1415 results about "Thermal protection" patented technology

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.

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 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 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 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 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.

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 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 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 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.

The invention discloses a high-strength heat-insulation aerogel structure with a dual multi-scale design and a preparation method thereof, relates to the technical field of thermal protection of spacecrafts, and in particular relates to the high-strength heat-insulation aerogel structure with the dual multi-scale design and the preparation method thereof. The invention aims to solve the problems of insufficient structural rigidity and poorer thermal protection property and thermal stability of the traditional spacecraft surface aerogel structure under the conditions of extreme high temperature and mechanical load. The structure is composed of an upper carbon-carbon composite material plate, a lower carbon-carbon composite material plate, an alloy skeleton and a filler. The method comprises the following steps: 1, preparing a multi-scale C / C composite material; 2, designing and preparing a titanium alloy honeycomb framework; 3, proportioning the multi-scale aerogel; 4, in-mold in-situ gelation; 5, drying and heat treatment; and 6, assembling and solidifying. The porosity of the aerogel is 85-98%, the structure has high rigidity and thermal protection performance, and the aerogel is applied to the field of aerospace thermal protection.

The invention relates to a multi-scale full-coupling transient prediction method for a porous interface phase change sweating cooling effect, and belongs to the technical field of thermal protection of a leading edge end material of a hypersonic aircraft. The problems that in aircraft phase change sweating cooling numerical simulation in the prior art, phase change porous interface full coupling information is lost, and heat and drag reduction effect prediction is not accurate are solved. The method comprises the following steps: determining physical parameters, boundary conditions and initial conditions of a porous interface; constructing a porous interface phase change sweating full-coupling model; obtaining a predicted porous interface phase change sweating multi-physical field based on a porous interface phase change sweating full coupling model; and based on the porous interface phase change sweating multi-physical field, the cooling efficiency of a porous interface phase change sweating blowing area and a blowing trailing edge area, the porous interface phase change sweating transient blowing ratio, the total consumption of a coolant and the porous interface phase change sweating friction resistance reduction effect are obtained.

The invention belongs to the technical field of hypersonic wind tunnel tests, and discloses a hypersonic wind tunnel test model support with an active cooling function and a test method thereof. A cooling water inlet and a cooling water outlet are formed in an inclined column of the bracket, so that the bracket has an active cooling function and can effectively support a test model to carry out a hypersonic wind tunnel test; the rear portion of the support is provided with a cavity defined by the cross beam, the skin and the like and a wire outlet hole of the tail cover, and the cavity and the wire outlet hole can be used for wiring of a test model measurement and control cable, the measurement and control cable is isolated from high-speed airflow, and thermal protection is achieved. The test method comprises the steps of mounting the batter post; installing a tail cover; the sealing plate is mounted on the inclined column; installing a left cross beam and a right cross beam; installing the skin; installing a test model; mounting a bracket; and performing a hypersonic wind tunnel test. The test method can meet the requirement of a hypersonic wind tunnel test, is compatible with an existing hypersonic wind tunnel test method, is stable and reliable, is high in efficiency, and has engineering practical value.

The invention relates to a method for testing the heat insulation temperature of a thermal protection coating. The method for testing the heat insulation temperature of the thermal protection coating comprises the following steps: S1, metal matrixes of a first sample and a second sample are the same, the surfaces of the metal matrixes of the first sample and the second sample are coated with metal bonding layers, the metal bonding layer of the second sample is coated with a ceramic layer, and thermocouples are arranged on the surfaces of the metal matrixes of the first sample and the second sample; s2, arranging the first sample and the second sample at mounting positions of a heating furnace; s3, starting the heating furnace, raising the temperature to a test temperature, and continuously introducing cooling air with set flow and flow velocity into the surfaces of the metal matrixes of the first sample and the second sample; and S4, after the temperature of the heating furnace is constant, the metal matrix surface temperatures of the first sample and the second sample are obtained through a thermocouple, and the temperature difference between the metal matrix surface temperatures is the actual heat insulation temperature of the thermal protection coating under the current condition. The invention provides a method for testing the thermal insulation temperature of a thermal protection coating. The thermal insulation performance of the thermal protection coating can be rapidly and accurately tested.

The invention discloses a normal-temperature curing ablation-resistant coating and a preparation method thereof, the normal-temperature curing ablation-resistant coating comprises a component A, a component B and a component C, the component A comprises the following raw materials by weight: 20-30 parts of epoxy resin, 30-40 parts of liquid nitrile rubber, 27-50 parts of an ablation-resistant filler, 10-20 parts of a heat-insulating functional filler, and 200-300 parts of toluene; the component B is a curing agent; and the component C is a curing accelerator. By selecting the high-performance polymer-based resin material and adding the ablation-resistant filler and the heat-insulating functional filler, the ablation-resistant and heat-insulating coating protection construction of the equipment under the normal-temperature curing condition is realized; the problems of ablation-resistant coating construction and thermal protection when a fragile cover, a launcher and the like in a launching system cannot use a high-temperature curing system ablation-resistant coating are solved. The coating can realize military technology and civil use, and can be widely applied to fire prevention and heat insulation construction of buildings.

The utility model provides a fault simulation training device for an overheat protection circuit of a civil aviation air conditioning component, comprising a stand body, a central console, and an air conditioning component overheat protection fault simulation circuit module comprising a dark box and a circuit panel; the dark box is connected to the central console in communication; the dark box includes an air conditioning component overheat protection fault simulation circuit, which includes a DC power supply, a thermal relay, a thermistor module, a reset switch module, and an overheat indication module; the DC power supply is connected to one end of the thermistor module via a coil of the thermal relay, the other end of the thermistor module is grounded, the reset switch module and the overheat indication module are respectively connected to the first make contact and the second make contact of the thermal relay; the circuit panel is provided with a connection schematic diagram corresponding to the fault simulation circuit and a plurality of fault measurement holes. The utility model can simulate various circuit fault points in the air conditioning component overheat protection fault simulation circuit module, and realize direct measurement of circuit faults.

The present invention provides a method for preparing a composite coating of a carbon fiber thermal insulation material, the method comprising: purifying the carbon fiber, then doping the surface with a composite metal element, placing the doped carbon fiber in a high-temperature furnace for surface carbon coating, cracking the organic carbon source on the surface of the carbon fiber under the catalytic action of the composite metal element to form amorphous carbon, then performing surface modification to obtain surface-modified carbon fiber, and then stirring and mixing with the components in the coating, applying, and obtaining a composite coating of a thermal insulation material containing carbon fiber. The carbon fiber of the present invention has good performance, and through the catalysis of the composite metal during the pyrolysis process, amorphous carbon with a porous structure is formed on the surface, which has good thermal insulation ability. The coating of the present invention has extremely high thermal insulation ability, corrosion resistance, wear resistance and adhesion ability, and can be used for thermal insulation coatings in various thermal protection systems.

The invention relates to the technical field of thermal protection coatings, and discloses a high-reflection thermal protection coating and a preparation method thereof.LSTN0.125 powder is subjected to spray granulation and then pre-sintered and sieved, and spherical LSTN0.125 ceramic powder is obtained; the alloy substrate is subjected to roughening treatment, and a bonding layer is prepared on the surface of the alloy substrate through the atmosphere plasma spraying technology; the spherical LSTN0.125 ceramic powder obtained in the step S1 is sprayed on the bonding layer through the atmosphere plasma spraying technology, and an LSTN0.125 ceramic layer is obtained; and the LSTN0.125 ceramic layer is subjected to heat treatment, and the LSTN0.125 thermal protection coating is obtained. According to the invention, the LSTN0.125 with high reflectivity is used as a powder raw material of the coating, so that the high reflection performance of the coating in a near-infrared band can be remarkably enhanced, and the effect of isolating high temperature is achieved.

The invention discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic thermal constraint, which comprises the following steps of: determining a rudder leading edge radius and a rudder leading edge sweepback angle by taking a current state nominal trajectory as a basis and taking a rudder leading edge thermal response temperature as a judgment standard; secondly, the cone angle of one conical surface of the folded rudder is determined through iterative calculation, so that the chordwise size of the high-temperature-resistant refractory metal material and the thickness of a skin and a heat-proof coating at the cone position are designed; and thirdly, through the design of the interference position of the arched shock wave of the heat-proof cup, the gap between the fixed rudder and the movable rudder avoids the interference area of the arched shock wave, so that the chordwise distance between the folding gap and the rudder shaft is determined. By means of the aerodynamic configuration design method, the radius of the front edge of the rudder can be as low as 3 mm, the taper angle is as low as 4 degrees, it can be guaranteed that ablation deformation does not occur when the maximum heat flow of the front edge of the rudder does not exceed 10 MW / m < 2 >, the resistance characteristic of a high-speed aircraft can be effectively reduced, the characteristic of near-zero ablation is achieved, and the good aerodynamic configuration is maintained.

The invention discloses an antioxidant carbon aerogel material and a preparation method thereof. The method comprises the following steps: sequentially adding a solvent, a boron-containing modifier and a silicon-containing modifier into phenolic resin for esterification reaction; sequentially curing, drying and carbonizing to obtain boron-silicon modified carbon aerogel; depositing a SiC inner coating on the boron-silicon modified carbon aerogel under the assistance of supercritical CO2; a Zr-Si ceramic oxygen barrier layer and a SiO2 sealing and filling layer are sequentially prepared through a sol-gel method, and the carbon aerogel material with the SiC-Zr / Si-SiO2 composite coating is prepared. The invention also provides the aerogel material prepared by the method, the oxidation resistance of the aerogel material is obviously improved, and the aerogel material can be used for preparing a thermal insulation material applied to the field of high-temperature thermal protection.

According to the graphene fiber heat conduction enhanced high-temperature-resistant composite material and the preparation method thereof, graphene fibers are adopted as a substrate and can be violently activated in a high-temperature environment, a large number of nucleation sites are formed to grow a carbide coating on a surface layer, and the anchoring effect is very strong through covalent bonds between the nucleation sites and the substrate. Furthermore, the modulus of the graphene fiber is low, and in the rapid cooling process, huge thermal stress generated in the cooling process due to the fact that the thermal expansion coefficients of the fiber substrate and the carbide coating are not matched can be effectively dissipated, so that stable combination between the coating and the substrate can be guaranteed. The obtained composite material has stable interface bonding and extremely high thermal conductivity, and can be applied to a spaceflight thermal protection system.

The utility model relates to a thermal protector which is fast in disconnection and strong in bearing capacity. The thermal protector is mainly used for over-temperature rise and over-current protection of a refrigeration compressor or other motors. The heating device comprises a pin, a cover plate, a movable contact spring, a bimetallic strip, a first static pin, a second static pin, a heating body, a third static pin and a base, the two ends of the heating body are welded to the first static pin and the second static pin respectively, one end of the movable contact spring is electrically connected with the first static pin, the other end of the movable contact spring is matched with the third static pin, the pin is electrically connected with the third static pin, and the bimetallic strip is arranged on the base. The heating device is structurally characterized in that a loading cavity used for installing the heating body is formed in the base, a limiting table used for placing the heating body is arranged at the bottom of the loading cavity, the heating body is a spiral heating body or a planar heating body, and when the heating body is the spiral heating body, the limiting table is of a plurality of semicircular protruding structures. When the heating body is a planar heating body, the limiting table is arranged to be of a structure with two strip-shaped protrusions.

The invention relates to a high-temperature-resistant temperature and vibration composite sensor applied to the field of measurement and testing, which comprises a composite sensor body and a composite circuit board arranged in the composite sensor body, and adopts the cooperation of a thermal insulation sleeve, a magnetic induction ring, a thermal insulation protection feedback unit and a ventilation regulation and control unit. According to the composite sensor, the adaptability of the composite sensor body to the environment is effectively promoted, the composite sensor body can be effectively isolated and protected under the high-temperature condition, and the situation that components in the composite sensor body are damaged due to continuous heat transfer is prevented, so that the durability of the composite sensor body is remarkably improved; meanwhile, heat exchange protection can be effectively carried out on the internal environment of the composite sensor body, the influence of heat on all components in the composite sensor body is further reduced, and the heat resistance of the composite sensor body is enhanced.