CN114707317A - Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge - Google Patents

Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge Download PDF

Info

Publication number
CN114707317A
CN114707317A CN202210305647.XA CN202210305647A CN114707317A CN 114707317 A CN114707317 A CN 114707317A CN 202210305647 A CN202210305647 A CN 202210305647A CN 114707317 A CN114707317 A CN 114707317A
Authority
CN
China
Prior art keywords
projectile
attitude
quaternion
matrix
measurement
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
CN202210305647.XA
Other languages
Chinese (zh)
Inventor
龙达峰
程福亨
黄恺健
赵振廷
孙俊丽
魏晓慧
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Huizhou University
Original Assignee
Huizhou University
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Huizhou University filed Critical Huizhou University
Priority to CN202210305647.XA priority Critical patent/CN114707317A/en
Publication of CN114707317A publication Critical patent/CN114707317A/en
Pending legal-status Critical Current

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B35/00—Testing or checking of ammunition
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F17/00—Digital computing or data processing equipment or methods, specially adapted for specific functions
    • G06F17/10—Complex mathematical operations
    • G06F17/11—Complex mathematical operations for solving equations, e.g. nonlinear equations, general mathematical optimization problems
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F17/00—Digital computing or data processing equipment or methods, specially adapted for specific functions
    • G06F17/10—Complex mathematical operations
    • G06F17/16—Matrix or vector computation, e.g. matrix-matrix or matrix-vector multiplication, matrix factorization

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • Mathematical Physics (AREA)
  • Data Mining & Analysis (AREA)
  • General Engineering & Computer Science (AREA)
  • Computational Mathematics (AREA)
  • Mathematical Analysis (AREA)
  • Mathematical Optimization (AREA)
  • Pure & Applied Mathematics (AREA)
  • Algebra (AREA)
  • Databases & Information Systems (AREA)
  • Software Systems (AREA)
  • Computing Systems (AREA)
  • Computer Hardware Design (AREA)
  • Evolutionary Computation (AREA)
  • Geometry (AREA)
  • Operations Research (AREA)
  • Navigation (AREA)
  • Gyroscopes (AREA)

Abstract

The invention relates to a method and a system for filtering flight parameters of a spinning projectile based on trajectory priori knowledge, which select a system state equation taking a quaternion q of the attitude of the spinning projectile as a state variable of a filter and a quaternion q of a magnetic measurement resolving systemMRSystematic observation equation as filter observation variable, through projectile attitude quaternion matrix
Figure DDA0003564864800000011
Solving the quaternion q of the magnetic survey solution systemMR(ii) a Further constructing a projectile attitude filtering model, and finishing the optimal filtering estimation of the state equation X (k) by adopting a Kalman filtering algorithm
Figure DDA0003564864800000012
Calculating to obtain the optimal estimation of the quaternion of the attitude of the spinning projectile
Figure DDA0003564864800000013
And substituting into the projectile attitude quaternion matrix
Figure DDA0003564864800000014
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure DDA0003564864800000015
And calculating projectile velocity VnAnd projectile flight position PnAnd the real-time parameter measurement of the spinning projectile flight is completed, so that the accuracy and the reliability of the testing of the spinning projectile state are effectively improved.

Description

Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge
Technical Field
The invention relates to the technical field of methods for measuring flight attitude, speed and position of an aircraft or a spinning projectile, in particular to a method and a system for measuring flight parameters of the spinning projectile based on trajectory prior knowledge.
Background
Due to the special application environment of 'three high' of high overload, high autogyration and high speed of the launching of the rotating bomb, the existing mature bomb-borne measuring system can not be directly transplanted and applied to the testing application of the rotating bomb, and the problems of difficult survival, incomplete measuring parameters, low precision, poor reliability and the like exist. Therefore, the high-precision attitude measurement technology for the high-speed spinning projectile is difficult to conduct guidance transformation, a technology which is low in cost, high in precision and suitable for flight attitude measurement of the spinning projectile is urgently needed to be found, and the high-precision attitude measurement technology has important theoretical value and practical significance for solving the problem of aerial flight attitude measurement in guidance transformation of the spinning projectile.
Disclosure of Invention
Aiming at the problems, the invention provides a method for filtering flight parameters of a spinning projectile based on trajectory prior knowledge, which comprises the following steps:
s1 selecting system state equation with rotating missile attitude quaternion q as filter state variable and system quaternion calculated by magnetic measurementNumber q ofMRSystem observation equation as filter observation variable, through projectile attitude quaternion matrix
Figure BDA0003564864780000011
Solving the quaternion q of the magnetic survey solution systemMR；
S2: constructing a missile body attitude filtering model, and finishing optimal filtering estimation of a state equation X (k) by adopting a Kalman filtering algorithm
Figure BDA0003564864780000012
Calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure BDA0003564864780000013
And substituting into the projectile attitude quaternion matrix
Figure BDA0003564864780000014
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure BDA0003564864780000015
S3: calculating projectile velocity VnAnd projectile flight position Pn；
S4: and finishing the real-time parameter measurement of the spinning projectile flight.
Defining the relation between the geomagnetic sensor measurement output and the geomagnetic vector as follows:
Figure BDA0003564864780000016
according to trajectory prior knowledge, let
Figure BDA0003564864780000021
Equal to the directive angle alpha, and calculating the pitch angle theta of the high-speed rotating bombmAnd roll angle gammamThe formula is as follows:
Figure BDA0003564864780000022
calculating projectile attitude transformation matrix
Figure BDA0003564864780000023
The method is simplified as follows:
Figure BDA0003564864780000024
in the above formula, the first and second carbon atoms are,
Figure BDA0003564864780000025
output geomagnetic field vector information is measured for the triaxial geomagnetic sensor,
Figure BDA0003564864780000026
three components of the geomagnetic field under a navigation coordinate system;
Figure BDA0003564864780000027
a projectile attitude transformation matrix;
Figure BDA0003564864780000028
theta, gamma are expressed as the yaw, pitch and roll angles of the projectile, respectively.
Wherein the magnetism is calculated the system quaternion qMRThe method specifically comprises the following steps:
let q be the attitude quaternion of the projectile coordinate system relative to the navigation coordinate system, then the projectile attitude quaternion matrix
Figure BDA0003564864780000029
Can be expressed as:
Figure BDA00035648647800000210
quaternion of the magnetic measurement solution system
Figure BDA00035648647800000211
The solving formula of (1) is as follows:
Figure BDA00035648647800000212
in the above formula
Figure BDA0003564864780000031
The symbol is defined as
Figure BDA0003564864780000032
Is determined by:
Figure BDA0003564864780000033
in the above formula, q is q0+q1i+q2j+q3k，q0,q1,q2,q3Is a real number, i, j, k are mutually orthogonal imaginary units, expressed as q ═ q0,q1,q2,q3]T；
Figure BDA0003564864780000034
And
Figure BDA0003564864780000035
all represent a projectile attitude transformation matrix,
Figure BDA0003564864780000036
the system state equation comprises:
selecting the attitude quaternion q as a filter state variable X (t) q0,q1,q2,q3]TThe state equation of the system is as follows:
Figure BDA0003564864780000037
Figure BDA0003564864780000038
in the above formula, w (t) is the system process noise with the average value E [ w (t)]0, variance E [ w (t), wT(τ)]Q (t), q (t) is the system noise sequence variance matrix, ωi(i ═ x, y, z) is expressed as the rotational bounce angle rate component.
The system observation equation is as follows: selecting quaternion q of the magnetic measurement resolving systemMRAs the observation variable z (t) of the filter, the system observation measurement equation is:
Figure BDA0003564864780000039
in the above formula, v (t) is the measurement noise of the system, and the average value is E [ v (t)]0, variance E [ v (t), vT(τ)]R (t), r (t) is a measurement noise sequence variance matrix.
The method for constructing the missile attitude filtering model specifically comprises the following steps:
simplifying the system state equation and the system observation measurement equation and then carrying out discretization treatment to obtain:
Figure BDA0003564864780000041
in the above formula, the first and second carbon atoms are,
Figure BDA0003564864780000042
is an identity matrix, phik,k-1For the purpose of the state transition matrix calculation formula,
Figure BDA0003564864780000043
Figure BDA0003564864780000044
x (k-1) is the next state variable, w (k) is the system process noise, and v (k) is the system measurement noise.
Further, the optimal filtering estimation of the state equation X (k) is completed by adopting the Kalman filtering algorithm
Figure BDA0003564864780000045
The method comprises the following steps: a time update process and a measurement update process; wherein the content of the first and second substances,
the time updating process is represented by the formula:
Figure BDA0003564864780000046
the measurement updating process has the formula:
Figure BDA0003564864780000047
calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure BDA0003564864780000048
And substituting into the projectile attitude quaternion matrix
Figure BDA0003564864780000049
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure BDA00035648647800000410
As spin projectile attitude parameter a;
wherein, in the formula, the first and second groups,
Figure BDA00035648647800000411
predicting the state in one step; p(k,k-1)Predicting the mean square error for one step; k iskRepresenting the filter gain; r (k) is a measurement noise array; q (k-1) is a system noise variance matrix at the k-1 moment; i is an identity matrix; pkTo estimate the mean square error.
And further. Characterized in that, S3 still includes:
calculating the projectile velocity Vn and the projectile flight position Pn according to the formula:
Figure BDA00035648647800000412
in the above formula, the quaternion matrix of the projectile attitude
Figure BDA00035648647800000413
Measuring an output component, g, for a missile-borne triaxial accelerometernIs the gravity component under the navigation coordinate system; vn＝[vx,vy,vz]TIs the projectile velocity component;
Pn＝[x,y,z]Tis the rotating projectile flight position component.
As another preferred aspect, the present invention further provides a rotating projectile flight parameter measurement system based on trajectory prior knowledge, which at least includes a signal acquisition module and a signal processing module, wherein:
the signal acquisition module comprises a triaxial geomagnetic sensor, a triaxial MEMS gyroscope and a triaxial accelerometer which are sequentially in communication connection with the signal conditioning module; the signal conditioning module sends the acquired data to the ADC data acquisition module to complete data acquisition;
the signal processing module comprises an FPGA configuration module, a DSP digital signal processor and a FLASH which are sequentially in communication connection with the FPGA unit; the FLASH has a computer program stored therein, which when invoked and executed by the DSP digital signal processor, is configured to implement the method for filtering parameters of a spinning projectile based on ballistic prior knowledge as claimed in any one of claims 1 to 8;
further comprising:
the missile-borne flight control computer is connected with the DSP and controls the rotating missile to fly according to the attitude parameter A of the rotating missile, the speed Vn of the missile and the flight position Pn of the missile;
and the upper computer is connected with the FPGA unit and is used for reading the data acquired by the signal acquisition module and reading the data in the FLASH.
The triaxial geomagnetic sensor, the triaxial MEMS gyroscope and the triaxial accelerometer are respectively and correspondingly installed on an Xb axis, a Zb axis and an axis center 0 point position in a high-speed rotating missile carrier coordinate system.
In summary, the present invention providesThe method and the system for filtering the flight parameters of the spinning projectile based on the trajectory priori knowledge select a system state equation with a quaternion q of the attitude of the spinning projectile as a state variable of a filter and solve the quaternion q of the system by magnetic measurementMRSystem observation equation as filter observation variable, through projectile attitude quaternion matrix
Figure BDA0003564864780000051
Solving the quaternion q of the magnetic survey solution systemMR(ii) a Further constructing a projectile attitude filtering model, and finishing the optimal filtering estimation of the state equation X (k) by adopting a Kalman filtering algorithm
Figure BDA0003564864780000052
Calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure BDA0003564864780000053
And substituting into the projectile attitude quaternion matrix
Figure BDA0003564864780000054
In the method, the quaternion matrix of the optimal projectile body attitude is obtained by calculation
Figure BDA0003564864780000055
And calculating projectile velocity VnAnd projectile flight position PnAnd the real-time parameter measurement of the spinning projectile flight is completed, so that the accuracy and the reliability of the testing of the spinning projectile state are effectively improved.
Drawings
Fig. 1 is a schematic diagram of a system for measuring flight parameters of a spinning projectile based on ballistic prior knowledge according to an embodiment.
FIG. 2 is a schematic diagram of a spinning ball navigation sensor mounting and coordinate system according to an embodiment.
Fig. 3 is a schematic diagram of a method for filtering flight parameters of a spinning projectile based on ballistic prior knowledge according to an embodiment.
Detailed Description
The method for filtering flight parameters of a spinning projectile based on ballistic prior knowledge according to the present invention will be described in further detail with reference to the following embodiments and accompanying drawings.
As shown in fig. 1, the system for measuring flight parameters of a spinning projectile based on ballistic prior knowledge at least comprises a signal acquisition module and a signal processing module, wherein:
the signal acquisition module comprises a triaxial geomagnetic sensor, a triaxial MEMS gyroscope and a triaxial accelerometer which are sequentially in communication connection with the signal conditioning module; the signal conditioning module sends the acquired data to the ADC data acquisition module to complete data acquisition;
the signal processing module comprises an FPGA configuration module, a DSP digital signal processor and a FLASH which are sequentially in communication connection with the FPGA unit; the FLASH is stored with a computer program for implementing the method for filtering parameters of a spinning projectile flight based on ballistic prior knowledge according to any one of claims 1 to 8 when the computer program is called and executed by the DSP digital signal processor;
further comprising:
a missile-borne flight control computer connected with the DSP and used for controlling the speed V of the missile body according to the attitude parameters of the rotating missilenAnd projectile flight position PnControlling the rotary bomb to fly.
And the upper computer is connected with the FPGA unit and is used for reading the data acquired by the signal acquisition module and reading the data in the FLASH.
As shown in fig. 2, the three-axis geomagnetic sensor, the three-axis MEMS gyroscope and the three-axis accelerometer are respectively installed in the X-axis coordinate system of the high-speed rotating missile carrier correspondinglybAxis, ZbAxis and axis center 0 point position.
As shown in fig. 3, the method for filtering flight parameters of a spinning projectile based on ballistic prior knowledge in the present invention includes the following steps:
s1 selecting the system state equation with rotating missile attitude quaternion q as the filter state variable and the quaternion q of the magnetic measurement resolving systemMRSystem observation equation as filter observation variable, through projectile attitude quaternion matrix
Figure BDA0003564864780000061
Solving the quaternion q of the magnetic survey solution systemMR；
S2: constructing a missile body attitude filtering model, and finishing optimal filtering estimation of a state equation X (k) by adopting a Kalman filtering algorithm
Figure BDA0003564864780000062
Calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure BDA0003564864780000063
And substituting into a projectile attitude quaternion matrix
Figure BDA0003564864780000064
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure BDA0003564864780000065
S3: calculating projectile velocity VnAnd projectile flight position Pn；
S4: and finishing the real-time parameter measurement of the spinning projectile flight.
Wherein, the relation between the geomagnetic sensor measurement output and the geomagnetic vector is defined as:
Figure BDA0003564864780000071
in the above-mentioned formula (1),
Figure BDA0003564864780000072
is the output of the triaxial geomagnetic sensor measurement,
Figure BDA0003564864780000073
three components of the geomagnetic field under a navigation coordinate system;
Figure BDA0003564864780000074
for a projectile attitude transformation matrix, abbreviated
Figure BDA0003564864780000075
θ, γ are expressed as the yaw, pitch and roll angles of the projectile, respectively.
The rotating projectile realizes stable flight by high-speed rotation according to the trajectory prior knowledge, and the change of the rotating projectile in a launching navigation coordinate system is not large according to the trajectory prior knowledge,
Figure BDA0003564864780000076
the yaw angle can be regarded as a known angle (
Figure BDA0003564864780000077
Equal to the firing angle alpha), for simplicity of calculation
Figure BDA0003564864780000078
Processing at an angle of zero degrees. Thus, let
Figure BDA0003564864780000079
Equal to the directive angle alpha, and calculating the pitch angle theta of the high-speed rotating projectilemAnd roll angle gammamThe formula is as follows:
Figure BDA00035648647800000710
calculating projectile attitude transformation matrix
Figure BDA00035648647800000711
The method is simplified as follows:
Figure BDA00035648647800000712
if q represents the attitude quaternion of the projectile coordinate system relative to the navigation coordinate system, the projectile attitude quaternion matrix can be obtained by strapdown inertial navigation
Figure BDA00035648647800000713
Can be expressed as:
Figure BDA0003564864780000081
in the above formula, q is q0+q1i+q2j+q3k，q0,q1,q2,q3Is a real number, i, j, k are mutually orthogonal imaginary units, and may also be expressed as q ═ q0,q1,q2,q3]T。
Figure BDA0003564864780000082
And
Figure BDA0003564864780000083
all represent a projectile attitude transformation matrix,
Figure BDA0003564864780000084
the quaternion of the attitude of the spinning projectile can be deduced
Figure BDA0003564864780000085
The solving formula of (1) is as follows:
Figure BDA0003564864780000086
in the above formula (5)
Figure BDA0003564864780000087
(symbol)
Figure BDA0003564864780000088
Optionally, the content of the compound can be selected,
Figure BDA0003564864780000089
is determined by the following formula (6):
Figure BDA00035648647800000810
the method comprises the following steps of selecting a system state equation taking a rotating missile attitude quaternion q as a filter state variable, wherein the system state equation comprises the following steps:
selecting the attitude quaternion q as a filter state variable X (t) q0,q1,q2,q3]TIf ω isi(i ═ x, y, z) is expressed as the rotational missile angular velocity component, and according to the strapdown inertial navigation theory, the missile attitude equation described by quaternion is:
Figure BDA00035648647800000811
the filter state variable x (t) q ═ q in the present invention0,q1,q2,q3]TThe corresponding system state equation is:
Figure BDA0003564864780000091
in the above equation (8), w (t) is the system process noise, which is assumed to be zero-mean white noise, and the mean value is E [ w (t)]0, variance E [ w (t), wT(τ)]Q (t), q (t) is the system noise sequence variance matrix, ωi(i ═ x, y, z) is expressed as the rotational bounce angle rate component.
Selecting quaternion q of the magnetic measurement resolving systemMRAs the observation variable z (t) of the filter, the system observation measurement equation is:
Figure BDA0003564864780000092
in the above formula (9), v (t) is the measured noise of the system, assuming zero mean white noise, and its mean is E [ v (t)]0, variance E [ v (t), vT(τ)]R (t), r (t) is a measurement noise sequence variance matrix.
Therefore, a missile body attitude filtering model of the system is formed by the state equation (8) and the observation equation (9), and the Kalman filtering method is adopted to realize the optimal estimation of the missile body attitude quaternion q
Figure BDA0003564864780000093
The spin-missile attitude filtering equations (8) and (9) are simplified to a general form:
Figure BDA0003564864780000094
simplifying the system state equation and the system observation measurement equation and then carrying out discretization treatment to obtain:
Figure BDA0003564864780000095
in the above-mentioned formula (11),
Figure BDA0003564864780000096
is an identity matrix, phik,k-1The state transition matrix calculation formula is:
Figure BDA0003564864780000097
in the above-mentioned formula (12),
Figure BDA0003564864780000101
is the rotational bounce angle rate component;
Figure BDA0003564864780000102
x (k-1) is the next state variable, w (k) is the system process noise, and v (k) is the system measurement noise.
The invention adopts the discrete Kalman filtering algorithm to complete the optimal filtering estimation of the state variable X (k)
Figure BDA00035648647800001011
The filtering algorithm comprises a time updating process and a measurement updating process, wherein:
(1) and (3) time updating process:
Figure BDA0003564864780000103
(2) and (3) measurement updating process:
Figure BDA0003564864780000104
calculating to obtain the optimal estimation of the quaternion of the attitude of the spinning projectile
Figure BDA0003564864780000105
Substituting the quaternion matrix into the projectile attitude of the formula (4)
Figure BDA0003564864780000106
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure BDA0003564864780000107
As spin projectile attitude parameter a;
in the above formula, the first and second carbon atoms are,
Figure BDA0003564864780000108
one-step prediction for state; p(k,k-1)Predicting the mean square error for one step; kkRepresenting the filter gain; r (k) is a measurement noise array; q (k-1) is a system noise variance matrix at the k-1 moment; i is an identity matrix; pkTo estimate the mean square error.
The S3, further comprising:
calculating projectile velocity VnAnd projectile flight position PnThe formula is as follows:
Figure BDA0003564864780000109
in the above formula, the matrix of quaternion of projectile attitude
Figure BDA00035648647800001010
Measuring an output component, g, for a missile-borne triaxial accelerometernIs the gravity component under the navigation coordinate system; vn＝[vx,vy,vz]TIs the projectile velocity component; pn＝[x,y,z]TIs the rotating projectile flight position component.
While the invention has been described in conjunction with the specific embodiments set forth above, it is evident that many alternatives, modifications, and variations will be apparent to those skilled in the art in light of the foregoing description. Accordingly, it is intended to embrace all such alternatives, modifications, and variations that fall within the spirit and scope of the appended claims.

Claims (10)

1. A method for filtering flight parameters of a spinning projectile based on ballistic prior knowledge is characterized by comprising the following steps:
s1 selecting the system state equation with rotating missile attitude quaternion q as the filter state variable and the quaternion q of the magnetic measurement resolving systemMRSystem observation equation as filter observation variable, through projectile attitude quaternion matrix
Figure FDA0003564864770000011
Solving the quaternion q of the magnetic survey solution systemMR；
S2: constructing a missile body attitude filtering model, and finishing optimal filtering estimation of a state equation X (k) by adopting a Kalman filtering algorithm
Figure FDA0003564864770000012
Calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure FDA0003564864770000013
And substituting into the projectile attitude quaternion matrix
Figure FDA0003564864770000014
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure FDA0003564864770000015
S3: calculating projectile velocity VnAnd projectile flight position Pn；
S4: and finishing the real-time parameter measurement of the spinning projectile flight.
2. The method of claim 1, further comprising: defining the relation between the geomagnetic sensor measurement output and the geomagnetic vector as follows:
Figure FDA0003564864770000016
according to trajectory prior knowledge, let
Figure FDA0003564864770000017
Equal to the directive angle alpha, and calculating the pitch angle theta of the high-speed rotating projectilemAnd roll angle gammamThe formula is as follows:
Figure FDA0003564864770000018
calculating projectile attitude transformation matrix
Figure FDA0003564864770000019
The method is simplified as follows:
Figure FDA0003564864770000021
in the above formula, the first and second carbon atoms are,
Figure FDA0003564864770000022
output geomagnetic field vector information is measured for the triaxial geomagnetic sensor,
Figure FDA0003564864770000023
three components of the geomagnetic field under a navigation coordinate system are obtained;
Figure FDA0003564864770000024
a projectile attitude transformation matrix;
Figure FDA0003564864770000025
θ, γ are expressed as the yaw, pitch and roll angles of the projectile, respectively.
3. The method of claim 1, wherein the magnetometric solution system quaternion q is a function of the number of spin projectile flight parameters to be filteredMRThe method specifically comprises the following steps:
let q be the attitude quaternion of the projectile coordinate system relative to the navigation coordinate system, then the projectile attitude quaternion matrix
Figure FDA0003564864770000026
Can be expressed as:
Figure FDA0003564864770000027
quaternion of the magnetic measurement solution system
Figure FDA0003564864770000028
The solving formula of (1) is as follows:
Figure FDA0003564864770000029
in the above formula
Figure FDA00035648647700000210
The symbol is defined as
Figure FDA00035648647700000211
Figure FDA00035648647700000212
Is determined by:
Figure FDA00035648647700000213
in the above formula, q is q0+q1i+q2j+q3k，q0,q1,q2,q3Is a real number, i, j, k are mutually orthogonal imaginary units, expressed as q ═ q0,q1,q2,q3]T；
Figure FDA00035648647700000214
And
Figure FDA00035648647700000215
all represent a projectile attitude transformation matrix,
Figure FDA00035648647700000216
4. the method of claim 1, wherein the system state equation comprises:
selecting the attitude quaternion q as a filter state variable X (t) q0,q1,q2,q3]TThe state equation of the system is as follows:
Figure FDA0003564864770000031
Figure FDA0003564864770000032
in the above formula, w (t) is the system process noise with the average value E [ w (t)]0, variance E [ w (t), wT(τ)]Q (t), q (t) is the system noise sequence variance matrix, ωi(i ═ x, y, z) is expressed as the rotational bounce angle rate component.
5. The method for filtering parameters of a spinning projectile flight based on ballistic prior knowledge as claimed in claim 1, wherein the system observation equation is: selecting quaternion q of the magnetic measurement resolving systemMRAs the observation variable z (t) of the filter, the system observation measurement equation is:
Figure FDA0003564864770000033
in the above formula, v (t) is the measurement noise of the system, and the average value is E [ v (t)]0, variance E [ v (t), vT(τ)]R (t), r (t) is a measurement noise sequence variance matrix.
6. The method for filtering parameters of a spinning projectile flight based on ballistic prior knowledge as claimed in claim 1, wherein the constructing of the projectile attitude filtering model specifically comprises:
simplifying the system state equation and the system observation measurement equation and then carrying out discretization treatment to obtain:
Figure FDA0003564864770000034
in the above formula, the first and second carbon atoms are,
Figure FDA0003564864770000035
is an identity matrix, phik,k-1For the calculation of the state transition matrix the formula,
Figure FDA0003564864770000036
x (k-1) is the next state variable, w (k) is the system process noise, and v (k) is the system measurement noise.
7. The method for filtering flight parameters of a spinning projectile based on ballistic prior knowledge as claimed in claim 1, wherein said applying kalman filtering algorithm completes equation of state X (X: (b))k) To estimate the optimal filtering
Figure FDA0003564864770000041
The method comprises the following steps: a time update process and a measurement update process; wherein the content of the first and second substances,
the time updating process is represented by the formula:
Figure FDA0003564864770000042
the measurement updating process has the formula:
Figure FDA0003564864770000043
calculating to obtain optimal estimation of rotating projectile attitude quaternion
Figure FDA0003564864770000044
And substituting into the projectile attitude quaternion matrix
Figure FDA0003564864770000045
In the middle, calculating to obtain quaternion matrix of optimal projectile attitude
Figure FDA0003564864770000046
As spin projectile attitude parameter a;
in the above formula, the first and second carbon atoms are,
Figure FDA0003564864770000047
predicting the state in one step; p(k,k-1)Predicting the mean square error for one step; k iskRepresenting the filter gain;
r (k) is a measurement noise array, and Q (k-1) is a system noise variance array at the k-1 moment; i is an identity matrix; pkTo estimate the mean square error.
8. The method for filtering parameters of a spinning projectile flight based on ballistic prior knowledge as claimed in claim 1, wherein said S3 further comprises:
calculating projectile velocity VnAnd projectile flight position PnThe formula is as follows:
Figure FDA0003564864770000048
in the above formula, the matrix of quaternion of projectile attitude
Figure FDA0003564864770000049
Figure FDA00035648647700000410
Measuring an output component, g, for a missile-borne triaxial accelerometernIs the gravity component under the navigation coordinate system; vn＝[vx,vy,vz]TIs the projectile velocity component; pn＝[x,y,z]TIs the rotating projectile flight position component.
9. The rotating projectile flight parameter measurement system based on trajectory prior knowledge is characterized by comprising a signal acquisition module and a signal processing module, wherein:
the signal acquisition module comprises a triaxial geomagnetic sensor, a triaxial MEMS gyroscope and a triaxial accelerometer which are sequentially in communication connection with the signal conditioning module; the signal conditioning module sends the acquired data to the ADC data acquisition module to complete data acquisition;
the signal processing module comprises an FPGA configuration module, a DSP digital signal processor and a FLASH which are sequentially in communication connection with the FPGA unit; the FLASH has a computer program stored therein, which when invoked and executed by the DSP digital signal processor, is configured to implement the method for filtering parameters of a spinning projectile based on ballistic prior knowledge as claimed in any one of claims 1 to 8;
further comprising:
a missile-borne flight control computer connected with the DSP and used for controlling the flight of the missile according to the attitude parameters of the rotating missileA, projectile velocity VnAnd projectile flight position PnControlling the rotating projectile to fly;
and the upper computer is connected with the FPGA unit and is used for reading the data acquired by the signal acquisition module and reading the data in the FLASH.
10. The system of claim 9, wherein the tri-axial geomagnetic sensor, the tri-axial MEMS gyroscope, and the tri-axial accelerometer are respectively corresponding to the Xb-axis, Zb-axis, and 0-axis center point positions installed in the coordinate system of the high-speed rotating projectile carrier.
CN202210305647.XA 2022-03-25 2022-03-25 Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge Pending CN114707317A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202210305647.XA CN114707317A (en) 2022-03-25 2022-03-25 Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202210305647.XA CN114707317A (en) 2022-03-25 2022-03-25 Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge

Publications (1)

Publication Number Publication Date
CN114707317A true CN114707317A (en) 2022-07-05

Family

ID=82170639

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202210305647.XA Pending CN114707317A (en) 2022-03-25 2022-03-25 Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge

Country Status (1)

Country Link
CN (1) CN114707317A (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116992553A (en) * 2023-05-25 2023-11-03 中国人民解放军32804部队 Whole-course trajectory estimation method of boosting gliding aircraft

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116992553A (en) * 2023-05-25 2023-11-03 中国人民解放军32804部队 Whole-course trajectory estimation method of boosting gliding aircraft
CN116992553B (en) * 2023-05-25 2024-02-06 中国人民解放军32804部队 Whole-course trajectory estimation method of boosting gliding aircraft

Similar Documents

Publication Publication Date Title
CN107314718B (en) High speed rotation bullet Attitude estimation method based on magnetic survey rolling angular rate information
CN109373833B (en) Combined measurement method suitable for initial attitude and speed of spinning projectile
CN104698485B (en) Integrated navigation system and air navigation aid based on BD, GPS and MEMS
CN109373832B (en) Method for measuring initial parameters of rotating projectile muzzle based on magnetic rolling
JP2006514258A (en) Projectile guidance using accelerometers and GPS receivers
CN104697521A (en) Method for measuring posture and angle speed of high-speed rotating body by gyro redundant oblique configuration mode
CN102853834A (en) High-precision scheme of IMU for rotating carrier and denoising method
CN111486870B (en) System-level calibration method for inclined strapdown inertial measurement unit
CN114707317A (en) Method and system for measuring flight parameters of spinning projectile based on trajectory prior knowledge
Hoang et al. Measurement optimization for orientation tracking based on no motion no integration technique
CN107807375B (en) Unmanned aerial vehicle attitude tracking method and system based on multiple GPS receivers
Parfiriev et al. Model of inertial navigation system for unmanned aerial vehicle based on MEMS
Cui et al. Calibration of MEMS accelerometer using kaiser filter and the ellipsoid fitting method
CN114964224B (en) Error autonomous suppression method for micro inertial navigation system
Wang et al. Functional iteration in-flight alignment method for projectiles MSINS
CN113932803B (en) Inertial/geomagnetic/satellite integrated navigation system suitable for high-dynamic aircraft
CN116125789A (en) Gesture algorithm parameter automatic matching system and method based on quaternion
CN114398827A (en) Virtual gyroscope construction method based on deep learning
CN112683265B (en) MIMU/GPS integrated navigation method based on rapid ISS collective filtering
Zhang et al. Research on UAV attitude data fusion algorithm based on quaternion gradient descent
Keke et al. Research on integrated attitude determination methods based on MEMS device for quadrotor UAVs
CN111397635A (en) Rotation modulation method for MEMS inertial navigation system
CN116070066B (en) Method for calculating rolling angle of guided projectile
CN113984042B (en) Series combined navigation method applicable to high-dynamic aircraft
CN114234971B (en) Noise-reducing full-attitude IMU attitude calculation method

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination