DE102019124397B4 - Acquisition system - Google Patents

Abstract

System (20) which can be installed on an artificial spacecraft (17) and is designed to detect space debris objects, with a detection device for determining a size, a number, a temporal distribution and / or a flight direction of space particles, the detection device at least one film-like, movable sensor part (1) and at least one detector part (6), the film-like, movable sensor part (1) having a sensor surface (13) for detecting at least one space particle caused by an impact (P1, P2) or puncture (14) (15) formed marking, wherein the film-like, movable sensor part (1) is designed to be movable relative to the detector part (6) in such a way that at least one marking made by the impact (P1, P2) or punch (14) on the film-like, movable Sensor part (1) is transported relative to the detector part (6), the detector part (6) for detecting the by the Impact (P1, P2) or the perforation (14) caused marking is carried out on the film-like, movable sensor part (1), characterized in that the film-like movable sensor part (1) is a sensor film (1) which is spaced between at least two Transport rollers or guide rollers (2, 3), between which the sensor surface (13) is arranged, the film-like movable sensor part (1) being a sensor film (1) which is designed as at least one endless belt (16, 16a, 16b) , with each endless belt (16, 16a, 16b) being assigned a detector part (6), with each endless belt being guided by at least four deflection rollers (4) in such a way that two parallel, closely spaced sensor surfaces (13a, 13b) are formed and wherein a first endless belt (16a) facing a tapering space particle (15) and a second, viewed in the direction of flight of the space particle (15), arranged behind the first endless belt (16a) and from d An endless belt (16b) spaced apart is provided in the first endless belt (16a).

Description

The invention relates to a system that can be installed on an artificial spacecraft and is designed to detect space debris objects, with a detection device for determining a size, a number, a temporal distribution and / or a flight direction of space particles. The invention also relates to a method for operating the system.

Since the start of operational space travel with the launch of the first artificial Earth satellite Sputnik 1 in October 1957, a number of rockets, satellites, probes and space shuttles have been launched into orbit. Garbage is produced at every single launch, e.g. by mission-related space debris objects, with the garbage remaining in orbit for a long time. This space debris endangers space travel in all orbits of the earth. It is therefore necessary to use all possible methods to record them. On February 10, 2009, for the first time in human history, two satellites collided in orbit. The collision involved the two communication satellites Iridium 33 and the Russian cosmos 2251, which were located over Siberia at an altitude of 789 km above the earth's surface. The mutual impact, which occurred at a relative speed of 11.6 km / s (= 41,760 km / h), resulted in a debris field consisting of a large number of newly created space garbage objects, most of which are still circling the earth as space debris .

Parts smaller than 5 cm can hardly be located from the ground in Low Earth Orbit (LEO). On the other hand, collisions of satellites with space debris objects with a diameter of more than approx. 1 cm generally lead to a complete loss of the mission. The number of space debris objects in orbits increases exponentially with decreasing diameter.

There are different approaches for measuring space debris objects in space. These can be divided into destructive and non-destructive processes. Non-destructive methods are characterized by the fact that the space debris object is remotely sensed by means of a measuring device (active interaction). Ground-based radar and telescope systems can be mentioned here as examples, which are used, among other things, to measure the orbit of space debris objects and thus for early warnings. Using ground-based systems, space debris objects with a diameter of approx. 10 cm in the LEO and approx. 1 m in the geostationary orbit (GEO) can be operationally measured. Smaller space debris objects, e.g. with a diameter of 5 cm in the LEO or with a diameter of 10 cm in the GEO, can be detected sporadically. Space debris objects with a diameter smaller than 2 cm cannot be detected in the LEO. The capture of space debris using ground-based radar and telescope systems is currently the most widely used method. Radar and telescope systems for the detection of space debris can be both ground-based and space-based.

However, most devices for the detection of space debris are currently ground-based.

In destructive processes, a space debris object to be measured hits the measuring device “intentionally”. Destructive processes act as "in-situ" measuring instruments directly in orbit and provide measured values for small space debris objects (diameter in the range from µm to cm) that cannot be recorded using radar and telescope systems. One example is the SOLID system developed at DLR Bremen that uses the subsystems of a satellite (solar panels and attitude control system). A summary of existing in-situ detectors can be found e.g. in the work "Bauer, Waldemar, Space Debris Detection for the Validation of Simulation Models, DLR Research Report, 2015-08".

Space debris objects with a diameter of approx. 1 cm are critical for unnamed satellites, since these space debris objects have a high probability of completing a mission in the event of a collision. Smaller space debris objects in the millimeter range can limit a mission. These space debris objects can break through the walls of a satellite and damage vulnerable components inside a satellite. Space debris objects in the µm range are critical, for example, for sensitive instruments (e.g. optical telescopes).

The European published application EP 2 236 422 A1 discloses a sensing device for a space flying object comprising a sensing plate member supporting a plurality of conductive sensing lines at a predetermined pitch on a non-conductive thin film that can be exposed in a space environment. Further, the detection device comprises a detection circuit connected to each detection line, detection of the object flying in space which has collided with the detection plate member is enabled by the detection circuit when a detection line on the detection plate member by a collision with a flying in space Object is interrupted.

From the US published patent application US 2012/0 318 925 A1 and the European patent EP 3 345 840 B1 Further solutions are known in which two or more electrically conductive layers are separated by an insulator in order to be able to measure voltage changes occurring in the event of a breakdown or impact.

Other known methods are based on an optical observation of the space particles by means of a CCD detection, as in the Chinese patent applications CN 105 910 583 A , CN 103 675 938 A and CN 101 846 511 A is shown.

Another common measuring method is characterized by the fact that ion detectors, laser and acoustic methods are used. For measurements with ion detectors, the Chinese patent application GN CN 107 808 817 A proposed a use of a mass spectrometer. Another from the Chinese Offenlegungsschrift CN 107 589 459 A Well-known ion measurement method uses a capacitance measurement on a three-layer, electrically insulated metal foil.

All known systems have in common that they cause high costs and are largely unable to capture the smallest particles. Today's predictions are based on simulation models that hardly use measured values for validation.

Known systems are also not able to record a size, a number, a temporal distribution or a flight direction of space particles.

The invention is based on the object of specifying a system for the detection of space debris objects of the type mentioned that can be installed on an artificial spacecraft and that is able to detect a size, a number, a temporal distribution and a flight direction of space particles. Furthermore, the system should be able to cost-effectively detect the smallest particles down to the µm range.

According to the invention, this object is achieved by a system with the features of independent claim 1. Advantageous further developments of the system result from subclaims 2 to 5.

Another object of the invention is to provide a method for detecting space debris objects in space by means of the system.

This further object is also achieved by a method according to claim 6. Advantageous embodiments of the method emerge from the dependent claims 7 to 8.

The invention is based on the idea of using a film through which space particles can penetrate or on which space particles impacting leave at least a mark in the form of a deformation or other change. These changes are recorded and evaluated by a detector part, in principle like a scanner. By using several parallel foils, it is even possible to record a flight direction of space particles.

The invention makes it possible to create a device for determining the size, number, temporal distribution and direction of flight of small space particles in a simple manner. Furthermore, the invention makes it possible to record the space debris which is produced as a result of space travel activities and which represents an increasing problem for space missions currently in orbit and planned future space missions.

• Der Begriff „Weltraummüll“ beschreibt dabei alle nicht mehr aktiv vom Boden aus kontrollierbaren Objekte, die sich als Folge von Start, Betrieb und nicht geplanten Ereignissen im Zusammenhang mit Raumfahrzeugen in einer Erdumlaufbahn befinden.

The following should be explained conceptually:

• The term “space debris” describes all objects that can no longer be actively controlled from the ground and that are in earth orbit as a result of launch, operation and unplanned events in connection with spacecraft.

The detection according to the invention helps to solve the problem that arises due to their high path speed, the high kinetic energy of these uncontrolled moving masses and the associated potential for damage in the event of a collision with other objects. These objects, summarized under the term “space debris”, are parts of very different sizes and masses.

In addition to burned-out rocket stages and complete satellites that are no longer operational, small parts such as fragments after explosions or remains of rocket fuel are also part of the problem. Without going into a complete description of all possible causes of space debris here, it is important to note that the number of objects forming a collision risk increases exponentially with decreasing size or mass. A further complicating factor is that it is precisely these small parts that can no longer be observed from the ground and therefore the existing strategy of evasive maneuvers for operational spacecraft cannot be carried out in the case of larger space debris with a known orbit.

The detection according to the invention can advantageously be carried out with a terrestrial object location be combined. From the ground, objects with a size of 0.1 m or more are recognizable in low earth orbit and objects with a size of approx. 0.3 m or more in geostationary distance from the earth. That is, while large parts can be located from the ground, the invention allows the smallest parts to be detected in space by attaching the system according to the invention to a missile.

For these parts of the currently existing space debris, the orbit parameters of the orbit can be determined through appropriate evaluation of the observations, and thus possible collisions with satellites that are still active can be recognized and prevented by planned orbit maneuvers. The detection according to the invention is very advantageous for smaller particles that cannot be seen from the ground, since even the smallest objects can cause various types of damage due to their high kinetic energy. The detection of these small parts provides knowledge of their frequency, size and spatial distribution, which is important for the design of future spacecraft.

For this purpose, a wide variety of sensors were developed in the past and placed in orbit. With the knowledge that can be obtained by the invention, however, it is possible to create models with regard to the statistical properties of small-sized space debris.

Another advantage of the invention is that the usable sensor measuring areas no longer have to be small due to the simple technology, but can be significantly more than 1 m ² .

In addition to the technology according to the invention for detecting space debris, an evaluation of damage to hardware that has returned to the ground and other damage is also possible. In the past there were problems with windows on the space shuttle, with the EURECA platform or with parts of the Hubble Space Telescope hardware. Compared to other detection solutions, the invention offers the advantage that a measurement duration, that is to say the period in which the hardware was exposed to space conditions, does not have to be short, which allows reliable statistical statements.

The known problem of a measurement area that is too small for statistical surveys arises because of the low probability of space debris particles striking a small measurement area.

For an orbit near the pole with 98 ° inclination and an orbit height of 800 km, for example, with a measuring area of 1 m ² and a particle size of 100 µm, only one hit in 0.8 days can be assumed. For an object size of 1 mm, this value even drops to just one hit in 50 years. This clearly shows the importance of a sufficiently large sensor area for statistically meaningful measurements.

In addition to the problem of a sensor surface that is too small, the invention also eliminates the difficulty of convincing the developers of future satellites to plan for the additional hardware required, since the hardware according to the invention is inexpensive. Another advantage provided by the invention is an uncomplicated spatial alignment of the sensor surfaces. If the measuring device known from the prior art is located, for example, on a solar generator of a satellite, it will always be aligned with the sun due to the mission and cannot be optimally adapted to the special mission requirements of measuring space debris particles.

The invention also offers the advantage over other measurement methods that excessive masses and energy requirements are not necessary, since the sensor foils can be thin and light. Another problem solved by the invention is that, in contrast to the currently usual method, with a subsequent analysis of damage to hardware that has returned to the ground, the exact time of the impact in orbit can be determined and thus conclusions about the flight direction of the space debris particles are possible.

Compared to other known methods and systems, the invention has the advantage that the sensor technology (detector part) does not have to be arranged on the entire surface exposed to space particles. For example, there is no need to use electrically insulated layers, which have high technical requirements and lead to large masses, especially in the case of larger measuring surfaces. With the present invention it becomes possible to easily realize even large measuring areas which are important for static analyzes, since in this case there is an extensive separation between the measuring area and the sensor technology (detector). The surface exposed to the space particles consists only of a film, whereby the markings made by the impacting space debris particles can easily be detected by the sensor.

The film or the films allow large measuring areas, a simple technology and have the advantage of a low required mass. An advantageous transport device for the film conveys the information contained on the measuring surface to a sensor technology arranged at a suitable location on the spacecraft and can there together with the Time, position and attitude information of the spacecraft are evaluated.

The invention provides a device which enables the properties, size, number, temporal distribution and direction of flight of small space debris particles to be determined in a simple manner with a large measuring surface.

In other words, the object is achieved according to the invention in that a system that can be installed on an artificial spacecraft and is designed to detect space debris objects has a detection device for determining a size, a number, a temporal distribution and / or a flight direction of space particles wherein the detection device comprises at least one film-like, movable sensor part and at least one detector part, wherein the film-like, movable sensor part forms a sensor surface for detecting at least one marking caused by an impact or puncture of a space particle, the film-like, movable sensor part relative to the detector part in such a way is designed to be movable that at least one marking made by the impact or puncture is transported on the film-like, movable sensor part relative to the detector part, the detector part for detecting the is carried out by the impact or the breakdown caused marking on the film-like, movable sensor part.

The moving film-like structure located on the detection device as a sensor film is ideally penetrated by the space debris and the marking thus produced is conveyed by a transport device to a detector part and evaluated there.

The basic idea of the invention is therefore to arrange a movable, film-like structure as a sensor film in such a way that it is penetrated by the space debris particles on its flight path. The trace of the breakthrough is transported to a sensor by means of rotatable transport rollers which are preferably present and evaluated there. In a simple embodiment of the invention, it is thus possible to determine the number of space debris particles for a given area within a certain time, as well as to determine the size of these space debris particles. The geometrical extent of the space debris particles can be deduced from the hole size. The temporal distribution of the space debris per area can be determined by knowing the transport speed of the measuring film.

In other words, the characteristic technical principle of the invention consists in that a sensor film is attached to the detector element, through which the space particles ideally penetrate. The sensor film is moved, for example, via suitably positioned transport rollers, which means that the breakdown point generated by the space particle is moved to one or more sensors and its position on the sensor film is detected there. In order to achieve this, the sensor film is provided with a marking in the longitudinal and transverse directions that can be detected by the sensor. The coordinates of the breakdown are recorded by the sensor and stored, for example, in an associated data memory. Since the position of each recorded measurement is saved, subsequent new breakthroughs can be distinguished from those that have already been recorded.

According to the invention it is provided that the film-like movable sensor part is a sensor film which is stretched between at least two spaced apart transport rollers, between which the sensor surface is arranged. With this simple first embodiment, a numerical detection of the space particles is possible. It is possible to reverse the direction of the transport rollers in order to enable the film to be moved back and forth and thus achieve a long service life.

For this it is particularly favorable if a single sensor surface is arranged between the two transport or guide rollers. The film is stored on both rolls. There is a first transport role at the beginning. The movement of the sensor film unrolls it from the transport roller. A second transport roller, on which the sensor film is rolled up, is available in the direction of the transport direction. This arrangement is characterized in that a single sensor surface is arranged between the two transport rollers, a first transport roller or guide roller being arranged in the opposite transport direction of the sensor film, the sensor film being unrolled from this transport roller or guide rollers or with the movement of the sensor film from the transport roller or guide roller is unrolled, and wherein a second transport roller or guide roller is present in the direction of the transport direction on which the sensor film is rolled up.

According to the invention it is provided that the film-like movable sensor part is a sensor film which is designed at least as an endless belt. A detector part is assigned to each endless belt. This advantageous embodiment of the invention also enables the point in time of the breakdown of the film to be determined. For this purpose, the roll-up device is replaced by a mechanism similar to a conveyor belt. The measuring film consists of two layers on top of each other. These move in opposite directions, as with a conveyor belt, and must be arranged as close to each other as possible in order to perform the task. The desired breakthrough of the space debris now takes place through both foils, practically at the same time, the two breakthrough points of the foil immediately moving apart as they move in opposite directions. By knowing the transport speed of the measuring film and by measuring the position of the two breakthroughs on the foil, the point in time of the breakthrough can now be determined in a simple manner. In this case, use is made of the fact that, due to the design, the two breakdown points have the same distance with respect to one of the respective axes of rotation of the conveyor belt mechanism. By knowing the speed of movement of the measuring film, the position of the conveyor belt mechanism at the time of the breakdown can thus be determined, and thus the time of the breakdown can be determined.

So that the two superimposed film sections are as close to one another as possible, each endless belt is guided by at least four deflection rollers in such a way that two parallel sensor surfaces that are close to one another are formed. This minimizes measurement errors when determining the point in time of the breakdown. To simplify the calculation of the point in time of the breakthrough, it can be assumed that the markings on the upper and lower foils are made simultaneously. The time lag between the penetration of the upper film and the lower film is negligible.

According to the invention there is a first endless belt facing a tapering space particle and a second endless belt, seen in the direction of flight of the space particle, arranged behind the first endless belt and spaced apart from the first endless belt. By combining two mechanisms similar to the conveyor belt described above, the invention also enables the direction vector of the impact to be determined from the trajectory of the space particle and the direction of the breakthrough. For this purpose, the two conveyor belt elements are arranged one above the other and parallel to one another, which means that the space particle now penetrates through four measuring foils. The assignment of the copies that belong together is easily possible by knowing the times of the copies. Now that two spatial coordinates are known, the trajectory of the space particle can be determined. For the final determination of the flight vector, the determination of the direction of the breakdown is still missing.

Since a determination by comparing the breakdown times does not appear to be possible due to the high flight speeds of the space debris particles and the comparatively small distance between the measuring foils, the direction must be determined by an optical analysis of the breakdown crater using a suitable sensor.

With regard to the scope of the intended measurements, a distinction can basically be made between three embodiments, as described. To enlarge the measurement area, several of these detector elements can be mounted on the spacecraft.

If the intention is only to record the number of space particles, then the first embodiment is sufficient, in which the sensor film is transported in only one layer, for example from a first roll to a second roll, and is rolled up there. This process can be repeated as often as required by reversing the direction of rotation, as long as the condition of the sensor film allows it. The disadvantage of this embodiment is that the point in time of the breakdown event cannot be determined, since only the location coordinates are recorded by the sensor.

If it is also intended to measure the point in time of the breakdown of the space particle, the second embodiment is required. The sensor film is arranged in a manner comparable to the mechanics of a conveyor belt. The resulting geometry means that the breakdown now takes place, practically simultaneously, at two points on the endless sensor film. For a correct measurement it is necessary to arrange the two parts of the sensor film moving in opposite directions as close to one another as possible. This is achieved through additional pulleys and other guide structure elements. The measuring principle used is that two breakdown points are detected by the sensor for each space particle, both of which, assuming a symmetrical arrangement of sensor and sensor film, are equidistant from the sensor. When evaluating the measurement coordinates, if the starting position and transport speed of the sensor film are known, the point in time at which the breakdown occurred can thus be determined. By comparing the known point in time with the recorded data from the position control of the spacecraft, the location of the event can be determined in addition to the point in time.

If, according to the invention, in addition to measuring the number, point in time and location, one also intends to determine the direction of the space particle, the following technical implementation is required. In this embodiment, two mechanisms according to the second embodiment are described in a simplified manner and are arranged one above the other. By at the appropriate distance arranged sensor film elements now result, according to the measuring method described according to the second embodiment, two separate measuring points with a known point in time of the breakdown. The breakdown time can therefore be assumed to be the same for both sensor elements due to the extremely high orbital speed of the space particles and the comparatively small distance between the sensor foils. Together with the recorded data from the position and attitude control of the spacecraft, an undirected straight line on which the space particle is moving can now be determined using the two known points. For a complete determination of the direction vector, an optical examination of the breakdown crater on the sensor or an evaluation of the disturbing torques occurring at the time of the event can also be used to monitor the position of the spacecraft. With regard to the requirements of a technical implementation, it should also be mentioned that the transport speed of the sensor film is very low and therefore no negative effects occur due to the friction effects of the two sides of the sensor film. The speed at which the two breakdown points move apart serves only to be able to determine the position of the sensor film at the time of the breakdown.

A particularly advantageous film dimensioning results if, according to a further preferred embodiment of the invention, the material of the sensor film is made and / or the film thickness is dimensioned such that it can be penetrated by space particles of at least 1 mm, in particular at least 0.1 mm. The resulting penetration hole can be easily and clearly detected using optical methods.

In a further advantageous embodiment of the system according to the invention it is provided that an effective sensor area of at least ^{1 m 2} , preferably of several m 2, is present. A large sensor area avoids the disadvantages described at the beginning. With a large sensor area, it is possible to get a precise overview of the space debris situation.

In order to achieve an evaluation on earth or in space, the system according to the invention has one or at least one evaluation electronics connected to each detector part for signaling purposes for evaluating the markings of the space particle impacts or punctures on each film-like, movable sensor part.

1. i. Anordnen des beweglichen Sensorteils an gegenüber im Weltraum fliegenden Weltraumteilchen exponierter Position eines künstlichen Weltraumflugkörpers,
2. ii. Bewegen des beweglichen Sensorteils zu einem Detektorteil, und
3. iii. Erfassen der Markierung eines Einschlags oder Durchschlags auf dem beweglichen Sensorteil mittels des Detektorteils.

The method according to the invention for determining a number of space particles with a system described above is characterized by the steps

1. i. Arranging the movable sensor part at a position of an artificial spacecraft that is exposed to space particles flying in space,
2. ii. Moving the movable sensor part to a detector part, and
3. iii. Detecting the marking of an impact or puncture on the movable sensor part by means of the detector part.

In an advantageous embodiment of the method, the detector part detects the position of the two breakthroughs and the point in time of the breakthrough is calculated by means of the transport speed of the film-like, movable sensor part.

In a further advantageous embodiment of the method, in addition to the time of the breakthrough event, the location of the breakthrough event can also be determined by means of recorded data from the position control of the spacecraft.

By means of statistical probability or statistical recording of space debris objects, conclusions can be drawn about larger objects. A method for the statistical recording of space debris objects is thus possible. On the basis of the space particles hitting the sensor surface, a statistical probability of the presence of larger space parts can be calculated.

Working examples

Exemplary embodiments of the invention are explained in more detail with reference to the drawings, further advantageous developments of the invention and advantages thereof being described.

• 1 eine schematische Darstellung einer ersten Ausführungsform des erfindungsgemäßen Systems zur Erfassung von Weltraummüll-Objekten,
• 2 eine schematische Darstellung einer zweiten Ausführungsform des erfindungsgemäßen Systems zur Erfassung von Weltraummüll-Objekten,
• 3 eine weitere schematische Darstellung der dritten Ausführungsform des erfindungsgemäßen Systems,
• 4 eine schematische Darstellung des Verfahrens am Beispiel der zweiten Ausführungsform des erfindungsgemäßen Systems zur Erfassung von Weltraummüll-Objekten, und
• 5 eine weitere schematische Darstellung der zweiten Ausführungsform des erfindungsgemäßen Systems mit weiteren Details.

Show it:

• 1 a schematic representation of a first Embodiment of the system according to the invention for the detection of space debris objects,
• 2 a schematic representation of a second embodiment of the system according to the invention for detecting space debris objects,
• 3 a further schematic representation of the third embodiment of the system according to the invention,
• 4th a schematic representation of the method using the example of the second embodiment of the system according to the invention for detecting space debris objects, and
• 5 a further schematic representation of the second embodiment of the system according to the invention with further details.

the 1 until 3 show schematically three embodiments of one on an artificial spacecraft 17th installable system 20 for the detection of space debris objects. All the illustrated embodiments have at least one film-like, movable sensor part 1 in the form of a sensor film. Furthermore, all of the illustrated embodiments have at least one detector part 6th on. In all three embodiments described in 1 until 3 each sensor film is shown 1 between at least two spaced apart transport or guide rollers 2 and 3 stretched. Between roles 2 and 3 is the sensor surface 13th as part of the sensor film 1 . On the sensor foils 1 , of all the embodiments shown, there are markings applied in the longitudinal and transverse directions 5 by the detector part 6th are recognizable in order to allow the coordinates of the through a carbon copy 14th or impact P1 , P2 to be able to determine caused markings.

In the in 1 depicted embodiment, the system 20 is able to size and number of on the film-like, movable sensor part 1 hitting space particles 15th to determine. The film-like, movable sensor part 1 is as a sensor film 1 executed. The sensor film 1 in this embodiment is a single layer from a first transport roller 2 unrolled and to a second transport roller 3 transported in the x direction and rolled up there. This process can be reversed by reversing the direction of rotation, as long as the condition of the sensor film 1 this can be repeated as often as desired. The sensor film 1 tensioned between the first transport roller 2 and the second transport roller 3 a sensor area 13th to capture at least one through an impact P1 , P2 or punch 14th of a space particle 15th evoked marking. On the second transport roller 3 , and / or on the transport roller 2 , is a detector part 6th arranged. This detector part 6th can for example be designed as a sensor. In 1 is a space particle 15th drawn on the sensor film 1 hits and a punch 14th caused. This breakthrough 14th forms a mark on the sensor film 1 , with the sensor film 1 relative to the detector part 6th can be moved in such a way that the through the punch 14th made marking relative to the detector part 6th is transported where the coordinates of the breakthrough 14th based on the markings 5 (only in 5 to see) recorded and, for example, in an associated data memory 18th be filed. This becomes the determination of the number of space debris particles 15th for a given area within a certain time, as well as the determination of the size of these space debris particles 15th enables. It can be based on the geometric extent of the space debris 15th by the hole size of the penetration 14th getting closed. The temporal distribution of the space debris particles 15th per area, can be achieved by knowing the transport speed of the sensor film 1 to be determined. Since the position of each recorded measurement is saved, subsequent new breakthroughs can be distinguished from those that have already been recorded. It is possible to reverse the direction of the transport rollers 2 , 3 to move the sensor foil back and forth 1 to enable and thus to achieve a long service life. As all marks from carbon copies 14th or impacts P1 , P2 are recorded and saved, are new copies 14th or impacts P1 , P2 distinguishable from already recorded.

In the in 2 In the illustrated embodiment, the system 20 is capable of the size and number as well as a temporal distribution and the direction of flight from onto the film-like, movable sensor part 1 hitting space particles 15th to determine. In this embodiment the sensor film is 1 as an endless belt 16 executed. The endless belt 16 wraps around the first and second guide role 2 , 3 each partially. By rotating the first and second guide rollers 2 , 3 the endless belt moves 16 in a circuit in such a way that it passes through the detector part with each revolution 6th is passed. The endless belt 16 forms between the first and second guide role 2 , 3 two sensor surfaces 13a , 13b , namely a first sensor surface 13a on its surface when traveling in the x direction between the first guide roller 2 and the second leadership role 3 , and a second sensor surface 13b on its surface when returning in the opposite x-direction between the second guide roller 3 and the first leadership role 2 . Pulleys 4th are located in the forward and backward directions and steer the endless belt 16 so that the endless belt is 16 two essentially parallel and closely spaced sensor surfaces in the flow and return 13a , 13b form. This advantageous embodiment of the invention also enables the point in time of the breakdown of the film to be determined. Forward and reverse of the endless film 16 move like a conveyor belt, i.e. in opposite directions. A breakthrough 14th of a space debris particle 15th now takes place through both sensor surfaces 13a and 13b . The breakthrough 14th through the two sensor surfaces 13a , 13b occurs because of the high orbital speed and the large kinetic energy of the space particles 15th practically at the same time, with the two carbon copies 14th the endless film 16 move them apart immediately by moving them in opposite directions. By knowing the transport speed of the endless film 16 and by measuring the location of the two breakthroughs 14th on the through Forward and reverse formed two layers of the continuous film 16 you can now easily determine the point in time of the breakdown 14th to be determined. The fact that the two breakdown points 14th Due to the design, based on one of the respective axes of rotation of the guide rollers 2 , 3 have the same distance. By knowing the speed of movement of the continuous film 16 can thus change the position of the guide rollers 2 , 3 at the time of the breakthrough 14th via a position control 19th and thus the point in time of the breakdown 14th be determined. By minimizing the distance between the forward and reverse lines of the endless film 16 , over the pulleys 4th , becomes a measurement error when determining the point in time of the breakdown 14th minimized. To simplify the calculation of the point in time of the breakthrough 14th it can be assumed that the markings on the two sensor surfaces 13a , 13b take place at the same time, because the time lag between the two breakdowns 14th of the sensor surfaces 13a , 13b is negligible.

As in 3 shown, two endless belts can also be used 16a , 16b used, each with endless belt 16a , 16b a detector part 6a , 6b assigned. The structure of this embodiment corresponds to the doubling of the structure of the embodiment according to FIG. 2 . To avoid repetition, the description of the structure of the embodiment according to FIG. 2 referenced. As a result, in this embodiment two sensor surfaces are essentially parallel and close to one another 13a , 13b and 13c , 13d formed. This embodiment of the invention additionally enables the determination of the direction vector of the breakdown by the combination of two of the conveyor belt-like mechanisms described above 14th from the trajectory of the space particle and the direction of the breakthrough. For this purpose, the two conveyor belt-like mechanisms are arranged one above the other and essentially parallel to one another, whereby the breakdown is achieved 14th of the space particle 15th now through four layers of sensor foils 1 he follows. The assignment of the copies that belong together 14th is through knowing the times of the breakthroughs 14th easily possible. Now that two spatial coordinates are known, the trajectory of the space particle 15th be determined. For the final determination of the flight vector, the determination of the direction of the breakdown is still missing 14th . As a determination by comparing the breakdown times, due to the high flight speeds of the space debris particles 15th and because of the comparably small spacing between the layers of the sensor foils 1 , does not appear possible, the direction must be determined by an optical analysis of the breakdown crater using a suitable sensor.

For all shown embodiments it is advantageous to use the material of the sensor film ( 1 ) and / or the thickness of the sensor film 1 so choose that they are from space particles 15th can be penetrated by at least 1 mm, in particular by at least 0.1 mm. Furthermore, it is advantageous for all the embodiments shown if the usable sensor area 13th respectively. 13a , 13b , 13c , 13d is at least 1m ² , preferably several m ² , and one with each detector part 6th , 6a , 6b Signal-connected evaluation electronics for evaluating the on each sensor film 1 striking impacts P1 , P2 or carbon copies 14th the space particles 15th is available.

4th is a schematic representation of the method using the example of the second embodiment of the system according to the invention for detecting space debris objects according to FIG. 2 . With regard to the designations of the illustrated components of the system 20, refer to 2 referenced so that the corresponding reference symbols in 4th has been omitted for the sake of clarity. Instead, the zero point 30 for the x coordinate is shown in this figure. P1 refers to an impact in the upper layer of the sensor film 1 with the x-coordinate x1 and the y-coordinate y1 aligned perpendicular thereto. P2 refers to an impact in the lower layer of the sensor film 1 with the x-coordinate x2 and the y-coordinate y2 aligned perpendicular thereto. The sensor film 1 thus has markings with the coordinates x1, y1 or x2, y2 viewed in the respective transport direction x, -x. The detector part 6th recognizes these coordinates as the sensor film passes by 1 . By knowing the transport speed of the endless film 16 and by measuring the position of the coordinates x1, y1 and x2, y2 you can now easily determine the point in time of the breakdown 14th to be determined. The fact that the two breakdown points 14th Due to the design, based on one of the respective axes of rotation of the guide rollers 2 , 3 , have the same distance. By knowing the speed of movement of the continuous film 16 can thus change the position of the guide rollers 2 , 3 at the time of the impacts P1 , P2 via a position control 19th and thus the point in time of the breakdown 14th be determined. To simplify the calculation of the point in time of the impacts P1 , P2 it is assumed that the markings on the two sensor surfaces 13a , 13b take place at the same time, because of the time lag between the two impacts P1 , P2 of the sensor surfaces 13a , 13b is negligible.

In 5 are the sensor film 1 who have favourited leadership roles 2 , 3 who have favourited pulleys 4th , the detector part 6th , Support elements 7th , a role 8th , a storage 9 the second leadership role 3 , Guide surfaces 10 and 11th and a spacecraft structure 12th the embodiment according to. 2 of the system 20 can be seen. Furthermore are the markings 5 in x and y directions on the sensor film 1 indicated in the figure. The sensor film 1 is via the transport or guide rollers 2 and 3 moves and their top and bottom over several pulleys 4th aligned in the measuring range at the smallest possible distance from one another. On the sensor film 1 there are markings made in the longitudinal and transverse directions by the detector part 6th can be recognized in order to determine the coordinates of a breakdown 14th or impact P1 , P2 to be able to determine. Support elements 7th serve to form a stiff frame structure and for receiving or rotatable mounting 9 the leadership role 3 . The support elements shown here with a rectangular cross-section 7th the frame structure each consist of one on a roll 8th Flattened, rolled-up structure, preferably made of CFRP, which unfolds independently into its predetermined shape after unrolling and thus obtains the required rigidity. For guiding on both sides and to ensure the smallest possible distance between the top and bottom of the sensor film 1 are on the supporting elements 7th Guide surfaces over the entire length of the measuring surface 10 and 11th attached, as a starting configuration together with the supporting elements 7th on the roll 8th are rolled up and because of their fixed connection with the support elements 7th unfold together at the beginning of the mission. All of the mechanics described above are based on a spacecraft structure 12th attached, with the support elements deployable in orbit 7th during takeoff on the rollers 8th are rolled up.

List of reference symbols

1

film-like, movable sensor part, sensor film

2

first guide roller, first transport roller

3

second guide roller, second transport roller

4th

Pulley

5

Markings in the XY coordinate system

6th

Detector part, sensor / scanner

6a, 6b

Detector part

7th

Support element

8th

role

9

storage

10

first guide surface

11th

second guide surface

12th

Spacecraft structure

13, 13a, 13b, 13c

Sensor area

14th

Breakthrough

15th

Space particles

16, 16a, 16b

Endless belt

17th

Spacecraft

18th

Storage

19th

Position control

P1

A hit

P2

A hit

X

direction

Claims (8)

System (20) which can be installed on an artificial spacecraft (17) and is designed to detect space debris objects, with a detection device for determining a size, a number, a temporal distribution and / or a flight direction of space particles, the detection device at least one film-like, movable sensor part (1) and at least one detector part (6), the film-like, movable sensor part (1) having a sensor surface (13) for detecting at least one space particle caused by an impact (P1, P2) or puncture (14) (15) formed marking, wherein the film-like, movable sensor part (1) is designed to be movable relative to the detector part (6) in such a way that at least one marking made by the impact (P1, P2) or punch (14) on the film-like, movable Sensor part (1) is transported relative to the detector part (6), the detector part (6) for detecting the through the Impact (P1, P2) or the perforation (14) caused marking is carried out on the film-like, movable sensor part (1), characterized in that the film-like movable sensor part (1) is a sensor film (1) which is spaced between at least two Transport rollers or guide rollers (2, 3), between which the sensor surface (13) is arranged, the film-like movable sensor part (1) being a sensor film (1) which is designed as at least one endless belt (16, 16a, 16b) Each endless belt (16, 16a, 16b) is assigned a detector part (6), each endless belt being guided by at least four deflection rollers (4) in such a way that two parallel, close-together sensor surfaces (13a, 13b) are formed and wherein there is a first endless belt (16a) facing a tapering space particle (15) and a second endless belt (16b), seen in the direction of flight of the space particle (15), arranged behind the first endless belt (16a) and spaced apart from the first endless belt (16a) is. System (20) according to Claim 1 , characterized in that a single sensor surface (13) is arranged between the two transport rollers or guide rollers (2, 3), a first transport roller or guide roller (2) being arranged in the opposite transport direction (X) of the sensor film (1), with from The sensor film (1) is unrolled from this transport roller or guide roller (2), and a second transport roller or guide roller (3) is present in the direction of the transport direction (X) on which the sensor film (1) is rolled up. System (20) according to one of the preceding claims, characterized in that the material of the sensor film (1) is such and / or the film thickness is dimensioned such that it is made up of space particles (15) of at least 1 mm, in particular of at least 0, 1 mm can be pierced. System (20) according to one of the preceding claims, characterized in that there is an effective sensor area (13) of at least 1 m ² , preferably of several m ² . System (20) according to one of the preceding claims, characterized in that the system (20) has evaluation electronics connected to each detector part (6) for signaling purposes for evaluating the markings of the impacts (P1, P2) striking each film-like, movable sensor part (1) or punctures (14) of the space particles (15). Method for determining a number of space particles (15) with a system (20) according to one of the preceding Claims 1 until 5 , characterized by the steps, i. Arranging the movable sensor part (1) at an exposed position of an artificial spacecraft (17) flying in space opposite space particles (15), ii. Moving the movable sensor part (1) to a detector part (6) iii. Detection of the marking of an impact (P1, P2) or puncture (14) on the movable sensor part (1) by means of the detector part (6). Procedure according to Claim 6 , characterized in that the detector part (6) detects the position of the two breakthroughs (14) and the time of the breakthrough (14) is calculated by means of the transport speed of the film-like, movable sensor part (1). Procedure according to Claim 6 characterized in that the location of the breakthrough event is determined by comparing the point in time of the breakthrough (14) with recorded data from the position control of the spacecraft (17).

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