needle, a plurality of output holes are formed in the free end of the steam needle, the outer sheath tube is arranged on the outer side of the steam needle in a sleeving mode, and the outer sheath tube can move relative to the steam needle so that the number of the output holes covered by the outer sheath tube can be adjusted. According to the steam gun, the problem that in the prior art, the ablation range of a steam needle cannot be flexibly adjusted can be solved, and the effect that the steam needle can conduct adaptive ablation on hyperplasia tissues of different depths and different sizes is achieved.

The present application provides a method and system for estimating the on-target laser power density required to burn through protective materials. The method of the present application includes: constructing a tightly coupled model between fluid, structural heat transfer, and laser through several sets of protective material ablation simulation tests; using the tightly coupled model to analyze the factors affecting the laser ablation characteristics of the protective material to obtain the main influencing factors; conducting simulation tests on the main influencing factors under different working conditions to obtain the laser ablation characteristics of the protective material under different main influencing factors, and then establishing a laser ablation database for protective materials; obtaining the ablation rate required to burn through the protective material, and obtaining the on-target laser power density required to burn through the protective material through the protective material laser ablation database and the required ablation rate. The present application can quickly predict the on-target laser power density required to burn through aircraft protective materials under flight conditions, thereby providing technical support for future aircraft thermal protection design.

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 present invention proposes a device for thermal environment assessment and testing of aircraft components and a method for using the device, which relates to the field of aircraft material testing technology. The device includes aircraft components, crucibles, refractory bricks, a test bench, an ablation system, a temperature measurement system, and a temperature control system. The device can provide a thermal environment higher than 1750 degrees to simulate the aerodynamic heating environment of aircraft components in service. The method for using the device includes the following steps: S1, setting up test equipment, performing heating environment pretreatment, and presetting the temperature control system; S2, maintaining the temperature control system configuration, and performing service thermal environment assessment tests; S3, measuring and recording the surface temperature distribution of aircraft components. By setting two working environments, including a heating environment and a service thermal environment, segmented heating is performed to the required test thermal environment, and the platinum-rhodium thermocouple is replaced with a smaller crystal sensor, thereby achieving synchronous online monitoring of high-temperature heating of aircraft components, non-destructive flow field, and in-situ measurement.

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 present invention proposes an intelligent planning method for emergency tasks of space debris cleaning. The planning platform applied by the method includes a contact debris cleaning tool task planning module based on reinforcement learning and a non-contact laser ablation deorbiting task planning module based on reinforcement learning, which are used to clean space debris in the scene. The contact debris cleaning tool task planning module based on reinforcement learning includes several contact capture and deorbiting removal aircraft and launch vehicles. The combined cleaning scheme of contact capture and deorbiting removal and non-contact laser ablation-driven deorbiting is adopted, which has the processing ability of "multiple to one", and can complete the cleaning task at a lower cost.

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 an efficient simulation method for continuous laser ablation of a CFRP laminated plate, and relates to the technical field of laser weapon damage high-speed aircrafts.The method comprises the steps that characteristic parameters of the CFRP laminated plate are designed based on dimensional analysis, and a three-dimensional geometric model considering material anisotropy is established; the method comprises the following steps: representing thermal-mechanical characteristics of a CFRP laminated plate in a high-temperature environment based on multi-scale thermal-mechanical analysis, and constructing a neural network model; defining a material constitutive relation related to temperature, and setting a displacement boundary and a multi-physics coupling loading condition by applying an ALE adaptive grid technology; establishing a multi-mechanism coupled laser ablation analysis model, and accurately representing an energy transfer process; high-precision tracking and evolution simulation of an ablation boundary are realized by adopting a method of combining a dynamic life-death unit and a self-adaptive grid reconstruction technology; and obtaining damage evaluation data through a restart analysis technology. The method provides important technical support for damage efficiency evaluation of laser weapons for intercepting high-speed aircrafts.

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 present invention provides a simulation method for the ablation effect of spacecraft surface materials in the autonomous acquisition mode, which relates to the technical field of spacecraft simulation calculation. The simulation method realizes the simulation calculation of the ablation effect of materials under the synergistic action of atomic oxygen and ultraviolet by means of autonomously selecting and acquiring the parameters affecting the ablation effect. The parameters affecting the ablation effect include polygon mesh parameters, atomic oxygen parameters, ultraviolet parameters, ablated material parameters and protective layer parameters. Based on the synergistic action of atomic oxygen and ultraviolet, the present invention can realize the acquisition mode of autonomously and freely selecting the parameters affecting the ablation effect, and realize the simulation experiment of the ablation effect of materials under different doses of atomic oxygen and ultraviolet irradiation, which can not only accurately simulate the actual situation, but also accelerate and accurately evaluate the ablation effect of materials under the synergistic action of atomic oxygen and ultraviolet.

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 discloses an engineering calculation method for predicting wall surface heat flow for different injection working media, and relates to the field of aerospace. According to the method, the wall surface heat flow under different ablation gas injection working media is obtained by calculating the mass injection coefficient, and the ratio of the average molecular weight (# imgabs0 #) of incoming flow gas to the average molecular weight (# imgabs1 #) of injection gas and the ratio of the unit mass enthalpy value of the incoming flow gas to the unit mass enthalpy value of the injection gas are introduced on the basis of an existing mass injection coefficient formula. Two conditions of # imgabs2 and # imgabs3 are considered at the same time, the wall surface heat flow can be quickly and accurately predicted under the action of different gas injection working media, and the precision of the method is remarkably higher than the calculation precision of an existing mass injection coefficient formula.

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 relates to the technical field of aircrafts, and discloses an ablative particle wall turbulent boundary layer frictional resistance and heat flow calculation method, equipment and medium, the method comprising the following steps: acquiring particle two-phase wall turbulent direct numerical simulation data; calculating each sub-effect corresponding to frictional resistance decomposition and each sub-effect corresponding to heat flow decomposition respectively according to the particle two-phase wall turbulent direct numerical simulation data; performing MPI reduction operation summation on each sub-effect corresponding to the frictional resistance decomposition to obtain the frictional resistance of the ablative particle wall turbulent boundary layer; and performing MPI reduction operation summation on each sub-effect corresponding to the heat flow decomposition to obtain the heat flow of the ablative particle wall turbulent boundary layer. The above method has strong compatibility and can be coupled with any compressible particle two-phase flow calculation mode; the method is efficient in parallel and can quickly process particle two-phase wall turbulent direct numerical simulation data to obtain the frictional resistance and heat flow of the ablative particle two-phase flow boundary layer, thereby providing a technical means for diagnosing and optimizing the design of high-speed ablative aircrafts.

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 invention discloses a thermal protection structure ablation path compensation synthesis imaging method, which belongs to the technical field of aerospace structure health monitoring, and comprises the following steps: acquiring a reference signal and a monitoring signal of each channel to obtain an ablation damage scattering signal, and calculating an ablation damage factor to obtain a path compensation result; different types and combinations of ablation damage factors are selected to carry out path compensation on the ablation damage positioning features, and all the ablation damage positioning features after path compensation are obtained; ablative damage positioning features in the monitoring area are extracted with the channel as the unit for synthetic imaging, and the position, where ablative damage occurs, of the thermal protection structure is positioned. The implementation process is simple, accurate and efficient; the problems that an existing thermal protection structure is high in temperature, high in complexity and low in reliability are solved, and the stability and reliability of thermal protection structure ablation damage monitoring under high-temperature and high-speed airflow are effectively improved.