Patent Description:
In particular, the invention allows a first spacecraft component to detach from a second spacecraft component upon occurrence of a defined condition, especially upon earth atmospheric re-entry.

Earth atmospheric re-entry is the movement of an object from outer space through the gases of earth's atmosphere. Aerodynamic heating and atmospheric drag are two phenomena experienced by the object during atmospheric re-entry, which can cause its loss of mass and its fragmentation into smaller objects. Therefore, the larger is the exposure of an object to aerodynamic heating and atmospheric drag, the larger are the chances that the object demises in the atmosphere and does not reach earth's surface.

Spacecraft are often too large to completely demise during their earth atmospheric re-entry phase at their end of life. To avoid associated debris endangering people and property on earth, large spacecraft are often de-orbited and brought down to earth in a controlled way using propulsion systems, which entails significant additional costs for the overall mission and increased debris on earth. As an alternative approach, the spacecraft can be separated or broken-up into smaller pieces, early during the atmospheric re-entry phase, which would ensure its full demise.

It is known to separate spacecraft components with the use of active and/or passive separation mechanisms. Active separation mechanisms are generally complex and require the use of an electrical signal to initiate the separation functionality while their performance can degrade over time, which impacts overall system safety and disposal reliability. Passive separation mechanisms make use of natural phenomena (e.g. aerodynamic heating) seen by the spacecraft during its lifecycle in order to initiate the separation functionality.

An example for a two-part passive separation mechanism having passively separable connecting surfaces is disclosed in <CIT>. This separation mechanism shows flat geometries for the connecting surfaces.

An example for a multi-part passive separation mechanism comprising more than two parts having passively separable connecting surfaces is shown in <CIT>. It utilizes curved and flat geometries for the connecting surfaces and uses multiple joining materials.

Each of said known two-part and multi-part passive separation mechanisms requires that components of the separation mechanism for joining two spacecraft components can be mounted and dismounted repeatedly from the spacecraft components.

<CIT> discloses a Hold-down and Release Mechanisms (HRMs), which securely and strongly connect and hold down the deployable appendages of a space system during launch, and which release said deployable appendages upon reception of a command signal. The system includes a segmented tie rod that extends mainly parallelly to axis z and comprises a first segment and a second segment, that are soldered to one another by means of a solder alloy forming a solder joint between said first and second segments. The first and second segments have flat end surfaces connected together via the solder joint. An extractor spring is provided to facilitate the separation of between the first and second segments when the solder joint is molten.

<CIT> discloses passive devices designed to facilitate demise of space systems during re-entry into the Earth's atmosphere.

<CIT> discloses a heat discharger suitable for application in heat pipes intended for artificial space satellites. <CIT> discloses a connection via two conical mating surfaces either via a weld or via an adhesive.

It is an object of the present application to provide a passive separation mechanism for spacecraft components for which design freedom, compatibility and easiness for system implementation as well as operational capabilities are enhanced while manufacturing complexity and costs as well as performance and reliability risks are reduced.

This object is solved by the separation mechanism for passive separation of a first spacecraft component from a second spacecraft component as defined in claim <NUM> and the spacecraft comprising said passive separation mechanism as defined in claim <NUM>.

The separation mechanism for passive separation of a first spacecraft component from a second spacecraft component of claim <NUM> comprises first and second structural elements and joining means for mechanically joining the first and second structural elements together. The first structural element is adapted to be mounted to the first spacecraft component and has a receiving hole with an inner first connecting surface. The second structural element is adapted to be mounted to the second spacecraft component and is further adapted to be at least in part received within the receiving hole of the first structural element. The second structural element has an outer second connecting surface facing the inner first connecting surface when the second structural element is at least in part received within the receiving hole of the first structural element. The joining means for mechanically joining the first and second structural elements together via their first and second connecting surfaces is, in a joint state of the passive separation mechanism, sandwiched between the inner first connecting surface and the outer second connecting surface and connects these surfaces together when the second structural element is at least in part received within the receiving hole of the first structural element. The joining means is solder meltable at a temperature in the range of <NUM> to <NUM>. By choosing a solder having such a melting temperature, the first and second structural elements separate from each other when a temperature in the range of <NUM> to <NUM> is experienced by the separation mechanism. In the joint state of the passive separation mechanism, the joining means can be referred to as a solder joint.

A spacecraft usually is exposed to a temperature higher than <NUM>°, in particular higher than <NUM> during the early stage of the earth atmospheric re-entry phase. Thus, a separation of the passive separation mechanism of claim <NUM> can be triggered by exposure to the frictional heat experienced during earth atmospheric re-entry phase, especially by exposure to the frictional heat experienced by a spacecraft already during an early phase of earth atmospheric re-entry.

During the earth atmospheric re-entry phase, the spacecraft and, thus, its components, including the first and second components joint by the claimed passive separation mechanism, are heated. By a full or even already by a partial heat transfer from the components to the separation mechanism, the melting of the soldered connection, i.e., the solder joint, between the two structural elements of the separation mechanism is achieved. This entails the separation of the two structural elements of the separation mechanism and, thus, the separation of the first and second spacecraft components to which the first and second structural elements can be mounted respectively, which results in the spacecraft breaking into smaller pieces which can fully demise in the atmosphere.

Accordingly, the claimed separation mechanism works entirely passive and relies on the melting of a (metallic) alloy during the earth atmospheric re-entry phase. Therefore, no impact on spacecraft disposal reliability will occur during a mission of a spacecraft provided with the claimed separation mechanism.

The first and second connecting surfaces are of corresponding conical shape. Then, one can think of the first structural element to have the shape of a cone trunk, wherein the curved surface area of the cone trunk forms the second connecting surface.

In order to ensure a safe transfer of mechanical loads typically seen during operation by spacecraft components, the joining means and the shape and size of the connecting surfaces are to be chosen correspondingly to provide sufficient connecting strength for the joint of the first and second structural elements of the passive separation mechanism and, thus, for the first and second spacecraft components, throughout all operational temperature levels of the spacecraft mission prior to the spacecraft disposal phase, i.e., its earth atmospheric re-entry phase.

To ensure sufficient and uniform properties of the solder joint separation mechanism, i.e., strength during life time and/or mission of a spacecraft with which the separation mechanism can be used, and uniform melting of the solder at the end of life time and/or mission, the matching connecting surfaces of the structural elements are preferably aligned in the joint state or for realizing the joint state of the separation mechanism such that the (resulting) solder joint has a homogenous thickness.

The first structural element and the second structural element can be made of aluminium or aluminium alloys, such as <NUM>, <NUM>, <NUM> and <NUM> aluminium alloy series. Alternatively, the first structural element and the second structural element can be made of titanium or titanium alloys, such as Titanium grade <NUM> (Ti-6Al-4V). Also other materials such as composites or other metals can be used as long as these materials provide sufficient solder joint strength for the intended spacecraft requirement. , generally, the material choice for the first and second structural elements as well as for the solder preferably depends on requirements of the mission planned for the spacecraft.

Solder can be applied to the first connecting surface and the second connecting surface irrespective of the chosen (substrate) material, e.g., aluminium, aluminium alloy, titanium or titanium alloy. However, to allow adhesion of the solder to the first connecting surface and the second connecting surface, at least one, and preferably both, of the first and second connecting surfaces can be silver-plated. The silver-plating on the connecting surface(s) may be applied by a galvanic process or any other industrial process capable of doing so. It is also conceivable to use other plating materials instead of silver, wherein the choice of the plating material preferably depends on the material of the connecting surface(s) and the solder type used. To realize the joint state of the passive separation mechanism, a homogeneous layer of solder is applied to one of the first and second connecting surfaces, and the first and second connecting surfaces are then joined by a soldering process, such as vacuum vapor phase soldering.

During the atmospheric re-entry phase, the atmospheric drag and the aerodynamic heating phenomena heat up the spacecraft and, thus, also the separation mechanism, which results in melting of the solder joint. Once molten, the solder joint loses its strength and the first and second structural elements of the passive separation mechanism separate from another. As the first and second structural elements are connected to the first and second spacecraft components, respectively, the latter entails the separation of the first and second spacecraft components.

Generally, the first structural element can have a first axis and the second structural element can have a second axis. The first and second axes can be collinear in the joint state of the separation mechanism.

Further, it is conceivable that the first structural element has a first through hole and/or the second structural element has a second through hole. In case the first structural element has a first axis and/or the second structural element has a second axis, the first through hole can extend along the first axis of the first structural element and/or the second through hole can extend along the second axis of the second structural element.

Preferably, the receiving hole of the first structural element forms a section of the first through hole of the first structural element.

Claim <NUM> defines a spacecraft comprising a first spacecraft component, a second spacecraft component, and the above-described passive separation mechanism. The first structural element is mounted to the first spacecraft component and the second structural element is mounted to the second spacecraft component.

Preferably, the first structural element of the passive separation mechanism is formed integrally with the first spacecraft component. In other words, the first structural element can be an integral part of the design and manufacturing of the first spacecraft component in a manner that mounting to and dismounting from the first spacecraft component is not intended without compromising the integrity of the first spacecraft component.

The second structural element of the passive separation mechanism may be fixedly or removably mounted to the second spacecraft component.

The first structural element can extend through a wall of the first spacecraft component such that a first end face of the first structural element extends parallel to a first surface of the wall and a second end face of the first structural element opposite the first end face extends parallel to a second surface of the wall opposite the first surface. The receiving hole then is formed in the first or second end face of the first structural element.

Preferably, when the first structural element extends through a wall of the first spacecraft component, the first through hole of the first structural element extends completely through the wall of the first spacecraft component.

In sum, the separation mechanism presented herein works in an entirely passive manner and can make use of any connecting surface geometry allowing sufficient separation reliability and solder joint strength.

A detailed description of an exemplary embodiment of the claimed passive separation mechanism provided in a spacecraft is given in the following with reference to the figures which show schematically:.

<FIG> shows a passive separation mechanism <NUM> connecting a first spacecraft component <NUM>, which is a structural panel in <FIG>, to a second spacecraft component <NUM>. This connection is realized in the following manner. A first structural element <NUM> of the passive separation mechanism <NUM> is formed integrally with the first spacecraft component <NUM> and a second structural component <NUM> of the passive separation mechanism <NUM> is subsequently fastened to the second spacecraft component <NUM> by means of a screw or bolt <NUM> or other fastening means. In <FIG>, a joint layer of solder <NUM> is sandwiched between a first connecting surface <NUM> of the first structural component <NUM> and a second connecting surface <NUM> of the second structural component <NUM> to form a solder joint between the first and second connecting surfaces <NUM>, <NUM> of the first and second structural elements <NUM>, <NUM>.

Only sections of the first and second spacecraft components <NUM>, <NUM> are shown in <FIG>. The spacecraft components <NUM>, <NUM> can be of any desired shape departing from the shapes shown in <FIG>. In the oblique view of <FIG>, the first and second spacecraft components <NUM>, <NUM> as well as the screw <NUM> are omitted.

<FIG> shows a section of a wall <NUM> of the first spacecraft component <NUM>. The wall <NUM> has opposed first and second surfaces <NUM>, <NUM> and a core <NUM>. A through hole or through bore extends through the wall <NUM> from the first wall surface <NUM> to the second wall surface <NUM> and through the core <NUM>.

As shown in <FIG>, the first structural element <NUM> of the passive separation mechanism <NUM> has the shape of a spool with a main section <NUM> in the form of a cylinder limited by first and second flanges at the opposed end faces of the cylinder.

According to <FIG>, the spool-shaped first structural element <NUM> is inserted into the through hole or through bore through the wall <NUM> of the first spacecraft component <NUM>. When in place with its first and second flange end faces extending approximately flush with the first and second wall surfaces <NUM>, <NUM> of the first spacecraft component <NUM>, the first structural element <NUM> is fixed in the through hole or through bore of the wall <NUM> by injection of a fixing means, such as a two-component adhesive or potting material, through bores <NUM> extending through the flange second end face of the spool. The fixing means fix the first structural element <NUM> to the first spacecraft component <NUM> in an integral, non-releasable manner. Two of the through bores <NUM> are shown in <FIG> through the second end face flange of the spool-shaped first structural element <NUM>.

With further reference to <FIG>, the spool-shaped first structural element <NUM> has an inner through hole <NUM> extending completely through the cylinder main body and flange sections of the first structural element <NUM>. Said inner through hole <NUM> has a stepped inner surface. A step in the through hole <NUM> marks a transition between two hole sections of the through hole <NUM>, namely a cylindrical hole section extending from the second end face flange towards the first end face flange of the spool-shaped first structural element <NUM> and a conical hole section tapering inwardly to a longitudinal spool axis A<NUM> of the first structural element <NUM> when extending from the first end face flange towards the second end face flange of the spool-shaped first structural element <NUM>. An inner surface of the conical hole section forms the first connecting surface <NUM> of the first structural element <NUM>. An opening angle <NUM> between the first connecting surface <NUM> and the axis A<NUM> is in the range of <NUM>° to <NUM>°. The first connecting surface <NUM> has a shape and opening angle <NUM> providing sufficient separation reliability on the one hand and sufficient solder joint strength on the other hand.

The second structural element <NUM> has the shape of a conical trunk snuggly fitting into the conical hole section of the first structural element <NUM>, which acts as a receiving hole of the first structural element <NUM> for the second structural element <NUM>.

Also the second structural element <NUM> has a through hole <NUM> extending from a first end face <NUM> of the conical trunk shaped second structural element <NUM> to an opposed second end face. Also the through hole <NUM> of the second structural element <NUM> has a stepped inner surface <NUM>. Here a step in the through hole <NUM> marks a transition between a first cylindrical hole section and a second cylindrical hole section. The first cylindrical hole section has a smaller cross-section than the second cylindrical hole section and extends from the first end face <NUM> towards the second end face of the second structural element <NUM>. The second cylindrical hole section which is of larger cross-section than the first cylindrical hole section is adapted to receive a head of screw <NUM>. The transition step forms an abutment stop for the head of screw <NUM> and the first cylindrical hole section is adapted to receive a shaft section of screw <NUM>. An inner surface of the first cylindrical hole section can be threaded.

Claim 1:
A passive separation mechanism (<NUM>) for passive separation of a first spacecraft component (<NUM>) from a second spacecraft component (<NUM>), the passive separation mechanism (<NUM>) comprising:
- a first structural element (<NUM>) having a receiving hole with an inner first connecting surface (<NUM>), the first structural element (<NUM>) adapted to be mounted to the first spacecraft component (<NUM>), and
- a second structural element (<NUM>) adapted to be at least in part received within the receiving hole of the first structural element (<NUM>) and having an outer second connecting surface (<NUM>) facing the inner first connecting surface (<NUM>) when the second structural element (<NUM>) is at least in part received within the receiving hole of the first structural element (<NUM>), the second structural element (<NUM>) adapted to be mounted to the second spacecraft component (<NUM>),
wherein the passive separation mechanism (<NUM>) further comprises:
- joining means (<NUM>) for mechanically joining the first and second structural elements (<NUM>, <NUM>) together via their first and second connecting surfaces (<NUM>, <NUM>), the joining means (<NUM>), in a joint state of the passive separation mechanism (<NUM>), sandwiched between and connecting the inner first connecting surface (<NUM>) and the outer second connecting surface (<NUM>) together when the second structural element (<NUM>) is at least in part received within the receiving hole of the first structural element (<NUM>), wherein the joining means (<NUM>) is solder meltable at a temperature in the range of <NUM> to <NUM> so as to separate the first and second structural elements (<NUM>, <NUM>) when a melting temperature, i.e., a temperature in the range of <NUM> to <NUM>, is experienced by the separation mechanism (<NUM>),
- wherein the first and second connecting surfaces (<NUM>, <NUM>) are of corresponding conical shape.