CN107014386A - The disturbing acceleration measuring method that a kind of attitude of flight vehicle is resolved - Google Patents
The disturbing acceleration measuring method that a kind of attitude of flight vehicle is resolved Download PDFInfo
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- CN107014386A CN107014386A CN201710412386.0A CN201710412386A CN107014386A CN 107014386 A CN107014386 A CN 107014386A CN 201710412386 A CN201710412386 A CN 201710412386A CN 107014386 A CN107014386 A CN 107014386A
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Abstract
The invention discloses the disturbing acceleration measuring method that a kind of attitude of flight vehicle is resolved, the attitude quaternion of aircraft is obtained, the measurement acceleration under body axis system is obtained；The disturbing acceleration under reference frame is obtained by the measurement acceleration under attitude quaternion and body axis system, and then obtains discrete disturbing acceleration；Discrete disturbing acceleration is carried out to carry out with the measurement acceleration under body axis system subtracting each other under first-order low-pass ripple, reconvert to body axis system after segment processing to obtain the accurate body acceleration for filtering out disturbing acceleration.The present invention is with low cost relative to prior art, it is not necessary to adds GPS sensor and measures；The attitude measurement accuracy of follow-up multi-sensor fusion technology can be lifted.
Description
Technical Field
The invention relates to the technical field of control of aviation aircrafts, in particular to an interference acceleration measuring method for resolving an aircraft attitude.
Background
When the aircraft flies in large-sized motor flight, the accelerometer on the aircraft can measure gravity acceleration and also has interference acceleration. Accurate acceleration data is very necessary for multi-sensor fusion techniques such as complementary filtering or kalman filtering. If the adopted acceleration data contains disturbance acceleration, the accuracy of attitude calculation is influenced, and even flight accidents are caused.
For the measurement of the interference acceleration of the aircraft, a Global Positioning System (GPS) sensor is added, and the interference acceleration is calculated by using data of the GPS sensor so as to be fed back to an attitude measurement system. However, this makes the flight control system complicated and expensive, and is not suitable for use in a low-cost flight control system.
Disclosure of Invention
Aiming at the problems in the prior art, the method for measuring the interference acceleration of the aircraft attitude calculation is provided, the problems that when the aircraft flies in a large-sized motor, the subsequent attitude calculation is wrong and the obtained attitude angle information is not accurate enough due to the fact that the acceleration measured by the aircraft body interferes with the acceleration besides the required gravity acceleration are solved, and meanwhile, a GPS sensor is not used for measuring the interference acceleration, and the method has the advantages of being low in cost, simple in algorithm design and good in effect.
In order to solve the problems in the prior art, the invention adopts the technical scheme that:
an interference acceleration measurement method for aircraft attitude calculation comprises the following steps:
step 1, acquiring attitude angles in an Euler angle form of an aircraft in real time in a set sampling period, wherein the attitude angles comprise a rolling angle, a pitch angle and a yaw angle in the Euler angle form, and converting the attitude angles into attitude quaternions
Step 2, obtaining the measured acceleration a under the coordinate system of the bodymeasureBy attitude quaternionAnd the measured acceleration a under the coordinate system of the machine bodymeasureObtaining the disturbance acceleration a under the reference coordinate systemdrTo obtain discrete disturbance acceleration
Step 3, obtaining the interference acceleration acc through the following formuladr：
Wherein, | | ameasureI and grRespectively representing the measured acceleration a under the coordinate system of the machine bodymeasureAnd the gravity acceleration g under the reference coordinate systemrThe value of the modulus of the (c) component,setting a threshold value;
step 4, adopting a first-order low-pass filter to perform interference acceleration acc obtained in the step 3drAnd performing first-order low-pass filtering to obtain the final interference acceleration acc under the reference coordinate system.
Step 5, converting the final interference acceleration acc under the reference coordinate system obtained in the step 4 into a final interference sum under a body coordinate systemVelocity accfMeasuring acceleration a in a coordinate system of the bodymeasureAnd the final interference acceleration acc under the coordinate system of the machine bodyfAccurate body acceleration a to c for obtaining filtering interference acceleration by subtractingmeasure。
Compared with the prior art, the invention has the following beneficial effects:
1. the cost is low, and a GPS sensor is not required to be added for measurement;
2. the attitude measurement precision of the subsequent multi-sensor fusion technology can be improved.
Drawings
FIG. 1 shows a system block diagram of a method of measuring disturbance acceleration.
Detailed Description
The technical scheme of the invention is further specifically described by the following embodiments and the accompanying drawings.
Referring to fig. 1, a method for measuring disturbance acceleration of aircraft attitude calculation includes the following steps:
step 1, acquiring an attitude angle of an aircraft in real time in a fixed sampling period (the sampling period of the embodiment is 2ms), wherein the attitude angle is in an Euler angle form and comprises a rolling angle, a pitch angle and a yaw angle in the Euler angle form, and converting the attitude angle in the Euler angle form into a quaternion expression mode through the following formula, namely acquiring an attitude quaternion
Where cos represents a cosine operation, sin represents a sine operation, φ represents a roll angle in the form of an attitude Euler angle, θ represents a pitch angle in the form of an attitude Euler angle, ψ represents a yaw angle in the form of an attitude Euler angle,is an attitude quaternion, q0、q1、q2、q3Are respectively attitude quaternionsFour elements of (1).
And 2, operating a disturbance acceleration measurement algorithm. In this embodiment, first, the flight control system of the aircraft acquires acceleration data in the MPU6050 sensor on the aircraft through the IIC bus, and since the flight control system is fixedly connected to the aircraft, the acceleration data in the MPU6050 sensor is the measured acceleration a in the body coordinate systemmeasure. At this time, the input of the disturbance acceleration measurement algorithm is the measured acceleration a in the coordinate system of the bodymeasureThe output of the disturbance acceleration measurement algorithm is disturbance acceleration a under a reference coordinate systemdr。
First, the measured acceleration a in the body coordinate systemmeasureDefined in the form:
wherein, ameasureRepresenting the measured acceleration in the coordinate system of the body,is the quaternion of the attitude,representing the conjugate of the attitude quaternion, grRepresenting the gravitational acceleration of the reference coordinate system, adrTo representThe disturbance acceleration in the reference coordinate system, △ a, represents the measurement noise of the accelerometer.
The organism coordinate system in this application refers to the coordinate system of solid antithetical couplet on the aircraft, and the initial point of organism coordinate system is the barycenter of aircraft, and the directional organism the place ahead of x axle, the directional organism of y axle is right-hand, and the directional organism below of z axle. And with reference to a coordinate system, the origin is the center of mass of the aircraft, the xy plane is tangent to the earth, the x axis points to the north, the z axis points to the gravity direction, and the y axis points to the east along the local latitude line. Between the two coordinate systems, the coordinate system can be converted by the attitude matrix, which is used in this embodimentAndthe form represents a coordinate transformation matrix.
Then, the above formula is transformed to obtain the disturbance acceleration a under the reference coordinate systemdrThe solving formula of (2):
obtaining discrete disturbance acceleration according to the above
Wherein,the acceleration of the discrete disturbance is represented,is shown onAn attitude quaternion obtained at an attitude angle sampling instant,is composed ofConjugation of (1).
M in FIG. 1q、NqRespectively representing the above formulaeAndthe form, namely, the transformation matrix representing the reference coordinate system to the body coordinate system, and the transformation matrix representing the body coordinate system to the reference coordinate system.
Step 3, preferably, the obtained discrete disturbance acceleration is subjected toBy segmentation, i.e. according to discrete disturbance accelerationObtaining the disturbance acceleration accdr. In the static state of the aircraft, the aircraft can be considered to have no disturbance acceleration item, so that whether the aircraft is static or not is judged, and whether the disturbance acceleration is solved or not is determined. Wherein, | | ameasureI and grRespectively representing the measured acceleration a under the coordinate system of the machine bodymeasureAnd the gravity acceleration g under the reference coordinate systemrThe value of the modulus of the (c) component,indicating a set threshold value related to accelerometer performance.
Step 4, preferably, a first-order low-pass filter is adopted to perform interference acceleration acc obtained in the step 3drPerforming first-order low-pass filtering to remove interference acceleration accdrAnd (8) obtaining a final required interference acceleration acc in the reference coordinate system by using the noise term △ a.
Step 5, converting the final interference acceleration acc under the reference coordinate system obtained in the step 4 into the final interference acceleration acc under the body coordinate systemfMeasuring acceleration a in a coordinate system of the bodymeasureAnd the final interference acceleration acc under the coordinate system of the machine bodyfAccurate body acceleration a to c for obtaining filtering interference acceleration by subtractingmeasureAnd the method is used for an attitude calculation algorithm.
Wherein, a &measureI.e. the exact body acceleration for which the disturbance acceleration is filtered out,
Claims (1)
1. An interference acceleration measurement method for aircraft attitude calculation is characterized by comprising the following steps:
step 1, acquiring attitude angles in an Euler angle form of an aircraft in real time in a set sampling period, wherein the attitude angles comprise a rolling angle, a pitch angle and a yaw angle in the Euler angle form, and converting the attitude angles into attitude quaternions
Step 2, obtaining the coordinate system of the bodyMeasuring acceleration ameasureBy attitude quaternionAnd the measured acceleration a under the coordinate system of the machine bodymeasureObtaining the disturbance acceleration a under the reference coordinate systemdrTo obtain discrete disturbance acceleration
Step 3, obtaining the interference acceleration acc through the following formuladr：
Wherein, | | ameasureI and grRespectively representing the measured acceleration a under the coordinate system of the machine bodymeasureAnd the gravity acceleration g under the reference coordinate systemrThe value of the modulus of the (c) component,setting a threshold value;
step 4, adopting a first-order low-pass filter to perform interference acceleration acc obtained in the step 3drPerforming first-order low-pass filtering to obtain the final interference acceleration acc under a reference coordinate system,
step 5, converting the final interference acceleration acc under the reference coordinate system obtained in the step 4 into the final interference acceleration acc under the body coordinate systemfMeasuring acceleration a in a coordinate system of the bodymeasureAnd the final interference acceleration acc under the coordinate system of the machine bodyfSubtracting to obtain accurate body acceleration a of filtering interference acceleration～ measure。
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108318035A (en) * | 2018-01-08 | 2018-07-24 | 山东大学 | A method of avoiding universal joint deadlock in the attitude measurement algorithm based on Eulerian angles |
| CN109471192A (en) * | 2018-12-19 | 2019-03-15 | 湖南航天机电设备与特种材料研究所 | A kind of full-automatic gravity tester Dynamic High-accuracy data processing method |
| CN109635388A (en) * | 2018-11-29 | 2019-04-16 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of full machine acceleration parameter measuring method |
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| CN109635388A (en) * | 2018-11-29 | 2019-04-16 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of full machine acceleration parameter measuring method |
| CN109471192A (en) * | 2018-12-19 | 2019-03-15 | 湖南航天机电设备与特种材料研究所 | A kind of full-automatic gravity tester Dynamic High-accuracy data processing method |
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Denomination of invention: Measuring method for disturbance acceleration in attitude estimation of aircraft Effective date of registration: 20200518 Granted publication date: 20190830 Pledgee: Tongling Tianyuan Equity Investment Group Co.,Ltd. Pledgor: WUHAN YIHESMART TECHNOLOGY Co.,Ltd. Registration number: Y2020340000024 |
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