21 results about "Aerodynamic heating" patented technology

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 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.

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 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.

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 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 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.

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.

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.

The application relates to a method for simulating the internal and external heating of a solid engine, which comprises the following steps: using a quartz lamp to heat the inner and outer sides of an engine shell test piece to simulate the internal heat source condition Qn and the external heat source condition Qw of the engine shell; measuring the temperature of the engine shell test piece under the heating of the quartz lamp to form a simulated temperature history curve T, and providing a temperature boundary condition for the design of the engine shell according to the simulated temperature history curve T. The quartz lamp is used to simulate the heating history of the internal and external engine shells, and finally the temperature history of the shell under the combined action of the aerodynamic heating and the internal heating of the engine in the flight process is obtained, the temperature boundary condition is provided for the design of the engine shell, the design quality is improved, the heating weight is reduced, the overall performance of the aircraft is improved, and the range is expanded.

The invention provides a double-shaft variable-sweep upper single-wing layout of a speed-domain-crossing patrolling missile, which comprises a double-shaft variable-sweep wing mechanism arranged on the back of a fuselage and can form an M-shaped integral structure; each double-shaft variable-swept-wing mechanism comprises an inner wing section and an outer wing section, and the wing root of the inner wing section is pivotally connected with the fuselage, so that the sweep-forward angle of the inner wing section is adjusted; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, so that the sweepback angle of the outer wing section is adjusted. In the embodiment, the angles of the inner wing section and the outer wing section can be independently changed, the left side and the right side have asymmetric deformation capacity, and the overall aerodynamic configuration is changed to adapt to different flight speeds from low subsonic velocity to supersonic velocity. The upper single-wing layout is adopted, and space can be released for placement of an air inlet channel and task loads. The high-aspect-ratio wings bend and deform upwards to deviate from the fuselage under the aerodynamic load, and the problem of structural interference of the wings and the fuselage caused by aerobombs is solved. Most airfoils and the whole variable-sweep rotating mechanism are located in the leeward shadow area of the fuselage, so that the stealth performance is improved, and the aerodynamic heating problem of the deformed airfoils during hypersonic flight is relieved.

The invention relates to the technical field of aerodynamics, in particular to a carbon-carbon heat-proof material surface catalytic ablation coupling aerodynamic heat prediction method, which comprises the following steps: S100, calculating the catalytic reaction generation rate of the carbon-carbon heat-proof material surface according to the temperature, the pressure and the component proportion of the carbon-carbon heat-proof material surface; s200, according to the temperature, the pressure and the component proportion of the surface of the carbon-carbon heat-proof material, the injection mass flow rate and the ablation reaction generation rate of the surface of the carbon-carbon heat-proof material are calculated; s300, according to the catalytic reaction generation rate, the injection mass flow rate and the ablation reaction generation rate, the component mass fraction gradient of the surface of the carbon-carbon heat-proof material is determined; and S400, according to the catalytic reaction generation rate, the ablation reaction generation rate, the injection mass flow rate and the component mass fraction gradient, the heat flow of the surface of the carbon-carbon heat-proof material under the catalytic ablation coupling effect is determined.

The application relates to the technical field of heat flow meters, and discloses a heat flow meter for high-heat-flow-density aerodynamic heating tests and a testing method, wherein the heat flow meter for high-heat-flow-density aerodynamic heating tests comprises a coaxial thermocouple outer electrode, a coaxial thermocouple inner electrode, an insulating adhesive tape and a phase change material wrapping layer; the coaxial inner electrode is in the innermost layer, the coaxial thermocouple outer electrode is sleeved on the outer side, the two are in contact to form a measurement junction through a protrusion on the top of the coaxial thermocouple inner electrode, and the gap between the coaxial thermocouple inner electrode and the coaxial thermocouple outer electrode is filled with the insulating adhesive tape. The heat flow meter for high-heat-flow-density aerodynamic heating tests improves the measurement precision of the heat flow meter through the cold compensation device of the cold end of the coaxial thermocouple by the latent heat of phase change of the phase change material.