BACKGROUND

Technical Field

This application relates to exothermic reaction aided spacecraft demise during re-entry. In particular, the application relates to space vehicle elements configured to be destroyed during re-entry of the spacecraft (space vehicle) into the atmosphere and to methods of manufacturing such space vehicle elements.

Description of the Related Art

With a view of future space activities, it is desirable to preserve Earth's orbital environment as a safe zone, free of debris. To this end, active efforts are required to remove spacecraft (e.g., satellites or launch vehicles) from orbit. This can be done by means of a controlled re-entry, actively de-orbiting the spacecraft and controlling the area where any debris may end up avoiding populated areas, or uncontrolled re-entry, in which case the risk to people needs to be controlled and minimized. In order to allow for uncontrolled re-entry, spacecraft need to be designed in a way that makes them almost completely demise during re-entry.

Design for Demise (D4D) is a hot topic at the moment as satellite manufacturers try to comply to the recently imposed Space Debris Mitigation requirements. These requirements state that there must be a less than 1 in 10,000 chance of someone being hit by falling space debris. In view of this aim, the space systems industry has been looking at design alternatives that would cause the satellite to disintegrate (demise) during the re-entry into the atmosphere. This demise option presents advantages with respect to other more conventional solutions. For instance, the option of a controlled de-orbit into a safe ocean splash-down could cost as much as 4 or 5 times the propellant needed on board with the consequent increase in satellite size, mass and cost, potentially even requiring a more powerful launcher.

Critical elements (space vehicle elements) in satellite design in the context of demise include titanium propellant tanks, reaction wheels, optical payloads, and balancing masses, for example. Demise of these elements during re-entry could be improved by appropriately adjusting the main factors influencing the survivability of each element. These factors include the material and mass of the element, its geometry or wall thickness, as well as the heat exposure and potential shielding effects induced by other elements of the spacecraft. The D4D approach proposes different solutions ranging from changing the material that critical items are made of, to relocating items to places where they receive more heating effect earlier in the re-entry and even triggering a partial break-up of the satellite structure during re-entry to aid demise. As recent work has shown, doing so can provide real improvements, but does not yet allow to attain the above overall goal with regard to demisability.

Moreover, the current efforts to design for demise require significant changes to existing spacecraft equipment. This can lead to costly development and qualification programs, dedicated to each specific type of equipment. Furthermore, it may be required to make concessions on the performance of the unit. For example, an avenue investigated for reaction wheels is to use smaller wheels spinning at higher speeds. Whereas this does somewhat improve the demisability, it has an impact on performance parameters such as the exported microvibration and the power consumption.

As noted above, recent work has shown that the currently explored methods for making spacecraft equipment more demisable may not be sufficient. Specifically, for reaction wheels, it has turned out that the ball bearing assembly for flywheel suspension does survive, even for reaction wheel designs optimized for demisability. Thus, more drastic changes may be required in order to fully demise this type of equipment, which could imply moving away from design solutions with decades of heritage.

It is also clear that some units use materials which are hard to demise, for which no other option exists due to other design constraints. Mechanisms often use steel ball bearings, housed in steel or titanium parts. These materials are necessary to provide the required mechanical strength, stiffness and lifetime performance of these units and cannot be replaced. Similarly, optical payloads also often use materials which are hard to demise, and cannot be replaced as they are required to meet the desired performance, for example thermo-elastic stability.

Another approach to improve demisability could be to try to ensure early break-up of the spacecraft structure. Current methods to help break up the spacecraft structure may be insufficient, however. For example, one design methodology could be to make use of Shape Memory Alloys (SMA) to break structural joints. This is a fully passive system, activated by the heat generated during re-entry. The Technology Readiness Level (TRL) for this solution is currently still low, and the impact on structural performance of the spacecraft needs to be assessed in greater detail. Moreover, SMA actuator solutions do not address the issue of harness and propulsion pipework which may hamper a controlled break-up of the structure.

Thus, there is a need for a new and improved techniques to demise spacecraft equipment. There is particular need for such techniques that do not impose additional design constraints on space vehicle elements.

BRIEF SUMMARY

In view of some or all of these needs, the present disclosure proposes a space vehicle element and a method of manufacturing a space vehicle element, having the features of the respective independent claims.

An aspect of the disclosure relates to a space vehicle element configured to be at least partially destroyed or demised during re-entry of the space vehicle (spacecraft) into the atmosphere. Here, at least partial destruction or demise is understood to include, for example, full destruction, partial destruction, or severing the space vehicle element into two or more parts, possibly followed by destruction or partial destruction of the individual parts. In particular, it may be understood that any remainder of the space vehicle element upon impact on the ground may have a reduced impact energy, with an aim to reduce it below 15J per impacting part (the amount of energy considered critical in case of collision with a human being). The space vehicle may be a satellite, for example. The space vehicle element may include a heat generating part. The heat generating part may include (or in some cases consist of) a metallo-thermal composition for providing additional heat during re-entry of the space vehicle into the atmosphere. The additional heat may be provided by an exothermic reaction of the metallo-thermal composition. Destruction of the space vehicle element may be expedited by the additional heat provided by the heat generating part. The heat generating part may be integrated, fully or at least partially, within the space vehicle element. Accordingly, at least a portion of the heat generating part may be arranged inside of the space vehicle element, e.g., away from a surface of the space vehicle element. Alternatively, the heat generating part may be arranged to surround (e.g., encircle, enclose), fully or at least partially, a portion of the space vehicle element. The metallo-thermal composition may comprise a mixture of a metal (e.g., in powder form) and a metal oxide (e.g., in powder form), for example aluminum and iron oxide, and possibly a binder. As such, the metallo-thermal composition may be equivalent to a pyrotechnic composition of metal powder. One example of a metallo-thermal composition is thermite.

Configured as described above, the proposed solution offers a new technique to demise spacecraft equipment (space vehicle elements), by adding energy to the system in order to help demise elements of existing designs, although incorporating it in new designs is also perfectly possible and appropriate. This is achieved by providing a metallo-thermal composition that is capable of an exothermic reaction once ignited, at least partly integrated with the space vehicle element or at least partially surrounding a portion of the space vehicle element. The heat generating part can correspond to a cartridge of thermite or a thermite-like substance. The benefit of this approach is that the impact on the design should be minimal. The mass and volume taken up by such heat generating part should be sufficiently small to minimize the impact on the design of the equipment, while providing sufficient energy to ensure the demise. The actuation of the heat generating parts can be completely passive, as ignition relies on the heat generated during re-entry.

In some embodiments, the heat generating part may be arranged (e.g., inserted) in a recess or cavity provided in the space vehicle element. The heat generating part may be arranged in a hollow, clearance, opening, etc., of the space vehicle element. Accordingly, the footprint of the space vehicle element is not altered. Moreover, heat transfer from the heat generating part to the space vehicle element during re-entry is optimized, while minimizing the additional material added, thus improving demise of the space vehicle element.

In some embodiments, the heat generating part may be arranged in contact (including in thermal contact) with a connecting section between a first part of the space vehicle element and a second part of the space vehicle element. Then, the additional heat provided by the heat generating part may expedite severing the connecting section during re-entry of the space vehicle into the atmosphere. The connecting section may be a joint or an interface (including the regular/typical mechanical mounting interfaces of space vehicle equipment), for example. Specifically, the connecting section may be a wire harness, a propulsion pipe, or a heat pipe, etc. Severing the connecting section may separate the first part and the second part or at least contribute to severing the first part and the second part. The connecting section may relate to a structural connection, i.e., the mechanical interface between the element and the space vehicle, or to a secondary connection, such as harness, pipework, etc. The latter are not meant to take mechanical loads under normal circumstances, but can become the primary connection during re-entry if the structural connection has already failed. Notably, heat generating parts can be simply added to any existing type of structural joint with little impact on the overall design.

In some embodiments, the heat generating part may be shaped to at least partially enclose the connecting section between the first part of the space vehicle element and the second part of the space vehicle element. The heat generating part may have annular or semi-annular shape, for example. For instance, the heat generating part may be arranged to sever the space vehicle element from its connector (interface) that couples the space vehicle element to a harness or to pipework. By at least partially enclosing the connecting section, heat transfer from the heat generating part to the connecting section can be optimized and the connecting section can be severed in a predictable (e.g., controlled) and reliable manner. Also, such heat generating part can be applied to, for example, harness and pipework in a simple manner, such as by clamping.

In some embodiments, the second part may be a fixation part for fixing the space vehicle element to a portion of the space vehicle. Accordingly, the space vehicle element can be reliably and predictably severed from the space vehicle during re-entry.

In some embodiments, the space vehicle element may further include a second heat generating part. The second heat generating part may include a second metallo-thermal composition for providing additional heat during re-entry of the space vehicle into the atmosphere by an exothermic reaction of the second metallo-thermal composition. The second heat generating part may be arranged in contact (including in thermal contact) with one of the first part or the second part of the space vehicle element, for expediting the destruction of the one of the first part or the second part of the space vehicle element by the additional heat provided by the second heat generating part. In some embodiments, the space vehicle element may comprise more than the first and second heat generating parts. Accordingly, a large space vehicle element can be broken up during re-entry and the resulting part of the space vehicle element can be individually demised, thereby improving the overall demise of the space vehicle element.

In some embodiments, the second heat generating part may be configured such that the exothermic reaction of the second metallo-thermal composition is activated after the exothermic reaction of the first metallo-thermal composition is activated. Thereby, the second heat generating part may provide the additional heat for expediting the destruction of the one of the first part or the second part of the space vehicle element after the connecting section between the first part and the second part of the space vehicle element has been severed.

In some embodiments, the heat generating part may be arranged in a vicinity of a structural connection between the space vehicle element and the space vehicle. Thereby, severing the structural connection during re-entry of the space vehicle into the atmosphere may be expedited by the additional heat provided by the heat generating part. In some implementations, the heat generating part may be arranged in such way as to break the structural connection between the space vehicle element and the space vehicle. To this end, the heat generating part may be shaped to at least partially enclose the structural connection.

In some embodiments, the space vehicle element may further include a second heat generating part. The second heat generating part may include a second metallo-thermal composition for providing additional heat during re-entry of the space vehicle into the atmosphere by an exothermic reaction of the second metallo-thermal composition. Destruction of the space vehicle element may be expedited by the additional heat provided by the second heat generating part. The first and second heat generating parts may be configured such that the exothermic reactions of their respective metallo-thermal compositions are activated in a predetermined sequence and/or with predetermined relative timing (e.g., with a predetermined sequence between respective activations). In some embodiments, the space vehicle element may comprise more than two heat generating parts and these more than two heat generating parts may be arranged/configured to ignite in a predetermined sequence. Thereby, breakup of the space vehicle element into pieces and individual demise of those pieces during re-entry can be effected according to a predetermined, optimized scheme, ensuring optimum demise of the space vehicle element during re-entry.

In some embodiments, the heat generating part may be configured such that the metallo-thermal composition is ignited by a thermal flux created by atmospheric friction during re-entry of the space vehicle into the atmosphere. The heat generated during the re-entry will be conducted through the space vehicle element and will eventually create a sufficiently high temperature and heat density to achieve ignition of the metallo-thermal composition. The heat generating part may be specifically arranged and configured to achieve this aim. That is, the heat generating part may be designed (e.g., with regard to the composition of the metallo-thermal composition), partly in dependence on an arrangement of the heat generating part within the space vehicle (e.g., in dependence on its distance from a surface of the space vehicle element, etc.). This avoids the need for fuses or other means for actively igniting the metallo-thermal composition of the heat generating part.

In some embodiments, the space vehicle element may further include a fuse for igniting the metallo-thermal composition. The fuse may be thermally coupled to the heat generating part. The fuse may be activated by a thermal flux created by atmospheric friction during re-entry of the space vehicle into the atmosphere. Accordingly, the metallo-thermal composition can be reliably ignited during re-entry of the space vehicle into the atmosphere.

In some embodiments, the fuse may include a heat conducting part for transferring heat generated by the thermal flux to the heat generating part, for activating the exothermic reaction of the metallo-thermal composition. Such a fuse may be referred to as a passive fuse. The (passive) fuse may be or comprise a copper element (e.g., a rod), for example, or an element formed of any other material with good heat conductivity and sufficiently high melting/disintegration temperature (e.g., higher than the ignition temperature of the metallo-thermal composition it is supposed to ignite). One end of the (passive) fuse may reach closer towards a surface of the space vehicle than the heat generating part. The heat conducting element can be configured in such a way to optimize the ignition, for example by maximizing the contact area between the heat conducting element and the metallo-thermal composition (i.e., the heat generating part). Thereby, the metallo-thermal composition can be ignited in a simple and efficient manner.

In some embodiments, the fuse may include two separated chambers having contained therein respective agents (e.g., chemical agents). The fuse may further include a dividing section (e.g., a dividing wall) arranged between and separating the chambers. The agents may mix if the dividing section is made permeable or destroyed. The two agents may be chosen to spontaneously ignite and undergo an exothermic reaction when being brought into contact with each other. The agents may undergo a so-called hypergolic reaction. The exothermic reaction may generate sufficient heat/energy to start the reaction of the metallo-thermal composition or ignite the metallo-thermal composition. The fuse may be configured such that the dividing section is made permeable or destroyed by the heat generated by the thermal flux. The dividing section may be configured to be made permeable or destroyed at a predetermined, tunable temperature (e.g., tunable by choice of material, thickness, etc., of the dividing section). For example, the dividing section may have a predetermined melting temperature, which may be lower than the ignition temperature of the metallo-thermal composition. It is understood that the casing of the two separated chambers other than the dividing section is configured (e.g., by choice of material and/or thickness) to not be made permeable or destroyed by the thermal flux until after the dividing section is made permeable, so that the two agents can mix and react with each other. Using such a fuse, sufficient heat for igniting the metallo-thermal composition can be provided and the metallo-thermal composition can be ignited in a reliable and efficient manner at a point where the temperature of the space vehicle element and the metallo-thermal composition is still lower than the ignition temperature of the metallo-thermal composition.

In some embodiments, the fuse may include a pyrotechnic or metallo-thermal composition coating. Optionally, the fuse may further include a metal wire, at least partially coated by the pyrotechnic or metallo-thermal composition. In some sense, this fuse may be seen as resembling a sparkler. The metal wire can be formed as needed for the specific application, for example to reach from the heat generating part towards an outer surface of the space vehicle where the heat generated by atmospheric friction is sufficient to ignite the coating. Then, the coating would burn up along the metal wire, eventually igniting the metallo-thermal composition of the heat generating part. Using such a fuse, the metallo-thermal composition can be ignited in a simple and efficient manner at a point where the temperature of the space vehicle element and the metallo-thermal composition is still lower than the ignition temperature of the metallo-thermal composition.

In some embodiments, the space vehicle element may be a reaction wheel that is usable for maneuvering of the space vehicle. The reaction wheel may include a ball bearing assembly. The heat generating part may be arranged along a center of the ball bearing assembly. In this implementation, demise of the reaction wheel, which is a massive component of any space vehicle, can be reliably effected. Since the heat generating part is arranged along the center of the ball bearing assembly, no additional space is required for the heat generating part, and heat transfer to the ball bearing assembly is optimized.

In some embodiments, the space vehicle element may be a solar array drive mechanism for driving a solar array of the space vehicle. The heat generating part may be arranged in a rotor shaft of the solar array drive mechanism or between stiffening ribs of a structural component of the solar array drive mechanism. The structural component may be an adapter, for example. In this implementation, demise of the solar array drive mechanism can be reliably effected. Since the heat generating part is arranged in the rotor shaft of the solar array drive mechanism or between stiffening ribs of the structural component of the solar array drive mechanism, no additional space is required for the heat generating part. Also, heat transfer to the solar array drive mechanism can be optimized.

Another aspect of the disclosure relates to a method of manufacturing a space vehicle element. The space vehicle element may be configured to be destroyed during re-entry of the space vehicle into the atmosphere. The method may include providing a heat generating part comprising a metallo-thermal composition for providing additional heat during re-entry of the space vehicle into the atmosphere by an exothermic reaction of the metallo-thermal composition. Destruction of the space vehicle element may be expedited by the additional heat provided by the heat generating part. The method may further include integrating the heat generating part within the space vehicle element.

It will be appreciated that apparatus features and method steps may be interchanged in many ways. In particular, the details of the disclosed apparatus (e.g., space vehicle element) can be realized by the corresponding method of manufacturing the apparatus, and vice versa, as the skilled person will appreciate. Moreover, any of the above statements made with respect to the apparatus are understood to likewise apply to the corresponding manufacturing method, and vice versa.

