A method for compensating the angle measurement accuracy of star sensors

Abstract

The invention relates to a star sensor angle measurement accuracy compensation method, comprising: determining a generation mechanism of a star sensor angle measurement error, and establishing a high-dynamic star sensor angle measurement accuracy error model according to the mechanism; the generation mechanism of the star sensor angle measurement error is due to the mismatch between the star point energy center and the time center; using inertial navigation measurement data to solve carrier angle dynamic information in a flight mission; calculating the star sensor angle measurement error caused by dynamics online by combining the solved carrier angle dynamic information with the established high-dynamic star sensor angle measurement accuracy error model; and using the calculated star sensor angle measurement error caused by dynamics to compensate for the attitude angle output by the star sensor.

Description

Technical Field

[0001] The invention relates to improving the measurement accuracy of a star sensor used in a missile-borne high-dynamic environment. Background Art

[0002] The star sensor is a device that determines its own attitude by measuring the direction of the navigation star in the carrier coordinate system. Its working process is to first measure the vector of the navigation star in the carrier coordinate system, and then obtain the vector corresponding to the inertial coordinate system through star map recognition; by comparing the vector relationship of the corresponding navigation stars in the two coordinate systems, the transformation matrix from the inertial coordinate system to the carrier coordinate system can be obtained, that is, the attitude of the carrier in the inertial coordinate system.

[0003] Star sensors are high-precision attitude sensors with an accuracy of arc seconds. When used with a sunshade, they can achieve autonomous attitude determination all day long. They are widely used in navigation equipment for satellites, aircraft and ships. With the development of photosensitive detector technology, star sensors have the characteristics of high sensitivity, wide dynamic range, low power consumption, large image array, and high output frequency, and are suitable for use in high-dynamic missile-borne environments. Star sensors are directly strapped to the carrier. When the carrier moves at a certain angular velocity, the image formed by the navigation star after passing through the optical system moves on the photosensitive detector, forming an image shift, which in turn affects the angular measurement accuracy of the star sensor. Guo He et al. studied the relationship between the energy of dynamic star point pixels and the exposure time, and regarded the star points on the image sensor as uniform motion during the exposure time. It is believed that the influence of star point deformation caused by dynamic factors on the center of mass positioning can be ignored. Li Xiao et al. studied the accuracy of extracting the center of star image targets using the center of mass method when the angular velocity is less than 2° / s based on the APS pixel output model, and analyzed the influence of the calculation window and exposure time on the accuracy of extracting the center of star image targets on this basis. Zhao Shufang et al. imaged the navigation star twice continuously, calculated the center of mass of the dynamic star target using the centroid algorithm, and assumed that the star target moves in a uniform linear motion on the image sensor during the exposure process to compensate for the dynamic angle measurement error. Gong Dezhu et al. proposed a method to improve the dynamic performance of the sensor by changing the internal timing design of the CCD and the data processing software algorithm, but it is more difficult to solve the problem of long integration time and irregular star motion trajectory.

[0004] These methods studied the impact of dynamic environment on the accuracy of star sensors, and carried out research on accuracy compensation methods in combination with the application background. They all have certain reference significance, but there are still the following shortcomings in the application of high-dynamic missile engineering practice: First, the carrier attitude motion process is random, and the uniform angular motion assumption has limitations. Second, changing the internal photosensitive detector and data processing software of the star sensor increases the complexity of the navigation equipment. Third, the dynamic environment of different flight profiles varies greatly, the accuracy compensation algorithm depends on the flight mission, and the algorithm has poor versatility. Summary of the invention

[0005] The technical problem solved by the present invention is to overcome the deficiencies of the prior art and provide a method for compensating the angular measurement accuracy of a star sensor based on missile angular acceleration measurement, so as to compensate the angular measurement accuracy of the missile-borne star sensor during a flight mission.

[0006] The technical solution of the present invention is: a method for compensating the angle measurement accuracy of a star sensor, comprising:

[0007] Determine the generation mechanism of the star sensor angle measurement error, and establish a high dynamic star sensor angle measurement accuracy error model based on the mechanism; the generation mechanism of the star sensor angle measurement error is due to the mismatch between the star point energy center and the time center;

[0008] Inertial navigation measurement data is used to calculate the carrier angular dynamic information during flight missions;

[0009] The above-solved carrier angle dynamic information is combined with the established high-dynamic star sensor angle measurement accuracy error model to calculate the star sensor angle measurement error caused by the dynamics online.

[0010] The attitude angle output by the star sensor is compensated by using the star sensor angle measurement error caused by the dynamics calculated above.

[0011] Preferably, the angular measurement accuracy error model of the high dynamic star sensor is described as follows: when the carrier moves at a variable angular velocity, the influence of the carrier movement on the star point extraction accuracy is proportional to the angular acceleration and proportional to the square of the exposure time.

[0012] Preferably, the angular measurement accuracy error model of the high dynamic star sensor is:

[0013]

[0014] Where Δθ is the angle measurement error of the star sensor, is the angular acceleration of the carrier, and h is the exposure time of the star sensor.

[0015] Preferably, the formula for calculating the star sensor angle measurement error caused by dynamics online is:

[0016]

[0017] in, is the sampling period of the inertial navigation, i is the inertial navigation sampling period count during the exposure period of the star sensor, and t0 is the start exposure time of the star sensor.

[0018] Preferably, the grayscale value of the star sensor is represented in one dimension on a certain axis, and then the attitude angle output by the star sensor is determined according to the focal length of the star sensor lens; the certain axis is a pitch axis or a yaw axis.

[0019] Preferably, during the compensation process, the pitch axis and yaw axis attitude angles output by the star sensor are compensated respectively to complete the angle measurement accuracy compensation of the entire star sensor.

[0020] Preferably, the measured angular velocity increment is smoothed in solving the carrier angular dynamic information using inertial navigation measurement data to obtain the angular acceleration.

[0021] Preferably, the angular velocity increment is smoothed using a backward second-order difference formula.

[0022] Preferably, the formula for calculating the angular acceleration is:

[0023]

[0024] in, is the carrier angular acceleration of the current inertial navigation sampling period, is the carrier angular velocity increment of the current inertial navigation sampling period, is the carrier angular velocity increment of the previous inertial navigation sampling period, is the sampling period of the inertial navigation.

[0025] Preferably, the missile control system uses the inertial navigation measurement data during the exposure process to complete the correction of the star sensor attitude measurement, that is, to complete the star sensor angle measurement compensation.

[0026] The beneficial effects of the present invention compared with the prior art are:

[0027] In order to improve the accuracy of missile autonomous navigation, the navigation system accuracy is improved by combining the angle measurement information of star sensors with inertial devices. The missile-borne environment has the characteristics of high dynamic motion. Attitude changes will cause image shift of star sensors, affecting the angle measurement accuracy of star sensors, thereby reducing navigation accuracy. This paper studies the influence of carrier attitude motion on the angle measurement accuracy of star sensors, and proposes a star sensor angle measurement accuracy compensation method based on missile angular acceleration measurement. The influence of carrier angular motion on the measurement accuracy of star sensors is modeled, and the star sensor measurement accuracy is compensated by real-time angular motion. The angular motion information of missile-borne inertial navigation equipment has the characteristics of short-term high accuracy, and the sampling frequency of inertial navigation equipment is high. Multiple inertial navigation equipment sampling data can be obtained within the exposure time of a star sensor. Inertial navigation measurement data is used to solve high-precision carrier angular dynamic information in the flight mission to compensate for the angle measurement accuracy of missile-borne star sensors. Without changing the existing algorithm of the star sensor, the accuracy compensation algorithm is measured online and has nothing to do with the flight mission profile, which enhances the versatility of the algorithm implementation.

[0028] The present invention does not change the internal photosensitive detector and data processing software of the star sensor, and does not increase the complexity of the navigation equipment;

[0029] For different flight profiles with large differences in dynamic environments, the accuracy compensation algorithm of the present invention does not depend on the flight mission, and the algorithm has strong versatility. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 This is the imaging principle of star sensor;

[0031] Figure 2 This is a diagram showing the tailing of the star sensor imaging under dynamic conditions;

[0032] Figure 3 This is an illustration of the working sequence of Xingmin. DETAILED DESCRIPTION

[0033] The present invention will be further described below in conjunction with the embodiments.

[0034] Star sensors are high-precision attitude sensors with an accuracy of arc seconds. With the development of photosensitive detector technology, it has been successfully applied in the design of missile control systems as an all-day autonomous navigation device. The missile-borne environment has a high dynamic characteristic. Attitude changes will cause star point image shifts, which significantly reduces the angular measurement accuracy of star sensors compared to static conditions. The sampling frequency of inertial navigation equipment is high, and the attitude changes during the exposure process can be solved through the design of navigation algorithms. In summary, this paper studies the impact of high dynamic environments on the measurement accuracy of star sensors. Combined with the typical attitude adjustment process of the aircraft control system, a dynamic measurement accuracy compensation method for star sensors is proposed, which improves the measurement accuracy of star sensors in dynamic environments.

[0035] (1) Basic principles of star sensor imaging

[0036] like Figure 1 As shown, starlight is incident as parallel light, passes through the star sensor lens, and forms an image on the photosensitive detector. The photosensitive detector is installed at the focus of the lens, and the focal length of the lens is f.

[0037] like Figure 2 As shown, the coordinate system of the star sensor body is defined as follows: the origin of the coordinate is located at the center of the lens, the X axis is perpendicular to the detector target surface, and the Y and Z axes are orthogonal within the detector target surface, satisfying the right-hand rule. At the initial moment of exposure, the star point is located at O' at the imaging point of the detector. When the carrier has posture motion, the star sensor rotates around the Z axis by an angle θ at the end of exposure, and the star point is located at O" at the imaging point of the detector. O'O"≈fθ, that is, in the dynamic case, the image of the star point is linear.

[0038] In order to improve the detection magnitude of the missile-borne star sensor, the exposure time cannot be too short, usually tens of milliseconds. During the exposure process, based on the carrier speed of 4° / s and the exposure of 30ms, the attitude of the star sensor changes by 7.2′ during the exposure process.

[0039] (2) Star sensor working sequence

[0040] like Figure 3 As shown, exposure and data processing are completed in sequence within each star sensor working frame rate to calculate the attitude information.

[0041] The present invention provides a method for compensating the angle measurement accuracy of a star sensor, and the specific implementation method is as follows:

[0042] Use the gray value of the star sensor to extract the centroid of the star point:

[0043] During the star measurement process, the control system sends a command to the star sensor to start the star measurement. The exposure starts at time t0, the angle between the starlight vector and the optical axis is θ0, and the exposure process ends at time (t0+h), where h is the exposure time. The energy center of the star point is extracted through the centroid extraction algorithm and sent to the control system as the exposure center at time (t0+h / 2).

[0044] Assume that the star point occupies n pixels on the photosensitive detector, and record the coordinate value of the mth pixel as (u m ,v m ), the gray value of the mth pixel is G m , then the pixel coordinates u and v of the star point energy centroid can be calculated by formula (1):

[0045]

[0046] The gray value is a digital representation of the energy of starlight on the photosensitive detector. The number of electrons generated by starlight during the exposure time is converted into grayscale value after analog / digital conversion. When the optical system design of the star sensor is completed, the grayscale value is proportional to the exposure time, that is, G = Kh, and the proportional coefficient K is determined by the optical system design. The star point energy center of mass is represented on the Y axis of the star sensor body coordinate system The corresponding star sensor output attitude angle is Formula (3):

[0047]

[0048]

[0049] Among them, θ t is the carrier motion attitude angle. Similarly, the star point energy center of mass on the Z axis of the star sensor body coordinate system and the output attitude can be obtained.

[0050] Formula (3) can be used to obtain the attitude angle (including pitch angle and yaw angle) output by the star sensor in each working cycle.

[0051] Modeling of angular measurement accuracy error of high dynamic star sensor:

[0052] The true attitude angle θ1 of the carrier at the exposure center of the star sensor is given by equation (4), and the angle measurement error Δθ of the star sensor is given by equation (5): is the angular velocity of the carrier, θ(t0) is the initial angle value at the starting point of the carrier movement at time t0, and θ is the angle value measured by the star sensor.

[0053]

[0054]

[0055] Angular motion of carrier The modeling is as follows:

[0056] When the carrier moves at a variable angular velocity, but

[0057] Will Substituting into formulas (2) and (3), the centroid coordinates and attitude angles of the star points measured by the star sensor are:

[0058]

[0059] in is the initial value of the carrier's angular velocity.

[0060] Will Substituting into formula (4), the true attitude angle of the carrier at the exposure center is:

[0061]

[0062] Star sensor angle measurement error It can be seen that when the carrier moves at a variable angular velocity, the influence of the carrier movement on the accuracy of star point extraction is proportional to the angular acceleration and proportional to the square of the exposure time. The mechanism of the angle measurement error of the star sensor is due to the mismatch between the energy center and the time center of the star point.

[0063] Using inertial navigation to calculate dynamic environment:

[0064] The angular increment information of the missile-borne inertial navigation equipment has the characteristics of high accuracy in a short time, and the inertial navigation equipment has a high sampling frequency. Multiple sampling data can be obtained within the exposure time of the star sensor. In the flight mission, the inertial navigation measurement data is used to calculate the high-precision carrier angle dynamic information to compensate for the angle measurement accuracy of the star sensor.

[0065] The inertial navigation device samples at discrete time and measures the carrier angle increment. The sampling period Generally, it is 5ms, and the carrier attitude angle is solved by discrete time integration. Then, the inertial navigation sampling results are differentially calculated to obtain the carrier angular velocity for use by the control system. In order to improve the angular velocity measurement accuracy, the backward second-order difference formula is used to smooth the angular increment.

[0066] According to Taylor's formula, the attitude angle is calculated by the second-order backward difference:

[0067]

[0068] Thus, the angular velocity of the current navigation cycle is obtained

[0069]

[0070] in, Δθ i is the angle increment of the current navigation cycle, Δθ i-1 The angle increment of the previous navigation cycle.

[0071] Similarly, the angular velocity increment is smoothed to obtain the angular acceleration of the current navigation cycle

[0072]

[0073] in, is the angular velocity increment of the current navigation cycle, It is the angular velocity increment of the previous navigation cycle.

[0074] Star sensor angle measurement accuracy compensation:

[0075] The angular dynamic information of the carrier solved by the missile-borne inertial navigation equipment is used to calculate the angular measurement error of the star sensor caused by the dynamics. The attitude angular acceleration actually generated by the control system is not a constant value. With a small change over time, the angle measurement error of the star sensor is,

[0076]

[0077] The inertial navigation online output is a discrete quantity. By discretizing the expression of formula (10), we can get:

[0078]

[0079] Where i is the inertial navigation sampling cycle count during the star sensor exposure period. For example, when When is 5ms, h is 30ms, i is 1, 2, ...6.

[0080] Attitude angle obtained by the star sensor centroid extraction algorithm Perform accuracy compensation to obtain the corrected attitude angle

[0081]

[0082] Without changing the existing algorithm of the star sensor, the missile control system uses the inertial navigation measurement data during the exposure process to complete the correction of the star sensor attitude measurement. The precision compensation algorithm is independent of the flight mission profile through online measurement, which enhances the versatility of the algorithm implementation.

[0083] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0084] Parts not described in detail in the present invention belong to common knowledge of those skilled in the art.

Claims

1. A method for compensating the angle measurement accuracy of a star sensor, characterized in that include: Determine the generation mechanism of the star sensor angle measurement error, and establish a high dynamic star sensor angle measurement accuracy error model based on the mechanism; the generation mechanism of the star sensor angle measurement error is due to the mismatch between the star point energy center and the time center; Inertial navigation measurement data is used to calculate the carrier angular dynamic information during flight missions; The above-solved carrier angle dynamic information is combined with the established high-dynamic star sensor angle measurement accuracy error model to calculate the star sensor angle measurement error caused by the dynamics online. The attitude angle output by the star sensor is compensated by using the star sensor angle measurement error caused by the above-calculated dynamics; The high dynamic star sensor angular measurement accuracy error model is described as follows: when the carrier moves at a variable angular velocity, the influence of the carrier movement on the star point extraction accuracy is proportional to the angular acceleration and proportional to the square of the exposure time; The high dynamic star sensor angle measurement accuracy error model is: Where Δθ is the angle measurement error of the star sensor, is the angular acceleration of the carrier, h is the exposure time of the star sensor; The formula for calculating the angle measurement error of the star sensor caused by dynamics online is: in, is the sampling period of the inertial navigation, i is the inertial navigation sampling period count during the exposure period of the star sensor, and t0 is the start exposure time of the star sensor.

2. The method according to claim 1, characterized in that: The gray value of the star sensor is represented in one dimension on a certain axis, and then the attitude angle output by the star sensor is determined according to the focal length of the star sensor lens; the certain axis is a pitch axis or a yaw axis.

3. The method according to claim 2, characterized in that: During the compensation process, the pitch axis and yaw axis attitude angles output by the star sensor are compensated separately to complete the angle measurement accuracy compensation of the entire star sensor.

4. The method according to claim 1, characterized in that: The angular velocity increment measured in the inertial navigation measurement data is smoothed to obtain the angular acceleration when solving the carrier angular dynamic information.

5. The method according to claim 4, characterized in that: The backward second-order difference formula is used to smooth the angular velocity increment.

6. The method according to claim 5, characterized in that: The formula for calculating angular acceleration is: in, is the carrier angular acceleration of the current inertial navigation sampling period, is the carrier angular velocity increment of the current inertial navigation sampling period, is the carrier angular velocity increment of the previous inertial navigation sampling period, is the sampling period of the inertial navigation.

7. The method according to any one of claims 1 to 6, characterized in that: The missile control system uses the inertial navigation measurement data during the exposure process to correct the star sensor attitude measurement, that is, to complete the star sensor angle measurement compensation.

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