CN113722931B - Spacecraft station measurement guiding calculation method based on combined deviation orbit prediction - Google Patents

Spacecraft station measurement guiding calculation method based on combined deviation orbit prediction Download PDF

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CN113722931B
CN113722931B CN202111063244.0A CN202111063244A CN113722931B CN 113722931 B CN113722931 B CN 113722931B CN 202111063244 A CN202111063244 A CN 202111063244A CN 113722931 B CN113722931 B CN 113722931B
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黄静琪
孙山鹏
王彦荣
陈俊收
景方
李菲菲
赵力文
苏杨宋裔
赵哲龙
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Abstract

The invention discloses a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which comprises the following steps: step 1: carrying out orbit prediction on the autonomous control spacecraft according to the autonomous control program and the orbit dynamics model, marking a prediction result as a spacecraft nominal ephemeris, and calculating nominal guiding data by using the nominal ephemeris; step 2: converting the errors of the space environment model and the dynamic model into control deviation, and forecasting to generate a plurality of deviation ephemeris according to the range of the control deviation; step 3: calculating deviation guiding data according to the deviation ephemeris, and calculating an angle difference value and a time difference value between the deviation guiding data and a nominal station entering point; step 4: and comprehensively calculating the guide information according to the guide data of the deviation ephemeris. When the guiding data calculated by the method is captured, the guiding data is only required to be waited at a theoretical waiting point, the waiting time is a determined interval, the station measurement operation is simple, and the tracking condition and the state of the spacecraft can be judged.

Description

Spacecraft station measurement guiding calculation method based on combined deviation orbit prediction
Technical Field
The invention belongs to the technical field of spacecraft station measurement tracking and guiding, and provides a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction.
Background
The autonomous control of the spacecraft refers to the process of autonomously completing space specified actions or tasks by means of own sensors and control devices under the condition that the spacecraft does not need to annotate control instructions on the ground, and comprises on-orbit autonomous control, reentry and return of the spacecraft and the like. In the autonomous control process or after the spacecraft is finished, in order to receive telemetry data and judge the working state, a ground station is sometimes required to track the spacecraft to acquire related data. In order to provide tracking guidance for the ground station, the ground system needs to simulate the autonomous movement process of the spacecraft, forecast the ephemeris of the spacecraft, and then convert the forecasted ephemeris of the spacecraft into azimuth and pitching data to provide tracking guidance information for the ground station. However, the ground system cannot accurately simulate the autonomous control process of the spacecraft, and dynamic model errors, engine faults and the like also influence actual control results, so that deviation exists between the forecast value of the guiding information and the actual value. The beam width of the ground station measurement and control antenna is limited, and the guiding data deviation can directly influence the tracking capture and fault discrimination of the ground station, so that a station measurement and control method for synthesizing various orbit prediction deviations is required to be researched in an autonomous spacecraft control mode.
Disclosure of Invention
The invention aims to provide a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which solves part of problems of ground station tracking and guiding under the condition of autonomous control deviation of a spacecraft.
The technical proposal adopted by the invention is that,
a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction comprises the following steps:
step 1: carrying out orbit prediction on the autonomous control spacecraft according to an autonomous control program and an orbit dynamics model, marking a prediction result as a spacecraft nominal ephemeris, and calculating nominal guiding data, namely nominal arrival time, azimuth and pitch angle by using the nominal ephemeris;
step 2: converting the errors of the space environment model and the dynamic model into control deviation, and forecasting to generate a plurality of deviation ephemeris according to the range of the control deviation;
step 3: calculating deviation guiding data according to the deviation ephemeris, and calculating an angle difference value and a time difference value between the deviation guiding data and a nominal station entering point;
if the angle difference of the guide data calculated by a certain deviation ephemeris is larger than or equal to the half-wave beam range of the antenna, the equipment cannot conduct guide capturing through the method under the deviation;
step 4: and comprehensively calculating the guide information according to the guide data of the deviation ephemeris.
The present invention is also characterized in that,
step 1 specifically includes, under the newton frame of reference, the equations of motion of the detector can be written as the following equations (1), (2) and (3):
wherein, the liquid crystal display device comprises a liquid crystal display device,for the two gravitation and other various ingenuities of spacecraft, +.>For thrust force applied to autonomous control process of spacecraft, initial value state r of probe is given 0 And->The nominal ephemeris of the detector can be obtained by integrating the above formula.
In step 2, the method for calculating the plurality of offset ephemeris specifically includes:
setting the total number of the offset ephemeris as N;
adding the control deviation, which is converted into other deviations, to the formula (1) to obtain the formula (4):
wherein, the liquid crystal display device comprises a liquid crystal display device,for different control offsets, a plurality of offset ephemeris may be calculated according to equation (4).
The step 3 is specifically as follows: calculating station tracking azimuth angle A corresponding to ith deviation ephemeris i Pitch angle E i And interpolate E i ＝E 0 At the corresponding epoch time T i And azimuth angle A i Calculating the departure time deviation delta T according to formulas (12) and (13) i And an inbound azimuth deviation ΔA i ：
ΔT i ＝T i -T 0 (12)，
ΔA i ＝A i -A 0 (13)，
If delta A i Less than the half-wave beam range of the antenna, recording the corresponding delta T of the offset ephemeris i Step 4 is entered;
if delta A i Greater than or equal to the half-wave beam range of the antenna, then interpolate to calculate A i ＝A 0 At the corresponding epoch time T i And pitch angle E i Calculating an arrival time deviation and an arrival pitch angle deviation according to formulas (14) and (15);
ΔT i ＝T i -T 0 (14)，
ΔE i ＝E i -E 0 (15)，
if delta E i Less than the half-wave beam range of the antenna, recording the corresponding delta T of the offset ephemeris i Step 4 is entered;
if delta E i If the difference is greater than or equal to the half-wave beam range of the antenna, the difference is too large to be suitable for the method.
In step 3, the angular deviation includes an azimuth deviation corresponding to the same pitch angle of the deviation ephemeris as the nominal approach point, or a pitch deviation corresponding to the same azimuth angle.
Step 4, specifically, calculating the minimum and maximum arrival deviation moments according to the following formulas (16) and (17);
ΔT min ＝min{ΔT i } (16)，
ΔT max ＝max{ΔT i } (17)，
guidance information { T for ground station 0 ,A 0 ,E 0 ,ΔT min ,ΔT max }, T therein 0 For theoretical arrival time, A 0 、E 0 Wait for azimuth and elevation, delta T, respectively, of the antenna min 、ΔT max Respectively the leading and trailing edges of the latency.
The beneficial effects of the invention are: the spacecraft station measurement guiding calculation method based on the combined deviation orbit prediction is suitable for the station measurement guiding problem under the condition that the orbit prediction has deviation during autonomous control of the spacecraft; when the station is used for capturing the guide data calculated by the method, the station only needs to wait at a theoretical waiting point, the waiting time is a determined interval, the station is easy to operate, and the method is beneficial to judging the tracking condition and the state of the spacecraft.
Detailed Description
The invention relates to a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which is further described in detail below with reference to the accompanying drawings and the detailed description.
In order to complete the heaven and earth cooperative tasks during autonomous control of the spacecraft, the ground system needs to simulate the autonomous control process of the spacecraft to forecast the ephemeris of the spacecraft and is used for guiding the station measurement tracking. In a general task, an epoch and an azimuth angle corresponding to a certain pitch angle observed by a ground station on a spacecraft are mainly used as guiding information, the ground station points an antenna to the azimuth and pitch a plurality of minutes before the epoch moment, and the spacecraft enters an antenna beam at the epoch moment. The range of allowed pilot data errors is limited due to the limited antenna beam width. For the autonomous control spacecraft, the ground system cannot accurately simulate the movement process of the spacecraft, and when the predicted ephemeris of the spacecraft has a certain deviation and the ground station does not find the spacecraft at the theoretical station-entering moment, the working state of the spacecraft cannot be judged, and then the ground station is difficult to guide to develop subsequent work. For the ground station, the fewer and better the tracking guidance process is operated, if multiple waiting points are set for different deviation tracks, the ground station needs to frequently control the antenna to the different waiting points, which is rather unfavorable for tracking capture.
In order to solve the technical problems, the invention provides a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which comprises the following steps: for the spacecraft capable of being controlled autonomously, a control program is designed before the spacecraft is launched, a ground system simulates the spacecraft autonomous control program, and a spacecraft orbit dynamics model is combined to forecast the spacecraft autonomous control process, so that a nominal ephemeris is generated. In the autonomous control process of the spacecraft, the theoretical control efficiency deviates from the actual control efficiency, the deviation is called as control deviation, and the spacecraft development can estimate the control deviation range through a correlation test. In addition to the control deviation, the spacecraft is influenced by the space environment during the movement, and the orbit forecast based on dynamics has deviation, which is generally much smaller than the control deviation and can be converted into the control deviation. And according to the synthesized control deviation, estimating the deviation range of the orbit ephemeris during the autonomous control of the spacecraft.
The invention discloses a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which is implemented according to the following steps:
step one: and carrying out orbit forecasting on the autonomous control spacecraft according to the autonomous control program and the orbit dynamics model. The prediction result is called nominal ephemeris. Nominal guidance data, i.e. nominal arrival time and azimuth, pitch angle (pitch angle is typically a fixed value of 3 ° or 5 °) is calculated using the nominal ephemeris.
Under the Newton frame of reference, the equation of motion of the detector can be written as:
in the formula (1), the components are as follows,for the two gravitation and other various ingenuities of spacecraft, +.>For thrust force applied to autonomous control process of spacecraft, initial value state r of probe is given 0 And->The nominal ephemeris of the detector can be obtained by integrating the above formula.
Step two: and converting the errors of the space environment model and the dynamic model into control deviation, forecasting and generating a plurality of deviation ephemeris according to the range of the control deviation, and setting the total number of the deviation ephemeris as N.
Adding the controlled deviation to formula (1) gives formula (4):
in the formula (4), the amino acid sequence of the compound,for different thrust deviations. A plurality of offset ephemeris may be calculated according to equation (4).
Step three: calculating the elevation angle of the station to be E according to the nominal ephemeris and the station address of the ground station 0 Time corresponding epoch time T 0 And azimuth angle A 0 。
Defining station site coordinates [ x ] under the station-finding ground fixed system ei y ei z ei ]The spacecraft position isRectangular coordinates of spacecraft in horizon are +.>
Wherein B is θ 、The matrix is a conversion matrix from an inertial system to a ground system, and B is a conversion matrix from the ground system to a ground system.
(7) In (8), θ g The star angle is Greenner. Lambda (lambda) i 、The latitude and longitude of the geodetic center are measured.
Weighing T 0 Azimuth angle A for theoretical arrival time 0 Pitch angle E 0 Is the nominal inbound site.
Step four: and calculating the guide data deviation of the deviation ephemeris.
1) Calculating station azimuth angle A corresponding to ith deviation ephemeris i Pitch angle E i And interpolate E i ＝E 0 At the corresponding epoch time T i And azimuth angle A i . Calculating the departure time deviation DeltaT according to formulas (12) and (13) i And an inbound azimuth deviation ΔA i 。
ΔT i ＝T i -T 0 (12)，
ΔA i ＝A i -A 0 (13)，
If delta A i Less than the half-wave beam range of the antenna, recording the corresponding delta T of the offset ephemeris i ；
If delta A i Greater than or equal to the half-wave beam range of the antenna, then 2).
2) Interpolation calculation A i ＝A 0 At the corresponding epoch time T i And pitch angle E i . Calculating the departure time deviation DeltaT according to formulas (14) and (15) i And an approach pitch angle deviation ΔE i 。
ΔT i ＝T i -T 0 (14)，
ΔE i ＝E i -E 0 (15)，
If delta E i Less than the half-wave beam range of the antenna, recording the corresponding delta T of the offset ephemeris i 。
If delta E i If the antenna is larger than or equal to the half-wave beam range of the antenna, the ephemeris deviation is too large to be suitable for the method. The booting may be performed in a manner of setting a plurality of waiting points.
Step five: guide information is calculated.
The minimum and maximum arrival deviation times are calculated according to equations (16), (17).
ΔT min ＝min{ΔT i } (16)，
ΔT max ＝max{ΔT i } (17)，
The guidance information to the ground station may be noted as { T } 0 ,A 0 ,E 0 ,ΔT min ,ΔT max }。
Based on this information, the ground station antenna is at T 0 +ΔT min The antenna is pointed at the theoretical waiting point A 0 、E 0 At the earliest from T 0 +ΔT min Starting to wait for spacecraft to enter antenna tracking beam at moment and waiting to T at the latest 0 +ΔT max Time of day. If the ground station is at T 0 +ΔT max After the moment, no target is found yet, and the method can be used as one of the marks of the abnormal state of the spacecraft to carry out subsequent emergency treatment.
Project the spacecraft station measurement guiding calculation method based on the combined deviation orbit forecast is further described in detail through a specific embodiment;
the satellite for the experiment of the thin atmosphere science is automatically controlled by the propeller according to a pulse period of 1 minute on and 3 minutes off when the satellite and the rocket are separated for 441 minutes until the height Hrpm < = 180km of the flat near place. And calculating the guiding data of the measuring station 1 and the measuring station 2 in the 7 th turn, wherein the half-wave beam width of the antennas of the measuring station 1 and the measuring station 2 is 3.0 degrees.
Step one: calculating a nominal ephemeris;
1) The kinetic model used for orbital ephemeris calculation is shown in table 1:
table 1 initial epoch and kinetic parameters
Project Value of
Epoch time (BJT) 2016-8-16 9:15:43.254
SatelliteQuality (kg) 105.4
Atmospheric damping coefficient 2.2
Frontal area (m) 2 ) 0.223
Atmospheric model MISS 90
2) The control deviation working conditions comprise the following four types,
a) The normal thrust force is applied to the cylinder,
b) The thrust force is 0.8 times of that of the piston,
c) The thrust force is 1.2 times of that of the piston,
d) Unpowered extrapolation.
When the satellite attitude control, orbit control, GPS receiving equipment, software and the like have faults, the satellite can enter a stop control state, and the orbit prediction is unpowered extrapolation. The control error of + -20% is reduced by the comprehensive errors such as atmospheric error, control error, initial value error and the like.
TABLE 2 comparison of different control error forecast results (pitch angle 5 degree)
TABLE 3 comparison of different control errors with forecast results at nominal thrust
In conclusion, the method comprises the steps of,
the guidance information for the measuring station 1 is:
{2016-8-16 10:41:39.597,117.0878°,5°,-0.3s,0.8s}；
the guidance information for the measuring station 2 is:
{2016-8-16 10:42:36.449,165.4680°,5°,-4s,22.4s}
as shown in tables 2 and 3, the station 1 arrival time deviation ± 0.8s, the arrival azimuth deviation ± 2.0 ° were within the range of the pilot antenna beam at 0.8, 1.2 times error compared to the nominal case. If the satellite stops controlling in a whole circle, the station 1 enters the station with 8s of time deviation and 7.7 degrees of azimuth deviation; the station 2 is at the station arrival time deviation 23s and the station arrival azimuth deviation 0.7 deg.. Thus if the satellite is deactivated after the 6 th turn, the 7 th turn station 2 station approach azimuth deviation is within the beam range. The autonomous control effect of the spacecraft can be primarily judged according to the actual tracking time of the measuring station.
The invention relates to a spacecraft station measurement guiding calculation method based on combined deviation orbit prediction, which aims to complete a space-earth cooperative task during autonomous control of a spacecraft, a ground system is required to simulate the autonomous control process of the spacecraft and calculate the ephemeris of the spacecraft by considering various orbit prediction deviations, and the guiding number is calculated and the station measurement is guided to be captured through the deviation ephemeris. The method improves the reliability of operation and has better applicability and operability.

Claims (3)

1. A spacecraft station measurement guiding calculation method based on combined deviation orbit prediction is characterized by comprising the following steps:
step 1: carrying out orbit prediction on the autonomous control spacecraft according to an autonomous control program and an orbit dynamics model, marking a prediction result as a spacecraft nominal ephemeris, and calculating nominal guiding data, namely nominal arrival time, azimuth and pitch angle by using the nominal ephemeris;
step 2: converting the space environment deviation and the dynamic model error into control deviation, and forecasting to generate a plurality of deviation ephemeris according to the range of the control deviation;
step 3: calculating deviation guiding data according to the deviation ephemeris, and calculating an angle difference value and a time difference value between the deviation guiding data and a nominal station entering point;
if the angle difference of the guide data calculated by a certain deviation ephemeris is larger than or equal to the half-wave beam range of the antenna, the equipment cannot conduct guide capturing through the method under the deviation;
step 4: comprehensively calculating guide information according to the guide data of the deviation ephemeris;
the step 3 is specifically as follows: calculate the firstiStation tracking azimuth corresponding to strip deviation ephemerisPitch angle->And interpolate +.>At the corresponding epoch timeT i And azimuth->According to formulas (12) and (13), calculating the departure time deviation +.>And the incoming azimuth deviation->：
（12），
（13），
If it isLess than the half-wave beam range of the antenna, the corresponding +.>Step 4 is entered;
if it isGreater than or equal to the half-wave beam range of the antenna, interpolation is calculated +.>At the corresponding epoch timeT i And pitch angle->Calculating the departure time deviation and the departure azimuth deviation according to formulas (14) and (15)>；
（14），
（15），
If it isLess than the half-wave beam range of the antenna, the corresponding +.>Step 4 is entered;
if it isIf the ephemeris is larger than or equal to the half-wave beam range of the antenna, the ephemeris deviation is too large to be suitable for the method;
step 4, specifically, calculating the minimum and maximum deviation moments of the arrival according to the following formulas (16) and (17);
(16)，
(17)，
guidance information to ground stationsWherein->For theoretical arrival time, +.>、/>Wait for the antenna to be azimuth and elevation, respectively, +.>、/>Respectively the leading and trailing edges of the latency.
2. The spacecraft station surveying guidance calculation method based on combined deviation orbit prediction according to claim 1, wherein step 1 specifically comprises that under the newton reference frame, the motion equation of the detector can be written as the following formulas (1), (2) and (3):
under the Newton frame of reference, the equation of motion of the detector can be written as:
(1)，
(2)，
(3)，
in the formula (1),for spacecraft position->For the two gravitation and other various ingenuities of spacecraft, +.>For the thrust exerted during autonomous control of the spacecraft, the initial state of the detector is given +.>And->The nominal ephemeris of the detector can be obtained by integrating the above formula, and the nominal guidance data of the ground station can be calculated according to the nominal ephemeris.
3. The spacecraft station surveying guidance calculation method based on combined bias orbit prediction according to claim 2, wherein in step 2, the calculation of the plurality of bias ephemeris predictions is specifically:
set the total number of the deviation forecast ephemeris asN；
Adding the control deviation to equation (1) yields equation (4):
（4），
wherein, the liquid crystal display device comprises a liquid crystal display device,for different control offsets, a plurality of offset ephemeris are calculated according to equation (4).
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Citations (3)

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Publication number Priority date Publication date Assignee Title
JP2016180729A (en) * 2015-03-25 2016-10-13 国立大学法人 和歌山大学 Satellite tracking antenna device and satellite tracking method
JP2016223781A (en) * 2015-05-27 2016-12-28 三菱電機株式会社 Satellite tracking device
CN112540390A (en) * 2020-11-26 2021-03-23 陕西星邑空间技术有限公司 Tracking forecast calculation method and device for spacecraft

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8099186B2 (en) * 2006-12-22 2012-01-17 The Boeing Company Satellite navigation using long-term navigation information and autonomous orbit control

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2016180729A (en) * 2015-03-25 2016-10-13 国立大学法人 和歌山大学 Satellite tracking antenna device and satellite tracking method
JP2016223781A (en) * 2015-05-27 2016-12-28 三菱電機株式会社 Satellite tracking device
CN112540390A (en) * 2020-11-26 2021-03-23 陕西星邑空间技术有限公司 Tracking forecast calculation method and device for spacecraft

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