34 results about "Aerodynamic heating" patented technology

The invention provides an inclined variable-angle aircraft interstage separation ground test device. The inclined variable-angle aircraft interstage separation ground test device comprises an aircraft separation device, an angle adjusting device, a temperature control and acquisition device, a vibration control and impact acquisition device, an ignition control device and a device base. The angle adjusting device comprises a device side plate, an arc groove, an angle ruler, a separation mechanism mounting plate, a supporting hinge, a supporting rod, an adjusting sliding rail, an adjusting base, a locking device, a device base gradienter, a rotating bearing and an auxiliary fixing block. The temperature control device comprises a quartz lamp array heating cage, an outer wall temperature control thermocouple and an inner wall temperature acquisition thermocouple, and can simulate a pneumatic heating environment at a separation moment, so that the thermal deformation matching capability of the separation mechanism in a high-temperature environment and the operation reliability of the separation mechanism are obtained. The vibration control and impact acquisition device is composed of a vibration table, a vibration control sensor and an impact sensor.

The invention discloses a method for quickly calculating transient thermal stress of a ceramic radome along a flight path, which comprises the following steps of: determining calculation input which comprises an aerodynamic thermal environment of the ceramic radome and a to-be-evaluated structure scheme; setting a circumferential discrete strategy of the ceramic radome to determine a plurality of meridian planes of the ceramic radome; constructing an aerodynamic thermal environment interpolation model, and determining a meridian plane where aerodynamic heating of the ceramic antenna housing is most strict in the whole flight process; constructing a two-dimensional feature structure model for transient heat transfer and thermal stress analysis of the ceramic radome, and determining an outer contour line of the two-dimensional feature structure model; constructing an axial thermal response calculation interval of the ceramic radome and performing interpolation calculation to obtain aerodynamic thermal environment parameters of the outer contour line at each flight moment; performing heat transfer calculation grid division on the two-dimensional feature structure model, and calculating structure transient temperature distribution of the two-dimensional feature structure model; calculating the temperature gradient of the structure based on the transient temperature distribution of the structure; and calculating transient thermal stress distribution of the ceramic radome by using the structural temperature gradient.

A thermal gradient management assembly comprising a functional component. The functional component includes an exterior facing segment, where the exterior facing segment is configured to experience aerodynamic heating, and an interior segment with the exterior facing segment, where the interior segment is configured for isolation from the aerodynamic heating. The functional component includes a thermal gradient extending between the exterior facing segment and the interior segment. The thermal gradient management assembly also comprises a gradient mitigating heating system includes a power source and a gradient heating element in communication with the power source, wherein the gradient heating element is coupled with the interior segment. The gradient heating element is configured to diminish the thermal gradient between the exterior facing and the interior segments.

The application discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, determines a rudder leading edge radius and a rudder leading edge sweepback angle based on a current state nominal trajectory and a rudder leading edge thermal response temperature as a judgment standard; secondly, a conical angle of a folding rudder-cone surface is determined through iterative calculation, so that a chordwise dimension of a high-temperature-resistant refractory metal material, and a thickness of a skin and a heatproof coating at the cone are designed; thirdly, through interference position design of a heatproof cup bow shock wave, a gap between a fixed rudder and a movable rudder avoids the bow shock wave interference zone, so that a chordwise distance between a folding gap and a rudder shaft is determined. Through the aerodynamic configuration design method, the rudder leading edge radius can be as low as 3mm, and the cone angle can be as low as 4°, so that ablation deformation does not occur when the maximum heat flow of the rudder leading edge is less than 10MW / m 2 2, the resistance characteristics of the high-speed aircraft can be effectively reduced, the aircraft has a near-zero ablation feature, and a good aerodynamic shape is maintained.

The invention discloses an inflatable returner structure safety analysis method and system considering a thermo-pneumatic elastic effect, and the method comprises the steps: building a structural dynamics sub-model, a fluid domain calculation sub-model and a structural thermodynamics sub-model of an inflatable returner, and comprehensively considering the structural thermal deformation caused by aerodynamic heat and the thermal expansion effect of an internally inflated gas; and a bidirectional coupling relationship among aerodynamic force, aerodynamic heat, structural deformation, temperature and internal charging pressure is constructed through the feedback effect of structural deformation on a flow field and an aerodynamic load. And carrying out multi-physics field simultaneous solution by adopting a time step propulsion method, and finally realizing comprehensive safety evaluation of the temperature, stress and deformation characteristics of the structure. According to the method, aerodynamic load distribution, temperature field change, stress response and vibration characteristics of the inflatable returner in the reentry process can be accurately predicted, so that the thermal protection performance and structural safety of the inflatable returner are comprehensively evaluated, and reliable analysis means and engineering basis are provided for optimization design and structural stability improvement of a thermal protection system.

A missile telemetry antenna array based on feed phase difference and screw coupling, comprising: inverted F antenna, missile cabin, rectifier block, RF cable and transmitter; a plurality of inverted F antennas are evenly distributed around the central axis of the missile cabin, and the inverted F antennas are fixedly installed on the outer wall of the missile cabin; the inverted F antennas are connected with the transmitter through the RF cable; a rectifier block is arranged above the head of each inverted F antenna; the rectifier block is fixedly installed on the outer wall of the missile cabin; and the rectifier block is used for flow guiding, so that no stagnation point of aerodynamic heating appears on the inverted F antenna. The present application optimizes the gain depression of the axis direction of the missile body, realizes the protection of the aerodynamic heating and force of the antenna by installing the rectifier block in front of the antenna, generates electromagnetic resonance and induced current through the installation screw of the rectifier block, realizes the secondary radiation of electromagnetic waves and the reflection guiding effect.

A multifunctional integrated Mars reentry detector relates to the technical field of Mars detection and mainly comprises a detector body, a rotor wing, a nose cone cover, a rotating device, a heat shield and the like. During launching, the rotor wings are in a compressed and folded state, and the detector body, the nose cone cover and the heat-proof cover serve as a whole device fairing. In the initial stage of reentry, the heat shield is used for protecting the rotor blade from pneumatic heating, can be further unfolded to a certain angle after descending at a certain speed, and serves as a speed brake for further speed reduction. And when the speed is reduced to subsonic speed, the rotor blades are unlocked and released, the blades are unfolded through an internal telescopic mechanism, the blades are driven to rotate by utilizing relative incoming flow, and rotation kinetic energy is converted into electric energy to be stored while lift force is provided for deceleration. In the landing stage, the rotor wings are switched to an active working mode, stored electric energy is used for providing rotating power for the rotating device, and autonomous hovering and obstacle avoidance are achieved; and the heat shield is completely unfolded to provide stable support and buffer for the detector body.

The application discloses a medium-high temperature heat power conversion working medium for high-speed aircrafts and a preparation method and application thereof, and comprises the following components in percentage by mass: 10-90% of ammonia, 10-70% of multi-component alcohol, and the rest of water; the multi-component alcohol comprises at least two of ethanol, isopropyl alcohol, ethylene glycol, n-propanol and n-butanol; and the mass of each alcohol in the multi-component alcohol accounts for 5-40% of the total mass of the medium-high temperature heat power conversion working medium. The ammonia, multi-component alcohol and water are combined for the first time, and a double-bed catalyst is combined, so that not only the physical heat absorption can be utilized, but also the heat absorption can be strengthened through chemical reaction, the heat sink can reach more than 7.0 MJ / kg, the application can be used for the aerodynamic heat absorption of the high-speed aircrafts with the speed higher than 5 Mach, the cooling capacity is high, the application is suitable for the heat protection system of the components with higher heat flow density and continuous long-time aerodynamic heating; meanwhile, the high-temperature high-pressure small-molecule gas is generated, the thrust can be generated to do work, and the maneuverability of the high-speed aircrafts is effectively strengthened.

The application discloses a kind of hypersonic gas-solid interface atomic collision event statistics method.The method comprises: simulating the high-speed flight of object in gas environment in NAMD software, output the position and speed information of all atoms in system at different time;Using VMD software, all gas atoms that can interact with the front surface in a certain period after system steady state are counted, and the coordinates, speed and potential energy of these gas atoms are output;Define the time when the gas atom turns in the gas-solid atomic repulsive force layer as the collision point, the incident point of the gas atom into the attractive force layer is obtained by pre-searching from the collision point, and the reflection point when the gas atom leaves is obtained by post-searching, which is recorded as a collision event, and the collision event information of different atoms and surface is obtained by circulation.The application proposes a kind of gas-solid interface atomic collision event statistics method under hypersonic strong shock wave condition, which can study the energy transfer mechanism of gas-solid interface from a deeper level and reveal the nature of aerodynamic heating.

The invention discloses an aerodynamic heat management configuration design method for a large-airspace and wide-speed-domain hypersonic aircraft. The aerodynamic heat management configuration design method comprises the steps that a reference aerodynamic layout is determined, and appearance design oriented to heat flow distribution optimization is carried out; establishing a thermal environment database through multi-working-condition numerical simulation; dividing the surface of the aircraft into thermal management areas of different grades based on the thermal environment database; configuring a thermal protection material and designing a gradient thermal protection structure for each thermal management area to realize conformal integration with the reference aerodynamic layout; the integrity and reliability of the integrated structure under the heat / force load are verified through heat / structure coupling simulation analysis; and carrying out performance evaluation and iterative optimization, and if the design indexes are not met, returning to corresponding steps for optimization until the design is converged. Integrated collaborative design and closed-loop iterative optimization of aerodynamic configuration and thermal management performance are realized from a design source, and the problems of overweight of a thermal protection system, poor thermal environment adaptability and low design efficiency caused by a traditional serial method are effectively solved.

The application discloses a kind of low-temperature heat conversion green working medium and its preparation method and application, and green working medium includes the following mass percentage components: 25%~90% alcohol, 1%~5% alcohol amine, 1%~5% aromatic compound, the rest is water;Alcohol includes at least three kinds of methanol, ethanol, isopropanol, ethylene glycol, n-propanol and n-butanol.The working medium of the application is coupled with physical phase change, endothermic chemical reaction, and the heat sink can reach more than 6MJ / kg, which can be used for the absorption of aerodynamic heat of high-speed aircraft with a speed higher than 5 Mach, has the advantages of low cost and strong cooling capacity, and is suitable for the thermal protection system of high heat flux density and long-time aerodynamic heating components.In addition, high-temperature and high-pressure small molecule gas is generated for work, effectively enhancing the maneuverability of high-speed aircraft.

The invention relates to the technical field of low-vacuum pipeline traffic thermal protection and thermal management, in particular to a low-vacuum pipeline traffic system pipe wall thermal buffering and enhanced heat transfer design method which comprises the steps that geometric parameters of a low-vacuum pipeline traffic system are obtained, and a geometric model is constructed; determining operation parameters, structure parameters and external environment parameters of the low-vacuum pipeline traffic system; performing numerical simulation calculation on the basis of the geometric model, the operation parameters, the structure parameters and the outside-pipe environment parameters to obtain in-pipe aerodynamic thermal environment dynamic change characteristics; based on the dynamic change characteristics of the aerodynamic thermal environment, selecting or preparing a composite phase change material; and constructing a pipe wall model for coupling the composite phase change material and the heat pipe. Effective buffering of strong heat flow impact and rapid heat guiding-out are achieved, the temperature and thermal stress of the inner wall of the pipeline are reduced, and the thermal safety and operation reliability of a system are improved.

The invention relates to an arc wind tunnel conduit test wall surface shear force determination method, and belongs to the field of aircraft ground aerodynamic heat test research, the conduit test wall surface shear force determination method provided by the invention can rapidly and accurately obtain the wall surface shear force of an arc wind rectangular conduit test based on actually measured heat flow and pressure; according to the correction method based on development, rapid evaluation of the arc wind tunnel rectangular conduit test wall surface shear force can be realized, and shear force determination of a typical aerodynamic heating environment can be realized within seconds; based on heat flow and pressure measurement, a developed correction method is combined, the wall shear force of the arc wind tunnel rectangular conduit test can be evaluated more accurately, and the error is increased by at least one time compared with that of a conventional engineering method.

The invention discloses a fractal structure function integrated thermal protection system containing a composite phase change material, which comprises a fractal structure web fixedly arranged on the inner side of an outer plate, an inner plate fixedly arranged on the lower side of the fractal structure web, and a middle partition plate fixedly arranged on the fractal structure web between the outer plate and the inner plate, a cavity between the middle partition plate and the outer plate is filled with a heat insulation material, and a cavity between the middle partition plate and the inner plate is filled with a composite phase change material; through the nano wetting effect of the modified low-melting-point metal-based alloy and the synergistic effect of a micron-sized framework of the expanded graphite, an alloy-graphite two-stage heat conduction network is constructed, the heat conduction coefficient is greatly increased, heat of the fractal-structure web can be rapidly transferred into the composite phase change material to be stored, the problem of thermal short circuit is avoided, and the fractal-structure web is suitable for large-scale production. The thermal protection performance and the temperature uniformity are improved, and reliable guarantee is provided for safe flight of an aircraft in an environment with serious aerodynamic heating.

A gas heater test bench, a test section is provided with a workpiece to be tested, a gas generator is communicated with the inlet of the test section, an exhaust section is communicated with the outlet of the test section, an igniter is arranged in the gas generator; the outlet of the exhaust section is communicated with a sound attenuation tower through an exhaust pipeline; a medium-pressure gas source is communicated with the gas generator and the igniter through a gas source supply system respectively; a high-pressure circulating water source is communicated with the test section, the gas generator and the exhaust section through a high-pressure water supply system respectively; a low-pressure water spraying source is communicated with the exhaust section through a low-pressure water spraying system; an alcohol source is communicated with the gas generator through an alcohol supply system; a nitrogen gas source and a liquid oxygen source are communicated with the gas generator through a liquid oxygen and nitrogen supply system; a kerosene supply source of the igniter is communicated with the igniter. The gas heater test bench satisfies the working state of the local missile head cover and the air inlet under the simulated flight state of the hypersonic aircraft, and provides a basis for the material selection and structure design of the type under the aerodynamic heating condition.

The application provides a method for assigning aerodynamic heat boundary conditions of variable sweep wing structure heat transfer simulation, which comprises defining a wing surface position with a first sweep angle in a whole trajectory as a reference position, rotating heat environment data to the reference position, and dividing a structure heat transfer grid at the reference position; then, creating a "point method" equation for the outer edge of a wing box, establishing a structure heat transfer grid surface grid point space position criterion based on the equation, and dividing the grid points into a set of aerodynamic heating points and a set of non-aerodynamic heating points; taking the set of aerodynamic heating points as an aerodynamic heat data interpolation target area, performing heat environment interpolation, and assigning 0Kw / m 2 to the cold wall heat flux of the set of non-aerodynamic heating points, restoring the constant value of enthalpy, and forming an adiabatic boundary condition. The method solves the problems that aerodynamic heat environment data cannot be directly interpolated to the structure heat transfer calculation grid surface due to the change of the spatial position of the variable sweep wing structure, and the structure heat transfer grid spatial position changes, resulting in that heat transfer analysis cannot be continuously performed along the whole trajectory.

The application discloses a method for analyzing interference aerodynamic heat of a booster, comprising the following steps: obtaining air flow parameters at different time according to flight trajectory calculation results; calculating shock wave angles of a booster nose at different time according to the air flow parameters at different time; determining a shock wave direct impact area, a first area and a second area according to the calculated shock wave angles of the booster nose; simplifying the first area into a nose stagnation point and performing boundary layer outer edge parameter calculation and cold wall heat flux density calculation; simplifying the second area into a conical surface and performing boundary layer outer edge parameter calculation and cold wall heat flux density calculation; and forming a bundled launch vehicle conical booster interference aerodynamic heating environment according to the boundary layer outer edge parameter calculation results and the cold wall heat flux density calculation results. The application can quickly provide a round of interference aerodynamic heating calculation results in the launch vehicle demonstration, scheme initial stage or multi-trajectory iteration process, so as to carry out missile body heat protection design and evaluate load and strength problems.

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

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

The application discloses a kind of ceramic radome transient thermal stress fast calculation method along flight trajectory, comprising: determining calculation input, including the aerodynamic thermal environment of ceramic radome and the structure scheme to be evaluated;Set the circumferential discrete strategy of ceramic radome, to determine the multiple meridional plane of ceramic radome;Interpolation model is constructed to aerodynamic thermal environment, determines the most severe meridional plane of ceramic radome aerodynamic heating in flight;Build the two-dimensional feature structure model of ceramic radome transient heat transfer and thermal stress analysis, and determine the external contour line of two-dimensional feature structure model;The axial thermal response calculation interval of ceramic radome is constructed and interpolation calculation is carried out, and the aerodynamic thermal environment parameter of each flight time of external contour line is obtained;The heat transfer calculation grid division is carried out to two-dimensional feature structure model, and the structure transient temperature distribution of two-dimensional feature structure model is calculated;Based on the structure transient temperature distribution, calculate structure temperature gradient;Utilize structure temperature gradient to calculate the transient thermal stress distribution of ceramic radome.

The invention relates to the technical field of wind tunnel aerodynamic thermal tests, in particular to an aerodynamic thermal parameter measurement test device and a use method thereof. Comprising a test chamber, a high-frequency induction heater is arranged on one side of the test chamber, a spray pipe is arranged at the air outlet end of the high-frequency induction heater, an outlet of the spray pipe is located in the test chamber, a water-cooling support is arranged in the test chamber, a rotary feeding arm is arranged on the water-cooling support, and a test model is installed on the rotary feeding arm; a rotatable flow blocking cover is arranged at the tail end of the rotary feeding arm, and when the flow blocking cover is in a closed state, the test model is protected from being scoured by high-temperature airflow; the flow blocking cover is located on one side of the test model when in an open state, so that the test model is exposed in high-temperature test airflow. According to the test device provided by the invention, by arranging the rotatable flow blocking cover, the test model is protected from being heated and scoured by high-temperature gas in the initial starting stage of the high-frequency induction heater, the test model is prevented from being scoured and heated by invalid airflow, and the reliability of cold wall heat flux density measurement is improved.

The application provides a heat sealing structure for aerodynamic heating of a separation surface and an aircraft, comprising: a sealing structure, a piston exhaust structure and an internal heat insulation structure; the side of the heat protection layer of the aircraft facing away from the atmosphere outside the aircraft is the interior of the aircraft, one end of a separation device is installed in the heat protection layer of the aircraft, and the other end is installed in the interior of the aircraft; a sealing structure is arranged between the separation device and the heat protection layer of the aircraft, an internal heat insulation structure is arranged between the separation device and the interior of the aircraft, and a piston exhaust structure is arranged at the end of the separation device facing the atmosphere outside the aircraft. The sealing structure, the piston exhaust structure and the internal heat insulation structure can not only avoid the direct inflow of aerodynamic heat from the outside of the separation surface into the interior of the aircraft, but also gradually release the internal pressure to balance the aerodynamic heat, thereby solving the heat flow problem of the aircraft during long-time flight after the fairing is separated.

The invention discloses an aviation aircraft skin heat exchange coefficient estimation method and an aircraft thermal design simulation method. The aviation aircraft skin heat exchange coefficient estimation method comprises the steps of obtaining a working condition calculation instruction and a cabin body characteristic length of an aircraft; if the working condition is the flight working condition, calculating ambient atmospheric parameters based on an atmospheric thermodynamic model, and judging the flow state of the skin surface according to a Reynolds number; selecting a temperature recovery factor according to the flow state, introducing a gas adiabatic index, and calculating a stagnation temperature considering the aerodynamic heating effect; introducing a reference temperature to correct the air physical property of the boundary layer; based on the flow state, selecting a Nusselt number correlation formula to calculate a convective heat transfer coefficient; and if the working condition is ground, calculating based on ground wind speed. Through the parameterization algorithm integrating the atmospheric physical model and the heat transfer theory correlation, decoupling calculation of flow and heat transfer can be realized without constructing an external fluid grid, and the calculation cost is reduced while the estimation precision of the high-speed flight working condition is ensured.

The invention discloses an aerodynamic thermal structure response prediction method based on Delaunay triangulation interpolation, relates to the technical field of aerospace, and is used for solving the problems that a traditional coupling method is difficult in calculation and a decoupling method is insufficient in precision. The method comprises the following steps: firstly, acquiring aerodynamic heat flow and pressure data of the surface of an object at different wall temperatures through parameterization calculation; further utilizing Delaunay triangulation interpolation to establish a continuous function relationship of heat flow and pressure with respect to wall surface temperature for each surface node of the structural finite element grid; and finally, the function relationship is used as an accurate boundary condition to drive finite element calculation, so that the thermal deformation, the thermal stress and the natural vibration frequency are efficiently and accurately predicted. According to the method, effective decoupling of the flow field and the structure field is realized through the intelligent agent model, and the calculation efficiency and the engineering practicability are greatly improved while the precision of the close coupling method is realized.

The application discloses a kind of pneumatic pressure and pneumatic hot environment simulation test device and test method thereof, high-temperature high-pressure gas is generated by fuel and air premixed combustion, after gas compression nozzle, release into test chamber, the test projectile of the position posture of test chamber center is pre-adjusted, erosion is carried out, and the environment of aerodynamic pressure field and aerodynamic heat temperature field is formed on the surface of test projectile, while heating coil is arranged around test projectile, and local heating is carried out to realize the temperature gradient distribution around projectile, test chamber is provided with observation window, and the temperature field distribution of projectile is obtained by infrared camera, and multiple columns of patch type pressure sensors are distributed on the surface of test projectile along axial direction, and the aerodynamic pressure suffered by each part of test projectile is collected.The application can simultaneously load the aerodynamic pressure and pneumatic hot environment of test projectile in laboratory environment, and the dynamic test range of aerodynamic pressure and pneumatic hot environment is wide, and the controllability of test process is strong.

This invention discloses a method for thermo-aerodynamic servo-elastic coupling analysis of an aircraft. Using the time-step size of unsteady aerodynamic forces as the time increment, a reduced-order aerodynamic-thermal model is employed at the current time step to calculate the temperature field of the complete aircraft structure at each discrete moment, obtaining the aircraft's thermal modes. The reduced-order unsteady aerodynamic model is then used to solve for the aerodynamic forces, calculating the aeroelastic response of the aircraft's thermal modes after aerodynamic heating. This aeroelastic response information is then transferred to the servo control system's servo dynamics model, outputting the control force for the next time step. Subsequently, the aero-servo-elastic response under the simultaneous action of aerodynamic forces and control forces at the next time step is solved, until the thermo-aerodynamic-servo-elastic calculation of the entire flight trajectory is completed. This invention considers the influence of friction and clearance in the aircraft's servo control system on the aeroelastic calculation, employing a reduced-order model to calculate aerodynamic forces and aerodynamic heat, thus improving computational efficiency while ensuring solution accuracy.

The application discloses a stealthy ultrahigh-temperature thermal protection coating as well as a preparation method and application thereof. The ultrahigh-temperature thermal protection coating is composed of a ceramicizable polymer precursor and an ultrahigh-temperature ceramic powder. The ceramicizable polymer precursor is prepared by reacting at least one metal element complex with a silicon-based polymer. During service, the polymer precursor is converted into an ultrahigh-temperature nano composite ceramic composed of transition metal carbon / nitride and Si(C / N) ceramic phases. The ultrahigh-temperature ceramic powder is composed of transition metal carbide, boride or nitride and is uniformly distributed in the polymer precursor. The coating can be directly coated on the outer surface of a high-speed aircraft structure. In the initial stage of flight, the coating can exhibit excellent electromagnetic wave absorption performance. In the high-speed flight process, the coating can be converted into an ultrahigh-temperature (nano) composite ceramic coating in situ through aerodynamic heating, thereby forming effective thermal protection for the aircraft and exhibiting the functions of stealth and thermal protection integration.

The application relates to the technical field of numerical simulation of the intersection of computational fluid dynamics and computational heat transfer, and discloses a temperature field prediction method based on multi-model cooperation, which comprises the following steps: training a neural network using an aerodynamic heat data set to obtain an aerodynamic heat rapid prediction model; selecting multiple different flight trajectories within the flight envelope of an aircraft and discretizing the flight trajectories into flight operating points; using the aerodynamic heat rapid prediction model to obtain the cold-wall heat flux of all flight operating points of the flight trajectories under a preset wall surface temperature; using the cold-wall heat flux as a boundary condition to solve a transient heat conduction equation of a thermal protection structure, obtaining the temperature field change of the thermal protection structure along the flight trajectory under aerodynamic heating conditions, and forming a structure temperature field data set; training a neural network using the structure temperature field data set to obtain a structure heat transfer prediction model; and using the structure heat transfer prediction model to predict the temperature field. The application can realize the coupled rapid analysis of aerodynamic heat and wall surface structure heat transfer under any operating condition within the full flight envelope.

The present invention discloses an aircraft air inlet lip and a method for cooling and insulating the air inlet lip. The air inlet lip is mainly composed of a mounting plate, an inlet pipe, a lip structure, a lip front end, and an outlet pipe. The leading edge of the aircraft air inlet lip forms a high-temperature surface due to aerodynamic heating. Aircraft fuel is used as a cooling medium and flows into the upper and lower independent flow channels in the heat dissipation core through the inlet pipe, exchanges heat with the lip structure, and finally flows out through the outlet pipe, taking away the heat, so that the temperature of the lip structure is reduced to within the temperature range allowed by the material. The present invention reduces the temperature of the lip structure by adopting a functional integration design concept. The lip heat insulation and cooling structure, form, and flow channel layout of the present invention are reasonable. Aircraft fuel is used as a cooling medium to actively reduce the temperature of the lip structure and improve the surface temperature distribution of the lip structure.

Athermal gradient management assembly comprising a functional component. The functional component includes an exterior facing segment, where the exterior facing segment is configured to experience aerodynamic heating, and an interior segment with the exterior facing segment, where the interior segment is configured for isolation from the aerodynamic heating. The functional component includes a thermal gradient extending between the exterior facing segment and the interior segment. The thermal gradient management assembly also comprises a gradient mitigating heating system includes a power source and a gradient heating element in communication with the power source, wherein the gradient heating element is coupled with the interior segment. The gradient heating element is configured to diminish the thermal gradient between the exterior facing and the interior segments.