阅读说明：本技术 能够编队飞行的小型卫星和数颗小型卫星的编队 (Formation of mini-satellites and several mini-satellites capable of flying in formation ) 是由 克劳斯·席林 于 2018-02-08 设计创作，主要内容包括：本发明涉及用于LEO应用的能够以编队(10)飞行的小型卫星,特别是质量为10kg或者更小的纳米卫星或者皮米卫星,其包括壳体(12)和设置在壳体(12)中的具有预定功能的至少一个插入板(14),以及用于在飞行轨迹T<Sub>k</Sub>的方向上产生定向冲力的推进系统(16)。小型卫星(10)包括独立且自主工作的防撞系统(18),当预期到与物体(30)发生碰撞时,防撞系统(18)能够通过推进系统(16)调整轨迹T<Sub>k</Sub>的轨迹校正T<Sub>kk</Sub>。在另一个方面,本发明涉及由数个能够以编队(10)飞行的小型卫星组成的编队(100),其中每个小型卫星(10)的相对位置和飞行轨迹T<Sub>k</Sub>可由独立且自主工作的防撞系统(18)进行修改。(The invention relates to a minisatellite capable of flying in formation (10) for LEO application, in particular a nano-satellite or pico-satellite with a mass of 10kg or less, comprising a housing (12) and at least one insert plate (14) with a predetermined function arranged in the housing (12), and a method for flying in a flight trajectory T k A propulsion system (16) generating a directional impulse in the direction of (a). The small satellite (10) comprises an independent and autonomous working collision avoidance system (18), the collision avoidance system (18) being capable of adjusting the trajectory T by means of the propulsion system (16) in anticipation of a collision with an object (30) k Track correction of (T) kk . In another aspect, the invention relates to a flight system consisting of several flying trains (10)Formation (100) of mini-satellites, wherein the relative position and flight trajectory T of each mini-satellite (10) k Can be modified by a separate and autonomously operating collision avoidance system (18).)

1. A small satellite (10) capable of flying in formation for LEO applications, in particular a nano-or pico-satellite with a mass of 10kg or less, comprising: a housing (12) and at least one insert plate (14), said at least one insert plate (14) being arranged in said housing (12) and having a predetermined function, and a propulsion system (16) for a trajectory T in a track_kCharacterized in that said small satellite (10) comprises an autonomous and independently operating collision avoidance system (18), said collision avoidance system (18) being capable of adjusting the flight trajectory T by means of said propulsion system (16) in anticipation of a collision with a foreign object (30)_kTrack correction of (T)_kk。

2. Small satellite (10) capable of flying in formation according to claim 1, characterized in that said independent and autonomous collision avoidance system (18) comprises object detection means (20), collision prediction means (22) and avoidance means (24), wherein said object detection means (20), collision prediction means (22) and avoidance means (24) are providedThe object detection device (20) has at least one or more optical or radio-based object detection sensors (84) for detecting the relative position and relative speed of a foreign object (30) in the direction of a conical impact tube (26), the conical impact tube (26) comprising the trajectory T (84)_kWherein the collision predicting means (22) is for determining a possible collision risk in the collision tube (26), and wherein the avoiding means (24) is for correcting T with respect to the trajectory_kkTo control the propulsion system (16).

3. Microsatellite (10) capable of flying in formation according to claim 2, characterised in that said object detection device (20) autonomously selects the minimum diameter d of said crash tube (26)_kIn such a way that at least the minisatellite (10) is included, in particular at least a multiple of the diameter of the minisatellite (10), and the collision prediction device (22) assigns a foreign object tube (32) to a foreign object (30), and selects a minimum diameter d of the foreign object tube (32) in such a way that at least the foreign object (30) is included_fIn particular in a manner that includes at least a multiple of the diameter of the object (30), wherein preferably the opening angle of the collision tube (26) and/or the foreign object tube (32) is selected as a function of the relative speed between the minisatellite (10) and the foreign object (30), and in an overlap region A of the collision tube (26) and the foreign object tube (32)_kfIn the case of medium overlap, the avoidance device (24) determines a trajectory correction T_kkIn particular trajectory correction T for simulating several trajectory corrections by means of a cost function and determining a minimum cost_kkAnd controlling the propulsion system to follow the trajectory correction T_kk。

4. Microsatellite (10) capable of flying in formation according to one of the preceding claims wherein said insert plate (14) comprises a plurality of functional cores (34) for providing predetermined functions, in particular an even number of at least two or more comparable functional cores (34) for redundantly providing functions, wherein a monitoring device (36) monitors the corrective action of said functional cores (34), wherein preferably said monitoring device (36) monitors the function of said functional cores (34) by testing a functional sequence and when a fault is detected said monitoring device (36) selects a fault correcting activity of one or a group of functional cores (34) for providing functions continuously and uninterruptedly.

5. Microsatellite (10) capable of flying in formation according to claim 4, characterised in that said monitoring means (36) perform FDIR algorithms (fault detection, fault isolation and fault recovery techniques) and in case of a fault perform a power reset, a switching between preferably two functional cores (34) and/or a software reset of at least one of said functional cores (34).

6. Microsatellite (10) capable of flying in formation according to one of the previous claims, characterised in that said propulsion system (16) comprises at least one reaction wheel (40) and at least two magnetic field coils (38) in a magnetotorquer device (116) for combined attitude control in any direction, preferably comprising one reaction wheel (40) and at least three, in particular six magnetic field coils (38).

7. Microsatellite (10) capable of flying in formation according to claim 6 characterized in that at least two of the group of at least one star sensor, at least one sun sensor, at least one gyroscope, preferably one MEMS gyroscope, and/or at least one magnetometer, in particular one 3D magnetometer, one 3D gyroscope, six two axis sun sensors and six two axis star sensors are arranged on the backside of the insert plate (14) and/or on one or more housing walls (46), wherein preferably said reaction wheel (40) is arranged on said insert plate (14) as a micro reaction wheel for attitude control correction.

8. Microsatellite (10) capable of flying in formation according to one of the preceding claims, characterized in that said propulsion system (16) comprises at least one electric propulsion, in particular arc jet, preferably at least one FEEP (field effect electric propulsion) thruster (50), and in particular four FEEP thrusters (50), said four FEEP thrusters (50) further being preferably arranged in an edge region or corner region (52), preferably in a casing frame (44), of a preferably cubic casing (12).

9. Microsatellite (10) able to fly in formation according to claim 8, characterised in that the edge area of the casing frame (44) comprises a hollow profile or a profile with a porous internal structure in which fuel for electric propulsion, in particular gallium, ammonia or hydrazine, is stored.

10. Microsatellite (10) capable of flying in formation according to one of the preceding claims, characterised in that a housing frame (44), several housing walls (46) and a floor (48) with at least two insert plates (14) are comprised in said housing (12), wherein said insert plates (14) are inserted in said floor (48) and said insert plates (14) are connected via a data and power bus (56) with a further insert plate (14) and/or said propulsion system (16) and/or at least one sensor and/or actuator device (140) and power supply device (60), said data and power bus (56) supporting communication protocol standards UART, SPI, CAN, space line and/or I²C, especially a plurality thereof.

11. Microsatellite (10) capable of flying in formation according to claim 10, characterised in that the backplane (48) comprises a multi-stage and scalable power supply means (60), said power supply means (60) being powered by at least one photovoltaic cell (66) and/or at least one accumulator (64) as energy source (62), wherein various voltage converters (68), charge controllers (70) and energy monitoring means (72) and energy switching means (74) are included to detect, distribute and control both the energy output from the energy source (62) and the energy consumption from the energy sink (76), in particular the energy consumption of the patch panel (14) or the propulsion system (16).

12. Microsatellite (10) capable of flying in formation according to claim 10 or 11, characterised in that said backplane (48) is planar and an I/O board (78) having at least one analogue and/or digital interface port (82) is insertable into a mating side (80) of said backplane (48) having a plurality of plug-in sockets (54), wherein preferably said mating side (80) comprises a data and power bus (56) and said plug-in sockets (54) and said backplane comprises an interface to a housing wall connector (58) at the side to electrically connect with said housing wall (46).

13. Microsatellite (10) capable of flying in formation according to one of claims 10 to 12, characterised in that said insertion board (14) is at least one communication board (COMM), one computer board (OBDH), one power board (EPS) and/or one attitude determination and control board (ADCS), and in that at least one housing wall (46) comprises at least one photovoltaic cell (66) and/or at least one magnetic field coil (38) and/or at least one optical sensor (86) and/or one antenna (88), and in that said housing frame (44) comprises at least part of said propulsion system (16), in particular at least one FEEP propeller (50).

14. Formation (100) consisting of mini-satellites (10) capable of flying in formation according to one of the preceding claims, characterized in that the flight trajectory T of each mini-satellite (10)_kAnd the relative position can be varied via an independently and autonomously operating collision avoidance system (18).

15. A convoy (100) according to claim 14, characterized in that the flight trajectory T of each mini satellite (10) is such that_kExceeds the formation track T_STrack correction of (T)_kkCan execute a specific trajectory T of said minisatellite (10)_kkOr other small satellites (10) located in the formation (100)Formation trajectory correction T_SkFor maintaining or replanning the formation track T_SWherein preferably said minisatellites (10) exchange their relative positions and/or their trajectories T bidirectionally_k。

Technical Field

The present invention relates to very small satellites capable of being formed into formation flight, in particular nano-or pico-satellites with a mass of 10kg or less for LEO (low earth orbit) applications, and the formation of several small satellites according to the independent claims.

Background

Since the first satellite of the earth, which was launched by a mass of about 80kg in 1957, small satellites were known as the prior art. Due to the rapid development of technology, it is now possible to realize small satellites with a transmission mass below 10kg, preferably less than 2kg, into a Low Earth Orbit (LEO) around the earth at an altitude between 200 km and a maximum of 1000 km.

In terms of space technology, the obvious trend is that a traditional large satellite will be supplemented by a satellite constellation consisting of distributed small satellites. These will be placed on a low earth orbit with high inclination to achieve global coverage and low latency satellite systems. In particular, near the poles, a given orbital mechanics leads to a particularly high concentration of satellite channels, which requires real-time detour maneuvers to avoid collisions. The combination of the large number of existing objects in the early days of space flight, which are no longer operational, with the large number of modern distributed satellite systems requires new methods to ensure safe satellite operation in orbit, which will be patented here.

LEO orbits can be reached by inexpensive transmission systems where the relative velocity of the satellite with respect to the earth's surface to compensate for the central gravity is important and therefore takes about 100 minutes to orbit the earth once. At this point, radio contact with a ground station above the horizon is for a maximum of 15 minutes. The common minisatellite LEO is used in particular for earth or weather observation, for radio broadcasting or in global satellite communication systems, and for research or technical demonstrations. In addition, such satellites may be used for surveillance in military environments or for local analysis of the earth's surface, for example for environmental monitoring, including storm or flood observation.

In LEO satellite formation, multiple satellites are typically deployed in circular orbits at different locations and take advantage of symmetry. There may be several separate orbital planes that are offset from each other and occupied by multiple satellites. In particular, the Walker constellation is often used here. The usual polar orbits intersect near the poles, so there is a particularly high risk of satellites colliding in the cross-orbits in the polar region. Each orbital plane may be occupied by one or more satellites. In particular, these satellite constellations with orbits characterized by inclinations between 50 ° and 100 ° are of interest for communication systems (usually implemented as Walker-Delta constellations). Furthermore, a highly elliptical configuration with a useful life is typically implemented in which the satellite passes near the earth in only a portion of its orbit and is further from the earth in the remainder of its orbit.

Recently, for LEO applications, the number of proposed satellite constellations has increased significantly. When the ad-hoc control based on-board sensor measurements keeps the satellites in a constant distance topology, this is called satellite formation. In many cases, the constellation of satellites is used for global coverage, such as in satellite navigation or telecommunication systems, and in other cases also for dense local coverage, with a constellation of satellites cooperating in the vicinity being used to map and view the earth's surface. Thus, the earth's surface can be uniformly covered by a formation of satellites with high space-time resolution.

The small satellite can realize series production, has low cost and can be matched with a traditional satellite to be launched into the orbit. A general class of pico-and nano-satellites with an emission mass of less than 10kg, which are available for many applications, was defined by the end of the nineties by the definition of the cubic satellite standard. The miniaturization, energy efficiency, modular structure and increasing autonomy of these satellite systems play an important role and are constantly being improved.

However, cubic satellites are limited in their ability to supply energy and store fuel for manipulation. This technique exhibits limited performance and accuracy for satellite formation. Another critical aspect is the life of the mini-satellites, which also means high failure rates, since system redundancy is generally not possible due to quality limitations. In particular, a critical drawback is limited fuel and radiation damage, since conventional radiation protection using shielding against hostile space radiation cannot be accommodated in a limited volume. In particular, for satellites with masses below 2kg to about 1kg, redundant systems cannot be installed, so that only one functional system can usually be accommodated on a single board.

Recently, the importance of collision avoidance strategies for micro-satellites has been enhanced by the increase in the number of space fragments and the formation of satellite fleets, particularly the number of satellite constellations in polar orbits. Recently, the rate of losses caused by collisions with objects in space (e.g., space debris, other satellites, or satellites in the same formation) has increased significantly. Up to now, in orbit, multiple satellites are almost implemented in the form of constellations, where each satellite is individually controlled by a ground control station. In addition, objects in the flight path are also detected by the ground station and included in the path plan. However, there is no connection to the ground station for about 90% of the orbit period, so the planning must be done in advance for the constellation-related part of the orbit. This is not feasible in the future in view of the expected large increase in the number of objects. In order to shorten the response time of a collision maneuver, an orbit control method is implemented directly in the on-board data processing system in order to determine the relative distances to other objects (cooperative and uncooperative) on the satellite to determine and autonomously implement (optionally in cooperation with cooperative objects) strategies for avoiding a collision. Therefore, there is an urgent need to develop small satellites, particularly with respect to increasing satellite density, collision avoidance strategies and relative navigation capabilities in LEO formation.

Furthermore, there is a need to provide a flexible hardware design for small satellites to be used in different problem areas, which hardware design must exhibit high reliability and operational lifetime.

Furthermore, the testability of such satellites is simplified, especially in mass production and standardization of hardware design, especially the basic equipment of small satellites, to reduce production costs and cycle times.

Finally, high fault tolerance and robustness to radiation failures and faults, as well as energy-efficient designs for adding small satellite functionality within limited energy resources, are to be developed.

JPH07-89497a shows a satellite with a collision avoidance device. Through buffer stop, can avoid the collision with different objects in the track. The collision avoidance system includes a microwave-based one-dimensional distance sensor for measuring the scalar distance between the satellite and other objects in orbit. In addition, means for collision prediction are described to determine potential collision risk and for controlling the detour means of the propulsion system. However, adaptive collision determination is not described, since only the direct distance to the object in orbit can be measured, and the direction from the satellite to the object cannot be determined. Therefore, no indication of the flight trajectory of the object can be derived, and therefore no energy-efficient circumvention maneuver strategy can be implemented. Thus, with parallel flight trajectories, as considered for satellite constellations, distance-based collision avoidance alone results in unpredictable motion, significant energy waste, and complete formation losses. For the special case of very small satellites with very limited fuel storage, in addition to distance, it is also necessary to know the direction towards the object in order to achieve collision avoidance in an energy-saving manner. Furthermore, this document does not outline how such a collision avoidance device can be implemented in very small satellites. In particular, how effectively it can be implemented for small satellites in a constellation, and how a constellation of interdependent satellites uses pure distance measurements as input to implement a collision avoidance strategy.

EP3095713 describes a carrier plate for a device for satellites, which may comprise an attitude control system, a storage container and/or a radio system. Each panel is equipped with a main bus comprising means for power control and a converter for solar cell power. In addition, a plug called ARINC is provided as an electrical interface plug-in system for the panel. In this case, at least the individual and specific assignment of the individual functions of the satellite to a specific board cannot be identified, and additional cables are still required.

In DE202005015431, U1 describes a reaction wheel for a microsatellite comprising a unit with a magnetic rotor and an inertial mass. In this case, the combination of the field coils does not describe a combination for relative attitude control. Furthermore, the necessary sensors for attitude alignment are not mentioned, such as gyroscopes, star sensors or sun sensors, so it remains unclear how to use such reaction wheels to control the relative attitude without interacting with other actuators such as magnetic field coils (for generating fields in case of wheel saturation) and how to use such propulsion systems and sensors to power save the overall control of attitude and direction of motion.

An electric drive system for a micro satellite is shown in EP0903487a 2. The resistive element is positioned adjacent to or within the chamber. When the fluid is introduced into the chamber, the fluid expands due to the thermal energy generated by the resistive element, thereby achieving gas emission propulsion or arc jet propulsion. If the pressure in the chamber rises to a certain pressure, the membrane ruptures, causing fluid to flow out of the chamber, so that the propulsion system is well suited for single use, but not for repeated attitude control actions. For such micro propulsion systems, pressure is increased by heating and impingement is generated by the jet of gas, but no details are described about the reuse of FEEP propulsion.

In DE102010045232a1, a formation of several miniature satellites is proposed, wherein the relative position and orbital trajectory of each miniature satellite is adjusted by an independent and autonomously operating position control system. Preferably, all individual satellites are equipped with a position control system. With an attitude control system, each individual satellite can target a region on the surface of the earth, and with a position control system, the distance to other satellites can be detected and adjusted. In addition, by means of a position control system, arbitrarily variable formation can be realized. However, collision avoidance with an adaptable collision tube is not described herein, and the system is only able to navigate relative to similar satellites, but is unprotected against collision with foreign objects. The publication does not provide any evidence of how such a change in position or attitude is achieved, and does not suggest a collision avoidance feature.

The present invention relates to the design of a microsatellite which can be flown in formation, in particular a microsatellite having a mass of less than 10kg, in particular a microsatellite in the range of 2kg to 1kg or less, for LEO applications which meet the above outlined requirements. In particular, a small satellite capable of flying in formation and formation of these small satellites are proposed, enabling autonomous manoeuvres for collision avoidance on the basis of relative navigation.

Such small satellites that can fly in a formation and the formation of such small satellites are the subject of the independent claims. Advantageous developments of the invention are the subject matter of the dependent claims.

Disclosure of Invention

According to the invention, a small satellite for use in an LEO, which can be flown in formation, in particular a nanosatellite or a picosatellite having a weight of 10kg or less, is proposed, comprising a housing and at least one plug-in board (plug-in board) arranged in the housing, preferably having a predeterminable function and being intended for use in a trajectory T_kA propulsion system generating a thrust in the direction of (a). It is proposed that the microsatellite comprises an autonomous and independently operating collision avoidance system capable of adjusting the flight trajectory T by means of a propulsion system in anticipation of a collision with a foreign object_kTrack T of_kk. The flying object may be any type of foreign object, in particular a passive object, in particular a space debris, micrometallite, asteroid or other passive object. The flying object may also be another small satellite or a part of a rocket with autonomous control or a single satellite that can fly in the same or a different formation. In this case it would be advantageous if not only the small satellites that can be flown in formation include a collision avoidance system, but also other active bodies in orbit include a similar collision avoidance system, and in the best case a two-way data exchange is established between the objects in orbit and the small satellites to achieve collision avoidance by trajectory correction, so that the maximum distance and the minimum possible collision probability can be provided.

Collision avoidance with known satellite constellations by remote control from ground stationsIn contrast to the prior art, an autonomous operating system within a small satellite is proposed that can initiate collision avoidance behavior without contacting a ground station to avoid collisions. Since in LEO applications, a small satellite has its orbital visibility to the ground station for only a short time, not one hundred percent collision avoidance is achieved by a constellation-based collision avoidance system. Autonomous operation collision avoidance systems within small satellites are capable of autonomously performing a pair of trajectories T at any position in orbit with respect to active and passive bodies in orbit_kkTo reduce the likelihood of collisions and to maintain the life and functionality of the satellite fleet. Particularly for polar satellite orbits near the pole, the probability of collision of satellites that can fly in formation is relatively high, since they are very close to each other in the environment around the pole area. Especially in this area, it is important to avoid collisions with autonomous operating systems, since the density of possible ground stations there is very low.

Similar to the crash situation in road or air traffic, but in space different sensors and environmental conditions (vacuum, low temperature.) must be considered. The implementation of avoidance maneuvers is based on completely different dynamics compared to road and air traffic. Although road transport is determined by the friction of the two-dimensional earth surface and the wheels during movement, and air traffic is essentially dominated by aerodynamics, in space the gravitational force is the dominant force, which determines the three-dimensional reaction obtained with respect to the movement of the propulsion system with respect to position changes.

From the relative distance data between the measured objects in space, predictions are derived from future paths and potential collision probabilities with other known objects, by means of a suitable orbit model of the satellite. For relative distance measurement, optical and radio measurement methods may be used. In addition, new measurements of the object and their trajectory predictions will be included. Thus, according to the current dynamics, the trajectory will be determined, which can be achieved without collision with other objects. Corresponding attitude and position control activities are determined and implemented. Objects that can be contacted via a communication link use a self-organizing process by a closed control loop from the sensor to the actuator via the communication link in order to safely coordinate the trajectories with respect to each other.

According to the invention, the function of the conventional wiring harness is replaced by a backplane, wherein one or more plug-in boards, each for a specific purpose, are used so that all power distribution and data transmission links of the satellite are realized in one board, and the subsystem boards are effectively connected by being suitably plugged into the backplane. Neither large nor small satellites have achieved this approach.

In an advantageous embodiment, the collision avoidance system can comprise an object detection device with at least one or more optical or radio-based object detection sensors, in particular an object detection device with microwave-based object detection sensors, for detecting the relative position and speed of a foreign object in the direction of a conical impact tube, which comprises the trajectory T_k. Furthermore, the collision avoidance system may comprise collision prediction means for determining a potential collision risk in the collision tube and for correcting T for the trajectory_kkControls the avoidance device of the propulsion system. Here, it is proposed that the collision avoidance system comprises an object detection device which is equipped with an optical-based, radio-based object detection sensor, in particular a microwave-based or radar-based object detection sensor, and which is capable of identifying active and passive objects in the track. The object detection sensor thus detects objects in the trajectory in the direction of a crash tube, which has a predeterminable size and possibly a predetermined spread angle (opening angle), and determines other objects in the crash tube. By means of a collision prediction device, which can determine the trajectory of an object in the trajectory and can detect an intersection with the collision tube, the risk of a collision can be determined, and control commands for the propulsion system can be derived by the collision avoidance device to perform trajectory correction T_kkThereby providing the smallest possible collision risk with the smallest possible energy consumption. Therefore, collision avoidance systems are based on identifying objects in the track with active object sensors, such as optical sensors, cameras or electromagnetic sensors, for example microwave, IR sensors orRadar detection to specifically identify passive objects. The object detection means may further be coupled to transmitting and receiving sensors which communicate with other objects that are active, in particular other small satellites that can be flown in formation, in order to retrieve their current position and their current trajectory. In particular, such communication links may exhibit a limited directional transmit and receive domain around a small satellite to communicate only with nearby active foreign objects. The collision prediction means may determine the risk of collision if an object in the trajectory is identified as a trajectory oriented towards the collision tube direction. Depending on how the trajectory of the object in the trajectory is oriented in the direction of the collision tube. In the case of a trajectory of an object in the trajectory intersecting the collision tube, the risk of collision can be deduced. In this case, the circumvention device may determine the trajectory T_kkIs corrected by the trajectory T_kkDefining a trajectory T_kTo eliminate collisions with the least possible effort in attitude and propulsion energy.

In a further advantageous implementation of the above-described embodiment, the object detection device may autonomously select the minimum diameter d of the collision tube_kSo as to include at least the mini-satellite, in particular chosen so as to include at least a multiple of the diameter of the mini-satellite. This ensures an environment where the crash tube comprises a small satellite, in order to ensure a certain safety distance when encountering foreign objects and small satellites. Further, the collision predicting apparatus may assign a foreign object tube to the foreign object, and select the minimum diameter d of the foreign object tube including at least the foreign object_fIn particular chosen so as to include at least multiples of the diameter of the object. It is therefore proposed to have both a collision tube comprising a mini-satellite and aligned along the flight trajectory of the mini-satellite and a foreign object tube determined so as to comprise a foreign object oriented in the trajectory direction of the foreign object, in particular a multiple of the diameter of the foreign object. The avoidance device can be arranged in the overlapping area A of the collision pipe and other main pipes_kfDetermining trajectory correction T in case of medium overlap_kkSo as to simulate a particular plurality of trajectory corrections with a cost function and to determine therefrom the trajectory T_kkAnd controlling the propulsion system to effect the trajectory modification T_kk. A least cost correction of the trajectory in the context of the present invention means that the trajectory correction is performed with as little effort as possible to reliably provide collision avoidance with the detected foreign object. This means a low energy consumption for correcting the trajectory with respect to attitude and acceleration thrusts, which maximizes the distance between the collision tube and the tube surrounding the foreign object as much as possible. In this example, heuristic methods such as set-valued functions, control engineering methods, fuzzy logic, or similar simulation and determination methods may be used to achieve the minimum cost for trajectory correction.

Space environments are characterized by intense radiation because, absent the shielding effect of the magnetic layer, the magnetic layer protects the earth's surface from radiation. Typical are single event effects (SEU) and latch-up (latch-up), which affect electronic components, among others. The more compact the electronic components are realized, the more sensitive to radiation effects. Reliable on-board electronics represent a particular challenge, especially for small satellites requiring particularly extreme miniaturization, since conventional approaches prohibit the use of radiation-resistant components (based on electronic components with a particularly thick silicon layer, and therefore very "old" technologies) or shielding by lead plates on the electronic components. In this regard, there is a need for other approaches to ensure reliable operation of small satellites through integrated software/hardware solutions.

For this reason, in a further embodiment the insertion plate comprises a plurality of functional cores for providing predeterminable functions, in particular an even number of comparable functional cores of the at least two or more, for redundantly providing the functions, wherein the monitoring (watch-dog) device monitors the correct operation of the functional cores, and wherein preferably the monitoring device monitors the functions of the functional cores in accordance with a test function sequence, and upon detection of a fault, the monitoring device selects a fault correction action for one or a group of the functional cores for the continuous, uninterrupted provision of the functions.

It is therefore proposed that the insert plate or at least one insert plate of a mini satellite comprises an even number of functional cores, in particular two, four or six functional cores, each capable of performing the basic function of the insert plate. Typically, the functional cores work autonomously and in parallel with each other. A monitoring device continuously and/or periodically monitors the function of each functional core and compares the inputs and outputs and the proper operation of each functional core. The monitoring device simulates a test function sequence in which the function of each functional core can be tested by given the known output parameters of the input parameters. When one of the functional cores is not functioning properly as a result of the monitoring or testing sequence of functions, the function will be shut down and a fault recovery action initiated. These may be a reset or deactivation of the functional core. In addition, the memory contents of the functional core with obvious errors can be reprogrammed or transferred. Therefore, during the failure correcting action, the running program can be rewritten. Furthermore, the interposer board or the functional core may be restarted, wherein errors in the adjustment may occur in both the storage content as well as the processing program. The monitoring device may monitor, for example, a periodic signal of the functional core or may check the specific result of the test functional process. Combinations of different monitoring methods are also conceivable. It can monitor both the CPU and RAM of each functional core, reading in and reading out the predetermined bit pattern and memory area, for example by checksum formation or by means of a software implemented fault injection algorithm (swift). In particular, the error rate can be reduced in this way.

The monitoring means may be hardware as well as implemented software and perform a cascaded monitoring function. In this case, in the case of thermal redundancy, for example during continuous operation, the normal operation of the functional core is continuously monitored. The functional core may operate in a master-slave operation, wherein the order of the master and slave operations may be interchanged in the event of an error. Thus, the local master program can work as a slave in parallel with the local copy of the second region, and the checksum can be supervised by coherency analysis. If a discrepancy or error is found here, a fault recovery step can be initiated, in particular of the main functional core.

In a previously implemented development, the monitoring can implement the FDIR algorithm (fault detection, identification and recovery technique) and in particular activate a power reset, a switching between functional cores and/or a software reset, wherein this is in particular implemented efficiently for two functional core energies in the thermal redundancy and wherein at least one functional core can be reset by software or hardware. The FDIR algorithm corresponds to a self-healing algorithm in which software is able to autonomously detect defects (fault detection, monitoring), determine faults (fault isolation, e.g., shutting down or resetting defective components), and perform appropriate corrections (recovery, e.g., switching to a second functional core or restarting the system). For example, incorrect bits may be detected by checksum generation and may be corrected. As the defect density increases, defective memory locations may be identified and corrected or blocked, or switched to a different functional core or reset. It is possible to achieve an enhanced robustness to faults, not in the case of majority voting, but with a minimum number of redundant functional cores, in particular two, so that radiation resistance is achieved even without the conventionally used shields and lead plates. Thus, the durability of the technical function of the small satellite is realized, the realization of the radiation resistance of the small satellite is largely ignored, and only the function and the reliability of the small satellite are ensured through the software technology.

Therefore, it is proposed to use energy-saving, highly miniaturized components in thermal redundancy, which are monitored by software on "intelligent" monitoring devices. The advanced FDIR software (fault detection, identification and recovery) enables a fast switching procedure to be initiated subsequently to the faultless functioning component and then a restart procedure of the faulty component to be initiated immediately after the occurrence of the radiation effect by the fast detection of a defect, so fast an internal response that no external observer perceives any change in the function of the electronic component. Such an implementation is particularly suitable for high reliability-related electronic components, is particularly important in the field of on-board data processing and attitude/orbit control, and therefore represents a necessary basis for cooperative, distributed, autonomously reacting satellites (e.g. for formation flight).

In an advantageous further development of the miniature satellite, the propulsion system may comprise at least one reaction wheel and at least two magnetic field coils, preferably one reaction wheel and at least three, in particular six magnetic field coils, of the magnetic torquer device for combined attitude control in any direction. The reaction wheel may be miniaturized and may, for example, be arranged on an insert plate. The at least two, in particular at least four and preferably six magnetic field coils may be arranged on the rear side of the housing surface of the miniature satellite. Mechanical thrust along at least one axis for attitude alignment is provided in an axial direction using reaction wheels. The field coils of the magnetic torquer device may enable alignment of the small satellite along the earth's magnetic field, wherein the field coils align themselves in the direction of the earth's magnetic field and thus enable rotation of the attitude. With at least one reaction wheel and two magnetic field coils of 90 ° rotation axis, attitude control can be achieved substantially with a minimum number of components and very low energy consumption for attitude control. Since six housing panels form the sides of a cuboidal minisatellite, it is advantageous for the same configuration of the housing walls to accommodate one magnetic field coil on each housing rear side, providing six magnetic coils and at least one reaction wheel to achieve attitude control of the minisatellite with rapid reaction and minimal energy consumption, with the two magnetic coils aligned in pairs in each axial direction.

In a further development of the above-described implementation of the propulsion system, it may be at least two of the group of at least one star sensor, at least one sun sensor, at least one gyroscope (preferably a MEMS gyroscope) and/or at least one magnetometer, in particular at least one 3D magnetometer, a 3D gyroscope, six two-axis sun sensors and six two-axis star sensors on the interposer board and/or one or more housing walls, in particular arranged on each housing wall, further preferably the reaction wheels are arranged as micro reaction wheels for attitude control correction on the interposer board or motherboard. The general structure of the reaction wheel is basically known, wherein particularly for use in small satellites it is proposed to use a particularly energy-efficient implementation with an energy consumption of 150mW or less at particularly high rotational speeds of 19,000U/min or more. For attitude control in propulsion systems, it is necessary to determine the relative attitude, e.g. the attitude of a small satellite relative to other satellites in the formation and relative to the earth's surface. For this purpose, a set of at least two sensors selected from a star sensor, a sun sensor, a gyroscope or a magnetometer may be used. The sun sensor is capable of determining the direction of the sun relative to the surface of the small satellite. Since the sun is only visible in a portion along the orbit of a small satellite, a star sensor may additionally be included, which can determine attitude relative to the constellation or star. Gyroscopes can determine attitude in space through the gyroscope principle, and magnetometers can determine attitude relative to the earth's magnetic field. By combining the individual sensors, in particular when arranged on the housing walls, it is possible to accommodate at least one solar sensor, at least one gyroscope and a 3D magnetometer and a 3D gyroscope, for example in the form of a MEMS gyroscope (micro-electro-mechanical system), by the same implementation of all six housing panels of the cubic satellite on each housing wall. Such a MEMS gyroscope may use an integrated circuit, in particular three-dimensional and comprising oscillating components, which can recognize accelerations and directional changes. The magnetometer may be implemented as a magneto-resistive semiconductor, in particular a 3D magnetic field compass. The placement of the various sensors on the rear wall of the housing or on the front side of the housing of the mini-satellite is preferred and allows for energy saving and simple attitude control of the propulsion system.

In a preferred further development of the satellite, the propulsion system comprises at least one electric propulsion device, in particular at least one FEEP thruster (field emission electric thruster) or arc jet (micro arc thruster), in particular four electric thrusters, which are preferably embodied as FEEP thrusters, further preferably arranged in edge regions or corner regions of preferably cubic housings, preferably in or on the housing frame. The casing frame comprises as structural elements four separate edges, made of light metal such as aluminum, present and defining a support structure for the casing surface of the small satellite. In the edge element, a FEEP propeller may be integrated, which may generate propulsion power in one direction. A FEEP thruster is a specific form of an electrothermal thruster in which electricity is used to heat a working gas to a high temperature and break it up into charged particles (ions and electrons). Here, the magnetic field spans between the cathode and the anode to a field that injects the charged fuel particles at high velocity. Due to conservation of momentum, the satellites move in the opposite direction to the injected fuel. The required power for establishing the magnetic field can be generated, for example, by solar cells which are arranged on the surface of the housing wall of the minisatellite. The thrust generated is relatively low and in the range of thousandths of a newton, but since the mass of a small satellite is less than 2kg, preferably less than 1kg, a low thrust energy is sufficient to produce attitude correction or trajectory correction, in particular to avoid relative collisions and to maintain the trajectory. Only little fuel is required and the running time of the FEEP propeller is long. A super linear propulsion effect is thus produced, although the total mass is even more reduced due to the associated low performance propulsion compared to the conventional propulsion concept of existing mini-satellite designs, a surprisingly high thrust/mass ratio can be achieved by the very small mass of the mini-satellite despite the relatively low thrust used by the thruster. In the prior art, in most cases of single use, the pressure is increased by heating and the injected gas generates thrust, unlike the prior art, the novel approach is to use acceleration of charged particles in an applied magnetic field, which is completely controllable, repeated many times and can be used at low thrust performance. Preferably, as a refinement of the preceding embodiment, the edge region of the housing frame may comprise a hollow profile or a profile with a porous internal structure, in which fuel, in particular gallium, ammonia or hydrazine, is stored for electric propulsion. Preferably, the fuel is stored refrigerated during launch and can be liquefied upon reaching the target orbit. It is therefore proposed that the four corner portions of the housing frame each have a FEEP thruster at its end portion, wherein the edge portions of the housing frame are designed as hollow profiles in which fuel, in particular gallium, ammonia or hydrazine in solid form at room temperature for the electric thruster, is stored during the launch or initialization phase. It is therefore proposed that the four corner profiles of the housing frame each have an electric thruster at its end, wherein the edge portions of the housing frame are designed as hollow profiles in which the fuel is stored during the launching or initialization phase. Instead of hollow-profile structural elements, porous internal structures are also foreseen, for example metal sponge structures with high rigidity at high hollow space density for fuel storage. The fact that the frozen fuel fills the hollow profile or sponge structure makes it possible to achieve a high mechanical stability, in particular during the launch phase of the mini-satellite. If the satellite is in its orbital position, the fuel may be liquefied and used to generate thrust performance. By this realization, no separate fuel tank is needed, but the fuel is integrated constructively into the mechanically stable part of the housing frame and it is used as a mechanically stable structure in the launch phase.

In an advantageous further development, in the housing, the housing frame CAN comprise various housing walls, in particular six housing walls and a base plate with at least two plug-in sockets, wherein the plug-in boards are plugged into the base plate and communicate via a data bus and a power bus for connecting further plug-in boards and/or a propulsion system and/or at least one sensor and/or actuator device and a power supply device, the bus supporting in particular at least one of a plurality of communication protocol standards, such as UART (universal asynchronous receiver transmitter), SPI (serial peripheral interface), CAN (controller area network), space wire and/or I (universal asynchronous receiver transmitter), CAN (serial peripheral interface), CAN (controller area network), space wire and/or I (interface)²C (inter integrated circuit), in particular a serial information bus. In this way no additional wiring is needed, as it is still prior art, and all power supply and data transmission lines of the satellite are implemented in the backplane. This backplane solution is critical for using microsatellites with a mass of less than 10 kg.

In this embodiment, a modular structure of the electrical system of the small satellite is proposed. The central part is a base plate which comprises at least two, in particular a plurality of sockets, into which respective plug-in boards can be plugged. The add-in boards communicate with each other through data and power buses that take into account at least one or more communication standard protocols. In addition, a space line standard communication protocol can be supported. The space line bus is a field bus designated by the ESA and is capable of transmitting serial and full duplex data at high speed. It has high robustness and low power consumption, and in particular high EMC tolerances, and is compliant to space requirements. This allows multiple patch panels to contact each other through the motherboard in a standardized patch system and exchange data between individual backplanes and between the power supply and the propulsion system and data external to the sensor system of the mini-satellite. Preferably, at least parts of the propulsion system and the sensor system, in particular the underlying system components thereof, are arranged in the housing wall or in the housing frame. Each patch panel may be adapted to different tasks including inter alia Communication (COMM), central data processing system (OBDH), propulsion control system (ADCS) and energy supply system (EPS). Furthermore, a plug-in system (SENS) with one or more add-in boards can be used for various scientific and technical tasks of small satellites, such as radar surveillance, visual monitoring of the earth's surface, providing communication services or similar services. Thus, small satellites in a minimal configuration are fully capable of orbital motion and control, and are particularly configured with autonomously operating collision avoidance systems. By inserting additional function boards, specific functions of the satellite can be provided for different application fields.

In a further development of the aforementioned embodiment, the backplane may comprise a multi-stage and scalable power supply means providing energy in at least one photovoltaic cell and/or at least one storage battery as an energy source, including various voltage converters, charge controllers and energy monitoring and power switches to detect, distribute and control the energy output of the energy source and the energy consumption of the energy sink, in particular of the plug-in board or propulsion system.

In this embodiment, a power system of a small satellite, in particular a pico-satellite, according to the invention is proposed, in which power generation is provided, for example based on solar cells which are arranged on the outside of the housing wall and, furthermore, on the inside, are arranged energy storage means, for example rechargeable batteries or fuel cells or the like. The energy storage device may be charged by a solar cell. Both the solar cell and the energy storage device can provide power to the electrical system of the small satellite, wherein different cascade voltage levels can be provided, which can be individually switched off in case of a fault. This provides redundant energy generation, storage, conversion and distribution. Several DC/DC converters for providing an operating voltage from the can photovoltaic cells may be provided here. For rechargeable energy storage devices, several charge controllers are foreseen for charging and energy output therefrom. The energy storage device may provide voltages at different voltage levels, with an additional DC/DC converter providing the different voltage levels. The DC/DC converters may include high efficiency, energy efficient boost and buck converters. Also, higher or lower voltage levels may be derived from the voltage levels of the further DC/DC converter.

Such an EPS (power system) can supply the various subsystems of the small satellite in various ways. Thus, a photovoltaic cell can be arranged on each housing wall surface, so that six individually operating photovoltaic cells are connected to each other to generate regenerative energy. On the insert plate, one, two or more accumulators may be arranged as energy storage means to store or provide energy. The photovoltaic cells on each housing wall may be divided into two and may include a space between them to accommodate, for example, an attitude sensor. Behind the housing wall, the magnetic field coil may be placed as a magnetic torquer or magnetometer. Magnetometers may be implemented electrically efficiently as integrated electromagnetic semiconductors. Using a DC/DC converter, the higher voltage of the photovoltaic cell can be reduced to a lower voltage to charge the battery or accumulator. For example, lithium ion batteries may be used at capacities of a few amp hours. Here a voltage of about 3.4V to 3.9V may be provided and increased to 5V by a DC/DC converter. In this case, a 3.3V bus and a 5V voltage bus may be provided. The energy distribution may be provided by circuit switches or current fuses. Over-voltage and over-current protection mechanisms are envisioned. In this way, a redundant power system with high reliability can be provided for small satellites.

According to a further embodiment of the above-described variant of the miniature satellite, the backplane may be planar and comprise an I/O board having at least one analog and/or digital interface socket, which is insertable into an insertion side of the backplane having a plurality of plug-in sockets. Here, the insertion side may include data and power buses, and a receptacle for electrical connection with the housing wall is included at the side of the backplane. Thus, the back plane of a small satellite is designated as the base of the patch panel with a flat surface and includes various receptacles. A single I/O board is used for external contact with the electrical system of the small satellite and includes analog and/or digital interface connectors. On the backplane, data and power buses are provided, which interconnect the various outlets. Laterally, for example in the plane of the base plate, it is foreseen to have at least one, preferably two, three or four on different sides of the base plate, connector to the housing wall to connect to the adjacent housing wall, which houses the photovoltaic cell, the magnetic field coil, the solar or star sensor, and the optical sensor, and which can be coupled with the interposer board by means of a power and data bus. In this way, a plug-in system is enabled that adapts the mini-satellite to various tasks. The housing walls may provide mutual contact, for example by flat cable connectors, so that a single connector to the housing frame is sufficient to contact all housing walls.

In a further development of the aforementioned embodiment, the add-on board may comprise at least one communication board (COMM), a data processing board (OBDH), a power board (EPS) and/or an attitude control board (ADCS). The housing wall comprises at least one photovoltaic cell and/or at least one magnetic field coil and/or at least one optical sensor and/or an antenna, whereby the housing frame comprises at least a part of the propulsion system, in particular at least one FEEP thruster.

This embodiment defines a minimum number of plug-in boards comprising at least one communication board-COMM-communication, one data processing board-OBDH-onboard data processing, one energy supply board-EPS-power system and/or one attitude control board-ADCS attitude determination and control system. At least OBDH and onboard data processing systems adopt a double redundancy design, and the electric power system EPS has redundancy and expandability so as to distribute electric energy to all subsystems of the small satellite. The COMM plug-in board is a fully redundant UHF communication subsystem for communicating with adjacent mini-satellites, but also for communicating with ground stations, in particular for receiving control data and transmitting sensor data. The ADCS is used to control the attitude and propulsion of the small satellites and includes, for example, collision avoidance systems.

In another independent aspect, a combination of several small parts is providedFormation of formations of satellites of type, in which the relative position and flight trajectory T of each mini-satellite_kCan be adjusted by an independent and autonomous working collision avoidance system. This makes it possible to control the formation of a local concentration of small satellites designed to detect with high resolution a locally limited area of the earth's surface, or the formation of small satellites distributed on different orbits, for example crossing near the polar region, so that no collisions between small satellites occur. The collision avoidance system can utilize two-way communications between the small satellites to achieve collision avoidance and maintain formation, and can detect passive other objects such as space debris, asteroids, meteorites, or other space vehicles that do not support two-way communications. This allows formation of small satellites in LEO orbits to operate with long life and low risk of failure.

In an advantageous development of the formation, if the trajectory T of the mini-satellite is exceeded_kCan execute the formation track T_STrack correction of (T)_kkOr for maintaining or rearranging the formation track T in the formation_SOther small satellites of the formation trajectory correction T_SkWherein preferably the minisatellites exchange their relative position/attitude and/or their trajectory T bidirectionally_k. In this further development, it is proposed that the flight path T is as initial_kInitiating trajectory correction T by a collision avoidance strategy in the presence of substantial deviations_kkThen, the pair formation track T can be executed_SOr, where appropriate, the formation trajectory T is changed in this way for all the mini-satellites formed in the formation_SSo that the formation can be maintained and thereby collision avoidance is provided for the entire formation. The following are useful and advantageous: if the minisatellites are switched bidirectionally and communicate their relative positions and their trajectories T_kOr their trajectory correction T_kkSo that it can be decided whether or not the formation track T can be changed_SOr if a single trajectory is corrected for T_kkThe avoided satellites are brought back into the formation at their earlier locations so that the formation can be maintained.

Drawings

Further advantages can be seen from the presented figure description. In the drawings, examples of embodiments of the invention are shown. The figures, description and claims contain many combinations of features. The expert will consider these features appropriately and individually and group them into meaningful further combinations.

Shown is:

figure 1 is a first representative schematic of a formation of miniature satellites according to the present invention,

figure 2 is another perspective view of another representation of a formation of miniature satellites according to the present invention,

figure 3 is a schematic view of a representative collision avoidance system for a small satellite according to the present invention,

figure 4 is a schematic diagram of a block diagram of a collision avoidance system in an embodiment of a small satellite according to the present invention,

figure 5 is an exploded view of an embodiment of a minisatellite according to the invention,

figure 6 is a detailed view of the constituent extension stages of an embodiment of a minisatellite according to the present invention,

figure 7 is a backplane of one embodiment of a minisatellite according to the present invention,

figure 8 is a view of the various insert plates and housing components of an embodiment of a mini satellite according to the present invention,

figure 9 is a photograph of an insert board for various functions according to an embodiment of the minisatellite of the present invention,

figure 10 is an onboard data processing system of one embodiment of a microsatellite according to the present invention,

figure 11 is a block diagram of a monitoring device in an embodiment of a minisatellite according to the invention,

figure 12 is a schematic diagram of a power supply system in an embodiment of a small satellite according to the invention.

In these figures, similar elements are numbered with the same reference numerals. The drawings are only examples and should not be construed as limiting.

Detailed Description

In fig. 1a, a formation 100 of microsatellites 10 is shown. Formation 100 includes various ones of several spatially adjacentThe minisatellites 10, wherein each minisatellite is capable of observing the surface 108 of the earth up to the earth's horizon 106 within a single detection zone 102. By combining the individual detection zones 102 of each of the minisatellites 10, the formation covers a larger detection zone 104, which can map a relatively large surface area of the earth. The centroid of the formation 100 follows an orbit 150T above the earth's surface 108_SAnd (6) moving.

In this figure, the formation 100 is represented as a spatially finite collection of minisatellites 10 to map a large portion of the earth's surface to the formation's detection area 104 by combining the individual detection areas. Thus, a large area of the earth's surface 108 can be imaged to the earth's horizon 106. A two-way exchange 110 of information of the relative attitude/position and trajectory of the respective satellite 10 is established between the respective satellites 10. As each satellite 10 approaches or encounters a foreign object 30, each satellite, and the entire convoy 100, may then follow a collision avoidance strategy to control direction and individual trajectory in a manner that can avoid the foreign object 30. Thus, the rail 150T_SMay be changed and may then be corrected again to continue the preselected trajectory.

Fig. 1b shows another formation 100 of miniature satellites 10 in polar orbits 150. In this case, in the "string of pearls" 154, the minisatellite 10 orbits the earth on the orbit 150, wherein various orbits 150 along the circle can be provided, and thus various strings of pearls 154 consisting of the minisatellite 10 orbits the earth. At the earth poles 152, the minisatellites 10 in each orbit 150 meet, resulting in an increased probability of collision. In particular, the use of collision avoidance systems 18 near the earth poles 152 is useful for allowing targeted avoidance maneuvering of the various minisatellites 10 with respect to one another. Here, in order to achieve collision avoidance with minimal energy consumption, a two-way exchange of information 110 between the encountered minisatellites 10 in each orbit 150 may be useful.

Figure 2 shows a first embodiment of a small satellite configured in a predominantly cubic configuration. The miniature satellite 10 includes a housing 12, the housing 12 being comprised of a frame of six housing walls 46 and four poles as housing supports 44. Each housing wall 46 has two photovoltaic cells 66 which are spaced apart from one another and in the gap illustrated by an object detection sensor 84 in the form of an optical sensor 86, a star sensor or a sun sensor can be arranged. Furthermore, the space between the photovoltaic cells 66 provides the possibility of connection to an interface port 82 via the I/O board 78 for reading out data or for contact and programming prior to transmission. Furthermore, at least in the corner regions of the box-shaped housing 12, the antenna 88 is intended for radio reception, in particular for UHF reception, which is used for contacting ground stations and adjacent minisatellites 10.

At least four edges of the housing 12 are formed by frame bars 44, the frame bars 44 defining a housing frame. In each frame bar 44, a FEEP pusher 50 may be placed at one end of the bar, wherein each frame bar 44 may have a hollow profile and in this hollow chamber, fuel for the FEEP pusher may be stored. In particular, the fuel may be cooled before being emitted, in order to mechanically fill the hollow frame and contribute to the mechanical stability of the housing frame. In the rail, at the housing frame, a heating device may be intended for heating the fuel for liquefaction, thereby providing fueling for the FEEP thruster 50. Each of the four frame bars 44 forms an independent thruster for the miniature satellite 10, so that both thrust and attitude direction changes of the miniature satellite 10 can be achieved by controlling only a single FEEP thruster 50. By activating all the FEEP pushers 50 simultaneously, it is possible to follow the trajectory T of the microsatellite 10_kA linear push is generated. On the inwardly oriented side of each housing wall 46, a magnetic field coil 38, a magnetic moment device 116, and a magnetometer may be arranged to measure the magnetic field attitude toward the azimuth of the magnetic field or attitude relative to the minisatellite 10. A highly compact design can thereby be achieved which is robust in the transmission phase and which allows the arrangement of all the individual functional components within the satellite to be possible with a minimum of volume and mass.

Fig. 3 shows a first example of an embodiment of a collision avoidance and related navigation procedure. The microsatellite 10 may acquire a foreign object 30 or another microsatellite 10 by an object detection sensor 84 wherein at least a relative velocity and a relative trajectory may be determined. The foreign object 30 may be, for example, a small objectPlanets or space debris, or components such as those of a rocket's burn-out stage or artificial celestial bodies. The bumper system 18 is produced to have a diameter D_kThe diameter D of the collision tube 26_kIs a multiple of the diameter of the minisatellite 10 and therefore includes it the crash tube 26 also includes a flare angle α which flare angle α can be widened to increase or decrease the crash tube depending on the relative velocity between the minisatellite 10 and the foreign object 30 or an adjacent minisatellite 10 the collision avoidance system 18 also determines the trajectory T of the minisatellite 10 or foreign object 30_fAnd is defined to have a diameter D_fIncluding on the one hand the dimensions of the foreign object 30 or of the adjacent mini-satellite 10 and on the other hand a widening angle which can be adjusted depending on the relative speed between the mini-satellite 10 and the foreign object, wherein a high relative speed and/or a reduced distance leads to an increase in the angle of the respective tube 32.

The collision avoidance system 18 may calculate the intersection of the collision tube 26 and the foreign object tube 32 to detect the risk of collision. In this case, by continuing the trajectory T_kAnd assuming that the foreign object continues its specific trajectory T_fIn the case of (2), collision cannot be excluded. Thus, the collision avoidance system 18 determines the trajectory correction T_kSo that the foreign object tube 32 and the collision tube 26 do not overlap. In this way, the propulsion system 16, which includes inter alia attitude/position control and propeller control, enables trajectory correction T to be achieved with minimal energy consumption_k. Collisions with foreign objects 30 that may be very close to the trajectory of the minisatellite 10 may thereby be effectively excluded.

Fig. 4 shows another embodiment of the collision avoidance system 18 of the minisatellite 10. The collision avoidance system 18 includes an object detection device 20, and an object detection sensor 84 (such as an exemplary optical sensor 86) or an antenna 88 may be disposed on the object detection device 20 as a radio or radar sensor. Example optical sensors 86 may be optical cameras and/or infrared cameras. The object detecting device 20 calculates the trajectory (also referred to as trajectory T)_k) The collision tube 26. Further, by the object detection sensor 84, a nearby foreign object 30 can be identified, and the foreign object 30 can be observedDetermining the trajectory T of the distance_f. From knowledge of their own trajectory T_kAnd a foreign object trajectory T_fThe collision tube 26 and the foreign object tube 32 can be determined. In the collision prediction device 22, the collision tube 26 can be compared with the foreign-object tube 32, and from the intersection of these tubes it can be recognized that the trajectory T needs to be changed_kTo avoid collisions.

The collision predicting device 22 identifies the risk of collision, and the avoiding device 24 may determine the trajectory T_kkBy which the attitude/position and thrust corrections are performed with the lowest possible energy consumption and direction change so that the collision tube 26 no longer intersects the foreign object tube 32. The avoidance device 24 determines control information for controlling the propulsion and attitude system 16 (in particular the reaction wheels 40 for changing the relative attitude or the magnetic moment devices 116 for changing the pointing direction to the microsatellite 10) and then applies thrust to the microsatellite 10 through the FEEP thruster 50, thereby effecting the trajectory correction T_kk. Therefore, collision with the adjacent foreign object 30 can be avoided. If the foreign object 30 is another microsatellite 10, the relative attitude and trajectory of the microsatellite 10 can be exchanged, for example, by two-way exchange of information 110 between the microsatellites 10, and coordinated behavior to avoid collisions between the microsatellites 10 can be derived.

Figure 5 shows an exploded view of the minisatellite 10 of figure 2. The miniature satellite 10 includes six housing walls 46 with photovoltaic cells 66 disposed on the housing walls 46, wherein between two adjacent photovoltaic cells 66 a central strip-shaped area is recessed to accommodate the sensor system. An object detection sensor 84, such as a camera or radar sensor, may be disposed in the area. Six housing walls 46 are fixed to four frame bars, wherein each frame bar 44 comprises an arc jet impeller 50 and the fuel is supported in the hollow profile of the frame bar 44. Each frame bar 44 may include an electric heater that allows the fuel in the frame bar 44 to be heated to provide the necessary fuel supply for operation of the arc jet impellers 50. Inside the mini satellite 10, the backplane 48 is arranged with a socket on the mating side 80 into which the interposer board 14 is to be inserted. Each interposer board 14 may handle different tasks and may provide, for example, power, attitude control, coordinate processing computers, or provide communication functionality. In addition, the I/O board 78 plugs into the backplane 48, and the backplane 48 has I/O interface ports 82 for reading data and external programming prior to transmission. The interface port 82 may present, for example, analog and digital connections to read in and read out analog and digital data.

Figure 6 shows in respective figures 6a to 6d the steps of assembly of an embodiment of a microsatellite 10 according to the invention. In fig. 6a, the backplane 48 is shown as having a mating side 80, with the individual receptacles 54 arranged on the mating side 80, which are connected to each other by the data and power bus 56. At the boundary receptacle 54, an I/O board 78 is plugged in, which provides two I/O interface ports 82 for programming access to the mini-satellite prior to transmission. Both the backplane 48 and the I/O board 78 include housing-wall connectors 58, with the housing-wall connectors 58 in a position to electrically contact adjacent housing walls 46 to receive exemplary power from the photovoltaic cells 66 thereon and to contact magnetic field coils that may function as a magnetic torquer or magnetometer.

In fig. 6b, further add-on boards 14 are inserted, in particular an EPS board for providing power supply, an ADCS board for providing attitude and propulsion control, an OBDH board for providing higher-level computer functions and a COMM board providing communication capabilities. In addition, a sensor board SENS for providing sensor capabilities is inserted in the bottom plate 48, for example for earth observation, weather observation and various monitoring functions.

In fig. 6c, another expansion stage is shown in which the frame bars 44 are arranged orthogonal to the bottom plate 48 and parallel to the edges of the interposer board 14. In addition, reaction wheels 40 can be seen, which are located on the ADCS board, and the antenna 88 is connected to the COMM board. The individual functional insert plates 14 are mechanically connected to each other by means of stabilizing elements 28 in the form of screws.

Finally, fig. 6d shows the assembly of the housing wall 46, with a photovoltaic cell arranged on the housing wall 46, and including a further object detection sensor 84 and a housing wall cutout 112, for example for the I/O interface port 82. The housing wall 46 provides regenerative power by means of photovoltaic cells 66 and includes attitude controlled components with magnetic field coils 38 and magnetometers of magnetic moment devices 116 on the rear side of its housing wall and sensor elements for monitoring the surrounding area in the direction of the trajectory for collision avoidance and detection of relative attitude, as well as star and sun sensors. The trajectory direction is opposite to the side of the housing on which the arc jet impellers are arranged.

In fig. 7, three panels show the structure of the backplane 48 and the I/O board 78, which form the backbone of the mini-satellite 10. The base plate 48 has various receptacles 54 into which different insert plates 14 may be inserted. The receptacle 54 is disposed on the mating side 80 of the backplane 48. On the underside of the bottom plate 48, there is arranged a data and power bus 56, which connects the individual contacts of the socket 54 to each other. An I/O board 78 is inserted over the receptacle 54 at the edge of the backplane 48. This includes an I/O interface port 82 for acquiring data from the satellite dish 10 and programming prior to transmission, and for configuring, encoding and testing the functionality of the electrical system. The backplane 48 and the I/O board 78 have housing wall connectors 58 for electrically contacting the housing walls 46 to electrically connect the energy, sensor and actuator systems to the housing walls 46.

In fig. 8, various component sets of the mini-satellite 10 are shown as building blocks. The building block consists of five insert plates 14 and a bottom plate 48, which represents the standard configuration of the mini satellite 10. The backplane 48 has various receptacles 54 that are interconnected by a data and power bus 56. Each interposer board 14 is comprised of an I/O board 78, the I/O board 78 having I/O interface ports 82 for connecting the microsatellite system and OBDH board, EPS board having a battery 64 for providing power supply, ADCS board for attitude control, and COMM board having HF components for transmitting and receiving data via radio waves. Each insert plate 14 has an identically configured connecting strip 114 that may be inserted into the socket 54 of the base plate 48. In addition, four frame bars 44 are arranged, each frame bar integrating an arc jet thruster 50, which is the main propulsion means of the small satellite. Furthermore, six housing wall panels 46 are provided, having two photovoltaic cells 66 on their outer sides with a space between the two photovoltaic cells 66 for accommodating sensors (in particular the object detection sensor 84) and a housing wall cutout 112 for contacting the I/O interface port 82. Attached to the rear wall of the housing wall 46 is a field coil 38 that can be used both as part of a magnetic torquer and as a magnetometer to measure the earth's magnetic field position and to align the attitude/position of the small satellite with respect to the earth's magnetic field lines by means of electrical currents. Through the housing wall cut-out 112, the I/O interface port 82 can be accessed, as well as an outwardly directed sensor, which is for example arranged on the interposer board 14.

In fig. 9a, the insert plate for the ADCS is shown in more detail. This includes a reaction wheel 40 with which an impulse for changing the orientation of the minisatellite 10 can be generated in alignment with the direction of its axis. For this purpose, an attitude-controlled, associated reaction wheel control unit 122 is provided, which may be configured redundantly anyway and may be monitored via the monitoring device 36, to enable an enhanced robustness and radiation tolerance of the ADCS system. On two opposite edges of the interposer board 14, housing wall connectors 58 are arranged, and the data and power buses 56 of the backplane board 48 may be contacted by connector strips 114.

Fig. 9b shows the rear of the housing wall 46, which is constructed in a sandwich structure and has an aluminum core 120 in its interior for increased stability, for cooling and for shielding. The aluminum core 120 serves to increase mechanical rigidity and to a small extent protect against radiation and dissipate thermal energy. On the front side of the housing wall 46, two photovoltaic cells 66 are arranged, while on the rear side thereof, the magnetic field coil 38 is arranged, which is connected to a control unit 118 for the magnetic coil. The control unit 118 for the magnetic coils operates the magnetic field coils 38 as a magnetic torquer device 116 and thus effects attitude alignment of the microsatellite along the earth's magnetic field, and also operates the magnetic field coils 38 as magnetometers to provide magnetic attitude sensors. Mechanically enhanced aluminum cores 120 appear at the four corners of the housing wall 46 to provide mechanical stability of the housing wall 46 for power and attitude control of the small satellite 10 by means of the magnetic torquer devices 116. At least on one longitudinal side and one transverse side of the housing wall 46, the housing wall connector 58 is arranged in electrical contact with the adjacent housing wall 46 and with the base plate 48 and/or the internal insertion plate 14.

In fig. 10a, a perspective view of an insertion plate 14 of an OBDH, e.g. an onboard data processing system, a supervisory process computer, is shown. The OBDH includes two functionally identical functional cores 34a, each functional core 34a having its own independent storage device 134a and 134 b. The monitoring device 36 monitors the correct functioning of the two functional cores 34a and 34b, for example working in master-slave operation, and can switch between the two functional cores 34a and 34b in a thermally redundant operating mode and reset the two functional cores 34a, 34b to ensure reliable operation. Contact with the data and power bus 56 of the backplane 48 is made through the connector strip 114.

A block diagram of the functional distribution of the OBDH is presented in fig. 10 b. The two functional cores 34a and 34b communicate with each other and exchange data and store information. The supervisory monitoring device 36 monitors the input and output data of the two functional cores 34 and their correct operation and may initiate a test sequence to detect indications of inconsistent behavior of the functional cores 34a, 34 b. In this case, one or both functional cores 34a, 34b may be restarted, for example, may be reset to hardware or software, or the results of both functional cores 34a, 34b may be corrected. The functional cores 34a, 34b may operate in a master-slave mode, or individually and autonomously, but may also operate in synchronization, in parallel, and independently of each other. The functional cores 34a, 34b are connected to the data and power bus 56 via an interface device 132. At the data and power bus 56, there may be further connected a storage device 134, a time and clock generator device 136, an interface 138 to the rear wall of the housing panel and a sensor or actuator device 140, such as the propulsion and attitude control system 16, and a radio or optical sensor. In any event, some of these components may also be placed on the insert plate 14. The functional cores 34a, 34b of the monitoring device 36 may be functionally decoupled from the on-board data system and restarted. Here, the monitoring device 36 may run the FDIR algorithm to achieve high robustness and radiation tolerance even without conventional shielding techniques against spatial radiation.

The monitoring means 36 can operate in several levels and the individual functional cores are reset by software, for example in a first level. In the second stage, a hardware reset may be initiated, for example by a short-term interruption of the power supply of the functional core 34 or one of all the functional cores 34a, 34 b. In the third stage, the output results of the functional core 34 are monitored via software so that increased flexibility can be provided at different levels.

Fig. 11 schematically shows a so-called switch monitoring unit (TWU) as the monitoring device 36 of the minisatellite 10. In this case, the output of the functional core 34 is monitored and if it does not emit an active signal within a predetermined time frame, it is assumed that the functional core 34 has crashed. In this case, the reset of the failed functional core 34a or 34b is performed, and finally the master-slave configuration is replaced, so that the current master core becomes the slave core, and the previous slave core serves as the master functional core. Through the same logic, the interface device 132 is activated to connect the respective functional cores 34a, 34b to the data and power bus 56. TWU of monitoring device 36 in this case includes logic gate unit 146, monitoring unit 142 and FPGA unit 144 for providing switching actions. This allows the OBDH to monitor itself and recover itself in the event of a failure.

Finally, fig. 12 shows a power supply device 60ESP of the mini satellite 10. The power supply apparatus 60 includes a photovoltaic power system 124, a battery power source 126, and an energy control subsystem 130. By the separate architecture of the power supply means 60, a high robustness of the power supply of the various functional components of the mini-satellite 10 can be achieved.

The photovoltaic power system 124 includes an energy source 62, one or more photovoltaic cells 66. The photovoltaic cells 66 are energized at different voltage levels by various DC/DC voltage converters 68. In this case, an energy monitoring device 72, such as a current or voltage monitoring device 72, may determine the amount of energy delivered by the photovoltaic cells 66. The electrical energy is forwarded to the battery power supply 126. This includes two or more batteries 64 arranged on the ESP insertion plate, which are charged by the power of photovoltaic cells 66 in order to provide energy, for example to power supply device 60 in the earth shadow. An energy switching device 74 is provided to turn the battery 64 on and off. Different levels of output voltage may be provided by the cascaded voltage converters 68. In this case, the battery power source 126 includes two batteries 64 operating in parallel, which may provide power bi-directionally, in parallel, and independently of each other. Power is delivered to the energy control subsystem 130 at three voltage levels. With more power switching devices 74 powering the various subsystems of the small satellite. Furthermore, other voltage levels may be derived from the individual voltage levels with the further DC/DC voltage converter 68. Thus, different independently operating power circuits are provided, wherein energy can be extracted from the photovoltaic cells 66 as well as from the storage battery 64. Different voltage levels are provided so that different subsystems can be provided with independent and different voltage levels. Even if a defect associated with one voltage level occurs, many energy switching devices 74 and voltage converters 68 may be bridged by the voltage conversion and disconnection of the affected energy circuit. Thus, continuous operation is ensured even in the event of failure of one or more storage batteries 64 or one or more photovoltaic cells 66, or in the event of short-circuiting of one or more functional elements of the microsatellite 10.

The small satellite according to the invention can achieve a limited power reserve during its lifetime. This is achieved in particular by the multi-stage energy supply concept. In the area of redundancy and fault tolerance, "majority voting" can be ignored because advanced FDIR technology can only exploit the redundancy of two complementary working systems. For example, a memory area in a database may also be corrected or an error may be detected therein. During the run time of the system, a switch between the master system and the slave system can be performed, already making it possible to implement a switch of the failed system without significant delays, thus avoiding any operational disturbances of the mini-satellite. With the FDIR-based monitoring device, even without the conventional shielding of the mini-satellites, high operational reliability and radiation resistance can be ensured by a plurality of only two redundant functional cores, so that small mass can be achieved.

Powerful miniature satellite systems capable of flying in formation can be provided with small volume, small mass and low energy. Commercially available electrical components are used which are not radiation resistant. The propulsion system can be minimized by a new type of propulsion system based on arc spraying, in which the fuel is contained in the structural parts of the casing. The precise attitude control system can detect the relative position by means of magnetometers, sun and star sensors and gyroscopes, in combination with a propulsion system, which can ensure a long life of the inventive collision avoidance system according to the invention on its track.

The mini-satellite is designed as a modular system, similar to the modular system in automotive production, and can be provided for different tasks at low cost and easily by a basic configuration. With a high proportion of identical components, a low cost per component can be achieved, so that the proposed satellite platform exhibits extremely high durability, low cost and high flexibility in use. Due to relative navigation and collision avoidance, a multi-satellite system can be implemented that can operate autonomously without contact with the ground to accomplish desired tasks. No lead plates are required for radiation shielding and specialized development for different satellite missions can be avoided.

Small satellite systems are characterized by their self-organizing capabilities and high robustness. By using industry standards such as energy and data bus standards, industry available standard components for miniaturized systems and components can be used in difficult off-ground environments with high interference levels. The modular system architecture of each component supports flexible integration and production. A single miniature satellite can be produced in an automated manner, for example by means of a robot. Through automatic testing, the function and the performance of the miniature satellite before transmission can be ensured.

Thus, a low cost distributed satellite system for different purposes may be provided. These can be used, for example, for mapping, for locating tasks or for different tasks in the IT department. Furthermore, such small satellite formations may be used in commercial enterprises, for example for fleet management or for remote maintenance systems or for government tasks such as early warning systems, earth reconnaissance after environmental disasters or military applications. By means of this satellite formation, a highly secure and highly miniaturized system is provided, which can be applied in particular in telematics systems and industrial environments, but also for remote diagnostics and remote maintenance of mobile and stationary equipment. Possible applications are, for example, the automotive industry, the positioning and autonomous driving of fleet vehicles, in global automation and production logistics, in particular for mobile systems, remote control and positioning in the military field and earth observation, in research and space exploration, which offer the opportunity for cost-effective innovative tests to be carried out under extreme conditions, for example, for data providers whose occupancy analysis of parking lots, roads or traffic systems is used as an indicator of economic trends or weather service providers may be potential customers.

List of reference numerals

10 small satellite capable of flying in formation

12 casing

14 insertion plate

16 propulsion system

18 collision avoidance system

20 object detection device

22 collision prediction device

24 circumvention device

26 Collision tube

28 stabilizing element

30 foreign body

32 foreign body pipe

34 functional core

36 monitoring device

38 magnetic field coil

40 reaction wheel

42 shell

44 frame bar

46 casing wall

48 bottom plate

50 propulsion device, propeller

52 corner region of the housing

54 socket

56 data and power bus

58 housing wall connector

60 power supply device

62 energy source

64 accumulator

66 photovoltaic cell

68 Voltage converter

70 charging controller

72 energy monitoring device

74 energy switching device

76 energy well

78I/O board

80 mating side

82 interface port

84 object detecting sensor

86 optical sensor

88 antenna

Formation of 100 mini-satellites

102 individual detection zones

104 formation detection area

106 earth's horizon

108 earth's surface

110 two-way information exchange

112 housing wall cutout

114 connector strip

116 magnetic moment device

118 control unit for magnetic coil

120 aluminum core

122 reaction wheel control unit

124 photovoltaic power system

126 accumulator power supply system

128 voltage control

130 energy control subsystem

132 interface device

134 storage device

136 time and clock generator device

138 interface of rear wall of housing panel

140 sensor or actuator

142 monitoring unit

144 FPGA cell

146 logic gate unit

150 orbit

152 earth pole

154 Pearl string

156 insertion contact row for a plug-in socket

T_kTrajectory of satellite k

T_kkTrajectory correction

T_SFormation track

T_SkFormation trajectory correction

COMM communication board

OBDH computer board

EPS power panel

ADCS position control panel

SENSOR PLATE FOR SENS EARTH-OBJECTION