Attitude angle control system and method of carrier rocket

Abstract

The invention provides an attitude angle control system and method for a carrier rocket, and relates to the technical field of carrier rocket attitude measurement, and the system comprises a first antenna which is vertically installed close to the head of the carrier rocket and is used for determining a first coordinate; a second antenna vertically mounted away from the head of the carrier rocket for determining a second coordinate; the inertial measurement unit is arranged in the carrier rocket and is used for providing an inertial measurement result; the processor is mounted in the carrier rocket and used for acquiring the first coordinates and / or the second coordinates and determining an attitude angle calculation result of the carrier rocket based on the first coordinates and the second coordinates; and the processor is also used for acquiring the inertial measurement result and determining a target attitude angle of the carrier rocket based on the attitude angle calculation result and / or the inertial measurement result.

Description

Technical Field

[0001] The present invention relates to the technical field of attitude measurement of launch vehicles, and particularly to an attitude angle control system and method for a launch vehicle. Background Art

[0002] The launch vehicle market is showing a booming trend. With the diversified development of demands, maintaining high attitude accuracy and controlling costs during long-term flights is the future development trend of the attitude determination strategy of launch vehicles. In addition, to meet the requirements of rapid response, the rocket needs to have the ability to complete self-alignment in a short time.

[0003] Currently, during the flight of a launch vehicle, the motion state of the rocket body mainly comes from the changes in the speed and angle of the carrier sensed by the inertial measurement unit, obtaining navigation and positioning information including motion speed, attitude angle changes, and the position where it is located, etc. However, the calculation principle of this method causes the error to accumulate over time, resulting in a continuous decline in the navigation and positioning and attitude determination accuracy after the long-term flight of the launch vehicle. Summary of the Invention

[0004] The present invention provides an attitude angle control system and method for a launch vehicle, which are used to solve the problem of low accuracy in attitude angle control of a launch vehicle by an inertial measurement unit in related technologies.

[0005] In a first aspect, an embodiment of the present invention provides an attitude angle control system for a launch vehicle, and the system includes:

[0006] A first antenna vertically installed near the head of the launch vehicle, which is used to determine a first coordinate, where the first coordinate is the coordinate of a first positioning center in the geocentric coordinate system, and the first positioning center is the positioning center on the first antenna;

[0007] A second antenna vertically installed away from the head of the launch vehicle, which is used to determine a second coordinate, where the first antenna and the second antenna are both located on the target longitudinal axis, the target longitudinal axis is parallel to the longitudinal axis of the rocket body of the launch vehicle, the second coordinate is the coordinate of a second positioning center in the geocentric coordinate system, and the second positioning center is the positioning center on the second antenna;

[0008] An inertial measurement unit installed in the launch vehicle, which is used to provide inertial measurement results;

[0009] A processor installed in the launch vehicle, which is used to obtain the first coordinate and / or the second coordinate, and is used to determine the attitude angle calculation result of the launch vehicle based on the first coordinate and the second coordinate;

[0010] The processor is further configured to obtain the inertial measurement result, and determine the target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial measurement result.

[0011] In a second aspect, an embodiment of the present invention provides an attitude angle control method for a launch vehicle, which is applied to the system described in the first aspect. The method includes:

[0012] Obtain the first coordinate and / or the second coordinate;

[0013] Based on the first coordinate and / or the second coordinate, determine the attitude angle calculation result of the launch vehicle, where the attitude angle calculation result includes the heading angle calculation result and / or the pitch angle calculation result of the launch vehicle;

[0014] Obtain the inertial navigation result;

[0015] Based on the attitude angle calculation result and / or the inertial navigation result, determine the target attitude angle of the launch vehicle.

[0016] Optionally, the determining the attitude angle calculation result of the launch vehicle based on the first coordinate and / or the second coordinate includes:

[0017] Based on the first coordinate and / or the second coordinate, determine a third coordinate, where the third coordinate is the coordinate of the target baseline vector in the geocentric coordinate system, and the target baseline vector is the baseline vector between the first positioning center and the second positioning center;

[0018] Convert the third coordinate to a fourth coordinate, where the fourth coordinate is the coordinate of the target baseline vector in the local geodetic coordinate system, and the origin of the local geodetic coordinate system is the centroid position of the launch vehicle;

[0019] Based on the fourth coordinate, determine the attitude angle calculation result of the launch vehicle.

[0020] Optionally, the converting the third coordinate to the fourth coordinate includes:

[0021] Obtain the longitude and / or latitude of the second positioning center;

[0022] Based on the longitude and / or latitude of the second positioning center, convert the third coordinate to the fourth coordinate through a coordinate transformation matrix.

[0023] Optionally, the determining the attitude angle calculation result of the launch vehicle based on the fourth coordinate includes:

[0024] Based on the fourth coordinate, use trigonometric function relationships to determine the attitude angle calculation result of the launch vehicle.

[0025] Optionally, determining the attitude angle calculation result of the launch vehicle based on the fourth coordinate by using trigonometric function relationships includes:

[0026] Determining the attitude angle calculation result of the launch vehicle through the following formula:

[0027]

[0028] where ψ is the calculated result of the course angle of the launch vehicle, φ is the calculated result of the pitch angle of the launch vehicle, x is the value of the fourth coordinate on the X-axis, y is the value of the fourth coordinate on the Y-axis, and z is the value of the fourth coordinate on the Z-axis.

[0029] Optionally, determining the target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial navigation result includes:

[0030] Obtaining the state equation of the launch vehicle, where the state equation characterizes the motion state of the launch vehicle;

[0031] Based on the attitude angle calculation result and / or the inertial navigation result, constructing the observation equation of the launch vehicle;

[0032] Performing Kalman filter solution according to the state equation and the observation equation to determine the target attitude angle of the launch vehicle.

[0033] Optionally, constructing the observation equation of the launch vehicle based on the attitude angle calculation result and / or the inertial navigation result includes:

[0034] Performing lever arm correction processing on the second coordinate according to the inertial navigation result to obtain the fifth coordinate;

[0035] Determining the calculated attitude result in the local geodetic coordinate system according to the attitude angle calculation result;

[0036] Constructing the observation equation of the launch vehicle according to the fifth coordinate and the calculated attitude result.

[0037] Optionally, the observation equation includes:

[0038]

[0039] where Z is the observation result, is the second coordinate, is the predicted value of the second positioning center obtained according to the inertial navigation result, is the calculated attitude result, For the inertial navigation result, H is the observation equation matrix, x is the target attitude angle, and V is the observation error.

[0040] In a third aspect, an embodiment of the present invention provides a launch vehicle, including:

[0041] A launch vehicle body;

[0042] A system disposed on the launch vehicle body, and the system is the attitude angle control system of the launch vehicle as described in the first aspect.

[0043] An embodiment of the present invention provides an attitude angle control system for a launch vehicle, including a first antenna vertically installed near the head of the launch vehicle, which is used to determine a first coordinate, where the first coordinate is the coordinate of a first positioning center in the geocentric coordinate system, and the first positioning center is the positioning center on the first antenna; a second antenna vertically installed away from the head of the launch vehicle, which is used to determine a second coordinate, where both the first antenna and the second antenna are located on the target longitudinal axis, the target longitudinal axis is parallel to the longitudinal axis of the launch vehicle body, the second coordinate is the coordinate of a second positioning center in the geocentric coordinate system, and the second positioning center is the positioning center on the second antenna; an inertial measurement unit installed in the launch vehicle, which is used to provide an inertial measurement result; a processor installed in the launch vehicle, which is used to obtain the first coordinate and / or the second coordinate, and is used to determine the attitude angle calculation result of the launch vehicle based on the first coordinate and the second coordinate; the processor is further used to obtain the inertial measurement result, and based on the attitude angle calculation result and / or the inertial measurement result, determine the target attitude angle of the launch vehicle. In this way, the attitude angle calculation result of the launch vehicle can be determined by using the first antenna and / or the second antenna disposed on the launch vehicle, and the target attitude angle can be jointly calculated by combining the attitude angle calculation result and / or the inertial measurement result, so that the obtained target attitude angle has higher accuracy compared with the target attitude angle obtained based on the inertial measurement result, and to a certain extent, solves the problem of low accuracy of attitude angle control of the launch vehicle by the inertial measurement unit in the related art. Description of the Drawings

[0044] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0045] Figure 1 It is a conceptual diagram of an attitude angle control system for a launch vehicle provided by an embodiment of the present invention;

[0046] Figure 2 This is a flowchart of an attitude angle control method for a launch vehicle provided by an embodiment of the present invention;

[0047] Figure 3 This is a schematic diagram of a target baseline vector provided by an embodiment of the present invention;

[0048] Figure 4 This is a conceptual diagram of an attitude angle control method for a launch vehicle provided by an embodiment of the present invention. Detailed implementation manners

[0049] As described in the background art, the launch vehicle market is showing a booming trend. With the diversified development of demands, maintaining high attitude accuracy during long-term flight and controlling costs are the trends of the future attitude determination strategy for launch vehicles. And in order to meet the requirements of rapid response, the rocket needs to have the ability to complete self-alignment in a short time.

[0050] During the current flight of a launch vehicle, the motion state of the rocket body mainly comes from the changes in the speed and angle of the carrier sensed by the Inertial Navigation System (INS for short), and navigation and positioning information including motion speed, attitude angle change, and position is obtained. The inertial navigation system does not need to interact with the outside world and is not restricted by the surrounding environmental conditions, with good safety and robustness. However, due to its calculation principle, the error of this method accumulates over time, resulting in a continuous decline in navigation and positioning and attitude determination accuracy after the long-term flight of the launch vehicle.

[0051] An embodiment of the present invention provides an attitude angle control system for a launch vehicle, including a first antenna vertically installed near the head of the launch vehicle, configured to determine a first coordinate, where the first coordinate is the coordinate of a first positioning center in a geocentric coordinate system, and the first positioning center is the positioning center on the first antenna; a second antenna vertically installed away from the head of the launch vehicle, configured to determine a second coordinate, where both the first antenna and the second antenna are located on a target longitudinal axis, the target longitudinal axis is parallel to the longitudinal axis of the launch vehicle body, the second coordinate is the coordinate of a second positioning center in the geocentric coordinate system, and the second positioning center is the positioning center on the second antenna; an inertial measurement unit installed inside the launch vehicle, configured to provide an inertial measurement result; a processor installed inside the launch vehicle, configured to obtain the first coordinate and / or the second coordinate, and configured to determine an attitude angle calculation result of the launch vehicle based on the first coordinate and the second coordinate; the processor is further configured to obtain the inertial measurement result, and determine a target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial measurement result. In this way, the attitude angle calculation result of the launch vehicle can be determined by using the first antenna and / or the second antenna provided on the launch vehicle, and the target attitude angle is jointly calculated by combining the attitude angle calculation result and / or the inertial measurement result, so that the obtained target attitude angle has higher accuracy compared with the target attitude angle obtained based on the inertial measurement result, and to a certain extent solves the problem of low accuracy of attitude angle control of the launch vehicle by the inertial measurement unit in the related art.

[0052] Next, specific embodiments will be used to detail the technical solution of the present invention and how the technical solution of the present invention solves the above technical problems. These specific embodiments below can be combined with each other, and the same or similar concepts or processes may not be repeated in some embodiments. The embodiments of the present invention will be described below with reference to the drawings.

[0053] Figure 1 It is a conceptual diagram of an attitude angle control system for a launch vehicle provided by an embodiment of the present invention. As Figure 1 shown, the attitude angle control system for a launch vehicle provided by an embodiment of the present invention includes:

[0054] The first antenna is vertically installed near the head of the launch vehicle and is used to determine the first coordinate. Here, the first coordinate is the coordinate of the first positioning center in the geocentric coordinate system, and the first positioning center is the positioning center on the first antenna. The second antenna is vertically installed away from the head of the launch vehicle and is used to determine the second coordinate. Here, both the first antenna and the second antenna are located on the target longitudinal axis, and the target longitudinal axis is parallel to the longitudinal axis of the launch vehicle body. The second coordinate is the coordinate of the second positioning center in the geocentric coordinate system, and the second positioning center is the positioning center on the second antenna. The inertial measurement unit installed in the launch vehicle is used to provide inertial measurement results. The processor installed in the launch vehicle is used to obtain the first coordinate and / or the second coordinate, and is used to determine the attitude angle calculation result of the launch vehicle based on the first coordinate and the second coordinate. The processor is further used to obtain the inertial measurement result, and determine the target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial measurement result.

[0055] In an embodiment of the present invention, the first antenna and / or the second antenna may be related components of the Global Navigation Satellite System (GNSS) satellite navigation positioning technology. As a direct positioning technology based on radio signals, GNSS can provide high-precision timekeeping and absolute spatial position coordinates without being restricted by time and space. The navigation errors between each epoch (i.e., each observation time) are independent of each other, which can make up for the disadvantage that the inertial navigation error accumulates over time.

[0056] In an embodiment of the present invention, the geocentric coordinate system is the O-XYZ coordinate system established within the geoid, such as the WGS-84 coordinate system. The origin O is set at the mass center of the geoid and is represented by three mutually perpendicular axes X, Y, and Z. The X-axis coincides with the intersection line of the prime meridian plane and the equatorial plane, with the positive direction to the east; the Z-axis coincides with the earth's rotation axis, with the positive direction to the north; the Y-axis is perpendicular to the XZ plane to form a right-handed system. The processor can obtain the first coordinate and / or the second coordinate through GNSS. The first coordinate is used to describe the coordinate of the first positioning center on the first antenna in the geocentric coordinate system. The first positioning center can be at any position on the first antenna, such as the center point, etc. In actual situations, the specific position of the first positioning center can refer to the instruction manual of the first antenna, and the embodiments of the present invention do not make specific restrictions on this. Correspondingly, the second positioning center described by the second coordinate can also be at any position on the second antenna, and the embodiments of the present invention do not make specific restrictions on this.

[0057] In an embodiment of the present invention, the first antenna and / or the second antenna may both be located on the target longitudinal axis, that is, the baseline (target longitudinal axis) formed by the contact point of the first antenna with the launch vehicle and the contact point of the second antenna with the launch vehicle, and the target longitudinal axis is parallel to the longitudinal axis direction of the launch vehicle body. The longitudinal axis of the vehicle body is the axis pointing from the tail of the rocket body to the head of the rocket body. That is to say, the first antenna and / or the second antenna are perpendicular to the longitudinal axis of the vehicle body.

[0058] In an embodiment of the present invention, the inertial measurement unit installed in the launch vehicle may be any device capable of measuring the three-axis attitude angle and acceleration of the launch vehicle. The processor can directly obtain the inertial measurement results through the inertial measurement unit in the INS.

[0059] After the processor obtains the first coordinate and / or the second coordinate through GNSS technology, since the baseline formed by the contact point of the first antenna with the launch vehicle and the contact point of the second antenna with the launch vehicle is parallel to the direction of the launch vehicle body, the coordinates of the baseline parallel to the launch vehicle formed by the first antenna and / or the second antenna can be described by the first coordinate and the second coordinate, and since this baseline is parallel to the launch vehicle, the change in the attitude angle of the launch vehicle can be described, so as to obtain the calculation result of the attitude angle of the launch vehicle. After the processor obtains the calculation result of the attitude angle of the launch vehicle, it can combine with the inertial measurement results obtained by using the inertial measurement unit in the INS, for example, by means of Kalman filter solution, to jointly determine the target attitude angle of the launch vehicle, and control the launch vehicle to navigate according to the target attitude angle to complete the attitude angle control of the launch vehicle.

[0060] In this way, the calculation result of the attitude angle of the launch vehicle can be determined by using the first antenna and / or the second antenna provided on the launch vehicle, and the target attitude angle can be jointly calculated by combining the calculation result of the attitude angle and / or the inertial measurement result, so that the obtained target attitude angle has higher accuracy compared with the target attitude angle obtained based on the inertial measurement result, and to a certain extent solves the problem of low accuracy of attitude angle control of the launch vehicle by using the inertial measurement unit in the related art.

[0061] Figure 2 It is a flowchart of a method for controlling the attitude angle of a launch vehicle provided by an embodiment of the present invention. As Figure 2 shown, the method for controlling the attitude angle of a launch vehicle provided by an embodiment of the present invention includes steps 210 to 240. The method for controlling the attitude angle of a launch vehicle provided by an embodiment of the present invention can be executed by the processor in the attitude angle control system described above.

[0062] Step 210, obtain the first coordinate and / or the second coordinate.

[0063] In an embodiment of the present invention, the first coordinate may be the coordinate of the first positioning center in the geocentric coordinate system in the attitude angle control system described above. Correspondingly, the second coordinate may be the coordinate of the second positioning center in the geocentric coordinate system in the attitude angle control system described above.

[0064] The processor may perform single-point positioning on the first antenna and the second antenna through GNSS to respectively obtain the first coordinate and the second coordinate. It may also perform single-point positioning and relative positioning between the two antennas on the first antenna or the second antenna through GNSS to obtain the first coordinate and / or the second coordinate. After obtaining the first coordinate and / or the second coordinate, step 220 may be executed. For example, the second antenna far from the head of the launch vehicle may be used as the main antenna, and single-point positioning and relative positioning are performed on the main antenna to obtain the first coordinate and / or the second coordinate.

[0065] Step 220: Based on the first coordinate and / or the second coordinate, determine the attitude angle calculation result of the launch vehicle, where the attitude angle calculation result includes the heading angle calculation result and / or the pitch angle calculation result of the launch vehicle.

[0066] In an embodiment of the present invention, the attitude angle calculation result of the launch vehicle may represent the attitude angle of the launch vehicle at this time, and can be used to correct the attitude angle of the launch vehicle so that the launch vehicle runs along a predetermined ideal attitude angle. During the operation of the launch vehicle, multiple attitude angle parameters are involved. For example, Pitch Angle: The angle between the longitudinal axis of the rocket body and the reference plane (such as the horizontal plane of the local geodetic coordinate system, etc.), which is used to control the flight altitude and trajectory of the rocket. The change in the pitch angle can affect the lift direction of the rocket, thereby adjusting the flight trajectory. Yaw Angle: The angle between the projection of the longitudinal axis of the rocket body in the reference plane and the reference direction (such as the predetermined orbit reference direction, etc.), which is mainly used to control the flight direction of the rocket. The change in the yaw angle can ensure that the rocket flies along a predetermined route.

[0067] In an embodiment of the present invention, through the first coordinate and / or the second coordinate, the coordinates of the baseline parallel to the launch vehicle formed by the first antenna and / or the second antenna can be described by the first coordinate and the second coordinate. And since this baseline is parallel to the launch vehicle, the change in the attitude angle of the launch vehicle can be described, thereby obtaining the attitude angle calculation result of the launch vehicle.

[0068] Step 230: Obtain the inertial navigation result.

[0069] In an embodiment of the present invention, the processor may obtain the inertial navigation result through an inertial measurement unit in the launch vehicle.

[0070] Step 240: Determine the target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial navigation result.

[0071] In the embodiment of the present invention, the attitude angle calculation result and the inertial navigation result can be combined for integrated navigation through Kalman filtering to determine the target attitude angle of the launch vehicle. It can also be combined through a trained neural network model to judge the weighting value between the attitude angle calculation result and the inertial navigation result. In this way, the target attitude angle is calculated by combining the attitude angle calculation result and / or the inertial measurement result, so that the obtained target attitude angle has higher accuracy compared with the target attitude angle obtained based on the inertial measurement result, and to a certain extent solves the problem of low accuracy of attitude angle control of the launch vehicle through the inertial measurement unit in the related art.

[0072] In step 220 provided in the embodiment of the present invention, a specific implementation manner may include the following steps: Determine a third coordinate based on the first coordinate and / or the second coordinate, where the third coordinate is the coordinate of the target baseline vector in the geocentric coordinate system, and the target baseline vector is the baseline vector between the first positioning center and the second positioning center; Convert the third coordinate to a fourth coordinate, where the fourth coordinate is the coordinate of the target baseline vector in the local geodetic coordinate system, and the origin of the local geodetic coordinate system is the centroid position of the launch vehicle; Determine the attitude angle calculation result of the launch vehicle based on the fourth coordinate.

[0073] In the embodiment of the present invention, the origin of the local geodetic coordinate system can be the centroid of the launch vehicle, the Y-axis points to the geographic north, the X-axis points to the geographic east, and the Z-axis is perpendicular to the local reference ellipsoid and points to the sky.

[0074] In the embodiment of the present invention, relative positioning can be performed on the first antenna and the second antenna, and the relative positioning result can be resolved to obtain a third coordinate representing the position of the target baseline vector. Further, for the convenience of calculation, the third coordinate can be converted to a fourth coordinate through a coordinate transformation matrix, so that the target baseline vector can more accurately represent the attitude angle deviation of the launch vehicle, thereby facilitating the determination of the attitude angle calculation result of the launch vehicle.

[0075] In the embodiment of the present invention, the origin of the fourth coordinate is the rocket centroid. The direction deviation of the target baseline vector can be analogized to the deviation of the attitude angle of the launch vehicle, and the attitude angle calculation result of the launch vehicle can be obtained through trigonometric functions and the like.

[0076] In an embodiment of the present invention, during the process of converting the third coordinate to the fourth coordinate, the position of the second positioning center can be used for assisted conversion. Specifically, the longitude and / or latitude of the second positioning center are obtained; based on the longitude and / or latitude of the second positioning center, the third coordinate is converted to the fourth coordinate through a coordinate transformation matrix.

[0077] In an embodiment of the present invention, the processor can obtain the longitude and / or latitude of the second positioning center through GNSS technology. After obtaining the longitude and / or latitude of the second positioning center, the third coordinate can be converted to the fourth coordinate through a coordinate transformation matrix. A specific implementation manner of the coordinate transformation matrix can be shown as the following formula (1):

[0078]

[0079] where, X LLS represents the fourth coordinate in the local geodetic coordinate system and the position of the second positioning center in the local geodetic coordinate system, X WGS84 is the third coordinate in the geocentric coordinate system and the position of the second positioning center, λ is the longitude of the second positioning center, is the latitude of the second positioning center.

[0080] In an embodiment of the present invention, after determining the fourth coordinate, the position of the target baseline vector in the local geodetic coordinate system can be obtained. For example, Figure 3 as shown, the vector in the figure can be understood as the target baseline vector. Correspondingly, based on the fourth coordinate, the attitude angle calculation result of the launch vehicle can be determined using trigonometric function relationships, such as inverse trigonometric functions.

[0081] In Figure 3 the figure shown, the heading angle calculation and pitch angle calculation results of the launch vehicle can be determined respectively through the following formula (2):

[0082]

[0083] where, ψ is the heading angle calculation result, φ is the pitch angle calculation result, x is the X-axis value of the target baseline vector in the local geodetic coordinate system, y is the Y-axis value of the target baseline vector in the local geodetic coordinate system, and z is the Z-axis value of the target baseline vector in the local geodetic coordinate system.

[0084] In an embodiment of the present invention, after obtaining the attitude angle calculation result of the launch vehicle, the target attitude angle of the launch vehicle can be determined based on the attitude angle calculation result and / or the inertial navigation result. A specific implementation method includes the following steps: obtaining the state equation of the launch vehicle, where the state equation characterizes the motion state of the launch vehicle; constructing the observation equation of the launch vehicle based on the attitude angle calculation result and / or the inertial navigation result; and performing Kalman filter solution according to the state equation and the observation equation to determine the target attitude angle of the launch vehicle.

[0085] In an embodiment of the present invention, the state equation characterizes the motion state of the launch vehicle, where the parameters can be assumed according to the physical motion state of the rocket and the inertial navigation principle, and characterize the error between the finally calculated target attitude angle and the inertial navigation result. The state equation can be expressed by the following formula (3):

[0086]

[0087] where x is the motion state error of the launch vehicle; is the derivative of x;; G is the process noise matrix, w is the error of the gyroscope and accelerometer; F is the system dynamic matrix; is the inertial navigation position error; is the inertial navigation speed error, Φ is the inertial navigation attitude error, b g is the gyro zero bias; b a is the accelerometer zero bias; s g is the dropout scale factor error; s a is the accelerometer scale factor error.

[0088] Furthermore, in an embodiment of the present invention, in the process of constructing the observation equation of the launch vehicle based on the attitude angle calculation result and / or the inertial navigation result, the lever arm correction process can be performed on the second coordinate according to the relationship between the inertial measurement unit and the rocket centroid position to obtain the fifth coordinate; the resolved attitude result in the local geodetic coordinate system can be determined according to the attitude angle calculation result; and the observation equation of the launch vehicle can be constructed according to the fifth coordinate and the resolved attitude result.

[0089] Specifically, the satellite navigation positioning result is the position coordinate of the second antenna phase center, and the inertial navigation result is the navigation result of the inertial measurement center. Generally, the two center points do not coincide, and the lever arm correction should be performed during the integrated navigation calculation. The lever arm is the vector between the first antenna and the second antenna. The calculation process of predicting the position of the GNSS antenna phase center (i.e., the second positioning center) from the lever arm measurement value can be expressed by the following formula (4):

[0090]

[0091] Among them, is the GNSS antenna phase center (i.e., the second positioning center) estimated and predicted based on the inertial navigation result; is the inertial navigation result described above; is the transformation matrix from the inertial measurement unit (IMU) carrier coordinate system to the local geodetic coordinate system, is the lever arm vector from the IMU to the satellite navigation antenna in the IMU carrier coordinate system (with the centroid of the inertial navigation element as the origin, the forward direction of the element, which can be understood as parallel to the longitudinal direction of the launch vehicle as the Z-axis, and the right direction of the element as the X-axis).

[0092] In the embodiment of the present invention, according to the calculated attitude angle result, the resolved attitude result in the local geodetic coordinate system can be determined by the following formula (5):

[0093]

[0094] Among them, is the resolved attitude result in the local geodetic coordinate system; ψ is the calculated heading angle result described above; φ is the calculated pitch angle result described above.

[0095] In the embodiment of the present invention, in order to make the calculation result more accurate and considering the error in the GNSS calculation process, the position of the main antenna (i.e., the second antenna described above) obtained by GNSS positioning can be represented by the following formula (6):

[0096]

[0097] Among them, is the second coordinate described above, n r,G is the position error of GNSS, modeled as a white noise sequence; is the true value of the second coordinate.

[0098] According to the fifth coordinate and the resolved attitude result, the observation equation of the launch vehicle is constructed. The observation equation represents the motion state result calculated based on the measured data (e.g., through satellite navigation measurement), and can be as shown in formula (7):

[0099]

[0100] Among them, Z is the observation result, is the second coordinate obtained by measurement, is the predicted value of the second positioning center obtained according to the inertial navigation result in formula (4), and can be the estimated value of the second center obtained according to the inertial navigation result in formula (4), is the resolved attitude result, is the inertial navigation result, H is the observation equation design matrix, x is Formula (3), and V is the observation error. Specifically, H can be the following Formula (8):

[0101]

[0102] Wherein, is the transformation matrix from the inertial measurement unit (IMU) carrier coordinate system to the local geodetic coordinate system, is the lever arm vector from the IMU to the satellite navigation antenna in the IMU carrier coordinate system.

[0103] After the above calculations, the Kalman filter solution can be performed according to the state equation (i.e., Formula (3) described above) and the observation equation (i.e., Formula (7) described above) to obtain the attitude angle correction amount result x after dual-antenna correction, and x is compensated to the inertial navigation result to complete the attitude angle correction of the dual-antenna GNSS-aided launch vehicle.

[0104] To better understand the attitude angle control method of the launch vehicle provided by the embodiments of the present invention, an example is given below. It should be understood that the example is not restrictive. Figure 4 is the conceptual diagram of the attitude angle control method of the launch vehicle provided by the embodiments of the present invention. As Figure 4 shown, the control method of the launch vehicle provided by the embodiments of the present invention may include the following steps:

[0105] The first step: Install two GNSS antennas along the longitudinal axis of the rocket body on the rocket body, that is, a short baseline is formed. The antenna far from the rocket head is regarded as the base station, and single-point positioning of the main antenna (i.e., the second antenna described above) and relative positioning between the two antennas are performed to obtain the baseline vector (i.e., the target baseline vector described above) between the two antennas and the position of the main antenna (i.e., the second coordinate described above) in the WGS-84 coordinate system.

[0106] The second step: The baseline vector (i.e., the target baseline vector described above) and the position of the main antenna (i.e., the second coordinate described above) obtained by relative positioning solution are converted from the WGS-84 coordinate system to the local geodetic coordinate system (with the rocket's center of mass as the origin, the Y-axis pointing to the geodetic north, the X-axis pointing to the geodetic east, and the Z-axis perpendicular to the local reference ellipsoid and pointing to the sky) through the coordinate transformation matrix, and the transformation matrix is Formula (1) described above.

[0107] The third step: Convert the target baseline vector to the local geodetic coordinates, and then obtain the heading angle calculation result and the pitch angle calculation result according to Formula (2) above.

[0108] Step 4: Combine the single-point positioning result of the main antenna obtained by GNSS solution (i.e., the second coordinate), the heading angle of the rocket body (i.e., the heading angle calculation result), and the pitch angle (i.e., the heading angle calculation result) with the inertial navigation result in the local geodetic coordinate system through the formulas (3) to (8) described above for integrated navigation.

[0109] Through the attitude angle control system of the carrier rocket provided by the embodiments of the present invention, the drawback of the cumulative inertial navigation error over time can be compensated while reducing costs. To a certain extent, it solves the problem in the related art that the inertial navigation components are expensive, and simply relying on improving the performance of the inertial navigation components to compensate for the lack of navigation accuracy cannot effectively control costs.

[0110] In the embodiments of the present invention, a carrier rocket is further provided, including: a carrier rocket body; and a system disposed on the carrier rocket body, and the system is the attitude angle control system of the carrier rocket described above.

[0111] In the above description, technical details such as the composition of each layer are not elaborated in detail. However, those skilled in the art should understand that various technical means can be used to form layers, regions, etc. of the required shapes. In addition, to form the same structure, those skilled in the art can also design methods that are not exactly the same as the methods described above. In addition, although the above embodiments are described separately, this does not mean that the measures in each embodiment cannot be used in combination advantageously.

[0112] Although the preferred embodiments of the present invention have been described, those skilled in the art can make additional changes and modifications once they know the basic creative concept. Therefore, the appended claims are intended to be construed to include the preferred embodiments as well as all changes and modifications falling within the scope of the present invention.

[0113] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalent technologies, the present invention is also intended to include these modifications and variations.

Claims

1. An attitude angle control system for a launch vehicle, characterized in that, The system includes: A first antenna vertically installed near the head of the launch vehicle for determining a first coordinate, where the first coordinate is the coordinate of a first positioning center in the geocentric coordinate system, and the first positioning center is the positioning center on the first antenna; A second antenna vertically installed away from the head of the launch vehicle for determining a second coordinate, where both the first antenna and the second antenna are located on a target longitudinal axis, the target longitudinal axis is parallel to the longitudinal axis of the launch vehicle's body, the second coordinate is the coordinate of a second positioning center in the geocentric coordinate system, and the second positioning center is the positioning center on the second antenna; An inertial measurement unit installed inside the launch vehicle for providing inertial measurement results; A processor installed inside the launch vehicle for acquiring the first coordinate and / or the second coordinate, and for determining the attitude angle calculation result of the launch vehicle based on the first coordinate and the second coordinate; The processor is further configured to acquire the inertial measurement results, and determine the target attitude angle of the launch vehicle based on the attitude angle calculation result and / or the inertial measurement results.

2. A method for controlling the attitude angle of a launch vehicle, applied to the system as described in claim 1, characterized in that, The method includes: Acquiring a first coordinate and / or a second coordinate; Based on the first coordinate and / or the second coordinate, determining the attitude angle calculation result of the launch vehicle, where the attitude angle calculation result includes the heading angle calculation result and / or the pitch angle calculation result of the launch vehicle; Acquiring an inertial navigation result; Based on the attitude angle calculation result and / or the inertial navigation result, determining the target attitude angle of the launch vehicle.

3. The method according to claim 2, wherein The determining the attitude angle calculation result of the launch vehicle based on the first coordinate and / or the second coordinate includes: Based on the first coordinate and / or the second coordinate, determining a third coordinate, where the third coordinate is the coordinate of a target baseline vector in the geocentric coordinate system, and the target baseline vector is the baseline vector between the first positioning center and the second positioning center; Converting the third coordinate to a fourth coordinate, where the fourth coordinate is the coordinate of the target baseline vector in the local geodetic coordinate system, and the origin of the local geodetic coordinate system is the centroid position of the launch vehicle; Based on the fourth coordinate, determining the attitude angle calculation result of the launch vehicle.

4. The method according to claim 3, characterized in that The converting the third coordinate to the fourth coordinate includes: Acquiring the longitude and / or latitude of the second positioning center; Based on the longitude and / or latitude of the second positioning center, converting the third coordinate to the fourth coordinate through a coordinate transformation matrix.

5. The method according to claim 3, characterized in that, The determining the attitude angle calculation result of the launch vehicle based on the fourth coordinate includes: Based on the fourth coordinate, determining the attitude angle calculation result of the launch vehicle using trigonometric relationships.

6. The method according to claim 5, wherein The determining the attitude angle calculation result of the launch vehicle using trigonometric relationships based on the fourth coordinate includes: Determining the attitude angle calculation result of the launch vehicle through the following formula: Wherein, ψ is the calculated result of the course angle of the launch vehicle, φ is the calculated result of the pitch angle of the launch vehicle, x is the value of the fourth coordinate on the X-axis, y is the value of the fourth coordinate on the Y-axis, and z is the value of the fourth coordinate on the Z-axis.

7. The method according to claim 3, characterized in that, Determining the target attitude angle of the launch vehicle based on the calculated result of the attitude angle and / or the inertial navigation result includes: Obtaining the state equation of the launch vehicle, where the state equation characterizes the motion state of the launch vehicle; Constructing the observation equation of the launch vehicle based on the calculated result of the attitude angle and / or the inertial navigation result; Performing Kalman filter solution according to the state equation and the observation equation to determine the target attitude angle of the launch vehicle.

8. The method according to claim 7, characterized in that, Constructing the observation equation of the launch vehicle based on the calculated result of the attitude angle and / or the inertial navigation result includes: Performing lever arm correction processing on the second coordinate according to the inertial navigation result to obtain the fifth coordinate; Determining the calculated attitude result in the local geodetic coordinate system according to the calculated result of the attitude angle; Constructing the observation equation of the launch vehicle according to the fifth coordinate and the calculated attitude result.

9. The method according to claim 8, characterized in that, The observation equation includes: where Z is the observation result, is the second coordinate, is the predicted value of the second positioning center obtained from the inertial navigation result, is the attitude solution result, is the inertial navigation result, H is the observation equation matrix, x is the target attitude angle, and V is the observation error.

10. A launch vehicle, characterized in that, Including: The launch vehicle body; A system disposed on the launch vehicle body, and the system is the attitude angle control system of the launch vehicle described in claim 1.

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