RU2778262C1 - Spacecraft power supply system - Google Patents

Abstract

FIELD: space technology.

SUBSTANCE: invention relates to the design of the power supply system (PSS) of near-Earth spacecraft (SC) and can be used in the design and creation of power supply systems for automatic near-Earth low-orbit spacecraft based on solar panels (SP) and batteries (B). The power supply system of the spacecraft contains a battery (B) made of electrically connected battery modules (BM) in series, consisting of parallel connected single batteries or pairs of single batteries; on-board load connected to the output of the B; solar panels (SP). SPs are made of several solar generators electrically connected in parallel and connected via a blocking device (BD) to the corresponding BM. Solar generators consist of series-connected photovoltaic converters. The BDs are placed on a thermostatically controlled panel together with the corresponding BM and are made in the form of electrically connected semiconductor diodes installed on a single base using a parallel or series-parallel circuit, preventing the flow of reverse currents from the BM to the SP. In parallel to each BM, a battery module protection unit from overcharging is connected via a switching device.

EFFECT: reliability of the PSS increases.

13 cl, 9 dwg

Description

Scope of use

The invention relates to the design of a power supply system (SES) of near-Earth spacecraft (SC) and can be used in the design and creation of power supply systems for automatic near-Earth low-orbit spacecraft based on solar panels (SB) and storage batteries (AB).

Prior Art

Known SB for SES SC [patent US3952324], consisting of panels with placed rows of photoelectric converters (PVC), in which two or three adjacent rows of a plurality of series-connected PV cells are connected in parallel. Then the series-parallel groups are connected through the appropriate busbars to the energy use systems (AB and on-board load). At the same time, blocking diodes are installed on the panels, connected in series between each group of solar cells and the bus. Blocking diodes make it possible to prevent AB discharge in the absence of SB illumination, as well as panel failure in the event of short circuits in the solar cell.

The disadvantage of placing blocking diodes on the SB panels for near-Earth low-orbit spacecraft is a decrease in their reliability and service life due to thermomechanical loads due to constant exposure to cyclic temperature changes with a period of 90-100 minutes in the range from minus 110 to plus 90°C.

Known AB KA [utility model 178488], containing at least two modules electrically connected in series, each of which includes several cylindrical batteries connected electrically in parallel with the help of (+) and (-) current collectors located near their ends, at the same time, the connection of batteries to one of which is carried out by fuses.

Known AB KA [RF patent 2667905] consisting of several pairs of series-connected batteries with a one-time built-in current protection element that breaks its internal electrical circuit when the pressure inside the battery increases, electrically connected in a parallel-series circuit. As an additional protection against fires due to current overloads in the circuit of a pair of series-connected batteries, a fusible insert can be used, which is located between them. The proposed technical solution improves the safety of operation and the stability of the battery to degradation of battery characteristics.

Known SES and its control method [RF patent 2699764], containing solar panels connected to the load through a stabilized voltage converter, batteries connected to the solar panels through chargers, and to the load through discharge devices. Stabilized voltage converter perform two types: parallel and serial types. According to the method, the load voltage from the primary source of limited power (solar battery) is stabilized at the beginning of the spacecraft operation life using a parallel voltage stabilizer, and after the resource degradation of the solar battery characteristics and if its power is insufficient to power the load, they switch to load voltage stabilization with using a series voltage regulator. In this case, switching the stabilization mode is provided automatically or on command from the Earth.

Known SES and the method of its control [RF patent 2636384], containing SB, n AB and n chargers and discharge devices, which are controlled depending on the illumination, the degree of charge of all ABs, the input and output voltage of the SES. The method provides for a ban on the operation of the corresponding charger when the maximum level of charge of this AB is reached and the removal of this ban when the level of charge of this AB decreases. Similarly, a ban on the operation of the corresponding discharge device is introduced when the established minimum level of charge of a given AB is reached and it is removed when the level of charge of this AB increases. It also provides for the formation of a control signal to the monitoring and control unit (BCU) of the spacecraft to turn off part of the onboard equipment in the event of an emergency discharge of several batteries to a minimum level of charge. In this case, communication with the BCU is carried out via a duplicated trunk serial interface using a terminal device with a microcontroller.

Known SES KA [RF patent 2297706] in which a sectioned SB is used, an individual serial voltage regulator with an extreme power regulator is installed at the input of each section. At the input of the stabilizer (at the output of the SB section), a voltage is set equal to the optimal voltage of the current-voltage characteristic of the section, and at the output - a stabilized voltage equal to the voltage at the load. All stabilizer outputs are connected to the load busbar.

Closest to the claimed invention is a combined DC source (publication number WO 2008/127137 A1), containing a single secondary electrochemical current sources (for example, a module of parallel-connected lithium-ion batteries (LIA)) electrically connected in series to each of which, with the help of a switching device, a primary source of direct current (for example, a solar cell battery) is connected through an electronic unit for matching operating voltages. The monitoring and control unit monitors the voltage of each LIB module and turns off the corresponding PV battery when the LIB module is charged to a predetermined voltage level, or disconnects the combined DC source from the load when the LIB module is discharged below a predetermined voltage level.

A disadvantage of the known SES of near-Earth space vehicles is the use of complex electronic equipment to maintain the energy balance and ensure their long-term operation. This leads to an increase in the likelihood of failures or single failures in it, including the complete loss of operability of the SES. One of the main reasons for the occurrence of failures or malfunctions in the operation of the SES is the use of microcontrollers operating according to complex algorithms that are sensitive to various radiation effects in the control equipment for its operation.

Disclosure of invention

The objective of the present invention is to improve the reliability, fault tolerance and radiation resistance of the spacecraft SES by excluding from its composition microprocessor power electronic units and other equipment using semiconductor elements with a high degree of integration.

The specified technical result is achieved by the fact that:

1. In the power supply system of the spacecraft, containing a battery made of electrically connected in series n battery modules, consisting of single batteries connected in parallel or pairs of single batteries; n solar batteries, each of which is made of several solar generators electrically connected in parallel, consisting of series-connected photovoltaic converters, and connected through a unit for matching operating voltages to one of the battery modules; switching devices; switching devices; a monitoring and control unit that controls the voltage of each battery module and generates control signals for switching devices; the onboard load connected to the battery output, the operating voltage matching blocks are made in the form of blocking devices installed on a temperature-controlled panel together with the corresponding battery module and preventing the flow of reverse currents from the battery modules to solar generators; In parallel with each battery module, a battery module protection unit against overcharging is connected through a switching device.

2. In the power supply system of the spacecraft, the blocking device is made in the form of at least two semiconductor diodes electrically connected in parallel installed on a single base to ensure operability in the event of a break in one of them.

3. In the power supply system of the spacecraft, the blocking device is made in the form of at least four semiconductor diodes electrically connected in a series-parallel circuit, installed on a single base, ensuring that the device remains operational in any type of failure of one of the diodes.

4. In the power supply system of the spacecraft, Schottky diodes are used as semiconductor diodes, which have a minimum voltage drop compared to other types of diodes.

5. In the power supply system of the spacecraft, the switching device that connects the overcharge protection unit of the battery module is made in the form of an electronic key switching number that has a practically unlimited resource.

6. In the power supply system of the spacecraft, the switching device that connects the overcharge protection unit of the battery module is made in the form of an electromechanical contactor that is simpler in circuit implementation.

7. In the power supply system of the spacecraft, the onboard control complex of the spacecraft is used as a control and management unit.

8. In the power supply system of the spacecraft, an analog comparator with hysteresis (Schmitt trigger) is used as a monitoring and control unit, tuned to the value of the maximum charging voltage of the battery module.

9. In the power supply system of the spacecraft, the battery overcharge protection device is made in the form of at least two ballast resistors electrically connected in parallel, placed on the outer surface of the spacecraft and providing a discharge of excess generated power by radiation.

10. In the power supply system of the spacecraft, the ballast resistor is made in the form of a metal resistive element placed on the heat-insulating panel.

11. In the power supply system of the spacecraft, the battery is made of lithium-ion batteries of size 18650, which makes it possible to create batteries with increased reliability using various electrical configurations.

12. In the power supply system of the spacecraft, the battery is made in a parallel-series (PS) configuration.

13. In the power supply system of the spacecraft, the battery is made in a series-parallel-series (SPS) configuration.

An example of the implementation of SES KA.

The block diagram of the power supply system of the spacecraft is shown in Fig. one.

It includes a battery AB 1 placed on a temperature-controlled panel inside the non-hermetic housing of the spacecraft, to the terminals of which an onboard load 2 is connected. -mi single batteries 4 size 18650. The current-voltage characteristics of AM 3 at a temperature of 20°C are shown in Fig. 2. At the end of the SAS KA (5 years), the energy intensity of AM 3 is reduced by 50%.

Three solar panels SB 5 (5a, 5b, 5c) are connected to each AM 3 (Za, 36, Sv), each of which is made of 100 solar generators SG 6 electrically connected in parallel, consisting of two photovoltaic converters FEP 7 connected in series. The characteristics of SB 5 under illumination conditions of 1367 W/m2 at the beginning of the spacecraft's active lifetime (SAS) and at the end of a five-year SAS are shown in Fig. 3 and FIG. 4, respectively.

Each SB 5 is connected to the corresponding AM 3 using an electrical bus 8 with a resistance of 1 mOhm through a blocking device 9 placed together with AM 3 on a temperature-controlled panel, made in the form of 10 electrically connected parallel semiconductor diodes 10 with current-voltage characteristics installed on a single base, shown in FIG. 5.

An overcharge protection device 12 is connected to each AM 3 through an electronic key 11, made in the form of 2 ballast resistors 13 electrically connected in parallel. radiation. Connection/disconnection to AM 3 ballast resistors 13 is controlled by an electrical signal supplied to the electronic key 11 from the on-board control complex 14, which is not part of the SES.

A variant of the SES KA version is possible, in which an analog comparator with hysteresis (Schmitt trigger) is used to control the electronic key 10, tuned to the value of the maximum charging voltage AM 3.

Thus, in the spacecraft power supply system shown in Fig. 1, the onboard load 2 is powered in all segments of the SC orbit by AB 1, and the SB 5 in the SES in the illuminated segment of the orbit act as recharging devices for the corresponding AM 3. SB circuit 5. In emergency modes of operation of the spacecraft associated with long-term power consumption of onboard equipment at a minimum level, AM 3 is protected from overcharging in the illuminated part of the orbit by connecting ballast resistors 13 to them.

As an example demonstrating the operability of the proposed technical solution for constructing a solar power plant, for a standard cyclogram of the energy-loaded operation mode of the onboard equipment of the spacecraft, shown in Fig. 6., graphs obtained by calculation of the change in voltage and depth of discharge AM 3 in AB 1 during 24 hours of the flight of the spacecraft at the beginning and end of the ACS of the spacecraft are presented: Fig. 7 and FIG. 8 respectively. To demonstrate the automatic voltage balancing of AM 3 in AB 1, in the calculations, the initial depth of discharge of AM 3a, 3b, 3c is set to 0%, 10% and 20%, respectively.

The analysis of the graphs shows that with the above electrical characteristics of the components of the SES of the spacecraft, which do not contain power electronic units operating under the control of microcontrollers and other equipment using semiconductor elements with a high degree of integration, and the power consumption of the onboard equipment according to Fig. 6, at the beginning of the SAS SC, the proposed scheme for constructing the SES SC ensures stable maintenance of the daily energy balance with the initially fully charged battery 4 - at a maximum charging voltage of about 12.1 V, which corresponds to the zero depth of its discharge.

At the end of the SAS SC due to the degradation of the electrical characteristics of SB 5 and AB 1, stable maintenance of the daily energy balance is achieved at lower voltages of AB 4 - about 11.4 V at the end of the charge, which corresponds to 20% of the discharge depth.

With a long-term decrease in the power consumption of onboard equipment to a minimum (standby flight mode with a power consumption of 300 W) in order to avoid dangerous overcharging of AM 3 (above 4.2 V, which corresponds to a depth of discharge minus 20%) at the command of the onboard control complex 14 using the electronic key 11 an overcharge protection device 12 is connected in parallel to AM 3 and the maximum charging voltage of battery 4 is limited to 4.1 V (Fig. 9).

All the features listed in the description directly affect the achievement of the specified technical result through the proposed scheme for constructing the power supply system. Connecting the on-board load directly to the battery, and the solar battery to it through an electrical bus and a blocking device installed on a temperature-controlled panel together with the battery, as well as parallel connection to the battery through a key element of the battery overcharge protection device, is an essential feature of alternative elements aimed at achieving the technical result of the utility model: increasing the reliability, radiation resistance and fault tolerance of the spacecraft power supply system.

Claims (13)

1. The power supply system of the spacecraft, containing a battery made of electrically connected in series n battery modules consisting of single batteries connected in parallel or pairs of single batteries; n solar batteries, each of which is made of several solar generators electrically connected in parallel, consisting of series-connected photovoltaic converters, and connected through a unit for matching operating voltages to one of the battery modules; switching devices; a monitoring and control unit that controls the voltage of each battery module and generates control signals for switching devices; an on-board load connected to the battery output, characterized in that the operating voltage matching blocks are made in the form of blocking devices installed on a temperature-controlled panel together with the corresponding battery module and preventing the flow of reverse currents from the battery modules to the solar panels; In parallel with each battery module, a battery module protection unit against overcharging is connected through a switching device. 2. The spacecraft power supply system according to claim 1, characterized in that the blocking device is made in the form of at least two semiconductor diodes electrically connected in parallel installed on a single base. 3. The spacecraft power supply system according to claim 1, characterized in that the blocking device is made in the form of at least four semiconductor diodes electrically connected in series-parallel circuit installed on a single base. 4. The power supply system of the spacecraft according to any one of paragraphs. 2 and 3, characterized in that Schottky diodes are used as semiconductor diodes. 5. The power supply system of the spacecraft according to any one of paragraphs. 1-4, characterized in that the switching device connecting the battery module protection unit against overcharging is made in the form of an electronic key. 6. The power supply system of the spacecraft according to any one of paragraphs. 1-4, characterized in that the switching device that connects the overcharge protection unit of the battery module is made in the form of an electromechanical contactor. 7. The power supply system of the spacecraft according to any one of paragraphs. 1-6, characterized in that the onboard control complex of the spacecraft is used as a monitoring and control unit. 8. The power supply system of the spacecraft according to any one of paragraphs. 1-6, characterized in that an analog comparator with hysteresis is used as a monitoring and control unit, namely a Schmitt trigger, tuned to the value of the maximum charging voltage of the battery module. 9. The power supply system of the spacecraft according to any one of paragraphs. 1-8, characterized in that the battery overcharge protection device is made in the form of at least two ballast resistors electrically connected in parallel, placed on the outer surface of the spacecraft and providing a discharge of excess generated power by radiation. 10. The spacecraft power supply system according to claim 9, characterized in that the ballast resistor is made in the form of a metal resistive element placed on a heat-insulating panel. 11. The power supply system of the spacecraft according to any one of paragraphs. 1-10, characterized in that the battery is made of lithium-ion batteries of size 18650. 12. The power supply system of the spacecraft according to claim 11, characterized in that the battery is made in a parallel-series configuration. 13. The power supply system of the spacecraft according to claim 11, characterized in that the battery is made in a series-parallel-series configuration.

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