High-dynamic pointing time alignment method for spacecraft

Abstract

The invention discloses a high-dynamic pointing time alignment method of a spacecraft, and provides a time alignment concept on the basis of a relative measurement scheme of optical angle measurement information and microwave distance measurement information, and the method comprises the following steps: time alignment of a star sensor and an orbit; time alignment of the star sensor and the microwave distance measurement; time alignment of the track and the microwave distance measurement; microwave ranging is temporally aligned with the track. Compared with the prior art, the method has the beneficial effects that reasonable time sequence arrangement and time compensation are carried out on the process of the algorithm, and the relative pointing error caused by distance measurement data delay and track attitude data asynchronization is overcome, so that high-precision pointing under a high-dynamic working condition is carried out.

Description

Technical Field

[0001] The invention relates to a high dynamic pointing time alignment method for a spacecraft, belonging to the technical field of spacecraft attitude determination. Background Art

[0002] For high dynamic conditions with large anisotropy, due to the different working principles of different measurement units, optical measurement is to image the target and identify the target by controlling the detector exposure, and finally give the azimuth and pitch angle of the target relative to the main satellite; microwave radar is to measure the angle and distance of the target by microwave pulse reflection. In order to complete the high dynamic conditions of anisotropy, not only the relative motion information of the two satellites is required, but also the absolute orbit and attitude determination of the main satellite is required, which involves star sensitivity, gyroscope and orbit calculation. Among them, star sensitivity also realizes the measurement of satellite inertial attitude through the exposure of the detector, and the gyroscope obtains the angular increment information within a period of time through continuous integration to obtain the average angular velocity within the period of time. Therefore, in order to complete high dynamic pointing with large angular velocity, it is necessary to arrange the timing of each sensor unit reasonably and compensate the corresponding data appropriately. Therefore, research in this area needs to be solved urgently and becomes a key issue affecting pointing accuracy. Therefore, how to ensure the continuity and real-time performance of relative navigation data is the key to achieving high-precision pointing maintenance. Summary of the invention

[0003] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and propose a high-dynamic pointing time alignment method for a spacecraft, which, through reasonable timing arrangement and precise time alignment, calculates the relevant quantities of time compensation, ensures the continuity and real-time nature of relative navigation data, and realizes high-precision pointing to the target.

[0004] The technical solution of the present invention is:

[0005] A high dynamic pointing time alignment method for a spacecraft, comprising:

[0006] Get the time difference between the star sensor attitude timestamp and the microwave ranging timestamp, and then calculate the star sensor attitude Q″ when the star sensor and microwave ranging times are aligned based on the gyro angular velocity. ibX ;

[0007] Obtain the time difference between the orbit corresponding timestamp and the microwave ranging timestamp, perform orbit compensation, and obtain the six orbit numbers at the microwave ranging time after compensation;

[0008] According to the star sensor attitude Q″ when the star sensor is aligned with the microwave ranging time ibX , the six orbital numbers at the time of microwave ranging after compensation, and the orbital system attitude at the time of microwave ranging are calculated; combined with the measurement results of microwave ranging, relative navigation calculation is performed to obtain the relative position and speed of the two satellites in the orbital system at the time of microwave ranging.

[0009] Furthermore, calculate the star sensor attitude Q″ when the star sensor and microwave ranging are time-aligned. ibX The method is as follows:

[0010] Calculate the star sensor deviation compensation quaternion required for the star sensor and microwave ranging to achieve time alignment:

[0011]

[0012] In the formula, ω i is the gyro angular velocity, and ΔT_OH_2 is the time difference between the star sensor attitude timestamp and the microwave ranging timestamp;

[0013] According to the star sensor deviation compensation quaternion Δq ω _2, obtain:

[0014]

[0015] In the formula, Q ibX is the inertial quaternion of the star sensor.

[0016] Furthermore, obtain the time difference between the orbit corresponding timestamp and the microwave ranging timestamp, and perform orbit compensation to obtain the six orbital elements at the compensated microwave ranging moment, specifically:

[0017] Calculate the average orbital angular velocity before compensation:

[0018] where μ = 3.986004415e14

[0019] Calculate the six orbital elements after compensation:

[0020] a k = a g , e k = e g , i k = i g

[0021] Ω k = Ω g , ω k = ω g

[0022] Take the modulo of 2π;

[0023]

[0024] where, a g , e g , i g , Ω g , ω g , fg , M g are respectively the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, true anomaly, and mean anomaly of the orbit before compensation; a k , e k , i k , Ω k , ω k , f k , M k are respectively the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, true anomaly, and mean anomaly of the orbit at the microwave ranging moment after compensation.

[0025] Furthermore, the latitude depression angle at the microwave ranging moment after compensation is:

[0026] u k = ω k + f k , modulo 2π

[0027] where u k is the latitude depression angle at the microwave ranging moment after compensation.

[0028] Furthermore, perform relative navigation calculation to obtain the relative position of the two satellites in the orbit system at the microwave ranging moment:

[0029] First, calculate the time difference between the microwave ranging timestamp and the orbit corresponding timestamp:

[0030] ΔT dh = T_GD - T_LD

[0031] According to the relative velocity, perform relative position compensation to align to the orbit corresponding moment:

[0032]

[0033] where is the relative position of the two satellites after compensation, is the relative position of the two satellites output by relative navigation, is the relative velocity of the two satellites output by relative navigation.

[0034] Furthermore, obtain the time difference between the star sensor attitude timestamp and the orbit corresponding timestamp, and then calculate the star sensor attitude Q' when the star sensor is aligned with the orbit time according to the gyro angular velocity ibX ; then, according to the relative position of the two satellites after compensation and the relative velocity of the two satellites output by navigation, solve for the attitude angle and angular velocity for control.

[0035] Furthermore, the method for calculating the star sensor attitude Q' when the star sensor is aligned with the orbit time ibX is:

[0036] Calculate the star sensor deviation compensation quaternion Δq required for the time alignment of the star sensor and the orbit ω _1:

[0037]

[0038] where ω i is the gyro angular velocity, and ΔT_OH_1 is the time difference between the star sensor attitude timestamp and the corresponding orbit timestamp;

[0039] According to the star sensor deviation compensation quaternion Δq ω _1, we get:

[0040]

[0041] where Q ibX is the inertial quaternion of the star sensor.

[0042] A computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the above method are implemented.

[0043] An electronic device includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, the steps of the above method are implemented.

[0044] The advantages of the present invention compared with the prior art are as follows:

[0045] (1) The present invention designs the time alignment of the orbit and ranging, which is beneficial to the accuracy of relative navigation input, can calculate more accurate relative positions and speeds, can reduce the relative navigation error caused by orbit delay, and improve the relative navigation accuracy.

[0046] (2) The present invention designs the time alignment of ranging and orbit, which is beneficial to the accuracy of relative navigation output, reduces the pointing accuracy error caused by ranging delay, and improves the pointing accuracy.

[0047] (3) The present invention designs the time alignment of the star sensor and the orbit, which is beneficial to absolute attitude determination, can calculate a more accurate orbit system attitude, reduces the absolute attitude determination error caused by star sensor delay, and improves the attitude determination accuracy. Description of the Drawings

[0048] By reading the following detailed description of the preferred embodiments, various other advantages and benefits will become clear to those of ordinary skill in the art. The drawings are only for the purpose of showing the preferred embodiments and are not considered to be a limitation of the present invention. Moreover, throughout the drawings, the same reference numerals are used to represent the same components. In the drawings:

[0049] Figure 1 This is the relative position error curve of the relative navigation output after time alignment compensation in the embodiment of the present invention;

[0050] Figure 2 This is the relative velocity error curve of the relative navigation output after time alignment compensation in the embodiment of the present invention;

[0051] Figure 3 This is the relative pointing angle error curve calculated after time alignment compensation in the embodiment of the present invention;

[0052] Figure 4 This is the relative pointing angular velocity error curve calculated after time alignment compensation in the embodiment of the present invention;

[0053] Figure 5 This is the relative pointing angular acceleration error curve calculated after time alignment compensation in the embodiment of the present invention;

[0054] Figure 6 This is the flowchart of the high-dynamic pointing time alignment method for the space vehicle of the present invention. Detailed implementation manners

[0055] Hereinafter, exemplary embodiments of the present disclosure will be described in more detail with reference to the drawings. Although the exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided so that the present disclosure can be more thoroughly understood and the scope of the present disclosure can be completely conveyed to those skilled in the art.

[0056] The present invention proposes a high-dynamic pointing time alignment method for a space vehicle, as Figure 6 shown, including the following steps:

[0057] 1) Time alignment between star sensor and orbit

[0058] Using the time stamp of the star sensor, calculate the time difference ΔT_OH_1(k) between the star sensor attitude time stamp and the corresponding orbit time stamp:

[0059] ΔT_OH_1 = T_GD - T_OH (1)

[0060] In the formula, T_GD is the corresponding orbit time stamp, and T_OH is the time stamp corresponding to the star sensor attitude;

[0061] According to the gyro angular velocity, calculate the deviation compensation quaternion of the star sensor and calculate the compensated attitude:

[0062]

[0063] In the formula, Q ibXThe inertial quaternion of the star sensor is ω i The angular velocity of the gyroscope is

[0064] 2) Time alignment between the star sensor and microwave ranging

[0065] Using the time stamp of the star sensor, calculate the time difference ΔT_OH_2 between the attitude time stamp of the star sensor and the time stamp of microwave ranging:

[0066] ΔT_OH_2 = T_LD - T_OH (4)

[0067] In the formula, T_LD is the time stamp corresponding to microwave ranging;

[0068] According to the angular velocity of the gyroscope, calculate the deviation compensation quaternion of the star sensor and calculate the compensated attitude:

[0069]

[0070]

[0071] 3) Time alignment between the orbit and microwave ranging

[0072] Calculate the time difference between the time stamp corresponding to the orbit and the time stamp of microwave ranging:

[0073] ΔT = T_LD - T_GD (7)

[0074] Use the simple orbit recurrence to perform orbit compensation:

[0075] First, calculate the average orbit angular velocity:

[0076]

[0077] Six orbital elements:

[0078] a k = a g , e k = e g , i k = i g

[0079] Ω k = Ω g , ω k = ω g

[0080] , modulo 2π;

[0081]

[0082] u k = ω k + f k , modulo 2π

[0083] Among them, a is the semi-major axis of the orbit, e is the eccentricity, i is the orbital inclination, Ω is the right ascension of the ascending node, ω is the argument of perigee, f is the true anomaly, M is the mean anomaly, u is the latitude depression angle. Those with subscript g are the orbital elements before recursion, and those with subscript k are the orbital elements after recursion.

[0084] 4) Time alignment between microwave ranging and the orbit

[0085] After the above steps are completed, use the star sensor attitude Q″ at the microwave ranging moment ibX and the six orbital elements at the microwave ranging moment to calculate the attitude of the orbital system at the microwave ranging moment. Then, combined with the measurement results of microwave ranging, perform relative navigation calculation to obtain the relative position between the two satellites in the orbital system and velocity The time stamp corresponding to this result is the microwave ranging moment. Next, it is necessary to align the microwave ranging moment with the orbit calculation moment:

[0086] First, calculate the time difference between the microwave ranging time stamp and the corresponding time stamp of the orbit:

[0087] ΔT dh = T_GD - T_LD (10)

[0088] According to the relative velocity, perform compensation for the relative position and align it to the corresponding orbit moment:

[0089]

[0090] Among them, is the compensated relative position, is the relative position before compensation output by relative navigation, is the relative velocity output by relative navigation.

[0091] Then, combined with the result of step 1), use the star sensor attitude Q′ at the orbit calculation moment ibX , and the compensated relative position and relative velocity output by relative navigation to calculate the attitude angle and angular velocity for control.

[0092] The present invention will be further described below through specific embodiments:

[0093] The simulated orbit condition is a high-dynamic condition, as shown in Table 1:

[0094] Table 1 Simulated orbit condition information table

[0095] Initial condition Primary star Target star <![CDATA[Semi-major axis a0]]> 42047199.99m 42166699.99m <![CDATA[Eccentricity e0]]> 0.000011 0.003184 <![CDATA[Orbital inclination i0]]> 0.1° 11.254° <![CDATA[Right ascension of the ascending node Ω0]]> 0.07° 23.3739° <![CDATA[Argument of perigee ω0]]> 159.09° 161.491° <![CDATA[True anomaly f0]]> 35.227° 9.4689°

[0096] The simulation setting delay is: star sensor delay 300 ms, tracking and ranging delay 200 ms

[0097] First, calculate the exposure compensation time difference ΔT_OH_1 between the star sensor and the orbit. According to the gyro angular velocity, calculate the deviation compensation quaternion Δq ω _1 of the star sensor, and calculate the compensated attitude Q′ ibX .

[0098] Then, calculate the exposure compensation time difference ΔT_OH_2 between the star sensor and the microwave ranging. Then, according to the gyro angular velocity, calculate the deviation compensation quaternion Δq ω _2 of the star sensor, and calculate the compensated attitude Q″ ibX .

[0099] Furthermore, calculate the time difference ΔT between the orbit corresponding timestamp and the microwave ranging timestamp. Then, use the simple orbit recurrence to perform orbit compensation:

[0100] First, calculate the average orbit angular velocity n g and the six orbital elements. After calculation by the relative navigation algorithm, the output relative position error results are shown in Figure 1 , and the output relative velocity results are shown in Figure 2 . It can be seen from the simulation curve that after time alignment compensation, the relative navigation position accuracy is 20 m and the velocity accuracy is 0.15 m / s.

[0101] Finally, calculate the time difference between the microwave ranging timestamp and the orbit corresponding timestamp, and then perform relative position compensation according to the relative velocity, align it to the orbit corresponding moment to obtain the relative position output by relative navigation. Then, calculate the relative pointing angle, relative pointing angular velocity, and relative pointing angular acceleration, and their errors are respectively as shown in Figure 3 , Figure 4 , Figure 5 . It can be seen from the simulation curve that after time alignment compensation, the pointing angle accuracy is 0.001°, the pointing angular velocity accuracy is better than 0.0001° / s, and the pointing angular acceleration accuracy is better than 0.000001° / s 2 .

[0102] The above-described embodiments are only relatively preferred specific embodiments of the present invention. The general changes and substitutions made by those skilled in the art within the scope of the technical solution of the present invention should be included in the protection scope of the present invention.

Claims

1. A high dynamic pointing time alignment method for a spacecraft, characterized in that: include: Get the time difference between the star sensor attitude timestamp and the microwave ranging timestamp, and then calculate the star sensor attitude Q when the star sensor and microwave ranging times are aligned based on the gyro angular velocity. i ' b ' X ; Obtain the time difference between the orbit corresponding timestamp and the microwave ranging timestamp, perform orbit compensation, and obtain the six orbit numbers at the microwave ranging time after compensation; According to the star sensor attitude Q when the star sensor is aligned with the microwave ranging time i ″ bX , calculate the orbital attitude at the time of microwave ranging based on the six orbital elements at the time of microwave ranging after compensation; Combined with the measurement results of microwave ranging, relative navigation calculations are performed to obtain the relative positions and velocities of the two satellites in the orbital system at the time of microwave ranging.

2. A method for high dynamic pointing time alignment of a spacecraft according to claim 1, characterized in that: Calculate the attitude Q of the star sensor when the star sensor is aligned with the microwave ranging time i ″ bX The method is: Calculate the star sensor deviation compensation quaternion required for time alignment between star sensor and microwave ranging: In the formula, ω i is the gyro angular velocity, ΔT_OH_2 is the time difference between the star-sensing attitude timestamp and the microwave ranging timestamp; Compensate the quaternion Δq based on the star sensitivity deviation ω _2, we get: In the formula, Q ibX is the inertial quaternion of star sensitivity.

3. The high dynamic pointing time alignment method for a spacecraft according to claim 1, characterized in that: Get the time difference between the orbit corresponding timestamp and the microwave ranging timestamp, perform orbit compensation, and get the six orbit numbers at the microwave ranging time after compensation, specifically: Calculate the average orbital angular velocity before compensation: where μ = 3.986004415e14 Calculate the six orbital numbers after compensation: to k =a g ,And k =and g ,the k =i g Oh k =Oh g ,oh k =ω g modulo 2π; Among them, a g 、e g 、i g ,Ω g ,ω g 、f g 、M g are the orbital semi-major axis, eccentricity, orbital inclination, right ascension of ascending node, argument of perigee, true anomaly, and mean anomaly before compensation; a k 、e k 、i k ,Ω k ,ω k 、f k 、M k They are respectively the orbital semi-major axis, eccentricity, orbital inclination, right ascension of ascending node, argument of perigee, true anomaly and mean anomaly at the moment of microwave ranging after compensation.

4. A method for high dynamic pointing time alignment of a spacecraft according to claim 3, characterized in that: The latitude depression angle of the microwave ranging measurement after compensation is: u k =ω k +f k , modulo 2π In the formula, u k is the latitude depression angle at the time of microwave ranging after compensation.

5. The high dynamic pointing time alignment method for a spacecraft according to claim 1, characterized in that: Perform relative navigation calculations to obtain the relative positions of the two satellites in the orbital system at the time of microwave ranging: First, calculate the time difference between the microwave ranging timestamp and the orbit corresponding timestamp: ΔT dh =T_GD-T_LD According to the relative speed, the relative position is compensated and aligned to the corresponding time of the track: in, The relative position of the two stars after compensation. The relative positions of the two stars output by relative navigation, The relative speed between the two stars outputted by relative navigation.

6. The high dynamic pointing time alignment method for a spacecraft according to claim 1, characterized in that: Get the time difference between the star sensor attitude timestamp and the orbit corresponding timestamp, and then calculate the star sensor attitude Q when the star sensor and orbit time are aligned according to the gyro angular velocity. i ' bX ; Then, according to the relative position of the two satellites after compensation and the relative speed of the two satellites output by the navigation, the attitude angle and angular velocity for control are calculated.

7. A method for high dynamic pointing time alignment of a spacecraft according to claim 6, characterized in that: Calculate the star sensitivity attitude Q when the star sensitivity is aligned with the orbit time i ' bX The method is: Calculate the star sensitivity deviation compensation quaternion Δq required for time alignment between star sensitivity and orbit ω _1: In the formula, ω i is the gyro angular velocity, ΔT_OH_1 is the time difference between the star-sensing attitude timestamp and the orbit corresponding timestamp; Compensate the quaternion Δq based on the star sensitivity deviation ω _1, we get: In the formula, Q ibX is the inertial quaternion of star sensitivity.

8. A computer-readable storage medium storing a computer program, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

9. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the steps of the method according to any one of claims 1 to 7 are implemented.

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