CN107826269A - 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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- B64G1/36—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, using earth sensor, inertial attitude gyro, orbit maneuver motor, in the case of height of entering the orbit after satellite transmitting deficiency, implement perigean Satellite Orbit Maneuver.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 change that drift compensation value control satellite operation pattern is injected to control computer, reach continuous change rail of the satellite in stationary orbit.It the method overcome that existing change rail technology propellant expenditure is excessive, and satellite perigee posture determines the problem of difficult, can easier establish firing attitude, has a whole day ball capture ability, reliability height.
Description
Technical field
The present invention relates to a kind of perigee orbit changing method suitable for geostationary orbit satellite platform, belong to satellite control
Field.
Background technology
In the launch mission of the geostationary orbit satellite platform of prior art, the view rail of star sensor is configured without
In the case that road satellite is introduced into planned orbit after transmitting, highly less than perigee, it is unsatisfactory for earth sensor and uses height
During limitation, the apogee of satellite becomes rail and realizes extremely difficult, can not realize that correctly igniting promotes and become rail.
In conventional launch mission both domestic and external, due to having when rocket failure fails satellite being sent into the situation of planned orbit
Occur.Such as Indonesia's Palapa-D satellites 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 launchs, rocket
Location higher is 200km, and altitude of the apogee is 16400km track, well below planned orbit height.If satellite can not
Stationary orbit is timely entered, then can not carry out regular traffic, it will causes huge economic loss.With Palapa-D satellites not
Together, CHINASAT-9A satellites are configured without star sensor, and this determines to bring very big difficulty to satellite perigee posture.
In the case of 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 cause to significantly affect on the service life of satellite.Together
When, when becoming rail at perigee, the satellite of unassembled star sensor, the capture ability without whole day ball, it is non-to establish firing attitude
Often difficult, loss of ignition, which easily occurs, in change rail operation to be caused to become rail failure.
The content of the invention
Present invention solves the technical problem that it is：The static rail to be broken down for being configured without star sensor or star sensor
Road satellite, it is introduced into after its transmitting in the case of planned orbit, there is provided 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 demarcation is carried out under earth search pattern, determines gyro
Constant value drift, control satellite raise 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 pattern is transferred to earth directing mode；
(4) satellite, which is transferred to after earth directing mode, carries out gyroscopic drift demarcation, completes after demarcation 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 specified location in orbit plane, the specified location be using centroid of satellite in
The heart, X-axis points to the satellite health longitudinal axis, and towards the direction of the earth when Z axis points to satellite motion, Y axles press the orthogonal rule definition of the right hand
The XOZ planes of the coordinate system of foundation and orbit plane parallel position；
(6) inject gyroscopic drift offset, satellite operation pattern be transferred to full gyro earth directing mode, satellite continue to
Apogee track rise；
(7) before satellite reaches apogee, pitch attitude biasing is carried out, establishes firing attitude；
(8) satellite is moved along track to perigee, and gyroscopic drift offset is re-injected when reaching near the track of perigee；
(9) satellite operation pattern is transferred to full gyro apogee pattern, satellite booster device ignition operation；
(10) after igniting terminates, control satellite operation pattern is transferred to sun acquisition pattern.
Gyroscopic drift demarcating steps are as follows in the step (1)：
(a) after satellite is transferred to earth search pattern, satellite pitch attitude biasing is arranged to 20 °, roll attitude is biased
0 ° is arranged to, records current time t after attitude stabilization10Sun sensor rolls pitch attitude output φ on moment-Z face10And θ10
And three axis accelerometer integration output
(b) flight time after 20 minutes, records current time t11Sun sensor rolls pitching appearance on moment-Z face
State exports φ11And θ11And three axis accelerometer integration output φRIGA11,θRIGA11,ψRIGA11；
(c) satellite pitch attitude biasing is now arranged to -20 ° again, roll attitude biasing is arranged to 0 °, records posture
Current time t after stable20Sun sensor rolls pitch attitude output φ on moment-Z face20And θ20And three axis accelerometer integration
Output
(d) flight time after 20 minutes, records current time t21Sun sensor rolls pitching appearance on moment-Z face
State exports φ21And θ21And three axis accelerometer integration output
Wherein, two constant value drift amounts being calculated according to above-mentioned formula average to obtain constant value drift amount Bgx0：
(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 caused drift of track constant value drift, orbit angular velocity in the step (6), specifically
Calculation formula is as follows：
The compensation value calculation formula is：
Wherein, drift calculation formula is caused by orbit angular velocity：
For track constant value drift；
CboFor the three-axis attitude angle of corresponding body series relative orbit systemTransition matrix, ωbyFor track system
Along Y-axis orbital drift angular speed, centered on centroid of satellite, Z axis points to the earth's core for its middle orbit system, and Y-axis is put down perpendicular to track
Face, X-axis judge according to the right-hand rule.
The sun acquisition pattern is that satellite searches for the sun in space, quick with 2, faces of 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, faces of 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 point to earth center direction, earth sensor
Measure roll angle, the angle of pitch, sun sensor measurement yaw angle, the axis angular rate of gyro to measure three.
The full gyro earth directing mode be measured using gyro integration carrying out satellite body coordinate system three-axis attitude,
Angular velocity measurement.
The full gyro apogee pattern is to carry out satellite body coordinate using gyro integration during engine ignition
It is three-axis attitude measurement and angular velocity measurement.
The present invention compared with prior art the advantages of be：
(1) the invention provides a kind of perigee orbit changing method suitable for satellite platform, make full use of and defend
The existing sensor configuration of star and application state, without any extra change, can enter under conditions of without star sensor
The change rail of row satellite, the operation such as demarcation, igniting, dramatically reduce implementation and become the preoperative preparation of rail and checking work, simultaneously
Without in-orbit software flight course is modified i.e. can be achieved satellite by perigee enter it is apogean change rail motion；
(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, realize Satellite Orbit Maneuver at perigee, reduce the consumption of propellant,
The life-span of satellite operation is extended, method is easily achieved, and has versatility and portability.
Brief description of the drawings
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；
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, angle of pitch input, and sun sensor can be used as rolling, pitching and yaw angle input, gyro
Tri-axis angular rate can be surveyed, gyro integration can be used as three-axis attitude angle to input.After transmitting, satellite transit is in elliptic orbit, such as Fig. 2 institutes
Show, the constraint of the elliptic orbit is：Altitude of the apogee must meet that earth sensor can use condition.
Need to complete following step as shown in figure 1, the present invention provides satellite platform perigee orbit changing method
Suddenly.
(1) after entering the orbit, control system is operated under sun acquisition pattern, 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, while enters under earth search pattern
Row gyroscopic drift is demarcated, and determines gyroscope constant value drift.
Gyroscope constant value drift computational methods are as follows：
2 hours before apogee is reached, satellite is transferred to earth search pattern by sun acquisition pattern.After attitude stabilization, if
Put satellite pitch attitude and bias 20 °, roll attitude biases 0 °, after waiting attitude stabilization, writes down now (t10Moment) on-Z faces too
Positive sensor rolls pitch attitude output φ10And θ10, after 20 minutes, then write down now (t11Moment) sun sensor on-Z faces
Roll pitch attitude output φ11And θ11And three axis accelerometer integration output φRIGA11,θRIGA11,ψRIGA11.The constant value drift of gyro
Bgx0, Bgy0, Bgz0There is following relation：
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
Attitude of satellite biasing is refilled, sets pitch attitude to bias -20 °, roll attitude biases 0 °, waits attitude stabilization
Afterwards, now (t is write down20Moment) sun sensor rolls pitch attitude output φ on-Z faces20And θ20And three axis accelerometer integration is defeated
Go outAfter 20 minutes, then write down now (t21Moment) sun sensor rolls pitch attitude on-Z faces
Export φ21And θ21And three axis accelerometer integration outputThe constant value drift B of gyrogx0, Bgy0, Bgz0Have
Following relation：
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)
Wherein, two constant value drift amounts being calculated according to above-mentioned formula average to obtain constant value drift amount Bgx0：
(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 that earth sensor can use condition, in such as Fig. 2 institutes
(the x shown0, y0, z0) before correspondence position, earth sensor start, the orbital position according to residing for satellite establishes angle of pitch biasing
And the axis angular rate biasing that rolls and go off course, proceed by earth search.
(3) after earth sensor searches earth signal, send instruction and mode of operation is transferred to earth directing mode.This
When roll, the angle of pitch measured by earth sensor, the Z axis of satellite points to the earth's core；Yaw angle is measured by sun sensor, is pointed to too
Positive direction.Such as (x in Fig. 20, y0, z0) shown in position.
(4) under earth directing mode, gyroscopic drift demarcation is carried out.After the completion of demarcation, compare ground gyroscopic drift demarcation
As a result, earth search pattern Gyro Calibration result and earth directing mode Gyro Calibration result.Now, 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 computational methods are as follows：
After three-axis attitude stabilization, the roll angle and the angle of pitch, too of the earth sensor measurement that every frame remote measurement 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 integration output is designated as respectivelyT is recorded by terminatingnCorresponding 4 gyros integration output point
It is not designated asConstant value drift B can be tried to achievegx0, 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 the sampling period, T=0.512.
(5) under earth directing mode, injection driftage sun posture coefficient of issuing an order, the XOZ faces of satellite body are adjusted to
In orbit plane, such as (x in Fig. 21, y1, z1) shown in position, now, y1It is just vertical with orbit plane.
(6) gyroscopic drift offset is injected, this offset, which includes gyro itself constant value drift and orbit angular velocity, to be influenceed.Turn
Enter full gyro earth directing mode, firing attitude is established in preparation.Now, satellite three-axis attitude angle integrates by gyro angular speed and surveyed
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：
Drifted about caused by orbit angular velocity and be
Turn sequence Eulerian anglesThe three-axis attitude angle of corresponding body series relative orbit system, then formula 7. in
By CboExpression formula is substituted into drift value caused by orbit angular velocity, is obtained：
Therefore, compensation value calculation formula is as follows：
(7) about 4~5 minutes before satellite reaches apogee, pitch attitude biasing is carried out, satellite is turn 90 degrees around-Y axles,
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 refers to
To track direction of advance ,+x refers to ground.
(8) gyroscopic drift offset is refilled, this offset only includes the constant value drift of gyro itself, and purpose is mainly
The attitude of satellite is set to keep constant in inertial coodinate system.
(9) as shown in Fig. 2 during the entire process of satellite moves to perigee from apogee, posture begins in inertial space
(x is maintained eventually2, y2, z2) sensing.Perigee satellite is reached near perigee, re-injects drift compensation value, this offset bag
Containing gyroscope constant value drift and orbit angular velocity.
(10) at engine ignition point, control system mode of operation is referred to pattern by full gyro and is transferred to full gyro apogee
Pattern, now, satellite three-axis attitude angle is still by gyro angular speed integral measurement.After thruster sinks to the bottom 4 minutes, orbit maneuver motor
Start to light a fire.
(11) after igniting terminates, satellite is transferred to sun acquisition pattern, it is ensured that energy security.Earth station starts to survey rail, calculates
Track after control, check 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 realize it is easy and effective, be very easy to engineering use, can promote the use in all kinds of satellites fly
In row control task.
The content not being described in detail in description of the invention belongs to the known technology 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 step is as follows：
(1) control satellite is transferred to earth search pattern, and gyroscopic drift demarcation is carried out under earth search pattern, determines gyroscope constant value
Drift, control satellite raise 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 pattern is transferred to earth directing mode；
(4) satellite, which is transferred to after earth directing mode, carries out gyroscopic drift demarcation, completes after demarcation relatively and records ground gyro drift
Move calibration result, earth search pattern Gyro Calibration result and earth directing mode Gyro Calibration result；
(5) control satellite body is adjusted to specified location in orbit plane, and the specified location is the X centered on centroid of satellite
Axle points to the satellite health longitudinal axis, and towards the direction of the earth when Z axis points to satellite motion, Y-axis is pressed the orthogonal rule definition of the right hand and established
Coordinate system XOZ planes and orbit plane parallel position；
(6) gyroscopic drift offset is injected, satellite operation pattern is transferred to full gyro earth directing mode, satellite continues to far
Point track rise；
(7) before satellite reaches apogee, pitch attitude biasing is carried out, establishes firing attitude；
(8) satellite is moved along track to perigee, and gyroscopic drift offset is re-injected when reaching near the track of perigee；
(9) satellite operation pattern is transferred to full gyro apogee pattern, satellite booster device ignition operation；
(10) after igniting terminates, control satellite operation pattern is transferred to sun acquisition pattern.
2. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, it is special
Sign is：Gyroscopic drift demarcating steps are as follows in the step (1)：
(a) after satellite is transferred to earth search pattern, satellite pitch attitude biasing is arranged to 20 °, roll attitude is biased and set
For 0 °, current time t after attitude stabilization is recorded10Sun sensor rolls pitch attitude output φ on moment-Z face10And θ10And
Three axis accelerometer integration output φRIGA10,θRIGA10,
(b) flight time after 20 minutes, records current time t11It is defeated to roll pitch attitude for sun sensor on moment-Z face
Go out φ11And θ11And three axis accelerometer integration output φRIGA11,θRIGA11,ψRIGA11；
(c) satellite pitch attitude biasing is now arranged to -20 ° again, roll attitude biasing is arranged to 0 °, records attitude stabilization
Current time t afterwards20Sun sensor rolls pitch attitude output φ on moment-Z face20And θ20And three axis accelerometer integration output
φRIGA20,θRIGA20,
(d) flight time after 20 minutes, records current time t21It is defeated to roll pitch attitude for sun sensor on moment-Z face
Go out φ21And θ21And three axis accelerometer integration output φRIGA21,θRIGA21,
Wherein, two constant value drift amounts being calculated according to above-mentioned formula average to obtain constant value drift amount Bgx0：
(t11-t10)Bgy0=(θRIGA11-θRIGA10)-(θ11-θ10)
(t21-t20)Bgy0=(θRIGA21-θRIGA20)-(θ21-θ20)；
Calculate constant value drift amount Bgx0, Bgz0Formula is as follows：
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<mi>sin&theta;</mi>
<mn>21</mn>
</msub>
<mo>-</mo>
<mfrac>
<mrow>
<msub>
<mi>&phi;</mi>
<mn>21</mn>
</msub>
<mo>-</mo>
<msub>
<mi>&phi;</mi>
<mn>20</mn>
</msub>
</mrow>
<mrow>
<msub>
<mi>cos&theta;</mi>
<mn>21</mn>
</msub>
</mrow>
</mfrac>
<mo>.</mo>
</mrow>
3. a kind of perigee orbit changing method suitable for geostationary orbit satellite platform according to claim 1, it is special
Sign is：Gyroscopic drift offset includes the caused drift of track constant value drift, orbit angular velocity in the step (6), specifically
Calculation formula is as follows：
The compensation value calculation formula is：
<mrow>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>x</mi>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>y</mi>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>z</mi>
</mrow>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>=</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>x</mi>
<mn>0</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>y</mi>
<mn>0</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>z</mi>
<mn>0</mn>
</mrow>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>+</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>x</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>y</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>z</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>.</mo>
</mrow>
Wherein, drift calculation formula is caused by orbit angular velocity：
<mrow>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>x</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>y</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>B</mi>
<mrow>
<mi>g</mi>
<mi>z</mi>
<mn>1</mn>
</mrow>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>=</mo>
<msub>
<mi>C</mi>
<mrow>
<mi>b</mi>
<mi>o</mi>
</mrow>
</msub>
<mo>&CenterDot;</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<mn>0</mn>
</mtd>
</mtr>
<mtr>
<mtd>
<mo>-</mo>
<msub>
<mi>&omega;</mi>
<mrow>
<mi>b</mi>
<mi>y</mi>
</mrow>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<mn>0</mn>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>;</mo>
</mrow>
For track constant value drift；
CboFor the three-axis attitude angle of corresponding body series relative orbit systemθ, ψ transition matrix, ωbyIt is track system along Y-axis track
Drift about angular speed, and its middle orbit system is centered on centroid of satellite, and Z axis points to the earth's core, and Y-axis is perpendicular to orbit plane, and X-axis is according to the right side
Hand rule judges.
4. a kind of perigee suitable for geostationary orbit satellite platform according to claim 1 or 2 or 3 becomes rail side
Method, it is characterised in that：The sun acquisition pattern is that satellite searches for the sun in space, with 2, satellite body coordinate system-Z faces
The output signal of sun sensor is used as three axles respectively as the Angle Position feedback signal of the axis of rolling and pitch axis using gyro signal
Speed feedback signals.
5. a kind of perigee suitable for geostationary orbit satellite platform according to claim 1 or 2 or 3 becomes rail side
Method, it is characterised in that：The earth search pattern is that satellite searches for the earth in space, with 2, satellite body coordinate system-Z faces
The output signal of sun sensor is used as three axles respectively as the Angle Position feedback signal of the axis of rolling and pitch axis using gyro signal
Speed feedback signals.
6. a kind of perigee suitable for geostationary orbit satellite platform according to claim 1 or 2 or 3 becomes rail side
Method, it is characterised in that：The earth directing mode satellite body coordinate system-Z axis maintains to point to earth center direction, the earth
Sensor measurement roll angle, the angle of pitch, sun sensor measurement yaw angle, the axis angular rate of gyro to measure three.
7. a kind of perigee suitable for geostationary orbit satellite platform according to claim 1 or 2 or 3 becomes rail side
Method, it is characterised in that：The full gyro earth directing mode is to carry out the axle appearance of satellite body coordinate system three using gyro integration
State measurement, angular velocity measurement.
8. a kind of perigee suitable for geostationary orbit satellite platform according to claim 1 or 2 or 3 becomes rail side
Method, it is characterised in that：The full gyro apogee pattern is to carry out satellite sheet using gyro integration during engine ignition
Body coordinate system three-axis attitude measures and angular velocity measurement.
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| CN109460049A (en) * | 2018-11-14 | 2019-03-12 | 北京控制工程研究所 | Geo-synchronous orbit satellite apogee orbit changing method based on inertia directing mode |
| CN111077512A (en) * | 2019-11-26 | 2020-04-28 | 歌尔股份有限公司 | TOF module calibration method and system |
| CN111637876A (en) * | 2020-05-15 | 2020-09-08 | 北京控制工程研究所 | Implementation method of high-bandwidth high-precision rate integral gyro simulator |
| CN113415441A (en) * | 2021-06-29 | 2021-09-21 | 北京控制工程研究所 | Gas-liquid mixing variable thrust emergency orbit control method for geostationary orbit satellite |
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| CN113415441A (en) * | 2021-06-29 | 2021-09-21 | 北京控制工程研究所 | Gas-liquid mixing variable thrust emergency orbit control method for geostationary orbit satellite |
| CN113968360A (en) * | 2021-08-09 | 2022-01-25 | 中国空间技术研究院 | Satellite autonomous electric propulsion orbit transfer method for stationary orbit satellite |
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| CN113739772A (en) * | 2021-08-30 | 2021-12-03 | 北京控制工程研究所 | Fire point sensor longitude and latitude measuring method |
| CN113739772B (en) * | 2021-08-30 | 2023-09-29 | 北京控制工程研究所 | Fire sensor longitude and latitude measuring method |
| CN113830332A (en) * | 2021-10-09 | 2021-12-24 | 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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