2041 results about "Thermal protection" patented technology

The disclosed technology is directed to a computing device for detecting and preventing melting of a component of the computing device. In some examples, the computing device includes a cable that connects a power supply unit and an add-on card, and a thermal protection controller. Based on a sensor signal from a temperature sensor of the cable, the thermal protection controller determines that a temperature associated with the cable exceeds a threshold temperature. Responsive to determining that the temperature associated with the cable exceeds the threshold temperature, the thermal protection controller causes the power supply unit to cease supplying power to the add-on card by transmitting an overtemperature signal through the cable.

The invention relates to the technical field of communication and network transmission, in particular to a terminal data acquisition and transmission method for a long-distance heat supply pipeline. The method comprises the following steps that long-distance heat supply pipeline laying area data are obtained, pipe network laying form analysis is carried out, and single-pipe branched heat supply pipe network data are obtained; constructing a pipeline laying model based on the data of the single-pipe branched heat supply network; performing intelligent heat energy flow simulation based on the pipeline laying model, and calculating heat loss in the heat energy flow simulation process to obtain heat loss data; performing thermal protection thermal insulation material integration on the pipeline laying model based on the heat loss data to obtain composite thermal insulation material data; and identifying a straight pipe section of the pipeline laying model, and calculating the minimum thick wall of the straight pipe section to obtain minimum thick wall data. The heat efficiency, the stability and the intelligent management capability of the pipeline system are improved based on the communication and network transmission technology.

The invention provides a thermal protection coating as well as a preparation method and application thereof. The thermal protection coating comprises a NiCoCrAlY first bonding layer, a NiCoCrAlY second bonding layer, a CeYSZ first ceramic layer and a GdYbYSZ second ceramic layer which are sequentially arranged on the surface of a base material, and the second NiCoCrAlY bonding layer is made of CeO2 (CeO2) and Al2O3 (Al2O3). According to the invention, a thermal protection coating structure with double bonding layers and double ceramic layers is constructed on the surface of the base material, and CeO2 and Al2O3 nano-particle reinforced phases are introduced, so that the use temperature of the thermal protection coating is increased, and the thermal strain tolerance, the crack resistance and the oxidation resistance are greatly improved.

The invention discloses a micro-channel device for phase change sweating and an application method, and belongs to the technical field of thermal protection of aircrafts. The structure sequentially comprises a solid inner wall, a transportation phase change layer and a sweating outer wall surface from inside to outside, and is mainly composed of a bottom plate, a porous rib wall, a solid rib wall, a solid outer wall, an inlet flow channel, a main flow channel, a branch flow channel and the like. The bottom plate is a solid inner wall and is used for supporting and conveying; the flow channel can be divided into an inlet flow channel, a main flow channel and a branch flow channel, and a cooling medium sequentially passes through the inlet flow channel, the main flow channel and the branch flow channel; a cooling medium is injected into an external flow field through the porous rib walls which are periodically arranged; the solid rib walls isolate the adjacent main runners, so that a cooling medium flows out through the porous rib walls; the outer surfaces of the porous rib walls of the solid outer wall are flush to form a sweating outer wall surface; the device has the advantages of being integrated in structure, small in thickness, uniform in sweating and efficient, and a new way can be provided for long-time cross-domain sailing thermal protection of high-speed aircrafts.

The corrugated plate type integrated thermal protection system based on the porous dot matrix composite phase change material comprises an outer plate and an inner plate, the outer plate and the inner plate are arranged in a spaced mode, and a web, a thermal insulation material and the composite phase change material are arranged in a spaced cavity formed by the outer plate and the inner plate; the multiple web plates are arranged in the interval cavities in a penetrating mode and used for forming a supporting structure for connecting the outer plate and the inner plate, and the multiple web plates are inclined to one another; the spacing cavity comprises a first area and a second area, the first area is close to the outer plate and used for containing a heat insulation material, and the second area is close to the inner plate and used for containing a composite phase change material; wherein the composite phase change material is arranged in a porous dot matrix mode, and the porous dot matrix is adopted as a heat conduction reinforcing body. The high heat storage efficiency of the porous dot matrix composite phase change material is utilized to quickly absorb and store heat of the web, so that the heat transferred into the aircraft is reduced, and the system has higher heat insulation performance.

A direct current fast charging (DCFC) control system for a range extended electrified vehicle (REEV) includes a set of DCFC connectors electrically connecting a high voltage battery system of an electrified powertrain of the REEV to a charging port of the REEV and a control system configured to determine an ambient temperature external to the REEV, access a temperature model configured to model a temperature associated with the set of DCFC connectors, wherein the temperature model accounts for heat energy generated by an engine of the electrified powertrain, determine, using the temperature model, a modeled temperature associated with the set of DCFC connectors, and control, based on the ambient temperature and the modeled temperature associated with the set of DCFC connectors, a DCFC current provided by an external DCFC station to a high voltage battery system via a charging port and the set of DCFC connectors.

The invention discloses a supersonic aircraft head thermal protection structure, and belongs to the technical field of hot end component cooling, the supersonic aircraft head thermal protection structure comprises a cooling system, a preheating system, a CBC power generation system and a secondary cooler, the cooling system comprises a coolant liquid storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall surface of the aircraft head, a cooling medium flowing in a direction opposite to that of external high-temperature gas is arranged in the cooling channel; the preheating system comprises a fuel tank, a fuel adjustable pump and a fuel channel; the secondary cooler is used for transferring heat in the cooling channel into the fuel channel; the CBC power generation system comprises a heater, a primary cooler and a circulating working medium, the heater is used for transferring heat in the cooling channel to the circulating working medium, and the primary cooler is used for transferring heat in the circulating working medium into the fuel channel. Efficient thermal protection of the head of the supersonic aircraft can be achieved, the risk of structural failure caused by high temperature is reduced, and the safety of the aircraft is improved.

The invention provides a spacecraft quartz powder heat insulation tile which is prepared by adopting high-purity quartz powder (SiO2) as a main raw material and combining a high-temperature-resistant ceramic fiber reinforcement and a nano aerogel heat insulation layer through a high-temperature sintering process. The heat insulation tile has the following advantages: 1, the heat insulation tile has excellent high temperature resistance, and can tolerate 1600 DEG C high temperature for a long time and tolerate 1800 DEG C instantaneous thermal shock for a short time; 2, repeated use: the structure is stable, and high strength is still kept after multiple thermal cycles; light weight: the density is low (less than or equal to 0.5 g / cm < 3 >), and the load of a spacecraft is reduced; and 4, the thermal conductivity is low (less than or equal to 0.05 W / m.K), and heat transfer is effectively isolated. The method is suitable for thermal protection systems (TPS) of spacecrafts, such as thermal insulation of reentry capsules and rocket engine nozzles.

The invention belongs to the field of novel composite materials, and particularly relates to a strong-ablation-erosion-resistant ceramizable flexible thermal protection material and a preparation method thereof.According to the material, ethylene propylene diene monomer serves as a matrix, polysilazane resin of a special structure serves as a ceramization modifier, and aerogel powder serves as heat insulation reinforcing filler. The thermal protection material provided by the invention can form a compact high-strength ceramic protection layer on the surface in a high-temperature environment, improves the structural strength of an ablation carbonization layer, effectively reduces the ablation degree of the material, has the characteristics of light weight, ablation resistance, scouring resistance, efficient heat insulation and the like, and can meet the protection requirements of high-performance engines of novel weaponry in the future.

The invention discloses a high-entropy alloy silicon carbide nanofiber composite material and a preparation method and application thereof, a high-entropy alloy is loaded on silicon carbide nanofibers, compared with a carbon fiber material carrier, SiC fibers can form a compact SiO2 protective layer in a high-temperature aerobic environment, further oxidation failure of the internal structure can be prevented, and the high-entropy alloy silicon carbide nanofiber composite material is prepared. And meanwhile, SiC also has relatively high high-temperature stability and weather resistance, so that the high-entropy alloy silicon carbide nanofiber composite material can stably work in an extreme high-temperature oxidation environment (such as a spacecraft thermal protection system and an engine tail gas pipeline) of over 1000 DEG C, and perfectly adapts to a high-temperature wave-absorbing scene.

The invention relates to a double-rare-earth crystal inhibition reinforced modified ultra-high-temperature protective coating as well as a preparation method and application thereof. The preparation method comprises the following steps: preparing a CVD-SiC inner coating on the surface of a matrix material by adopting a chemical vapor deposition method; a slurry sintering method is adopted, and a double rare earth (DRE) doped silicon-based structure anti-oxidation middle layer is prepared on the surface of the CVD-SiC inner coating; and preparing the PVD-ZrC thermal ablation prevention outer coating on the rare earth doped silicon-based anti-oxidation layer by adopting a physical vapor deposition method. According to the SiC / dRe coated Si-ZrSi2-ZrB2 / ZrC coating provided by the invention, the crystallization growth of SiO2-TGO in a silicon-based component structure is inhibited by doping rare earth oxide (REO) and / or rare earth carbide (REC), so that the stress accumulation and thermal expansion mismatch effect of the coating in a thermal service process is greatly relieved. Meanwhile, the complementary advantages of silicon-based ceramic and ultra-high-temperature ceramic in the aspects of oxidation resistance and ablation resistance are fully combined, the thermal protection capacity of the coating under the ultra-high-temperature condition is improved, the service life of the coating is remarkably prolonged, and the service stability of the coating is remarkably improved.

The invention discloses a ceramic-metal composite structure with a low thermal expansion coefficient and a middle layer for assisting two-step temperature control brazing and a preparation method of the ceramic-metal composite structure. The composite structure is formed by metallurgical bonding of a ceramic matrix, a titanium-containing active brazing filler metal layer, a low-thermal-expansion-coefficient (2 * 10 <-6 > / K-9 * 10 <-6 > / K) middle layer, a low-temperature brazing filler metal layer and metal base metal in sequence. During preparation, a step-by-step temperature control brazing process is adopted, wherein in the first step, ceramic / intermediate layer connection is achieved through titaniferous active brazing filler metal with the solidus larger than or equal to 780 DEG C under the vacuum condition smaller than or equal to 5 * 10 <-3 > Pa; and secondly, under the vacuum or inert atmosphere, brazing filler metal with the solidus being smaller than or equal to 650 DEG C is adopted for completing connection of the middle layer and the metal base metal. The low-stress and high-strength ceramic-metal composite structure is successfully constructed on the basis of the thermal matching design of the low-thermal-expansion-coefficient middle layer and the ceramic and the stress gradient release mechanism of the two-step temperature control brazing process in a cooperative coupling mode, and the low-stress and high-strength ceramic-metal composite structure is particularly suitable for the high-temperature and high-reliability fields of aerospace thermal protection systems, high-power electronic device packaging and the like.

The invention discloses a high-temperature-resistant optical endoscope integrating thermal protection and pollutant purging functions, and relates to the technical field of optical endoscopes, an endoscope main body comprises an endoscope front end part, an endoscope middle part and an endoscope rear end part from front to back, and the endoscope main body comprises an inner layer sleeve and an outer layer sleeve; an optical imaging module and an optical machine assembly module are mounted on the inner side of the inner-layer sleeve; a thermal protection and pollutant purging module is arranged between the inner-layer sleeve and the outer-layer sleeve and comprises a cooling gas channel, a cooling liquid inflow channel and a cooling liquid backflow channel, a cooling gas inlet channel and a cooling gas outlet channel are formed in the two ends of the cooling gas channel, a cooling liquid inlet channel is formed in one end of the cooling liquid inflow channel, and a cooling liquid outlet channel is formed in the other end of the cooling liquid backflow channel; a cooling liquid outlet channel is arranged at one end of the cooling liquid backflow channel, and the cooling liquid inflow channel is communicated with the end, located at the front end of the endoscope, of the cooling liquid backflow channel. According to the endoscope cooling device, efficient cooling and pollutant purging of the endoscope can be achieved.

The invention discloses a heat flow real-time inversion method and device based on an annealing gradient algorithm, and relates to the technical field of numerical calculation, and the method comprises the steps: carrying out the numerical calculation based on an external surface heat flow initial value, obtaining a temperature field, selecting a temperature value of a measurement point position, constructing an objective function, randomly generating a new solution, comparing the objective function constructed by the new solution and an old solution, and obtaining a heat flow real-time inversion result; if yes, the new solution serves as a candidate solution, if not, the new solution serves as the candidate solution according to the Monte Carlo judgment criterion by means of probability decision, and local iterative optimization is carried out by means of the gradient information optimization search process till the target function value reaches a predefined threshold value or the temperature parameter is reduced to the lower temperature limit; and ending under the condition that an end condition is met to obtain a final thermal flow vector. Through the randomness of the annealing algorithm and the Monte Carlo judgment criterion, local optimum is avoided, the multi-peak problem of nonlinear heat flow boundary condition identification can be solved, and an accurate identification result of the heat flow on the surface of the thermal protection structure is obtained.

The invention discloses a nanostructure thermal protection coating resistant to long-time high-temperature oxidation and a preparation method thereof.The nanostructure thermal protection coating takes a ceramic matrix composite material as a matrix, and a bonding layer and a composite component coating are sequentially sprayed to the surface of the matrix. Preparing a nano-structure self-healing additive / Re2Si2O7 spherical powder body; preparing a bonding layer on the surface of the ceramic-based composite material serving as a matrix; and 3, depositing the nano-structure self-healing additive / Re2Si2O7 spherical powder on the surface of the bonding layer to form the composite component coating. According to the preparation method, the self-healing additive / Re2Si2O7 is prepared into the nano-structure spherical powder suitable for thermal spraying for the first time, and a method is provided for implementation of the high-performance material on a thermal spraying nano-structure coating. The prepared coating is stable in phase structure at high temperature, phase change of powder is avoided in the spraying process, and the coating has good thermal expansion coefficient matching performance with a base body.

According to the transition prediction method suitable for the wide Mach domain of the high-speed aircraft provided by the invention, temperature and Mach number coupling correction is performed on key parameters in time scale modeling of a second mode of a transition mode, and simplified processing of a flow compression effect and a heat transfer process by traditional single parameter association is broken through; the influence of an incoming flow disturbance effect is embedded in the mode, and dynamic association between test environment parameters and transition triggering criteria is achieved; according to the improved mode, accurate prediction results are obtained on a plurality of examples with the Mach number range of 4.7-14, the research results provide a more reliable theoretical tool for aerodynamic configuration optimization of a high-speed aircraft with a wider speed range, and the method has important engineering guiding significance for thermal protection system design of a new-generation reusable aircraft.

The invention relates to a heat protection structure of a dredging coupling semi-open type dissipation structure, which belongs to the technical field of heat protection of high-speed aircrafts and comprises a dredging structure shell, a dredging structure capillary core body, a dredging structure working medium, a semi-open type dissipation structure, a semi-open type dissipation working medium, a semi-open type dissipation structure exhaust pipe and a semi-open type dissipation structure control switch. Aiming at a local high-heating part of a hypersonic aircraft, on the basis of semi-active dredging type uniform-temperature heat protection, phase change heat absorption and discharge heat dissipation of an active dissipation working medium are combined, and heat absorption and dissipation ways are increased, so that the total heat amount is effectively reduced while the internal heat distribution of the structure is optimized, and the heat dissipation efficiency is improved. Internal heat control and management of the structure are realized through coupling of dredging semi-active heat protection and active dissipation, the problem of non-ablative heat protection in an extreme environment is solved, and a feasible heat protection technical approach is provided for performance improvement of high-speed aircrafts in the future.

A heat protective sleeve includes a body being cylindrical extending along a longitudinal direction and comprising a triply periodic minimal surfaces (TPMS) structure, an open cell foam structure, or a lattice structure. A strip extends along the longitudinal direction of the body. The strip is a same material as the TPMS structure, the open cell foam structure, or the lattice structure. The material of the strip is solid and without holes through it. An endcap removably couples to the body. In some examples, the body has fastening threads and the endcap has mating threads. The mating threads of the endcap are configured to engage with the fastening threads of the body to releasably couple the endcap to the body.

The invention discloses a gas-phase sweating cooling thermal protection system with self-adaptive flow control. The gas-phase sweating cooling thermal protection system comprises a porous structure, a compact material structure, a porous structure inner cavity, a sweating agent supply pipeline, a pressure regulating device, a sweating agent gas source and an electronic control system. The invention further discloses a flow self-adaptive control gas phase sweating cooling thermal protection method, and when the thermal environment state is stabilized, the sweating agent supply flow q of the sweating agent gas source is determined according to the pneumatic heat flow Q on the outer side of the porous structure, the pneumatic pressure P1 on the outer side of the porous structure and the pressure P2 of an inner cavity of the porous structure; in an unstable thermal environment state, pre-judging changes of Q and P1 at the next moment, and adjusting q in advance; meanwhile, the temperature T of the porous structure is monitored, and when T rises to be higher than the upper limit of the optimal temperature interval, the sweating agent flow increment is started till T falls to the lower limit of the optimal temperature interval. The pneumatic thermal protection requirement can be efficiently met with the lower sweating agent flow requirement, and the pneumatic thermal protection device has wide application prospects in the technical field of pneumatic thermal protection.

The invention discloses a preparation method of an interface self-reaction strong chemical bonding radiant heat protection multilayer gradient coating, and belongs to the technical field of coating surface modification. The invention aims to solve the problems that the existing coating is high in preparation temperature, long in period, high in cost and not suitable for industrial batch production, the interface bonding between the coating and a nickel-based high-temperature alloy matrix is poor, the coating is easy to crack and even lose efficacy due to mismatching of thermal expansion under high-temperature service, and the thermal protection performance of the traditional coating is insufficient. The interface self-reaction strong-chemical-bonding radiant heat protection multilayer gradient coating and a nickel-based high-temperature substrate are subjected to self-reaction at the interface to form strong chemical bonding, the bonding strength of the coating reaches 15 MPa or above, the emissivity value of the coating reaches up to 0.9, the structure of the coating is kept complete and does not fall off after the coating is oxidized in 1250 DEG C high-temperature static air for 30 h, and the coating has the advantages of being simple in preparation process, low in cost and the like. The method has application potential in a nickel-based high-temperature alloy anti-oxidation radiation thermal protection multifunctional integrated metal thermal protection system for a high-speed aircraft.

The invention discloses a stable coolant supply structure for a sweating cooling system of a hot end component of an aircraft, the hot end component is arranged to be of a porous structure, that is, directional pore channels are densely distributed in the hot end component of the aircraft in the advancing direction; meanwhile, the inner space of the aircraft hot end component is divided into one or more coolant cavities, the front end of each coolant cavity corresponds to a directional pore channel of the corresponding cooling area, and the rear end of each coolant cavity is communicated with a branch pipeline with a control valve. The branch pipelines are gathered and then communicated with an output port of the coolant driving system, and all the control valves are connected with the regulator. The invention further discloses a stable coolant supply method of the sweating cooling system for the hot end part of the aircraft. The invention belongs to the technical field of thermal protection of aerospace aircrafts, and solves the problem that a coolant in a sweating cooling system of a hot end component of an aircraft cannot be stably supplied in a variable incoming flow environment in the prior art.

The invention discloses a thermally-induced flexible composite phase change material coupled multi-crosslinking network aerogel material and thermal protection application, and belongs to the technical field of new materials. The coupling material is composed of multiple cross-linked network aerogel, a thermally-induced flexible composite phase change material and a binder. Wherein the thermally-induced flexible composite material is a hydrated inorganic salt phase-change material, the binder is a flame-retardant inorganic binder, and the multi-crosslinked network aerogel is a sodium alginate / gelatin composite material and is crosslinked through a double-crosslinking agent to form a stable multi-network structure for improving the flame retardance and mechanical strength of the aerogel. The prepared coupling material has high flame retardance and good heat insulation performance, and is very suitable for large-scale application in the field of thermal protection in buildings, batteries and electronic devices in the future.

The invention belongs to the technical field of composite material forming, and relates to a cabin thermal protection structure forming method which comprises the following steps: designing a needled felt; the needled felt is prepared; loading the needled felt into a mold and sealing; hybrid phenolic resin is injected into the mold; heating the mold to cure the hybrid phenolic resin; obtaining, heating and drying the heat-proof layer in the mold; the bonding face of the heat-proof layer and the structural part is coated with a primer; coating a silicon-based adhesive on the bonding surface; bonding and curing the heat-proof layer and the structural member; and processing the bonding allowance. According to the cabin thermal protection structure forming method provided by the invention, the heat-proof sleeve and the force-bearing structure are separately prepared, and then the heat-proof sleeve and the force-bearing structure are connected into the integral part in a sleeving manner, so that the heat-proof sleeve and the structural part can be produced in parallel, the production efficiency of products is greatly improved, and the production period is shortened.

The utility model discloses a template embedded type dam face permanent heat preservation protection module. Comprising a heat preservation core layer, a facing anti-seepage protection layer is arranged on the side, close to a concrete pouring formwork, of the heat preservation core layer, an internal pressure-bearing protection layer is arranged on the side, close to dam surface layer crack limiting steel bars, of the heat preservation core layer, and a thermal broken bridge supporting keel is arranged in the heat preservation core layer; an anchoring device is arranged in the area between the internal pressure-bearing protection layer and the dam surface layer crack limiting steel bars. The modular prefabrication and construction method is adopted, the prefabrication and installation method of a single module is simple and efficient, details such as anchoring protection measures are fully considered, the engineering construction operability is high, convenience and high efficiency are achieved, and the construction time of the heat preservation layer is greatly saved.

The invention discloses a preparation method of a toughening environmental barrier coating, and relates to a preparation method of a coating. The invention aims to solve the problem of poor fracture toughness of the existing environmental barrier coating. The method comprises the following steps: 1, pretreating a base material; 2, preparing a bonding layer; and 3, preparing the environmental barrier layer. The high-melting-point metal silicide prepared on the surface of the base material or the bonding layer compounded by the high-melting-point metal silicide and the rare earth silicate has excellent temperature resistance and good interface bonding strength, and due to the addition of the rare earth silicate, the expansion coefficient of the bonding layer is highly matched with that of the base material; according to the prepared yttrium oxide stabilized zirconia fiber, mullite fiber or silicon carbide fiber reinforced rare earth silicate environmental barrier layer, the fracture toughness of pure Yb2SiO5 is improved to 6.5 MPa.m < 1 / 2 >-8.0 MPa.m < 1 / 2 > from about 1.85 MPa.m < 1 / 2 >. The strengthening and toughening environmental barrier coating prepared by the invention has very important significance on development of thermal protection coatings of hot end parts of aircrafts with high thrust-weight ratios in China.

The thermal protection structure comprises a first porous medium layer arranged on the arc-shaped wall face of the front edge of the aircraft and a plurality of second porous medium layers arranged on the upper wall face and the lower wall face of the front edge of the aircraft at intervals in the flow direction. Solid walls are arranged between the first porous medium layers and the second porous medium layers and between every two adjacent second porous medium layers; the porosity of each second porous medium layer is firstly increased and then reduced along the flow direction, and the porosity of the first porous medium layer is smaller than that of the second porous medium layer at the most upstream. The method is applied to the field of thermal protection of aircrafts, the problem of local overheating existing in a traditional sweating cooling technology is effectively solved through the synergistic effect of the intermittent sweating layout and the gradient porosity material, and the situation that the cooling efficiency is reduced due to overheating of a cooling agent is avoided.

The invention provides a preparation method of an ultrahigh-temperature oxidation ablation resistant tantalum-based ceramic and silicon carbide composite thermal protection coating, which comprises the following steps: (1) mixing a polymer binder and a solvent to form a solution, and carrying out ball-milling mixing on the solution, a tantalum-containing compound, silicon carbide, silicon and other components to prepare precursor slurry; (2) uniformly coating the surface of a carbon substrate with the slurry obtained in the step (1) by a dip-coating method, and drying; (3) realizing curing and pyrolysis of the coating through segmented heat treatment; and (4) the coating is subjected to vacuum siliconizing treatment, and the composite thermal protection coating is obtained. According to the tantalum-based ceramic and silicon carbide composite thermal protection coating prepared through the method, the ultrahigh-temperature oxidation resistance, ablation resistance, thermal shock resistance and heat insulation performance of the coating are remarkably improved. After the coating is ablated for 2400s under oxyacetylene flame at the temperature of 1650 DEG C, the integrity and excellent thermal protection effect of the coating can still be kept.

The invention discloses a medium-high temperature heat-power conversion working medium for a high-speed aircraft as well as a preparation method and application of the medium-high temperature heat-power conversion working medium. The medium-high temperature heat-power conversion working medium comprises the following components in percentage by mass: 10-90% of ammonia, 10-70% of multi-component alcohol and the balance of water, the multi-component alcohol comprises at least two of ethanol, isopropanol, ethylene glycol, n-propyl alcohol and n-butyl alcohol; and the mass of each alcohol in the multi-component alcohol is 5-40% of the total mass of the medium-high temperature heat-work conversion working medium. Ammonia, multi-component alcohol and water are combined for the first time, a double-bed catalyst is combined, physical heat absorption can be utilized, heat absorption can be enhanced through chemical reaction, the heat sink can reach 7.0 MJ / kg or above, and the cooling agent can be used for pneumatic heat absorption of a high-speed aircraft with the speed higher than 5 Mach and is high in cooling capacity and low in cost. The method is suitable for a thermal protection system of a pneumatic heating component with higher heat flux and continuous long time; meanwhile, high-temperature and high-pressure small-molecule gas is generated, thrust can be generated to do work, and the maneuverability of the high-speed aircraft is effectively enhanced.

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.

A design and optimization method for a thermal protection system of a high-speed aircraft considers uncertainty of physical property parameters, heat flux density, geometric shape and the like and probability distribution characteristics of the parameters in design and application of the thermal protection system, the uncertainty and the probability distribution characteristics are embedded into a design process, and uncertainty analysis and design optimization of the thermal protection system are achieved; the method comprises the following steps: performing uncertainty modeling on small subsample characteristic parameters such as material attribute parameters, input environment parameters and machining process parameters in thermal protection design by utilizing a nuclear density theory to realize accurate modeling of input parameter uncertainty distribution characteristics; design parameters are optimized through the agent model based on physical information, so that a design model and a result have interpretability, and very high extrapolation and generalization capabilities are achieved.