CN106933094B - Dual-redundancy airborne flight control computer - Google Patents
Dual-redundancy airborne flight control computer Download PDFInfo
- Publication number
- CN106933094B CN106933094B CN201710117227.8A CN201710117227A CN106933094B CN 106933094 B CN106933094 B CN 106933094B CN 201710117227 A CN201710117227 A CN 201710117227A CN 106933094 B CN106933094 B CN 106933094B
- Authority
- CN
- China
- Prior art keywords
- board card
- logic control
- control board
- main
- standby
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B9/00—Safety arrangements
- G05B9/02—Safety arrangements electric
- G05B9/03—Safety arrangements electric with multiple-channel loop, i.e. redundant control systems
Landscapes
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- Automation & Control Theory (AREA)
- Safety Devices In Control Systems (AREA)
- Hardware Redundancy (AREA)
Abstract
The invention discloses a dual-redundancy airborne flight control computer, which comprises a main logic control board card, a standby logic control board card and a scheduling board card; the dispatching board card is provided with a main slot and a standby slot, the main logic control board card is installed in the main slot, and the standby logic control board card is installed in the standby slot. The main logic control board card and the standby logic control board card are packaged by adopting different software architectures and hardware, and are respectively supplied with power by two independent power modules. The power supply module of the main logic control board card is uniformly controlled by the standby logic control board card and the dispatching board card, and the power supply module of the standby logic control board card is uniformly controlled by the main logic control board card and the dispatching board card. The invention provides redundancy for core complex components in the flight control computer based on heterogeneous redundancy technology, reduces the failure rate of the flight control computer, isolates the failure by using an independent power module and control logic, ensures the availability of the system under the condition of single-point failure and improves the reliability of the unmanned aerial vehicle.
Description
Technical Field
The invention belongs to the technical field of avionics, and particularly relates to a dual-redundancy airborne flight control computer, which is used for improving the reliability of the airborne flight control computer.
Background
With the continuous development of unmanned aerial vehicles, the complexity of airborne avionics systems of the unmanned aerial vehicles is continuously improved, and the problem of the increased failure probability is brought. The flight control computer is used as a key component of the unmanned aerial vehicle, and the stability of the flight control computer directly influences the reliability of the flight control system, so that the flight safety and the task execution capacity of the unmanned aerial vehicle are influenced. Therefore, designing a highly reliable flight control computer is one of the key elements for improving the overall performance of the unmanned aerial vehicle.
Generally speaking, in addition to selecting highly reliable components, the most effective method for improving the reliability of the flight control computer is to use redundancy technology, that is, two or more identical components or systems are used to correctly and coordinately complete a unified task. For the flight control computer, the complexity of the system is increased while the redundancy unit is added, and the reliability of the system is reduced if the redundancy is not properly added. From the angle of redundancy configuration, most of the existing flight control computers adopt isomorphic redundancy, namely, all redundancy units of the flight control computers adopt the same hardware platform and control software, so that the problem is that the fault occurring in a certain element is also likely to occur in the isomorphic redundancy unit; and the heterogeneous redundancy utilizes different channels to select different hardware platforms and control software, so that common-mode faults are avoided. Therefore, designing a proper margin and its configuration is very necessary to improve the reliability of the unmanned aerial vehicle.
Disclosure of Invention
In order to solve the problems, the invention discloses a dual-redundancy airborne flight control computer which is mainly used for solving the following technical problems: (1) the reliability of the airborne flight control computer is improved, and the condition that the whole system is paralyzed due to the fault of the flight control computer is avoided; (2) by utilizing heterogeneous redundancy technology, the common-mode fault is avoided, and a fault switching method of the dual-redundancy airborne flight control computer is provided; (3) the use range of the airborne flight control computer is expanded by utilizing the scheduling board card, so that the scheduling board card can be applied to different types of unmanned aerial vehicles such as fixed wings, helicopters, aerostats and the like.
Specifically, the invention provides a dual-redundancy airborne flight control computer, which comprises a main logic control board card, a standby logic control board card and a scheduling board card; the dispatching board card is provided with a main slot and a standby slot, the main logic control board card is installed in the main slot, and the standby logic control board card is installed in the standby slot.
Preferably, as for the dual-redundancy airborne flight control computer, the main and standby logic control boards receive the position, attitude and other related information of the airborne flight control computer, perform calculation according to a preset flight path or flight mission, and finally output control instructions to each execution mechanism of the airborne flight control computer through the bus.
Preferably, the dual-redundancy airborne flight control computer is characterized in that the scheduling board card is a fixed-wing scheduling board card, a helicopter scheduling board card or an aerostat scheduling board card.
Preferably, as for the dual-redundancy airborne flight control computer, the main logic control board card and the standby logic control board card are packaged by adopting different software architectures and hardware, and are respectively powered by two independent power modules.
Preferably, as for the dual-redundancy airborne flight control computer, the power module of the main logic control board is controlled by the standby logic control board and the scheduling board, and the power module of the standby logic control board is controlled by the main logic control board and the scheduling board.
Preferably, as for the dual-redundancy airborne flight control computer, the main logic control board card, the standby logic control board card and the scheduling board card are independently communicated with the CAN bus through the CAN transceiver and respond to the instruction from the telemetry system.
Preferably, the CAN transceiver is powered by an independent DC-DC, as described above for a dual-redundancy airborne flight control computer.
According to another aspect of the present invention, the present invention further provides a fault handling method using the dual-redundancy airborne flight control computer as described above, including the following steps:
when the main logic control board card has a fault, firstly, the ground control system sends an instruction for switching to a silent mode to the main logic control board card through the CAN bus, and if the fault degree is not high, the main logic control board card responds to the instruction to enter the silent mode; if the fault degree is too high, the main logic control board card cannot respond to the command, a main board power-off command is sent to the dispatching board card and the standby logic control board card through the CAN bus, and the dispatching board card and the standby logic control board card respectively send high-level and no-output disconnection signals in response to the command, so that a power module of the main logic control board card is cut off; then, the ground control system sends an instruction through the CAN bus to enable the standby logic control board card to enter a working mode;
when the dispatching board card has a fault, if the main logic control board card has a fault, the ground control system sends an instruction for switching to the silent mode to the main logic control board card through the CAN bus, and if the fault degree is not high, the main logic control board card responds to the instruction and enters the silent mode; if the fault degree is too high, the main logic control board card cannot respond to the command, the main logic control board card is powered off and the fault is isolated by sending a main board power-off command to the dispatching board card and the standby logic control board card.
According to another aspect of the present invention, the present invention further provides a power-on method for the dual-redundancy airborne flight control computer, including the following steps:
(1) the main logic control board card is inserted into the main slot, and the standby logic control board card is inserted into the standby slot;
(2) after the computer is powered on and the scheduling board card is initialized, the main logic control board card and the standby logic control board card are powered on and initialized; the scheduling board card identifies whether the board card is inserted into the main slot or not, and if only the main slot is identified to have the board card inserted, the single redundancy mode is entered; if the fact that only the standby slot is inserted with the board card is recognized, alarm information is sent out through the bus, and the user is prompted that the airborne flight control computer is abnormal; if the fact that the board cards are inserted into the two slots is recognized, first communication is respectively initiated to the board cards of the main and standby slots, the configuration information of the board cards on the main and standby slots is read, and software in the board cards is dispatched to be switched to a main and standby working mode;
(3) the main logic control board card automatically identifies the current slot as a main slot through configuration information in first communication initiated by the scheduling board card, and the software in the board is switched to a main controller mode; the standby logic control board card automatically identifies the current slot as a standby slot, and the software in the board is switched to a standby controller mode;
(4) the main logic control board card and the standby logic control board card read the configuration information of the scheduling board card, the configuration information is identified as the scheduling board card of the unmanned aerial vehicle of a certain type, and the software is switched to a mode under the type;
(5) the computer self-check is carried out, the computer self-check is communicated with a ground station, and the current computer state including a hardware version, a software version, a main and standby working modes, and the configuration information, voltage, temperature and air pressure information of the identified scheduling board card are downloaded;
(6) and computers wait for ground commands.
The invention has the advantages that: compared with the prior art, the method is based on the heterogeneous redundancy technology, redundancy is provided for core complex components in the flight control computer, the fault rate of the flight control computer is reduced, the fault is isolated by using the independent power module and the control logic, the availability of the system under the condition of single-point fault is ensured, and the reliability of the unmanned aerial vehicle is improved.
Drawings
Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments. The drawings are only for purposes of illustrating the preferred embodiments and are not to be construed as limiting the invention. Also, like reference numerals are used to refer to like parts throughout the drawings. In the drawings:
figure 1 shows a hardware block diagram of an on-board flight control computer according to an embodiment of the invention.
Figure 2 illustrates a logical schematic of an on-board flight control computer according to an embodiment of the present invention.
Detailed Description
Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be embodied in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.
Firstly, a hardware structure:
as shown in fig. 1, the airborne flight control computer of the present invention mainly comprises three boards: a main logic control board card (hereinafter also referred to as logic control board card 1), a standby logic control board card (hereinafter also referred to as logic control board card 2) and a scheduling board card. The dispatching board card is provided with a main slot and a standby slot, the main logic control board card is installed in the main slot, and the standby logic control board card is installed in the standby slot.
The main logic control board card and the standby logic control board card are responsible for receiving the position, the attitude and other related information of the system, resolving according to a preset flight path or a flight mission, and finally outputting control instructions to each actuating mechanism of the system through the bus.
In order to realize the heterogeneous dual-redundancy design, the main logic control board and the standby logic control board are respectively packaged by completely different hardware and software architectures on the premise that the external interface definitions are kept consistent.
The dispatching board card is connected with the main logic control board card and the standby logic control board card, so that the communication function with the control board card is realized, and the external interface of the flight control computer is expanded. The pin definitions of the main slot and the standby slot are completely the same. When used as a single redundancy computer, the main slot is provided with a main logic control board card, and the standby slot can be used for expanding other self-research modules, such as an inertial navigation module. The scheduling board card can be used as a fixed wing scheduling board card, a helicopter scheduling board card or an aerostat scheduling board card.
The main and standby positions of the logic control board cards 1 and 2 are determined by the positions of the slots into which the logic control board cards are inserted, namely, a 1 card (or a 2 card) is inserted into the main slot and automatically becomes a main card, and a standby logic control board card is inserted into the other slot and becomes a standby logic control board card. If 1 and 2 are two identical board cards, the system forms isomorphic dual redundancy, otherwise, isomeric dual redundancy is formed; if dual redundancy is not needed, the system can work normally as long as a single logic control board card is inserted into the main slot.
The two logic control board cards 1 and 2 adopt different software architectures and hardware packages and are powered by two independent power supply modules. The main and standby positions of the two logic control board cards are determined by the slots into which the two logic control board cards are inserted, and the two logic control board cards are inserted into the main slots to work as the main logic control board cards, and vice versa. The two logic control board cards are powered by two independent power supply modules, the on-off of each power supply module is controlled by other logic control board cards and the dispatching board card together, for example, the power supply module in charge of the main card is uniformly controlled by the standby board card and the dispatching board card. The three main parts of the flight control computer CAN independently communicate with the CAN bus through the communication interface and respond to the instruction from the telemetering system.
The board cards 1 and 2 are respectively supplied with power by independent power supply modules, the on-off of each power supply module is jointly determined and controlled by other logic control board cards and the dispatching board card, namely, the power supply module in charge of the main card can be uniformly controlled by the standby card and the dispatching board card, namely, the power supply module in charge of the standby card can be uniformly controlled by the main card and the dispatching board card. This is designed to achieve fault isolation.
By designing different scheduling board cards, the system can be applied to different types of unmanned aerial vehicles such as fixed wings, helicopters and aerostatics.
II, hardware logic:
aiming at the hardware framework, the invention provides a scheduling control method for main-standby switching and fault handling, which can process the faults of a scheduling board card (including two conditions of no output and logic control board card (including three conditions of no output, output and input abnormity).
As shown in fig. 2: the main and standby logic control board cards adopt respective independent power supply paths and power supply modules so as to avoid system faults caused by power supply abnormity; the power supply of the scheduling board card is also from the two power supplies and is independent of the main and standby circuits.
Because the function of the dispatching board card is single, STM32 series high-reliability micro control unit MCU is selected as a core to realize most of logic functions of the dispatching board card.
The CAN bus is used as a communication bus, and a CAN bus communication interface with complete electric appliance isolation is designed on the scheduling board card, so that the CAN bus communication interface CAN be applied to strong-current and strong-interference peripherals.
A CAN transceiver (transceiver) chip of a CAN bus communication interface adopts independent DC-DC power supply to isolate the influence of current change on a main slot (main logic control board) and a standby slot (main logic control board) power supply circuit. If DC-DC is not adopted for isolation, the change of the current on the transceiver chip can directly influence the power supply of the main slot and the standby slot.
The power module P1 in charge of the main circuit is switched on and off by a Control signal (Control)2 signal of the logic Control board card 2 and a scheduling signal (Schedule)1 signal of a scheduling chip on the scheduling board card, only when the Control2 and the Schedule1 signals are simultaneously controlled to be switched off, the P1 is switched off, and the other states P1 keep a path; the power module P2 of the responsible circuit is turned on or off by the Control1 signal of the logic Control board 1 and the Schedule2 signal of the scheduling chip.
The OutputDisable1 and OutputDisable2 signals are silent mode control signals of the master/slave logic control board card, and through the two signals, the master/slave CAN close the output of the corresponding CAN bus transceiver chip.
Thirdly, the power-on working process of the flight control computer of the invention comprises the following steps:
1. the logic control board card 1 is inserted into the main slot, and the logic control board card 2 is inserted into the standby slot.
2. After the computer is powered on and the scheduling board card is initialized, both default output Schedule1 (high level) and Schedule2 (high level) commands are power switches closed, and both Control1 (no output) and Control2 (no output) commands from the opposite board card are power switches closed, so that the logic Control board cards 1 and 2 are powered on and initialized. The MCU on the scheduling board card identifies whether the board card is inserted into the main slot or not, and if the MCU identifies that only the main slot has the board card inserted into the main slot, the single redundancy mode is entered; if the fact that only the standby slot is inserted with the board card is recognized, alarm information is sent out through the bus, and the user is prompted that the airborne flight control computer is abnormal; if the fact that the board cards are inserted into the two slots is recognized, first communication is respectively initiated to the board cards of the main and standby slots, the configuration information of the board cards of the main and standby slots is read, and software in the board cards is dispatched to be switched to a main and standby working mode.
3. The logic control board card 1 automatically identifies the current slot as a main slot through configuration information in first communication initiated by a scheduling board card, and the software in the board is switched to a main controller mode; the logic control board card 2 automatically recognizes the current slot as the standby slot, and the software in the board is switched to the controller mode.
4. The logic control board card 1 and the logic control board card 2 read the configuration information of the scheduling board card, the configuration information is identified as the scheduling board card of the unmanned aerial vehicle of a certain type, and the software is switched to a mode under the type.
5. The computer self-checks, communicates with the ground station, and downloads the current computer state, including the hardware version, the software version, the main and standby working modes, the configuration information of the identified dispatch board card, the voltage, the temperature, the air pressure and other information.
6. The computer waits for a ground command.
Fourthly, the invention realizes the processes of switching between the computer master and the computer slave and isolating the fault:
for unmanned aerial vehicle applications, where control real-time requirements are very high, a hot-standby mechanism must be used to switch quickly. Therefore, after the initial power-on, the main and standby logic control boards are both in a power-on state, and are identified as the working mode of the main logic control board and the silent mode of the standby logic control board (except that no communication application is initiated to the bus, no difference is found from the working mode), and at this time, the system is in a dual-redundancy hot backup state. The dispatching circuit is powered by double power supplies and is directly hung on an external bus. When the ground station finds that the airborne flight control computer is abnormal, the ground station can directly send an instruction to the dispatching board card to close the main logic control board card (or the standby logic control board card) and open the standby logic control board card (or the main logic control board card).
Aiming at the problem that the main logic control board card fails in the working process: firstly, the ground control sends an instruction for switching to a silent mode to the main logic control board card through a bus, and if the fault degree is not high, the main logic control board card can respond to the instruction and enter the silent mode; if the fault degree is too high, the main logic Control board card cannot respond to the instruction at all, a main board power-off instruction is sent to the dispatching board card and the standby logic Control board card through the bus, the dispatching board card and the standby logic Control board card respond to the instruction and send out Control2 (high level) and Schedule1 (no output) disconnection signals respectively, at the moment, a power supply system of the main logic Control board card is cut off, and the fault of the main logic Control board card is isolated. And then, the ground control system sends an instruction through the bus, so that the standby logic control board card enters a working mode and takes over relevant peripheral equipment and the bus. At this time, because the main logic Control board card is powered off, the Control1 signal is not output at this time, which means that the power supply condition of the standby logic Control board card is not affected no matter what signal is sent by the scheduling board card or a fault occurs, and the standby logic Control board card is in a power-on state.
Aiming at the problem that the scheduling board card fails in the working process: at this time, if the ground does not execute the main/standby switching instruction, the fault of the scheduling board card cannot be found, but the signals of Schedule1 and Schedule2 are in a non-output state due to the fault of the scheduling board card, which represents an open instruction, but because the main/standby logic control board cards mutually send a power switch closing instruction to each other, the power supply of the main/standby logic control board cards is not cut off, and the main/standby logic control board cards can still work normally. At this time, if the main logic control board card has a fault, the ground control system detects that the flight control computer is abnormal, firstly, an instruction for switching to the silent mode is sent to the main logic control board card through the bus, and if the fault degree is not high, the main logic control board card can respond to the instruction and enter the silent mode; if the fault degree is too high, the main board cannot respond to the command at all, a main board power-off command is sent to the dispatching board card and the standby logic Control board card, no signal is output due to the fault of the dispatching system, namely, the signal represents a disconnection command, the standby logic Control board card responds to the main logic Control board card power-off command, Control2 (high level) is sent, the main logic Control board card is powered off, and the fault is isolated.
In addition, if the energy system (i.e. each power module) has its own power switching function, the switching controller on the interface board may also be shielded, and the active/standby switching is performed by the energy system directly receiving a ground command.
The mode of performing redundancy switching by powering on and powering off is the final solution, which is used for thoroughly isolating a faulty computer and solving the problem that the condition of peripheral equipment and a bus is influenced when the computer is not powered off. No matter the fault board card enters the silent mode or the power supply is directly cut off, careful consideration is needed, the fault board card is required to be sequentially carried out according to the priority, and the irreversible situation is avoided because the ground station does not know the working state of the standby redundancy at all.
The above description is only for the preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the appended claims.
Claims (8)
1. A fault handling method using a dual-redundancy airborne flight control computer comprises a main logic control board card, a standby logic control board card and a scheduling board card; the dispatching board card is provided with a main slot and a standby slot, the main logic control board card is installed in the main slot, and the standby logic control board card is installed in the standby slot, and the dispatching board card comprises the following steps:
when the main logic control board card has a fault, firstly, the ground control system sends an instruction for switching to a silent mode to the main logic control board card through a CAN bus, and if the fault degree is not high, the main logic control board card responds to the instruction to enter the silent mode; if the fault degree is too high, the main logic control board card cannot respond to the command, a main board power-off command is sent to the dispatching board card and the standby logic control board card through the CAN bus, and the dispatching board card and the standby logic control board card respectively send high-level and no-output disconnection signals in response to the command, so that a power module of the main logic control board card is cut off; then, the ground control system sends an instruction through the CAN bus to enable the standby logic control board card to enter a working mode;
when the dispatching board card has a fault, if the main logic control board card has a fault, the ground control system sends an instruction for switching to the silent mode to the main logic control board card through the CAN bus, and if the fault degree is not high, the main logic control board card responds to the instruction and enters the silent mode; if the fault degree is too high, the main logic control board card cannot respond to the command, the main logic control board card is powered off and the fault is isolated by sending a main board power-off command to the dispatching board card and the standby logic control board card.
2. A power-on method of a dual-redundancy airborne flight control computer is disclosed, wherein the dual-redundancy airborne flight control computer comprises a main logic control board card, a standby logic control board card and a scheduling board card; the dispatching board card is provided with a main slot and a standby slot, the main logic control board card is installed in the main slot, and the standby logic control board card is installed in the standby slot, and the dispatching board card comprises the following steps:
(1) the main logic control board card is inserted into the main slot, and the standby logic control board card is inserted into the standby slot;
(2) after the computer is powered on and the scheduling board card is initialized, the main logic control board card and the standby logic control board card are powered on and initialized; the scheduling board card identifies whether the board card is inserted into the main slot or not, and if only the main slot is identified to have the board card inserted, the single redundancy mode is entered; if the fact that only the standby slot is inserted with the board card is recognized, alarm information is sent out through the bus, and the user is prompted that the airborne flight control computer is abnormal; if the fact that the board cards are inserted into the two slots is recognized, first communication is respectively initiated to the board cards of the main and standby slots, the configuration information of the board cards on the main and standby slots is read, and software in the board cards is dispatched to be switched to a main and standby working mode;
(3) the main logic control board card automatically identifies the current slot as a main slot through configuration information in first communication initiated by the scheduling board card, and the software in the board is switched to a main controller mode; the standby logic control board card automatically identifies the current slot as a standby slot, and the software in the board is switched to a standby controller mode;
(4) the main logic control board card and the standby logic control board card read the configuration information of the scheduling board card, the configuration information is identified as the scheduling board card of the unmanned aerial vehicle of a certain type, and the software is switched to a mode under the type;
(5) the computer self-check is carried out, the computer self-check is communicated with a ground station, and the current computer state including a hardware version, a software version, a main and standby working modes, and the configuration information, voltage, temperature and air pressure information of the identified scheduling board card are downloaded;
(6) and computers wait for ground commands.
3. The method of claim 1 or 2, wherein:
and the main logic control board card and the standby logic control board card receive the position, the attitude and other related information of the airborne flight control computer, resolve according to a preset flight path or a flight mission, and finally output control instructions to each executing mechanism of the airborne flight control computer through a bus.
4. The method of claim 1 or 2, wherein:
the scheduling board card is a fixed wing scheduling board card, a helicopter scheduling board card or an aerostat scheduling board card.
5. The method of claim 2, wherein:
the main logic control board card and the standby logic control board card are packaged by adopting different software architectures and hardware, and are respectively supplied with power by two independent power modules.
6. The method of claim 5, wherein:
the power module of the main logic control board card is controlled by the standby logic control board card and the dispatching board card together, and the power module of the standby logic control board card is controlled by the main logic control board card and the dispatching board card together.
7. The method of claim 1 or 5, wherein:
the main logic control board card, the standby logic control board card and the scheduling board card are independently communicated with the CAN bus through the CAN transceiver and respond to instructions from the telemetry system.
8. The method of claim 7, wherein:
the CAN transceiver is powered by an independent DC-DC.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710117227.8A CN106933094B (en) | 2017-03-01 | 2017-03-01 | Dual-redundancy airborne flight control computer |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710117227.8A CN106933094B (en) | 2017-03-01 | 2017-03-01 | Dual-redundancy airborne flight control computer |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN106933094A CN106933094A (en) | 2017-07-07 |
| CN106933094B true CN106933094B (en) | 2020-08-18 |
Family
ID=59424591
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710117227.8A Active CN106933094B (en) | 2017-03-01 | 2017-03-01 | Dual-redundancy airborne flight control computer |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106933094B (en) |
Families Citing this family (19)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108011791A (en) * | 2017-11-10 | 2018-05-08 | 中国航空工业集团公司西安航空计算技术研究所 | A kind of airborne dual-redundancy CAN communication system configuration |
| CN108196569B (en) * | 2017-12-20 | 2021-02-09 | 中国航天空气动力技术研究院 | General unmanned aerial vehicle ground satellite station |
| CN108363644B (en) * | 2017-12-20 | 2021-08-03 | 中国航空工业集团公司西安飞行自动控制研究所 | Main and standby selection method for distributed telex asynchronous working flight control computer |
| CN108919636A (en) * | 2018-06-11 | 2018-11-30 | 北京三快在线科技有限公司 | Aircraft and its control method and computer readable storage medium |
| CN108768754B (en) * | 2018-06-28 | 2021-03-16 | 西安微电子技术研究所 | High-reliability fault-tolerant system based on bus network |
| CN109587020B (en) * | 2018-11-08 | 2020-11-27 | 中国电子科技集团公司第三十八研究所 | Dual-redundancy airborne radar communication method and device with isolation |
| CN109917897B (en) * | 2019-03-20 | 2021-01-22 | 浪潮商用机器有限公司 | Redundant board power management system and method |
| CN109976237A (en) * | 2019-04-12 | 2019-07-05 | 西安爱生技术集团公司 | A kind of unmanned aerial vehicle onboard computer remaining control circuit |
| CN112714893A (en) * | 2020-04-21 | 2021-04-27 | 深圳市大疆创新科技有限公司 | Double-flight control switching method, flight control system and aircraft |
| CN112147880B (en) * | 2020-08-26 | 2022-08-05 | 山东航天电子技术研究所 | Remote control instruction selection control method |
| CN112115659A (en) * | 2020-09-16 | 2020-12-22 | 天津津航计算技术研究所 | High-reliability system packaging integrated chip of redundancy technology |
| CN112162509B (en) * | 2020-09-25 | 2024-01-02 | 中国直升机设计研究所 | Active control rotor real-time control system based on FPGA+CPU architecture |
| CN112859579B (en) * | 2021-01-25 | 2022-11-15 | 北京大学 | Redundancy fault-tolerant control method for remote sensing networking of unmanned aerial vehicle |
| CN113590518B (en) * | 2021-08-03 | 2023-07-28 | 北京北航天宇长鹰无人机科技有限公司 | Synchronization system and method for dual redundancy data buses |
| CN113534656B (en) * | 2021-09-07 | 2022-01-21 | 中国商用飞机有限责任公司 | Telex flight backup control system and telex flight backup control method |
| CN113859318A (en) * | 2021-11-16 | 2021-12-31 | 中国铁道科学研究院集团有限公司 | Train braking device and method |
| CN114104251A (en) * | 2021-12-22 | 2022-03-01 | 北京临近空间飞艇技术开发有限公司 | Reciprocating type stratospheric airship capable of being repeatedly used during overlong navigation |
| CN114326824B (en) * | 2022-03-16 | 2022-06-03 | 电子科技大学 | Heterogeneous high-density hybrid unmanned aerial vehicle cluster topology control method based on bionic algorithm |
| CN116931415B (en) * | 2023-09-18 | 2023-12-19 | 西北工业大学 | Autonomous redundancy management method for dual-redundancy electromechanical actuating system controller |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103116285A (en) * | 2011-11-17 | 2013-05-22 | 上海航天测控通信研究所 | Double central processing unit (CPU) unibus computer system for deep space probe |
| CN103455005A (en) * | 2013-09-06 | 2013-12-18 | 北京四方继保自动化股份有限公司 | Controller redundancy and switching method |
| CN105187248A (en) * | 2015-09-16 | 2015-12-23 | 浙江众合科技股份有限公司 | Redundancy switching system |
| CN105550077A (en) * | 2015-12-10 | 2016-05-04 | 中国航空工业集团公司西安飞机设计研究所 | Backup control system |
| CN105718336A (en) * | 2014-11-30 | 2016-06-29 | 上海航空电器有限公司 | Dual-redundancy dispatching control system for aircraft alarm computer |
| CN105717787A (en) * | 2014-11-30 | 2016-06-29 | 上海航空电器有限公司 | Dual-redundancy control system and control method for intelligent power distribution device |
-
2017
- 2017-03-01 CN CN201710117227.8A patent/CN106933094B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103116285A (en) * | 2011-11-17 | 2013-05-22 | 上海航天测控通信研究所 | Double central processing unit (CPU) unibus computer system for deep space probe |
| CN103455005A (en) * | 2013-09-06 | 2013-12-18 | 北京四方继保自动化股份有限公司 | Controller redundancy and switching method |
| CN105718336A (en) * | 2014-11-30 | 2016-06-29 | 上海航空电器有限公司 | Dual-redundancy dispatching control system for aircraft alarm computer |
| CN105717787A (en) * | 2014-11-30 | 2016-06-29 | 上海航空电器有限公司 | Dual-redundancy control system and control method for intelligent power distribution device |
| CN105187248A (en) * | 2015-09-16 | 2015-12-23 | 浙江众合科技股份有限公司 | Redundancy switching system |
| CN105550077A (en) * | 2015-12-10 | 2016-05-04 | 中国航空工业集团公司西安飞机设计研究所 | Backup control system |
Also Published As
| Publication number | Publication date |
|---|---|
| CN106933094A (en) | 2017-07-07 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106933094B (en) | Dual-redundancy airborne flight control computer | |
| US8760004B2 (en) | Electrical power distribution | |
| CN106462147B (en) | Remote Terminal Unit (RTU) with universal input/output (UIO) and related configuration method | |
| RU2572565C2 (en) | Aircraft electric energy distribution network | |
| CN107347018A (en) | A kind of triple redundance 1553B bus dynamic switching methods | |
| US20130299642A1 (en) | Multifunctional Controller for a Satellite | |
| CN108693866B (en) | Device and method for updating control software in high-voltage control equipment | |
| US11155341B2 (en) | Redundant fly-by-wire systems with fault resiliency | |
| EP3738872B1 (en) | Redundant fly-by-wire systems with fault resiliency | |
| CN112714725B (en) | Vehicle control system | |
| CN105122152A (en) | Control of aircraft systems with at least two remote data concentrators for control of an aircraft system component | |
| US5560570A (en) | Automatic piloting device for aerodynes | |
| US10756566B2 (en) | Method and apparatus for supplying electric power to a device | |
| CN104583077A (en) | Electric circuit for cutting off an electric power supply comprising transistors and fuses having redundant logic | |
| TWI590987B (en) | The control system of fail-safe device and method for unmanned aerial vehicle | |
| US10054998B2 (en) | Automatically enabled airplane-mode portable power source | |
| US11787355B2 (en) | Protection circuitry and a method for protecting a vehicle power supply | |
| US10268613B2 (en) | Redundant control system devoid of programmable devices | |
| CN107302393B (en) | Main-standby interlocking system for ensuring remote control single machine | |
| EP3005005A2 (en) | Fault tolerant electronic control architecture for aircraft actuation system | |
| CN112670946B (en) | Power supply system, control method of power supply system, and electrical device | |
| US20230159038A1 (en) | Control device for controlling a vehicle function for a vehicle, and method for operating a control device | |
| CN115743522B (en) | Method and system for activating backup flight control system of large-sized conveyor | |
| CN114114894B (en) | Fly-by-wire backup control system and fly-by-wire backup control method | |
| CN112131055B (en) | Multi-mode three-motor dynamic fault-tolerant system |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CB03 | Change of inventor or designer information | ||
| CB03 | Change of inventor or designer information |
Inventor after: Zhao Lei Inventor after: Guo Ta Inventor before: Zhao Lei |