CN107826269B - A kind of perigee orbit changing method suitable for geostationary orbit satellite platform - Google Patents
A kind of perigee orbit changing method suitable for geostationary orbit satellite platform Download PDFInfo
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64G1/242—Orbits and trajectories
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
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- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
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Abstract
A kind of perigee orbit changing method suitable for satellite platform in the insufficient situation of height of entering the orbit after satellite transmitting, implements perigean Satellite Orbit Maneuver using earth sensor, inertial attitude gyro, orbit maneuver motor.Point nearby captures the earth in distant first, earth sensor is activated by way of being transferred to posture over the ground and obtaining gyro sieve calibration result, by the variation to control computer injection drift compensation value control satellite operation mode, reach continuous change rail of the satellite in stationary orbit.The method overcome existing change rail technology propellant expenditure is excessive, satellite perigee posture determines difficult problem, can easier establish firing attitude, has whole day ball capture ability, high reliablity.
Description
Technical field
The present invention relates to a kind of perigee orbit changing methods suitable for geostationary orbit satellite platform, belong to satellite control
Field.
Background technique
In the launch mission of the geostationary orbit satellite platform of the prior art, the view rail of star sensor is not configured
In the case that road satellite does not enter planned orbit after transmitting, highly it is lower than perigee, is unsatisfactory for earth sensor and uses height
When limitation, the apogee change rail realization of satellite is extremely difficult, cannot achieve correctly igniting and promotes change rail.
In previous launch mission both domestic and external, due to having when rocket failure fails the case where satellite is sent into planned orbit
Occur.Such as Indonesia's Palapa-D satellite of transmitting in 2009, due to rocket failure, apogee only has 2.1 ten thousand kilometers, by implementing
Multiple perigee becomes rail strategy, just realizes fixed point.This time satellite is only fed into closely by CHINASAT-9A satellite launch, rocket
Location higher is 200km, and altitude of the apogee is the track of 16400km, well below planned orbit height.If satellite cannot and
When enter stationary orbit, then can not carry out regular traffic, it will cause huge economic loss.It is different from Palapa-D satellite,
CHINASAT-9A satellite does not configure star sensor, this brings very big difficulty to the determination of satellite perigee posture.
When satellite is not sent into planned orbit by rocket, satellite can be sent to ground using the orbit maneuver motor of itself
Ball stationary orbit, but satellite need to additionally consume more propellants.This will significantly affect the service life of satellite.Together
When, when perigee becomes rail, it is non-to establish firing attitude for the satellite of unassembled star sensor, the capture ability without whole day ball
Often difficult, change rail operation, which is easy to appear loss of ignition, to be caused to become rail failure.
Summary of the invention
Technical problem solved by the present invention is for the static rail to break down without configuration star sensor or star sensor
Road satellite when not entering planned orbit after its transmitting, provides a kind of perigee suitable for satellite platform
Orbit changing method.
The present invention solves above-mentioned technical problem and is achieved by following technical solution:
A kind of perigee orbit changing method suitable for geostationary orbit satellite platform, this method operating procedure are as follows:
(1) control satellite is transferred to earth search pattern, and gyroscopic drift calibration is carried out under earth search pattern, determines gyro
Constant value drift, control satellite are increased to apogee track；
(2) earth search is carried out after satellite reaches earth sensor available track segmental arc height；
(3) after earth sensor searches earth signal, control satellite operation mode is transferred to earth directing mode；
(4) satellite, which is transferred to after earth directing mode, carries out gyroscopic drift calibration, completes after calibration relatively and records ground top
Spiral shell drift calibration result, earth search pattern Gyro Calibration result and earth directing mode Gyro Calibration result；
(5) control satellite body is adjusted to designated position in orbit plane, and the designated position, which is with centroid of satellite, is
The heart, X-axis are directed toward the satellite health longitudinal axis, the direction towards earth when Z axis is directed toward satellite motion, and Y-axis presses the orthogonal rule definition of the right hand
The XOZ plane and orbit plane parallel position of the coordinate system of foundation；
(6) inject gyroscopic drift offset, satellite operation mode be transferred to full gyro earth directing mode, satellite continue to
Apogee track increases；
(7) before satellite reaches apogee, pitch attitude biasing is carried out, firing attitude is established；
(8) satellite is moved along track to perigee, re-injects gyroscopic drift offset when reaching near the track of perigee；
(9) satellite operation mode is transferred to full gyro apogee mode, satellite booster device ignition operation；
(10) after lighting a fire, control satellite operation mode is transferred to sun acquisition mode.
Gyroscopic drift demarcating steps are as follows in the step (1):
(a) after satellite is transferred to earth search pattern, 20 ° is set by satellite pitch attitude biasing, roll attitude is biased
It is set as 0 °, records current time t after attitude stabilization10Sun sensor rolls pitch attitude and exports φ on moment-Z face10And θ10
And three axis accelerometer integral output
(b) flight time is by after twenty minutes, recording current time t11Sun sensor rolls pitching appearance on moment-Z face
State exports φ11And θ11And three axis accelerometer integral output φRIGA11,θRIGA11,ψRIGA11；
(c) -20 ° are set by satellite pitch attitude biasing again at this time, roll attitude biasing is set as 0 °, records posture
Current time t after stabilization20Sun sensor rolls pitch attitude and exports φ on moment-Z face20And θ20And three axis accelerometer integral
Output
(d) flight time is by after twenty minutes, recording current time t21Sun sensor rolls pitching appearance on moment-Z face
State exports φ21And θ21And three axis accelerometer integral output
Wherein, it averages to obtain constant value drift amount B according to two constant value drift amounts that above-mentioned formula is calculatedgx0:
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)；
Calculate constant value drift amount Bgx0, Bgz0Formula is as follows:
Gyroscopic drift offset includes the drift of track constant value drift, orbit angular velocity generation in the step (6), specifically
Calculation formula is as follows:
The compensation value calculation formula are as follows:
Wherein, the drift calculation formula that orbit angular velocity generates are as follows:
For track constant value drift；
CboFor the triaxial attitude angle of corresponding this system relative orbit systemTransition matrix, ωbyFor track system
Along Y-axis orbital drift angular speed, centered on centroid of satellite, Z axis is directed toward the earth's core for middle orbit system, Y-axis perpendicular to orbit plane,
X-axis judges according to the right-hand rule.
The sun acquisition mode is that satellite searches for the sun in space, quick with 2, the face the satellite body coordinate system-Z sun
The output signal of sensor respectively as the axis of rolling and pitch axis Angle Position feedback signal, it is anti-using gyro signal as three axle speed rates
Feedback signal.
The earth search pattern is that satellite searches for the earth in space, quick with 2, the face the satellite body coordinate system-Z sun
The output signal of sensor respectively as the axis of rolling and pitch axis Angle Position feedback signal, it is anti-using gyro signal as three axle speed rates
Feedback signal.
The earth directing mode satellite body coordinate system-Z axis maintains to be directed toward earth center direction, earth sensor
Roll angle, pitch angle are measured, sun sensor measures yaw angle, three axis angular rate of gyro to measure.
The full gyro earth directing mode be carried out using gyro integral satellite body coordinate system three-axis attitude measurement,
Angular velocity measurement.
The full gyro apogee mode is to carry out satellite body coordinate using gyro integral during engine ignition
It is three-axis attitude measurement and angular velocity measurement.
The advantages of the present invention over the prior art are that:
(1) it the present invention provides a kind of perigee orbit changing method suitable for satellite platform, makes full use of and defends
The configuration of star existing sensor and application state, without any additional change, can under conditions of being not necessarily to star sensor into
The change rail of row satellite, calibration, the operation such as igniting dramatically reduce implementation and become the preoperative preparation of rail and verifying work, simultaneously
Modify without in-orbit software to flight course can be realized satellite by perigee into apogean change rail move；
(2) present invention during completing to become rail, passes through note in the case where satellite is not sent into planned orbit by rocket
Enter constant value drift amount, reduce the change rail strategy of orbit inclination angle, realizes Satellite Orbit Maneuver at perigee, reduce the consumption of propellant,
The service life of satellite operation is extended, method is easily achieved, and has versatility and portability.
Detailed description of the invention
Fig. 1 is the satellite perigee orbit changing method flow chart that invention provides；
Fig. 2 is the satellite orbital position and pose adjustment schematic diagram that invention provides；
Specific embodiment
Satellite is configured with earth sensor, inertial attitude sensor (gyro), sun sensor and orbit maneuver motor；Software
On, earth sensor can be used as rolling, pitch angle input, and sun sensor can be used as rolling, pitching and yaw angle input, gyro
Tri-axis angular rate can be surveyed, gyro integral can be used as triaxial attitude angle input.After transmitting, satellite transit is in elliptic orbit, such as Fig. 2 institute
Show, the constraint of the elliptic orbit is: altitude of the apogee, which must meet earth sensor, can use condition.
As shown in Figure 1, the present invention, which provides satellite platform perigee orbit changing method, needs to complete following step
Suddenly.
(1) after entering the orbit, control system works under sun acquisition mode, keeps Direct to the sun.To be established a little using the earth
Fiery posture simultaneously obtains accurate gyroscope constant value drift, and the first step is transferred to earth search pattern, at the same under earth search pattern into
Row gyroscopic drift calibration, and determine gyroscope constant value drift.
Gyroscope constant value drift calculation method is as follows:
2 hours before reaching apogee, satellite is transferred to earth search pattern by sun acquisition mode.After attitude stabilization, if
It sets satellite pitch attitude and biases 20 °, roll attitude biases 0 °, after waiting attitude stabilization, writes down (t at this time10Moment) on the face-Z too
Positive sensor rolls pitch attitude and exports φ10And θ10, after twenty minutes, then write down (t at this time11Moment) sun sensor on the face-Z
It rolls pitch attitude and exports φ11And θ11And three axis accelerometer integral output φRIGA11,θRIGA11,ψRIGA11.The constant value drift of gyro
Bgx0, Bgy0, Bgz0There is following relationship:
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
Attitude of satellite biasing is refilled, setting pitch attitude biases -20 °, and roll attitude biases 0 °, waits attitude stabilization
Afterwards, (t at this time is write down20Moment) sun sensor rolls pitch attitude and exports φ on the face-Z20And θ20And three axis accelerometer integral is defeated
OutAfter twenty minutes, then (t at this time is write down21Moment) sun sensor rolls pitch attitude on the face-Z
Export φ21And θ21And three axis accelerometer integral outputThe constant value drift B of gyrogx0, Bgy0, Bgz0Have
Following relationship:
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)
Wherein, it averages to obtain constant value drift amount B according to two constant value drift amounts that above-mentioned formula is calculatedgx0:
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)；
Calculate constant value drift amount Bgx0, Bgz0Formula is as follows:
(2) close near apogee, determine that the segmental arc orbit altitude meets earth sensor and can use condition, in such as Fig. 2 institute
(the x shown0, y0, z0) before corresponding position, earth sensor booting, the orbital position according to locating for satellite establishes pitch angle biasing
And axis angular rate biasing is rolled and yaws, start to carry out earth search.
(3) after earth sensor searches earth signal, instruction is sent by operating mode and is transferred to earth directing mode.This
When roll, pitch angle measured by earth sensor, the Z axis of satellite is directed toward the earth's core；Yaw angle is measured by sun sensor, is directed toward too
Positive direction.Such as (x in Fig. 20, y0, z0) shown in position.
(4) under earth directing mode, gyroscopic drift calibration is carried out.After the completion of calibration, compare ground gyroscopic drift calibration
As a result, earth search pattern Gyro Calibration result and earth directing mode Gyro Calibration result.At this point, can be to gyroscope constant value drift
Progress is further accurate, should be consistent with acquired results in step (1) if flight control process is normal, and calculation method is as follows:
After three-axis attitude stabilization, the roll angle and pitch angle, too of the earth sensor measurement that every frame telemetering passes down are recorded
Positive sensor output, measurement moment.The sampling period of earth sensor is 0.512 second.By start recording moment t0Corresponding 4
Gyro integral output is denoted as respectivelyT is recorded by terminatingnCorresponding 4 gyros integral output difference
It is denoted asConstant value drift B can be acquiredgx0, Bgy0, Bgz0。
In formula
Wherein, ωoxi, ωoyi, ωoziThe projection for being orbit angular velocity in body coordinate system, ωbxi, ωbyi, ωbziFor
Satellite body angular speed, T are sampling period, T=0.512.
(5) under earth directing mode, the face XOZ of satellite body is adjusted to by injection yaw sun posture coefficient of issuing an order
In orbit plane, such as (x in Fig. 21, y1, z1) shown in position, at this point, y1It is just vertical with orbit plane.
(6) gyroscopic drift offset is injected, this offset includes that gyro itself constant value drift and orbit angular velocity influence.Turn
Enter full gyro earth directing mode, firing attitude is established in preparation.It is surveyed at this point, satellite triaxial attitude angle is integrated by gyro angular speed
Amount.It need to ensure in (4)~(6) operation implementation procedure, the exportable correct posture of earth sensor.
Wherein, gyroscopic drift compensation value calculation method is as follows:
Orbit angular velocity generate drift be
Turn sequence Eulerian anglesThe triaxial attitude angle of corresponding this system relative orbit system, then formula 7. in
By CboExpression formula substitutes into the drift value that orbit angular velocity generates, and obtains:
Therefore, compensation value calculation formula is as follows:
(7) about 4~5 minutes before satellite reaches apogee, pitch attitude biasing is carried out, turn 90 degrees satellite around-Y-axis,
Firing attitude is established in completion.When satellite reaches apogee, just at (x in Fig. 22, y2, z2) shown in posture, i.e. ,-z-axis be directed toward
Track direction of advance ,+x refer to ground.
(8) gyroscopic drift offset is refilled, this offset only includes the constant value drift of gyro itself, and purpose is mainly
Remain unchanged the attitude of satellite in inertial coodinate system.
(9) as shown in Fig. 2, posture begins in inertial space during the entire process of satellite moves to perigee from apogee
(x is maintained eventually2, y2, z2) direction.Perigee satellite reaches near perigee, re-injects drift compensation value, this offset packet
Containing gyroscope constant value drift and orbit angular velocity.
(10) at engine ignition point, control system operating mode is referred to that mode is transferred to full gyro apogee by full gyro
Mode, at this point, satellite triaxial attitude angle is still by gyro angular speed integral measurement.After thruster sinks to the bottom 4 minutes, orbit maneuver motor is opened
Initial point fire.
(11) after lighting a fire, satellite is transferred to sun acquisition mode, it is ensured that energy security.Earth station starts to survey rail, calculates
Track after control checks perigee transfer orbital control effect.
The present invention relies only on earth sensor, gyro, sun sensor and orbit maneuver motor, is without in-orbit software modification
It can be achieved.Method realization is simple and effective, greatly facilitates engineering use, can promote the use and fly in all kinds of satellites
In row control task.
The content that description in the present invention is not described in detail belongs to the well-known technique of those skilled in the art.
Claims (8)
1. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform, it is characterised in that steps are as follows:
(1) control satellite is transferred to earth search pattern, and gyroscopic drift calibration is carried out under earth search pattern, determines gyroscope constant value
Drift, control satellite are increased to apogee track；
(2) earth search is carried out after satellite reaches earth sensor available track segmental arc height；
(3) after earth sensor searches earth signal, control satellite operation mode is transferred to earth directing mode；
(4) satellite, which is transferred to after earth directing mode, carries out gyroscopic drift calibration, completes after calibration relatively and records the drift of ground gyro
Move calibration result, earth search pattern Gyro Calibration result and earth directing mode Gyro Calibration result；
(5) control satellite body is adjusted to designated position in orbit plane, and the designated position is the X centered on centroid of satellite
Axis is directed toward the satellite health longitudinal axis, the direction towards earth when Z axis is directed toward satellite motion, and Y-axis is pressed the orthogonal rule definition of the right hand and established
Coordinate system XOZ plane and orbit plane parallel position；
(6) gyroscopic drift offset is injected, satellite operation mode is transferred to full gyro earth directing mode, satellite continues to far
Point track increases；
(7) before satellite reaches apogee, pitch attitude biasing is carried out, firing attitude is established；
(8) satellite is moved along track to perigee, re-injects gyroscopic drift offset when reaching near the track of perigee；
(9) satellite operation mode is transferred to full gyro apogee mode, satellite booster device ignition operation；
(10) after lighting a fire, control satellite operation mode is transferred to sun acquisition mode.
2. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: gyroscopic drift demarcating steps are as follows in the step (1):
(a) after satellite is transferred to earth search pattern, 20 ° is set by satellite pitch attitude biasing, roll attitude is biased and is arranged
It is 0 °, records current time t after attitude stabilization10Sun sensor rolls pitch attitude and exports φ on moment-Z face10And θ10And
Three axis accelerometer integral output φRIGA10,θRIGA10,
(b) flight time is by after twenty minutes, recording current time t11It is defeated to roll pitch attitude for sun sensor on moment-Z face
φ out11And θ11And three axis accelerometer integral output φRIGA11,θRIGA11,
(c) -20 ° are set by satellite pitch attitude biasing again at this time, roll attitude biasing is set as 0 °, records attitude stabilization
Current time t afterwards20Sun sensor rolls pitch attitude and exports φ on moment-Z face20And θ20And three axis accelerometer integral output
φRIGA20,θRIGA20,
(d) flight time is by after twenty minutes, recording current time t21It is defeated to roll pitch attitude for sun sensor on moment-Z face
φ out21And θ21And three axis accelerometer integral output φRIGA21,θRIGA21,
Wherein, it averages to obtain constant value drift amount B according to two constant value drift amounts that above-mentioned formula is calculatedgx0:
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)；
Calculate constant value drift amount Bgx0, Bgz0Formula is as follows:
3. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: gyroscopic drift offset includes the drift of track constant value drift, orbit angular velocity generation in the step (6), specifically
Calculation formula is as follows:
The compensation value calculation formula are as follows:
Wherein, the drift calculation formula that orbit angular velocity generates are as follows:
For track constant value drift；
CboFor the triaxial attitude angle of corresponding this system relative orbit systemThe transition matrix of θ, ψ, ωbyIt is track system along Y-axis track
Drift about angular speed, and middle orbit system is centered on centroid of satellite, and Z axis is directed toward the earth's core, and Y-axis is perpendicular to orbit plane, and X-axis is according to the right side
Hand rule judgement.
4. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: the sun acquisition mode is that satellite searches for the sun in space, with 2, the face satellite body coordinate system-Z sun sensitivity
The output signal of device respectively as the axis of rolling and pitch axis Angle Position feedback signal, using gyro signal as three axis Rate Feedbacks
Signal.
5. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: the earth search pattern is that satellite searches for the earth in space, with 2, the face satellite body coordinate system-Z sun sensitivity
The output signal of device respectively as the axis of rolling and pitch axis Angle Position feedback signal, using gyro signal as three axis Rate Feedbacks
Signal.
6. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: the earth directing mode satellite body coordinate system-Z axis maintains to be directed toward earth center direction, earth sensor measurement
Roll angle, pitch angle, sun sensor measure yaw angle, three axis angular rate of gyro to measure.
7. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: the full gyro earth directing mode be carried out using gyro integral satellite body coordinate system three-axis attitude measurement,
Angular velocity measurement.
8. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, special
Sign is: the full gyro apogee mode is to carry out satellite body coordinate system using gyro integral during engine ignition
Three-axis attitude measurement and angular velocity measurement.
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| CN108820256A (en) * | 2018-06-23 | 2018-11-16 | 中国卫星海上测控部 | A kind of spacecraft bias track design method |
| CN109460049B (en) * | 2018-11-14 | 2021-11-16 | 北京控制工程研究所 | Geosynchronous orbit satellite apogee orbit transfer method based on inertial pointing mode |
| CN111077512B (en) * | 2019-11-26 | 2023-12-26 | 歌尔光学科技有限公司 | TOF module calibration method and system |
| CN111637876B (en) * | 2020-05-15 | 2021-12-07 | 北京控制工程研究所 | Implementation method of high-bandwidth high-precision rate integral gyro simulator |
| CN113415441B (en) * | 2021-06-29 | 2022-12-13 | 北京控制工程研究所 | Gas-liquid mixing variable thrust emergency orbit control method for geostationary orbit satellite |
| CN113968360B (en) * | 2021-08-09 | 2024-03-15 | 中国空间技术研究院 | Autonomous electric propulsion orbital transfer method on stationary orbit satellite |
| CN113739772B (en) * | 2021-08-30 | 2023-09-29 | 北京控制工程研究所 | Fire sensor longitude and latitude measuring method |
| CN113830332B (en) * | 2021-10-09 | 2023-04-07 | 北京控制工程研究所 | Ignition attitude establishment and dynamic tracking method for electric propulsion orbit transfer |
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| CN102424116A (en) * | 2011-12-08 | 2012-04-25 | 中国空间技术研究院 | Method for optimizing orbital transfer strategy of geostationary orbit satellite |
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