Monitoring System and Method for the Motion Status of Low-Altitude Near-Earth Vehicles

Abstract

The present invention belongs to the field of avionics and aviation instrument technology, and relates to a monitoring system and method for the motion state of a near-ground low-altitude aircraft. The present invention includes a differential pressure sensing system, a temperature sensor, a pressure sensor, and a single-chip microcomputer. The sensing ends of the pressure sensor and the temperature sensor are both arranged in the skin of the aircraft, and the signal output ends of the pressure sensor and the temperature sensor are connected to the single-chip microcomputer; the single-chip microcomputer is connected to the flight control system of the aircraft through a communication line and a protocol, and sends information to the ground command and monitoring center and receives monitoring; it is helpful for accurate measurement of the altitude, multi-directional air flow velocity, lifting speed, etc. of the aircraft. Relying on the characteristics of low-altitude atmospheric parameters, the aircraft's dependence on GPS and differential stations can be greatly reduced, the difficulty of low-altitude control by personnel can be reduced, and it is conducive to improving the intelligence level of low-altitude operations and automatic and safe landing of aircraft, especially unmanned aerial vehicles.

Description

Technical Field

[0001] The invention belongs to the technical field of avionics and aviation instruments, and relates to a monitoring system and method for the motion state of a near-ground low-altitude aircraft. Background Art

[0002] More than 100 years ago, the Wright brothers invented the fixed-wing aircraft and various aircraft were subsequently developed based on the basic principles of various fluid mechanics laws discovered 300 years ago, including the Bernoulli equation. At the same time, people also found that the atmospheric density in the troposphere (about 10km above sea level), especially in low altitudes (below 2km), basically changes linearly, so diaphragm instruments were used in aircraft and are still in use today.

[0003] On December 17, 1903, the Wright brothers of the United States completed the first human flight. It has been nearly 120 years since then. The pneumatic mechanical diaphragm instrument has been continuously improved and used with the development of aircraft. In order to make the indication data of the altimeter, airspeed indicator and vertical speed indicator on the aircraft close to the actual motion posture of the aircraft, a set of differential pressure pitot tubes (pitot tubes) with complex mechanical structures and pressure-taking pipelines are designed. In order to prevent icing in the air, an electric heating device is also designed in the differential pressure pitot tube. The dynamic pressure taking port faces the airflow direction to withstand the relative airflow impact pressure (dynamic pressure). The airspeed meter diaphragm is connected through the dynamic pressure pipeline to drive the transmission mechanism to make the pointer indicate on the instrument panel. Several static pressure taking ports are opened around the differential pressure pitot tube to ensure that it is not affected by the moving airflow, so that the air pressure in the altimeter diaphragm is consistent with the field pressure at the height of the aircraft. The air pressure in the altimeter diaphragm changes with the external air pressure to agitate the diaphragm and drive the altimeter pointer to indicate the altitude. The diaphragm of the vertical speed indicator is connected to the static pressure and is connected to the sealed instrument housing. The speed of the pressure difference change is used to indicate the vertical speed. The above-mentioned diaphragm instrument system is affected by many factors and has a large indication error. For this reason, a radio altimeter is installed to judge the relative altitude by the length of time for radio wave transmission and reception. However, due to technical equipment limitations and environmental influences, the airspeed indicator, vertical speed indicator and two altimeters cannot objectively reflect the flight attitude of the aircraft. During low-altitude flight, especially during landing, the flight manual requires that the altitude, airspeed and descent speed should be visually measured without looking at the instrument at an altitude below 300m during the landing process. It mainly depends on the visual judgment of the flight personnel, which is difficult to ensure accuracy and affects safety. Moreover, once the pitot tube freezes or the pitot tube heating device fails, the flight safety will be endangered if it is not repaired in time.

[0004] With the advancement of semiconductor technology and numerical control technology, thyristor sensors have begun to replace diaphragm sensors on aircraft. CNC technology is also used in the flight control systems of new aircraft models, but its pressure-taking structure has remained basically unchanged, and has not produced a qualitative improvement compared to diaphragm instruments.

[0005] With the popularization of satellite navigation technology, unmanned and manned aircraft are equipped with GPS or Beidou navigation systems. With the support of ground satellite navigation three-dimensional reference differential stations, the height difference can be within 1 meter, and the aircraft plane displacement accuracy can be within 5 cm. However, in areas without ground differential stations, the position height error still cannot meet the needs of low-altitude flight of aircraft, especially the low-altitude operation of unmanned aircraft (such as agricultural and forestry plant protection) still requires visual inspection and manual intervention, which cannot meet the needs of automatic driving. Over-reliance on Starlink transmission is even more dangerous for satellite signal interference and interruption. In summary, aircraft (manned or unmanned) now urgently need to overcome the disadvantages of relying mainly on visual inspection and signal transmission during low-altitude operations and landings.

[0006] In short, the reasons why the aircraft's own perception ability is not strong are: first, due to the complex and changeable low-altitude near-ground electromagnetic field environment, there is a technical bottleneck in improving the sensitivity of its own wireless signal transmission sensing equipment. Second, the low-altitude atmospheric parameters are relatively stable, but the selection of sensors, sensing structure layout and algorithms, etc., follow the membrane box instrument architecture thinking, which limits the application and development of new technologies. Summary of the invention

[0007] The present invention provides a method for monitoring the motion state of a near-ground low-altitude aircraft, which is helpful for accurately measuring the altitude, multi-directional airspeed, and ascent and descent speed of the aircraft. Relying on the characteristics of low-altitude atmospheric parameters, the aircraft's dependence on GPS and differential stations can be greatly reduced, and the difficulty of low-altitude control by personnel can be reduced, which is conducive to improving the intelligence level of low-altitude operations and automatic safe landing of aircraft, especially unmanned aircraft.

[0008] The present invention solves the technical problem through the following technical solutions:

[0009] A monitoring system for the motion state of a near-ground low-altitude aircraft, the monitoring system comprising a differential pressure sensor group, a temperature sensor, a pressure sensor, a monitoring system single-chip computer U1, a flight control system single-chip computer FK, a wireless communication module FK and a ground monitoring station; the sensing ends of the pressure sensor and the temperature sensor are both arranged in the skin of the aircraft, and the signal output ends of the differential pressure sensor group, the pressure sensor and the temperature sensor are connected to the monitoring system single-chip computer U1; the monitoring system single-chip computer U1 is connected to the flight control system single-chip computer FK of the aircraft through a communication line and a protocol to control the flight of the aircraft, and the monitoring system single-chip computer U1 also sends information to a computer of the ground monitoring station in a wireless manner and receives monitoring;

[0010] The differential pressure sensor group is provided with two groups with exactly the same structure, which respectively detect the longitudinal and lateral air pressure differences of the aircraft. Each group of differential pressure sensor groups is provided with a differential pressure pitot tube and a differential pressure sensor. The two differential pressure pitot tubes of the two differential pressure sensor groups are arranged orthogonally, and the axial direction of one differential pressure pitot tube is arranged along the longitudinal direction of the aircraft and is consistent with the direction indicated by the longitude and latitude directions of the gyroscope on the aircraft; the differential pressure pitot tube includes a tube body, a throttle plate, a front pressure-taking tube and a rear pressure-taking tube. The tube body is a straight circular tube with openings at both ends, and the throttle plate is vertically arranged in the middle of the tube body; a throttle hole is opened in the center of the throttle plate, and the output ends of the front pressure-taking tube and the rear pressure-taking tube are connected to the differential pressure sensor, and the differential pressure sensor measures the pressure difference on both sides of the throttle plate; the output signal of the differential pressure sensor is connected to the input end of the single-chip microcomputer, and the output end of the single-chip microcomputer is connected to the flight control system.

[0011] In the above-mentioned monitoring system for the motion status of near-ground low-altitude aircraft, the outer periphery of the throttle plate is fixedly connected to the inner wall of the tube body, the throttle plate divides the interior of the tube body into a front pressure-taking section and a rear pressure-taking section, the front pressure-taking tube and the rear pressure-taking tube are respectively located on both sides of the throttle plate and close to the throttle plate, the input end of the front pressure-taking tube is vertically connected to the front pressure-taking section, and the input end of the rear pressure-taking tube is vertically connected to the rear pressure-taking section.

[0012] In the above-mentioned monitoring system for the motion status of low-altitude aircraft near the ground, both sides of the throttle hole are chamfered to form a knife-like edge portion, and both sides of the hole edge are smoothly connected with the inner wall of the tube body.

[0013] In the above-mentioned monitoring system for the motion state of a near-ground low-altitude aircraft, the front pressure-taking section and the rear pressure-taking section have the same length, the length of the front pressure-taking section is D, the inner diameter of the tube body is R, and D>10R.

[0014] A method for monitoring the motion state of a near-ground low-altitude aircraft, wherein the aircraft is provided with a monitoring system as described in any one of claims 1 to 3, wherein air enters a tube body through a front pressure-taking section, flows out from a rear pressure-taking section after passing through a throttle hole, and a pressure difference is formed on both sides of the throttle hole, and a differential pressure sensor receives a pressure difference signal measured by the front pressure-taking tube and the rear pressure-taking tube;

[0015] The differential pressure sensor transmits the differential pressure signal to the monitoring system single chip U1 in real time. At the same time, the temperature sensor measures the atmospheric temperature signal of the current position of the aircraft, and the pressure sensor measures the atmospheric pressure signal of the current position of the aircraft. The single chip collects differential pressure parameters, static pressure parameters and temperature parameters. The temperature parameters correct the air density and process it through the mathematical model stored in the single chip to obtain the real-time air flow rate, which is the real-time flight speed of the aircraft.

[0016] The air velocity in the differential pressure pitot tube is processed according to the following function:

[0017] A. Functional relationship between pressure difference and volume flow rate in differential pressure pitot tube:

[0018]

[0019] The parameters in the formula are as follows: F: volume flow rate; K: correction coefficient, the value of K is 1.0-1.5; C: outflow coefficient of the medium; β: ratio of the diameter of the throttle hole to the diameter of the tube body; d: diameter of the throttle hole; ε: expansion coefficient of the medium; △P: pressure difference before and after the throttle plate; ρ: medium density;

[0020] B. The volume flow rate F is calibrated to the aircraft's airspeed.

[0021] The above-mentioned method for monitoring the motion state of near-ground low-altitude aircraft calculates the relative altitude of the aircraft based on the pressure and temperature parameters measured by the monitoring system, in accordance with the air density table and the ideal gas equation.

[0022] The above-mentioned method for monitoring the motion state of a near-ground low-altitude aircraft calculates the ascent and descent speed of the aircraft: the ascent and descent speed of the aircraft is calculated by the change in the altitude of the aircraft per unit time.

[0023] The above-mentioned monitoring system for the motion state of a near-ground low-altitude aircraft includes a pressure difference sensor CG1 and a pressure difference sensor CG2, a temperature sensor CG3, a pressure sensor CG4, a monitoring system single-chip computer U1, a flight control system single-chip computer FK, a wireless communication module WX and a ground monitoring computer. The parameters of the pressure difference sensor CG1 and the pressure difference sensor CG2 are the same. The output signals of the pressure difference sensor CG1, the pressure sensor CG2, the temperature sensor CG3 and the pressure sensor CG4 are respectively connected to the P0.1 port to the P0.4 port of the single-chip computer in sequence; the processing software of the above-mentioned signals is pre-stored in the single-chip computer, and the processing results of each signal are connected and transmitted to the flight control system single-chip computer FK through the output port P1 of the single-chip computer U1, and the flight control system single-chip computer controls the flight state of the aircraft. The flight control system single-chip computer is connected to the wireless communication module WX, and the wireless communication module WX communicates with the monitoring computer of the ground monitoring station through the wireless transmission system. In addition to completing the monitoring task, the monitoring computer of the ground monitoring station can also directly send instructions to control the flight state of the aircraft.

[0024] The single-chip microcomputer is pre-installed with real-time database software, which can store the records of the aircraft's three-dimensional spatial motion trajectory, attitude parameters, and real-time temperature and air pressure on the horizontal, longitudinal and vertical axes. At the same time, it can be connected to the flight control system through digital communication lines and protocols, and displayed on the display of the flight control system and the ground monitoring station as needed.

[0025] Compared with the prior art, the differential pressure airspeed tube adopts a straight tube structure, which has a simple structure and allows both positive and negative airflows to flow through the throttle hole. Compared with the traditional method of measuring airspeed with a pressure sensor, this method can get rid of the range limitation and improve accuracy.

[0026] The present invention utilizes reliable low-altitude atmospheric characteristics to iteratively improve the aircraft's own attitude perception capability, which can greatly reduce the aircraft's dependence on external signals and reduce the difficulty of low-altitude control by personnel. It is beneficial to improving the intelligence level of unmanned aerial vehicles' low-altitude operations and automatic and safe landings, and also improves the accuracy of control.

[0027] This method can reduce the error of the existing aircraft altimeter and air velocity meter from several meters to tens of meters to centimeter level, making the three main parameter indicators of aircraft altitude, air velocity and ascent and descent speed closer to the actual attitude of the aircraft, which is conducive to real-time control of aircraft attitude and reduces the difficulty of control. It can also meet the requirements of forward and reverse bidirectional air velocity measurement and improve the control performance of unmanned aerial vehicles. The main parameters such as temperature, air pressure, air velocity, altitude and ascent and descent speed are more accurate.

[0028] The application of modern technologies such as sensor selection, combined use methods, single-chip microcomputers and algorithms, digital communications, etc. in this technology will greatly improve the accuracy and reliability of flight attitude parameters, and is an innovative achievement in the digitalization and intelligence of avionics. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] Figure 1 Schematic diagram of the structure of the differential pressure pitot tube of the present invention;

[0030] Figure 2 It is a top view of the structure of the present invention;

[0031] Figure 3 It is the electrical schematic diagram of the present invention;

[0032] The marks in the attached drawings indicate: 1. tube body, 2. throttle plate, 3. front pressure taking tube, 4. rear pressure taking tube, 5. differential pressure airspeed tube, 6. front pressure taking section, 7. rear pressure taking section, 8. blade, 9. throttle hole, CG3. temperature sensor, CG4. pressure sensor, U1. single chip microcomputer, FK. flight control system, WX. wireless communication module. DETAILED DESCRIPTION

[0033] The specific implementation of the present invention is further described in detail below in conjunction with the accompanying drawings and examples. The following examples are used to illustrate the present invention.

[0034] See attached Figure 1 To Attachment Figure 3The monitoring system of the present invention comprises a differential pressure sensor group, a temperature sensor, a pressure sensor, a monitoring system single chip U1, a flight control system single chip FK, a wireless communication module FK and a ground monitoring station; the sensing ends of the pressure sensor and the temperature sensor are arranged in the skin of the aircraft, and the signal output ends of the differential pressure sensor group, the pressure sensor and the temperature sensor are connected to the monitoring system single chip U1; the monitoring system single chip U1 is connected to the flight control system single chip FK of the aircraft through a communication line and a protocol to control the flight of the aircraft, and the monitoring system single chip U1 also sends information to the computer of the ground monitoring station in a wireless manner and receives monitoring;

[0035] The differential pressure sensor group is provided with two groups with completely identical structures, which respectively detect the longitudinal and lateral air pressure differences of the aircraft. Each group of differential pressure sensor groups is provided with a differential pressure pitot tube 5 and a differential pressure sensor. The two differential pressure pitot tubes 5 of the two differential pressure sensor groups are arranged orthogonally, and the axial direction of one differential pressure pitot tube 5 is arranged along the longitudinal direction of the aircraft and is consistent with the direction indicated by the longitude and latitude of the gyroscope on the aircraft; the differential pressure pitot tube 5 includes a tube body 1, a throttle plate 2, a front pressure-taking tube 3 and a rear pressure-taking tube 4. The tube body 1 is a straight round tube with openings at both ends, and the throttle plate 2 is vertically arranged in the middle of the tube body 1; a throttle hole 9 is opened in the center of the throttle plate 2, and the output ends of the front pressure-taking tube 3 and the rear pressure-taking tube 4 are connected to the differential pressure sensor, and the differential pressure sensor measures the pressure difference on both sides of the throttle plate 2; the output signal of the differential pressure sensor is connected to the input end of the single-chip microcomputer, and the output end of the single-chip microcomputer is connected to the flight control system. The above-mentioned pressure differential method has a simple structure, avoids errors caused by complex structures, and reduces potential faults.

[0036] The outer periphery of the throttle plate 2 is fixedly connected to the inner wall of the tube body 1, and the throttle plate 2 divides the interior of the tube body 1 into a front pressure-taking section 6 and a rear pressure-taking section 7. The front pressure-taking tube 3 and the rear pressure-taking tube 4 are respectively located on both sides of the throttle plate 2 and close to the throttle plate 2. The input end of the front pressure-taking tube 3 is vertically connected to the front pressure-taking section 6, and the input end of the rear pressure-taking tube 4 is vertically connected to the rear pressure-taking section 7.

[0037] Because the flight direction of the UAV operation is 360°, in order to facilitate pressure taking when flying in any direction, the differential pressure pitot tube 5 and the differential pressure sensor are each provided with two, the differential pressure pitot tube 5 is arranged horizontally, the two differential pressure pitot tubes 5 are perpendicular to each other and fixedly connected, and the direction of the differential pressure pitot tube is along the longitudinal and transverse direction of the aircraft, which is consistent with the direction indicated by the longitude and latitude of the gyroscope. The length of the front pressure taking section and the rear pressure taking section is the same, the length of the front pressure taking section is D, the inner diameter of the tube body is R, D>10R. In this way, both ends of the differential pressure pitot tube can enter and exit the air, and the pressure taking direction is not restricted, so as to match the pressure taking when the aircraft flies in different directions. The two differential pressure pitot tubes 5 are respectively connected to the signal input ends of the two differential pressure sensors. This arrangement allows the single-chip microcomputer to collect the pressure difference in two vertical directions, and then the software of the single-chip microcomputer can calculate the flight direction and flight speed through the synthesis and decomposition relationship of the speed.

[0038] In order to prevent water from entering the throttle hole 9 or condensing water vapor to freeze in a low-temperature working environment, chamfers are provided on both sides of the edge of the throttle hole 9, and the edge of the throttle hole 9 forms a knife-like cutting edge 8. The sharp cutting edge 8 can prevent water in the airflow from adhering to the edge of the throttle hole 9, thereby preventing the throttle hole 9 from freezing, thereby avoiding monitoring failures caused by freezing and improving the stability of monitoring.

[0039] In order to ensure the accuracy of pressure measurement, the differential pressure pitot tube is installed on the outer surface of the aircraft, preferably on the top of the aircraft.

[0040] The flight control system transmits information to the ground monitoring station by radio transmission. This configuration avoids dependence on base stations, differential stations and satellite signals. Even in areas without third-party signal coverage, within the operating range, information transmission and control can be performed only through signal transmission between the drone and the ground monitoring station.

[0041] In order to protect the differential pressure sensor from wind and sun, the differential pressure sensor is arranged inside the aircraft. The differential pressure sensor uses a silicon capacitor sensor chip, the pressure sensor uses a silicon resistor sensor chip, and the temperature sensor uses PT100 or PT1000, and is treated with high and low temperature resistance, salt spray, vibration, moisture-proof and the like; the sensor structure packaging uses composite materials and light metal materials with good strength, sealing, corrosion resistance and shock absorption, and small size and light weight after packaging; the pressure channel layout design is to carry out air channel layout design according to the design specifications and fluid design calculation book required to match the differential pressure sensor; the single-chip computer, software and digital communication technology are single-chip computer integrated circuits and algorithms for communication, calculation, storage, etc. designed according to the output signals of different sensors and the display and recording requirements of the aircraft attitude.

[0042] When the above device is used to monitor the motion state of a near-ground low-altitude aircraft, the air flow enters the tube body through the front pressure-taking section, flows out from the rear pressure-taking section after passing through the throttle hole, and a pressure difference is formed on both sides of the throttle hole. The pressure is taken out from the front pressure-taking tube and the rear pressure-taking tube to the differential pressure sensor, and the differential pressure sensor can measure the pressure difference.

[0043] The monitoring system includes differential pressure sensors CG1 and CG2, temperature sensor CG3, pressure sensor CG4, monitoring system single chip microcomputer U1, flight control system single chip microcomputer FK, wireless communication module WX and ground monitoring computer. The parameters of differential pressure sensors CG1 and CG2 are the same. The output signals of differential pressure sensor CG1, differential pressure sensor CG2, temperature sensor CG3 and pressure sensor CG4 are connected to the P0.1 port to the P0.4 port of the single chip microcomputer in sequence. The processing software of the above signals is pre-stored in the single chip microcomputer. The processing results of each signal are connected and transmitted to the flight control system single chip microcomputer FK through the output port P1 of the single chip microcomputer U1. The flight control system single chip microcomputer controls the flight state of the aircraft. The flight control system single chip microcomputer is connected to the wireless communication module WX, and the wireless communication module WX communicates with the monitoring computer of the ground monitoring station through the wireless transmission system. In addition to completing the monitoring task, the monitoring computer of the ground monitoring station can also directly send instructions to control the flight state of the aircraft.

[0044] The differential pressure sensor transmits the differential pressure signal to the monitoring system microcontroller U1 in real time. At the same time, the temperature sensor measures the atmospheric temperature signal of the current position of the aircraft, and the pressure sensor measures the atmospheric pressure signal of the current position of the aircraft. The microcontroller collects differential pressure parameters, static pressure parameters and temperature parameters. The temperature parameters correct the air density and process it through the mathematical model stored in the microcontroller to obtain the air flow. After the air flow is calibrated, it is displayed as the airspeed of the aircraft.

[0045] The air flow velocity in the differential pressure pitot tube 5 is processed according to the following function:

[0046] A. Functional relationship between pressure difference and volume flow rate in differential pressure pitot tube:

[0047] The functional relationship between pressure difference and volume flow rate under the influence of temperature on air density is:

[0048]

[0049] F: volume flow rate, K: correction coefficient, C: outflow coefficient of the medium, β: ratio of the diameter of the throttle hole to the diameter of the tube body, d: diameter of the throttle hole, ε: expansion coefficient of the medium, △P: pressure difference before and after the throttle plate, ρ: medium density, K is 1.0-1.5;

[0050] B. The volume flow rate F is calibrated to the aircraft's airspeed.

[0051] Since the general calculation standard in the field of airspeed monitoring is currently under an ideal state, it is suitable for ideal environments such as a single medium and a constant temperature. However, in the actual working environment and working process, the air components in the airflow are complex and the temperature will change. If it is calculated according to the general calculation standard, errors will occur. The correction coefficient K in the present invention is calculated after multiple experiments and industry standards, and the value range of the correction coefficient K is 1.1-1.5. The air velocity result obtained in this way is more accurate.

[0052] Calculate the aircraft's altitude: Calculate the relative altitude of the aircraft based on the temperature and pressure parameters according to the software's preset air density table and ideal gas equation algorithm model.

[0053] Calculate the aircraft's ascent and descent speed: The aircraft's ascent and descent speed can be calculated by the aircraft's altitude change per unit time.

[0054] According to the compass on the aircraft, the axes of the two differential pressure pitot tubes are set to be north-south and east-west respectively. When the flight direction of the aircraft deviates from the axes of the two differential pressure pitot tubes, the two differential pressure pitot tubes respectively collect the flight speed of the aircraft in two vertical directions. When the flight direction of the aircraft is located between the two differential pressure pitot tubes, the pressure difference of the two differential pressure pitot tubes is calculated respectively to obtain the speed of the horizontal axis and the vertical axis, and then the flight speed of the aircraft when flying in any direction is calculated by the preset algorithm;

[0055] The wireless communication module U4 in the present invention adopts CC1101 PA2, the two differential pressure sensors both adopt Fuji differential pressure sensor model FKCW35V5-**D*Y-BA, and the single-chip microcomputer U1 adopts C8051F340 series chips, including 4 16-bit counters / timers, two full-duplex serial ports UART with extended baud rate configuration and up to 40 I / O pins, which can fully meet the needs of communication and control with other modules; the single-chip microcomputer FK can also adopt C8051F340 or other series chips.

[0056] The single-chip microcomputer is pre-installed with real-time database software, which can record and store the three-dimensional spatial motion trajectory of the aircraft's horizontal, longitudinal and vertical axes, attitude parameters, and real-time temperature and air pressure. At the same time, it can be connected to the flight control system through digital communication lines and protocols, and displayed on the display of the flight control system and ground monitoring station as needed.

[0057] The reasons for the errors of traditional airspeed indicators and altimeters are: First, the sensor uses a pressure sensor to take the dynamic pressure and static pressure respectively, and then calculates the difference. The highest accuracy is 0.01%, that is, the error fluctuation of one atmospheric pressure (i.e.: 101325Pa) is within 20Pa, and the range is above 10Pa. Second, the analog quantity is taken and the analog-to-digital conversion is performed in the flight control system hardware. The error caused by the hardware crosstalk is about 0.3%. The dynamic pressure is proportional to the square of the airspeed. It is greater in low-speed flight (below 100km / h), especially below 28m / s. When the dynamic pressure is about 60KPa, the error exceeds 20%. The altitude is calculated as 10Pa per meter. Under standard atmospheric pressure, the laboratory measured 0.01% error is 10Pa, that is, plus or minus 1m, excluding packaging and transmission. The hardware crosstalk is 0.3%, that is, a standard atmospheric pressure plus or minus error of 300Pa, that is, 30m. Therefore, the low-altitude 300-meter flight manual requires visual flight. The present invention effectively reduces the errors caused by traditional methods and structural problems and improves monitoring accuracy.

Claims

1. A monitoring system for the motion state of a near-ground low-altitude aircraft, characterized in that: The monitoring system includes a differential pressure sensor group, a temperature sensor, a pressure sensor, a monitoring system single chip U1, a flight control system single chip FK, a wireless communication module FK and a ground monitoring station; the sensing ends of the pressure sensor and the temperature sensor are arranged in the skin of the aircraft, and the signal output ends of the differential pressure sensor group, the pressure sensor and the temperature sensor are connected to the monitoring system single chip U1; the monitoring system single chip U1 is connected to the flight control system single chip FK of the aircraft through a communication line and a protocol to control the flight of the aircraft, and the monitoring system single chip U1 also sends information to the computer of the ground monitoring station in a wireless manner and receives monitoring; The differential pressure sensor group is provided with two groups with completely identical structures, respectively detecting the longitudinal and lateral air pressure differences of the aircraft. Each differential pressure sensor group is provided with a differential pressure pitot tube (5) and a differential pressure sensor. The two differential pressure pitot tubes (5) of the two differential pressure sensor groups are arranged orthogonally. The axial direction of one differential pressure pitot tube (5) is arranged along the longitudinal direction of the aircraft and is consistent with the direction indicated by the longitude and latitude directions of the gyroscope on the aircraft. The differential pressure pitot tube (5) comprises a tube body (1), a throttle plate (2), a front pressure taking tube (3) and a rear pressure taking tube (4). The tube body (1) is a straight circular tube with openings at both ends. The throttle plate (2) is vertically arranged in the middle of the tube body (1). A throttle hole (9) is provided in the center of the throttle plate (2). The output ends of the front pressure taking tube (3) and the rear pressure taking tube (4) are connected to the differential pressure sensor, and the differential pressure sensor measures the pressure difference on both sides of the throttle plate (2). The output signal of the differential pressure sensor is connected to the input end of the single-chip microcomputer, and the output end of the single-chip microcomputer is connected to the flight control system.

2. The system for monitoring the motion state of a near-ground low-altitude aircraft as claimed in claim 1, characterized in that: The outer periphery of the throttle plate (2) is fixedly connected to the inner wall of the tube body (1); the throttle plate (2) divides the interior of the tube body (1) into a front pressure-taking section (6) and a rear pressure-taking section (7); the front pressure-taking tube (3) and the rear pressure-taking tube (4) are respectively located on both sides of the throttle plate (2) and close to the throttle plate (2); the input end of the front pressure-taking tube (3) is vertically connected to the front pressure-taking section (6); and the input end of the rear pressure-taking tube (4) is vertically connected to the rear pressure-taking section (7).

3. The system for monitoring the motion state of a near-ground low-altitude aircraft as claimed in claim 2, characterized in that: Both sides of the hole edge of the throttling hole (9) are chamfered so that the hole edge forms a blade-like cutting edge (8), and both sides of the hole edge are smoothly kissed with the inner wall of the tube body (1).

4. The system for monitoring the motion state of a near-ground low-altitude aircraft as claimed in claim 3, characterized in that: The front pressure-taking section (6) and the rear pressure-taking section (7) have the same length, the length of the front pressure-taking section (6) is D, the inner diameter of the tube body is R, and D>10R.

5. A method for monitoring the motion state of a near-ground low-altitude aircraft, characterized in that: The aircraft is provided with a monitoring system as claimed in any one of claims 1 to 3, wherein air enters the tube body (1) through the front pressure-taking section (6), flows out from the rear pressure-taking section after passing through the throttle hole (9), and a pressure difference is formed on both sides of the throttle hole (9), and a differential pressure sensor receives a pressure difference signal measured by the front pressure-taking tube (3) and the rear pressure-taking tube (4); The differential pressure sensor transmits the differential pressure signal to the monitoring system single chip U1 in real time. At the same time, the temperature sensor measures the atmospheric temperature signal of the current position of the aircraft, and the pressure sensor measures the atmospheric pressure signal of the current position of the aircraft. The single chip collects differential pressure parameters, static pressure parameters and temperature parameters. The temperature parameters correct the air density and process it through the mathematical model stored in the single chip to obtain the real-time air flow. The air flow is calibrated and displayed as the airspeed of the aircraft. The air velocity in the differential pressure pitot tube (5) is processed according to the following function: A. Functional relationship between pressure difference and volume flow rate in differential pressure pitot tube: The parameters in the formula are as follows: F: volume flow rate; K: correction coefficient, the value of K is 1.0-1.5; C: outflow coefficient of the medium; β: ratio of the orifice diameter to the tube diameter; d: orifice diameter; ε: expansion coefficient of the medium; △P: pressure difference before and after the throttle; ρ: medium density; B. Volume flow rate is calibrated to the aircraft's airspeed.

6. The method for monitoring the motion state of a near-ground low-altitude aircraft according to claim 5, characterized in that: Based on the pressure and temperature parameters measured by the monitoring system, the relative altitude of the aircraft is calculated according to the air density table and the ideal gas equation.

7. The method for monitoring the motion state of a near-ground low-altitude aircraft according to claim 6, characterized in that: Calculate the aircraft's ascent and descent speed: Calculate the aircraft's ascent and descent speed by the aircraft's altitude change per unit time.

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