Infrared heating device for vacuum thermal testing

Abstract

This disclosure relates to an infrared heating device for vacuum thermal testing. The infrared heating device includes: multiple plate units, heating wires, and connectors. The surface of each plate unit has multiple mesh openings and wiring grooves distributed between the mesh openings; the heating wires are embedded in the wiring grooves via a heat-conducting medium; the connectors are located on the plate units near the ends of the heating wires; the connectors are used to detachably electrically connect the heating wires to conductors, which are connected to an external power source; the multiple plate units can be detachably assembled to form a continuous radiating surface and enclose a heating chamber adapted to the shape of a spacecraft; the heating wires generate heat when energized and transfer the heat to the plate units via the heat-conducting medium; the plate units transfer external heat flow to the spacecraft surface through thermal radiation. This disclosure, through modular design and assembly, enables the infrared heating device to possess high flexibility and adaptability.

Description

Technical Field

[0001] This disclosure relates to the field of spacecraft thermal testing technology, specifically to an infrared heating device for vacuum thermal testing. Background Technology

[0002] During the satellite development phase, vacuum thermal testing is a crucial step in ensuring the stable and reliable operation of the satellite in the complex thermal environment of space. Accurately simulating the heat flow absorbed by the satellite in orbit is a key factor in improving the precision of thermal balance testing, and the infrared heating cage is a crucial device for simulating extraterrestrial heat flow through thermal radiation. Traditional infrared heating cages generally employ a fixed nickel-chromium alloy heating belt and a metal frame integrated structure; however, this structure has several problems.

[0003] Specifically, the integrated structure of the fixed metal frame and heating band makes it difficult to adjust the shape of the device. When the size or shape of the test object changes, the entire frame must be remade, leading to a significant increase in cost. Furthermore, the rigid connection between the heating band and the frame results in an unadjustable heat source distribution, easily creating heating blind spots at the frame connection points, severely affecting the accuracy of heat flow simulation. Moreover, the heating elements of traditional infrared heating cages are exposed, resulting in numerous and complex cables during testing, increasing wiring difficulty and posing significant electrical safety hazards. In addition, the maintenance cost of traditional infrared heating cages is high; the integrated structure means that partial damage requires replacement of the entire heating unit.

[0004] It is evident that there is an urgent need in this field to develop an infrared heating device that combines high adaptability, high thermal field uniformity, and high reliability to meet the stringent requirements of vacuum thermal testing for next-generation spacecraft. Utility Model Content

[0005] This disclosure provides an infrared heating device for vacuum thermal testing in order to address the problems existing in the prior art.

[0006] According to a first aspect of this disclosure, an infrared heating device for vacuum thermal testing is provided, comprising:

[0007] Multiple plate units, each plate unit having multiple mesh holes on its surface and wiring grooves distributed between the multiple mesh holes;

[0008] A heating wire, wherein the heating wire is configured to be embedded in the wiring groove via a heat-conducting medium;

[0009] A connector is disposed on the plate unit near the end of the heating wire; the connector is configured to detachably electrically connect the heating wire to a conductor configured to connect to an external power source.

[0010] The plate units are detachable and spliced ​​to form a continuous radiating surface and enclose a heating chamber adapted to the shape of the spacecraft; the heating wire is configured to generate heat when energized and transfer the heat to the plate units through the heat-conducting medium; the plate units are configured to transfer external heat flow to the spacecraft surface through thermal radiation.

[0011] In one embodiment of this disclosure, the plate unit is an aluminum plate, and at least one side of the aluminum plate is processed to form a radiating surface, wherein the thermal emissivity of the radiating surface is greater than or equal to 0.7.

[0012] In one embodiment of this disclosure, one side of the aluminum plate is processed to form a first radiating surface, and the other side of the aluminum plate is processed to form a second radiating surface; wherein the first radiating surface and the second radiating surface have different thermal emissivity.

[0013] In one embodiment of this disclosure, the total area of ​​the mesh in the continuous radiating surface formed by splicing multiple plate units ranges from 45% to 75%.

[0014] In one embodiment of this disclosure, the thermally conductive medium is thermally conductive silicone grease.

[0015] In one embodiment of this disclosure, the distance between the inner surface of the heating chamber and the outer surface of the spacecraft ranges from 10 to 30 cm.

[0016] In one embodiment of this disclosure, a single plate unit is composed of at least two plate sub-units; the at least two plate sub-units are joined at their edges on a plane to form the plate unit.

[0017] In one embodiment of this disclosure, multiple heating wires are laid on a plate unit and / or a plate sub-unit, wherein the ends of at least two of the heating wires are arranged adjacent to each other and are configured to be connected to a conductor via the same connector.

[0018] In one embodiment of this disclosure, one of the heating wires is configured to be laid across at least two plate units and / or at least two plate sub-units.

[0019] In one embodiment of this disclosure, the heating wire includes a heating wire and an insulating outer sheath wrapped around the heating wire; the connector includes a bayonet structure for wiring, the surface of which is configured as an insulating surface.

[0020] One beneficial effect of this disclosure is that by forming a continuous radiating surface through detachable and splicable plate units, the shape of the heating chamber can be dynamically adapted to different spacecraft shapes, thereby solving the problem of poor versatility of traditional infrared heating cages. This disclosure, through modular design in production and assembly, enables the infrared heating device to possess high flexibility and adaptability, facilitating rapid assembly and adjustment under different environments and experimental requirements, reducing experimental costs, and simplifying replacement and maintenance.

[0021] Furthermore, the mesh size, shape, and density on the plate unit can be adjusted according to experimental needs, thereby adjusting the heat exchange area between the spacecraft and the heat sink, and improving the efficiency and accuracy of thermal experiments. In addition, according to different experimental needs, wiring grooves can be specifically opened on the plate unit to adjust the specific wiring method; the heating wires are embedded in the wiring grooves through a heat-conducting medium, which improves the efficiency of heat conduction, reduces the temperature gradient, and ensures the consistency of the experiment.

[0022] Other features and advantages of this disclosure will become clear from the following detailed description of exemplary embodiments with reference to the accompanying drawings. Attached Figure Description

[0023] The accompanying drawings, which are incorporated in and form a part of this specification, illustrate embodiments of the present disclosure and, together with their description, serve to explain the principles of the present disclosure.

[0024] Figure 1 This is a schematic diagram of a plate unit and heating wire structure provided in an embodiment of the present disclosure;

[0025] Figure 2 This is a partial cross-sectional view of a plate unit provided in an embodiment of this disclosure;

[0026] Figure 3 This is a schematic diagram of a plate unit and heating wire structure provided in an embodiment of the present disclosure;

[0027] Figure 4 This is a schematic diagram of an application scenario of the infrared heating device provided in an embodiment of this disclosure;

[0028] Figure 5 This is a schematic diagram of the structure of a plate unit formed by splicing four plate sub-units according to an embodiment of the present disclosure;

[0029] Figure 6 This is a schematic diagram of the plate unit partitioning in a verification embodiment of this disclosure.

[0030] Figures 1 to 6 The one-to-one correspondence between the component names and the reference numerals in the figures is as follows:

[0031] 100. Infrared heating device; 1. Plate unit; 10. Plate sub-unit; 11. Mesh; 12. Wiring trough; 2. Heating wire; 21. Heating wire; 22. Insulating outer sheath; 3. Heat-conducting medium; 4. Wiring component; 5. Wire; 200. Vacuum tank; 300. Wave-absorbing heat sink. Detailed Implementation

[0032] Various exemplary embodiments of the present disclosure will now be described in detail with reference to the accompanying drawings. It should be noted that, unless otherwise specifically stated, the relative arrangement, numerical expressions, and values ​​of the components and steps set forth in these embodiments do not limit the scope of the present disclosure.

[0033] The following description of at least one exemplary embodiment is merely illustrative and is in no way intended to limit this disclosure or its application or use.

[0034] Techniques, methods, and equipment known to those skilled in the art may not be discussed in detail, but where appropriate, such techniques, methods, and equipment should be considered part of the specification.

[0035] It should be noted that similar labels and letters in the following figures indicate similar items; therefore, once an item is defined in one figure, it does not need to be discussed further in subsequent figures.

[0036] In this article, terms such as "up," "down," "front," "back," "left," and "right" are used only to indicate the relative positional relationship between related parts, rather than to define the absolute position of these related parts.

[0037] In this article, "first," "second," etc., are used only to distinguish one another, and not to indicate degree of importance, order, or prerequisite for each other.

[0038] In this document, terms such as “equal” and “same” are not strict mathematical and / or geometric limitations, but also include errors that are understandable to those skilled in the art and permissible in manufacturing or use.

[0039] Vacuum thermal testing is a core technology for simulating the high vacuum, cold black background, and extreme temperature environment of space, primarily used to verify the on-orbit operational reliability of spacecraft and their components. This technology creates a high vacuum environment by evacuating gas from a vacuum pump array within a container, and combines this with a heating and cooling system to achieve cyclical temperature control, thereby replicating the alternating heating and cooling processes experienced by the spacecraft. (Reference) Figure 4 The vacuum thermal test system mainly consists of a vacuum tank 200, an infrared heating device 100, a microwave absorbing heat sink 300, and a supporting structure. All components work together to accurately simulate the complex working conditions in outer space.

[0040] Specifically, the vacuum chamber 200, as the main test container, is typically made of stainless steel and designed as a cylindrical or cuboid structure. It is equipped with an air extraction port, heat-conducting pipes, and safety monitoring devices, capable of withstanding extremely low pressure while maintaining stable thermal environment control. Its core function is to provide a sealed vacuum space for the test specimen while maintaining background temperature uniformity through a thermal control system. The infrared heating device 100, as the core heat source of the test system, precisely regulates the specimen temperature through radiative heat transfer. Typical forms include infrared heating cages, lamp arrays, and thin-film heaters. The microwave absorbing heat sink 300, as a multifunctional composite component, is usually installed around the infrared heating device 100. Its main body adopts an absorbing cone structure and is coated with a high-temperature resistant coating, combining microwave absorption and thermal radiation enhancement functions. This design can both shield the inner wall of the vacuum chamber 200 from thermal interference and absorb the microwave energy released by the test specimen, preventing high-power signal reflection from interfering with the test data.

[0041] During the test, the infrared heating device 100, equipped with the microwave absorbing heat sink 300, needs to be placed entirely inside the vacuum chamber 200. The fixing method is selected according to the size of the specimen; for example, an adjustable bracket or trolley support can be used, or a suspension system combined with flexible connectors can be used for fixation. The test procedure mainly includes three stages: vacuuming, thermal cycling, and data acquisition. The vacuum system first reduces the pressure inside the chamber to the target value. Then, the infrared heating device 100 applies a temperature load according to a preset program. The built-in sensors and control system of the microwave absorbing heat sink 300 provide real-time temperature data feedback and dynamically adjust the heat flow input to maintain stable test conditions. Vacuum thermal testing is widely used for the thermal coupling verification of precision components such as spaceborne antennas and optical payloads, providing key technical support for the on-orbit reliability assessment of spacecraft.

[0042] As described in the background art, traditional infrared heating cages typically employ a fixed nickel-chromium alloy heating belt and a metal frame integrated structure. This structure suffers from problems such as poor versatility, high cost, and inconvenience in storage. To address these issues, this disclosure provides an infrared heating device 100 for vacuum thermal testing and its assembly method.

[0043] refer to Figures 1 to 3This disclosure provides an infrared heating device 100 for vacuum thermal testing, comprising: multiple plate units 1, heating wires 2, and connectors 4. The multiple plate units 1 are detachable and can be assembled to form a continuous radiating surface, enclosing a heating chamber adapted to the shape of a spacecraft. The plate units 1 can be modularly and flexibly designed; for example, various sizes and shapes of plate units 1 can be manufactured according to requirements and assembled together to enclose the heating chamber. The shape of the heating chamber can be highly adapted to the shape of the spacecraft. Suitable plate units 1 can be selected for assembly based on different spacecraft shapes, or new plate units 1 can be designed specifically for this purpose. The internal space of the three-dimensional structure formed by the plate units 1 is the heating chamber, and its inner surface is a continuous radiating surface adapted to the shape of the spacecraft. This allows for full-coverage thermal radiation simulation of the spacecraft placed inside the heating chamber, avoiding heating blind spots (dead zones) and improving the accuracy of vacuum thermal testing.

[0044] In one embodiment of this disclosure, the distance between the inner surface of the heating chamber and the outer surface of the spacecraft ranges from 10 to 30 cm, preferably 20 cm. This ensures the uniformity of thermal radiation. Specifically, in a vacuum, heat is transferred only through radiation. If the distance is too close (e.g., less than 10 cm), the local radiation from the infrared heating cage will be too concentrated, disrupting the uniformity of temperature distribution; while if the distance is too far (e.g., more than 30 cm), the radiation energy will attenuate significantly, potentially failing to effectively simulate the target heat flow. An interval of approximately 20 cm has been experimentally verified to achieve a balance between radiation intensity and coverage, ensuring uniform heating of the spacecraft surface and meeting the simulation requirements of extraterrestrial heat flow.

[0045] Furthermore, the heat sink 300 is typically mounted on plate unit 1 to absorb the heat released by the spacecraft. If the distance between the spacecraft and the inner surface of the heating chamber is too small, the heat emitted may be blocked by plate unit 1 itself, preventing the heat sink 300 from efficiently absorbing heat and affecting the simulation effect of the cold black environment. An interval of approximately 20cm provides a sufficient path for the radiative transfer of heat to the heat sink 300, ensuring heat dissipation efficiency.

[0046] refer to Figures 1 to 3The plate unit 1 has multiple mesh holes 11 on its surface, and wiring grooves 12 distributed among the mesh holes 11. The mesh holes 11 are formed through the plate unit 1, and can be various shapes such as round holes, oblong holes, elliptical holes, rectangular holes, rhomboid holes, and triangular holes. It should be noted that this disclosure does not specifically limit the shape, number, density, or size of the mesh holes 11. Different mesh holes can be selectively formed according to specific experimental requirements, thereby changing the obstruction of the spacecraft surface and meeting the required heat exchange area between the spacecraft and the heat sink 300, thereby improving the accuracy and efficiency of the experiment.

[0047] In one embodiment of this disclosure, the total area of ​​the mesh 11 in the continuous radiating surface formed by splicing multiple plate units 1 ranges from 45% to 75%. The coverage factor for the spacecraft surface in vacuum thermal testing ranges from approximately 0.45 to 0.75. For example, when the test requires a coverage factor of 0.5, the specific opening method of the mesh 11 can be designed so that the total area of ​​the mesh 11 accounts for approximately 50% of the continuous radiating surface of the heating chamber.

[0048] The heating wire 2 is constructed to be embedded in the wiring groove 12 via a heat-conducting medium 3. The design of the wiring groove 12, which is formed between the mesh openings 11, should meet the test requirements to ensure that the heating wire 2 can be reasonably arranged and achieve the best heat conduction effect. For example, the wiring can be laid out in a zigzag or S-shape, so that the heating wire 2 can be evenly distributed on the board unit 1. This disclosure does not limit the specific size of the groove, and the groove can be made according to the size of the heating wire 2 selected for this test. When laying the wiring, the heating wire 2 needs to be placed in the wiring groove 12, and it should be ensured that the heating wire 2 is straightened in the groove to avoid bending or damage. By setting the embedded heating wire 2, sufficient contact area between the heating wire 2 and the board unit 1 is ensured to optimize the heat conduction effect. Then, aviation tape can be used to fix the heating wire 2 in the wiring groove 12 and seal it to ensure that the heating wire 2 is firmly held in the correct position.

[0049] In one specific embodiment of this disclosure, the thermally conductive medium 3 is thermally conductive silicone grease. Before installing the heating wire 2, thermally conductive silicone grease needs to be evenly applied in the wiring groove 12 to enhance the heat conduction between the board unit 1 and the heating wire 2, improve heat utilization, and ensure efficient heat transfer. Besides thermally conductive silicone grease, other materials can also be selected as the thermally conductive medium 3. Specifically, the thermally conductive medium 3 needs to meet requirements such as high temperature resistance, low volatility, high thermal conductivity, and long-term stability. For example, materials such as graphite heat sinks, thermally conductive gels, and thermally conductive silicone pads can be selected to replace thermally conductive silicone grease.

[0050] refer to Figure 1 and Figure 3A connector 4 is provided on the plate unit 1 near the end of the heating wire 2. The connector 4 is configured to detachably connect the heating wire 2 to the conductor 5, which is configured to connect to an external power source. The heating wire 2 is configured to generate heat when energized and transfer the heat to the plate unit 1 through the thermally conductive medium 3. The plate unit 1 is configured to transfer external heat flow to the spacecraft surface through thermal radiation.

[0051] In one specific embodiment of this disclosure, such as Figure 2 As shown, the heating wire 2 includes a heating wire 21 and an insulating outer sheath 22 wrapped around the heating wire 21. The connector 4 includes a bayonet structure for wiring, the surface of which is constructed as an insulating surface. The insulating outer sheath 22 on the surface of the heating wire 2 reduces the safety hazards of the heating wire 2. It is fixed to the plate unit 1 by insulating aviation tape, thereby achieving secondary insulation and improving the safety of the test. The bayonet structure is used to connect the heating wire 2 and the wire 5, thereby enabling convenient and quick wire connection by plugging in. Compared with the traditional welding method, this greatly reduces the wiring processing time and effectively improves the assembly efficiency. The surface of the bayonet structure is constructed as an insulating surface, thereby providing insulation protection for the cable connection points and effectively reducing the safety hazards caused by the numerous cables inside the tank.

[0052] This disclosure utilizes detachable and connectable plate units 1 to form a continuous radiating surface, allowing the heating chamber shape to dynamically adapt to different spacecraft shapes, thereby solving the problem of poor versatility of traditional infrared heating cages. Through modular design in production and assembly, this disclosure enables the infrared heating device to possess high flexibility and adaptability, facilitating rapid assembly and adjustment under different environments and experimental requirements, reducing experimental costs, and facilitating replacement and maintenance.

[0053] Furthermore, the size, shape, and density of the mesh 11 on the plate unit 1 can be adjusted according to experimental needs, thereby adjusting the heat exchange area between the spacecraft and the heat sink and improving the efficiency and accuracy of thermal experiments. In addition, according to different experimental needs, wiring grooves 12 can be specifically opened on the plate unit 1 to adjust the specific wiring method; the heating wire 2 is embedded in the wiring groove 12 through the heat-conducting medium 3, which improves the efficiency of heat conduction, reduces the temperature gradient, and ensures the consistency of the experiment.

[0054] In one embodiment of this disclosure, the plate unit 1 is an aluminum plate, at least one side of which is treated to form a radiating surface, the thermal emissivity of which is greater than or equal to 0.7. Specifically, the surface treatment of the aluminum plate can be a high emissivity treatment process, preferably an anodizing process or a black paint spraying process. Untreated aluminum plates have very low thermal emissivity; only in a small number of tests where maintaining a low emissivity coefficient for the plate unit 1 is required can tests be conducted directly without surface treatment. In most cases, tests require the thermal emissivity of the plate unit 1 to reach 0.7 or higher, thus necessitating a high emissivity treatment process on the aluminum plate surface.

[0055] Anodizing processes generate a uniform black anodic oxide film on the surface of aluminum plates through an electrolytic reaction, while black paint spraying can also treat the surface of aluminum plates to be black. Those skilled in the art know that aluminum itself is a highly reflective material, strongly reflecting infrared radiation, resulting in low emissivity. Treating the surface of aluminum plates black can suppress reflection and enhance absorption, thereby increasing its thermal emissivity. This characteristic is particularly important in vacuum thermal testing, enabling efficient heat transfer and reducing energy waste.

[0056] In one embodiment of this disclosure, one side of the aluminum plate is treated to form a first radiating surface, and the other side of the aluminum plate is treated to form a second radiating surface. The first and second radiating surfaces have different thermal emissivity, and generally, the difference in thermal emissivity between the first and second radiating surfaces can be less than or equal to 0.1. In some experimental scenarios, it is necessary for the two sides of the aluminum plate to have different emissivity. In this case, the two sides of the aluminum plate can be treated differently, for example, different surface treatment processes can be used, or the same process can be used for different degrees of treatment. This disclosure allows for flexible adjustment of the radiation characteristics of the two sides of the aluminum plate, improving heat utilization efficiency, avoiding heating blind spots, and achieving uniform thermal radiation coverage.

[0057] In one embodiment of this disclosure, reference is made to Figure 5 Each plate unit 1 is composed of at least two plate sub-units 10; at least two plate sub-units 10 are joined at their edges on a plane to form plate unit 1. For example, four plate sub-units 10 can be spliced ​​together to form a complete plate unit 1. This disclosure can design the shape and size of each plate unit 1, as well as the specific structure of the mesh 11 and wiring groove 12 on the plate unit 1, according to the shape of the spacecraft and the experimental requirements; each plate unit 1 can be further divided into different partitions, and the coverage coefficient, shape, size, and other requirements of each partition can be independent of each other; according to the requirements of each partition, different plate sub-units 10 can be prepared separately, and then spliced ​​together to form a complete plate unit 1.

[0058] This disclosure further enhances the flexibility and adaptability of the infrared heating device. By selecting and cutting plate sub-units 10 of different shapes and sizes, plate units 1 of arbitrary shapes and sizes can be spliced ​​together. Mesh holes 11 of different sizes, shapes, and densities can be opened on different plate sub-units 10, and different heating wires 2 can be laid out, thereby achieving different degrees of heating on different areas of a surface of a spacecraft, thus improving the accuracy of the experiment.

[0059] The infrared heating device disclosed herein can achieve different wiring methods. Specifically, in one embodiment of this disclosure, such as... Figure 1 As shown, when a heating wire 2 is laid on a plate unit 1 and / or a plate sub-unit 10, connectors 4 can be installed at both ends of the heating wire 2.

[0060] In another embodiment of this disclosure, such as Figure 3 As shown, multiple heating wires 2 are laid on a plate unit 1 and / or a plate sub-unit 10, wherein at least two heating wires 2 are arranged with their ends adjacent to each other and are configured to be connected to the wire 5 through the same connector 4. For example, two heating wires 2 can be laid on a plate unit 1. In this case, three connectors 4 can be set on the plate unit 1, one of which is connected to the ends of two heating wires 2. This allows the two heating wires 2 to be connected to the wire 5 respectively. Thus, after the external power is turned on, the wire 5 can supply power to all the heating wires 2 on the plate unit 1, so that the heating wires 2 heat up evenly on the plate unit 1.

[0061] In another embodiment of this disclosure, a heating wire 2 is configured to be laid across at least two plate units 1 and / or at least two plate sub-units 10. Specifically, the heating wire 2 can be laid across adjacent plate units 1 and / or plate sub-units 10. In this case, the two ends of the heating wire 2 can be located on two plate units 1 and / or plate sub-units 10 respectively. Thus, only one connector 4 can be provided on each plate unit 1 and / or plate sub-unit 10.

[0062] This disclosure also provides a method for assembling an infrared heating device for vacuum thermal testing, used to assemble the aforementioned infrared heating device for vacuum thermal testing. The assembly method includes the following steps:

[0063] S1: Based on the shape and size of the spacecraft to be tested, multiple plate units 1 are prepared, and multiple mesh holes 11 and wiring grooves 12 distributed among the mesh holes 11 are formed on each plate unit 1. Specifically, the plate unit 1 can be an aluminum plate. In this case, step S1 also includes: performing surface treatment on at least one side of the aluminum plate to improve its thermal emissivity. The surface treatment can specifically be anodizing or black paint spraying, thereby improving the thermal emissivity of the aluminum plate surface to meet the specific requirements of the test.

[0064] Further, step S1 includes: designing the density and / or shape and / or size of multiple mesh openings 11 based on the target coverage coefficient of this experiment, and opening multiple mesh openings 11 on the plate unit 1. Depending on the specific experimental requirements, different mesh openings can be selectively opened to change the obstruction of the spacecraft surface, meeting the required heat exchange area between the spacecraft and the absorbing heat sink 300, thereby improving the accuracy and efficiency of the experiment. In addition, based on the experimental requirements, an appropriate heating wire 2 can be selected, and the specific routing method can be designed accordingly; according to the selected heating wire 2 model, routing grooves 12 can be correspondingly opened between the mesh openings 11.

[0065] In one embodiment of this disclosure, the plate unit 1 in step S1 is prepared by the following steps: based on the target shape and target size of the plate unit 1, at least two plate sub-units 10 are obtained; the at least two plate sub-units 10 are joined together at their edges on a plane to form the plate unit 1. This disclosure allows for the design of the shape and size of each plate unit 1 according to the shape of the spacecraft and experimental requirements. Each plate unit 1 can be further divided into different partitions, and the coverage coefficient, shape, size, and other requirements of each partition can be independent of each other. Different plate sub-units 10 can be prepared separately according to the requirements of each partition, and then spliced ​​together to form a complete plate unit 1.

[0066] S2: Fill the wiring groove 12 with thermally conductive medium 3. Specifically, the thermally conductive medium 3 can be thermally conductive silicone grease. Before installing the heating wire 2, thermally conductive silicone grease needs to be evenly applied to the wiring groove 12 to enhance the heat conduction between the board unit 1 and the heating wire 2, improve the heat utilization rate, and ensure efficient heat transfer.

[0067] S3: Embed and secure the heating wire 2 within the wiring groove 12. Specifically, place the heating wire 2 into the wiring groove 12, ensuring it is straightened within the groove to prevent bending or damage. By using an embedded heating wire 2, sufficient contact area is ensured between the heating wire 2 and the board unit 1, optimizing heat conduction. Then, aviation tape can be used to secure the heating wire 2 within the wiring groove 12 and seal it, ensuring the heating wire 2 remains firmly in the correct position.

[0068] S4: Install a connector 4 on the plate unit 1 near the end of the heating wire 2, and connect the heating wire 2 and the conductor 5 through the connector 4. Specifically, this disclosure connects the heating wire 2 and the conductor 5 through the connector 4, ensuring a reliable and stable electrical connection. Compared with traditional welding, the connector 4 connection method simplifies the manufacturing process and reduces processing time.

[0069] S5: Multiple plate units 1 are assembled and enclosed to form a heating chamber adapted to the shape of the spacecraft. Specifically, according to the shape of the spacecraft to be heated, multiple plate units 1 are assembled into a three-dimensional shape and enclosed to form a heating chamber. The distance between the inner surface of the heating chamber and the outer surface of the spacecraft can be about 20cm, thereby ensuring that the spacecraft surface is heated evenly and meets the simulation requirements of heat flow outside space.

[0070] S6: All wires 5 of the plate unit 1 are brought together to the same terminal and connected to an external power source. Specifically, after modular assembly, all wires 5 can be guided to the same position to achieve circuit conduction. When the external power source is turned on, all wires 5 can simultaneously supply power to all heating wires 2, and the heating wires 2 laid on each plate unit 1 can heat up synchronously and evenly, ensuring uniform heating of the spacecraft surface.

[0071] Example

[0072] 1. Purpose: This embodiment aims to verify the feasibility of splicing modular plate units and their heating uniformity, confirm the temperature control accuracy of each zone under the target external heat flow conditions, and ensure that the relative error is ≤5%, thereby providing experimental data support for the engineering application of the infrared heating device provided in this disclosure.

[0073] 2. Method:

[0074] 2.1 Modular assembly: Reference Figure 6 The edges of the four plate sub-units (partitions ① and ② are rectangles, and partitions ③ and ④ are right trapezoids) are joined together to form a shape as shown below. Figure 6 The complete plate unit shown ensures that the geometry matches the actual experimental requirements.

[0075] 2.2 Mesh Formation and Heating Wire Arrangement: Mesh forms are created on the plate unit with a coverage factor of 0.5. Wiring grooves (2mm in diameter) are created on the plate unit, and heating wires with a resistance of 6.2Ω / m are selected. Two 7.8-meter heating wires are arranged in each of zones ① and ②, and two 10.8-meter heating wires are arranged in each of zones ③ and ④, with the lengths evenly distributed to balance the heat load.

[0076] 2.3 Vacuum Environment Setup: The wired board unit is suspended in a vacuum chamber by an insulating rope. Heat flow meters are suspended at the center of each partition, 20cm away from the surface of the board unit, to simulate the distance between the outer surface of the spacecraft and the radiating surface of the board unit in actual experiments. Four heat flow meters are used to monitor the temperature of the four partitions in real time.

[0077] 2.4 Thermal Cycling Test: The target heat flux density for the test is 403.24 W / m³. 2 199.81W / m 2 The corresponding heat flux meter temperatures were 25℃ and -23℃, respectively. The temperature was maintained for 5 hours after reaching the target temperature each time, and this cycle was repeated 4 times. The measured temperatures of the heat flux meters in each zone were recorded, and the relative errors between these temperatures and the target values ​​were calculated.

[0078] 3. Results and Analysis:

[0079] Table 1. Test Results Data of Verification Examples

[0080]

[0081] See Table 1, at 403.24 W / m 2 and 199.81W / m 2 Under both heat flux conditions, the measured temperatures of the four zones remained stable near the target values, with a maximum relative error of only 0.32%, far below 5%, demonstrating the temperature uniformity of the plate unit. Furthermore, the plate unit did not deform or loosen its interfaces under vacuum suspension, maintaining its geometric integrity and ensuring reliable assembly.

[0082] The experimental results demonstrate that the splicing of different shaped partitions and the uniform heating verify the feasibility of the modular design of the infrared heating device disclosed herein, which can highly flexibly adapt to the thermal testing requirements of various complex-shaped spacecraft. Furthermore, a relative error within 5% proves the effectiveness of the heating wire arrangement, the uniformity of heat flow distribution meets expectations, and the thermal control accuracy is up to standard.

[0083] The various embodiments of this disclosure have been described above. These descriptions are exemplary and not exhaustive, and are not limited to the disclosed embodiments. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments. The terminology used herein is chosen to best explain the principles, practical application, or technical improvements to the embodiments in the market, or to enable others skilled in the art to understand the embodiments disclosed herein. The scope of this disclosure is defined by the appended claims.

Claims

1. An infrared heating device for vacuum heat testing, characterized by comprising: include: Multiple plate units (1), the surface of the plate unit (1) is provided with multiple mesh holes (11) and wiring grooves (12) distributed between the multiple mesh holes (11); Heating wire (2), the heating wire (2) is configured to be embedded in the wiring groove (12) through a heat-conducting medium (3); A connector (4) is disposed on the plate unit (1) near the end of the heating wire (2); the connector (4) is configured to detachably electrically connect the heating wire (2) to a conductor (5), the conductor (5) being configured to connect to an external power source; Among them, multiple plate units (1) can be disassembled and spliced ​​to form a continuous radiating surface and enclose a heating chamber adapted to the shape of the spacecraft; the heating wire (2) is configured to generate heat when energized and transfer the heat to the plate unit (1) through the heat-conducting medium (3); the plate unit (1) is configured to transfer the target external heat flow to the spacecraft surface through thermal radiation.

2. The infrared heating device of claim 1, wherein, The plate unit (1) is an aluminum plate, and at least one side of the aluminum plate is processed to form a radiating surface, the thermal emissivity of the radiating surface being greater than or equal to 0.

7.

3. The infrared heating device according to claim 2, characterized in that, One side of the aluminum plate is processed to form a first radiating surface, and the other side of the aluminum plate is processed to form a second radiating surface; wherein the first radiating surface and the second radiating surface have different thermal emissivity.

4. The infrared heating device according to claim 1, characterized in that, In the continuous radiating surface formed by splicing multiple plate units (1), the total area of ​​the mesh (11) accounts for 45%-75%.

5. The infrared heating device according to claim 1, characterized in that, The thermally conductive medium (3) is thermally conductive silicone grease.

6. The infrared heating device according to claim 1, characterized in that, The distance between the inner surface of the heating chamber and the outer surface of the spacecraft ranges from 10 to 30 cm.

7. The infrared heating device according to claim 1, characterized in that, Each of the plate units (1) consists of at least two plate sub-units (10); at least two plate sub-units (10) are joined at their edges on a plane to form the plate unit (1).

8. The infrared heating device according to claim 7, characterized in that, Multiple heating wires (2) are laid on a plate unit (1) and / or a plate sub-unit (10), wherein at least two of the heating wires (2) are arranged adjacent to each other and are configured to be connected to a wire (5) through the same connector (4).

9. The infrared heating device according to claim 7, characterized in that, One of the heating wires (2) is configured to be laid across at least two plate units (1) and / or at least two plate sub-units (10).

10. The infrared heating device according to claim 1, characterized in that, The heating wire (2) includes a heating wire (21) and an insulating outer sheath (22) wrapped around the heating wire (21); the connector (4) includes a bayonet structure for wiring, the surface of which is configured as an insulating surface.

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