301 results about "Thermal test" patented technology

A power measuring apparatus including a voltage transformer, a voltage-current measuring circuit, and an operation unit is provided. The voltage transformer converts an input voltage signal into an output voltage signal. The voltage-current measuring circuit includes an inductor, a first resistor, a first capacitor and a second capacitor. A first end of the first inductor herein receives the output voltage signal and a second end thereof generates a voltage feedback signal. The operation unit performs an operation on the voltage feedback signal and a terminal current signal between the first terminal and the second terminal of the first capacitor so as to calculate a power signal. In this way, the accuracy of thermal test is effectively increased by the presented invention.

The invention relates to a ground zero-gravity test device for a small-space one-dimension extension mechanism for a spacecraft to overcome the defects of various test devices in the prior art and to solve the problem that the small-space one-dimension extension mechanism simulates the zero-gravity state of in-orbit operation in the ground gravity environment and to further solve the problem about compatibility of the small-space one-dimension extension mechanism with the comprehensive operation of the vacuum test and the thermal test. The ground zero-gravity test device adopts the principle of square teeth meshing and combines the real-time synchronism of a synchronous belt to vertically hang a tested mechanism. The ground zero-gravity test device is simple and feasible in design, small in size and light in weight, convenient to move, and strong in adaptability to the environment, can effectively operate in a vacuum or thermal environment condition, and can be used for testing the expanding characteristic of one-dimension extension mechanisms like a coiled extension arm and a thin-wall tubular extension arm.

The invention discloses a fault on-line repairing method for satellite power supply and distribution test system, which can fast repair the fault of power supply and distribution test equipments on line in thermal tests, aging tests and other tests in which the satellite needs continuous powering up. The internal components of the power supply and distribution test equipments are encapsulated in the plug-in or guide-way form, and simultaneously modular structure designing method is adopted to lead a magnetic latching relay can be fast replaced when in fault; the digital quantity output and input adopt dual redundant configuration, and the analog quantity output signals are displayed by a measurement display unit; the solar cell array simulator output is directly loaded on the satellite via independent interface; a flow backward resisting diode is connected in series between a voltage stabilization source power supply positive line, and an analog quantity outgoing line is connected with a protective resistor in series to prevent the on-satellite products from being influenced by cable plugging when the satellite is powered up; on the basis of satellite-ground synergic design, the redundant resources can be started by fast replacing the faulted modules one faults occur to the equipments to realize equipment on-line repairing.

The invention relates to an infrared radiation heat flow density reinforcement device for a high temperature pneumatic thermal simulating test of a missile, comprising a quartz lamp heating array, a blower, a variable frequency speed controller, a quartz glass airflow isolation board, a high temperature radiation test piece, a water cooling box, a heat flux sensor, a temperature sensor and a computer. A one-way airflow guiding isolation channel is formed by the quartz glass airflow isolation board, and flow guiding blades are driven to rotate at high speed by a motor of the blower to form a one-way strong airflow; and in the high temperature transient pneumatic thermal test of simulating the hypersonic speed flying of the missile, the strong airflow flows the surface of the quartz lamp heating array in the isolation channel such that the temperature is reduced below the softening temperature of the quartz glass. Because the heat flow density received by the surface of the high temperature radiation test piece of the missile in a heat intensity experiment exceeds the limit of 1 Mw / m<2> when the wind cooling is not carried out, the transient heat flow density can reach 2 Mw / m<2>. The invention provides a more efficient dynamic thermal radiation high temperature testing means for developing the hypersonic speed missile with faster flying speed.

The invention discloses an ultrahigh-temperature thermal current simulation system used for a spacecraft vacuum thermal test in a vacuum container. The ultrahigh-temperature thermal current simulation system mainly comprises a semi-cylindrical graphite heating array, high-reflectivity high-temperature thermal insulation assembly units, a graphite heating array installing support, an ultrahigh temperature measuring unit and a power-regulator power control unit. A plurality of graphite heating rods are arranged and combined into the semi-cylindrical graphite heating array according to the requirements for the density and uniformity of thermal currents. Each graphite heating rod is composed of a graphite heating body and a graphite electrode, alternating-current electricity supply is adopted, and wiring electrode leads penetrate through a high-current cabin-penetrating electricity supply flange arranged on a container wall of the vacuum low-temperature environment simulation container and are electrically connected with the temperature measuring unit and the power-regulator power control unit respectively. Through the unique design of the graphite heating array, the high-reflectivity high-temperature thermal insulation assembly units and the ultrahigh temperature measuring unit, measurement of and control over temperatures over 1800 DEG C in the space environment simulation container are achieved, and the ultrahigh-temperature thermal current simulation system is also suitable for high-temperature and high-thermal-current environment simulation requirements in vacuum thermal tests of all types of spacecraft.

The invention relates to a high-temperature radiation temperature enhancement device of a quartz lamp used for a guided missile thermal test. The high-temperature radiation temperature enhancement device comprises an interlayer water-cooled quartz lamp heating array, a water-cooled type cooper electrode base, a water flow channel of a lamp wall, a tungsten thermal fuse, a quartz lamp electrode, cooling water, a through water-cooled channel and a metal connecting pipe. A wall surface of the quartz lamp used for heating is designed into an interlayer form, part of the heat quantity on the surface of the quartz lamp can be removed by the cooling water flowing through the water flow channel of an interlayer lamp wall, so that the surface temperature of quartz glass is dropped to below softening temperature 1600 DEG C, an infrared light generated by the tungsten thermal fuse of the quartz lamp can penetrate through a wall surface of a heater of a transparent interlayer water-cooled quartz lamp and the transparent cooling water to radiate outwards. In the high-temperature radiation temperature enhancement device, a radiating capacity of the quartz lamp heating array exceeds the primary use limit temperature 1200 DEG C, and the maximum radiation temperature is increased to 1450 DEG C. Therefore, an effective high-temperature pneumatic thermal test analog method is provided for researching a hypersonic missile with faster flying speed, and the high-temperature radiation temperature enhancement device has an important application value in military engineering.

The invention relates to the technical field of spacecraft environmental simulation, and discloses an infrared lamp array heat-flow density calibration device. The infrared lamp array heat-flow density calibration device is characterized in that the material and shape of a set simulation piece is in accordance with a product to be calibrated, the outer side of the simulation piece is provided with a heating piece, a thermal control coating is sprayed on the outer side of the heating piece, and a thermal-protective coating is arranged on the inner side of the simulation piece; and a thermocouple is arranged between the simulation piece and the thermal-protective coating, and an infrared lamp array is arranged near the outer side of the simulation piece. The invention also discloses an infrared lamp array heat-flow density calibration method, wherein the simulation piece is respectively heated by energizing the infrared lamp array and by utilizing the heating piece, and the temperature fields in the two heating methods are compared; when the two temperature fields are the same, the heating power of the heating piece and the energizing current of the infrared lamp array are calibrated. The infrared lamp array heat-flow density calibration device provided by the invention solves the problem that the precision cannot be tested when the external heat flow is simulated by the infrared lamp array in a complicated surface spacecraft vacuum thermal test, thereby improving the stability of the test results; and through the calibration of the lamp array, the consumption of the heat flow sensor is lowered, and the test cost is lowered.

This invention discloses one thermal resistant network module and discloses one method to computer micro electron part by use of thermal resistance network, which comprises the following steps: according to the resistance network module to get unknown parameters temperature expression; processing thermal test to get micro electron part relative temperature data; extracting micro electron part parameters; establishing electron part limit element artificial module; correcting the limit module to ensure limit module accuracy to get effect module for limit artificial data; establishing optimization function to process optimization process to determine the unknown parameters value to get each boundary condition temperature.

The invention relates to the field of spacecraft ground thermal tests and discloses a spacecraft thermal vacuum testing temperature control method. The method comprises the steps of measuring the position of a laser heating center with two thermodetectors with different measurement ranges at the same time to obtain average temperature T1 and average temperature T2 within two measurement ranges; taking the specific value T1 / T2 of the two average temperatures as the setting parameter of a proportion and integral time coefficient in a PID controller, conducting real-time online setting on the proportion and integral time coefficient according to the change of T1 / T2, and obtaining the change value of current heating power through the PID controller by means of the set coefficient; adjusting laser output power according to the change value. By the adoption of the method, under the condition that the overshoot of controlled temperature is reduced greatly, controlled temperature adjusting time is not increased, and the robustness and stability of a spacecraft thermal testing temperature control system are improved.

The invention discloses a method used for simulating the thermal resistance value of a thermal test chip, evaluating the correlation between the thermal resistance value of a heating block and the thermal resistance value of the thermal test chip and finding the size of the heating block that fits for heat generating situations of actual chips. Furthermore, after implementing the radiating performance test of the heating block, the method also can improve the reliability of test results by validating that whether the relation between a transient response value of the thermal resistance of the heating block and a steady response value of the thermal resistance of the thermal test chip lies in a preset variation range.

The invention provides a wave-absorbing temperature control type external heat flow simulating device under a thermal vacuum environment. The wave-absorbing temperature control type external heat flow simulating device comprises a structural silicon carbide wave-absorbing component 1, a temperature control system 2, a wave-absorbing box 3 and a fixing mounting plate 4. The fixing mounting plate 4 is connected to the inner surface of the wave-absorbing box 3. The temperature control system 2 is disposed between the wave-absorbing box 3 and the fixing mounting plate 4 and connected with the fixing mounting plate 4. The structural silicon carbide wave-absorbing component 1 is connected to the inner surface of the fixing mounting plate 4. The wave-absorbing temperature control type external heat flow simulating device under the thermal vacuum environment has the advantages that the device integrates a thermal heat flow simulating function and a wave-absorbing function, microwaves can be prevented from reflecting back to the surface of a product effectively, vacuum thermal test simulation can be performed under a real on-track thermal coupling state, and test reliability is increased.

The invention discloses an atomic force microscope scanning thermal probe and a preparation method thereof. The obtained atomic force microscope scanning thermal probe comprises a probe cantilever, a probe tip, a graphene thin film layer and a low thermal conductivity layer, and the thermal conductivity of the low thermal conductivity layer ranges from 0.2 W / mK to 2 W / mK; the probe tip is located at one end of the probe cantilever, and the probe tip is coated with the graphene thin film layer which is coated with the low thermal conductivity layer, only the portion, corresponding to a probe tip body, of the graphene thin film layer is coated with the low thermal conductivity layer, and the portion, corresponding to the point of the probe tip, of the graphene thin film layer is not coated with the low thermal conductivity layer. By means of the atomic force microscope scanning thermal probe and the preparation method thereof, the accuracy and the resolution ratio of atomic force microscope thermal tests can be improved.

A method of thermally testing for whether an integrated circuit die is attached to a die pad, is provided. Heat is applied from an external heat source to a first side of the die pad. The die is attached to a second side of the die pad. The second side of the die pad is opposite the first side of the die pad. A temperature of the die is measured at a first die location on the die. The die pad is located between the external heat source and the first die location. This method may be performed using a thermal test apparatus having a socket and the external heat source. The socket is adapted to receive and retain one or more leads of a die package. The apparatus may further include a mechanism adapted to move the heat source toward the first side of the die pad.

The invention relates to a test bracket which is used in a satellite vacuum thermal test. The test bracket is an L-shaped framework structure, and the aim of adopting the framework structure is to reduce the barrier to the satellite body by the test bracket as much as possible. The bracket is formed by welding 1Cr18Ni9Ti rectangular steel tubes and can bear 1500kg of satellite weight. The A surface and the B surface of the bracket can both touch the ground, when the satellite is arranged in the bracket, the B surface touches the ground, and when the test is done in a space simulator, the A surface touches the ground. An installing flange in the B surface of the test bracket is connected with a satellite coupling flange by a coupling frame which adopts an inverted cone stainless steel truss structure in order to reduce the barrier to the heat dispersing surface of the satellite body, and moreover, the test bracket has the function of regulating the levelness of the satellite. The bracket increases the whole strength and stiffness of the bracket and can avoid violent sloshing of the satellite in the process of lifting and switching. In the test process, adopting the bracket can reduce the influence to the radiation on the surface of the satellite as much as possible and ensure the levelness index of the satellite within 1mm / 1m.

The invention discloses an ordinary-pressure thermal test system of a manned spacecraft, and a method for performing ordinary-pressure thermal test by the system. The system comprises an inner insulating sealing chamber for setting the manned spacecraft, and a vacuum unit and a vacuum tank which are arranged outside, wherein the vacuum tank is used for providing a vacuum environment for inner part of the manned spacecraft; a control valve for controlling the vacuum unit to operate is arranged between the vacuum tank and a sealed cabin of the spacecraft; a refrigeration dehumidifier unit and a refrigerating unit are arranged outside the insulating sealing chamber; the refrigerating unit is connected with an immediate heat exchanger in the to-be-tested manned spacecraft by a liquid loop; and the liquid loop exchanges heat with the loop in the spacecraft in the insulating sealing chamber. The invention provides a new method in the ground thermal test of the large manned spacecraft, which can check performance of the whole space station assembly, has high safety and flexibility, reduces test expenses and improves thermal test level of the spacecraft in China.

A soft-sided thermal test chamber that is readily mountable upon / over a shaker test table. The test chamber includes four interconnected insulated side walls and a top which define an enclosure. A two-piece bottom receives the side walls to fully enclose the test object upon the table. The two-piece bottom includes an interchangeable center piece which is sized to accommodate various tables or test objects. The side and top pieces are hung from a frame that is readily lifted using an overhead crane or chain fall hoist to allow easy access to the table during set up. The bottom and side walls are interconnected via hook and loop fasteners, which allows for rapid alteration of the set up of the test equipment while maintaining a thermally sealed chamber.

The invention discloses a testing system for temperature gradients of a large spacecraft structure under atmospheric pressure. The system comprises an atmospheric pressure temperature circulating device provided with a temperature gradient control system, a temperature measuring system, a mechanical property measuring system, a gravity unloading system and the like. The atmospheric pressure temperature circulating device is used for providing stable background temperature. The temperature gradient control system is used for locally heating a test specimen based on a PID algorithm to generate set temperature gradients. The temperature measuring system comprises a thermocouple and an infrared thermography non-contact temperature measuring system. The mechanical property measuring system comprises a strain gage and a non-contact deformation measurement system. A simulation system is capable of simulating temperature gradients of the large spacecraft structure on an orbit and measuring force and thermal properties corresponding to the structure on the environment of ground atmospheric pressure. The testing system for temperature gradients of the large spacecraft structure under atmospheric pressure has following beneficial effects: a new method is provided for simulation of temperature gradients of the ground; an atomospheric pressure device is used for checking the large spacecraft structure; safety and flexibly are high; test cost is reduced; and level and efficiency of a thermal test are improved.

The invention discloses an analysis and test system and a test method for the junction temperature of a semiconductor lamp. The system comprises a radiation power tester, an electrical parameter generating and measuring apparatus, a temperature tester, a variable-environment lamp test integrating sphere, a test reference point and light source module temperature distribution computing module, a light source module thermal resistance distribution and junction temperature computing module, a multi-working point optimization and analysis module and a central monitoring and processing computer. The test system provided by the invention can effectively test the junction temperature distribution of the lamp under conditions of normal working state of the semiconductor lamp and no damage to the structure of the lamp. The system is formed based on a concept that: the whole semiconductor lamp is used as an object of study, a proper part of a light module is used as a test reference point, and the junction temperature of the semiconductor lamp is analyzed by a method combining physical test, a numerical model, a physical model and computer thermal test.

The invention provides an inner-wall non-sectional type high-temperature thermal test device of a large high-speed aircraft round-shell structure. The device comprises an L-shaped electrode support, an annular flat copper bracket, a quartz burner, an electrode, a metal pressing piece, a water-cooling copper tube, a large high-speed aircraft round-shell structure, an annular ceramic locating rack and a round locating hole. The device can achieve a non-sectional type high-temperature thermal test for a 2-m large high-speed aircraft inner cabin structure, prevents wave-shaped peak-valley variation of a large elastomer temperature field caused by a sectional splicing heating mode, improves the temperature uniformity, reality and safe reliability of the large elastomer temperature test, and provides an important means for researching high-temperature characteristic and heat-insulation performance of the large high-speed aircraft inner cabin structure.

The invention relates to an infrared lamp array heating system in the infrared lamp array heating application technical field. The infrared lamp array heating system comprises a test product and a tool, wherein the test product is fixed by the tool; an infrared lamp array is arranged out ofoutside the test product; one side of the infrared lamp array, which is back on the test product, is provided with a thermal-insulation lampshade; and the reflection layer of the thermal-insulation lampshade is used for transferring diffusion heat to the test product. According to the equipment disclosed by the invention, the problem of overlong temperature rise and reduction of a vacuum thermal test of a spacecraft assembly is solved, a product with a low absorptivity can effectively obtain an anticipative heating rate, and the pollution problem caused when the a spacecraft is heated with a traditional method is solved. Meanwhile, the thermal-insulation lampshade is arranged between the infrared lamp and the a frame to avoid thermal deformation due to the fact that the frame is heated, and the test precision is improved. Further, the test reliability is improved, and the test cost is greatly lowered.

The invention discloses a thermal test solution for an engine, which consists of the following components in weight percent: 70-80 percent of ethylene glycol, 0.1-5.0 percent of aliphatic monocarbocylic, 1.0-5.0 percent of straight-chain aliphatic dicarboxylic acid, 0.1-1.0 percent of triazine skeleton tricarboxylic acid, 0.1-4.5 percent of aromatic acid, 0.1-1.0 percent of composite inhibitor, 0.1-0.5 percent of hydrolytic polymaleic anhydride, 0.2-4.0 percent of alkenyl amines, 0.2-2.0 percent of cyclohexylamine carbonate, 0.001-0.2 percent of antifoaming agent and the balance of deionized water. The invention can inhibit and prevent the corrosion during the thermal test process of the engine, has a unique gaseous phase corrosion inhibition technology, can be used for preventing the corrosion generated by potential corrosion which is formed by the contacting of the residual liquid and the air during the placing process after the thermal test of the engine, and has both a fully organic acid type engine coolant corrosion inhibition technology and a gaseous phase corrosion inhibition technology as well as good stability.

The invention discloses a Mars surface thermal environment simulation system. The Mars surface thermal environment simulation system includes a low pressure wind speed simulation system, a low pressure wind speed measurement system, a gas temperature control system, a gas pressure control system and a Mars rover attitude control system, wherein the low pressure wind speed simulation system utilizes a vacuum container to perform modification, adds an air channel, a fan, a flow deflector, a rectification net and other devices in the vacuum container, and employs the mode of direct current gas absorption to realize 0-20m / s uniform wind speed simulation. The Mars surface thermal environment simulation system completes low pressure wind thermal balance test of the Mars rover, completes the thermal design verification in the special environment, and can provide thermal test environment conditions for a detector or load of the Mars environment to perform the related test verification.

The invention discloses a temperature control method of a wave-absorbing external heat flow simulation system for a large-scale spacecraft microwave antenna vacuum thermal test, surface temperature of a wedge inside a wave-absorbing box is used as a controlled object, and a thin film heater on the outer surface of the wave-absorbing box serves as a heater, thereby realizing temperature control of an antenna arranged inside the wave-absorbing box. According to the method, on the basis of a multistage PID algorithm, a stable-state and dynamic model of wave-absorbing material surface temperature-antenna surface arrived heat flow is established to set ratio parameters, integral parameters and differential parameters at different temperatures; algorithm target curves are given in a segmented manner; a fuzzy control algorithm is adopted to improve the balancing speed of the controlled object; and a neural network model is introduced to predict the test piece surface temperature. The temperature control method of the wave-absorbing external heat flow simulation system for the large-scale spacecraft microwave antenna vacuum thermal test realizes uniformity and high-efficiency and high-precision control of antenna temperature in a large-scale spacecraft microwave antenna thermal test, is suitable for a large-scale microwave antenna vacuum thermal test, improves test effectiveness and coverage of the test, and is also suitable for a vacuum thermal test of a spacecraft equipped with a large-scale microwave antenna.

The invention discloses a data collection system for a vacuum low-temperature environment. The data collection system comprises a vacuum container, a rotary table, a conductive sliding ring, a test piece, a temperature measurement sensor, a test piece measurement cable, a data collector, a rotary table movable disc, a comprehensive controller and an upper computer. The rotary table and the temperature measurement sensor are located inside the vacuum container, the comprehensive controller and the upper computer are located outside the vacuum container, the test piece is placed on the rotary table movable disc, the temperature measurement sensor adhering to the surface of the test piece is connected to the data collector through the test piece measurement cable, the data collector sends data to the comprehensive controller in an RS485 bus mode through the conductive sliding ring, and the comprehensive controller is in communication with the upper computer in a LAN mode. The data collection system is used for measuring the temperature of a rotary spacecraft test piece, and the gap in measurement for the temperature of a vacuum thermal test rotary test piece of a domestic spacecraft is filled.

The invention discloses a test method for thermophysical parameters based on a multilayer composite material, wherein the method comprises the following steps: heating a multilayer composite materialtested object by the system, and measuring the time-temperature curve T (t) of the tested object; calculating to obtain thermal resistance and interface thermal resistance of various layers of materials in the vertical direction of the multilayer composite material tested object by the system; carrying out analogue simulation by the system, to obtain heat transfer disturbance between the layers ofmaterials; and measuring planar thermophysical parameters via an Angstrom method by the system, and converting into related thermophysical parameters. A test system for the thermophysical parametersbased on the multilayer composite materials includes a time-temperature curve obtaining module, a thermal resistance conversion module, a system analogue simulation and a thermophysical parameter calculation module. The method can measure the vertical-direction thermal resistance and the horizontal-direction thermal diffusion coefficient of the multilayer composite material, and simulates the thermal crosstalk direction of heat flow among multiple layers of materials. The method can be widely applied in the field of material heat testing.

The invention discloses an identification method for lumped thermal parameters of a lithium ion battery. The identification method for the lumped thermal parameters is based on an lumped thermal equivalent circuit and special signal stimulation and by combination with a mathematical algorithm; according to the method, temperature rise of the battery is simulated based on the lumped thermal equivalent circuit; a thermal time constant in the environment is obtained by a temperature change experiment; a functional relation between impedance and temperature is established according to impedance at different temperatures; and by applying sine alternating current with single frequency and amplitude, relatively high accuracy of a battery thermal model is ensured; according to a battery temperature rise model and the functional relation between impedance and temperature, temperature change of the battery under special stimulation is circulated; and an actually-tested battery temperature is approached through a search iterative algorithm, so as to identify specific heat capacity and heat resistance of the battery. According to the method, complex test equipment, such as a heat insulating calorimeter, an isothermal calorimeter and other expensive thermal test equipment are not needed; and in addition, the method has the advantages of simplicity and reliability in identifying thermal parameters of the lithium ion battery, easy realization of engineering and the like.

The invention belongs to the technical field of a spacecraft thermal test, and relates to a ground normal-pressure equivalent thermal test method for the space radiation heat exchange law. Through the invention, the shortcoming that the simulation of the space radiation heat exchange law can not be realized in the ground normal-pressure environment can be overcome, and the simulation of the cooling law of a space radiator can be realized. In the invention, the basic working principle is that the cooling heat flows are equivalent, namely that the cooling heat flow of a ground cooling system is equal to that of a space cooling system; and under the condition that the ground cooling system and the space cooing system have the same inlet temperature, if the outlet temperatures of the branches of the two cooling systems are kept consistent, the cooling heat flows of the two systems can be deemed to be equivalent according to the principle of energy conservation.

The invention provides a multi-functional experimental device for emergency discharge foundation experimental study in runaway reaction, wherein the device is used for carrying out the overall process of simulation of the runway reaction, providing useful data for the engineering design and model selection for a safe discharge device, and providing basic data for danger estimation response. The simulation experimental device mainly comprises a reaction kettle, a bearing kettle, a feed system, a discharge control system, a temperature control system, a pressure control system, a discharge collecting and processing system, a data collecting and storing system, an automatic operation and control system and a safe auxiliary appliance. The simulation experimental device is inventive in structural design, has great functions, is convenient to test, is time-saving, has high reactant dosages, can be used for implementing adiabatic calorie thermal test, and has less thermal inertia; the tested experimental data can be directly used for engineering application; and the method can be used for overall process simulation of the runaway reaction, and provides useful data for the estimation of reaction danger and the emergency discharge device.