86 results about "Aerodynamic heating" patented technology

The invention relates to hydromechanical device, in particular to a wind tunnel used for comprehensive climatic environment tests and aerodynamic tests of railway vehicles. The wind tunnel comprises a tunnel body. The tunnel body is an annular flow channel. A ventilation system used for generating air flow is arranged in the tunnel body. The tunnel body is further provided with an environmental simulation system and an aerodynamic simulation system. The environmental simulation system is used for controlling the weather conditions in the tunnel body. The aerodynamic simulation system is used for enabling the air flow to form a pneumatic flow field. By the adoption of the wind tunnel, the climatic environment tests of the whole real railway vehicle can be carried out, and the aerodynamic tests of a scale model of the railway vehicle can also be carried out. On the basis that the demands for the railway vehicle tests can be met, the wind tunnel can also be used for the related environment tests and the related aerodynamic tests in the fields such as the aerospace field, the automobile field and the construction field.

The invention provides a hypersonic speed guided missile multi-field coupling dynamics integrated simulation analysis method, which comprises the following steps of: establishing a coupled structural motion equation; improving an aerodynamic heating calculation method; improving an unsteady aerodynamics calculation method; establishing an integrated simulation analysis technique process; designing a relationship between an integrated simulation analysis operation model and data; and realizing an integrated simulation analysis system. Compared with the prior art, the aerodynamic heating calculation method and the unsteady aerodynamics calculation method are improved; various techniques of flutter analysis, thermal flutter analysis, pneumatic servo elastic analysis and thermal pneumatic servo elastic analysis are integrated; and an effective integration method for integrally solving and optimally designing a high-speed winged missile all-missile combination body coupling dynamics problem is provided.

The invention relates to a high-speed aircraft lifting surface aerodynamic heating structure multidisciplinary optimization design platform belonging to the technical field of modern high-speed aircraft designs. In the invention, by aiming at the demands of lifting surface heat vibration analysis in an aerodynamic heating environment when a high-speed aircraft lifting surface is designed and considering the positive correlation among aerodynamic heating structures, an aerodynamic heating structure integrated analysis method including aerodynamic heating calculation, transient heat conduction analysis, structural heat model analysis and non-constant aerodynamic force and heat vibration analysis is researched, and the aerodynamic heating structure multidisciplinary optimization design taking the most dangerous heat vibration speed as the constraint and taking the lifting surface structural quality as a target function is researched to achieve the goal of the optimization design on the high-speed aircraft lifting surface. The invention breaks through the problem that the multidisciplinary design optimization is carried out without combining with the influence of three aspects of aerodynamic heating on the structure in the traditional high-speed aircraft design, provides a set of high-speed aircraft lifting surface aerodynamic heating structure optimization design methods and waysand has the advantages of simple scheme and high efficiency.

The invention provides a heat-insulation stealthy high-temperature-resistant gas inlet duct. The gas inlet duct comprises a wave-transparent layer, a heat-insulation stealth layer, an electromagneticshielding layer and a low emissivity coating from inside to outside. The wave-transparent layer is formed by a continuously oxide-enhanced oxide-based composite material; the heat-insulation stealth layer is a multilayer composite material which is formed by alternately arranging heat-insulation core layers and electromagnetic periodic structure wave absorbing layers in a matching optimizing design manner, every heat-insulation core layer is a ceramic fiber mat reinforced aerogel composite heat-insulation core layer, and the electromagnetic periodic structure wave absorbing layer is composed of a high-temperature electromagnetic periodic structure cell pattern; and the electromagnetic shielding layer is formed by carbon fibers, and the low infrared emissivity coating is sprayed on the external surface of the electromagnetic shielding layer. The heat-insulation stealthy high-temperature-resistant gas inlet duct has the advantages of good heat insulation effect, wide wave absorbing frequency band, good low-frequency wave absorbing performance and certain infrared stealth performance, and can be applied to high-speed aircrafts to solve the problems of heat insulation and radar and infrared stealth of aircraft gas inlet ducts under severe aerodynamic heating conditions.

The present airship-shaped space craft has a middle fuselage extending in a fore-and-aft direction, and a pair of two outer fuselages extending in the fore-and-aft direction located symmetrically on both sides of the middle fuselage. In the above fuselages, gas of a specific gravity of which is smaller than that of air is filled, and the middle fuselage is connected with the outer fuselages by a horizontal wing. The horizontal wing is provided with propelling devices supported in a gimbal fashion for generating thrust in any optional direction, jet engines with backwards directed nozzles to be controlled within a range from a slantwise upward direction to a slantwise downward direction, and rocket engines with ejection nozzles to be controlled in right-and-left and up-and-down directions. During ascent of the space craft, at first the propelling devices, then the jet engines and at last the rocket engines are actuated so as to make the space craft reach and fly along a satellite orbit. Upon return of the space craft, the rocket engines, the jet engines and the propelling devices are actuated in reverse order, and aerodynamic heating at reentry into the atmospheric space can be reduced by making the space craft descend slowly. Also, control of the flight in the atmospheric space becomes easy, so that the space craft can safely and easily land on a predetermined narrow area of the ground.

An improved method and device for the reduction of aerodynamic drag and for improved performance of bluff base vehicles by increasing the pressure on the bluff base of the vehicle by controlling the wake flow and the interaction of the wake flow with the vehicle bluff base region. An improved method and device for generating a reduction in drag force on the bluff base of a body moving through a fluid. The apparatus consists of two opposing panels attached to the bluff base and aligned approximately parallel to the side edge of the bluff base and lying in a plane that is parallel to the vehicle centerline. The drag force reduction results from controlling the flow entering the bluff base trailing wake from the left side surface, right side surface, bottom surface, and top surfaces of the vehicle. The objects and advantages also extend to other applications in which an object, body, or vehicle is moving through either a gas or a fluid.

The invention provides a heat flow control method, device and system for an aerothermodynamic experiment of an aircraft. The method comprises the steps that preset flying height information and preset flying speed information of the aircraft are acquired; the surface temperature of an aircraft experiment sample is collected in real time; a target heat flow density value is calculated according to the preset flying height information, the preset flying speed information and the surface temperature; a heat flow density value of the aircraft experiment sample is collected in real time; feedback control is performed according to the heat flow density value and the target heat flow density value, and a heat flow heating device is controlled to heat the aircraft experiment sample to reach the target heat flow density value. The simulation accuracy of an aerodynamic heating environment of the aircraft is effectively improved.

The invention relates to a method for determining temperature field distribution of a cabin of a carrier rocket in a flight phase in an atmospheric layer. The method comprises a first step of determining an aerodynamic heating thermal flow qh changing along with the flight height and acting on the outer wall of the cabin; a second step of determining an average natural convection heat transfer coefficient alpha n of the shell surface of an instrument in a closed cavity of the cabin changing along with the flight height; a third step of determining a forced convection heat transfer coefficient alpha f of interior of the cabin caused by flight accelerated speed and constant leakage of gas in the cabin; and a fourth step of building a cabin node thermal network model, finishing thermal coupling analysis, and obtaining the cabin temperature field distribution. The method comprehensively considers the impact of aerodynamic heating outside the cabin, air natural convection in the cabin and forced convection on the cabin thermal environment, thereby effectively solving the problem of determination of temperature field distribution of the cabin of the carrier rocket in the flight phase in the atmospheric layer.

The invention discloses a hypersonic aerocraft leading edge fluid-heat-solid integrated calculation method, and belongs to the field of aerocraft pneumatic calculation. For solving the problem of complex coupling of structure heat transfer and hypersonic speed flow aerodynamic heating, the method avoids the complex data exchange and calculation amount caused by flow field and structure coupling alternate iteration calculation performed within a time domain through a traditional aerodynamic heating / structure heat transfer coupling solving method, the flow field and the structure are adopted asa physical field, and a uniform control equation is adopted. Physical parameters of a fluid and solid interface are defined again, a finite volume method is performed on a full physical field for spatial dispersion, and hidden time iteration is adopted for time stepping. Compared with a coupling algorithm, the method needs no extra data exchange or coupling strategy, the calculation result and anexperiment value are closer, the calculation amount and the grid dependency are smaller, and the better stability and calculation precision are achieved.

The invention provides an ultra-wide-band wave absorbing structure integrating temperature resistance and mechanical properties and a preparation method of the ultra-wide-band wave absorbing structure. The ultra-wide-band wave absorbing structure comprises an inorganic anti-oxidation layer, a ceramic matrix composite layer I, aerogel layers, periodic structure layers, a ceramic matrix composite layer II, a magnetic wave absorbing material layer and an electromagnetic shielding layer from outside to inside. The electrical loss type wave absorbing structure with heat insulation performance and high-temperature resistance is designed on the outer surface of the ultra-wide-band wave absorbing structure by using periodic structures, effective absorption of high-frequency electromagnetic waves,heat insulation and cooling are realized, a bottom magnetic wave absorbing material is enabled to be at lower operating temperature, the stealth performance of the structure is basically unchanged, and finally, the ultra-wide-band wave absorbing structure integrating the temperature resistance and the mechanical properties is obtained. The problem about radar stealth for high-temperature strong scattering parts such as an inlet of a high-speed flight vehicle under intense aerodynamic heating condition can be solved.

The invention provides a simulation method for an elastic waverider hypersonic flight vehicle. The method includes the following steps that considering influences of aerodynamic heating and variable section inertia moment, an elastic model of a hypersonic flight vehicle free beam structure is established; a model superposition method is used for solving an elastic vibration equation, the intrinsic frequency, damping ratio and intrinsic vibration mode of each mode are solved, and on this basis, an elastic vibration generalized coordinate equation of the waverider flight vehicle is obtained; a computational fluid dynamics method is used for obtaining aerodynamic force and engine propulsion of the flight vehicle; on the basis of analyzing thrust, aerodynamic force and aeroelasticity, a rigid body-elastic coupling model of the hypersonic flight vehicle is established. The simulation method is suitable for performing modeling and simulation on the waverider flight vehicle. On this basis, the established model of the waverider hypersonic flight vehicle is more accurate, and the aerodynamic / propelling / elastic coupling characteristics of the flight vehicle can be reflected better in the process of model simulation.

The invention discloses an aerodynamic heat flow measuring device and method for the plume field of an engine in a vacuum chamber. A support is fixedly mounted on an installation platform of a four-axis movement mechanism, the rear end of a heat flow sensor penetrates the rear side of the support, the heat flow sensor is sleeved with an insulated end cover, and an insulation sleeve is arranged between the heat flow sensor and the support. A signal line of the heat flow sensor is placed in a signal line protective piece, and connected to a signal converter by penetrating the front side surface of the support and then lead to a signal collector outside the vacuum chamber via a chamber-through flange. Voltage signals measured by the heat flow sensor are converted into current signals via the signal converter, and then collected by the signal collector, heat flow values are obtained by conversion according to the coefficient of the hot flow sensor, and thus, the aerodynamic heat flow of the plume field of the engine in the vacuum chamber is obtained. The method and device have the advantages that a three-dimensional movement device is cooperated, total-flow-field measurement is realized, and the problem that interference among sensors influences the flow-field form and further leads to inaccurate measurement is completely solved.

The invention provides an ultra-low orbit satellite with an aerodynamic heating protection function. The ultra-low orbit satellite comprises a satellite body. The satellite body comprises a satellite head, a satellite bottom plate and a satellite side surface structure plate. The satellite side surface structure plate is divided into a satellite conical segment and a satellite columnar segment. The ultra-low orbit satellite further comprises a heat insulation gasket, aerogel, a first heat insulation layer and a second heat insulation layer, wherein the heat insulation gasket is arranged in the satellite body and located between the satellite head and the satellite conical segment; the aerogel is located on the outer surface of the satellite head and the outer surface of the satellite conical segment; the first heat insulation layer is located on the outer surface of the aerogel and the outer surface of the satellite columnar segment; and the second heat insulation layer is located in the satellite body and covers the inner side of the satellite head and the inner side of the satellite side surface structure plate. According to the ultra-low orbit satellite with the aerodynamic heating protection function, the problem that the satellite head and the side edge are high in temperature due to aerodynamic heating when the satellite operates on a low orbit at a high speed is effectively solved, a single machine in the satellite can work within an appropriate temperature range, and the heat protection problem of the satellite operating in an ultra-low orbit region for a long time in the aerospace field is solved.

A method for calculating a three-dimensional temperature field of an ablation heat-proof structure relates to a method for calculating the temperature field of a heat protection structure, in particular to the calculation of the temperature field of an ablation heat-proof structure. The invention overcomes the deficiency of the existing heat-proof structure temperature field analysis method, and provides a three-dimensional temperature field analysis method of the ablation heat-proof structure with high calculation accuracy and simple operation. The present invention is mainly realized through the following technical schemes: (1) calculation of one-dimensional transient heat conduction of the ablation heat-proof layer, (2) calculation of three-dimensional transient heat conduction of the structure, (3) temperature of the ablation heat-proof layer and the structure Field coupling calculations. The invention can complete the analysis of the three-dimensional temperature field of the structure of the hypersonic vehicle equipped with an ablation heat-proof layer under the condition of aerodynamic heating, and the calculation result can be used as the temperature boundary condition of the thermoelastic coupling dynamics analysis. The invention can be used in the heat-proof layer structure design of hypersonic aircraft.

The invention relates to a hypersonic aircraft dredging-type rudderpost heat protection structure. The rudderpost is designed as a hollow structure, the upper surface of the hollow cavity is positioned above the rudderpost mounting gap and is isolated from the control surface through a heat-insulation layer, and the lower surface is positioned inside the aircraft body and is isolated from the control actuator through a phase-change material or heat sink. The side wall of the hollow cavity is provided with a channel which is parallel to the rudderpost bus, a capillary core is welded with the surface of the channel, the hollow cavity is filled with a working medium, and the heat quantity of aerodynamic heating at the rudderpost gap is dredged to the inner wall of the whole hollow cavity through the phase change and dispersion of the working medium in the hollow cavity. The part of the rudderpost inside the aircraft body and the external surface of the control surface part are respectively provided with a heat-insulation layer.

The invention discloses a semi-physical simulation method and a simulation device for the aerodynamic effect in a high-altitude and high-speed environment. The method comprises the following steps: in a high-altitude and high-speed working environment of an airborne camera, pressure and temperature data of the surface of an optical window is calculated through a mathematical method, and the optimal thickness of the optical window is optimized; then, the pressure and temperature data and the optimal thickness of the optical window directly act on an aerodynamic effect physical simulation device, differential pressure and temperature adjustment is carried out directly on the surface of the optical window, and physical simulation experiment is carried out; and finally, feedback to a mathematical simulation model is formed based on the physical experiment result. By adjusting pressure and temperature systems of inside-cabin and outside-cabin environment simulation chambers, a differential pressure environment and a high-temperature environment are established on the surface of the optical window, and an aerodynamic effect caused by high-altitude and high-speed air flowing across the surface of the optical window is simulated. Scientific experiment basis and conditions are created for the ground research of the influence of aerodynamic pressure and aerodynamic heating effect in a high-altitude and high-speed environment on the performance of the optical window.

The invention discloses an aerothermoelasticity calculation method of multi-field coupling of trajectory status of hypersonic aircraft, and belongs to the field of aircraft aerodynamic-calculation. Inthe engineering fields of aerospace and the like, aircraft structures suffer severe aerodynamic heating effects with increasing of flying speed of the supersonic aircraft; and at the same time, flying trajectories are more complex due to more and more complex mission requirements, and aerodynamic heat environments are also more difficult to predict. Therefore, the aerothermoelasticity problem under the complex trajectories deserves attention. According to the method of the invention, a method of combination of inviscid-external-flow solutions and an engineering method is adopted for aerodynamic heating calculation of the high-speed aircraft, compared with a pure numerical method, the method of combination can save massive memory and calculation time, and reduces calculation costs, and compared with a pure engineering method, the same is wider in an application range, can be adapted to relatively complex shapes, and can also provide relatively-accurate aerodynamic heat solutions at thesame time.

The invention provides an aerodynamic heating furnace. An impeller is located in a furnace body. A guide plate is installed in the furnace body through a guide plate support and is fixed to the inner surface of a rear wall plate of the furnace body. Two surfaces, adjacent to the impeller, of the guide plate are spaced. An air inlet hole is formed in the center of the guide plate, and the air inlet hole and the center of the impeller are concentric. A plurality of airflow holes are evenly distributed in the circumference of an annular impeller support. The sum of the sectional areas of the airflow holes is 1-2 times the circumference area of the impeller. A guide edge protruding in the axial direction is arranged on the outer circumference of the guide plate and the guide edge faces towards an inner wall of the furnace body. Turbulence is increased by the right-angled guide edge and heating uniformity and the heating depth are improved by the right-angle guide edge. According to the aerodynamic heating furnace, wind energy is used as a heating medium, the impeller is driven by a motor, so that the wind energy is converted into thermal energy, high-uniformity of a temperature field is achieved through flowing of the wind energy, heating elements such as a heating tube can be omitted, the number of parts used for keeping the uniformity of the temperature field is reduced, the number of partial high temperature points is reduced, the furnace body heat capacity is reduced, the number of furnace body heat bridges is reduced, and therefore the purposes that energy consumption is reduced and the energy-saving effect is achieved are achieved.

The invention provides an aerodynamic heat measuring method suitable for a blunt body reentry vehicle. The aerodynamic heat measuring method is used for measuring of surface heat flow distribution ofthe side walls, the large bottom, the stagnation point area and the boundary layer transition area of the blunt body reentry vehicle and measuring of total heating capacity of the windward side and the leeward side. The aerodynamic heat measuring method comprises the following steps that according to the appearance and a reentry trajectory of the blunt body reentry vehicle, simulation analysis onaerodynamic heat load on the surface of the blunt body reentry vehicle is conducted; according to the sizes and distribution of aerodynamic heating heat flows, sensor selecting and measuring range determining are conducted; according to the aerodynamic heat measuring task requirement of the blunt body reentry vehicle, configuration demand analysis of sensors is conducted; sensor layout design is conducted; and sensor calibration and mounting are conducted. According to the aerodynamic heat measuring method, the real reentry heat flow load can be obtained, particularly, real-time data of transition of the boundary layer on the surface of the vehicle can be provided, and guiding significance is achieved on a current calculating analysis model and the ground test technical scheme.

The invention discloses a high-temperature-resistant air intake passage with both radar stealth and infrared stealth and a preparation method thereof. The high-temperature air intake passage is composed of radar stealth and infrared stealth compatibility zone coating, ceramic coating, wave-absorbing and bearing integrated layer, electromagnetic shielding layer, ceramic coating and low-emissivity infrared stealth coating from inside to outside. The radar stealth and infrared stealth compatibility zone coating has periodic patterns and is made of a precious metal and SiO2 by bonding; the ceramiccoating is composed of a glass layer or a mullite layer; the wave-absorbing and bearing integrated layer is composed of a fiber-reinforced high-temperature-resistant ceramic-based wave-absorbing composite material; and the electromagnetic shielding layer is composed of high-conductivity carbon fiber or silicon carbide fiber; the infrared stealth coating is composed of a low-emissivity noble metalfilm. The high-temperature-resistant air intake passage has high working temperature, strong designability and the advantages of both radar stealth and infrared stealth and can solve the problem of radar stealth and infrared stealth of a hypersonic vehicle air intake passage under severe aerodynamic heating conditions.

The invention relates to the technical field of hot wall correction methods and discloses an improved hot wall correction method considering a hypersonic chemical unbalance effect for aerodynamic heating numerical prediction of a high-speed aircraft. According to the method, convection heat flow is decomposed according to contribution of a physical process, a heat flow part caused by temperature gradient meets a traditional linear hot wall correction method, a heat flow part caused by reaction diffusion is related to wall surface high-temperature chemical action, and corresponding treatment iscarried out according to a wall surface chemical reaction mechanism. Respectively solving the heat flow caused by the temperature gradient and the heat flow caused by the reaction diffusion, and adding the solved heat flows to obtain the total heat flow of the wall surface. According to the method provided by the invention, only the hypersonic flow field under the conditions of the cold wall andthe heat insulation wall needs to be solved, fluid-solid coupling iterative solution is not needed, and the calculation efficiency is high.

The invention discloses a method and system for indicating characteristics of air intake systems under back pressure of hypersonic aircrafts. The method comprises the following steps of: verifying thecorrectness of a hypersonic aircraft air intake system calculation method; simulating a hypersonic aircraft air intake system flow field to obtain air intake system schlieren figure under progressiveincrease of back pressure, judging an appointed range of distance between a shock wave forward pass position corresponding to each air intake schlieren figure and an inlet of an isolated section of the aircraft, and considering that the air intake systems are located in a non-starting boundary; and simulating the hypersonic aircraft air intake system flow field to obtain air intake system upper wall surface heat flow distribution figures under progressive increase of back pressure, judging an upper wall surface heat flow value of the inlet point at the isolated section according to each air intake system upper wall surface heat flow distribution figure, and if the upper wall surface heat flow value is greater than an appointed value, considering that the air intake systems are located inthe non-starting boundary and aerodynamic heating degrades. According to the method and system, back pressure preventing performance of the air intake systems can be pre-judged accurately, and designof the air intake systems is improved.

This method proposes a complex aircraft aerodynamic heating calculation method that can be used in engineering design. This method uses the CFD method to solve the Euler equation for the external inviscid flow field to calculate the complex shape; determine the outer edge parameters of the aircraft boundary layer through CFD; and then in Inside the boundary layer where the viscosity plays a dominant role, the engineering reference enthalpy method is used to calculate the surface heat flow, and various forms of the reference enthalpy method are synthesized. This method not only overcomes the disadvantage that the pure engineering algorithm is difficult to solve the complex shape and surface parameters, but also the calculation amount is much smaller than the pure numerical algorithm, which effectively improves the accuracy and calculation efficiency of the outer edge parameters of the aircraft boundary layer. This method and software use the CFD method to establish a surface flow field database for the complex shape of the aircraft, and can calculate the aerodynamic heating of the aircraft for a long time along the flight trajectory. The calculation efficiency is much higher than the direct numerical solution of the N-S equation to calculate the aerodynamic heat. It can be applied to engineering design.

The invention relates to a preparation method for a sharp leading edge possessing a fiber prefabricated body inner core surface coated with a two-dimension laminated carbon-fiber protection layer. The two-dimension laminated carbon-fiber protection layer is used to coat the prefabricated body inner core, so that local hot-fluid concentration of the prefabricated body inner core caused by fiber cracking and needle dropping because of mechanic processing is eliminated, and the erosion-corrosion resistant performance of the prepared composite material sharp leading edge is improved. Because the fiber prefabricated body inner core surface is coated with the two-dimension laminated carbon-fiber protection layer, the protection layer is tightly attached to the processed prefabricated body inner core, the protection layer does not need mechanical processing, the surface is complete and has no tiny projections and depressions, and by taking the protection layer surface as the radial surface of the sharp leading edge, the influence on aerodynamic heating caused by needle dropping and fiber cracking of the prefabricated body inner core is eliminated, and the prepared composite material sharp leading edge has good erosion-corrosion resistant performance.

The invention discloses a non-probability aerothermoelasticity reliability designing method for a hypersonic control surface. The method is characterized in that uncertainties in aerothermoelasticity designing for the hypersonic control surface are considered, and a sample space for uncertain input parameters is established according to a vertex method; the effects of thermal stress generated by aerodynamic heating on inherent attributes of the control surface are considered, and a bidirectional fluid-structure coupling method is achieved to accurately simulate a dynamic response process of the control surface under the effects of gust loads; on such a basis, reliability indexes are introduced, and reliability analysis of a control surface structure is carried out according to upper and lower bounds of analysis results. Based on the steps, with the control surface structure weight as an optimization objective, and with reliabilities of the maximum displacement of the control surface, the maximum root restraint force and the maximum root restraint moment as restraint conditions, non-probability aeroelasticity reliability optimization designing for the hypersonic control surface can be achieved. The method provided by the invention has the advantages that high reliability of the hypersonic control surface structure is ensured while the structure mass is reduced, so that a new thought is provided for the aerothermoelasticity designing for the hypersonic control surface.

The invention relates to a liquid oxygen temperature zone cold helium pressurization system. In the system, a pressurized gas cylinder 3 is immersed in a liquid oxygen storage tank 1, and can fully exchange heat with liquid oxygen. The b pressurized gas cylinder 3 is arranged, in the liquid oxygen storage tank 1, near the upper part of the storage tank, after a few seconds of flight, the pressurized gas cylinder 3 is exposed from the liquid oxygen, at the end of the flight, the residual gas in the pressurized gas cylinder 3 is smaller and the utilization rate of medium is improved. After the pressurized gas flows out of the cylinder, the enthalpy value of the pressurized gas is increased by making full use of the heat exchange of pipelines, aerodynamic heating and the heat exchange betweenthe gas and fuel in the storage tank. The parallel redundant design of a pressurization solenoid valve can improve the work reliability and fault tolerance ability of the system. The pressurization system avoids the coupling with an engine, omits a heat exchanger of the engine, and saves the development cost, and at the same time, the pressurization system can be self-verified and improves the accuracy of the pressurization design.

The invention provides a material high-temperature mechanical performance test device in an aerodynamic heat-force coupling environment, which belongs to the technical field of the test of material high-temperature mechanical performance. The material high-temperature mechanical performance test device comprises a main machine framework, a mechanical loading system and a clamp water cooling system, wherein the main machine framework adopts a double-stand-column design, the lower end of an oil cylinder mechanism of the mechanical loading system penetrates through an upper crossbeam of the mainmachine framework and is fixed with the upper crossbeam in a thread manner, an upper drawing bar, an upper clamp and the like of the mechanical loading system are connected to the lower end of the oilcylinder mechanism, and other parts of the mechanical loading system are successively connected with a lower crossbeam. The clamp water cooling system is used for cooling a sample clamping mechanismin the mechanical loading system. The material high-temperature mechanical performance test device designs the main machine which can be matched with a coherent jet aerodynamic-heat coupling environment simulation test cabin, has a unique structure design, integrates multiple advanced test technologies, and can well implement the ultrahigh-temperature mechanical performance test of a ceramic material in the coherent jet aerodynamic-heat coupling environment.

An embodiment of the invention discloses a structure temperature analysis method used for a semi-closed region of a hypersonic aircraft, relates to a structure temperature analysis technology used forthe semi-closed region of the hypersonic aircraft, and solves the problem of not taking precision and efficiency into consideration at the same time for an aerodynamic heating / multi-part radiated interference / three-dimensional heat transfer analysis result of the semi-closed complex interference region. The method comprises the steps of firstly generating an aerodynamic heating grid of a concerned part; carrying out aerodynamic heating analysis by utilizing the aerodynamic heating grid; carrying out convective heat flow analysis under different wall temperatures; obtaining a convective heat flow database under a high wall temperature; generating a three-dimensional heat transfer analysis grid considering convective heat exchange under the high wall temperature and multi-part radiated interference; and finally carrying out three-dimensional heat transfer calculation considering the convective heat exchange under the high wall temperature and the multi-part radiated interference by utilizing the three-dimensional heat transfer analysis grid. The method is mainly used for structure temperature analysis of the semi-closed region of the hypersonic aircraft.

The invention discloses an analyzing method for a temperature field of an integral fuel tank, and belongs to the technical field of aircraft designs. The method includes the following steps: step 1, establishing a simulation model of an integral fuel tank structure, and determining a fuel space area inside the integral fuel tank structure; step 2, filling the fuel space area with simulation fuel so that the fuel space area is full of fuel; step 3, further dividing the simulation fuel in the fuel space area into n layers; step 4, applying aerodynamic heating data, changing along with the time, on the aerodynamic surface of the integral fuel tank structure; step 5, carrying out fuel consumption and filling gas simulation through a ladder simulation method; and step 6, obtaining a change result of the temperature of the integral fuel tank structure along with the time after the fuel consumption simulation through the ladder simulation structure. According to the method, temperature field changes of the structure can be effectively obtained without being limited by test equipment or test environments, the product design period is short, and the design efficiency is improved. The analyzing method is easy to promote and apply, and has great practical values.