Method and device for installing and calibrating inertial navigation device
By establishing a calibration measurement coordinate system and performing geometric calculations, the installation and calibration process of inertial navigation devices is simplified, solving the problems of complexity and low accuracy of existing methods. This achieves efficient and convenient calibration results, applicable to different aircraft structures.
Patent Information
- Application Number
- CN202511416493.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-09-30
- Publication Date
- 2025-12-19
AI Technical Summary
Existing methods for installing and calibrating inertial navigation devices are complex to operate, have low accuracy, limited applicability, and are subject to limitations imposed by external conditions, making it difficult to achieve efficient and convenient calibration.
The coordinate system measurement method is adopted. By establishing a calibrated measurement coordinate system, the coordinates of the installation point of the inertial navigation device are measured. The roll angle, pitch angle and heading angle errors are obtained by geometric calculation, which simplifies the operation, eliminates the use of adapter plates and flat plates, and is adaptable to different aircraft structures.
It improves calibration accuracy and ease of use, shortens measurement time, enhances the versatility and environmental adaptability of the method, reduces equipment requirements, and makes field maintenance more convenient.
Smart Images
Figure CN121163552A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application belongs to the technical field of installation adjustment of integrated navigation system, and particularly relates to an installation calibration method and device of an inertial navigation device. BACKGROUND
[0002] An inertial navigation system is an aided navigation system which uses accelerometers and gyroscopes to measure the acceleration and angular velocity of an object, and uses a computer to continuously estimate the position, attitude and velocity of the moving object. Before use, it needs to be provided with initial position and velocity so that the system can detect the position change (such as movement east or west), velocity change (velocity size or direction) and attitude change (rotation around each axis) of the aircraft. Therefore, the attitude of the installed inertial navigation device must be consistent with the current attitude of the aircraft so as to be able to provide the aircraft attitude (X-aircraft roll axis 13, Y-aircraft pitch axis 12, Z-aircraft heading axis 11) information, such as Figure 1 . Since there is an installation error between the two, calibration is required before installation. In order to meet this technical requirement, the current technical method is as follows: Step one, adjust the aircraft body to a horizontal state.
[0003] Step two, as shown in Figure 2 , install the adapter plate 21 to the installation position of the inertial navigation device on the aircraft, the adapter plate 21 is provided with two heading measurement reference holes 24 and roll measurement reference line 23, pitch measurement reference line 22, and the aircraft heading normal line 25 is marked in Figure 2 .
[0004] Step three, use the inclinometer to measure the roll angle error α and pitch angle error β of the installation surface at the roll measurement reference line 23 and the pitch measurement reference line 22, and determine the positive and negative signs according to the inclination direction.
[0005] Step four, find two reference points on the heading axis of the aircraft, set a flat plate at a distance h from the front of the aircraft and adjust its position to ensure that the flat plate plane is perpendicular to the line connecting the two reference points on the aircraft, install a target mirror on the adapter plate of the inertial measurement component, and observe the distance L between the center of the target mirror and the center line of the flat plate, as shown in Figure 3 . Figure 3 , which includes the measurement flat plate 31, the aircraft attitude heading coordinate axis 32, the inertial navigation component 33, the target mirror installation position 34, the target mirror center line 35, the aircraft heading normal line 36, the heading angle error 37, the distance L between the target mirror center and the flat plate center line 38, and the distance 39 between the target mirror and the measurement flat plate. According to the formula, the heading angle error δ = arctan (L / h) is calculated, and finally the positive and negative signs are determined according to the position of the target point on the flat plate center line.
[0006] Step five, according to the installation error of the inertial navigation device measured in the above steps, the roll angle error α, the pitch angle error β and the heading angle error δ of the aircraft are input into the inertial navigation assembly system, and the system is corrected so as to accurately provide the aircraft attitude information.
[0007] The existing installation calibration technology method measures the roll angle error and the pitch angle error by using the inclinometer on the adapter plate. Due to the tooling error and the measurement error of the inclinometer, the measurement result is not high in accuracy and has a large error. When measuring the heading angle error, the flat plate is placed in front of the aircraft, and the position of the flat plate is repeatedly adjusted to ensure that it is perpendicular to the aircraft heading reference line. The operation is complex, the working space requirement is large, and the calibration time is long. At the same time, after the structure of part of the aircraft changes, the target is blocked and cannot be observed, and the method is not strong in universality. The flat plate adjustment and the target measurement distance can also cause low measurement accuracy. Finally, due to the limitation of the external field conditions, the method brings great difficulty to the implementation of the external field. In view of the above situation, it is necessary to design a new inertial navigation device installation calibration method which is convenient, simple, high in accuracy and reliability, and strong in applicability. SUMMARY
[0008] The present application provides an inertial navigation device installation calibration method and device, which solves the problems of the existing method, such as complex operation, difficult implementation, low accuracy and poor applicability.
[0009] The first aspect of the present application provides an inertial navigation device installation calibration method, comprising: S1, establishing a calibration measurement coordinate system and an aircraft horizontal plane a, a longitudinal section c and a transverse section b; S2, measuring the measurement point coordinates of the motion direction of the inertial navigation device in the calibration measurement coordinate system, and obtaining the roll angle error α, the pitch angle error β and the heading angle error δ according to the measurement point coordinates of the motion direction of the inertial navigation device; S3, correcting the inertial navigation device according to the obtained roll angle error α, pitch angle error β and heading angle error δ.
[0010] Optionally, the measurement point coordinates of the motion direction of the inertial navigation device are measured, and the roll angle error α, the pitch angle error β and the heading angle error δ are obtained according to the measurement point coordinates of the motion direction of the inertial navigation device, comprising: The roll installation hole points r1 and r2 of the inertial navigation device are measured and obtained, the distance from the point r1 to the aircraft horizontal plane is recorded as Z1, the distance from the point r1 to the aircraft longitudinal section is recorded as Y1, the distance from the point r2 to the aircraft horizontal plane is recorded as Z2, and the distance from the point r2 to the aircraft longitudinal section is recorded as Y2; The pitch installation hole points p1 and p2 of the inertial navigation device are measured and obtained, the distance from the point p1 to the aircraft horizontal plane is recorded as Z1 ′Let X1 be the distance from point p2 to the aircraft's cross-section, and Z2 be the distance from point p2 to the aircraft's horizontal plane. ′ The distance to the aircraft's cross-section is denoted as X2; Measure and obtain the mounting points y1 and y2 of the inertial navigation device's heading hole, and denote the distance from point y1 to the aircraft's longitudinal section as Y1. ′ The distance to the aircraft's cross-section is denoted as X1. ′ Let Y2 be the distance from point y2 to the horizontal plane of the aircraft. ′ The distance to the aircraft's cross-section is denoted as X2. ′ ; The roll angle error α is obtained as arctan(ΔZ1 / ΔY1). When Z1 > Z2, ΔZ1 = Z1 - Z2; otherwise, when Z1 < Z2, ΔZ1 = Z2 - Z1. When Y1 > Y2, ΔY1 = Y1 - Y2; otherwise, when Y1 < Y2, ΔY1 = Y2 - Y1. To obtain the pitch angle error β = arctan(ΔZ2 / ΔX1), when Z1 ′ >Z2 ′ At that time, ΔZ2=Z1 ′ -Z2 ′ Conversely, Z1 ′ <Z2 ′ When, ΔZ2=Z2 ′ -Z1 ′ When X1 > X2, ΔX1 = X1 - X2; conversely, when X1 < X2, ΔX1 = X2 - X1. Obtain the heading angle error δ=arctan(ΔY2 / ΔX2), when Y1 ′ >Y2 ′ At that time, ΔY2=Y1 ′ -Y2 ′ Conversely, Y1 ′ <Y2 ′ At that time, ΔY2=Y2 ′ -Y1 ′ When X1 ′ >X2 ′ At that time, ΔX2=X1 ′ -X2 ′ Conversely, X1 ′ <X2 ′ When, ΔX2=X2 ′ -X1 ′ .
[0011] Optionally, the sign of the roll angle error α is determined according to Z1 > Z2 or Z2 > Z1; The sign of the pitch angle error β is determined by Z1. ′ >Z2 ′ or Z2 ′ >Z1 ′determine; the positive or negative of the heading angle error δ is determined according to Y1 ′ > Y2 ′ or Y2 ′ > Y1 ′ determine.
[0012] Optionally, the coordinates of the measurement points of the inertial navigation device moving direction are measured, and the roll angle error α, the pitch angle error β and the heading angle error δ are obtained according to the coordinates of the measurement points of the inertial navigation device moving direction, comprising: In the measurement coordinate system, the coordinate data of four measurement points of the inertial navigation component are measured by using the laser tracker; the four measurement points include: the front two points along the heading of the aircraft and the rear two points along the heading of the aircraft; the distance from the measurement point 1 to the cross section b of the aircraft is denoted as x1, the distance from the measurement point 1 to the longitudinal section c of the aircraft is denoted as y1, the distance from the measurement point 1 to the horizontal plane a of the aircraft is denoted as z1, the distance from the measurement point 2 to the cross section b of the aircraft is denoted as x2, the distance from the measurement point 2 to the longitudinal section c of the aircraft is denoted as y2, the distance from the measurement point 2 to the horizontal plane of the aircraft is denoted as z2, the distance from the measurement point 3 to the cross section b of the aircraft is denoted as x3, the distance from the measurement point 3 to the longitudinal section c of the aircraft is denoted as y3, the distance from the measurement point 3 to the horizontal plane of the aircraft is denoted as z3, the distance from the measurement point 4 to the cross section b of the aircraft is denoted as x4, the distance from the measurement point 4 to the longitudinal section c of the aircraft is denoted as y4, and the distance from the measurement point 4 to the horizontal plane of the aircraft is denoted as z4; When (z1+z4)>(z2+z3), the roll angle error α=60arctan(Δz1 / Δy) is obtained according to Δz1=(z1+z4-z2-z3) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is positive; when (z1+z4)<(z2+z3), the roll angle error α=60arctan(Δz2 / Δy) is obtained according to Δz2=(z2+z3-z1-z4) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is negative; When (z1+z2)>(z3+z4), the pitch angle error β=60arctan(Δz3 / Δx) is obtained according to Δz3=(z1+z2-z3-z4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is positive; when (z1+z2)<(z3+z4), the pitch angle error β=60arctan(Δz4 / Δx) is obtained according to Δz4=(z3+z4-z1-z2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is negative; When (y1+y2)>(y3+y4), according to Δy2=(y1+y2-y3-y4) / 2, Δx=(x3+x4-x1-x2) / 2, the heading angle error δ=60arctan(Δy2 / Δx) is obtained, and the sign of δ is positive; when (y1+y2)<(y3+y4), according to Δy3=(y3+y4-y1-y2) / 2, Δx=(x3+x4-x1-x2) / 2, the heading angle error δ=60arctan(Δy3 / Δx) is obtained, and the sign of δ is negative.
[0013] Optionally, S1 comprises: S11, establishing a calibration measurement coordinate system according to the reference measurement points on the roll axis, the pitch axis and the heading axis of the aircraft; the calibration measurement coordinate system comprises the mutually perpendicular roll axis, the pitch axis and the heading axis; S12, establishing an aircraft horizontal plane a through the points on the horizontal plane of the aircraft, establishing an aircraft longitudinal section c through the Y-axis connecting line of the roll axis of the aircraft and perpendicular to the horizontal plane, and establishing an aircraft cross section b through the X-axis point of the pitch axis and perpendicular to the aircraft longitudinal section.
[0014] The second aspect of the application provides an inertial navigation device installation calibration device, comprising: a standard establishment module, configured to establish a calibration measurement coordinate system and an aircraft horizontal plane a, a longitudinal section c and a cross section b; a measurement module, configured to measure the coordinates of the measurement points in the motion direction of the inertial navigation device in the calibration measurement coordinate system, and obtain a roll angle error α, a pitch angle error β and a heading angle error δ according to the coordinates of the measurement points in the motion direction of the inertial navigation device; S3, correcting the inertial navigation device according to the obtained roll angle error α, the pitch angle error β and the heading angle error δ.
[0015] Optionally, the measurement module is specifically configured to: measure and obtain the roll installation hole points r1 and r2 of the inertial navigation device, record the distance from the point r1 to the aircraft horizontal plane as Z1 and the distance to the aircraft longitudinal section as Y1, record the distance from the point r2 to the aircraft horizontal plane as Z2 and the distance to the aircraft longitudinal section as Y2; measure and obtain the pitch installation hole points p1 and p2 of the inertial navigation device, record the distance from the point p1 to the aircraft horizontal plane as Z1 ′ , and the distance to the aircraft cross section as X1, record the distance from the point p2 to the aircraft horizontal plane as Z2 ′ , and the distance to the aircraft cross section as X2; measure and obtain the heading installation hole points y1 and y2 of the inertial navigation device, record the distance from the point y1 to the aircraft longitudinal section as Y1 ′ , and the distance to the aircraft cross section as X1′ the distance from point y2 to the plane of the aircraft is Y2 ′ the distance to the cross section of the aircraft is X2 ′ ; obtain the roll angle error α = arctan (ΔZ1 / ΔY1), when Z1>Z2, ΔZ1=Z1-Z2; otherwise Z1 obtain the pitch angle error β = arctan (ΔZ2 / ΔX1), when Z1 ′ >Z2 ′ , ΔZ2=Z1 ′ -Z2 ′ ; otherwise Z1 ′ ′ >Z2 ′ , ΔZ2=Z2 ′ -Z1 ′ ; when X1>X2, ΔX1=X1-X2; otherwise X1 ′ <X2, ΔX1=X2-X1; obtain the heading angle error δ = arctan (ΔY2 / ΔX2), when Y1 ′ ′ ′ ′ ′ ′ ′ ′ ′ ′ ′ ′ ′ ′ .
[0016] Optionally, the positive or negative of the roll angle error α is determined according to Z1>Z2 or Z2>Z1; the positive or negative of the pitch angle error β is determined according to Z1 ′ >Z2 ′ or Z2 ′ >Z1 ′ ; the positive or negative of the heading angle error δ is determined according to Y1 ′ ′ ′ ′ .
[0017] Optionally, the measurement module is specifically configured to: In the measurement coordinate system, the coordinates of four measurement points of the navigation component are measured by using the laser tracker; the four measurement points include: two points in front of the aircraft along the heading and two points behind the aircraft along the heading; the distance from measurement point 1 to the cross section b of the aircraft is denoted as x1, the distance from measurement point 1 to the longitudinal section c of the aircraft is denoted as y1, the distance from measurement point 1 to the horizontal plane a of the aircraft is denoted as z1, the distance from measurement point 2 to the cross section b of the aircraft is denoted as x2, the distance from measurement point 2 to the longitudinal section c of the aircraft is denoted as y2, the distance from measurement point 2 to the horizontal plane of the aircraft is denoted as z2, the distance from measurement point 3 to the cross section b of the aircraft is denoted as x3, the distance from measurement point 3 to the longitudinal section c of the aircraft is denoted as y3, the distance from measurement point 3 to the horizontal plane of the aircraft is denoted as z3, the distance from measurement point 4 to the cross section b of the aircraft is denoted as x4, the distance from measurement point 4 to the longitudinal section c of the aircraft is denoted as y4, and the distance from measurement point 4 to the horizontal plane of the aircraft is denoted as z4; When (z1+z4)>(z2+z3), the roll angle error α=60arctan(Δz1 / Δy) is obtained according to Δz1=(z1+z4-z2-z3) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is positive; when (z1+z4)<(z2+z3), the roll angle error α=60arctan(Δz2 / Δy) is obtained according to Δz2=(z2+z3-z1-z4) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is negative; When (z1+z2)>(z3+z4), the pitch angle error β=60arctan(Δz3 / Δx) is obtained according to Δz3=(z1+z2-z3-z4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is positive; when (z1+z2)<(z3+z4), the pitch angle error β=60arctan(Δz4 / Δx) is obtained according to Δz4=(z3+z4-z1-z2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is negative; When (y1+y2)>(y3+y4), the heading angle error δ=60arctan(Δy2 / Δx) is obtained according to Δy2=(y1+y2-y3-y4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is positive; when (y1+y2)<(y3+y4), the heading angle error δ=60arctan(Δy3 / Δx) is obtained according to Δy3=(y3+y4-y1-y2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is negative.
[0018] Optionally, the standard establishing module is specifically configured to establish a calibration measurement coordinate system according to reference measurement points on the lateral roll axis, the pitch axis and the heading axis of the aircraft; the calibration measurement coordinate system comprises mutually perpendicular lateral roll axis, pitch axis and heading axis; under the calibration measurement coordinate system, an aircraft horizontal plane a is established by points on the horizontal plane of the aircraft, an aircraft longitudinal section c is established by a Y-axis connecting line of the aircraft lateral roll axis and perpendicular to the horizontal plane, and an aircraft cross section b is established by an X-axis point of the pitch axis and perpendicular to the aircraft longitudinal section.
[0019] The application provides an inertial navigation device installation calibration method and device, and relates to an inertial navigation system installation calibration process. Compared with the traditional inertial navigation device calibration method, the application omits the installation of an adapter plate and the use of a flat plate, simplifies the operation steps, and saves calibration time. The application establishes a coordinate system by using external reference points of an aircraft, is not limited by structural differences of the aircraft, can adapt to different structural measurement conditions, can be applied to inertial navigation system measurement of various types of platforms, and greatly improves the universality of the aircraft. The application uses spatial geometric calculation, provides valuable experience for calibration work of other systems, and has a far-reaching influence on subsequent calibration methods. BRIEF DESCRIPTION OF DRAWINGS
[0020] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or the prior art description. Obviously, the drawings in the following description are some embodiments of the application, and for those skilled in the art, other drawings can also be obtained without creative labor.
[0021] Figure 1 Fig. 1 is a schematic diagram of an aircraft attitude axis; Figure 2 Fig. 2 is a schematic diagram of an inertial navigation assembly calibration adapter plate; Figure 3 Fig. 3 is a schematic diagram of a flat plate calibration measurement method; Figure 4 Fig. 4 is a schematic diagram of an inertial navigation device installation point measurement of the application; Figure 5 Fig. 5 is a flowchart of an inertial navigation device installation calibration technical method of the application; MARKED DESCRIPTION: 11-aircraft heading axis, 12-aircraft lateral roll axis, 13-aircraft pitch axis; 21 - adapter plate, 22 - pitch reference line, 23 - roll reference line, 24 - heading reference hole, 25 - aircraft heading normal; 31 - measurement plate, 32 - aircraft attitude heading coordinate axis, 33 - inertial navigation assembly, 34 - target mirror mounting position, 35 - target mirror center line, 36 - aircraft heading normal, 37 - heading angle error, 38 - distance L between target mirror center and plate center line, 39 - distance between target mirror and measurement plate; 41 - reference coordinate system pitch axis, 42 - reference coordinate system heading axis, 43 - reference coordinate system roll axis, 44 - inertial measurement assembly mounting point, 45 - distance between inertial measurement assembly mounting point and horizontal plane, 46 - distance between inertial measurement assembly mounting point and longitudinal section, 47 - distance between inertial measurement assembly mounting point and transverse section, b - aircraft transverse section, a - aircraft horizontal plane, c - aircraft longitudinal section. DETAILED DESCRIPTION
[0022] The inertial navigation device installation calibration method and device provided by the present application will be described in further detail below with reference to the accompanying drawings.
[0023] The present application is distinguished from conventional inertial navigation installation calibration technology in that a coordinate system measurement method is adopted, reference points on the aircraft are first collected to establish a coordinate system, the inertial navigation device mounting point is then measured, and finally the installation attitude angle error is obtained through geometric calculation.
[0024] To make the purpose of the implementation of the present application and the advantages of the technical solutions more clear, the technical solutions of the present application will be described in more detail below by introducing the content of an inertial navigation system component installation calibration work.
[0025] As shown in Figure 4 and Figure 5 , the present application provides an inertial navigation device installation calibration method, comprising: Step 1: establishing a calibration measurement coordinate system according to the directions of the reference measurement points on the reference coordinate system roll axis 43, the reference coordinate system pitch axis 41 and the reference coordinate system heading axis 42; Step 2: establishing the aircraft horizontal plane a through the points on the aircraft horizontal plane under the calibration measurement coordinate system, establishing the aircraft longitudinal section c through the Y-axis line and perpendicular to the horizontal plane, and establishing the aircraft transverse section b through the X-axis point and perpendicular to the aircraft longitudinal section; Step 3: under the calibration measurement coordinate system: measuring and obtaining the inertial measurement assembly mounting point 44, obtaining the distance 45 between the inertial measurement assembly mounting point and the horizontal plane, the distance 46 between the inertial measurement assembly mounting point and the longitudinal section, and the distance 47 between the inertial measurement assembly mounting point and the transverse section; Exemplarily, in a certain specific embodiment; Measure and obtain the inertial navigation device roll mounting hole point r1, point r2, the distance from point r1 to the plane of the aircraft is recorded as Z1, and the distance to the longitudinal section of the aircraft is recorded as Y1, the distance from point r2 to the plane of the aircraft is recorded as Z2, and the distance to the longitudinal section of the aircraft is recorded as Y2; Measure and obtain the inertial navigation device pitch mounting hole point p1, point p2, the distance from point p1 to the plane of the aircraft is recorded as Z1 ′ , and the distance to the transverse section of the aircraft is recorded as X1, the distance from point p2 to the plane of the aircraft is recorded as Z2 ′ , and the distance to the transverse section of the aircraft is recorded as X2, Measure and obtain the inertial navigation device heading mounting hole point y1, point y2, the distance from point y1 to the longitudinal section of the aircraft is recorded as Y1 ′ , and the distance to the transverse section of the aircraft is recorded as X1 ′ , the distance from point y2 to the longitudinal section of the aircraft is recorded as Y2 ′ , and the distance to the transverse section of the aircraft is recorded as X2 ′ , as Figure 4 .
[0026] Step four: after the above measurement, the inertial navigation device roll angle, pitch angle, heading angle error calculation is carried out according to the trigonometric function, the detailed method is as follows: 1. Roll angle error calculation When Z1>Z2, ΔZ1=Z1-Z2; otherwise Z1 < Y2, ΔY1=Y2-Y1; Roll angle error α=arctan(ΔZ1 / ΔY1), according to Z1>Z2 or Z2>Z1 to define the sign of positive or negative.
[0027] 2. Pitch angle error calculation When Z1 ′ >Z2 ′ , ΔZ2=Z1 ′ -Z2 ′ ; otherwise Z1 ′ ′ < Z2 ′ , ΔZ2=Z2 ′ -Z1 ′ ; When X1>X2, ΔX1=X1-X2; otherwise X1 Pitch angle error β=arctan(ΔZ2 / ΔX1), according to Z1 ′ >Z2 ′ or Z2 ′ >Z1 ′ to define the sign of positive or negative.
[0028] 3. Heading angle error calculation When Y1 ′ > Y2 ′ , ΔY2 = Y1 ′ - Y2 ′ ; otherwise Y1 ′ < Y2 ′ , ΔY2 = Y2 ′ - Y1 ′ ; When X1 ′ > X2 ′ , ΔX2 = X1 ′ - X2 ′ ; otherwise X1 ′ < X2 ′ , ΔX2 = X2 ′ - X1 ′ ; Heading angle error δ = arctan(ΔY2 / ΔX2), according to Y1 ′ > Y2 ′ or Y2 ′ > Y1 ′ , define the sign of the symbol positive or negative.
[0029] It can be understood that in actual measurement, Z1 and Z2, Y1 and Y2, X1 and X2 in the above formula must exist errors, and the case of the same value is ignored.
[0030] Step five: the finally calculated roll angle error α, pitch angle error β, heading angle error δ input into the system for correcting the consistency of the inertial navigation device and the aircraft attitude. See Figure 5 .
[0031] The inertial navigation system component is a cuboid structure, which is installed on the aircraft equipment frame plane by four mounting points. The specific implementation steps of the calibration method are as follows: Step one: set up reference measurement points in the direction of the heading axis, roll axis and pitch axis outside the aircraft, and measure the coordinate data of the reference measurement points using a laser tracker to establish a coordinate system.
[0032] Step two: select three points 1, 2, 3 on the aircraft structure on the aircraft horizontal plane and points 6, 7 on the center line of the aircraft Y direction. Measure the coordinate data of the above five points using a laser tracker. Select points 1, 2, 3 to establish the aircraft horizontal plane a. Draw the aircraft longitudinal section c perpendicular to the aircraft horizontal plane through points 6, 7, and draw the aircraft transverse section b perpendicular to the longitudinal section through point 6.
[0033] Step three: In the measurement coordinate system, use the laser tracker to measure the coordinate data of the four mounting points 1, 2, 3, and 4 of the component. Among them, points 1 and 2 are the first two points in the forward direction of the aircraft, and points 3 and 4 are the last two points in the forward direction of the aircraft.
[0034] The distance from measuring point 1 to the aircraft cross section b is denoted as x1, the distance from measuring point 1 to the aircraft longitudinal section c is denoted as y1, and the distance from measuring point 1 to the aircraft horizontal plane a is denoted as z1. The distance from measuring point 2 to the cross section b is denoted as x2, the distance from measuring point 2 to the aircraft longitudinal section c is denoted as y2, and the distance from measuring point 2 to the aircraft horizontal plane is denoted as z2. The distance from measuring point 3 to the cross section b is denoted as x3, the distance from measuring point 3 to the aircraft longitudinal section c is denoted as y3, and the distance from measuring point 3 to the aircraft horizontal plane is denoted as z3. The distance from measuring point 4 to the cross section b is denoted as x4, the distance from measuring point 4 to the aircraft longitudinal section c is denoted as y4, and the distance from measuring point 4 to the aircraft horizontal plane is denoted as z4.
[0035] Step four: Through the above measurement process, the coordinate data of the four mounting points of the inertial measurement assembly is obtained. Since there are two roll, pitch and heading angle errors in the four mounting points, the coordinate values of the two mounting hole points in the roll, pitch and heading directions are obtained by averaging the two values, and the angle error values in each direction are calculated. According to the trigonometric function calculation method as follows: 1. Roll angle error calculation When (z1+z4)>(z2+z3), Δz1=(z1+z4-z2-z3) / 2, Δy=(y2+y3-y1-y4) / 2 Roll angle error α=60arctan(Δz1 / Δy), the self-defined symbol is positive; When (z1+z4)<(z2+z3), Δz2=(z2+z3-z1-z4) / 2, Δy=(y2+y3-y1-y4) / 2 Roll angle error α=60arctan(Δz2 / Δy), the self-defined symbol is negative; 2. Pitch error calculation When (z1+z2)>(z3+z4), Δz3=(z1+z2-z3-z4) / 2, Δx=(x3+x4-x1-x2) / 2 Pitch angle error β=60arctan(Δz3 / Δx), the self-defined symbol is positive; When (z1+z2)<(z3+z4), Δz4=(z3+z4-z1-z2) / 2, Δx=(x3+x4-x1-x2) / 2 Pitch angle error β=60arctan(Δz4 / Δx), the self-defined symbol is negative; 3. Heading angle error calculation When (y1+y2)>(y3+y4), Δy2=(y1+y2-y3-y4) / 2, Δx=(x3+x4-x1-x2) / 2 Course error δ=60arctan(Δy2 / Δx), the sign is positive by definition. When (y1+y2)<(y3+y4), Δy3=(y3+y4-y1-y2) / 2, Δx=(x3+x4-x1-x2) / 2 Course error δ=60arctan(Δy3 / Δx), the sign is negative by definition. Step five: input the calculated roll angle error α, pitch angle error β, course angle error δ into the system to correct the consistency of the inertial navigation device.
[0036] The present application relates to a kind of inertial navigation system installation calibration method, using laser tracker selects aircraft outer reference point to establish measurement coordinate system and constructs reference plane. The coordinate system is measured on the measurement point coordinate of the direction of motion of inertial navigation device, by geometric analysis, the required direction of motion angle error is calculated using trigonometric function, to bind compensation inertial navigation component installation error value. The present application is more accurate compared with traditional method, simple, measurement time is short, work efficiency is high, environmental adaptability is strong, improve the universality of method, simultaneously required equipment is less, field maintenance is convenient, can be applicable to other system calibration work, with extensive popularization significance.
Claims
1. An inertial navigation device installation calibration method, characterized by, The method comprises the following steps: S1, establishing a calibration measurement coordinate system and an aircraft horizontal plane a, a longitudinal section c and a transverse section b; S2, measuring the coordinates of the measurement points in the direction of motion of the inertial navigation device in the calibration measurement coordinate system, and obtaining the roll angle error α, the pitch angle error β and the heading angle error δ according to the coordinates of the measurement points in the direction of motion of the inertial navigation device; S3, correcting the inertial navigation device according to the obtained roll angle error α, the pitch angle error β and the heading angle error δ.
2. The method of claim 1, wherein, The method for measuring the coordinates of the measurement points in the direction of motion of the inertial navigation device and obtaining the roll angle error α, the pitch angle error β and the heading angle error δ according to the coordinates of the measurement points in the direction of motion of the inertial navigation device comprises the following steps: measuring and obtaining the roll installation hole points r1 and r2 of the inertial navigation device, recording the distance between the point r1 and the aircraft horizontal plane as Z1, the distance between the point r1 and the aircraft longitudinal section as Y1, recording the distance between the point r2 and the aircraft horizontal plane as Z2, and the distance between the point r2 and the aircraft longitudinal section as Y2; Measure and obtain the pitch mounting hole points p1, p2 of the inertial navigation device, the distance from point p1 to the horizontal plane of the aircraft is recorded as Z1 ′ , the distance to the cross section of the aircraft is recorded as X1, the distance from point p2 to the horizontal plane of the aircraft is recorded as Z2 ′ , and the distance to the cross section of the aircraft is recorded as X2; Measure and obtain the inertial navigation device heading mounting hole points y1, y2, the distance from point y1 to the aircraft longitudinal section is Y1 ′ , the distance to the aircraft transverse section is X1 ′ , the distance from point y2 to the aircraft horizontal plane is Y2 ′ , the distance to the aircraft transverse section is X2 ′ ; obtaining the roll angle error α = arctan(ΔZ1 / ΔY1), when Z1 > Z2, ΔZ1 = Z1 - Z2; when Z1 < Z2, ΔZ1 = Z2 - Z1; when Y1 > Y2, ΔY1 = Y1 - Y2; when Y1 < Y2, ΔY1 = Y2 - Y1; The pitch angle error β = arctan(ΔZ2 / ΔX1) is obtained, when Z1 ′ > Z2 ′ , ΔZ2 = Z1 ′ - Z2 ′ ; otherwise, when Z1 ′ < Z2 ′ , ΔZ2 = Z2 ′ - Z1 ′ ; when X1 > X2, ΔX1 = X1 - X2; otherwise, when X1 < X2, ΔX1 = X2 - X1; The heading angle error δ = arctan(ΔY2 / ΔX2) is obtained when Y1 ′ > Y2 ′ , ΔY2 = Y1 ′ - Y2 ′ ; otherwise Y1 ′ < Y2 ′ , ΔY2 = Y2 ′ - Y1 ′ ; when X1 ′ > X2 ′ , ΔX2 = X1 ′ - X2 ′ ; otherwise X1 ′ < X2 ′ , ΔX2 = X2 ′ - X1 ′ .
3. The method of inertial navigation unit installation calibration according to claim 2, wherein, the positive and negative of the roll angle error α is determined according to Z1 > Z2 or Z2 > Z1; the sign of the pitch error β depends on Z1 ′ > Z2 ′ or Z2 ′ > Z1 ′ determination; The positive or negative of the course angle error δ is determined according to Y1 ′ > Y2 ′ or Y2 ′ > Y1 ′ is determined.
4. The method of claim 1, wherein, The method for measuring the coordinates of the measurement points in the direction of motion of the inertial navigation device and obtaining the roll angle error α, the pitch angle error β and the heading angle error δ according to the coordinates of the measurement points in the direction of motion of the inertial navigation device comprises the following steps: under the measurement coordinate system, using a laser tracker to measure the coordinate data of four measurement points of the inertial navigation component; the four measurement points comprise two front points along the heading direction of the aircraft and two rear points along the heading direction of the aircraft; the distance between the measurement point 1 and the transverse section b of the aircraft is recorded as x1, the distance between the measurement point 1 and the longitudinal section c of the aircraft is recorded as y1, the distance between the measurement point 1 and the horizontal plane a of the aircraft is recorded as z1, the distance between the measurement point 2 and the transverse section b of the aircraft is recorded as x2, the distance between the measurement point 2 and the longitudinal section c of the aircraft is recorded as y2, the distance between the measurement point 2 and the horizontal plane of the aircraft is recorded as z2, the distance between the measurement point 3 and the transverse section b of the aircraft is recorded as x3, the distance between the measurement point 3 and the longitudinal section c of the aircraft is recorded as y3, the distance between the measurement point 3 and the horizontal plane of the aircraft is recorded as z3, the distance between the measurement point 4 and the transverse section b of the aircraft is recorded as x4, the distance between the measurement point 4 and the longitudinal section c of the aircraft is recorded as y4, and the distance between the measurement point 4 and the horizontal plane of the aircraft is recorded as z4; when (z1 + z4) > (z2 + z3), the roll angle error α = 60arctan(Δz1 / Δy) is obtained according to Δz1 = (z1 + z4 - z2 - z3) / 2 and Δy = (y2 + y3 - y1 - y4) / 2, and the sign of α is positive at this time; when (z1 + z4) < (z2 + z3), the roll angle error α = 60arctan(Δz2 / Δy) is obtained according to Δz2 = (z2 + z3 - z1 - z4) / 2 and Δy = (y2 + y3 - y1 - y4) / 2, and the sign of α is negative at this time; When (z1+z2)>(z3+z4), the pitch angle error β=60arctan(Δz3 / Δx) is obtained according to Δz3=(z1+z2-z3-z4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is positive; when (z1+z2)<(z3+z4), the pitch angle error β=60arctan(Δz4 / Δx) is obtained according to Δz4=(z3+z4-z1-z2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is negative. When (y1+y2)>(y3+y4), the heading angle error δ=60arctan(Δy2 / Δx) is obtained according to Δy2=(y1+y2-y3-y4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is positive; when (y1+y2)<(y3+y4), the heading angle error δ=60arctan(Δy3 / Δx) is obtained according to Δy3=(y3+y4-y1-y2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is negative.
5. The method of claim 1, wherein, S1 comprises: S11, establishing a calibration measurement coordinate system according to reference measurement points in the directions of the lateral roll axis, the pitch axis and the heading axis of the aircraft; the calibration measurement coordinate system comprises mutually perpendicular lateral roll axis, pitch axis and heading axis; S12, establishing an aircraft horizontal plane a through points on the horizontal plane of the aircraft, establishing an aircraft longitudinal section c through the Y-axis connecting line of the aircraft lateral roll axis and perpendicular to the horizontal plane, and establishing an aircraft transverse section b through the X-axis point of the pitch axis and perpendicular to the aircraft longitudinal section.
6. An inertial navigation device installation calibration apparatus, characterized by, Comprise: a standard establishment module for establishing a calibration measurement coordinate system and an aircraft horizontal plane a, a longitudinal section c and a transverse section b; a measurement module for measuring the coordinates of measurement points in the motion direction of the inertial navigation device in the calibration measurement coordinate system, and obtaining a roll angle error α, a pitch angle error β and a heading angle error δ according to the coordinates of the measurement points in the motion direction of the inertial navigation device; S3, correcting the inertial navigation device according to the obtained roll angle error α, pitch angle error β and heading angle error δ.
7. An inertial navigation device installation calibration apparatus according to claim 6, wherein, The measurement module is specifically used for: measuring and obtaining the roll installation hole points r1 and r2 of the inertial navigation device, recording the distance from point r1 to the aircraft horizontal plane as Z1 and the distance to the aircraft longitudinal section as Y1, recording the distance from point r2 to the aircraft horizontal plane as Z2 and the distance to the aircraft longitudinal section as Y2; Measure and obtain the pitch mounting hole points p1, p2 of the inertial navigation device, the distance from point p1 to the horizontal plane of the aircraft is recorded as Z1 ′ , the distance to the cross section of the aircraft is recorded as X1, the distance from point p2 to the horizontal plane of the aircraft is recorded as Z2 ′ , and the distance to the cross section of the aircraft is recorded as X2; Measure and obtain the inertial navigation device heading mounting hole points y1, y2, the distance from point y1 to the aircraft longitudinal section is Y1 ′ , the distance to the aircraft transverse section is X1 ′ , the distance from point y2 to the aircraft horizontal plane is Y2 ′ , the distance to the aircraft transverse section is X2 ′ ; obtaining a roll angle error α=arctan(ΔZ1 / ΔY1), when Z1>Z2, ΔZ1=Z1-Z2; when Z1Z2, ΔY1=Y2-Y1; The pitch angle error β = arctan(ΔZ2 / ΔX1) is obtained, when Z1 ′ > Z2 ′ , ΔZ2 = Z1 ′ - Z2 ′ ; otherwise, when Z1 ′ < Z2 ′ , ΔZ2 = Z2 ′ - Z1 ′ ; when X1 > X2, ΔX1 = X1 - X2; otherwise, when X1 < X2, ΔX1 = X2 - X1; The heading angle error δ = arctan (ΔY2 / ΔX2) is obtained when Y1 ′ > Y2 ′ , ΔY2 = Y1 ′ - Y2 ′ ; otherwise Y1 ′ < Y2 ′ , ΔY2 = Y2 ′ - Y1 ′ ; when X1 ′ > X2 ′ , ΔX2 = X1 ′ - X2 ′ ; otherwise X1 ′ < X2 ′ , ΔX2 = X2 ′ - X1 ′ .
8. An inertial navigation device installation calibration apparatus according to claim 7, wherein, the positive and negative of the roll angle error α is determined according to Z1>Z2 or Z2>Z1; the sign of the pitch error β depends on Z1 ′ > Z2 ′ or Z2 ′ > Z1 ′ determined; The positive or negative of the course angle error δ is determined according to Y1 ′ > Y2 ′ or Y2 ′ > Y1 ′ is determined.
9. The inertial navigation device installation calibration apparatus of claim 6, wherein, the measurement module is specifically used for: In the measurement coordinate system, the coordinates of the four measurement points of the navigation component are measured by using a laser tracker; the four measurement points include: the first two points in the forward direction of the aircraft and the last two points in the rear direction of the aircraft; the distance from the measurement point 1 to the cross section b of the aircraft is denoted as x1, the distance from the measurement point 1 to the longitudinal section c of the aircraft is denoted as y1, the distance from the measurement point 1 to the horizontal plane a of the aircraft is denoted as z1, the distance from the measurement point 2 to the cross section b is denoted as x2, the distance from the measurement point 2 to the longitudinal section c of the aircraft is denoted as y2, the distance from the measurement point 2 to the horizontal plane of the aircraft is denoted as z2, the distance from the measurement point 3 to the cross section b is denoted as x3, the distance from the measurement point 3 to the longitudinal section c of the aircraft is denoted as y3, the distance from the measurement point 3 to the horizontal plane of the aircraft is denoted as z3, the distance from the measurement point 4 to the cross section b is denoted as x4, the distance from the measurement point 4 to the longitudinal section c of the aircraft is denoted as y4, and the distance from the measurement point 4 to the horizontal plane of the aircraft is denoted as z4; When (z1+z4)>(z2+z3), the roll angle error α=60arctan(Δz1 / Δy) is obtained according to Δz1=(z1+z4-z2-z3) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is positive; when (z1+z4)<(z2+z3), the roll angle error α=60arctan(Δz2 / Δy) is obtained according to Δz2=(z2+z3-z1-z4) / 2 and Δy=(y2+y3-y1-y4) / 2, and the sign of α is negative; When (z1+z2)>(z3+z4), the pitch angle error β=60arctan(Δz3 / Δx) is obtained according to Δz3=(z1+z2-z3-z4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is positive; when (z1+z2)<(z3+z4), the pitch angle error β=60arctan(Δz4 / Δx) is obtained according to Δz4=(z3+z4-z1-z2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of β is negative; When (y1+y2)>(y3+y4), the heading angle error δ=60arctan(Δy2 / Δx) is obtained according to Δy2=(y1+y2-y3-y4) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is positive; when (y1+y2)<(y3+y4), the heading angle error δ=60arctan(Δy3 / Δx) is obtained according to Δy3=(y3+y4-y1-y2) / 2 and Δx=(x3+x4-x1-x2) / 2, and the sign of δ is negative.
10. The inertial navigation device installation calibration apparatus of claim 6, wherein, The standard establishment module is specifically configured to establish a calibration measurement coordinate system according to the reference measurement points in the directions of the roll axis, the pitch axis and the heading axis of the aircraft; the calibration measurement coordinate system includes mutually perpendicular roll axis, pitch axis and heading axis; in the calibration measurement coordinate system, the aircraft horizontal plane a is established through a point on the aircraft horizontal plane, the aircraft longitudinal section c is established through a line of the roll axis Y axis and perpendicular to the horizontal plane, and the aircraft cross section b is established through a point of the pitch axis X axis and perpendicular to the aircraft longitudinal section.