A satellite antenna deployment device comprising a hot knife pressing and a telescopic rod support

Abstract

The present invention discloses a satellite antenna deployment device comprising a hot knife pressing and a telescopic rod support, which relates to the technical field of satellite antenna pressing and deployment. It provides an alternative technical approach for the pressing, releasing, deployment and locking of the deployable antenna mechanism of a spacecraft. The satellite antenna deployment device is installed on the spacecraft body and comprises a top antenna array carbon fiber frame, a side outer antenna array carbon fiber frame, a side inner antenna array carbon fiber frame, a telescopic rod mechanism and a hot knife locking mechanism. At the moment when the hot knife locking mechanism cuts off the binding rope, the side inner antenna array carbon fiber frame and the side outer antenna array carbon fiber frame start to rotate under the torque action of the hinge and the composite hinge, and at the same time, the telescopic rod extends driven by the movement of the side outer antenna array carbon fiber frame. When the upper and lower surfaces of the side inner antenna array carbon fiber frame and the side outer antenna array carbon fiber frame are correspondingly parallel, they are locked, and the antenna completes the entire deployment process.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite antenna clamping and deployment, and is a non-explosive spacecraft mechanism locking and deployment device. Background Art

[0002] With the development of space technology, more and more spacecraft need to be deployed in orbit, and corresponding locking and release devices need to be set for these deployable spacecraft structures. This device is a mechanical device used to achieve the firm connection between the main body and accessories or between components during spacecraft launch, and to release the restraint according to the established requirements after entering orbit. Due to the different masses of spacecraft components and their sensitivities to the impact environment, different locking and release devices are required. According to different unlocking methods, the locking and release devices can be divided into explosive devices and non-explosive devices. The explosive device is the earliest and most commonly used locking and release device on spacecraft, with outstanding advantages such as light weight, small volume, high specific energy, high reliability, and the ability to achieve multi-point synchronous unlocking and release. However, it also has some insurmountable disadvantages, such as: generating a large unlocking impact on the structure during unlocking, the gas generated after explosion having chemical pollution, being disposable, and having a high cost, and it is no longer suitable for some spacecraft. Compared with the explosive device, the non-explosive device generally has the advantages of small impact, no pollution, reusable, and low cost.

[0003] Therefore, the applied research on non-explosive devices has received more and more attention. The thermal knife type locking and release device is a typical non-explosive device, which has significant advantages compared with the explosive unlocking device, such as: being insensitive to electromagnetic interference, having a small unlocking impact, little pollution, simple structure, and being able to be repeatedly tested many times, and has high application value. The thermal knife type locking and release technologies at home and abroad mainly use high-temperature co-fired ceramic electrothermal elements (blade-shaped structures) to cut Kevlar ropes (melting point greater than 500 °C), but its structure is relatively complex, and its mass, volume, required power, and working voltage are all large. Due to the relatively sharp blade-shaped structure of the thermal knife, when contacting the binding rope, vibration friction is likely to cause damage to each other, and it is sensitive to the creep and tension of the binding rope. These factors seriously restrict the application of the thermal knife type locking and release device. Summary of the Invention

[0004] In view of the above problems, the present invention proposes a satellite antenna deployment device including a thermal knife pressing and a telescopic rod support, providing an alternative technical approach for the pressing, release, deployment, and locking of the deployable antenna mechanism of the spacecraft.

[0005] The technical solution of the present invention is as follows: The satellite antenna deployment device is installed on the satellite body 8, and includes a top antenna array carbon fiber frame 1, a side outer antenna array carbon fiber frame 7, a side inner antenna array carbon fiber frame 9, a telescopic rod mechanism, and a thermal knife locking mechanism;

[0006] The carbon fiber frame 1 of the top antenna array is fixedly connected to the star body 8. One end of the carbon fiber frame 9 of the inner side of the side end antenna array is hinged to the carbon fiber frame 1 of the top antenna array, and the other end is hinged to the carbon fiber frame 7 of the outer side of the side end antenna array;

[0007] A hot knife locking mechanism is arranged between the star body 8 and the carbon fiber frame 9 of the inner side of the side end antenna array to lock the carbon fiber frame 9 of the inner side of the side end antenna array before the antenna is deployed;

[0008] A hot knife locking mechanism is also arranged between the star body 8 and the carbon fiber frame 7 of the outer side of the side end antenna array to lock the carbon fiber frame 7 of the outer side of the side end antenna array before the antenna is deployed;

[0009] The telescopic rod mechanism is connected between the star body 8 and the carbon fiber frame 7 of the outer side of the side end antenna array, extends and guides during the antenna deployment process, and locks the carbon fiber frame 7 of the outer side of the side end antenna array after the antenna is deployed.

[0010] Further, the hot knife locking mechanism includes a hot knife 11, a pulley 12 and a rope 14. The hot knife 11 is fixedly connected to the star body 8. The pulley 12 is fixedly connected to the carbon fiber frame 7 of the outer side of the side end antenna array or the carbon fiber frame 9 of the inner side of the side end antenna array. The rope 14 is wound around the hot knife 11 and the pulley 12 in a loop. Thus, the rope can be cut off after the hot knife heats up to release the locking of the carbon fiber frame 9 of the inner side of the side end antenna array.

[0011] Further, the telescopic rod mechanism includes a telescopic rod 10, a telescopic rod sliding pin 4 and a slideway 5. One end of the telescopic rod 10 is hinged to the star body 8, and the other end is fixedly connected to the telescopic rod sliding pin 4. The slideway 5 is fixedly connected to the carbon fiber frame 7 of the outer side of the side end antenna array. The telescopic rod sliding pin 4 is slidably connected in the slideway 5.

[0012] Further, the telescopic rod 10 includes a telescopic inner cylinder and a telescopic outer cylinder that are inserted into each other. The bottom end of the telescopic outer cylinder is hinged to the star body 8. The telescopic rod sliding pin 4 is fixedly connected to the top end of the telescopic inner cylinder. A ball head lock perpendicular to it is fixedly connected to the top opening of the telescopic outer cylinder. A card slot adapted to the ball head lock is opened at the bottom end of the telescopic inner cylinder. Thus, when the telescopic rod 10 extends to the fully deployed state, the ball head in the ball head lock extends into the card slot to lock the length of the telescopic rod 10.

[0013] Further, the carbon fiber frame 1 of the top antenna array and the carbon fiber frame 9 of the inner side of the side end antenna array are hinged through a hinge 3. The hinge 3 includes two hinges hinged through a hinge shaft. The two hinges are respectively fixedly connected to the carbon fiber frame 1 of the top antenna array and the carbon fiber frame 9 of the inner side of the side end antenna array;

[0014] A torsion spring one is also sleeved on the hinge shaft, and both ends of the torsion spring one respectively abut against the top antenna array carbon fiber frame 1 and the inner side antenna array carbon fiber frame 9 at the side end, so that the two maintain a movement trend towards a mutually horizontal state; alternatively, a first rotary power source is installed on the top antenna array carbon fiber frame 1, the output shaft of the first rotary power source is connected to the hinge shaft, and the hinge shaft is fixedly connected to the inner side antenna array carbon fiber frame 9 at the side end through one of the hinges. Thus, after the thermal knife cuts the rope, a torque is generated to make the top antenna array carbon fiber frame 1 and the inner side antenna array carbon fiber frame 9 at the side end move towards a mutually horizontal state.

[0015] Furthermore, the inner side antenna array carbon fiber frame 9 at the side end and the outer side antenna array carbon fiber frame 7 at the side end are hinged through a composite hinge 13;

[0016] The composite hinge 13 includes multiple pairs of support rods. The two support rods in the same pair are hinged through a rotating shaft and are respectively hinged to the inner side antenna array carbon fiber frame 9 at the side end and the outer side antenna array carbon fiber frame 7 at the side end;

[0017] A torsion spring two is also sleeved on the rotating shaft, and both ends of the torsion spring two respectively abut against the outer side antenna array carbon fiber frame 7 at the side end and the inner side antenna array carbon fiber frame 9 at the side end, so that the two maintain a movement trend towards a mutually horizontal state; alternatively, a second rotary power source is installed on the inner side antenna array carbon fiber frame 9 at the side end, the output shaft of the second rotary power source is connected to the rotating shaft, and the rotating shaft is fixedly connected to the outer side antenna array carbon fiber frame 7 at the side end through one of the support rods. Thus, after the thermal knife cuts the rope, a torque is generated to make the inner side antenna array carbon fiber frame 9 at the side end and the outer side antenna array carbon fiber frame 7 at the side end move towards a mutually horizontal state.

[0018] Furthermore, the top antenna array carbon fiber frame 1, the outer side antenna array carbon fiber frame 7 at the side end, and the inner side antenna array carbon fiber frame 9 at the side end are all made of carbon fiber materials.

[0019] When the satellite antenna needs to be deployed in the present invention, at the moment when the thermal knife locking mechanism cuts the binding rope, the inner side antenna array carbon fiber frame at the side end and the outer side antenna array carbon fiber frame at the side end start to rotate under the torque action of the hinge and the composite hinge. At the same time, the telescopic rod extends under the drive of the movement of the outer side antenna array carbon fiber frame at the side end. When the upper and lower surfaces of the inner side antenna array carbon fiber frame at the side end and the outer side antenna array carbon fiber frame at the side end are correspondingly parallel, the hinge and the composite hinge lock the antenna mechanism through their own locking mechanisms. At the same time, the telescopic rod extends to the longest limit position and is locked through its own locking mechanism, improving the stiffness and surface accuracy of the deployed antenna, and the antenna completes the entire deployment process. Description of the Drawings

[0020] Figure 1 Schematic structural diagram of the present invention;

[0021] Figure 2a Schematic structural diagram of the hinge of the present invention;

[0022] Figure 2b Side schematic structural diagram of the hinge of the present invention;

[0023] Figure 3a Schematic structural diagram of the composite hinge of the present invention;

[0024] Figure 3b Side schematic structural diagram of the composite hinge of the present invention;

[0025] Figure 4a Schematic structural diagram of the hot knife of the present invention;

[0026] Figure 4b Side schematic structural diagram of the hot knife of the present invention;

[0027] Figure 5 Schematic diagram of the chute of the present invention;

[0028] Figure 6 Partial schematic diagram of the hot knife, antenna frame and pressing rope under the pressing state

[0029] Figure 7 Schematic diagram of the retracted state of the telescopic support rod, chute mechanism and antenna frame

[0030] Figure 8 Schematic diagram of the deployed state of the telescopic support rod, chute mechanism and antenna frame

[0031] In the figure: 1 - Top antenna array carbon fiber frame, 2 - Antenna array carbon fiber frame support, 3 - Hinge, 4 - Telescopic rod sliding pin, 5 - Slideway, 6 - Telescopic rod, 7 - Outer side end antenna array carbon fiber frame, 8 - Star body, 9 - Inner side end antenna array carbon fiber frame, 10 - Telescopic rod, 11 - Hot knife, 12 - Pulley, 13 - Composite hinge, 14 - Rope. Detailed implementation manners

[0032] To clearly illustrate the technical features of this patent, the following will elaborate on this patent in detail through specific implementation manners and in conjunction with its attached drawings.

[0033] As shown in the present invention Figure 1 The satellite antenna deployment device including hot knife pressing and telescopic rod support shown includes a top antenna array carbon fiber frame 1, an outer side end antenna array carbon fiber frame 7, an inner side end antenna array carbon fiber frame 9, a star body 8, a telescopic rod mechanism, a hot knife locking mechanism, and a hinge mechanism;

[0034] The hot knife locking mechanism includes a hot knife 11 and a pulley 12, as Figure 4a , Figure 4b shown. The hot knife 11 is bolted to the star body 8, and a rope 14 is wound around the hot knife heating rod and the pulley 12 for pressing, and the pulley 12 is bolted to the carbon fiber frame 9 of the inner side of the side end antenna array surface.

[0035] The hinge mechanism includes a hinge 3 between the carbon fiber frame 1 of the top antenna array surface and the carbon fiber frame 9 of the inner side of the side end antenna array surface, and a composite hinge 13 between the carbon fiber frame 9 of the inner side of the side end antenna array surface and the carbon fiber frame 7 of the outer side of the side end antenna array surface. The hinge of the hinge 3 is bolted to the surfaces of the carbon fiber frame 1 of the top antenna array surface and the carbon fiber frame 9 of the inner side of the side end antenna array surface. The composite hinge is embedded in the carbon fiber frame 7 of the outer side of the side end antenna array surface and the carbon fiber frame 9 of the inner side of the side end antenna array surface. The schematic diagram of the composite hinge is as Figure 3a , Figure 3b shown.

[0036] The telescopic rod mechanism includes a telescopic rod 10, a telescopic rod sliding pin 4 and a slideway 5. The telescopic rod 10 is hinged to the star body 8. The telescopic rod sliding pin 4 is fixed on the telescopic rod 10. The slideway 5 is welded to the carbon fiber frame 7 of the outer side of the side end antenna array surface. The telescopic rod sliding pin 4 and the slideway 5 form a pin connection. The slideway is as Figure 5 shown.

[0037] The hinge 3 is fixedly connected to the carbon fiber frame 1 of the top antenna array surface and the carbon fiber frame 9 of the inner side of the side end antenna array surface, or a power source for driving the active rotation of the hinge 3 is provided in the connecting seat.

[0038] At the moment when the hot knife locking mechanism cuts off the binding rope, the carbon fiber frame 9 of the inner side of the side end antenna array surface and the carbon fiber frame 7 of the outer side of the side end antenna array surface start to rotate under the moment action of the hinge and the composite hinge. At the same time, the telescopic rod extends under the drive of the movement of the carbon fiber frame 7 of the outer side of the side end antenna array surface. When the upper and lower surfaces of the carbon fiber frame 9 of the inner side of the side end antenna array surface and the carbon fiber frame 7 of the outer side of the side end antenna array surface are parallel to each other, the hinge and the composite hinge lock the antenna mechanism through their own locking mechanisms, that is, after unfolding, the two hinges rotate along the axis. When reaching the fully unfolded state, the planes of the hinges connected to the axis are mutually pressed to achieve seamless self-locking; at the same time, the telescopic rod extends to the longest limit position and is locked through its own locking mechanism to improve the stiffness and surface accuracy of the antenna after unfolding, and the antenna completes the entire unfolding process.

[0039] The top antenna array carbon fiber frame 1, the side outer antenna array carbon fiber frame 7, and the side inner antenna array carbon fiber frame 9 are all made of carbon fiber materials. The antenna installed on the satellite is required to maintain a stable shape in a space environment with rapidly changing temperature, which requires the use of materials with extremely small linear expansion coefficients, that is, materials with good thermal stability. The rigidity, strength and extremely small linear expansion coefficient required by the antenna can be obtained by selecting the single-layer laying angle, ply ratio and ply sequence of carbon fiber. The antenna generally adopts a high-strength and high-rigidity carbon fiber honeycomb sandwich structure that can withstand the static and dynamic loads of the active section, as well as good microwave reflection characteristics. The density is small. The density of carbon fiber is basically equivalent to that of magnesium and plating, which is 0.20~0.57 times the density of several other metal materials (calculated according to carbon fiber M40JB). Generally speaking, the use of carbon fiber as a structural material can reduce the structural mass by 30%~40%. High specific strength and specific modulus. The comparison of specific strength (ratio of tensile strength to density of materials) and specific modulus (ratio of elastic modulus to density) well illustrates the superiority of carbon fiber in terms of light weight and high strength. The light weight and high strength performance of carbon fiber is the most significant. Its specific strength is 5 times higher than that of steel and 4 times higher than that of aluminum alloy; the specific modulus is 1.3~12.3 times that of other structural materials. Advanced design, carbon fiber reinforced composite material is an anisotropic material, showing significant anisotropy, and there are obvious differences in electrical, magnetic, thermal conductivity, specific heat, thermal expansion coefficient and mechanical properties along the fiber axis and perpendicular to the fiber axis. Composite materials have much higher fatigue resistance than metal materials. Under normal circumstances, the fatigue strength of metal materials is 40%~50% of the ultimate tensile strength, while the fatigue limit of carbon fiber reinforced polymer composite materials can reach 70%~80% of the tensile strength, indicating that the life of composite components is higher than that of traditional material components when working under long-term alternating load conditions. Good high temperature performance. At a temperature of 400℃, the elastic modulus of aluminum alloy drops almost to zero, and the strength also drops significantly. At a high temperature of 400°C, the strength and elastic modulus of carbon fiber remain basically unchanged.

[0040] The hinge 3 and the composite hinge 13 are hidden hinges. When the antenna is fully unfolded, all parts of the hinge are not exposed, which can reduce the difficulty of installing the effective load on the antenna working surface, improve the overall shape accuracy of the antenna working surface, and improve the antenna detection performance.

[0041] The telescopic rod mechanism includes two sections of telescopic sleeve rods. When the antenna is folded, the two small-diameter sleeve rods are stored inside the large-diameter sleeve rod; when the antenna is fully deployed, the small-diameter sleeve rod is fully extended from the large-diameter sleeve rod and is locked by a spherical lock at its end. The entire telescopic rod mechanism supports the antenna array surface, thereby improving the rigidity and surface accuracy of the antenna working surface.

[0042] The bottom end of the carbon fiber frame 1 of the top antenna array is welded to the satellite body 8 through the support 2 of the antenna array carbon fiber frame.

[0043] The hinge of the hinge 3 is bolted to the carbon fiber frame 1 of the top antenna array and the carbon fiber frame 9 of the inner side of the side end.

[0044] The telescopic rod 10 is hinged to the satellite body 8, and the sliding pin 4 on the telescopic rod 10 is pin-connected to the slideway 5 on the carbon fiber frame 7 of the outer side of the side end.

[0045] The composite hinge 13 is embedded in the carbon fiber frame 7 of the outer side of the side end and the carbon fiber frame 9 of the inner side of the side end.

[0046] There are many specific implementation ways of the present invention. The above are only the preferred implementation manners of the present invention. It should be pointed out that for those of ordinary skill in the art, without departing from the principle of the present invention, several improvements can be made, and these improvements should also be regarded as the protection scope of the present invention.

Claims

1. A satellite antenna deployment device comprising a hot knife pressing and a telescopic rod support, characterized in that, The satellite antenna deployment device is installed on the satellite body (8), and it includes a top antenna array carbon fiber frame (1), a side outer antenna array carbon fiber frame (7), a side inner antenna array carbon fiber frame (9), a telescopic rod mechanism, and a hot knife locking mechanism; The top antenna array carbon fiber frame (1) is fixedly connected to the satellite body (8). One end of the side inner antenna array carbon fiber frame (9) is hinged to the top antenna array carbon fiber frame (1), and the other end is hinged to the side outer antenna array carbon fiber frame (7); A hot knife locking mechanism is arranged between the satellite body (8) and the side inner antenna array carbon fiber frame (9) to lock the side inner antenna array carbon fiber frame (9) before the antenna is deployed; A hot knife locking mechanism is also arranged between the satellite body (8) and the side outer antenna array carbon fiber frame (7) to lock the side outer antenna array carbon fiber frame (7) before the antenna is deployed; The telescopic rod mechanism is connected between the satellite body (8) and the side outer antenna array carbon fiber frame (7), extends and guides during the antenna deployment process, and locks the side outer antenna array carbon fiber frame (7) after the antenna is deployed; The hot knife locking mechanism includes a hot knife (11), a pulley (12), and a rope (14). The hot knife (11) is fixedly connected to the satellite body (8). The pulley (12) is fixedly connected to the side outer antenna array carbon fiber frame (7) or the side inner antenna array carbon fiber frame (9). The rope (14) is looped around the hot knife (11) and the pulley (12); The telescopic rod mechanism includes a telescopic rod (10), a telescopic rod sliding pin (4), and a slideway (5). One end of the telescopic rod (10) is hinged to the satellite body (8), and the other end is fixedly connected to the telescopic rod sliding pin (4). The slideway (5) is fixedly connected to the side outer antenna array carbon fiber frame (7). The telescopic rod sliding pin (4) is slidably connected in the slideway (5); The telescopic rod (10) includes a telescopic inner cylinder and a telescopic outer cylinder that are inserted into each other. The bottom end of the telescopic outer cylinder is hinged to the satellite body (8). The telescopic rod sliding pin (4) is fixedly connected to the top end of the telescopic inner cylinder. A ball head lock perpendicular to it is fixedly connected to the top opening of the telescopic outer cylinder. A card slot adapted to the ball head lock is opened at the bottom end of the telescopic inner cylinder; The top antenna array carbon fiber frame (1) and the side inner antenna array carbon fiber frame (9) are hinged through a hinge (3). The hinge (3) includes two hinges hinged by a hinge shaft. The two hinges are respectively fixedly connected to the top antenna array carbon fiber frame (1) and the side inner antenna array carbon fiber frame (9); A first torsion spring is also sleeved on the hinge shaft, and two ends of the first torsion spring respectively abut against the top antenna array carbon fiber frame (1) and the inner side antenna array carbon fiber frame (9) at the side end, so that the two maintain a movement trend towards a mutually horizontal state; alternatively, a first rotary power source is installed on the top antenna array carbon fiber frame (1), an output shaft of the first rotary power source is connected to the hinge shaft, and the hinge shaft is fixedly connected to the inner side antenna array carbon fiber frame (9) at the side end through one of the hinges; The inner side antenna array carbon fiber frame (9) at the side end and the outer side antenna array carbon fiber frame (7) at the side end are hinged through a composite hinge (13); The composite hinge (13) includes multiple pairs of support rods. Two support rods in the same pair are hinged through a rotating shaft and are respectively hinged to the inner side antenna array carbon fiber frame (9) at the side end and the outer side antenna array carbon fiber frame (7) at the side end; A second torsion spring is also sleeved on the rotating shaft, and two ends of the second torsion spring respectively abut against the outer side antenna array carbon fiber frame (7) at the side end and the inner side antenna array carbon fiber frame (9) at the side end, so that the two maintain a movement trend towards a mutually horizontal state; alternatively, a second rotary power source is installed on the inner side antenna array carbon fiber frame (9) at the side end, an output shaft of the second rotary power source is connected to the rotating shaft, and the rotating shaft is fixedly connected to the outer side antenna array carbon fiber frame (7) at the side end through one of the support rods.

2. The satellite antenna deployment device according to claim 1, comprising a hot knife pressing and a telescopic rod support, characterized in that, The top antenna array carbon fiber frame (1), the outer side antenna array carbon fiber frame (7) at the side end, and the inner side antenna array carbon fiber frame (9) at the side end are all made of carbon fiber materials.

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