Spacecraft and separation mechanism for a spacecraft

Abstract

The invention relates to a spacecraft (100) and a separation mechanism (130) for a spacecraft, wherein the separation mechanism (130) connects spacecraft components (110, 120). One of at least two structural elements of the separation mechanism (130) consists of a material whose melting point is lower than the temperature acting on the spacecraft (100) during reentry into a planet's atmosphere. This enables a passive, yet safe and controlled initiation or execution of fragmentation.

Description

[0001] The invention relates to a spacecraft comprising a first spacecraft component and a second spacecraft component and a passive separation mechanism for separating the first spacecraft component from the second spacecraft component, as well as a separation mechanism for a spacecraft.

[0002] To protect vital satellite orbits, space debris must be avoided. Future spacecraft in LEO (Low Earth Orbit) and MEO (Medium Earth Orbit) must be returned to Earth after their operational lifetime. To prevent damage to the Earth's surface during reentry, large spacecraft in particular must be fragmented early in the reentry process. The individual, small fragments then burn up in the upper atmosphere.

[0003] The separation of spacecraft components is possible through active measures. Passive separation mechanisms are also known. For example, early fragmentation of a spacecraft during reentry can be achieved using adapted metallic connectors that join the individual components of the satellite.

[0004] Spacecraft in which two spacecraft components are connected by a connecting element, wherein the connecting element is a solder that melts upon frictional heat generated during the spacecraft's entry into the Earth's atmosphere, are known from EP 3 085 627 B1 and EP 4 361 047 A1. The melting of the connecting element provides a passive separation mechanism.

[0005] To manufacture the spacecraft, the spacecraft components must be precisely aligned with each other. Subsequently, the connecting element must be formed and molded onto the two spacecraft components.

[0006] A simplification of the assembly process is known from EP 3 246 256 B1. This document discloses a fastening device for a spacecraft, consisting of a shaft and a head, the shaft and head being connected by a solder joint. The solder joint is meltable by frictional heat generated during the spacecraft's reentry into the Earth's atmosphere. A passive separation mechanism is provided by separating the shaft and head. This fastening device can be manufactured independently of, and even before, the assembly of the spacecraft components, thus simplifying the assembly process and significantly improving the possibilities for inspecting the individual components.

[0007] The breakdown of the bond is based on the thermodynamic phase transformation of melting and therefore requires no actuation mechanism (electronics, actuator, etc.). The necessary heating to break the bond occurs passively during reentry due to friction between the atmosphere and the spacecraft. The separation of the two spacecraft components is then caused by the spacecraft's tumbling motion during reentry.

[0008] Against this background, the object of the invention is to provide an alternative and / or improved method for the passive separation of spacecraft components upon re-entry into the atmosphere of a planet.

[0009] According to the invention, a spacecraft or a separation mechanism for a spacecraft is provided according to the features of the independent claims. The dependent claims describe further embodiments of the present invention.

[0010] The design according to the invention makes it possible, on the one hand, to use standardized fasteners, for example screws, and on the other hand, to provide a separation mechanism that can be prefabricated in series, thus enabling high quality at low cost. Furthermore, the separation mechanism according to the invention is easy to assemble.

[0011] By attaching a standardized connecting element, directly or indirectly, to a structural element, which melts upon reentry and thus breaks the bond with adjacent structural elements, the spacecraft components can be separated. The direct or indirect integration of the melting structural element into the structure of the spacecraft component provides good protection for the structural element during the spacecraft's operational lifetime. Simultaneously, the melting structural element provides ample space, making the removal of the third structural element from its mounting position safer and smoother compared to prior art designs.

[0012] In this context, the term "melting" refers to the process by which the solid, dimensionally stable, and structurally stable state of the structural element is dissolved. This occurs through heating. Between the solid and liquid states, there can be a transitional range in which the material's creep properties increase, the material becomes softer and more fluid, and finally liquid. Here, "melting" does not refer to the mere attainment of the liquid state, but rather encompasses the transitional phenomena, i.e., the entire melting process.

[0013] One aspect is that the passive separation mechanism is configured to transfer heat from the first spacecraft component and / or the second spacecraft component to the structural elements in order to melt at least one of the structural elements. This accelerates the initiation of fragmentation and makes the fragmentation process safer.

[0014] According to one embodiment, the spacecraft has at least one separation mechanism designed according to the invention. Another embodiment provides that the spacecraft has several separation mechanisms designed according to the invention, in particular that a spacecraft component is connected to at least two of the separation mechanisms.

[0015] The temperature acting on the separation mechanism, especially on the structural elements, during re-entry into an atmosphere is also referred to as the re-entry temperature.

[0016] One aspect is that the second structural element consists of a material whose melting point is lower than that of the first or third structural element. This ensures that the solid fastener easily slides out of the softening structural element, even with differing melting points. Both of these factors facilitate fragmentation.

[0017] The structural element made of the material with the lower melting temperature can, for example, be a plastic.

[0018] The structural elements, in particular the connecting element, are preferably made of a metal alloy.

[0019] The structural element containing the material with the higher melting point preferably consists of a metal alloy.

[0020] It is also possible that the structural element with the lower melting point, i.e., the second structural element, is made of a material such as a metal alloy. However, plastics or a combination of metal and plastic can also be used here.

[0021] For the first structural element and / or the third structural element and / or the connecting element, preferably one or more identical and / or different metal alloys such as aluminum, in particular 2024 aluminum, 6061 aluminum, 7075 aluminum, or titanium, in particular Ti6Al4V, or bismuth-tin, in particular BiSn42, or tin-silver, in particular Sn96Ag4, or tin-copper, in particular Sn99.3Cu0.7, are used.

[0022] For the structural element with the material having the lower melting temperature, in particular with a melting temperature which is lower than the re-entry temperature, alloys with tin, especially bismuth-tin or tin-silver or tin-copper or a combination thereof, have proven to be particularly advantageous.

[0023] A particular embodiment provides that the melting temperature of the structural element with the material having the lower melting temperature is less than 400°C, preferably less than 280°C, particularly preferably in a range between 120°C and 250°C, especially in a range between 100°C and 200°C, whereby the material with the lower melting temperature is meltable by frictional heat, which occurs when the first spacecraft component and the second spacecraft component enter the atmosphere of a planet.

[0024] One variant of an initial version of the separation mechanism envisions the first structural element being made of the material with the lower melting point, and the second structural element being made of the material with the higher melting point. In this variant, the separation mechanism would detach from the first spacecraft component upon reentry. For example, the first structural element could be made of a solder material and the second structural element of an aluminum alloy.

[0025] Another variant of the first version of the separation mechanism proposes that the second structural element be made of the material with the lower melting point and the first structural element be made of the material with the higher melting point. In this variant, the separation mechanism would disintegrate upon re-entry. For example, the second structural element might be made of a solder material and the first structural element of an aluminum alloy.

[0026] According to one aspect, the connecting element consists of two sections, referred to as the shaft and the head. This designation does not specify the shape. The head can be similar to or identical in design to the shaft. In the connected state, the head is associated with the second spacecraft component. In the connected state, the shaft is associated with the first spacecraft component.

[0027] One aspect is that the second or third structural element provides a cylindrical receiving space designed to accommodate at least a section of the fastener's shaft, with the second structural element having an inner surface. This ensures that at least a portion of the shaft is enclosed by the structural element, thus providing a large contact area. During operation, this large contact area ensures the necessary strength, and in the event of fragmentation, it ensures good heat transfer and, if necessary, a large area of ​​the fastener is exposed.

[0028] One aspect is that the connecting element is separate from the third structural element. In this case, the separation mechanism comprises the first, second, and third structural elements. The separate connecting element can be attached to the third structural element during the assembly of the spacecraft components, preferably in a detachable manner. Upon reentry, the first and third structural elements would separate, particularly through softening and / or melting of the second structural element. The connecting element would then remain attached to the third structural element.

[0029] One aspect is that the shank of the connector is cylindrical and has an outer surface. This facilitates assembly, but may also promote fragmentation.

[0030] Preferably, corresponding structures are formed on the outer surfaces of the shank of the fastener and the second or third structural element, enabling a force-fit and / or form-fit connection between the fastener and the second structural element, particularly by applying a relative rotational movement. It is especially recommended that corresponding threads are provided on the outer surfaces. Preferably, the fastener is designed as a threaded rod or screw. The connection can also be made via a bayonet fitting.

[0031] An intermediary threaded element, for example a so-called wire thread insert, can be provided between the structuring of the fastener and the structuring of the structural element, which, in the sense of the present invention, represents a connection, in particular an indirect one, between the fastener and the structural element. This threaded element is also referred to as a wire thread insert and preferably consists of a wire made of copper or a copper alloy.

[0032] An alternative design for the force-fit and / or form-fit connection of the fastener and the second or third structural element, using corresponding structuring of the lateral surfaces, is possible through a translational movement of the fastener into the receiving space of the second structural element. In this process, the structuring interlocks and suppresses the reversal of the movement.

[0033] The connection between the second or third structural element and the connecting element is preferably designed to be detachable without damage. However, it is also possible that the connection between the second or third structural element and the connecting element prevents non-destructive detachment after its creation.

[0034] One aspect is that the first spacecraft component has a recess or an outer edge to which the first structural element is attached, preferably permanently, ideally by a positive fit and / or by a material bond. By means of a recess or an outer edge, several differently oriented surfaces of the first spacecraft component are provided for connection with the first structural element, which can be used to create a particularly durable connection.

[0035] The second structural element is arranged within the first structural element. The first structural element is designed such that it completely encloses the second structural element perpendicular to the longitudinal extent of the receiving space of the second structural element.

[0036] Preferably, the average thickness of the structural element made from the material with the low melting point, perpendicular to its longitudinal extent, is equal to or less than the average thickness of the structural element made from the material with the high melting point. This provides a particularly large amount of material for melting and, with the melting, a large amount of space for relative movement of the structural elements to each other or of the connecting element to a structural element, thus facilitating fragmentation.

[0037] A further factor favoring separation and fragmentation is explained below.

[0038] One aspect is that the first structural element has at least one flange-like projection at its distal end, extending longitudinally and bearing against at least one surface of the first spacecraft component. This protects the edge of the recess or the outer edge of the spacecraft component. Furthermore, it increases the contact area between the spacecraft component and the structural element, resulting in an improved connection between the two. This also provides support for the structural element, preventing it from tilting within the recess or at its outer edge.

[0039] One aspect is that the connection between the first structural element and the spacecraft component is made by means of a material bond using a filler or adhesive.

[0040] According to an advantageous embodiment, the first structural element has at least one flange-like projection at each of its two opposite distal ends in the longitudinal direction. This allows the spacecraft component to be positively engaged by the structural element, thereby improving the connection between the structural element and the spacecraft component. Preferably, the projection of one distal end rests against one surface, and the projection of the opposite distal end rests against the opposite surface.

[0041] Another preferred embodiment provides that the outer diameter of the first structural element corresponds exactly to the recess in the spacecraft component in which the first structural element is arranged. This allows for a flush surface with a reduced gap. Such a flush mounting is used, for example, in the Fig. 10 and Fig. 11 shown.

[0042] One aspect is that, for attaching the separation mechanism to the spacecraft component, the recess in the spacecraft component is larger than necessary to accommodate the separation mechanism itself. A filler material is introduced between the spacecraft component and the first structural element of the separation mechanism, which, once cured, creates a material-bonded and / or form-fit connection between the first structural element of the separation mechanism and the spacecraft component.

[0043] One aspect is that the first structural element, with its flange-like projections on both sides, grips the surfaces of the first spacecraft component in the area of ​​the recess or the outer edge, thereby not only protecting the edge of the recess or the outer edge, particularly against damage, but also ensuring that the structural element is positively engaged and held against the spacecraft component.

[0044] Preferably, the projections are rotationally symmetrical. However, the projections can also have a different shape, for example, a rectangle or polygon.

[0045] One aspect is that at least the first spacecraft component consists of a multilayer structure, in particular a sandwich component. Preferably, the at least first spacecraft component consists of a three-layer composite material, wherein the outer layers, which provide the surfaces, are flat and / or closed, and the middle layer has a cellular structure with numerous cavities. The walls of the cellular structure are preferably oriented orthogonally to the extent of the outer layers, for example as webs, in particular a honeycomb structure. However, the walls of the cellular structure can also be curved and / or chaotically distributed, for example, if the middle layer is designed as a foam.

[0046] A particularly preferred embodiment has a honeycomb-shaped cell structure.

[0047] An alternative embodiment consists in at least one of the spacecraft components being made of a monolithic body, for example, a carbon fiber composite material or a metal.

[0048] The design of the structural element with projections on both sides has proven to be particularly advantageous in conjunction with a spacecraft component consisting of a multi-layered structure.

[0049] Spacecraft typically consist of sandwich structures. These sandwich structures—also known as spacecraft components—are bolted together to form the spacecraft's load-bearing structure. The sandwich structures usually incorporate one-piece, metallic inserts, also called structural elements. These serve as force transmission elements for mounting the load-bearing structure and for accommodating equipment. The connection is then achieved via bolted joints. In summary, the present invention is a temperature-sensitive bolted connection point that serves as a separation mechanism.

[0050] This separation mechanism, preferably metallic, consists essentially of two components: an insert made up of several structural elements and a connecting element. The insert consists of at least two parts, but can also consist of three. These parts are also called structural elements.

[0051] According to a preferred embodiment, at least one part of the separation mechanism consists of an alloy commonly used for connections in aerospace applications, for example with aluminum or titanium, which has a higher melting point. At least a second part consists of an alloy with a significantly lower melting point.

[0052] The melting point of both parts is higher than the temperature encountered during operation. However, the melting point of at least one part is significantly lower than the temperature it experiences during reentry into Earth's atmosphere. Upon reentry, the parts with a melting point below the reentry temperature melt, potentially causing the connection to detach, for example, by the screw coming loose from the insert. This allows for controlled separation.

[0053] Preferably, only one of the parts has a melting point lower than the re-entry temperature. However, it is also possible for several parts to have a melting point lower than the re-entry temperature.

[0054] This passive dissolution of the connection allows the components of the spacecraft to separate from each other promptly and in a controlled manner during reentry.

[0055] The first structural element of the separation mechanism according to the invention can be placed in through holes and blind holes of spacecraft components and also on edges of the spacecraft components.

[0056] The second or third structural element may have through holes, blind holes, through threads and / or blind threads of different sizes.

[0057] One aspect is that the first structural element has protrusions to increase the contact area with the spacecraft component.

[0058] Another aspect is that the first structural element and also the third structural element consist of a shaft and at least one, preferably two, flange sections. The shafts and the flange sections define a cavity.

[0059] The second structural element is arranged in the cavity, in particular formed, preferably pre-formed.

[0060] For the initial shaping of the second structural element into the cavity, at least one opening is provided in the separation mechanism, in particular in the first structural element.

[0061] One aspect is that the openings are spatially limited and / or not rotationally symmetrical with respect to the separation mechanism. A further function of the openings is that the second structural element also fills the openings during the initial forming process. Due to their monolithic construction, the openings act as an anti-rotation device for the second structural element.

[0062] This at least one opening is preferably provided on one of the head sides of the separating mechanism. In particular, all openings of a separating mechanism are provided on the same head side.

[0063] One aspect is that the openings are arranged on the head side which is opposite the head side with the preferred access for the connecting element.

[0064] The cavity, particularly for the initial forming of the second structural element, is bounded at the head end by the flange sections. Both the first and third structural elements have flange sections. These flange sections are arranged in pairs facing each other. This enlarges the cavity for the second structural element. As the size of the second structural element increases, the fragmentation of the separation mechanism, and thus the separation of the spacecraft components during reentry, is facilitated.

[0065] One aspect is that the cavity has a special design. Specifically, a flange section of the first structural element and a flange section of the third structural element are each wedge-shaped. This means that these flange sections have a surface oriented towards the cavity, which is inclined relative to the orientation of the cavity and / or the longitudinal direction of the receiving space, preferably at an angle between 20° and 70° relative to the orientation of the cavity and / or the longitudinal direction of the receiving space.

[0066] Another aspect is that the surfaces of the wedge-shaped flange sections are aligned approximately parallel to each other, in particular they enclose an angle of less than 30°, preferably an angle of less than 10°.

[0067] The design of the flange sections as wedge surfaces allows for a further improvement in separation properties upon re-entry. This prevents the first and third structural elements from becoming jammed or entangled.

[0068] One aspect is that the two wedge-shaped flange sections are positioned on opposite ends of the separation mechanism. Preferably, in a connection position of the structural elements with respect to a separation direction, the wedge-shaped flange section of the first structural element is arranged further forward, i.e., closer to a separation opening formed upon re-entry, than the wedge-shaped flange section of the third structural element.

[0069] The first structural element can also be designed in two parts, consisting of an outer part and an inner part. The outer part is positioned facing the space component or attached to it. The inner part is positioned facing away from the space component and is connected to the second structural element. To connect the inner and outer parts, these have positive locking elements, in particular threads or snap-fit ​​projections, on their facing surfaces.

[0070] The invention allows for different embodiments. Some examples are listed below. The invention is not limited to the embodiments shown. One or more features of the following embodiments can also be combined. Fig. Figure 1 shows a schematic representation of a spacecraft with multiple spacecraft components and multiple separation mechanisms; Fig. Figure 2 shows a schematic representation of a spacecraft component of the spacecraft; Fig. Figure 3 shows a schematic representation of the arrangement of a separation mechanism for connecting two spacecraft components of a spacecraft; Fig. Figure 4 shows a schematic sectional view of a first version of the spacecraft's separation mechanism; Fig. Figure 5 shows a schematic sectional view of the first version of the separation mechanism in a two-sided embodiment; Fig. Figure 6 shows a schematic sectional view of the first version of the separation mechanism in a one-sided embodiment; Fig. Figure 7 shows a schematic sectional view of the first version of the separation mechanism in an end-face embodiment; Fig. Figure 8 shows a schematic isometric representation of the first version of the separation mechanism; Fig. Figure 9 shows a schematic sectional view of the first version of the separation mechanism in a variant of the two-sided embodiment; Fig. Figure 10 shows a schematic sectional view of a first version of the separation mechanism in one variant; Fig. Figure 11 shows a schematic sectional view of the first version of the separation mechanism in a variant of the end-face embodiment; Fig. Figure 12 shows a schematic representation of a separation mechanism with multiple receiving chambers; Fig. Figure 13 shows a schematic representation of an isometric view of a second version of a separation mechanism in a two-sided embodiment; Fig. Figure 14 shows a schematic representation of an isometric view of a second version of a separation mechanism in the two-sided embodiment; Fig. Figure 15 shows a schematic representation of a cutaway isometric view of a second version of a separation mechanism in the two-sided embodiment; Fig. Figure 16 shows a schematic sectional view of a second version of a separation mechanism in the two-sided embodiment; Fig. Figure 17 shows a schematic representation of an isometric view of a second version of a separation mechanism in a one-sided embodiment; Fig. Figure 18 shows a schematic representation of an isometric view of a second version of a separation mechanism in the one-sided embodiment; Fig. Figure 19 shows a schematic representation of a cutaway isometric view of a second version of a separation mechanism in the one-sided embodiment; Fig. Figure 20 shows a schematic sectional view of a second version of a separation mechanism in the one-sided embodiment; Fig. Figure 21 shows a schematic representation of an isometric view of a second version of a separation mechanism in an end-face embodiment; Fig. Figure 22 shows a schematic representation of an isometric view of a second version of a separation mechanism in the end-face embodiment; Fig. Figure 23 shows a schematic representation of a cutaway isometric view of a second version of a separation mechanism in the end-face embodiment; Fig. Figure 24 shows a schematic sectional view of an isometric view of a second version of a separation mechanism in the end-face embodiment; Fig. Figure 25 shows a schematic representation of a separation mechanism in a top-side embodiment.

[0071] Fig. Figure 1 shows a schematic representation of a spacecraft 100. The spacecraft 100 consists of several spacecraft components 110, 120, two of which are shown with reference numerals as examples. These spacecraft components 110, 120 are connected by means of at least one separation mechanism 130. The spacecraft 100 has several separation mechanisms 130, which are shown as triangles as examples.

[0072] Fig. Figure 2 shows a schematic representation of one of the spacecraft components 110, 120 of the spacecraft 100, in particular the first spacecraft component 110, without the separation mechanism. This, for example, the first spacecraft component 110, has two opposing surfaces 113, 114. A recess 111 for receiving a separation mechanism is provided on at least one of the surfaces 113, 114. The recess 111 can also be configured as a through-hole extending from one surface 113 to the other surface 114 of the spacecraft component 110. The spacecraft component 110 also has at least one outer edge 112. A separation mechanism can also be arranged on the outer edge 112.

[0073] Fig. Figure 3 shows a schematic representation of an embodiment of a spacecraft 100 comprising a first spacecraft component 110 and a second spacecraft component 120. The first spacecraft component 110 and the second spacecraft component 120 are connected to each other via a separation mechanism 130. This passive separation mechanism 130 allows the first spacecraft component 110 to separate from the second spacecraft component 120, whereby at least parts of the separation mechanism 130 heat up during reentry of the spacecraft 100 into a planet's atmosphere and thus lose their structural integrity.

[0074] This occurs as early as possible, and therefore at the lowest possible separation temperature, whereby the separation temperature is higher than the maximum operating temperature. The maximum operating temperature is the temperature that is not exceeded during normal operation.

[0075] This separation temperature is also referred to as the melting temperature and is less than 400°C; in particular, it is provided that the melting temperature is less than 280°C, preferably less than 250°C, and most preferably less than 120°C. The melting temperature can also be greater than 100°C, in particular greater than 120°C, preferably greater than 190°C, and most preferably greater than 230°C.

[0076] Preferably, the melting temperature lies within the aforementioned limits. It has proven particularly advantageous for the melting temperature to be in a range between 120°C and 250°C. Further embodiments of the invention are described in more detail with the following figures.

[0077] The Fig. 4, Fig. 5, Fig. 6 to Fig. Figure 7 shows a first version of the separation mechanism 130. This separation mechanism 130 consists of a first structural element 140, a second structural element 150, and a third structural element which is designed as a connecting element 170. The different embodiments of the first version of the separation mechanism 130 are explained in more detail below.

[0078] Fig. Figure 4 shows a schematic representation of a separation mechanism 130 of the spacecraft (not shown here) for connection to a first spacecraft component (not shown here) and a second spacecraft component (not shown here). The separation mechanism 130 shown comprises a first structural element 140 and a second structural element 150, as well as a connecting element 170.

[0079] According to the embodiment shown here, the connecting element 170 consists of a shaft 175 and a head 179. The shaft 175 has an outer surface 172.

[0080] The second structural element 150 has a receiving space 174, which is designed to at least partially receive the shaft 175 of the connecting element 170. For this purpose, the second structural element 150 has a receiving space 174 with an inner surface 152, 162. Preferably, the receiving space 174 is cylindrical.

[0081] The inner surface 152 of the second structural element 150 and the outer surface 172 of the connecting element 170 are designed to correspond to each other and have structures for connection by means of positive locking and / or force locking. These structures are preferably threads.

[0082] The connection between the second structural element 150 and the connecting element 170 can be made directly or indirectly by using a threaded element 177. This threaded element 177 is also referred to as a wire thread insert and preferably consists of a wire made of copper or a copper alloy.

[0083] The connection between the second structural element 150 and the connecting element 170 is preferably designed to be detachable without damage. However, it is also possible that the connection between the second structural element 150 and the connecting element 170 prevents non-destructive detachment after its formation.

[0084] The second structural element 150 is surrounded by the first structural element 140. Preferably, the second structural element 150 is concentric with respect to the receiving space 174. In particular, the second structural element 150 is rotationally symmetrical. The first structural element can also be rotationally symmetrical. A particularly preferred embodiment provides that the second structural element 150 and the first structural element 140 are arranged concentrically with respect to each other.

[0085] Fig. Figure 5 shows a schematic representation of a separation mechanism 130 in a two-sided embodiment. The structural elements 140 and 150 are arranged in the spacecraft component 110 such that the receiving space 174 is accessible from both sides of the spacecraft component 110. This means that the connecting element (not shown) can be connected to the structural element 150 from both sides of the spacecraft component 110 or penetrate the receiving space 174 in the structural element 150. For this purpose, the spacecraft component 110 has a recess 111 that extends continuously from one surface 113 to the other surface 114.

[0086] Fig. Figure 6 shows a schematic representation of a separation mechanism 130 in a one-sided embodiment. The structural elements 140 and 150 are arranged in the spacecraft component 110 such that the receiving space 174 is accessible only from one side of the spacecraft component 110. This means that the connecting element (not shown) can only be connected to the structural element 150 from one side of the spacecraft component 110. For this purpose, the spacecraft component 110 has a recess 111 that extends continuously from one surface 113 to the other surface 114.

[0087] The separation mechanism 130, with its structural elements 140 and 150, is embedded in the recess 111 of the space component 110 and is bonded to it. This bond can be achieved by adhesive bonding. It is also possible that the space component 110 is machined to fit the separation mechanism 130 during manufacturing, particularly if the space component is made of a composite material.

[0088] The recess 111 in the spacecraft component 110 can be adapted almost exactly to the size of the separation mechanism 130. However, it is also possible that the recess 111 in the spacecraft component 110 is larger than the separation mechanism 130, which is indicated here by a dashed line. A filling material 115 is introduced between the edges of the recess 111 and the separation mechanism 130 to connect the separation mechanism 130 and the spacecraft component 110. This material hardens there and connects the separation mechanism 130 to the spacecraft component 110 at least materially and / or form-fittingly.

[0089] Fig. Figure 7 shows a schematic representation of a separation mechanism 130 in an end-face embodiment. The structural elements 140 and 150 are arranged on an outer edge 112 of the spacecraft component 110. The receiving space 174 is oriented with its longitudinal extent parallel to the plane of extension of the spacecraft component 110 and is accessible only from the end face of the spacecraft component 110. This means that the connecting element (not shown here) can only be connected to the structural element 150 from the end face of the spacecraft component 110. It is not absolutely necessary for the spacecraft component 110 to have a recess, as shown in the Fig. 5 and Fig. Figure 6 is shown. However, it is also possible to provide a recess on the outer edge 112 in which the separating mechanism 130 is positioned. For this purpose, reference is made to the Fig. 11 referred.

[0090] In the Fig. 5, Fig. 6 to Fig. In the embodiments of the separation mechanism 130 shown in Figure 7, the first structural element 140 has at least one flange-like projection 141, 142 at each of its two opposite distal ends in the longitudinal direction. This allows the spacecraft component 110 to be positively engaged by the structural element 140, thereby improving the connection between the structural element 140 and the spacecraft component 110. Preferably, the projection 141 of one distal end rests against one surface 113. Depending on the embodiment, the projection 142 of the opposite distal end rests against the opposite surface 114 or is arranged inside the structural element 110.

[0091] Especially in the embodiments of Fig. 5 and Fig. 7 The first structural element 140, with its flange-like projections 141, 142 on both sides, clamps the surfaces 113, 114 of the first spacecraft component 110 in the area of ​​the recess 111 and in the area of ​​the outer edge 112, respectively, thereby positively engaging and retaining the structural element 140 on the spacecraft component 110. The connection between the first structural element 140 and the spacecraft component 110 can additionally or alternatively be material-bonded.

[0092] At the in Fig. In the embodiment shown in Figure 6, one projection 141 is embedded in the surface 113 of the first spacecraft component 110; in particular, the projection 141 and the surface 113 are designed to be plane-flat to each other.

[0093] Fig. Figure 8 shows a schematic representation of an isometric view of the first version of a separation mechanism. The embodiment shown is similar to that in the Fig. 4, Fig. 5 to Fig. 6 shown.

[0094] Fig. Figure 9 shows a schematic sectional view of the first version of the separation mechanism 130 in a variant of the two-sided embodiment. The first structural element 140 of the separation mechanism 130 is designed in two parts. The first structural element 140 consists of an outer part 143, which is positioned facing the space component 110 and is, in particular, connected to it, and an inner part 144, which is positioned facing away from the space component 110 and is, in particular, connected to the second structural element 150. For connecting the inner part 144 and the outer part 143, these have positive locking elements 148, in particular threads or detent projections, on their mutually facing surfaces.

[0095] Fig. Figure 10 shows a schematic sectional view of a first version of the separation mechanism 130 in a variant analogous to the Fig. 4. In the recess 111 of the spacecraft component 110, a filling material is introduced between the spacecraft component 110 and the first structural element 140 of the separation mechanism 130. In its hardened state, this filling material creates a material-bonded and / or form-fit connection between the first structural element 140 and the spacecraft component 110. The flange-like projections 141, 142, which are also shown in the Fig. The 22 are shown, their positions indicated.

[0096] In this embodiment, the projections 141, 142 provided at opposite distal ends of the first structural element 140 are positioned flush with or contour-following the adjacent surfaces 113, 114 of the spacecraft component 110. This ensures that the separation mechanism 130 is mounted in the spacecraft component 110 without any protrusions relative to the surface 113, 114.

[0097] Fig. Figure 11 shows a schematic sectional view of the first version of the separation mechanism 130 in a variant of the end-face embodiment, analogous to the Fig. 7. In the recess 111 of the spacecraft component 110, a filling material is introduced between the spacecraft component 110 and the first structural element 140 of the separation mechanism 130, which, in its hardened state, creates a material-bonded and / or form-bonded connection between the first structural element 140 and the spacecraft component 110.

[0098] Fig. Figure 12 shows a schematic representation of an embodiment of the separation mechanism 130, in which several second structural elements 150, for example three second structural elements 150, are provided in a first structural element 140. In such an embodiment, the first structural element 140 is not rotationally symmetrical. This embodiment is possible with a two-sided separation mechanism, a one-sided separation mechanism, or an end-face separation mechanism.

[0099] The Fig. 13, Fig. 14, Fig. 15, Fig. 16, Fig. 17, Fig. 18, Fig. 19, Fig. 20, Fig. 21, Fig. 22, Fig. 23 to Fig. Figure 24 shows a second version of the separation mechanism 130. This separation mechanism 130 consists of a first structural element 140, a second structural element 150, a third structural element 160, and a connecting element 170. The connecting element is not shown in these figures, but is shown in the Fig. 3 and Fig. Figure 4 shows the different embodiments of the separation mechanism 130. These are explained in more detail below.

[0100] The Fig. 13, Fig. 14, Fig. 15 to Fig. Figure 16 shows a two-sided embodiment of the separation mechanism 130, wherein a central receiving space 174 for a connecting element 170 is accessible from two sides. The receiving space 174 is designed as a passage through the separation mechanism 130. The separation mechanism 130 is designed and configured to be arranged and fixed in the surface of a spacecraft component.

[0101] The Fig. 17, Fig. 18, Fig. 19 to Fig. Figure 20 shows a one-sided embodiment of the separation mechanism 130, wherein a central receiving space 174 for a connecting element 170 is accessible from only one side. The receiving space 174 is designed as a recess in the separation mechanism 130. The separation mechanism 130 is designed and configured to be arranged and fixed in the surface of a spacecraft component.

[0102] The Fig. 21, Fig. 22, Fig. 23 to Fig. Figure 24 shows an end-face embodiment of the separation mechanism 130, wherein a central receiving space 174 for a connecting element 170 is accessible from only one side. The receiving space 174 is designed as a recess in the separation mechanism 130. The separation mechanism 130 is designed and configured to be arranged and fixed on an edge of a spacecraft component.

[0103] The Fig. 13, Fig. 14, Fig. 17, Fig. 18, Fig. 21 and Fig. Figure 22 shows the embodiments of the separation mechanism 130 in isometric representation of two opposite head sides.

[0104] The Fig. 15, Fig. 19 and Fig. Figure 23 shows the embodiments of the separation mechanism 130 in a sectional isometric view.

[0105] The Fig. 16, Fig. 20 and Fig. Figure 24 shows the embodiments of the separation mechanism 130 in a sectional view.

[0106] In all embodiments of the second version of the separation mechanism 130, it consists of an externally arranged first structural element 140, a centrally arranged second structural element 150 and an internally arranged third structural element 160.

[0107] The first structural element 140 is designed for attachment to a spacecraft component, in particular for a permanent and non-destructively detachable connection to a spacecraft component. For this purpose, the first structural element 140 preferably has projections 141, 142 to increase the contact area with the spacecraft component (not shown here) and / or the filling material.

[0108] The third structural element 160 provides a recording room 174 for something not shown here, but which is included in the Fig. 3 and Fig. The third structural element 160 is connected to another spacecraft component by means of the connecting element 170, as shown in Figure 4. This connection can be non-destructive or non-destructive.

[0109] The first structural element 140 and also the third structural element 160 consist of a shaft 145, 165 and at least one, preferably two, flange sections 146, 147, 166, 167. The shafts 145, 165 and the flange sections 146, 147, 166, 167 define a cavity 158.

[0110] The second structural element 150 is arranged in the cavity, in particular formed, preferably pre-formed.

[0111] For the initial shaping of the second structural element 150 into the cavity 158, at least one opening 149 is provided in the separation mechanism 130, in particular in the first structural element 140. In the example shown here, there are four openings 149.

[0112] This at least one opening 149 is preferably provided on one of the end faces of the separating mechanism 130. In particular, all openings of a separating mechanism 130 are provided on the same end face. In the example shown, the openings 149 are arranged on the end face opposite the end face with the preferred access for the connecting element 170.

[0113] The cavity 158, particularly for the initial forming of the second structural element 150, is bounded at its head end by flange sections 146, 147, 166, 167. Both the first structural element 140 and the third structural element 160 have flange sections 146, 147. The flange sections 146, 166, 147, 167 are arranged in pairs facing each other. This enlarges the cavity 158 for the second structural element 150. As the size of the second structural element 150 increases, separation upon re-entry is facilitated.

[0114] In the example shown, the cavity 158 has a special configuration. A flange section 146 of the first structural element 140 and a flange section 166 of the third structural element are wedge-shaped. This means that the at least one flange section 146, 166 has a surface oriented towards the cavity 158. This surface is inclined towards the orientation of the cavity 158 and / or towards the longitudinal direction of the receiving space 174; preferably, it forms an angle between 20° and 70° with the orientation of the cavity 158 and / or the longitudinal direction of the receiving space 174.

[0115] One aspect is that the surfaces of the wedge-shaped flange sections 146, 166 are aligned approximately parallel to each other, in particular enclosing an angle of less than 30°, preferably an angle of less than 10°, to each other.

[0116] The design of the flange sections 146 and 166 as wedge surfaces improves the separation properties upon reentry. The first structural element 140 and the third structural element 160 cannot become jammed or interlocked. Furthermore, this design improves the force transmission between the structural elements 140, 150, and 160, at least until the structural element with the lower melting temperature, in this case the second structural element 150, reaches its melting temperature. In particular, forces in the axial direction can be transmitted as compressive forces, and the occurrence of shear forces is reduced.

[0117] As shown in the example, the two wedge-shaped flange sections 146, 166 are positioned on opposite head sides of the separation mechanism 130. Preferably, in a connection position of the structural elements 140, 150, 160 with respect to a separation direction 133, the wedge-shaped flange section 146 of the first structural element 140 is arranged further forward, i.e., closer to a separation opening 135 formed upon re-entry, than the wedge-shaped flange section 166 of the third structural element 160.

[0118] The third structural element 160 has a receiving space 174, which is designed to at least partially receive the shaft 175 of the connecting element 170. For this purpose, the third structural element 160 has a receiving space 174 with an inner surface 162. Preferably, the receiving space 174 is cylindrical.

[0119] The inner surface 162 of the third structural element 160 and the outer surface 172 of the connecting element 170 are designed to correspond to each other and have structures for connection by means of positive locking and / or force locking. These structures are preferably implemented as a thread.

[0120] The connection between the third structural element 160 and the connecting element 170 can be made directly or indirectly by using a threaded element 177. This threaded element 177 is also referred to as a wire thread insert and preferably consists of a wire made of copper or a copper alloy.

[0121] The connection between the third structural element 160 and the connecting element 170 is preferably designed to be detachable without damage. However, it is also possible that the connection between the third structural element 160 and the connecting element 170 prevents non-destructive detachment after its formation.

[0122] The second structural element 150 is surrounded by the first structural element 140. Preferably, the second structural element 150 is designed with a coarse structure concentric to the receiving space 174. In particular, the coarse structure of the second structural element 150 is rotationally symmetrical. The first structural element can also be rotationally symmetrical. A particularly preferred embodiment provides that the second structural element 150 and the first structural element 140 are arranged concentrically to each other.

[0123] The third structural element 160 is surrounded by the second structural element 150. Preferably, the third structural element 160 is designed concentrically with respect to the receiving space 174 in its basic structure. In particular, the third structural element 160 is designed rotationally symmetrically in its basic structure. A particularly preferred embodiment provides that the third structural element 160 and the second structural element 150 are arranged concentrically with respect to each other.

[0124] Fig.Figure 25 shows a schematic representation of a separation mechanism 130 in a surface-mounted embodiment. In this embodiment, the separation mechanism 130 is arranged and attached to the surface 113 of a space component 110 with the first structural element 140. The connection between the first structural element 140 and the space component 110 can be material-bonded, for example by adhesive bonding, and / or force-fit, for example by screw, and / or form-fit, for example by rivet. Otherwise, the separation mechanism 130 is identical to the embodiments described above. Reference sign 100 spacecraft 110 First spacecraft component 111 Exclusion 112 outer edge 113 surface 114 surface 115 Filling mass 120 Second spacecraft component 130 Separation mechanism 133 Separation direction 135 Separation opening 140 first structural element 141 lead 142 lead 143 Exterior part 144 Inner part 145 shaft 146 Flange section 147 Flange section 148 positive locking elements 149 Opening 150 second structural element 152 inner surface area 158 cavity 160 third structural element 162 inner surface area 165 shaft 166 Flange section 167 Flange section 170 fasteners 172 outer surface area 174 Recording room 175 shaft 177 Threaded element 179 heads

Claims

[1] Spacecraft (100) comprising: - a first spacecraft component (110), - a second spacecraft component (120), - a passive separation mechanism (130) for separating the first spacecraft component (110) from the second spacecraft component (120), wherein the two spacecraft components (110, 120) are connected to each other in a connected state, the passive separation mechanism (130) comprises: - a first structural element (140) which is attached to the first spacecraft component (110), - a second structural element (150) which is attached to the first structural element (140), - a third structural element (160) for mechanically connecting the first spacecraft component (110) with the second spacecraft component (120), wherein in the connected state the third structural element (160) and the second structural element (150) are mechanically connected to each other, characterized by , that - the third structural element (160) and the second structural element (150) are designed to form either a detachable mechanical connection, in particular a force-fit and / or form-fit connection, with the third structural element being designed as a connecting element (170), or - the third structural element (160) for a detachable mechanical connection, in particular for a force-fit and / or form-fit connection, is formed with the connecting element (170), and wherein the third structural element (160) has a shaft (165), wherein in the connected state the shaft (165) is at least partially enclosed by the second structural element (150), and wherein at least two of the structural elements (140, 150, 160) consist of different materials, wherein at least one of the structural elements (140, 150, 160), in particular the first structural element (140) or the second structural element (150), consists of a material whose melting temperature is lower than the melting temperature of the material from which at least one of the other structural elements (140, 150, 160), in particular the first structural element (140) and / or the third structural element (160), consists, wherein at least the material with the lower melting temperature has a melting temperature which is lower than the temperature acting on the structural elements (140, 150, 160) during a reentry of the spacecraft (100) into the atmosphere of a planet. [2] Spacecraft (100) according to the preceding claim, wherein the passive separation mechanism (130) is configured to transfer heat from the first spacecraft component (110) and the second spacecraft component (120) to the structural elements (140, 150, 160) so that, upon reentry, the structural elements (140, 150, 160), in particular the first structural element (140) or the second structural element (150), are heated at least to such an extent that at least one of the structural elements (140, 150, 160), in particular the first structural element (140) or the second structural element (150), loses its solid state and / or is melted. [3] Spacecraft (100) according to one of the two preceding claims, wherein in an assembled state of the separation mechanism (130) the first structural element (140) and the third structural element (160) enclose a cavity (158) which is filled with the second structural element (150). [4] Spacecraft (100) according to one of the preceding claims, wherein the first structural element (140) and the third structural element (160) each have a shaft (145, 165) which is bounded on both sides by at least one flange section (146, 147, 166, 167), preferably by two flange sections (146, 147, 166, 167). [5] Spacecraft (100) according to one of the preceding claims, wherein the flange sections (146, 147, 166, 167) define a cavity (158) in a longitudinal direction of the separation mechanism (130) in which the second structural element (150) is received and / or wherein the shafts (145, 165) define a cavity (158) transversely to a longitudinal direction of the separation mechanism (130) in which the second structural element (150) is received. [6] Spacecraft (100) according to one of the preceding claims, wherein at least one of the flange sections (146, 147, 166, 167), preferably a flange section (146, 147) of the first structural element (140) and a flange section (166, 167) of the third structural element (160) is wedge-shaped in the longitudinal direction, in particular wherein the surface of the wedge-shaped flange section (146, 147, 166, 167) has an angle to the longitudinal direction which is in the range between 20 degrees and 70 degrees. [7] Spacecraft (100) according to one of the preceding claims, wherein in the case of at least one of the structural elements (140, 150, 160) with the material having the lower melting temperature, in particular in the case of the second structural element (150), the melting temperature is less than 400°C, in particular greater than 100°C and / or less than 250°C, whereby the material with the lower melting temperature, in particular the first structural element (140) or the second structural element (150), can be heated and / or melted to the point of loss of strength and structural integrity by frictional heat which occurs when the first spacecraft component (110) and the second spacecraft component (120) enter an atmosphere of a planet. [8] Spacecraft (100) according to one of the preceding claims, wherein the second structural element (150) provides a cylindrical receiving space (174) which is designed to receive at least one section of the shaft (165) of the third structural element (160) designed as a connecting element (170), wherein the second structural element (150) has an inner surface (152) and the shaft (175) of the connecting element (170) is cylindrical and has an outer surface (172), and / or wherein the third structural element (160) provides a cylindrical receiving space (174) which is designed to receive at least one section of the shaft (175) of a connecting element (170), wherein the third structural element (160) has an inner surface (162) and the shaft (175) of the connecting element (170) is cylindrical and has an outer surface (172). [9] Spacecraft (100) according to the preceding claim, wherein on the inner surface (152, 162) of the second structural element (150) or the third structural element (160) on the one hand and on the outer surface (172) of the connecting element (170) on the other hand, structures are provided which enable a force-fit and / or form-fit connection of connecting element (170) and second structural element (150) or third structural element (160), in particular by applying a relative rotational movement. [10] Spacecraft (100) according to one of the preceding claims, wherein the first spacecraft component (110) has a recess (111) or an outer edge (112) to which the first structural element (140) is attached, in particular permanently, preferably by a form-fit and / or by a material-fit. [11] Spacecraft (100) according to one of the preceding claims, wherein the first structural element (140) surrounds the second structural element (150) perpendicular to the longitudinal extent of the receiving space (174) of the second structural element (150). [12] Spacecraft (100) according to one of the preceding claims, wherein the first structural element (140) has at least one flange-like projection (141, 142) in the longitudinal direction at at least one distal end which abuts at least one of the surfaces (113, 114) of the first spacecraft component (110). [13] Spacecraft (100) according to the preceding claims, wherein the first structural element (140) has at least one flange-like projection (141, 142) at each of its two opposite distal ends in the longitudinal direction, wherein the projection (141) of one distal end abuts one surface (113) and the projection (142) of the opposite distal end abuts the opposite surface (114). [14] Spacecraft (100) according to one of the two preceding claims, wherein the first structural element (140) with its flange-like projections (141, 142) clamps or encompasses the surfaces (113, 114) of the first spacecraft component (110) in the area of ​​the recess (111) or in the area of ​​the outer edge (112). [15] Spacecraft (100) according to one of the preceding claims, wherein the first structural element (140) consists of an outer part (143) and an inner part (144), wherein the outer part is positioned facing the spacecraft component (110) and is in particular connected to it, and wherein the inner part (144) is positioned away from the spacecraft component (110) and is in particular connected to the second structural element (150). [16] Separation mechanism (130) for use in a spacecraft (100), wherein the separation mechanism (130) is designed as a passive separation mechanism (130) and comprises the following: - a first structural element (140), for connection with a first spacecraft component (110) of the spacecraft (100), - a second structural element (150) which is attached to the first structural element (140), - a third structural element (160) which is designed and constructed for a detachable mechanical connection, in particular for a force-fit and / or form-fit connection, with the second structural element (150), wherein the first spacecraft component (110) can be connected to the second spacecraft component (120) by means of the third structural element (160), characterized by , that the third structural element (160) is designed and constructed as a connecting element (170) for a detachable mechanical connection, in particular for a force-fit and / or form-fit connection, with the second structural element (150). or the third structural element (160) and a connecting means (170) for a cooperating detachable mechanical connection, in particular for a force-locking and / or form-locking connection, are arranged and designed together, and wherein the third structural element (160) has a shaft (165), wherein in the connected state the shaft (165) is at least partially enclosed by the second structural element (150), wherein at least one of the three structural elements (140, 150, 160) consists of a material whose melting temperature is lower than the melting temperature of the material from which the other two of the three structural elements (140, 150, 160) consist, wherein the melting temperature of the material with the lower melting temperature is lower than the temperature acting on the spacecraft (100) during re-entry into the atmosphere of a planet, in particular the separation mechanism (130) is designed with the features according to one of claims 3 to 9.

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