System and method for measuring temperature of outer surface of spacecraft reentry capsule and computing device

Abstract

The invention provides a system and method for measuring the temperature of the outer surface of a spacecraft reentry capsule and computing equipment, the outer surface of the spacecraft reentry capsule is provided with a heat insulation layer, the system comprises a sensor assembly, the sensor assembly is buried in the heat insulation layer, and the sensor assembly is used for measuring the heat radiation intensity; the calculation unit is arranged in a bottom cabin of the spacecraft reentry capsule and is used for receiving the heat radiation intensity data from the sensor assembly and calculating the temperature of the outer surface of the reentry capsule according to the heat radiation intensity data; and the storage unit is in communication connection with the calculation unit and is used for storing the thermal radiation intensity data of the sensor assembly and the calculation result of the calculation unit. According to the technical scheme of the invention, the temperature of the outer surface of the reentry capsule moving at a high speed can be measured.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace temperature measurement, and particularly relates to a system and method for measuring the temperature of the outer surface of a spacecraft reentry capsule and a computing device. Background Art

[0002] Spacecraft are an important means of transportation for humans to travel back and forth to build space stations in space. The reentry capsule is the section that astronauts take when returning to Earth. Different from the orbital module, the reentry capsule still needs to ensure airtightness under the action of high temperature and high pressure. When the reentry capsule of a spacecraft enters the atmosphere at high speed, due to friction with the atmosphere, the surrounding gas molecules and the outer surface of the reentry capsule are in a state of burning and viscosity, and the surface temperature is not easily dissipated, forming a high-temperature area in this region, and the temperature can reach several thousand degrees Celsius.

[0003] The extreme environment faced by the spacecraft reentry capsule when re-entering the Earth's atmosphere poses extremely high requirements for the design of the thermal protection system and temperature monitoring technology. In this process, it is crucial to study and ensure accurate temperature measurement.

[0004] Therefore, a technical solution is needed to realize the temperature measurement of the outer surface of the high-speed moving reentry capsule. Summary of the Invention

[0005] The present invention aims to provide a system and method for measuring the temperature of the outer surface of a spacecraft reentry capsule and a computing device, to realize the temperature measurement of the outer surface of the high-speed moving reentry capsule and meet the high precision of measurement.

[0006] According to one aspect of the present invention, there is provided a system for measuring the temperature of the outer surface of a spacecraft reentry capsule, the system comprising:

[0007] A sensor assembly, the sensor assembly being buried in the heat insulation layer, the sensor assembly being used for measuring the thermal radiation intensity;

[0008] A computing unit, the computing unit being arranged in the bottom cabin of the spacecraft reentry capsule, for receiving the thermal radiation intensity data from the sensor assembly and calculating the temperature of the outer surface of the reentry capsule according to the thermal radiation intensity data;

[0009] A storage unit, the storage unit being communicatively connected to the computing unit, for storing the thermal radiation intensity data of the sensor assembly and the calculation result of the computing unit.

[0010] According to some embodiments, the sensor assembly comprises:

[0011] A photoelectric converter, the photoelectric converter being used for sensing the thermal radiation intensity and converting the thermal radiation intensity into an electrical signal;

[0012] An analog-to-digital converter that converts the electrical signal of the optoelectronic converter into a digital signal.

[0013] According to some embodiments, the calculation unit calculates the temperature of the outer surface of the return capsule according to the following formula:

[0014] ∮(γ, T) = αγ -5 ﹝exp(β / γT) - 1﹞ -1

[0015] Where ∮(γ, T) is the thermal radiation intensity of the outer surface of the return capsule, γ is the radiation wavelength, T is the temperature value corresponding to the thermal radiation intensity, and α and β are coefficients. The unit of α is watt·centimeter 2 , and the unit of β is centimeter·K, where α and β are pre-calibrated according to the temperature measurement range.

[0016] According to some embodiments, the heat insulation layer includes a first heat insulation layer and a second heat insulation layer, and the sensor assembly includes a first group of sensors and a second group of sensors;

[0017] The first group of sensors is buried in the first heat insulation layer along the circumferential direction at a predetermined interval angle;

[0018] The second group of sensors is buried in the second heat insulation layer along the circumferential direction at a predetermined interval angle, and the second group of sensors is arranged staggeredly with respect to the first group of sensors.

[0019] According to some embodiments, the first group of sensors includes four sensors, and the adjacent sensors are spaced 90° apart;

[0020] The first group of sensors is arranged staggeredly by 45° in the circumferential direction with respect to the second group of sensors.

[0021] According to another aspect of the present invention, there is provided a method for measuring the temperature of the outer surface of a spacecraft return capsule, including:

[0022] Obtaining the thermal radiation intensity, where the thermal radiation intensity comes from a sensor assembly buried in the heat insulation layer on the outer surface of the spacecraft return capsule;

[0023] Calculating the temperature of the outer surface of the return capsule according to the thermal radiation intensity data.

[0024] According to some embodiments, the formula for calculating the temperature of the outer surface of the return capsule is:

[0025] ∮(γ, T) = αγ -5 ﹝exp(β / γT) - 1﹞ -1

[0026] Wherein, ∮(γ, T) is the thermal radiation intensity on the outer surface of the return capsule, γ is the radiation wavelength, T is the temperature value, α and β are coefficients, the unit of α is watt·centimeter 2 , the unit of β is centimeter·K, wherein the radiation wavelength γ is determined according to the temperature measurement range, and α and β are pre-calibrated according to the temperature measurement range.

[0027] According to some embodiments, obtaining the thermal radiation intensity includes:

[0028] Sensing the thermal radiation intensity through the optoelectronic converter in the sensor assembly, and converting the thermal radiation intensity into an electrical signal.

[0029] Converting the electrical signal of the optoelectronic receiving element into a digital signal through the analog-to-digital converter in the sensor assembly.

[0030] According to some embodiments, the sensor assembly is buried in the heat insulation layer on the outer surface of the spacecraft return capsule, and includes:

[0031] The heat insulation layer includes a first heat insulation layer and a second heat insulation layer. The sensor assembly includes a first group of sensors and a second group of sensors. The first group of sensors is buried in the first heat insulation layer along the circumferential direction at a predetermined interval angle;

[0032] The second group of sensors is buried in the second heat insulation layer along the circumferential direction at a predetermined interval angle, and the second group of sensors is staggeredly arranged relative to the first group of sensors.

[0033] According to another aspect of the present invention, there is provided a computing device, including:

[0034] A processor; and

[0035] A memory storing a computer program, which when executed by the processor causes the processor to execute the method described in any one of the above.

[0036] According to the embodiments of the present invention, burying the sensor assembly in the heat insulation layer will not affect the state of the object to be measured. It can not only perform non-destructive detection of the temperature on the outer surface of the return capsule, but also effectively measure the thermal radiation intensity. The calculation unit receives the thermal radiation intensity data from the sensor assembly, and is used to calculate the temperature on the outer surface of the return capsule. The calculation unit is arranged in the bottom cabin of the return capsule and placed away from the high-temperature area, reducing the risk of failure caused by overheating and enhancing the safety of the system. The storage unit is communicatively connected to the calculation unit, receives the data collected by the sensor and the information processed by the calculation unit. The storage unit retains the original data and the processing results, enabling subsequent measurement data to be stored and retrieved.

[0037] According to some embodiments, the first group of sensors and the second group of sensors are respectively buried in the first thermal insulation layer and the second thermal insulation layer, and the second group of sensors is arranged staggeredly with respect to the first group of sensors. By burying the sensor assembly in the thermal insulation layer at a predetermined interval angle in the circumferential direction, it is possible to ensure the omnidirectional monitoring of the temperature of the outer surface of the return capsule. Such a layout helps to capture the temperature distribution of the entire surface and avoid errors in the calculation results caused by unmonitored local high-temperature areas. The second group of sensors is arranged staggeredly with respect to the first group of sensors, forming a redundant system. If a certain group of sensors fails or malfunctions, the other group of sensors can still provide effective temperature data, thereby improving the overall reliability and fault tolerance of the system.

[0038] According to some embodiments, the temperature of the outer surface of the return capsule is calculated by a formula, and the coefficients in the formula are determined by the data calibrated by the black body, which improves the accuracy of temperature measurement and ensures the precision of the temperature measurement system within a specific temperature range.

[0039] It should be understood that the above general description and the following detailed description are only exemplary and do not limit the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for the description of the embodiments.

[0041] Figure 1 The schematic diagram of the system for measuring the temperature of the outer surface of the spacecraft return capsule according to an exemplary embodiment is shown.

[0042] Figure 2A The cross-sectional schematic diagram of the placement of the first group of sensors on the outer surface of the spacecraft return capsule according to an exemplary embodiment is shown.

[0043] Figure 2B The cross-sectional schematic diagram of the placement of the second group of sensors on the outer surface of the spacecraft return capsule according to an exemplary embodiment is shown.

[0044] Figure 3 The flowchart of the method for measuring the temperature of the outer surface of the spacecraft return capsule according to an exemplary embodiment is shown.

[0045] Figure 4 The block diagram of the computing device according to an exemplary embodiment is shown. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0046] Example embodiments will now be described more fully with reference to the accompanying drawings. However, the example embodiments can be implemented in various forms and should not be construed as limited to the embodiments set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concept of the example embodiments to those skilled in the art. Like reference numerals in the figures denote like or similar parts, and thus their repetitive description will be omitted.

[0047] In addition, the described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a thorough understanding of the embodiments of the present invention. However, those skilled in the art will realize that the technical solutions of the present invention may be practiced without one or more of the specific details, or other methods, components, devices, steps, etc. may be employed. In other cases, well-known methods, devices, implementations, or operations are not shown or described in detail to avoid obscuring aspects of the present invention.

[0048] The block diagrams shown in the drawings are merely functional entities and do not necessarily correspond to physically independent entities. That is, these functional entities may be implemented in software form, or in one or more hardware modules or integrated circuits, or in different networks and / or processor devices and / or microcontroller devices.

[0049] The flowcharts shown in the drawings are merely illustrative and do not necessarily include all of the content and operations / steps, nor do they necessarily have to be executed in the order described. For example, some operations / steps may be decomposed, while some operations / steps may be combined or partially combined, so the actual execution order may change according to the actual situation.

[0050] It should be understood that although terms such as first, second, third, etc. may be used herein to describe various components, these components should not be limited by these terms. These terms are used to distinguish one component from another. Thus, the first component discussed below may be referred to as the second component without departing from the teachings of the concept of the present invention. As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0051] The user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data for analysis, stored data, displayed data, etc.) involved in the present invention are all information and data that have been authorized by the user or fully authorized by all parties, and the collection, use, and processing of the relevant data need to comply with the relevant laws, regulations, and standards of the relevant countries and regions, and corresponding operation entrances are provided for the user to choose to authorize or refuse.

[0052] Those skilled in the art can understand that the drawings are only schematic diagrams of the exemplary embodiments, and the modules or processes in the drawings are not necessarily essential for implementing the present invention, so they cannot be used to limit the protection scope of the present invention.

[0053] When the return capsule of a spacecraft enters the atmosphere at high speed, the thickness of the atmosphere is very large. Due to friction with the atmosphere, the surrounding gas molecules and the outer surface of the return capsule are in a state of burning and viscosity, and the surface temperature is not easily dissipated, forming a high-temperature area in this region, and the temperature can reach several thousand degrees Celsius.

[0054] Currently, there are mainly three methods for heat protection and heat dissipation of the return capsule internationally. The first one is mainly to prevent the outer surface of the return capsule from rubbing violently with the atmosphere and burning through the cabin wall. Ablative materials are coated on the surface of the return capsule. These polymer materials include composite materials synthesized from phenolic, quartz glass, asbestos, etc. Heat is dissipated by means of the melting, pyrolysis, evaporation, sublimation, gasification, etc. of the materials when heated at high temperature. The second is radiative heat protection, using composite materials such as ceramics with extremely high emissivity, titanium alloy materials, to radiate and dissipate heat. The third is endothermic heat protection. In some parts of the return capsule, metal materials with a large heat capacity, high melting point, and good thermal conductivity are used to absorb a large amount of aerodynamic heat.

[0055] Understanding the temperature conditions at various altitudes during the process of the return capsule entering the atmosphere, and according to the different characteristics of heat protection in the atmospheric segments at different altitudes of the return capsule, measuring the surface temperature of the return capsule during the reentry section can promote the research on materials such as heat-insulating polymers, ceramics, and titanium alloys. For the thermal stability and chemical stability of high-efficiency heat-insulating materials, high-temperature aerogel materials, and foam carbon materials, etc. under extreme high-temperature conditions; such as the configuration of high-thermal-conductivity carbon-carbon bond composite materials like carbon nanotubes; the thermochemical properties of high-temperature thermal control coating materials, and give targeted solutions to provide reliable data support.

[0056] Currently, most of the infrared temperature measurements at home and abroad adopt the scheme of emissivity correction. The scheme of emissivity correction sets the black body as the maximum emissivity of 1, and the emissivity of actual objects is less than 1. However, due to the influence of factors such as the material, color, and surface roughness of the actual object, the specific value is difficult to determine.

[0057] Therefore, the present invention proposes a system for measuring the outer surface temperature of a spacecraft return capsule to achieve the measurement of the temperature of the outer surface of the high-speed moving return capsule.

[0058] The exemplary embodiments of the present invention will be described below with reference to the drawings.

[0059] When the reentry capsule re - enters the atmosphere under the action of gravity, due to the drastic change of air flow, in order to keep the high - speed flying reentry capsule in a stable attitude, it is made into a bell - shaped with a large bottom and a small top to overcome the disturbance of high - altitude air flow. In order to avoid local overheating of the surface of the reentry capsule, the flight state is adjusted by autorotation and rolling to make the surface evenly heated. The outer surface of the spacecraft reentry capsule has a heat - insulation layer. When the reentry capsule enters the Earth's atmosphere at high speed, due to the intense friction with air molecules, the surface temperature will rise sharply, possibly reaching thousands of degrees Celsius. The main function of the heat - insulation layer is to protect the interior of the reentry capsule from the external high temperature and at the same time minimize the heat conduction to the interior.

[0060] Figure 1 The schematic diagram of the system for measuring the temperature of the outer surface of a spacecraft reentry capsule according to an exemplary embodiment is shown.

[0061] According to an exemplary embodiment, refer to Figure 1 , the system for measuring the temperature of the outer surface of a spacecraft reentry capsule includes a sensor assembly 101, a calculation unit 103, and a storage unit 105. The sensor assembly 101 is buried in the heat - insulation layer. The sensor assembly 101 is used to measure the thermal radiation intensity. The calculation unit 103 is arranged in the bottom cabin of the spacecraft reentry capsule and is used to receive the thermal radiation intensity data from the sensor assembly 101. The storage unit 105 is communicatively connected to the calculation unit 103, and the storage unit 105 is used to store the thermal radiation intensity data of the sensor assembly 101 and the calculation results of the calculation unit 103.

[0062] The sensor assembly 101 consists of a photoelectric converter and an analog - to - digital converter. When the spacecraft enters the atmosphere, due to the high - temperature environment generated by high - speed friction, the photoelectric converter is used to monitor the temperature distribution on the outer surface of the reentry capsule. The high - temperature radiation generated by the friction between the outer surface of the reentry capsule and the atmosphere is directly received by the photoelectric converter, and the thermal radiation intensity is converted into an electrical signal. The analog - to - digital converter processes and converts the electrical signal of the photoelectric converter. The analog quantity component converts the analog signal generated by the photoelectric converter into a more easily processed form, and the digital quantity acquisition component converts the pre - processed electrical signal into a digital signal.

[0063] Both the photoelectric converter and the analog - to - digital converter are calibrated in the vicinity with a black body. The calibration object is the complete circuit board of the photoelectric converter and the supporting analog - to - digital conversion module. Usually, after the photoelectric converter is soldered on the circuit board, before actual application, temperature calibration is carried out in front of the metrological black body. For example, when the temperature of the black body is m (unit: °C), the digital voltage of the circuit board corresponding to it is n (unit: V). The radiation intensity is converted into an electrical signal, and then the electrical signal is matched with the temperature, providing a standard reference point for the photoelectric converter.

[0064] Figure 2AA cross-sectional schematic diagram showing a first set of sensors according to an exemplary embodiment placed on the outer surface of a spacecraft reentry capsule.

[0065] Figure 2B A cross-sectional schematic diagram showing a second set of sensors according to an exemplary embodiment placed on the outer surface of a spacecraft reentry capsule.

[0066] The spacecraft heat insulation layer has a first heat insulation layer 201 and a second heat insulation layer 203. The sensor assembly may include a first set of sensors and a second set of sensors. The first set of sensors is buried in the first heat insulation layer 201 along the circumferential direction at a predetermined interval angle, and the second set of sensors is buried in the second heat insulation layer 203 along the circumferential direction at a predetermined interval angle. The first set of sensors is staggeredly arranged relative to the second set of sensor assemblies.

[0067] Each heat insulation layer of the spacecraft has a certain thickness. For example, each heat insulation layer is 3 cm thick. See Figure 2A , the first set of sensors is buried along the first heat insulation layer 201 of the outer diameter of the reentry capsule. Assuming there are four sensors 101 in the first set of sensors, the interval between adjacent sensors 101 is 90°; see Figure 2B , the second set of sensors is buried in the second heat insulation layer 203. The second set of sensors also has four sensors 101, and the interval between adjacent sensors is 90°. The first set of sensors is staggeredly arranged by 45° relative to the second set of sensors in the circumferential direction.

[0068] The physical model of an ideal black body is described by Planck's formula:

[0069] W(λ, T) = C1λ -5 ﹝exp(C2 / λT) - 1﹞ -1

[0070] In the formula, W(λ, T) is the spectral power intensity of black body radiation, with the unit of watt·centimeter 2 ·micrometer -1 ; C1 = 3.74×10 -12 is the first radiation constant, with the unit of watt·centimeter 2 ; C2 = 1.43 is the second radiation constant, with the unit of centimeter·K. λ is the spectral radiation wavelength, with the unit of micrometer; T is the black body temperature, with the unit of K.

[0071] In Planck's temperature measurement formula, C1 and C2 are just two constant terms applicable only when the radiator is a black body.

[0072] The inventor found that these two constants can be changed into variables that vary with factors such as the material composition, properties, and shape of the radiator, and the concept of the emissivity coefficient as a parameter is introduced. The temperature measurement formula still retains the core of Planck's formula, that is, the temperature measurement formula according to the present invention is a creative extension of Planck's formula.

[0073] According to an exemplary embodiment, the calculation unit calculates the return cabin outer surface temperature by a predetermined formula, and the calculation formula is as follows:

[0074] ∮(γ, T) = αγ -5 ﹝exp(β / γT)-1﹞ -1

[0075] Where ∮(γ, T) is the thermal radiation intensity of the outer surface of the return capsule, γ is the radiation wavelength, T is the temperature value corresponding to the thermal radiation intensity, α and β are coefficients, and the unit of α is watt cm 2 , and β has a unit of cm·K, wherein the radiation wavelength γ is determined according to the temperature measurement range, and α and β are pre-calibrated according to the temperature measurement range.

[0076] The coefficients α and β in this example embodiment are variable coefficients, and α and β vary with the temperature of the outer surface and the wavelength of radiation. In different temperature ranges, the values ​​of α and β also vary. According to some embodiments, the process of determining α and β, for example, first adjusts the black body to 1000°C and 1100°C, and the sensor component is at a receiving aperture of 2mm and a close distance of 5mm, 5mm being the distance between the sensor component and the surface of the material being measured, and measures the thermal radiation intensity at these two temperatures respectively. In engineering practice, the wavelength range and material corresponding to the photodetector are selected according to the temperature measurement range, such as short-wave measurement of high temperature and long-wave measurement of low temperature, and the radiation wavelength of 0.94um is used for testing in the temperature range of 1000°C and 1100°C. The radiation intensity at 1000°C is ∮1 (γ, 1000°C) and the radiation intensity at 1100°C is ∮2 (γ, 1100°C).

[0077] Substituting the calibration result into the above formula, the obtained ∮1(γ, 1000℃) and ∮2(γ, 1100℃) are combined to obtain:

[0078] ∮1(γ, 1000℃) = αγ -5 ﹝exp(β / γ1000℃)-1﹞ -1

[0079] ∮2(γ, 1100°C) = αγ -5 ﹝exp(β / γ1100℃)-1﹞ -1

[0080] Obtain the values ​​of coefficients α and β that vary with temperature and radiation wavelength.

[0081] By substituting the obtained α and β coefficient values ​​into the calculation formula, the relationship between thermal radiation intensity and temperature can be determined, and the outer surface temperature of the return capsule between 1000℃ and 2000℃ can be measured in real time.

[0082] Similarly, for higher temperatures on the outer surface of the reentry capsule, blackbody calibration sensor assemblies with temperatures of 2000°C and 2100°C can be selected to relatively accurately measure the temperature on the outer surface of the reentry capsule in the temperature range of 2000°C to 3000°C. Blackbody calibration sensor assemblies with temperatures of 2400°C and 2500°C can be selected to measure the temperature on the outer surface of the reentry capsule in the temperature range of 3000°C to 4000°C and higher temperatures.

[0083] According to some embodiments, multiple temperatures can be used to obtain multiple sets of parameter values and store them in a database. During actual temperature measurement, the corresponding parameter values can be obtained from the database based on the measured temperature and temperature change for temperature measurement calculations.

[0084] Using the traditional method requires physical calibration of a blackbody corresponding to the temperature. If a blackbody at 3000 degrees Celsius cannot be manufactured, calibration cannot be performed, and thus temperature measurement cannot be carried out. However, using the method of the present exemplary embodiment only requires calibration with a blackbody once, and the blackbody data at other temperatures can be deduced by combining data with a simple algorithm.

[0085] Through the blackbody calibration sensor assembly, two close but different temperature points are set for calibration to capture the response characteristics of the sensor within the operating temperature range. More close but different temperature points can also be set for calibration, which helps to identify and compensate for the non-linear response of the sensor. Selecting multi-point calibration helps to capture the response characteristics of the sensor within this temperature range and perform appropriate correction.

[0086] The strategy of segmented calibration is applicable to devices such as spacecraft reentry capsules that need to work in extreme environments. As the temperature rises, the radiation characteristics of materials may change. Therefore, for different temperature segments, re-calibration is performed to adapt to the new temperature conditions and ensure the linearity and accuracy of the sensor assembly within different temperature ranges. The reliability and accuracy of the temperature measurement system are improved through precise calibration.

[0087] Figure 3 A flowchart of a method for measuring the temperature on the outer surface of a spacecraft reentry capsule according to an exemplary embodiment is shown.

[0088] See Figure 3 , in S101, obtain the thermal radiation intensity, where the thermal radiation intensity comes from a sensor assembly buried in the thermal insulation layer on the outer surface of the spacecraft reentry capsule.

[0089] According to an exemplary embodiment, the thermal radiation intensity is obtained using a sensor assembly. The thermal radiation intensity is sensed by a photoelectric converter in the sensor assembly and converted into an electrical signal, and then the electrical signal of the photoelectric receiving element is converted into a digital signal by an analog-to-digital converter in the sensor assembly.

[0090] The sensor assembly is buried in the heat insulation layer on the outer surface of the spacecraft reentry capsule. The heat insulation layer has a first heat insulation layer and a second heat insulation layer, and the sensor assembly has a first group of sensors and a second group of sensors. The first group of sensors is buried in the first heat insulation layer at a predetermined angular interval along the circumferential direction, and the second group of sensors is buried in the second heat insulation layer at a predetermined angular interval along the circumferential direction. The second group of sensors is arranged staggeredly with respect to the first group of sensors.

[0091] For example, the first group of sensor assemblies has four sensors, and the interval between adjacent sensors is 90°. The second group of sensor assemblies is buried in the second heat insulation layer. The second group of sensor assemblies also has four sensors, and the interval between adjacent sensors is 90°. The first group of sensors is arranged staggeredly by 45° with respect to the second group of sensors in the circumferential direction.

[0092] In S103, calculate the temperature of the outer surface of the reentry capsule according to the heat radiation intensity data.

[0093] According to the exemplary embodiment, calculate the temperature of the outer surface of the reentry capsule through a predetermined formula. The calculation formula is as follows:

[0094] ∮(γ, T) = αγ -5 ﹝exp(β / γT) - 1﹞ -1

[0095] In the formula, ∮(γ, T) is the heat radiation intensity of the outer surface of the reentry capsule, γ is the radiation wavelength, T is the temperature value corresponding to the heat radiation intensity, and α and β are coefficients. The unit of α is watt·centimeter 2 , the unit of β is centimeter·K, where the radiation wavelength γ is determined according to the temperature measurement range, and α and β are pre-calibrated according to the temperature measurement range.

[0096] The α and β coefficients in this exemplary embodiment are variable coefficients, and α and β vary with the temperature and radiation wavelength of the outer surface. In different temperature ranges, the values of α and β also change.

[0097] According to some embodiments, the process of determining α and β, for example, first adjust the black body to 1000 °C and 1100 °C, and measure the heat radiation intensity at these two temperatures respectively. Select multi-point calibration. Setting two close but different temperature points helps to capture the response characteristics of the sensor within this temperature range and perform appropriate correction.

[0098] The radiation intensity at 1000°C is ∮1(γ, 1000°C) and the radiation intensity at 1100°C is ∮2(γ, 1100°C). The calibration results are sent to the calculation unit and the storage unit, and the obtained ∮1(γ, 1000°C) and ∮2(γ, 1100°C) are combined to obtain the values ​​of the coefficients α and β that vary with temperature and radiation wavelength. Substituting the obtained α and β coefficient values ​​into the calculation formula to clarify the relationship between the thermal radiation intensity and temperature, the outer surface temperature of the return module between 1000°C and 2000°C can be measured in real time.

[0099] For higher temperatures on the outer surface of the return capsule, the 2000℃ and 2100℃ blackbody calibration sensor assemblies can be used to measure the outer surface temperature of the return capsule in the temperature range of 2000℃ to 3000℃. The 2400℃ and 2500℃ blackbody calibration sensor assemblies can be used to measure the outer surface temperature of the return capsule in the temperature range of 3000℃ to 4000℃.

[0100] The method for measuring the outer surface temperature of a spacecraft return capsule proposed in the present invention can realize real-time measurement of the outer surface temperature of a high-speed moving return capsule, and has the advantages of high measurement accuracy, storage and playback of measurement data, etc.

[0101] Figure 4 A block diagram of a computing device is shown according to an exemplary embodiment.

[0102] like Figure 4 As shown, computing device 30 includes processor 12 and memory 14. Computing device 30 may also include bus 22, network interface 16, and I / O interface 18. Processor 12, memory 14, network interface 16, and I / O interface 18 may communicate with each other via bus 22.

[0103] The processor 12 may include one or more general-purpose CPUs (Central Processing Units, processors), microprocessors, or application-specific integrated circuits, etc., for executing relevant program instructions. According to some embodiments, the computing device 30 may also include a high-performance graphics card (GPU) 20 for accelerating the processor 12.

[0104] The memory 14 may include a machine system readable medium in the form of a volatile memory, such as a random access memory (RAM), a read-only memory (ROM) and / or a cache memory. The memory 14 is used to store one or more programs including instructions and data. The processor 12 can read the instructions stored in the memory 14 to execute the above-mentioned method according to the embodiment of the present invention.

[0105] The computing device 30 may also communicate with one or more networks via the network interface 16. The network interface 16 may be a wireless network interface.

[0106] The bus 22 may include an address bus, a data bus, a control bus, etc. The bus 22 provides a path for exchanging information between components.

[0107] It should be noted that, in the specific implementation process, the computing device 30 may further include other components necessary for normal operation. In addition, those skilled in the art can understand that the above devices may also only include the components necessary to implement the solutions of the embodiments of this specification, and do not necessarily include all the components shown in the figure.

[0108] The present invention also provides a computer-readable storage medium, on which a computer program is stored, and when the program is executed by a processor, the steps of the above method are implemented. The computer-readable storage medium may include, but is not limited to, any type of disk, including floppy disks, optical discs, DVDs, CD-ROMs, microdrives, and magneto-optical discs, ROMs, RAMs, EPROMs, EEPROMs, DRAMs, VRAMs, flash memory devices, magnetic or optical cards, nanosystems (including molecular memory ICs), network storage devices, cloud storage devices, or any type of medium or device suitable for storing instructions and / or data.

[0109] The embodiments of the present invention also provide a computer program product, which includes a non-transitory computer-readable storage medium storing a computer program, and the computer program is operable to cause a computer to execute some or all of the steps of any one of the methods described in the above method embodiments.

[0110] Those skilled in the art can clearly understand that the technical solutions of the present invention can be implemented by means of software and / or hardware. The "units" and "modules" in this specification refer to software and / or hardware that can independently complete or cooperate with other components to complete specific functions, where the hardware may be, for example, a field-programmable gate array, an integrated circuit, etc.

[0111] It should be noted that, for the foregoing method embodiments, for the sake of simple description, they are all expressed as a series of action combinations, but those skilled in the art should know that the present invention is not limited by the described action sequence, because according to the present invention, certain steps can be performed in other sequences or simultaneously. Secondly, those skilled in the art should also know that the embodiments described in the specification are all preferred embodiments, and the actions and modules involved are not necessarily essential to the present invention.

[0112] In the above embodiments, the descriptions of the various embodiments have their own emphases. For the parts not detailed in a certain embodiment, reference can be made to the relevant descriptions of other embodiments.

[0113] In several embodiments provided by the present invention, it should be understood that the disclosed device can be implemented in other ways. For example, the device embodiments described above are merely illustrative. For example, the division of units is only a logical function division. In actual implementation, there may be other division methods. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the displayed or discussed coupling or direct coupling or communication connection between each other can be through some service interfaces. The indirect coupling or communication connection of the device or unit can be in electrical or other forms.

[0114] The units described as separate components may or may not be physically separated. The components displayed as units may or may not be physical units, that is, they can be located in one place or distributed to multiple network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0115] In addition, in each embodiment of the present invention, the functional units can be integrated in a processing unit, or each unit can exist physically alone, or two or more units can be integrated in one unit. The above integrated units can be implemented in the form of hardware or in the form of software functional units.

[0116] If the integrated unit is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a computer-readable memory. Based on such an understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. The computer software product is stored in a memory and includes several instructions for causing a computer device (which can be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the methods in each embodiment of the present invention.

[0117] In the above embodiments, the descriptions of each embodiment have their own emphases. For the parts not detailed in a certain embodiment, reference can be made to the relevant descriptions of other embodiments.

[0118] The above specifically shows and describes the exemplary embodiments of the present invention. It should be understood that the present invention is not limited to the detailed structures, setting methods, or implementation methods described here; on the contrary, the present invention is intended to cover various modifications and equivalent settings included within the spirit and scope of the appended claims.

Claims

1. A system for measuring the temperature of the outer surface of a spacecraft return capsule, wherein the outer surface of the spacecraft return capsule has a heat insulation layer, characterized in that: The system comprises: A sensor component, wherein the sensor component is buried in the thermal insulation layer and is used to measure the intensity of thermal radiation; A calculation unit, the calculation unit is arranged in the bottom cabin of the spacecraft return capsule, and is used to receive the thermal radiation intensity data from the sensor assembly and calculate the outer surface temperature of the return capsule according to the thermal radiation intensity data; A storage unit is communicatively connected with the calculation unit and is used to store the thermal radiation intensity data of the sensor assembly and the calculation result of the calculation unit.

2. The system according to claim 1, characterized in that The sensor assembly comprises: A photoelectric converter, the photoelectric converter is used to sense the thermal radiation intensity and convert the thermal radiation intensity into an electrical signal; An analog-to-digital converter converts the electrical signal of the photoelectric converter into a digital signal.

3. The system according to claim 1, characterized in that The calculation unit calculates the outer surface temperature of the return capsule according to the following formula: ∮(γ,T)=γγ -5 ﹝exp(β / γT)-1﹞ -1 Where ∮(γ, T) is the thermal radiation intensity of the outer surface of the return capsule, γ is the radiation wavelength, T is the temperature value corresponding to the thermal radiation intensity, α and β are coefficients, and the unit of α is watt cm 2 The unit of , β is cm·K, where α and β are pre-calibrated according to the temperature measurement range.

4. The system according to claim 1, characterized in that The thermal insulation layer includes a first thermal insulation layer and a second thermal insulation layer, and the sensor assembly includes a first group of sensors and a second group of sensors; The first group of sensors are buried in the first thermal insulation layer along the circumferential direction at a predetermined interval angle; The second group of sensors are buried in the second heat insulation layer along the circumferential direction at a predetermined interval angle, and the second group of sensors are staggered relative to the first group of sensors.

5. The system according to claim 4, characterized in that The first group of sensors includes four sensors, and adjacent sensors are spaced 90° apart; The first group of sensors is arranged at a 45° offset in the circumferential direction relative to the second group of sensors.

6. A method for measuring the temperature of the outer surface of a spacecraft return capsule, characterized in that: include: Acquiring thermal radiation intensity, where the thermal radiation intensity comes from a sensor assembly, where the sensor assembly is buried in a thermal insulation layer on the outer surface of a spacecraft reentry module; The outer surface temperature of the reentry module is calculated based on the thermal radiation intensity data.

7. The method according to claim 6, characterized in that The formula for calculating the outer surface temperature of the return capsule is: ∮(γ,T)=γγ -5 ﹝exp(β / γT)-1﹞ -1 Where ∮(γ, T) is the thermal radiation intensity of the outer surface of the return module, γ is the radiation wavelength, T is the temperature value, α and β are coefficients, and the unit of α is watt cm 2 , and β has a unit of cm·K, wherein the radiation wavelength γ is determined according to the temperature measurement range, and α and β are pre-calibrated according to the temperature measurement range.

8. The method according to claim 6, characterized in that The sensor assembly comprises: A photoelectric converter, the photoelectric converter is used to sense the thermal radiation intensity and convert the thermal radiation intensity into an electrical signal; An analog-to-digital converter converts the electrical signal of the photoelectric converter into a digital signal.

9. The method according to claim 8, characterized in that The thermal insulation layer includes a first thermal insulation layer and a second thermal insulation layer, and the sensor assembly includes a first group of sensors and a second group of sensors; The first group of sensors are buried in the first thermal insulation layer along the circumferential direction at a predetermined interval angle; The second group of sensors are buried in the second heat insulation layer along the circumferential direction at a predetermined interval angle, and the second group of sensors are staggered relative to the first group of sensors.

10. A computing device, characterized in that: include: processor; as well as A memory storing a computer program, which, when executed by the processor, enables the processor to perform the method according to any one of claims 6 to 9.

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