CN216424729U - Flying tube system of transverse double-tilting rotor aircraft - Google Patents
Flying tube system of transverse double-tilting rotor aircraft Download PDFInfo
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- CN216424729U CN216424729U CN202123075470.9U CN202123075470U CN216424729U CN 216424729 U CN216424729 U CN 216424729U CN 202123075470 U CN202123075470 U CN 202123075470U CN 216424729 U CN216424729 U CN 216424729U
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Abstract
The utility model relates to a transverse type double-tilting rotor aircraft flies the pipe system, include: the aircraft comprises a flight control system arranged on an aircraft body, and a tilting control system and an information acquisition system which are connected with the flight control system through a bus; the flight control system comprises an FPGA chip, and a flight control machine-mounted computer, a sensor group, a digital frequency hopping radio station, a helicopter steering engine and a fixed wing steering engine which are respectively connected with the FPGA chip; the flight control system is connected with the ground station through a digital frequency hopping radio station; the tilting control system comprises a tilting airborne computer, a modulation and demodulation chip and a tilting motor, wherein the modulation and demodulation chip and the tilting motor are respectively connected with the tilting airborne computer, and the modulation and demodulation chip is further connected with a tilting angle sensor. All information acquisition and sending tasks in the flight control system are realized by the FPGA chip and peripheral circuits, and the service efficiency of the airborne computer is improved. Among the control system verts, through using rotary transformer to measure rotor nacelle angle of verting, have the installation simple, the precision is high, respond fast characteristics.
Description
Technical Field
The utility model relates to an aircraft control technical field particularly, relates to a transverse type two rotor craft that vert fly pipe system.
Background
The vertical take-off and landing fixed wing aircraft is an aircraft combining the characteristics that a helicopter can take off and land vertically and a fixed wing aircraft can cruise at high speed. The existing vertical take-off and landing fixed wing aircraft mainly comprises four-rotor composite wings, four-rotor tilt rotors, transverse double-tilt rotors and the like. The four-rotor composite wing is a mainstream vertical take-off and landing fixed wing aircraft due to simple structure and control. And the rotor that verts is two to the traverse mode, has lacked two rotors and rotary mechanism for four rotors, has lacked the consumption of power of two rotors, and holistic consumption reduces, and duration is stronger, and the direct benefit of bringing is that its carrying capacity has obvious promotion in comparison with traditional four rotors under the same volume, and the load-carrying capacity is also bigger.
However, the structure and composition of the tandem dual tiltrotor system is complex. The system of the flying pipe needs to control a helicopter state steering engine and a fixed wing state steering engine simultaneously, also needs to monitor and control the tilting angle of the rotor wing nacelle, and also needs to control the rotating speed of an engine in the states of the helicopter and the fixed wing. Therefore, an improved circuit structure is required to solve this problem.
SUMMERY OF THE UTILITY MODEL
The utility model provides a not enough to prior art, the utility model provides a two rotor aircraft flight system circuit structure that verts of horizontal formula can control helicopter state machine and fixed wing state steering wheel, can carry out data acquisition and control to the angle of verting of rotor nacelle, controls according to demand engine speed simultaneously, through improving circuit structure, and each subsystem circuit separately installs near being controlled the object, prevents the interference that the signal line overlength arouses, has reduced the complexity of single system.
The utility model discloses a realize that the technical scheme that above-mentioned purpose adopted is:
a flying tube system for a tandem double tiltrotor aircraft, comprising: the aircraft comprises a flight control system arranged on an aircraft body, and a tilting control system and an information acquisition system which are connected with the flight control system through a bus; the flight control system comprises an FPGA chip, and a flight control machine-mounted computer, a sensor group, a digital frequency hopping radio station, a helicopter steering engine and a fixed wing steering engine which are respectively connected with the FPGA chip; the flight control system is connected with the ground station through a digital frequency hopping radio station; the tilting control system comprises a tilting airborne computer, a modulation and demodulation chip and a tilting motor which are respectively connected with the tilting airborne computer, and the modulation and demodulation chip is also connected with a tilting angle sensor; the information acquisition system comprises an information acquisition onboard computer and an engine ECU (electronic control Unit) control unit, a nacelle gear box lubricating oil temperature sensor, a nacelle gear box lubricating oil pressure sensor and a throttle valve steering engine which are respectively connected with the information acquisition onboard computer.
The flight control system, the tilting control system and the information acquisition system are communicated through a 485 bus.
The sensor group comprises a differential pressure sensor and an airspeed tube module, an inertia measuring unit, an air pressure sensor, a satellite positioning module and a magnetic compass.
The FPGA chip is connected with the air pressure sensor, the satellite positioning module, the magnetic compass and the digital frequency hopping radio station through an RS232 interface; the pressure difference sensor and the airspeed tube module, the inertia measuring unit, the air pressure sensor and the flight control machine-mounted computer are connected through an SPI bus, and a helicopter steering engine and a fixed wing steering engine are connected through PWM ports.
The helicopter steering engine is connected with a helicopter steering engine tilting disk, the fixed wing steering engine is connected with a fixed wing control surface, and a PWM (pulse-width modulation) port outputs a PWM signal to control the helicopter steering engine tilting disk and the fixed wing control surface to rotate.
The tilting angle sensor is a rotary transformer and is installed on a rotating shaft of the nacelle.
And the lubricating oil temperature sensor and the lubricating oil pressure sensor are arranged in the nacelle gear box.
The tilting airborne computer is connected with a tilting motor through an RS232 interface, and is connected with a modulation and demodulation chip through an SPI bus, and the modulation and demodulation chip is connected with a tilting angle sensor through a signal line.
The information acquisition tilting computer is connected with an engine ECU (electronic control Unit) through an RS232 interface, is connected with a gear box lubricating oil temperature sensor and a nacelle gear box lubricating oil pressure sensor through an RS232 interface, and is connected with a throttle valve steering engine through a PWM (pulse-width modulation) port.
The utility model has the following beneficial effects and advantages:
1. the flight control system, the tilting control system and the information acquisition system are used for respectively improving the control circuit for the flight of the aircraft, the tilting of the rotor wing nacelle and the rotating speed of the engine, so that the complexity of a single system is reduced. Meanwhile, each system circuit is separately installed near the controlled object, and interference caused by overlong signal lines is prevented.
2. All information acquisition and sending tasks in the flight control system are realized by the FPGA chip and peripheral circuits, so that the use efficiency of the onboard computer of the flight control system is improved.
3. The tilting control system measures the tilting angle of the rotor nacelle by using the rotary transformer and has the characteristics of simple installation, high precision and quick response.
Drawings
FIG. 1 is a system block diagram of the present invention;
fig. 2 is a structural diagram of the flight control system of the present invention;
fig. 3 is a structural diagram of a tilting control system of the present invention;
FIG. 4 is a structural diagram of the information acquisition system of the present invention;
figure 5 is a schematic view of a transverse double tiltrotor aircraft.
Detailed Description
In order to make the above objects, features and advantages of the present invention more comprehensible, embodiments of the present invention are described in detail below with reference to the accompanying drawings. In the following description, numerous specific details are set forth in order to provide a thorough understanding of the present invention. The present invention can be embodied in many other forms than those specifically described herein, and it will be apparent to those skilled in the art that similar modifications can be made without departing from the spirit and scope of the invention.
Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. The terminology used herein in the description of the invention is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention.
As shown in fig. 1, the flight tube system of the transverse double-tilt rotor aircraft comprises a flight control system, a tilt control system and an information acquisition system; the flight control system is communicated with the tilting control system through an RS485 bus, the flight control system sends the expected tilting angle of the unmanned aerial vehicle nacelle to the tilting control system, and the tilting control system sends the actual tilting angle of the nacelle to the flight control system. The flight control system is communicated with the information acquisition system through an RS485 bus, the flight control system sends the expected rotating speed of the engine to the information acquisition system, and the information acquisition system sends acquired ECU data, gear box temperature and pressure data to the flight control system.
As shown in fig. 2, the flight control system includes an onboard computer, a Field Programmable Gate Array (FPGA) chip, a sensor, a digital frequency hopping radio, a helicopter steering engine, and a fixed wing steering engine. The airborne computer is STM32H743, the FPGA chip is EP4CE10E22, and the sensor is NEO-M8N satellite positioning module, HMR2300 magnetic compass, DO4525 differential pressure sensor + airspeed tube module, ADIS16448 inertial measurement unit, baroceptor. The helicopter steering engine is SF40035MG, and the fixed wing steering engine is DA 26-30.
As shown in fig. 2, the FPAG chip reads airspeed data of the DO4525 differential pressure sensor + airspeed tube module through the SPI bus, three-axis acceleration and angular velocity data of the inertial measurement unit, and air pressure height data of the air pressure sensor. And reading the position and speed data of the satellite positioning module through RS232, reading the heading data of the magnetic compass, and reading a ground personnel control instruction received by the frequency hopping radio station. And ECU data, gear box temperature and pressure data acquired by the information acquisition system are read through an RS485 bus, and the actual tilting angle of the nacelle acquired by the tilting control system is read. And all data is packaged and sent to the on-board computer over the SPI bus at a frequency of 50 HZ. The onboard computer analyzes the received data and commands, obtains the control quantity of the helicopter steering engine and the fixed wing steering engine, the expected tilting angle of the nacelle and the expected speed of the engine through the prior art, and sends the control quantity, the expected tilting angle of the nacelle and the expected speed of the engine to the FPGA chip through the SPI bus at the frequency of 50 HZ. The FPGA chip converts the control quantity into PWM waves, the PWM waves are used for controlling a helicopter steering engine and a fixed wing steering engine, finally a helicopter swashplate and a fixed wing control surface are controlled, and meanwhile, the expected tilting angle of the nacelle and the expected rotating speed of an engine are respectively sent to a tilting control system and an information acquisition system through a 485 bus.
As shown in fig. 3, the tilting control system includes an on-board computer, a tilting angle sensor, a modem chip, and a tilting motor controller. The airborne computer is STM32F103, and the angle sensor that verts is the resolver of TS2620N271 model, and the modem chip is AD2S 1200. The modulation and demodulation chip is connected with the rotary transformer through a signal wire to convert the cos/sin signal into an angle signal. The airborne computer is communicated with the modulation and demodulation chip through the SPI bus, and the tilting angle sensor data are collected to obtain the tilting angle of the nacelle. And the airborne computer is communicated with the flight control system through an RS485 bus, receives the expected tilting angle of the nacelle and sends the actual tilting angle of the nacelle. The tilting airborne computer calculates the rotation speed of the tilting motor through a classical PID control algorithm according to the expected tilting angle of the nacelle and the actual tilting angle of the nacelle, and controls the rotation of the tilting motor through RS 232.
As shown in FIG. 4, the information acquisition system comprises an onboard computer, a nacelle gearbox lubricating oil temperature sensor, a nacelle gearbox lubricating oil pressure sensor and a throttle steering engine. The onboard computer is STM32F103, and the onboard computer acquires data of a nacelle gear box lubricating oil temperature sensor and a nacelle gear box lubricating oil pressure sensor through an RS232 interface. And the onboard computer acquires monitoring data of an ECU (electronic control Unit) of the engine through an RS232 interface. And the onboard computer sends the data to the flight control system through an RS485 bus and receives an expected rotating speed instruction of the flight control system. And the information acquisition onboard computer obtains the control quantity of the throttle valve steering engine through a classical PID control algorithm according to the acquired rotating speed data and the expected rotating speed of the engine, converts the control quantity into PWM information to control the throttle valve steering engine, and keeps the rotating speed of the engine constant at a set value.
As shown in fig. 5, the schematic diagram of the transverse double-tilt rotor aircraft is shown, where 1 is a control surface of a fixed wing, 2 is a tilt disk, 3 is a tilt nacelle, and 4 is a tilt motor.
The utility model discloses a go up the preferred embodiment of the utility model, it should be pointed out, the FPGA chip that the utility model relates to, computer etc. are conventional selection, only protect structure technical characteristics such as hardware connection relation and positional relation, and the technical staff in the field passes through the utility model discloses the structural characteristics who records combine conventional programming logic to realize the utility model discloses the function solves the utility model discloses technical problem. For those skilled in the art, without departing from the principle of the present invention, several improvements and modifications can be made, and these improvements and modifications should be considered as the protection scope of the present invention.
Claims (9)
1. A flying-tube system for a tandem double tiltrotor aircraft, comprising: the aircraft comprises a flight control system arranged on an aircraft body, and a tilting control system and an information acquisition system which are connected with the flight control system through a bus; the flight control system comprises an FPGA chip, and a flight control machine-mounted computer, a sensor group, a digital frequency hopping radio station, a helicopter steering engine and a fixed wing steering engine which are respectively connected with the FPGA chip; the flight control system is connected with the ground station through a digital frequency hopping radio station; the tilting control system comprises a tilting airborne computer, a modulation and demodulation chip and a tilting motor which are respectively connected with the tilting airborne computer, and the modulation and demodulation chip is also connected with a tilting angle sensor; the information acquisition system comprises an information acquisition onboard computer and an engine ECU (electronic control Unit) control unit, a nacelle gear box lubricating oil temperature sensor, a nacelle gear box lubricating oil pressure sensor and a throttle valve steering engine which are respectively connected with the information acquisition onboard computer.
2. The flying tube system of a tandem double tilt rotor aircraft according to claim 1, wherein the flight control system, the tilt control system, and the information acquisition system communicate with each other via a 485 bus.
3. The flying tube system of a tandem double tiltrotor aircraft according to claim 1, wherein the sensor group comprises a differential pressure sensor + pitot tube module, an inertial measurement unit, an air pressure sensor, a satellite positioning module, and a magnetic compass.
4. The flying tube system of a tandem double tilt rotor aircraft according to claim 3, wherein the FPGA chip is connected to a barometric sensor, a satellite positioning module, a magnetic compass, and a digital frequency hopping radio station through RS232 interfaces; the pressure difference sensor and the airspeed tube module, the inertia measuring unit, the air pressure sensor and the flight control machine-mounted computer are connected through an SPI bus, and a helicopter steering engine and a fixed wing steering engine are connected through PWM ports.
5. The flying pipe system of the transverse double-tilt rotor aircraft according to claim 4, wherein the helicopter steering engine is connected with a helicopter steering engine swashplate, the fixed wing steering engine is connected with a fixed wing control surface, and the PWM port outputs PWM signals to control the helicopter steering engine swashplate and the fixed wing control surface to rotate.
6. The system of claim 1, wherein said tilt angle sensor is a resolver mounted on a nacelle rotating shaft.
7. The aircraft flight tube system of claim 1, wherein the oil temperature sensor and the oil pressure sensor are disposed within a nacelle gearbox.
8. The flying tube system of a tandem double tilt rotor aircraft according to claim 1, wherein said tilt onboard computer is connected to said tilt motor via an RS232 interface and to said modem chip via an SPI bus, said modem chip being connected to said tilt angle sensor via a signal line.
9. The flying pipe system of a transverse double-tilt rotor aircraft according to claim 1, wherein the information acquisition tilt computer is connected to the engine ECU control unit through an RS232 interface, connected to the gearbox lubricating oil temperature sensor and the nacelle gearbox lubricating oil pressure sensor through an RS232 interface, and connected to the throttle steering engine through a PWM port.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202123075470.9U CN216424729U (en) | 2021-12-08 | 2021-12-08 | Flying tube system of transverse double-tilting rotor aircraft |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202123075470.9U CN216424729U (en) | 2021-12-08 | 2021-12-08 | Flying tube system of transverse double-tilting rotor aircraft |
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| CN216424729U true CN216424729U (en) | 2022-05-03 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116700318A (en) * | 2023-08-03 | 2023-09-05 | 中国民用航空飞行学院 | Flight attitude measurement and control system and method based on FPGA collaboration |
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2021
- 2021-12-08 CN CN202123075470.9U patent/CN216424729U/en active Active
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116700318A (en) * | 2023-08-03 | 2023-09-05 | 中国民用航空飞行学院 | Flight attitude measurement and control system and method based on FPGA collaboration |
| CN116700318B (en) * | 2023-08-03 | 2023-10-20 | 中国民用航空飞行学院 | Flight attitude measurement and control system and method based on FPGA collaboration |
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