CN116483106A - Integrated unmanned aerial vehicle system for inspection and beating - Google Patents
Integrated unmanned aerial vehicle system for inspection and beating Download PDFInfo
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- CN116483106A CN116483106A CN202310480557.9A CN202310480557A CN116483106A CN 116483106 A CN116483106 A CN 116483106A CN 202310480557 A CN202310480557 A CN 202310480557A CN 116483106 A CN116483106 A CN 116483106A
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
- G05D1/106—Change initiated in response to external conditions, e.g. avoidance of elevated terrain or of no-fly zones
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- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
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- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention relates to a reconnaissance and beating integrated unmanned aerial vehicle system which comprises three parts, namely a flight sensor, a redundancy flight control computer and an executing mechanism. The flight control computer part adopts a three-redundancy hardware architecture based on FlexRay and Can buses. The actuating mechanism part comprises an aileron steering engine, a V-tail steering engine, a front wheel steering engine, an air door steering engine, an accelerator steering engine and a propeller speed regulation steering engine. The invention carries out corresponding redundancy design on each component part of the flight control system, wherein the three-redundancy flight control computer is based on the distributed architecture design of the FlexRay and the Can bus, so that the system eliminates the faults of the Bayesian general, improves the reliability, and has the advantages of strong expansibility, simple and flexible structure, low maintenance cost and the like. The system can meet the requirements of high reliability, low cost and high cost performance of the unmanned aerial vehicle.
Description
Technical Field
The invention relates to a reconnaissance integrated unmanned aerial vehicle system, in particular to a distributed flight control system redundancy architecture method based on a FlexRay and Can bus fusion similar and dissimilar redundancy concept, which is used for meeting the requirements of the reconnaissance integrated unmanned aerial vehicle on the reliability, the instantaneity, the maintainability, the universality, the expansibility and the cost performance of a flight control system under the condition of considering cost performance factors.
Background
Along with the increasingly wide application of unmanned aerial vehicles, the increasingly wide application field, the increasingly enhanced functions, the increasingly improved development, production, use and maintenance costs, and the increasingly higher requirements on the reliability of the flight control system. The reliability design of the flight control system directly affects the flight safety of the aircraft, so that the quality and the assembly process quality of components and parts are difficult to meet the system requirements simply by relying on improvement, the redundancy design technology can effectively improve the reliability and fault tolerance of the flight control system, and the redundancy technology is adopted to fundamentally improve the fault tolerance and the residual capacity of the system from the design of the architecture of the flight control system, so that the aim of softening faults to eliminate the influence of the faults on the normal operation of the system is fulfilled.
The redundant fault-tolerant flight control system is successfully applied to the organic machines such as civil airliners, fighters and the like, and the fault rate of the flight control system is reduced to 10 -7 -10 -10 /flight hours. However, the redundant fault-tolerant flight control system of the unmanned aerial vehicle cannot meet the requirements of the unmanned aerial vehicle on size, power consumption, price and the like, and cannot be directly applied to the unmanned aerial vehicle. With the development of microelectronic, electronic, computer, bus and other technologies, electronic devices have been more integrated, more powerful, and less bulky, lighter, less powerful, and cheaper. Industrial electronics are widely used, and generally develop at a much higher rate than avionics, but are less reliable. How to reasonably apply advanced industrial products to avionic equipment, fully utilize the progress of industrial technology to improve the product performance, meet high reliability and reduce cost, and is always a problem to be solved by unmanned aerial vehicle flight control system designers.
Disclosure of Invention
The invention aims to provide a reconnaissance and beating integrated unmanned aerial vehicle system, and aims to effectively improve the reliability and fault tolerance of a large-scale medium-high altitude long-endurance reconnaissance and beating integrated unmanned aerial vehicle flight control system under the condition of low cost.
The invention provides a reconnaissance and beating integrated unmanned aerial vehicle system, which comprises three parts, namely a flight sensor, a redundancy flight control computer and an actuating mechanism. The flight control computer part adopts a three-redundancy hardware architecture based on FlexRay and Can buses. The actuating mechanism part comprises an aileron steering engine, a V-tail steering engine, a front wheel steering engine, an air door steering engine, an accelerator steering engine and a propeller speed regulation steering engine.
Preferably, the redundant design of the flight sensor system adopts a multi-system multi-frequency point scheme, and sensor signal redundancy is realized through data fusion by installing dissimilar navigation equipment, and the method specifically comprises satellite navigation, inertial navigation, combined navigation, an atmospheric data system, a radar altimeter, a wind vane sensor, a ground contact switch and engine/fuel/power distribution/landing gear monitoring. In the aspect of inertial navigation, a high-precision fiber optic gyroscope is adopted, so that an unmanned aerial vehicle can throw a terminal guidance bomb with high precision. The same sensor can provide different navigation positioning and attitude determination information, and redundant backup relations exist between the sensors; altitude information sources include an atmospheric data system, satellite (GPS, BD2, GLONASS) navigation, and radar altimeter; the position information sources are satellite (GPS, BD2, GLONASS) navigation and inertial navigation; the gesture information source is provided with a high-precision optical fiber gyro and a movement differential gesture (gesture information in two directions of rolling and yaw is provided through movement differential of double antennae) provided by a satellite navigation system. Wherein, "altitude information source, position information source, attitude information source" refers to an onboard sensor device capable of providing altitude information, position information, and attitude information.
Preferably, the redundancy design of the actuator part is realized by the electrical redundancy of the servo system through the pneumatic control surface (aileron, V tail) slice design. The pneumatic control surface is designed in a slicing way, the control surfaces of ailerons and V tails on two sides of an unmanned aerial vehicle are respectively divided into two pieces, 4 control surfaces close to one side of a machine body are called inner control surfaces, 4 control surfaces far away from one side of the machine body are called outer control surfaces, and if the control surfaces are blocked due to failure of the inner steering engine, the outer control surfaces are controlled to bring an airplane back. The electrical redundancy of the servo system means that two sets of control circuits are designed in the steering engine controller, one set is effective by default in normal operation, and the other set is switched to after failure occurs.
Preferably, the redundancy design of the flight control computer part is in a triple-modular redundancy mode, and particularly similar redundancy is formed by configuring three identical flight control boards to run similar programs (only different bus drives). The bottom layer provides hardware synchronization (nanosecond level) for inter-board communication (including interface boards) over the FlexRay bus.
Preferably, in the three-mode redundancy mode of the flight control computer, three flight control boards acquire the same sensor data, and respectively calculate control laws to obtain control surface instructions to drive a steering engine to realize control closed loop. The control surface command can isolate single machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, and the fault board card is positioned according to accumulated deviation of the control command so as to enable dual-mode redundancy work.
Preferably, the flight control system is divided into two voting modes according to the difference of the voting points of the key pneumatic control surface control instructions to form redundancy. a) A steering engine end value voting mode, wherein a steering surface command is transmitted to a steering engine through a CAN bus, and a pneumatic steering surface is driven after median voting is carried out on a steering engine end; b) And in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, performs median voting, and then sends the control surface instruction to the steering engine through the RS422, and when the steering engine CAN bus fails/loses a signal, the data instruction of the RS422 is changed. The design of the two voting modes is mutually redundant, namely, after a fault occurs, the control instruction of the control surface still has triple redundancy, and the reliability is greatly improved.
Preferably, the dual-mode redundancy mode of the flight control computer is as follows: after judging the single machine fault, the steering engine end/interface board determines the on-duty machine according to the three-machine fault conditions indicated by each flight control board and the serial numbers of the flight control boards, drives the steering engine according to the control surface instruction of the on-duty machine, and does not vote on the control surface instruction any more. And determining a secondary fault board card according to the information such as heartbeat, board card self-detection and the like, and starting a single mode.
Preferably, in the single-mode of the flight control computer, if the on-duty machine fails again in the dual-mode redundancy mode, the control surface command/interface board switches the on-duty right to a normal single machine, and the flight control is independently completed. Compared with the prior art, the redundancy architecture scheme of the large-sized medium-high-altitude long-endurance integrated unmanned aerial vehicle flight control system provided by the invention integrates multiple redundancy design ideas, and corresponding redundancy designs are carried out in each component part of the flight control system, especially three-redundancy flight control computers, based on the distributed architecture design of FlexRay and Can buses, so that the system eliminates the faults of the Bayesian general, improves the reliability, and has the advantages of strong expansibility, simple and flexible structure, low maintenance cost and the like. The fault tolerance technology and the application of industrial mature products enable the system to simultaneously meet the requirements of high reliability, low cost and high cost performance of the unmanned aerial vehicle.
Drawings
Fig. 1 is a block diagram of a flight control system in a redundancy architecture scheme of a unmanned aerial vehicle system with integrated inspection and beating.
Fig. 2 is a schematic diagram of a redundant flight control computer in a redundant architecture scheme of a unmanned aerial vehicle system with integrated scouting and beating.
Fig. 3 is a working state transition diagram of a redundant flight control computer in a redundant architecture scheme of a beaten-to-beaten integrated unmanned aerial vehicle system.
Fig. 4 is a schematic diagram of the connection between a three-redundancy flight control computer and the outside in the redundancy architecture scheme of the unmanned aerial vehicle system with integrated scouting and beating.
Fig. 5 is a schematic diagram of a redundant design scheme of a pneumatic control surface in an actuator in a redundant architecture scheme of a man-machine integrated system.
Detailed Description
The present invention will be described in detail with reference to the accompanying drawings.
Fig. 1 is a block diagram of a flight control system in a redundancy architecture scheme of a unmanned aerial vehicle system with a supervisory-and-beating integrated system, wherein the whole flight control system comprises three parts, namely a three-redundancy flight control computer, a sensor and an executing mechanism. The key sensors are provided with a set of inertial navigation system, satellite navigation, an atmospheric data computer and a windvane attack angle/sideslip angle sensor; the general sensor comprises a radar altimeter and a grounding switch; other monitoring content includes engine monitoring, fuel system health monitoring, power distribution network monitoring, landing gear monitoring, etc. The actuating mechanism comprises ailerons, V tails, front wheel steering, air doors, an accelerator, a propeller speed regulation steering engine and a steering engine controller. The inertial navigation system is fused with satellite navigation system, atmospheric data system and angle of attack sideslip angle sensor data, and after resolving by the combined navigation algorithm, the combined navigation system is formed, the combined navigation data is output, the combined navigation system is communicated with the flight control computer, the radar altimeter, the ground contact switch and other monitoring devices are also communicated with the flight control computer, wherein the altitude information sources are the atmospheric data system, satellite (GPS, BD2, GLONASS) navigation and the radar altimeter; the position information source is satellite (GPS, BD2, GLONASS) navigation and inertial navigation; the gesture information source is a moving differential gesture provided by a high-precision fiber-optic gyroscope and a satellite navigation system.
Fig. 2 is a schematic diagram of a redundant flight control computer in a method for constructing a redundancy architecture of a unmanned aerial vehicle system with a scouting and beating integrated system, wherein the flight control computer hardware comprises 1 VPX chassis, 2 power boards, 3 flight control boards (flight control computers), 1 interface board, 1 communication bottom board, 1 communication back board and 3 aviation plugs (connectors).
The power panel is responsible for completing the conversion of external primary power input to form a 12V secondary power supply, and completing the power characteristic requirements such as voltage-withstanding surge and the like, and supplying power to the flight control panel and the interface panel through the bottom plate.
The three flight control boards have the same hardware and the same software (except bus driving) to form similar redundancy. The bottom layer provides hardware synchronization (nanosecond level) and FlexRay bus for inter-board communication (including interface board). The flight control board receives sensor signals through an onboard serial port and DI, carries out flight control law resolving, and a resolving result, namely a control surface instruction, is output to the interface board through FlexRay and then is converted into RS422/DA/PWM to each steering engine. Meanwhile, key pneumatic control surface instructions of each board card are directly output to a steering engine controller through the same CAN bus.
The interface board is responsible for simultaneously collecting information of three flight control boards through a FlexRay bus, and each control instruction is voted by a median theorem and converted into signals such as RS 422/DA/PWM/IO. And meanwhile, the secondary reconstruction after the primary failure can be performed.
The base plate provides signal pull-up and pull-down processing for each plate and a FlexRay bus; the back plate provides a cable transition from the patch panel to the connector.
Fig. 4 is a schematic diagram of the connection between a three-redundancy flight control computer and the outside of the redundancy architecture method of the large-scale middle-high altitude observation and beating integrated unmanned aerial vehicle flight control system. The three flight control boards realize synchronization through a hardware bottom layer, realize communication among flight control boards through a FlexRay bus, acquire sensor information through modes such as serial ports and IO, calculate control plane instructions according to a control law, and drive a steering engine through a CAN bus and the serial ports. In order to realize the system redundancy, the three flight control computers run the same program (except bus configuration difference), collect the same sensor data, give the steering surface deflection instruction at the same time, the steering engine controller carries on the median processing; the key sensors, such as integrated navigation, send data to three flight control boards simultaneously; the steering engine controller adopts a CAN bus interface, and simultaneously collects the instructions of 3 flight control boards, and drives the steering engine through median processing. Besides key sensors, interfaces of other steering engines such as throttle, front wheels and the like are arranged on an interface board in other states needing to be monitored. Wherein, the key sensor includes: inertial navigation, satellite navigation, atmospheric data system, angle of attack sideslip angle sensor.
The pneumatic servo adopts the configuration of a single-channel controller and a steering engine; the air door, the throttle and the paddle control steering engine adopt an integrated steering engine. The three flight control boards realize synchronous output through the CAN bus, meanwhile, the interface board collects the three flight control board control surface instructions according to the Flexray, and the three flight control board control surface instructions are converted into RS422 instructions after median voting and input, and one path of RS422 of each steering engine is realized.
Each steering engine controller simultaneously collects control surface instructions of three flight control boards through a CAN bus, and votes according to a median theorem to drive a steering engine; meanwhile, in order to ensure safety, a control surface instruction (output after voting by an interface board) of an RS422 serial port is used under the fault condition of the CAN bus; and other steering engines such as an air door, an accelerator, a paddle control and the like are voted on the interface board, and the steering engines are driven by PWM or DA signals.
Fig. 3 is a working state transition diagram of a redundant flight control computer in a redundant architecture scheme of a large-scale middle-high altitude long-endurance integrated unmanned aerial vehicle flight control system. Under normal conditions, the flight control computer works in a triple-mode redundancy mode, and is shifted to a double-mode redundancy mode when a single-machine fault occurs, and is shifted to a single-mode when a secondary fault occurs.
Triple modular redundancy mode: the three flight control boards acquire the same sensor data, respectively calculate control laws to obtain control surface instructions, and drive the steering engine to realize control closed loop. The control surface command can isolate single machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, and the fault board card is positioned according to accumulated deviation of the control command so as to enable dual-mode redundancy work. And the voting points are divided into the following groups according to the differences: a) The steering engine end value voting mode is that steering engine commands are transmitted to the steering engine through a CAN bus, and the pneumatic steering engine is driven by median voting at the steering engine end; b) And in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, performs median voting, and then sends the control surface instruction to the steering engine through the RS422, and when the steering engine CAN bus fails/loses a signal, the data instruction of the RS422 is changed.
And after judging the single machine fault in the dual-mode redundancy mode, the steering engine end/interface board determines the on-duty machine according to the serial numbers of the flight control boards and the three machine fault conditions indicated by each flight control, drives the steering engine according to the control surface instruction of the on-duty machine, and does not vote on the control surface instruction any more. And determining a secondary fault board card according to the information such as heartbeat, board card self-detection and the like, and starting a single mode.
In the single mode and the dual mode redundant mode, if the on-duty machine fails again, the control surface command/interface board switches the on-duty right to a normal single machine, and flight control is independently completed.
Fig. 5 is a schematic diagram of a redundant design scheme of a pneumatic control surface in an actuator in a redundant architecture scheme of a man-machine integrated system. As shown in the figure, the control surfaces of the ailerons and the V-tails on two sides of the unmanned aerial vehicle are respectively divided into two pieces, and 4 control surfaces close to one side of the unmanned aerial vehicle body comprise an inner left aileron, an inner right aileron, an inner left V-tail and an inner right V-tail; and 4 control surfaces far away from one side of the engine body, wherein the control surfaces comprise an outer left aileron, an outer right aileron, an outer left V tail and an outer right V tail, and if the control surfaces are blocked due to the failure of the inner steering engine, the outer control surfaces are controlled to bring the aircraft back. The electrical redundancy of the servo system means that two sets of identical circuits A and B are designed in the steering engine controller, signals output by a default A path are effective in normal operation, the circuit is switched to a B path after faults occur, and if the B path fails again, the system reports faults.
The invention has been described in detail in connection with the drawings, but it will be apparent to those skilled in the art that the description is intended to be construed in a limited sense only by the appended claims. The scope of the invention is not limited by the description. Any changes or substitutions that would be readily apparent to one skilled in the art within the scope of the present disclosure are intended to be encompassed within the scope of the present invention. Therefore, the protection scope of the present invention should be subject to the protection scope of the claims.
Claims (1)
1. An inspection and beating integrated unmanned aerial vehicle system, the system include three parts of flight sensor, surplus flight control computer and actuating mechanism, its characterized in that: the method for constructing the redundancy framework of the flight control system is characterized in that each part is subjected to corresponding redundancy design, a flight control computer part adopts a three-redundancy hardware framework based on FlexRay and Can buses, and an actuating mechanism part comprises an aileron steering engine, a V-tail steering engine, a front wheel steering engine, a throttle steering engine, an accelerator steering engine and a propeller speed regulation steering engine;
the redundant design of the flight sensor system adopts a multi-system multi-frequency point scheme, realizes sensor signal redundancy through data fusion by installing dissimilar navigation equipment, and specifically comprises satellite navigation, inertial navigation, combined navigation, an atmosphere data system, a radar altimeter, a weathervane sensor, a touchdown switch and engine/fuel/power distribution/landing gear monitoring; in the aspect of inertial navigation, the redundant design of the flight sensor system adopts a high-precision fiber optic gyroscope, so that an unmanned aerial vehicle can realize high-precision throwing of a terminal guidance bomb; the same sensor can provide different navigation positioning and attitude determination information, and redundant backup relation exists between the sensor and the sensor; the altitude information source comprises an atmospheric data system, a satellite navigation and a radar altimeter; the position information source is satellite navigation and inertial navigation; the gesture information source is provided with a moving differential gesture provided by a high-precision fiber-optic gyroscope and a satellite navigation system;
the redundant design of the actuating mechanism part is realized by the electrical redundancy of a servo system through the pneumatic control surface slicing design; the pneumatic control surface is designed in a slicing way, namely the control surfaces of ailerons and V tails on two sides of an unmanned aerial vehicle are respectively divided into two pieces, 4 control surfaces close to one side of a machine body are called inner control surfaces, 4 control surfaces far away from one side of the machine body are called outer control surfaces, and if the control surfaces are blocked due to the failure of the inner steering engine, the outer control surfaces are controlled to bring the aircraft back; the electrical redundancy of the servo system is that two sets of control circuits are designed in the steering engine controller, one set is effective by default in normal operation, and the other set is switched to after failure occurs;
the redundancy design of the flight control computer part is a triple-modular redundancy mode, and particularly, similar programs are operated by configuring three identical flight control boards to form similar redundancy; the bottom layer provides hardware synchronization, and inter-board communication is carried out through a FlexRay bus; the three-mode redundancy mode of the flight control computer is that three flight control boards collect the same sensor data, and respectively calculate control laws to obtain control surface instructions, and drive a steering engine to realize control closed loop; the control surface instruction can isolate single machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, and according to accumulated deviation of the control instruction, the fault board card is positioned, and dual-mode redundancy work is started;
the dual-mode redundancy mode is: after judging the single machine fault, the steering engine end/interface board determines the on-duty machine according to the three-machine fault conditions indicated by each flight control board and the serial numbers of the flight control boards, drives the steering engine according to the control surface instruction of the on-duty machine, and does not vote on the control surface instruction any more; determining a secondary fault board card according to the heartbeat and board card self-checking information, and starting a single-mode;
the single mode is that if the on-duty machine fails again in the dual-mode redundancy mode, the control surface command/interface board card switches the on-duty right to a normal single machine to independently complete flight control;
the median voting is divided into two voting modes according to the difference of the voting points of the control instructions of the key pneumatic control surfaces in the flight control system to form redundancy; a) A steering engine end value voting mode, wherein a steering surface command is transmitted to a steering engine through a CAN bus, and a pneumatic steering surface is driven after median voting is carried out on a steering engine end; b) And in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, performs median voting, and then sends the control surface instruction to the steering engine through the RS422, and when the steering engine CAN bus fails/loses a signal, the data instruction of the RS422 is changed.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202310480557.9A CN116483106A (en) | 2023-04-28 | 2023-04-28 | Integrated unmanned aerial vehicle system for inspection and beating |
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| CN202310480557.9A CN116483106A (en) | 2023-04-28 | 2023-04-28 | Integrated unmanned aerial vehicle system for inspection and beating |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117215177A (en) * | 2023-11-09 | 2023-12-12 | 北京控制工程研究所 | Heaven and earth round trip integrated control system and control method |
| CN118348881A (en) * | 2024-05-20 | 2024-07-16 | 西安格儒电子科技有限公司 | A flight control computer system with trimmable redundancy |
| CN118409543A (en) * | 2024-05-13 | 2024-07-30 | 陕西星汉磐芯科技有限公司 | Three-redundancy multi-core integrated electronic system |
-
2023
- 2023-04-28 CN CN202310480557.9A patent/CN116483106A/en not_active Withdrawn
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117215177A (en) * | 2023-11-09 | 2023-12-12 | 北京控制工程研究所 | Heaven and earth round trip integrated control system and control method |
| CN118409543A (en) * | 2024-05-13 | 2024-07-30 | 陕西星汉磐芯科技有限公司 | Three-redundancy multi-core integrated electronic system |
| CN118348881A (en) * | 2024-05-20 | 2024-07-16 | 西安格儒电子科技有限公司 | A flight control computer system with trimmable redundancy |
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Application publication date: 20230725 |