Mechatronics integrated bus for spacecraft and method of manufacturing the same

Abstract

This invention provides an integrated electromechanical-thermal bus suitable for spacecraft, comprising an integrated flexible conduit and integrated interfaces at both ends of the flexible conduit. The integrated flexible conduit contains spaced regions A, B, C, and D, with region B adjacent to region A and region C adjacent to region D. The integrated interface has fluid interfaces corresponding to regions A and C, a power interface corresponding to region B, and an optical fiber connector corresponding to region D. This invention integrates three previously separate wire, cable, and conduit systems into a single bus, significantly reducing structural weight, the number of mounting brackets, and the envelope space occupied by the installation. Simultaneously, it fundamentally avoids the negative impact of extreme space temperatures on electrical and optical performance, improving transmission efficiency and device lifespan. The manufacturing method of the integrated electromechanical-thermal bus provided by this invention enables the mass production of integrated buses.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft overall design technology, specifically, it relates to an electromechanical-thermal integrated bus suitable for spacecraft and its manufacturing method. Background Technology

[0002] As space missions become increasingly complex, spacecraft are evolving towards larger sizes, higher power densities, and higher data transmission rates. This presents significant challenges to spacecraft system design, primarily in three areas: First, in terms of thermal control systems, to ensure critical payloads operate at suitable temperatures, modern large spacecraft generally require active thermal control systems based on fluid loops (typically single-phase or two-phase fluids) to collect and transfer heat through circulating working fluids. Second, in terms of energy systems, with the surge in payload power demands, traditional cable transmission methods result in heavy cables, high voltage drop losses, and messy layouts. The adoption of centralized power buses (high-voltage DC) for power distribution has become a development trend, effectively reducing weight and improving efficiency. Third, in terms of information systems, the demand for high-speed data interaction has spurred the application of fiber optic signal buses, which offer advantages such as high bandwidth, strong electromagnetic interference resistance, and light weight.

[0003] Currently, the aforementioned fluid pipelines, power cables, and fiber optic cables are designed and laid separately on spacecraft. For example, the invention patent with publication number CN116137374A designs a thermal control scheme combining a pump-driven fluid circuit and a phase change thermal storage module for a spaceborne active phased array antenna, but it is not a whole-star bus system. While the invention patent with publication number CN119730181A addresses the heat dissipation problem inside spaceborne communication equipment, proposing an integrated design that fully integrates energy storage, liquid cooling circuits, and equipment structure, its integration is limited to thermal control and equipment housing, not a whole-star bus-type integration. Furthermore, the invention patent with publication number CN106484001B only addresses the temperature control of onboard cables, employing an inner and outer tank structure and maintaining cable temperature through heating elements and multi-layer thermal insulation components, but it does not involve the integration of thermal control fluid pipelines and fiber optic signals, and it has significant shortcomings in terms of lightweighting and space utilization.

[0004] This traditional approach has significant drawbacks. First, the harsh thermal environment affects electrical / optical performance: the space environment is extreme, with equipment temperatures reaching over +150°C when facing the sun and below -150°C when facing away from the sun. High temperatures significantly reduce the current-carrying capacity of power cables and increase resistance loss; while low temperatures can easily cause fiber optic coatings to become brittle and even develop microcracks, increasing signal attenuation and breakage risks. Although individual thermal insulation coatings, such as multi-layered insulation materials, can be applied to the cables, the effect is limited and further increases mass and complexity. Second, the large mass and envelopment space: the independent structural supports, fasteners, and protective layers of the three systems lead to an accumulation of total mass and space occupied, which is detrimental to the lightweighting and miniaturization of spacecraft. Third, layout conflicts and coordination difficulties: when three independent cables are laid within the confined spacecraft cabin, path conflicts are likely to occur, increasing assembly difficulty and potentially affecting the layout of other equipment.

[0005] Therefore, there is an urgent need to develop an electromechanical-thermal integrated bus suitable for spacecraft and its manufacturing method to overcome the shortcomings of existing technologies. Summary of the Invention

[0006] To address the shortcomings of existing technologies, the purpose of this invention is to provide an electromechanical-thermal integrated bus suitable for spacecraft and its manufacturing method.

[0007] According to the present invention, an electromechanical-thermal integrated bus suitable for spacecraft includes: an integrated flexible conduit 1 and an integrated interface disposed at both ends of the integrated flexible conduit 1;

[0008] The integrated flexible pipeline 1 includes a middle section 11 with a uniform cross-section and four manifolds 12 welded to both ends of the middle section 11. The two fluid channels within the integrated flexible pipeline 1 are connected at both ends to the fluid interfaces on the integrated interface via two manifolds 12.

[0009] Preferably, the intermediate section 11 includes an inner substrate piping layer 2 and a partition section 3; the partition section 3 is disposed inside the inner substrate piping layer 2, forming region A, region B, region C, and region D; Area A is used for cold fluid flow, area B is used for power cable installation, area C is used for hot fluid flow, and area D is used for signal fiber optic installation; area B is adjacent to area A, and area C is adjacent to area D. The integrated interface is equipped with a fluid interface corresponding to region A and region C, a power interface corresponding to region B, and an optical fiber interface connector corresponding to region D. The two ends of regions A and C are respectively connected to the fluid interface through two manifolds 12.

[0010] Preferably, the contact area between region A and region B is greater than the contact area between region B and region C, and the contact area between region C and region D is greater than the contact area between region D and region A; this ensures the cooling effect of region B and the heat preservation effect of region D.

[0011] Preferably, the electromechanical-thermal integrated bus suitable for spacecraft further includes a temperature regulation mechanism; The temperature regulation mechanism includes a controller, a flow valve installed in the fluid interface, and temperature sensors installed in areas B and D. The controller is electrically connected to the flow valve and the temperature sensor respectively. The temperature sensor is used to monitor the temperature in area B or area D in real time and send the data to the controller. The controller controls the opening size of the corresponding flow valve by temperature comparison to ensure that the ambient temperature in area B and area D is maintained within the optimal operating temperature range of the power cable and the signal fiber optic cable respectively.

[0012] Preferably, the optimal operating temperature range is -10°C to +40°C.

[0013] Preferably, the surface of the inner matrix piping layer 2 is wound with a continuous fiber reinforcement layer 4, which can enhance radial strength.

[0014] Preferably, the surface of the fiber reinforcement layer 4 is provided with vibration damping patterns 5 at intervals along its axial direction, which can provide vibration damping capability for the integrated flexible pipeline 1 in the axial direction and improve flexibility.

[0015] Preferably, the two manifolds 12 are connected to the fluid interface by fusion welding of the metal plunger joint 6.

[0016] A method for manufacturing an electromechanical-thermal integrated bus suitable for spacecraft, according to the present invention, includes the following steps: S1. Based on the characteristics of the integrated flexible pipeline cavity 1, an extrusion mold is designed, and the pipeline is prepared using continuous extrusion molding equipment; S2. Design and trial-produce a mold for fusion welding of the middle section 11 and the two end manifolds 12 of the integrated flexible pipeline 1, and perform fusion welding; S3. A continuous fiber reinforcement layer 4 is wound around the surface of the pipeline in the middle section 11, and the winding density is controlled. S4. The structure of the vibration damping texture 5 on the surface of the middle section 11 is prepared by using rotational 3D printing technology; S5. Post-process the damping texture 5 to eliminate microstructural defects that can easily lead to mechanical failure. S6. The two manifolds 12 are connected to the fluid interface through a fusion-welded metal plunger joint 6. S7. Finally, conduct performance testing and verification to evaluate the characteristics of the integrated flexible pipeline 1.

[0017] Preferably, in step S4, the vibration damping texture 5 is prepared by secondary printing of pure TPU on the outer surface of the pipe.

[0018] Compared with the prior art, the present invention has the following beneficial effects: 1. In this invention, by including four separate regions A, B, C, and D and an integrated interface, the original three independent lines, cables, and pipeline systems can be integrated into a single bus. This greatly reduces the structural mass, the number of mounting brackets, and the envelope space occupied by the installation, which is beneficial for the overall optimization design of the spacecraft. At the same time, by utilizing the spacecraft's inherent active thermal control fluid loop, the negative impact of extreme space temperatures on electrical and optical performance is fundamentally avoided, improving transmission efficiency and device lifespan. In addition, the flexible design allows the bus to adapt to the complex interior layout of the spacecraft, flexibly bypass obstacles, and reach various payload devices that need service.

[0019] 2. In this invention, by including a controller, a flow valve installed in the fluid interface, and a temperature regulation mechanism with temperature sensors installed in areas B and D, an active and controllable excellent thermal environment can be provided for temperature-sensitive power cables and optical fibers, further improving transmission efficiency and device lifespan.

[0020] 3. In this invention, the setting of fiber reinforcement layer and vibration damping texture can enhance radial strength and provide a certain vibration damping capacity for integrated flexible pipeline in the length direction, while also improving flexibility.

[0021] 4. In this invention, region B is set adjacent to region A, and region D is set adjacent to region C. A temperature regulation mechanism is configured to utilize the spacecraft’s inherent active thermal control fluid loop to monitor and actively regulate the operating environment temperature of power cables and optical fibers in real time, keeping them within the optimal operating temperature range, which significantly improves transmission efficiency and device lifespan. Attached Figure Description

[0022] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a schematic diagram of the overall structure of the electromechanical-thermal integrated bus applicable to spacecraft according to the present invention; Figure 2 This is a schematic cross-sectional view of the middle section of the electromechanical-thermal integrated bus applicable to spacecraft according to the present invention; Figure 3 This is a schematic diagram illustrating the manufacturing process of the electromechanical-thermal integrated bus intermediate segment applicable to spacecraft according to the present invention; Figure 4 This is a schematic diagram of the electromechanical-thermal integrated bus fusion welding process applicable to spacecraft according to the present invention; Figure 5 This is a schematic diagram showing the connection between the electromechanical and thermal integrated bus manifold and the metal plunger connector of the present invention applicable to spacecraft.

[0023] The diagram shows: Detailed Implementation

[0024] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0025] An electromechanical-thermal integrated bus suitable for spacecraft includes an integrated flexible conduit 1 and an integrated interface disposed at both ends of the integrated flexible conduit 1. The integrated flexible conduit 1 is provided with a region A for cold fluid flow, a region B for power cable installation, a region C for hot fluid flow, and a region D for installing signal optical fibers. Region B is adjacent to region A, and region C is adjacent to region D. The integrated interface is provided with a fluid interface corresponding to region A and region C, a power interface corresponding to region B, and an optical fiber interface connector corresponding to region D.

[0026] The integrated flexible pipeline 1 includes a middle section 11 with a uniform cross section and four manifolds 12 welded to both ends of the middle section 11. The two ends of region A and region C are respectively connected to the fluid interface through two manifolds 12.

[0027] The electromechanical-thermal integrated bus for spacecraft also includes a temperature control mechanism. This mechanism comprises a controller, flow valves installed within the fluid interface, and temperature sensors installed in zones B and D. The controller is electrically connected to both the flow valves and temperature sensors. The temperature sensors monitor the temperature in zone B or zone D in real time and transmit the data to the controller. The controller controls the opening size of the corresponding flow valves based on temperature comparisons, ensuring that the ambient temperature in zones B and D is maintained within the optimal operating temperature range for the power cables and signal fibers, respectively. The optimal operating temperature range is -10°C to +40°C.

[0028] The intermediate section 11 includes an inner matrix pipe layer 2 and a partition section 3 disposed inside the inner matrix pipe layer 2. The surface of the inner matrix pipe layer 2 is wrapped with a continuous fiber reinforcement layer 4, and the surface of the fiber reinforcement layer 4 is provided with vibration damping patterns 5 at intervals along its axial direction.

[0029] A method for manufacturing an electromechanical-thermal integrated bus suitable for spacecraft includes the following steps: S1. Based on the characteristics of the integrated flexible pipeline cavity 1, an extrusion mold is designed, and the pipeline is prepared using continuous extrusion molding equipment; S2. Design and trial-produce a mold for fusion welding of the middle section 11 and the two end manifolds 12 of the integrated flexible pipeline 1, and perform fusion welding; S3. A continuous polyimide fiber reinforcement layer 4 is wound around the surface of the pipeline in the middle section 11, and the winding density is controlled. S4. The structure of the vibration damping texture 5 on the surface of the middle section 11 is prepared by using rotational 3D printing technology; S5. Post-process the damping texture 5 to eliminate microstructural defects that can easily lead to mechanical failure. S6. The two manifolds 12 are connected to the fluid interface through a fusion-welded metal plunger joint 6. S7. Finally, conduct performance testing and verification to evaluate the characteristics of the integrated flexible pipeline 1.

[0030] In step S4, the vibration damping texture 5 is prepared by secondary printing of pure TPU on the outer surface of the pipe.

[0031] The invention will now be described in more detail with reference to the accompanying drawings.

[0032] like Figure 1 , Figure 2 As shown, an electromechanical-thermal integrated bus suitable for spacecraft includes an integrated flexible conduit 1 and integrated interfaces disposed at both ends of the integrated flexible conduit 1. The integrated flexible conduit 1 is provided with a region A for cold fluid flow, a region B for power cable installation, a region C for hot fluid flow, and a region D for installing signal optical fibers. Region B is adjacent to region A, and region C is adjacent to region D. The integrated interface is provided with a fluid interface corresponding to region A and region C, a power interface corresponding to region B, and an optical fiber interface connector corresponding to region D. The integrated flexible conduit 1 includes a middle section 11 with a uniform cross section and four manifolds 12 welded to both ends of the middle section 11. The two ends of region A and region C are respectively connected to the fluid interface through two manifolds 12.

[0033] By incorporating four separate zones (A, B, C, and D) and an integrated interface, the original three independent wire, cable, and conduit systems can be merged into a single bus. This significantly reduces structural mass, the number of mounting brackets, and the envelope space occupied by the cabling, which is beneficial for the overall optimization design of the spacecraft. At the same time, by utilizing the spacecraft's inherent active thermal control fluid loop, the negative impact of extreme space temperatures on electrical and optical performance is fundamentally avoided, improving transmission efficiency and device lifespan. Furthermore, the flexible design allows the bus to adapt to complex spacecraft cabin layouts, flexibly bypassing obstacles to reach various payload devices that require service.

[0034] like Figure 2 As shown, an electromechanical-thermal integrated bus suitable for spacecraft also includes a temperature regulation mechanism. The temperature regulation mechanism includes a controller, a flow valve installed in the fluid interface, and temperature sensors installed in zones B and D. The controller is electrically connected to the flow valve and the temperature sensor respectively. The temperature sensor is used to monitor the temperature in zone B or zone D in real time and send the data to the controller. The controller controls the opening size of the corresponding flow valve by temperature comparison to ensure that the ambient temperature in zone B and zone D is maintained within the optimal operating temperature range of the power cable and the signal fiber optic cable, respectively. The optimal operating temperature range is -10℃ to +40℃.

[0035] By incorporating a temperature regulation mechanism that includes a controller, a flow valve installed in the fluid interface, and temperature sensors installed in zones B and D, an active, controllable, and superior thermal environment is provided for temperature-sensitive power cables and optical fibers, further improving transmission efficiency and device lifespan.

[0036] like Figures 1 to 3 As shown, the intermediate section 11 includes an inner substrate piping layer 2 and a partition section 3 disposed inside the inner substrate piping layer 2. The partition section 3 effectively divides region A, region B, region D, and region C. In addition, to ensure the cooling effect of region B and the heat preservation effect of region D, the contact area between region A and region B is greater than the contact area between region B and region C. Similarly, the contact area between region C and region D is greater than the contact area between region D and region A. At the same time, a continuous fiber reinforcement layer 4 is wound around the surface of the inner substrate piping layer 2, and vibration damping patterns 5 are arranged at intervals along its axial direction on the surface of the fiber reinforcement layer 4.

[0037] By setting the fiber reinforcement layer 4 and the vibration damping pattern 5, the radial strength can be enhanced and a certain vibration damping capacity can be provided for the integrated flexible pipeline 1 in the axial direction, while also improving flexibility.

[0038] This invention also provides a method for manufacturing an electromechanical-thermal integrated bus suitable for spacecraft, such as... Figures 1 to 5 As shown, it includes the following steps: S1. Based on the characteristics of the integrated flexible pipeline cavity 1, an extrusion mold is designed, and a continuous extrusion molding equipment is used to prepare the pipeline, ensuring the uniformity of the structural dimensions and phase state of the main body of the pipeline, as well as ensuring the density and pressure resistance of the internal structure of the pipeline, thereby ensuring the reliability and stability of the mechanical properties of the pipeline. S2. Design and trial production of a mold for fusion welding of the middle section 11 and the two end manifolds 12 of the integrated flexible pipeline 1. To ensure the dimensional accuracy of the inner and outer diameters of the hose, a steel mandrel coated with polytetrafluoroethylene is proposed to be inserted into the pipeline. The two printed pipeline sections are clamped and fixed by welding mold. The corresponding melt is injected, cooled and shaped, and the middle section 11 and the two end manifolds 12 are connected. The melt is made of TPU material with a hardness of 80A to 90A, which has high strength, high toughness, high flexibility, high wear resistance, wide temperature adaptability and corrosion resistance, and good processability. S3. A continuous polyimide fiber reinforcement layer 4 is wound around the surface of the pipeline in the middle section 11. The winding density is controlled and the outer diameter of the pipeline is optimized to ensure that it has high toughness, high and low temperature resistance, radiation resistance and corrosion resistance, low density, low dielectric and low water absorption. S4. Use rotational 3D printing technology to prepare a special textured structure on the pipe surface, namely vibration damping texture 5. In view of the size and function of the textured structure, use pure TPU for secondary printing on the outer surface of the pipe. S5. Post-process the prepared outer 3D printed structure of the pipeline, namely the vibration damping texture 5, to eliminate microstructural defects that are prone to mechanical failure. S6. The two manifolds 12 are connected to the fluid interface via fusion welding of the metal plunger joints 6, specifically through the outer nut, thereby ensuring that the integrated bus can be connected to the fluid circuit through the fluid interface. S7. Finally, conduct performance testing and verification to evaluate the characteristics of the integrated piping system.

[0039] This invention integrates three separate lines, cables, and conduits into a single bus, significantly reducing structural mass, the number of mounting brackets, and the envelope space occupied by the cabling. This facilitates overall spacecraft optimization design. Furthermore, by utilizing the spacecraft's inherent active thermal control fluid loop, it fundamentally avoids the negative impact of extreme space temperatures on electrical and optical performance, improving transmission efficiency and device lifespan. In addition, the flexible design allows the bus to adapt to complex spacecraft cabin layouts, flexibly bypassing obstacles to reach various payload devices that require service.

[0040] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0041] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. An electromechanical-thermal integrated bus suitable for spacecraft, characterized in that, include: An integrated flexible pipeline (1) and integrated interfaces at both ends of the integrated flexible pipeline (1); The integrated flexible pipeline (1) includes a middle section (11) with equal cross-section and four manifolds (12) welded to both ends of the middle section (11). The two fluid channels within the integrated flexible pipeline (1) are connected at both ends to the fluid interfaces on the integrated interface via two manifolds (12).

2. The electromechanical-thermal integrated bus for spacecraft according to claim 1, characterized in that, The intermediate section (11) includes an inner substrate pipeline layer (2) and a partition section (3); the partition section (3) is located inside the inner substrate pipeline layer (2) to form region A, region B, region C and region D; Area A is used for cold fluid flow, area B is used for power cable installation, area C is used for hot fluid flow, and area D is used for signal fiber optic installation; area B is adjacent to area A, and area C is adjacent to area D. The integrated interface is equipped with a fluid interface corresponding to region A and region C, a power interface corresponding to region B, and an optical fiber interface connector corresponding to region D. The two ends of regions A and C are respectively connected to the fluid interface through two manifolds (12).

3. The electromechanical-thermal integrated bus for spacecraft according to claim 2, characterized in that, The contact area between region A and region B is greater than the contact area between region B and region C, and the contact area between region C and region D is greater than the contact area between region D and region A; this ensures the cooling effect of region B and the heat preservation effect of region D.

4. The electromechanical-thermal integrated bus for spacecraft according to claim 2, characterized in that, The electromechanical-thermal integrated bus for spacecraft also includes a temperature regulation mechanism; The temperature regulation mechanism includes a controller, a flow valve installed in the fluid interface, and temperature sensors installed in areas B and D. The controller is electrically connected to the flow valve and the temperature sensor respectively. The temperature sensor is used to monitor the temperature in area B or area D in real time and send the data to the controller. The controller controls the opening size of the corresponding flow valve by temperature comparison to ensure that the ambient temperature in area B and area D is maintained within the optimal operating temperature range of the power cable and the signal fiber optic cable respectively.

5. The electromechanical-thermal integrated bus for spacecraft according to claim 4, characterized in that, The optimal operating temperature range is -10℃ to +40℃.

6. The electromechanical-thermal integrated bus for spacecraft according to claim 2, characterized in that, The surface of the inner matrix pipeline layer (2) is wrapped with a continuous fiber reinforcement layer (4), which can enhance the radial strength.

7. The electromechanical-thermal integrated bus for spacecraft according to claim 6, characterized in that, The surface of the fiber reinforcement layer (4) is provided with vibration damping patterns (5) at intervals along its axial direction, which can provide vibration damping capability for the integrated flexible pipeline (1) in the axial direction and improve flexibility.

8. The electromechanical-thermal integrated bus for spacecraft according to claim 1, characterized in that, The two manifolds (12) are connected to the fluid interface by fusion welding of metal plunger joints (6).

9. A method for manufacturing an electromechanical-thermal integrated bus suitable for spacecraft, characterized in that, Includes the following steps: S1. Based on the cavity characteristics of the integrated flexible pipeline (1), an extrusion mold is designed, and the pipeline is prepared using a continuous extrusion molding equipment; S2. Design and trial production of a mold for fusion welding of the middle section (11) and the manifolds (12) of the integrated flexible pipeline (1), and perform fusion welding; S3. A continuous fiber reinforcement layer (4) is wound around the surface of the pipeline in the middle section (11), and the winding density is controlled. S4. The vibration damping texture (5) on the surface of the middle section (11) is prepared by using rotational 3D printing technology; S5. Post-process the damping texture (5) to eliminate microstructural defects that are prone to mechanical failure. S6. The two manifolds (12) are connected to the fluid interface through a fusion-welded metal plunger joint (6); S7. Finally, conduct performance testing and verification to evaluate the characteristics of the integrated flexible pipeline (1).

10. The method for manufacturing an electromechanical-thermal integrated bus suitable for spacecraft according to claim 9, characterized in that, In step S4, the vibration damping texture (5) is prepared by secondary printing of pure TPU on the outer surface of the pipe.

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