1158 results about "Thermal protection" patented technology

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.

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 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 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 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 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 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 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 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.

The invention provides a phase change enhanced ablation-resistant modified liquid silicone rubber material as well as a preparation method and application thereof, and belongs to the field of advanced materials. According to the invention, paraffin is adopted as a phase change filler, so that the ablation resistance and heat insulation performance of the expanded ceramic liquid silicone rubber material are remarkably improved, and the prepared silicone rubber composite material has the advantages of low density, low ablation surface temperature, strong carbon layer and high heat absorption capacity. The phase-change reinforced modified expanded ceramic flexible ablation-resistant material further improves the ablation resistance and high-temperature heat insulation performance of a liquid silicone rubber material, and has a good application prospect in the fields of aerospace thermal protection, building heat insulation and the like.

The present invention discloses a porous material transpiration cooling leading edge structure, which includes multiple layers of porous material layers and a cooling cavity provided on one side of the multiple layers of porous material layers. A coolant supply channel is provided in the cooling cavity, and the coolant supply channel is used to supply coolant to the cooling cavity. The multiple layers of porous material layers include an outer layer, an intermediate layer, and an inner layer, and the material of the inner layer is a shape memory alloy porous material. The present invention can adaptively adjust the porosity of the porous material, automatically adjust the usage amount of the coolant, effectively improve the usage efficiency of the coolant, and further enhance the thermal protection effect in the high-temperature tip region. Moreover, the shape memory alloy material can be reused. The porous material transpiration cooling leading edge structure can be used for structures such as the nose cone of an aircraft, the leading edge of a rudder, and the leading edge of a wing that are subjected to more high-temperature heat fluxes.

The invention discloses a variable geometry turbine guider blade adjusting structure and method and application, through cooperative adjustment of a slidable clamshell-shaped suction surface and a rotatable tail flap, the front thickness, the aerodynamic configuration and the throat area of a turbine blade are accurately controlled, and therefore wide-range flow adjustment is achieved. The pressure surface side fixing part provides a stable supporting and flowing path through a pressure surface section, a front end groove and a tail semicircular arc surface which are integrally formed. The slidable clamshell-shaped suction surface is connected with an external driving source and can slide along the front-end groove to adjust the shape of the blade; the rotatable tail flap further optimizes the flow regulation by rotating to regulate the shape of the throat area at the tail of the blade. According to different working condition requirements, the clamshell-shaped suction surface and the tail flap can be adjusted independently or in a linkage mode, high efficiency and stability of the turbine guider blade under different flight working conditions are guaranteed, and meanwhile the turbine guider blade has good thermal protection performance and high-temperature tolerance.

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 belongs to the technical field of phase change materials, and particularly relates to an erythritol and fumed silica composite phase change material and a preparation method thereof.The preparation method comprises the following process steps that fumed silica and erythritol are mixed according to the mass ratio of 10: 90-30: 70, and a mixture is obtained; heating the mixture at 130-160 DEG C for 2-4 hours, and stirring for 5 minutes every 30 minutes in the heating process; the heated mixture is subjected to dipping treatment under the vacuum condition, and composite phase change material powder is obtained; the composite phase-change material powder is subjected to compression molding under the pressure of 3-5 MPa, the pressure is maintained for 3-7 minutes, and the erythritol fumed silica low-heat-conductivity shaped composite phase-change material is obtained, the problem of phase-change material leakage is solved, the packaging cost is reduced, meanwhile, the heat conductivity of the phase-change material can be reduced, the phase-change material has the heat insulation and heat storage capacity, and the thermal conductivity of the phase-change material is improved. The application in the field of thermal protection is facilitated.

The invention discloses a chromium-doped indium oxide-based high-temperature heat flux sensor and a preparation method thereof, a screen printing technology is adopted, a slurry formula and printing process parameters are optimized, a thermopile structure is accurately formed on a high-temperature-resistant ceramic substrate, and then a thermopile and a thermal resistance layer are subjected to step-by-step high-temperature sintering treatment, so that the chromium-doped indium oxide-based high-temperature heat flux sensor is obtained. And firm combination between the thin film and the substrate and densification of a microstructure are ensured. The finally prepared sensor has excellent high-temperature oxidation resistance and long-term working stability, is stable and reliable in thermoelectric output signal, meets the strict requirements of aerospace, energy power and other fields on high-temperature heat flow measurement, is suitable for heat flow monitoring in high-temperature combustion environments, high-speed pneumatic heating and other scenes, and has wide application prospects. And a reliable test means is provided for design and performance evaluation of the thermal protection system in an extreme environment.

The invention relates to the technical field of textile materials, and particularly discloses a fabric capable of improving thermal protection performance and a preparation method and application of the fabric. Nano-scale aluminum oxide; polyethylene glycol diacrylate; a closed type waterborne polyurethane emulsion; a silane coupling agent containing a vinyl group; a photochromic dye; a thermochromic dye; the preparation method comprises the following steps: S1, preparing composite slurry; s2, applying the composite slurry to the surface of a fabric base material in a dot matrix form by adopting a screen printing process to form a coating; s3, the coating is irradiated with ultraviolet light; and S4, carrying out baking treatment on the fabric subjected to the ultraviolet light curing treatment at the temperature of 140-160 DEG C. According to the invention, the polyethylene glycol diacrylate, the closed waterborne polyurethane and the vinyl-containing silane coupling agent are adopted to form an interpenetrating polymer network structure in situ, so that when the obtained fabric is rubbed or washed and the like, the coating is not easy to crack and peel off.

The invention discloses a method for increasing the phase transition temperature of Al2O3 aerogel through grain boundary pinning and nano-structure strengthening, which comprises the following steps: co-hydrolyzing an aluminum source, yttrium salt, a silicon source and magnesium salt to form sol through the coordination of double mechanisms of grain boundary pinning and nano-domain isolation, and carrying out gradient heating aging, mixed medium supercritical drying and two-stage sintering to obtain the phase transition temperature of the Al2O3 aerogel. The gamma-alpha phase transition temperature of the Al2O3 aerogel is greater than or equal to 1300 DEG C. The specific surface area of the aerogel after heat treatment at 1300 DEG C is larger than 100 m / g, the process cost is controllable, and the aerogel is suitable for industrial production and can be applied to extreme environments such as a spacecraft thermal protection system and a high-temperature industrial furnace lining.

The utility model discloses a thermal protector temperature testing device which comprises a testing assembly, a control panel and a wind wheel motor, the testing assembly comprises a wind barrel, a barrel cover assembly, an air duct basin, a centrifugal wind wheel and a socket frame, the socket frame is provided with conductive clamping pieces used for being electrically connected with a thermal protector, and the conductive clamping pieces are uniformly distributed in the circumferential direction. The barrel cover assembly covers the air barrel, the centrifugal wind wheel is arranged on the lower side of the air duct basin, a heating pipe is arranged on the outer side of the centrifugal wind wheel, the wind wheel motor is in driving connection with the centrifugal wind wheel, the conductive clamping piece is arranged above the air supply duct correspondingly, a ventilation opening is formed in the bottom of the air duct basin, and the position of the air suction end of the centrifugal wind wheel corresponds to the ventilation opening. A temperature sensor is arranged on the lower side of the socket frame, on-off indicating lamps are arranged on the control panel, the conductive clamping pieces are electrically connected with the corresponding on-off indicating lamps respectively, and the control circuit is electrically connected with the temperature sensor and the heating tube. The temperature testing device for the thermal protector is beneficial to improving the working efficiency and improving the effectiveness of quality inspection.

Electrically heated motorcycle glove with temperature overload control and uniform heating, characterized in that it comprises a back of the hand region (1), a palm region (2), a lacing (3) and a wrist region (4), wherein the back of the hand region (1) has a back of the hand base surface (11) in the interior of which a heating layer (14) is installed, in the central region of which an adjusting element (15) is installed, wherein the palm region (2) has a palm base surface (21) in the interior of which a heating layer (14) is installed, wherein the wrist region (4) has an extended thermal protection layer (41), above which an operating switch (42) is attached, and in the lower region of the extended thermal protection layer (41) a battery compartment (45) is formed, in the interior of which a rechargeable battery (46) is accommodated;wherein the back of the hand base (11) and the palm base (21) are sewn with sewing thread, the back of the hand base (11) and the extended heat protection layer (41) are sewn with sewing thread, and the palm base (21) and the extended heat protection layer (41) are sewn with sewing thread;

A thermal protection structure of a high-temperature and supersonic aircraft belongs to the technical field of thermal protection of aircrafts under high heat flux and comprises a water absorption layer, a front edge area and two side areas, and the front edge area comprises a first evaporative cooling porous medium layer, a cross-scale porous medium layer and a high-temperature-resistant wall surface which are sequentially laid on the outer side of the front edge of the water absorption layer; the two side areas comprise a second evaporative cooling porous medium layer and a third evaporative cooling porous medium layer which are laid on the outer sides of the two sides of the water absorption layer respectively, and a coolant conveying pipeline communicated with the front edge and the two side areas is arranged on the inner side of the water absorption layer; fluid is sucked to the high-temperature-resistant wall face through the capillary acting force of micron-scale and nano-scale pores of the first evaporative cooling porous medium layer and the cross-scale porous medium layer, the millimeter-scale pores of the cross-scale porous medium layer provide steam escape channels, and the high-temperature-resistant wall face blocks direct disturbance of high-temperature main flow to the fluid in the porous medium. Steam outlets on two sides of a front edge stagnation point discharge steam; the device has the advantages of no need of an additional liquid supply power device, high temperature resistance and high cooling efficiency.

The invention discloses a ceramic component gradient modified C / C-SiC-ZrC composite material and a preparation method thereof, and belongs to the technical field of carbon / carbon composite material matrix modification preparation, a carbon fiber preform and a heating element are placed in chemical liquid phase vaporization deposition equipment, a precursor solution is added until the carbon fiber preform is fully soaked, and a ceramic component gradient modified C / C-SiC-ZrC composite material is obtained; and carrying out chemical liquid phase vaporization deposition reaction to obtain the low-density C / C composite material with porosity in unidirectional gradient distribution. And putting the mixture into a graphite crucible with a carbon-rich area facing downwards and a carbon-less area facing upwards, completely putting the graphite crucible into the ceramic powder for modification, reacting and infiltrating, and cooling to room temperature to obtain the C / C-SiC-ZrC composite material with the gradient modified ceramic components. The end with high ceramic content endows the material with a high-temperature thermal protection function, the end with high pyrolytic carbon content can effectively conduct heat generated in a ceramic-rich area, and the end with low ceramic content can effectively conduct heat generated in a ceramic-rich area. Due to the gradient design from the ceramic-rich area to the carbon-rich area, the material has high-temperature oxidation resistance and ablation resistance and is light, and lightweight design of an aerospace thermal protection structure is facilitated.

The present invention discloses a fireproof and explosion-proof cable, comprising a conductor, an inner shielding layer, an insulating layer, an outer shielding layer, a filling layer, a tape layer, an armor layer, and an outer sheath. The outer wall of the tape layer is provided with an isolation layer, and the outer wall of the isolation layer is provided with a protective layer. The isolation layer and the protective layer are provided in multiple groups and are equidistantly distributed along the length of the cable. Due to the provision of the protective layer, when a fire occurs and the temperature around the outer sheath rises significantly, the generated high temperature will be absorbed by the expanded graphite in the first expansion chamber. After the expanded graphite expands, the first expansion chamber drives the concave block to move toward the convex block, so that the vent holes and the connecting holes are staggered and blocked, forming a heat-insulating protective layer, preventing the external high temperature from entering the cable through the vent holes and the connecting holes, and preventing the high temperature from entering the conductor core, causing the conductor core to overheat, causing a short circuit, or even the risk of explosion, thereby improving the safety and service life of the cable.

Embodiments include thermal covers to help drain fluids away from dangerous environments, e.g., in an engine bay in a vehicle. Thermal covers can comprise a hat portion extending between two distal edges, support straps beneath the hat portion and extending between the distal edges, and thermal insulation disposed between the hat portion and the support straps. These can be coupled to the floor so as to allow space below for the flow of fluid. This allows the fluid to flow to holes in the floor of the engine bay, allowing the fluid (which could be flammable) to be drained away from potentially dangerous environments in the engine bay which may involve high temperatures.

The invention provides a graphene oxide and aramid nanofiber hybrid modified chopped fiber reinforced flexible polymer thermal protection material as well as a preparation method and application thereof, and belongs to the field of advanced materials. Aiming at the problems of huge modulus difference, poor interface compatibility and serious stress concentration between fibers and a flexible matrix, a hybrid interface phase composed of graphene oxide and aramid nanofibers is constructed, so that an inert surface microstructure of the chopped fibers is improved, and a multi-scale mechanical interlocking structure is promoted to be formed between the fibers and the flexible matrix; the modulus mismatch between the fiber and a rubber matrix is effectively relieved, and more uniform and efficient stress transfer of an interface area is realized. In the processes of tensile deformation, dynamic loading, high-temperature ablation and the like, the hybrid interface phase can keep tight interface combination between the fibers and the flexible matrix, the ablation anti-scouring performance and mechanical performance of the flexible composite material are remarkably improved, especially the flexible ablation performance in the high-dynamic and deformation state, and the flexible composite material has good application prospects.

The invention relates to the technical field of functional materials, in particular to phenolic aldehyde hybrid SiO2 mesoporous microspheres and application thereof. Mesopores with the pore diameter of 2-50 nm are formed in the surface and the interior of the phenolic aldehyde hybridized SiO2 mesoporous microsphere; the average diameter of the phenolic aldehyde hybrid SiO2 mesoporous microspheres is 10 to 100 microns; the phenolic hybrid SiO2 mesoporous microspheres can be applied to the fields of thermal protection, wave absorption, heat insulation and coatings. According to the phenolic hybrid SiO2 mesoporous microsphere disclosed by the invention, by virtue of a unique mesoporous structure and an organic-inorganic hybrid characteristic, excellent heat insulation performance and multifunctionality are realized, and meanwhile, the problems of insufficient protection performance, single function, poor processability, high cost and the like of an existing material are solved.

The invention relates to the technical field of bus duct thermal protection, in particular to a bus duct thermal protection method and system based on intelligent monitoring, and the method comprises the steps: obtaining a packet loss delay bandwidth generation state label, calling an unstable link collection path node congestion rate port number to generate a scheduling table, collecting a temperature rise rate displacement, calculating a thermal slope and combining field disturbance to obtain a trigger value, and comparing thermal triggering with thermal accumulation to identify a high-temperature-sensitive section mark, and calling the mark to compare a cutoff threshold value to generate a starting result. By collecting multi-dimensional data such as packet loss, delay and bandwidth for dynamic analysis, the state of a communication link can be evaluated in real time, an unstable link can be found, an accurate triggering boundary value is generated by combining parameters such as thermal load, electric field intensity and harmonic waves, cooling or power-off protection is started in time by matching with a heat flow trend before a high-temperature-sensitive area appears, and the reliability of the system is improved. Response lag of a traditional technology is avoided, an intelligent response mechanism is innovated to achieve accurate and flexible thermal protection, stable and safe operation of equipment is guaranteed, and overheating accidents are reduced.