Spacecraft black box full life cycle power supply management system

Abstract

The invention relates to the technical field of spacecraft safety, in particular to a spacecraft black box full-life-cycle power management system. Comprising a bus, a plurality of solar cell arrays, a storage battery pack, a bus voltage sampling circuit, a plurality of shunt circuits and a filter circuit, and the solar cell arrays and the shunt circuits are the same in number and are in one-to-one correspondence; the solar cell arrays are respectively connected with one end, far away from the load, of the bus; one end of the bus is connected with a second relay switch, a third relay switch and a fourth relay switch, the other end of the second relay switch is connected with the computer module through the power conversion circuit, and the other end of the third relay switch is connected with the Beidou short message module through the power conversion circuit. The fourth relay switch is connected with the UV measurement and control module through the power conversion circuit. The method supports the whole life cycle power management of the spacecraft black box, increases the support of active section work, prolongs the working time of the spacecraft black box, and prolongs the application range of the spacecraft black box.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft safety technology, and in particular to a full life cycle power management system for a spacecraft black box. Background Art

[0002] In recent years, space has become increasingly important as a core area of ​​future strategic competition, and competition has become increasingly fierce. As countries have stepped up their efforts to explore and utilize space, the boundaries of the battlefield are quietly expanding into space, which has made the space security situation increasingly severe and has become the focus of international attention. At the same time, with the frequent increase in human space activities, it has not only promoted the rapid development of space technology, but also brought about a problem that cannot be ignored - the sharp increase in the number of space debris. These fragments fly at high speeds in space, posing a serious threat to the safe operation of spacecraft and increasing the risks and costs of space missions.

[0003] In addition, the space environment is extremely harsh and full of various factors that are unfavorable to spacecraft. Large temperature fluctuations, strong electromagnetic radiation, and high-energy particle radiation all pose potential huge risks to the electronic systems and life support systems of spacecraft, which may lead to system performance degradation, data loss, and even equipment damage. These factors work together to make spacecraft face many safety hazards during their in-orbit operation. If they are not dealt with in a timely manner, they may cause spacecraft to malfunction or even fail, affecting the successful completion of the mission.

[0004] When a spacecraft is operating in orbit, it relies on the measurement and control system to communicate with the ground measurement and control station to ensure the smooth progress of the mission and the safety of the spacecraft. However, once a serious failure occurs in the key parts of the spacecraft such as the measurement and control subsystem, power subsystem or control subsystem, it may cause the interruption of communication with the ground, and the spacecraft will fall into a "lost connection" state. It will be difficult for ground personnel to obtain real-time status information of the spacecraft, thus missing the best time to repair the spacecraft. At the same time, due to the lack of specific data on the ground about the spacecraft failure, such as the operating status before and after the failure, abnormal parameters, etc., it will become extremely difficult to accurately locate the fault point and analyze the cause, further increasing the difficulty and time cost of repairing the spacecraft.

[0005] It is similar to the "black box" devices installed on airplanes and ships to record data during operation and analyze the causes of accidents after they occur. Black box devices are also needed for spacecraft in orbit. Different from black box devices such as airplanes and ships, spacecraft black boxes operate in space, and existing black box devices in other fields are difficult to apply to spacecraft. The spacecraft black box system is mainly used for functions such as spacecraft core operating parameter monitoring, spacecraft fatal fault detection, and emergency communications.

[0006] The Chinese patent application number CN202311670630.5 discloses a spacecraft black box system, including a short message module and its antenna, a processor module, a UHF / VHF communication module and its antenna, a power management module, a distributed body-mounted solar array, and a battery pack; Figure 1 However, the power management module of this patented product still uses existing mature technology without any improvement, and adopts a very simple and direct power-on strategy, which leads to the following limitations:

[0007] First, this patented product uses the satellite-rocket separation signal generated when the satellite and rocket separate to first power on the processor module and put it into working state, which results in all other modules of the spacecraft black box system being in a power-off state during this stage; that is, this patented product does not support the active launch phase to monitor the health status of the parasitic satellite, and can only be powered on and work after the satellite and rocket separate;

[0008] Second, when this patented product detects an abnormality in the key telemetry parameters of the parasitic spacecraft, it will power on the short message module and the UHF communicator at the same time without considering the black box's own energy status; in actual applications, the black box's own energy cannot support the simultaneous power supply of the short message module and the UHF communicator for a long time. Summary of the invention

[0009] The present invention provides a power management system for the entire life cycle of a spacecraft black box. Its purpose is to solve the problem that the prior art is unable to monitor the health status of a parasitic satellite during the active phase of launch into orbit by improving the power-on strategy and energy management strategy of the spacecraft black box system, and also to solve the problem of short working time of the spacecraft black box system under special circumstances.

[0010] To achieve the above object, the technical solution of the present invention is:

[0011] The present invention provides a full life cycle power management system for a spacecraft black box, comprising a bus, a plurality of solar cell arrays, a battery pack, a bus voltage sampling circuit, a plurality of shunt circuits and a filter circuit, wherein the number of the solar cell arrays and the shunt circuits is the same and corresponds one to one;

[0012] Several solar cell arrays are connected in parallel and are respectively connected to the end of the busbar far away from the load; a precision resistor 2 is arranged on the busbar, and the end of the precision resistor 2 close to the load is connected to the power supply of the satellite platform through a relay switch zero; the end of the busbar close to the load is respectively connected to one end of relay switch 2, one end of relay switch 3 and one end of relay switch 4, the other end of relay switch 2 is connected to the computer module through a power conversion circuit, the other end of relay switch 3 is connected to the Beidou short message module through a power conversion circuit, and the other end of relay switch 4 is connected to the UV measurement and control module through a power conversion circuit;

[0013] The battery pack is connected to the busbar via a relay switch 1, the relay switch 1 is connected in parallel with a diode 4, the positive electrode of the diode 4 is connected to the busbar, and the negative electrode of the diode 4 is connected to the positive electrode of the battery pack;

[0014] The on-off of the relay switch zero is controlled by ground commands or the satellite platform; the on-off of the relay switch one is controlled by ground commands or the satellite-rocket separation signal; the relay switch two is closed by default; the on-off of the relay switches three and four is controlled by the computer module;

[0015] One end of the bus voltage sampling circuit is connected to the bus, and the other end is connected to the negative of the primary bus; the bus voltage sampling circuit is connected to the shunt circuit and is used to drive the shunt circuit; one end of the shunt circuit is connected to the positive electrode of the solar cell array, and the other end is connected to the negative of the primary bus;

[0016] One end of the filter circuit is connected to the bus, and the other end is connected to the primary bus negative.

[0017] Furthermore, the power supply of the satellite platform is connected to the end of the precision resistor 2 close to the load through the relay switch zero, the diode one and the precision resistor one in sequence.

[0018] Further, the relay switch four is connected to the busbar through fuse three, the relay switch three is connected to the busbar through fuse four, and the relay switch two is connected to the busbar through fuse five.

[0019] Furthermore, the shunt circuit includes a PWM circuit, a drive circuit and a power switch, one end of the power switch is connected to the positive pole of the solar cell array, and the other end is connected to the negative pole of the primary bus; the PWM circuit is connected to the bus voltage sampling circuit, and the PWM is used to generate pulses according to changes in the bus voltage, and then control the on and off of the power switch through the drive circuit.

[0020] Furthermore, the filter circuit includes capacitor 1, capacitor 2, fuse 1 and fuse 2; capacitor 1 and capacitor 2 are connected in parallel, capacitor 1 is connected to the bus through fuse 1, and the other end of capacitor 1 is connected to the negative primary bus; capacitor 2 is connected to the bus through fuse 2, and the other end of capacitor 2 is connected to the negative primary bus.

[0021] The beneficial effects achieved by the present invention are:

[0022] The present invention supports the power management of the spacecraft black box throughout its life cycle, increases support for active segment operation, and extends the working time and application scope of the spacecraft black box.

[0023] The present invention improves the power-on strategy. For a parasitic spacecraft that is powered on in the active section, the black box system is also powered on in the active section synchronously and powered by the parasitic spacecraft. The working state of the parasitic spacecraft can be monitored in real time in the active section, thereby improving the power-on strategy of the spacecraft black box and making its application scope more universal.

[0024] The present invention proposes an energy management strategy for the black box system, so that it can select different working modes according to its own energy status, thereby extending the working time of the black box. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0026] Figure 1 This is a diagram of the composition of the existing spacecraft black box system.

[0027] Figure 2 It is a schematic diagram of the power supply system disclosed in the present invention.

[0028] Figure 3 It is a schematic diagram of an embodiment of the present invention applied to a spacecraft black box.

[0029] In the figure, K0 is relay switch zero; K1 is relay switch one; K2 is relay switch two; K3 is relay switch three; K4 is relay switch four; R1 is precision resistor one; R2 is precision resistor two; D1 is diode one; D2 is diode two; D3 is diode three; D4 is diode four; C1 is capacitor one; C2 is capacitor two; FU1 is fuse one; FU2 is fuse two; FU3 is fuse three; FU4 is fuse four; FU5 is fuse five. DETAILED DESCRIPTION

[0030] The following will be combined with the drawings in the specification of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention (the following will be combined with the embodiments of the present invention to describe the technical solutions of the present invention). Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0031] It should be noted that if the embodiments of the present invention involve directional indications (such as up, down, left, right, front, back, etc.), the directional indications are only used to explain the relative position relationship, movement status, etc. between the components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0032] In addition, if there are descriptions involving "first", "second", etc. in the embodiments of the present invention, the descriptions of "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "first" and "second" may explicitly or implicitly include at least one of the features. In addition, if the meaning of "and / or" appearing in the full text is to include three parallel schemes, taking "A and / or B" as an example, it includes scheme A, or scheme B, or a scheme that satisfies both A and B. In addition, the technical solutions between the various embodiments can be combined with each other, but it must be based on the ability of ordinary technicians in this field to implement. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such a combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0033] The spacecraft black box device is installed on the surface of the parasitic spacecraft, records the engineering telemetry data in real time during the operation of the spacecraft, and judges whether the spacecraft has a fatal fault through the core telemetry parameters of the key subsystems of the spacecraft. Once a fatal fault is detected, the early warning mechanism is started, the short message module and the UHF / VHF communication module are activated, and emergency short message telemetry information and early warning beacons are sent to the ground to help the ground quickly find the spacecraft and provide telemetry information before and after the fault, so as to help the ground analyze and locate the cause of the fault, and provide an emergency measurement and control channel for rescuing the spacecraft. The spacecraft black box system adopts a body-mounted solar cell array. When the attitude is uncontrolled due to a fatal fault of the parasitic spacecraft, the attitude of the black box system is in a slow rotation state, resulting in limited energy generated by the solar cell array. When the power supply of the parasitic spacecraft is unavailable, the energy state of the spacecraft black box system is difficult to achieve single-turn balance, so the present invention is needed for control.

[0034] In addition, since the spacecraft black box is a relatively independent system compared to the parasitic spacecraft, it has its own independent power supply components, data management, location positioning, emergency communication and other means. Therefore, it is limited by constraints such as volume and weight, and the power supply components of the spacecraft black box are required to be simple and reliable.

[0035] Therefore, the present invention is mainly used to achieve the following objectives:

[0036] (1) It can provide stable power supply for the black box system during the active phase;

[0037] (2) It can provide stable power supply to the black box during the normal operation of the parasitic spacecraft;

[0038] (3) It can provide stable power to the black box when a fatal failure occurs in the parasitic spacecraft, and ensure that the power supply can be long enough.

[0039] like Figure 2 As shown, the present invention provides a power management system for the entire life cycle of a spacecraft black box, comprising a bus, a plurality of solar cell arrays, a battery pack, a bus voltage sampling circuit, a plurality of shunt circuits and a filter circuit, wherein the number of the solar cell arrays is the same as that of the shunt circuits and they correspond one to one;

[0040] Several solar cell arrays are connected in parallel and are respectively connected to the end of the busbar far away from the load; a precision resistor 2 is arranged on the busbar, and the end of the precision resistor 2 close to the load is connected to the power supply of the satellite platform through a relay switch zero; the end of the busbar close to the load is respectively connected to one end of relay switch 2, one end of relay switch 3 and one end of relay switch 4, the other end of relay switch 2 is connected to the computer module through a power conversion circuit, the other end of relay switch 3 is connected to the Beidou short message module through a power conversion circuit, and the other end of relay switch 4 is connected to the UV measurement and control module through a power conversion circuit;

[0041] The battery pack is connected to the busbar via a relay switch 1, the relay switch 1 is connected in parallel with a diode 4, the positive electrode of the diode 4 is connected to the busbar, and the negative electrode of the diode 4 is connected to the positive electrode of the battery pack;

[0042] The on-off of the relay switch zero is controlled by ground commands or the satellite platform; the on-off of the relay switch one is controlled by ground commands or the satellite-rocket separation signal; the relay switch two is closed by default; the on-off of the relay switches three and four is controlled by the computer module;

[0043] One end of the bus voltage sampling circuit is connected to the bus, and the other end is connected to the negative of the primary bus; the bus voltage sampling circuit is connected to the shunt circuit and is used to drive the shunt circuit; one end of the shunt circuit is connected to the positive electrode of the solar cell array, and the other end is connected to the negative of the primary bus;

[0044] One end of the filter circuit is connected to the bus, and the other end is connected to the primary bus negative.

[0045] For example, the number of the solar cell arrays and the shunt circuits are both two, and the two solar cell arrays are connected to the end of the bus bar away from the load through diode 2 and diode 3 respectively.

[0046] Furthermore, the power supply of the satellite platform is connected to the end of the precision resistor 2 close to the load through the relay switch zero, the diode one and the precision resistor one in sequence.

[0047] Among them, the diode 1 is used to control the flow direction of the current so that the current can only flow from the power supply of the satellite platform to the bus; the precision resistor 1 is used to facilitate the remote measurement of the power supply current of the satellite platform; and the precision resistor 2 is used to facilitate the remote measurement of the power supply current of the solar cell array.

[0048] Further, the relay switch 2, relay switch 3 and relay switch 4 are connected in parallel, the relay switch 4 is connected to the busbar through fuse 3, the relay switch 3 is connected to the busbar through fuse 4, and the relay switch 2 is connected to the busbar through fuse 5.

[0049] Furthermore, the power conversion circuit is a TPS54560 series DC-DC.

[0050] The bus voltage sampling circuit is used to collect the bus voltage. The bus voltage sampling circuit adopts the existing technology, and the specific details and distance are not repeated in this article.

[0051] Furthermore, the shunt circuit includes a PWM circuit, a drive circuit and a power switch, one end of the power switch is connected to the positive pole of the solar cell array, and the other end is connected to the negative pole of the primary bus; the PWM circuit and the drive circuit are powered by a separate power supply, the PWM circuit is connected to the bus voltage sampling circuit, and the PWM is used to generate pulses according to changes in the bus voltage, and then control the on and off of the power switch through the drive circuit.

[0052] Among them, the PWM circuit, drive circuit and power switch are all existing technologies, and the specific details and principles are not repeated here.

[0053] The filter circuit includes capacitor 1, capacitor 2, fuse 1 and fuse 2; capacitor 1 and capacitor 2 are connected in parallel, capacitor 1 is connected to the bus through fuse 1, and the other end of capacitor 1 is connected to the negative primary bus; capacitor 2 is connected to the bus through fuse 2, and the other end of capacitor 2 is connected to the negative primary bus.

[0054] In general, the solar cell array is used to receive sunlight to generate energy. The shunt circuit shunts the excess current after the battery pack is fully charged. The filter circuit is mainly used to reduce the bus voltage ripple. The battery pack serves as an energy storage device. The bus voltage sampling circuit is a supporting part of the shunt circuit, which mainly controls PWM to generate pulses based on the collected bus voltage value.

[0055] The composition principle of the spacecraft black box power supply assembly is as follows Figure 2 and Figure 3 As shown, before the satellite is launched, the initial states of relay switch zero and relay switch two are closed, and the initial states of relay switch one, relay switch three and relay switch four are open; relay switch one is controlled by ground command or satellite-rocket separation signal to be closed, and the switch states of relay switch three and relay switch four are controlled by the black box computer module to be closed.

[0056] The power-on state of the spacecraft black box in the active segment is the same as that of the parasitic spacecraft. Specifically, if the parasitic spacecraft is powered on in the active segment, the black box system is also powered on in the active segment. If the active segment of the parasitic spacecraft is not powered on, the black box system is also not powered on; that is, the relay switch zero state is consistent with the power-on state of the parasitic satellite. Assuming that the active segment of the parasitic satellite is not powered on, the relay switch zero remains in the disconnected state. Relay switch two is closed by default, so the processor module starts working when it is powered on. Whether the relay switch zero relay state is powered on is controlled by the parasitic satellite, and the relay switch zero can be closed by the ground injection command. The power supply of the spacecraft black box is preferentially powered by the power supply provided by the parasitic satellite.

[0057] After entering orbit, the satellite-rocket separation signal triggers the relay switch to turn to the closed state. When the parasitic spacecraft power supply is working normally, the parasitic spacecraft power supply is used for power supply. When the parasitic spacecraft power supply fails and the power supply voltage provided is lower than the spacecraft black box solar cell array voltage or battery group voltage, it automatically switches to the black box solar cell array or battery group for power supply.

[0058] During the normal operation of the spacecraft black box, the parasitic spacecraft is powered, and only the computer module is turned on to complete the functions of CAN bus data monitoring, storage, and key telemetry collection of the parasitic spacecraft. When a fatal fault is detected in the parasitic spacecraft, the computer module controls the switch relay switch 3 and relay switch 4 to close, and the UV measurement and control module and Beidou short message module are powered on. The positioning information and a small amount of key telemetry information are transmitted to the ground through Beidou short messages, and the UV measurement and control module sends the telemetry of core subsystems such as power supply, control, measurement and control, and propulsion to the ground station.

[0059] The main core idea of ​​the present invention is to dynamically adjust the black box system operation mode according to the energy status of the spacecraft black box to adjust the energy consumption of the black box system, so as to adapt to the energy status of the spacecraft black box system. The bus voltage is selected as the main evaluation parameter for judging the energy status of the spacecraft black box system. The threshold judgment method is adopted, and nine thresholds from V0 to V8 are selected. These thresholds are generally determined based on empirical values ​​or a certain percentage ratio. The spacecraft black box system consists of a computer module, a Beidou short message module, and a UV measurement and control module. The operation modes of each module are as follows:

[0060] Table 1 Spacecraft black box system power management method

[0061]

[0062]

[0063] According to the energy status of the spacecraft black box, the operating mode is dynamically adjusted to reduce energy consumption, extend the life of the black box as much as possible, and provide more parasitic satellite telemetry, location and other information to the ground.

[0064] The above descriptions are only optional embodiments of the present invention, and are not intended to limit the patent scope of the present invention. All equivalent structural changes made using the contents of the present invention's specification and drawings, or directly / indirectly applied in other related technical fields, are included in the patent protection scope of the present invention.

Claims

1. A power management system for the entire life cycle of a spacecraft black box, characterized by: It includes a bus, a plurality of solar cell arrays, a battery pack, a bus voltage sampling circuit, a plurality of shunt circuits and a filter circuit, wherein the number of the solar cell arrays is the same as that of the shunt circuits and they correspond one to one; Several solar cell arrays are connected in parallel and are respectively connected to the end of the busbar far away from the load; a precision resistor 2 is arranged on the busbar, and the end of the precision resistor 2 close to the load is connected to the power supply of the satellite platform through a relay switch zero; the end of the busbar close to the load is respectively connected to one end of relay switch 2, one end of relay switch 3 and one end of relay switch 4, the other end of relay switch 2 is connected to the computer module through a power conversion circuit, the other end of relay switch 3 is connected to the Beidou short message module through a power conversion circuit, and the other end of relay switch 4 is connected to the UV measurement and control module through a power conversion circuit; The battery pack is connected to the busbar via a relay switch 1, the relay switch 1 is connected in parallel with a diode 4, the positive electrode of the diode 4 is connected to the busbar, and the negative electrode of the diode 4 is connected to the positive electrode of the battery pack; The on-off of the relay switch zero is controlled by ground commands or the satellite platform; the on-off of the relay switch one is controlled by ground commands or the satellite-rocket separation signal; the relay switch two is closed by default; the on-off of the relay switches three and four is controlled by the computer module; One end of the bus voltage sampling circuit is connected to the bus, and the other end is connected to the negative of the primary bus; the bus voltage sampling circuit is connected to the shunt circuit and is used to drive the shunt circuit; one end of the shunt circuit is connected to the positive electrode of the solar cell array, and the other end is connected to the negative of the primary bus; One end of the filter circuit is connected to the bus, and the other end is connected to the primary bus negative.

2. The spacecraft black box full life cycle power management system according to claim 1, characterized in that: The power supply of the satellite platform is connected to the end of the precision resistor 2 close to the load through the relay switch zero, the diode one and the precision resistor one in sequence.

3. The spacecraft black box full life cycle power management system according to claim 1, characterized in that: The relay switch four is connected to the busbar through fuse three, the relay switch three is connected to the busbar through fuse four, and the relay switch two is connected to the busbar through fuse five.

4. The spacecraft black box full life cycle power management system according to claim 1, characterized in that: The shunt circuit includes a PWM circuit, a drive circuit and a power switch, one end of the power switch is connected to the positive pole of the solar cell array, and the other end is connected to the negative pole of the primary bus; the PWM circuit is connected to the bus voltage sampling circuit, and the PWM is used to generate pulses according to changes in the bus voltage, and then control the on and off of the power switch through the drive circuit.

5. The spacecraft black box full life cycle power management system according to claim 1, characterized in that: The filter circuit includes capacitor 1, capacitor 2, fuse 1 and fuse 2; capacitor 1 and capacitor 2 are connected in parallel, capacitor 1 is connected to the bus through fuse 1, and the other end of capacitor 1 is connected to the negative primary bus; capacitor 2 is connected to the bus through fuse 2, and the other end of capacitor 2 is connected to the negative primary bus.

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