CN108776486B - Redundancy architecture method for flight control system of large-scale medium-high altitude scouting and printing integrated unmanned aerial vehicle - Google Patents
Redundancy architecture method for flight control system of large-scale medium-high altitude scouting and printing integrated unmanned aerial vehicle Download PDFInfo
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Abstract
The invention discloses a redundancy architecture method for a flight control system of a large-scale medium-high altitude scouting and printing integrated unmanned aerial vehicle, 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 a Can bus. 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 of the flight control system, wherein the three-redundancy flight control computer is based on the distributed architecture design of FlexRay and Can bus, so that the system eliminates the faults of the Byzantine general, improves the reliability, and has the advantages of strong expansibility, simple and flexible structure, low maintenance cost and the like. The system simultaneously meets the requirements of high reliability, low cost and high cost performance of the unmanned aerial vehicle.
Description
Technical Field
The invention relates to a redundancy architecture method for a flight control system of a large-scale medium-high altitude scouting and printing integrated unmanned aerial vehicle, in particular to a redundancy architecture method for a distributed flight control system based on the fusion similar and dissimilar redundancy ideas of FlexRay and Can buses, which is used for meeting the requirements of the large-scale medium-high altitude scouting and printing integrated unmanned aerial vehicle on the reliability, the real-time performance, the maintainability, the universality, the expansibility and the cost performance of the flight control system under the condition of considering cost performance factors.
Background
Along with the increasingly wide application of unmanned aerial vehicles, the application field is continuously enlarged, the function is continuously enhanced, the development, production, use and maintenance cost is continuously improved, and the requirement on the reliability of a flight control system is also increasingly high. The flight safety of the aircraft is directly influenced by the advantages and disadvantages of the reliability design of the flight control system, so that the system requirements are difficult to meet by simply depending on the improvement of the quality of components and parts and the quality of an assembly process, and the redundancy design technology can effectively improve the reliability and fault-tolerant capability of the flight control system, so that the fault-tolerant capability and residual capability of the system must be fundamentally improved by starting from the design of the architecture of the flight control system, and the fault softening is realized so as to achieve the purpose of eliminating the influence of faults on the normal work of the system.
The redundancy fault-tolerant flight control system is successfully applied to civil aircraft, fighters and other human-machines, and the fault rate of the flight control system is reduced to 10-7-10-10Hour of flight. However, the redundant fault-tolerant flight control system with an unmanned aerial vehicle cannot meet the requirements of the unmanned aerial vehicle on volume, power consumption, price and the like, and cannot be directly applied to the unmanned aerial vehicle. With the development of technologies such as microelectronics, electronics, computers, buses and the like, electronic devices have higher integration degree, stronger functions, smaller volume, lighter weight, smaller power and cheaper price. Industrial electronics is widely used, and its development speed is usually much higher than that of avionics, but its reliability is also low. How to reasonably apply advanced industrial products to avionic equipment and make full use of industrial technologyThe progress of art improves and produces the property ability, and reduce cost when satisfying high reliability is the problem that unmanned aerial vehicle flight control system designer needs to solve all the time.
Disclosure of Invention
The invention aims to provide a redundancy architecture method for a flight control system of a large-scale medium-high altitude survey and flight integrated unmanned aerial vehicle, aiming at effectively improving the reliability and fault-tolerant capability of the flight control system of the large-scale medium-high altitude long-endurance survey and flight integrated unmanned aerial vehicle under the condition of low cost.
The invention provides a redundancy architecture method for a flight control system of a large-scale medium-high altitude scouting and fighting integrated unmanned aerial vehicle. The flight control computer part adopts a three-redundancy hardware architecture based on FlexRay and a Can bus. 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 redundancy design of the flight sensor system adopts a multi-system multi-frequency point scheme, non-similar navigation equipment is installed, and sensor signal redundancy is realized through data fusion, wherein the redundancy design specifically comprises satellite navigation, inertial navigation, combined navigation, an atmospheric data system, a radar altimeter, a vane sensor, a ground contact switch and engine/fuel oil/power distribution/undercarriage monitoring. In the aspect of inertial navigation, a high-precision fiber-optic gyroscope is adopted, so that the unmanned aerial vehicle can realize high-precision terminal-throwing guidance bomb. The same sensor can provide different navigation positioning and attitude determination information, and redundant backup relations exist among the sensors; the altitude information source comprises an atmospheric data system, satellite (GPS, BD2, GLONASS) navigation and a radar altimeter; the position information source is satellite (GPS, BD2, GLONASS) navigation and inertial navigation; the attitude information source comprises a high-precision fiber-optic gyroscope and a moving differential attitude provided by a satellite navigation system (the attitude information in two directions of rolling and yawing is provided through the moving differential of double antennae). Wherein "height information source, position information source, attitude information source" refers to an onboard sensor device capable of providing height information, position information, and attitude information.
Preferably, the redundancy design of the actuator part is realized by the aid of an aerodynamic control surface (aileron, V tail) segment design and electrical redundancy of a servo system. The slicing design of the pneumatic control surface is that the control surfaces of ailerons and V tails on two sides of the unmanned aerial vehicle are respectively divided into two pieces, 4 control surfaces close to one side of the vehicle body are called inner side control surfaces, 4 control surfaces far away from one side of the vehicle body are called outer side control surfaces, and if the inner side steering engine fails to cause the control surfaces to be blocked, the outer side control surfaces are controlled to bring the safety belt of the vehicle back. The electrical redundancy of the servo system means that two sets of control circuits are designed in the steering engine controller, one set of control circuits is defaulted to be effective when the steering engine controller works normally, and the other set of control circuits is switched to after a fault occurs.
Preferably, the redundancy design of the flight control computer part is a triple modular redundancy mode, and specifically, similar redundancy is formed by configuring three identical flight control board cards and running similar programs (only bus drivers are different). The bottom layer provides hardware synchronization (nanosecond level) and inter-board communication (including interface boards) is performed through a FlexRay bus.
Preferably, the triple-modular redundancy mode of the flight control computer is that the three flight control boards acquire the same sensor data, respectively calculate a control law to obtain a control plane instruction, and drive the steering engine to realize a control closed loop. The control plane instruction can isolate single-machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, the fault board card is positioned according to accumulated deviation of the control instruction, and dual-mode redundancy work is started.
Preferably, the flight control system is divided into two voting modes according to different voting points of the control instruction of the key pneumatic control surface, so that redundancy is formed. a) In the steering engine end value voting mode, a control plane instruction is transmitted to the steering engine through a CAN bus, and a pneumatic control plane is driven after the steering engine end carries out medium value voting; b) and in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, carries out median voting and then sends the control surface instruction to the steering engine through RS422, and when the CAN bus of the steering engine has a fault or loses a signal, the data instruction of the RS422 is used instead. The two voting modes are mutually redundant in design, 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: and after the single machine fault is judged, the rudder machine end/interface board determines the current aircraft according to the sequence of the serial numbers of the flight control boards according to the three-machine fault conditions indicated by each flight control, drives the steering engine according to the control plane instruction of the current aircraft, and does not vote on the control plane instruction any more. And determining a secondary fault board card according to information such as heartbeat and board card self-checking, and starting a single mode.
Preferably, the single mode of the flight control computer is that if the on-duty aircraft fails again in the dual-mode redundancy mode, the control plane instruction/interface board card switches the on-duty right to a normal single machine to independently complete flight control.
Compared with the prior art, the redundancy architecture scheme of the flight control system of the large-scale medium-high altitude long-endurance integrated unmanned aerial vehicle integrates various redundancy design ideas, corresponding redundancy design is carried out in all components of the flight control system, particularly a three-redundancy flight control computer is designed on the basis of a distributed architecture of a FlexRay bus and a Can bus, so that the system eliminates the faults of the Byzantine general, the reliability is improved, and the redundancy architecture scheme has the advantages of strong expansibility, simple and flexible structure, low maintenance cost and the like. The fault-tolerant 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 redundant architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight inspection and flight integrated unmanned aerial vehicle provided by the invention.
Fig. 2 is a schematic diagram of a redundancy flight control computer in a redundancy architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight inspection and flight integrated unmanned aerial vehicle provided by the invention.
Fig. 3 is a working state transition diagram of a redundancy flight control computer in a redundancy architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight inspection and flight integrated unmanned aerial vehicle provided by the invention.
Fig. 4 is a schematic diagram of connection between a triple-redundancy flight control computer and the outside in a redundancy architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight survey and flight integrated unmanned aerial vehicle provided by the invention.
Fig. 5 is a schematic diagram of a redundant design scheme of a pneumatic control surface in an execution mechanism in a redundant architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight inspection and flight integrated unmanned aerial vehicle provided by the invention.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings.
Fig. 1 is a block diagram of a flight control system in a redundancy architecture scheme of a large-scale medium-high altitude long-endurance flight inspection and flight control integrated unmanned aerial vehicle flight control system provided by the invention, and the whole flight control system comprises three redundancy flight control computers, a sensor and an execution mechanism. The key sensors are provided with a set, and comprise an inertial navigation system, a satellite navigation system, an atmospheric data computer and a weathervane attack angle/sideslip angle sensor; common sensors include radar altimeters, touchdown switches; other monitoring content includes engine monitoring, fuel system health monitoring, distribution network monitoring, landing gear monitoring, and the like. The actuating mechanism comprises ailerons, a V-shaped tail, front wheel steering, an air door, an accelerator, a propeller speed regulation steering engine and a steering engine controller. The system comprises an inertial navigation system, a satellite navigation system, an atmospheric data system and an attack angle sideslip angle sensor, wherein the inertial navigation system is fused with the satellite navigation system, the atmospheric data system and the attack angle sideslip angle sensor, and after the data are resolved by a combined navigation algorithm, a combined navigation system is formed, combined navigation data are output and are connected and communicated with a flight control computer, a radar altimeter, a ground contact switch and other monitoring devices are also respectively connected and communicated with the flight control computer, wherein altitude information sources are an atmospheric data system, satellite (GPS, BD2, GLONASS) navigation and a radar altimeter; the position information source is satellite (GPS, BD2, GLONASS) navigation and inertial navigation; the attitude information source is a moving differential attitude provided by a high-precision fiber-optic gyroscope and a satellite navigation system.
Fig. 2 is a schematic diagram of a redundancy flight control computer in the redundancy architecture method of the flight control system of the large-scale medium-high altitude reconnaissance and flight integrated unmanned aerial vehicle, and the flight control computer hardware comprises 1 VPX chassis, 2 power supply 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 external primary power input conversion to form a 12V secondary power supply, completing power characteristic requirements such as voltage and surge resistance and the like, and supplying power to the flight control panel and the interface board through the bottom plate.
Three flight control boards are identical in hardware and identical in software (except bus drivers) to form similar redundancy. The bottom layer provides hardware synchronization (nanosecond level) and provides a FlexRay bus for inter-board communication (including interface boards). And the flight control board receives sensor signals through an onboard serial port and the DI, performs flight control law calculation, outputs a calculation result, namely a control plane instruction to an interface board through FlexRay, and then turns RS422/DA/PWM to each steering engine. Meanwhile, the key pneumatic control surface instruction of each board card is directly output to the steering engine controller through the same CAN bus.
The interface board is responsible for collecting information of the three flight control board cards through a FlexRay bus at the same time, and each control instruction is voted by using a median theorem and converted into signals such as RS422/DA/PWM/IO and the like. And meanwhile, secondary reconstruction after primary failure can be carried out.
The bottom board provides signal up-down pulling processing for each board and provides a FlexRay bus; the backplane provides patch panel to connector cable transition.
Fig. 4 is a schematic diagram of connection between a triple-redundancy flight control computer and the outside in the method for constructing the redundancy architecture of the flight control system of the large-scale medium-high altitude scouting and fighting integrated unmanned aerial vehicle provided by the invention. The three flight control boards are synchronized through a hardware bottom layer, communication between the flight control boards is achieved through a FlexRay bus, sensor information is acquired through serial ports, IO and other modes, a control surface instruction is resolved according to a control law, and a steering engine is driven through a CAN bus and a serial port. In order to realize system redundancy, the three flight control computers run the same program (except bus configuration difference), acquire the same sensor data, and simultaneously give a control plane deflection instruction, and a steering engine controller performs median processing; a key sensor such as combined navigation simultaneously sends data to three flight control panels; the steering engine controller adopts a CAN bus interface, simultaneously acquires the instructions of 3 flight control boards, and drives the steering engine through median processing. Except key sensors, other states needing monitoring, interfaces of other steering engines such as an accelerator, a front wheel and the like are arranged on an interface board. Wherein, key sensor includes: inertial navigation, satellite navigation, an atmospheric data system and an attack angle sideslip angle sensor.
The pneumatic servo adopts a configuration of a single-channel controller and a steering engine; the air door, the accelerator and the paddle control steering engine adopt an integrated steering engine. The three flight control boards realize synchronous output through a CAN bus, the interface board collects the control plane instructions of the three flight control boards according to Flexray, the instructions are converted into RS422 instructions to be input after median voting, and one RS422 is arranged on each steering engine.
Each steering engine controller simultaneously collects control plane instructions of three paths of flight control boards through a CAN bus, votes according to a median theorem and drives a steering engine; meanwhile, in order to guarantee safety, the CAN bus fault condition uses a control surface instruction (output after voting by an interface board) of an RS422 serial port; other steering engines such as an air door, an accelerator, a paddle control and the like are voted on an interface board, and the steering engines are driven through PWM or DA signals.
Fig. 3 is a working state transition diagram of a redundancy flight control computer in a redundancy architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance flight inspection and flight integrated unmanned aerial vehicle provided by the invention. Under normal conditions, the flight control computer works in a triple-modular redundancy mode, and is transferred to a double-modular redundancy mode when a single computer fails, and is transferred to a single-modular redundancy mode when a secondary failure occurs.
Triple modular redundancy mode: the three flight control boards collect the same sensor data, respectively calculate a control law to obtain a control plane instruction, and drive the steering engine to realize closed-loop control. The control plane instruction can isolate single-machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, the fault board card is positioned according to accumulated deviation of the control instruction, and dual-mode redundancy work is started. According to different voting points, the method is divided into the following steps: a) in the steering engine end value voting mode, a control plane instruction is transmitted to the steering engine through a CAN bus, and a pneumatic control plane is voted and driven in the steering engine end value voting mode; b) and in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, carries out median voting and then sends the control surface instruction to the steering engine through RS422, and when the CAN bus of the steering engine has a fault or loses a signal, the data instruction of the RS422 is used instead.
And after a single machine fault is judged in the dual-mode redundancy mode, the rudder machine end/interface board determines the current aircraft according to the three-machine fault conditions indicated by each flight control board and the sequence of the serial numbers of the flight control boards, drives the steering engine according to the control plane instruction of the current aircraft, and does not vote on the control plane instruction any more. And determining a secondary fault board card according to information such as heartbeat and board card self-checking, and starting a single mode.
And in the single-mode and the dual-mode redundancy mode, if the on-duty aircraft fails, the control surface instruction/interface board card switches the on-duty right to a normal single machine to independently complete flight control.
Fig. 5 is a schematic diagram of a redundancy design scheme of a pneumatic control surface in an execution mechanism in a redundancy architecture scheme of a flight control system of a large-scale medium-high altitude long-endurance unmanned aerial vehicle. 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 the 4 control surfaces close to one side of the body comprise an inner left aileron, an inner right aileron, an inner left V tail and an inner right V tail; keep away from 4 rudder faces of organism one side, including the left aileron in the outside, the right aileron in the outside, the left V tail in the outside, the right V tail in the outside, if inboard steering wheel trouble leads to the rudder face card to die, then control outside rudder face and take the aircraft safety belt back. The electric redundancy of the servo system refers to that two sets of circuits A and B which are the same are designed in the steering engine controller, the signal output by the path A is defaulted to be effective when the steering engine controller works normally, the steering engine controller is switched to the path B after a fault occurs, and if the path B fails again, the system reports the fault.
The present invention has been described in detail with reference to the accompanying drawings, but it should be understood by those skilled in the art that the description is only for the purpose of explaining the claims. The scope of the invention is not limited by the description. Any changes or substitutions that can be easily made by those skilled in the art within the technical scope of the present disclosure are intended to be included within the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (1)
1. The redundancy architecture method of the flight control system of the large-scale medium-high altitude reconnaissance and flight control integrated unmanned aerial vehicle comprises a flight sensor, a redundancy flight control computer and an execution mechanism, and is characterized in that: the redundancy construction method of the flight control system is characterized in that corresponding redundancy design is carried out on each part, the flight control computer part adopts a three-redundancy hardware architecture based on FlexRay and a Can bus, and the execution 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 redundancy design of the flight sensor system adopts a multi-system multi-frequency point scheme, non-similar navigation equipment is installed, and sensor signal redundancy is realized through data fusion, wherein the redundancy design specifically comprises satellite navigation, inertial navigation, combined navigation, an atmospheric data system, a radar altimeter, a vane sensor, a ground contact switch and engine/fuel oil/power distribution/undercarriage monitoring; in the aspects of redundancy design and inertial navigation of the flight sensor system, a high-precision fiber-optic gyroscope is adopted, so that the unmanned aerial vehicle can realize high-precision terminal throwing guidance bomb; the same sensor can provide different navigation positioning and attitude determination information, and redundant backup relations exist among the sensors; the altitude information source comprises an atmospheric data system, satellite navigation and a radar altimeter; the position information source comprises satellite navigation and inertial navigation; the attitude information source comprises a high-precision fiber-optic gyroscope and a moving differential attitude provided by a satellite navigation system;
the redundancy design of the executing mechanism part is realized by the fragmentation design of a pneumatic control surface and the electrical redundancy of a servo system; the slicing design of the pneumatic control surface is that the control surfaces of ailerons and V tails on two sides of the unmanned aerial vehicle are respectively divided into two pieces, wherein 4 control surfaces close to one side of the vehicle body are called inner control surfaces, 4 control surfaces far away from one side of the vehicle body are called outer control surfaces, and if the control surfaces are blocked due to the fault of the inner control machine, the outer control surfaces are controlled to bring the safety belt of the vehicle back; the electrical redundancy of the servo system is to design two sets of control circuits in the steering engine controller, one set of control circuits is defaulted to be effective during normal work, and the other set of control circuits is switched to after a fault occurs;
the redundancy design of the flight control computer part is a triple modular redundancy mode, and specifically, three identical flight control board cards are configured, and similar programs are operated 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 the three flight control boards acquire the same sensor data, respectively solve a control law to obtain a control plane instruction, and drive a steering engine to realize control closed loop; the control plane instruction can isolate single-machine faults through median voting, meanwhile, each flight control board card monitors the states of other board cards, the fault board card is positioned according to accumulated deviation of the control instruction, and dual-mode redundancy work is started;
the dual-mode redundancy mode is as follows: after the single machine fault is judged, the rudder machine end/interface board determines the current aircraft according to the three-machine fault condition of each flight control instruction and the sequence of the serial numbers of the flight control boards, drives the steering engine according to the control plane instruction of the current aircraft, and does not vote on the control plane instruction any more; determining a secondary fault board card according to the heartbeat and the board card self-checking information, and starting a single mode;
the single mode is that if the on-duty aircraft fails again in the dual-mode redundancy mode, the control surface instruction/interface board card switches the on-duty right to a normal single machine to independently complete flight control;
the median voting is that the flight control system is divided into two voting modes according to the difference of the voting points of the control instruction of the key pneumatic control surface, so as to form redundancy; a) in the steering engine end value voting mode, a control plane instruction is transmitted to the steering engine through a CAN bus, and a pneumatic control plane is driven after the steering engine end carries out medium value voting; b) and in the interface board voting mode, the interface board obtains a control surface instruction through FlexRay, carries out median voting and then sends the control surface instruction to the steering engine through RS422, and when the CAN bus of the steering engine has a fault or loses a signal, the data instruction of the RS422 is used instead.
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| CN110531787A (en) * | 2019-09-18 | 2019-12-03 | 朗星无人机系统有限公司 | A kind of unmanned plane drives into or out of control system automatically |
| CN111338265A (en) * | 2020-03-27 | 2020-06-26 | 浙江华奕航空科技有限公司 | Heterogeneous redundancy unmanned aerial vehicle autopilot |
| CN112346332A (en) * | 2020-11-20 | 2021-02-09 | 中国船舶工业集团公司第七0八研究所 | Fault-tolerant control system of underwater unmanned vehicle |
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| CN113608429A (en) * | 2021-06-16 | 2021-11-05 | 中电科芜湖通用航空产业技术研究院有限公司 | Distributed redundancy unmanned aerial vehicle |
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