CN104656654A - Aircraft posture control system - Google Patents
Aircraft posture control system Download PDFInfo
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- CN104656654A CN104656654A CN201310596232.3A CN201310596232A CN104656654A CN 104656654 A CN104656654 A CN 104656654A CN 201310596232 A CN201310596232 A CN 201310596232A CN 104656654 A CN104656654 A CN 104656654A
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Abstract
The invention discloses an aircraft posture control system, belongs to the field of control systems, in particular relates to improvement of an aircraft posture control system, and provides an aircraft posture control system which is small in error, low in noise and precise in control. The aircraft posture control system comprises a sensor, a memory, a time-delay judgment module, a calibration module, a filter circuit, an oscillator, a trigger, a timer and a power supply. The aircraft posture control system is characterized in that the sensor is connected with a filter, the memory, the time-delay judgment module, the calibration module, the filter circuit, the oscillator and the trigger in sequence; the filter circuit is connected with the oscillator and the trigger; the oscillator is connected with the power supply; the trigger is connected with the timer.
Description
Technical field
the invention belongs to control system field, especially relate to a kind of improvement of aircraft manufacturing technology system.
Background technology
mEMS inertial sensor has the advantage that volume is little, lightweight, power consumption is little, startup is fast, cost is low and reliability is high, so the miniaturization boat appearance measuring system based on MEMS sensor becomes the first-selection of aerial photography unmanned aerial vehicle.Accurately and in real time obtain the skyborne attitude angle of aircraft, be determine control accuracy and system stability key but, in the aerial survey system of low cost MEMS inertial sensor composition, because micro-mechanical gyroscope is by the impact of the factor such as temperature, mechanical vibration.There is larger static drift error, long integration can cause attitude angle to be dispersed, and therefore, is not suitable for the determination of long-time attitude angle.And accelerometer can export the accekeration produced when measuring system is moved, in situation that is static or uniform motion, the acceleration recorded is resolved and can also be used for measurement of dip angle.But because accelerometer is simultaneously to acceleration of gravity and acceleration of motion sensitivity, when measuring system strenuous exercise, larger measuring error can be produced.So, accelerometer be not suitable for short time attitude angle measurement.
Summary of the invention
the present invention is exactly for the problems referred to above, provides that a kind of error is little, noise is little and controls accurate a kind of aircraft manufacturing technology system.
for achieving the above object, the present invention adopts following technical scheme, the present invention includes sensor, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger, timer and power supply, it is characterized in that: sensor is connected successively with wave filter, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger; Filtering circuit is connected with oscillator, trigger, and oscillator is connected with power supply, and trigger is connected with timer.
as a kind of preferred version, processor of the present invention is STM32F103 single-chip microcomputer.
beneficial effect of the present invention.
sensor of the present invention is connected successively with wave filter, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger; Filtering circuit is connected with oscillator, trigger, and oscillator is connected with power supply, and trigger is connected with timer.Adopt kalman filter method, merge from accelerometer and gyrostatic signal, solve noise and attitude optimal estimation problem.Finally complete the aircraft manufacturing technology system with degree of precision and accuracy.
Accompanying drawing explanation
in order to make technical matters solved by the invention, technical scheme and beneficial effect clearly understand, below in conjunction with the drawings and the specific embodiments, the present invention is further elaborated.Should be appreciated that embodiment described herein only in order to explain the present invention, be not intended to limit the present invention.
fig. 1 is schematic block circuit diagram of the present invention.
Embodiment
as shown in the figure, the present invention includes sensor, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger, timer and power supply, it is characterized in that: sensor is connected successively with wave filter, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger; Filtering circuit is connected with oscillator, trigger, and oscillator is connected with power supply, and trigger is connected with timer.
as a kind of preferred version, processor of the present invention is STM32F103 single-chip microcomputer.
the present invention for core with 32 single-chip microcomputer STM32F103, adopts gyroscope (ENC-03) and acceleration transducer (MMA7260) to feed back as attitude angle, and the exact posture angle recorded is sent to host computer by RS232 universal serial bus shows aobvious.Whole system is made up of parts such as acceleration transducer, gyro sensor, filtering circuit, power circuit, MCU.
for axle acceleration sensor, when its sensing direction is consistent with acceleration of gravity direction, if be now zero angle of inclination, if acceleration transducer measurement result is F (θ), θ is angle of inclination, g is acceleration of gravity.
kalman filter is an optimization autoregression data processing algorithm.For the very most problem of solution, it is optimum, most effective or even the most useful.Its widespread use, more than 30 years, comprises robot navigation, controls, and Data Fusion of Sensor is even in the radar system of military aspect and guided missile tracking etc.
the present invention is directed to sensing data in multi-rotor aerocraft appearance measurement and control system and there is noise and measuring error, so that be used alone the problem that gyroscope and accelerometer can not obtain optimum attitude angle, set up the mathematical model of gyroscope and accelerometer error, adopt kalman filter method, realize data fusion, effectively improve accuracy of detection and the control accuracy of attitude detection system. the method is successfully applied in quadrotor attitude detection and control system, demonstrate the noise inhibiting ability that it is good, improve the adaptability of aircraft to environmental change.
be understandable that, above about specific descriptions of the present invention, the technical scheme described by the embodiment of the present invention is only not limited to for illustration of the present invention, those of ordinary skill in the art is to be understood that, still can modify to the present invention or equivalent replacement, to reach identical technique effect; Needs are used, all within protection scope of the present invention as long as meet.
Claims (2)
1. aircraft manufacturing technology system, comprise sensor, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger, timer and power supply, it is characterized in that: sensor is connected successively with wave filter, storer, time delay determination module, calibration module, filtering circuit, oscillator, trigger; Filtering circuit is connected with oscillator, trigger, and oscillator is connected with power supply, and trigger is connected with timer.
2. aircraft manufacturing technology system according to claim 1, is characterized in that: described processor is STM32F103 single-chip microcomputer.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201310596232.3A CN104656654A (en) | 2013-11-25 | 2013-11-25 | Aircraft posture control system |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201310596232.3A CN104656654A (en) | 2013-11-25 | 2013-11-25 | Aircraft posture control system |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN104656654A true CN104656654A (en) | 2015-05-27 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201310596232.3A Pending CN104656654A (en) | 2013-11-25 | 2013-11-25 | Aircraft posture control system |
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| CN (1) | CN104656654A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106444807A (en) * | 2016-09-29 | 2017-02-22 | 湖北航天技术研究院总体设计所 | Compound attitude control method of grid rudder and lateral jet |
| CN117724540A (en) * | 2024-02-18 | 2024-03-19 | 成都航空职业技术学院 | Automatic control method for aircraft motor |
-
2013
- 2013-11-25 CN CN201310596232.3A patent/CN104656654A/en active Pending
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106444807A (en) * | 2016-09-29 | 2017-02-22 | 湖北航天技术研究院总体设计所 | Compound attitude control method of grid rudder and lateral jet |
| CN106444807B (en) * | 2016-09-29 | 2019-04-12 | 湖北航天技术研究院总体设计所 | A kind of compound attitude control method of grid rudder and Lateral jet |
| CN117724540A (en) * | 2024-02-18 | 2024-03-19 | 成都航空职业技术学院 | Automatic control method for aircraft motor |
| CN117724540B (en) * | 2024-02-18 | 2024-04-19 | 成都航空职业技术学院 | Automatic control method for aircraft motor |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20150527 |