CN110209185A - A kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star - Google Patents
A kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star Download PDFInfo
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Abstract
A kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star of the present invention, include the following steps: that (1) carries out Attitude rate estimator and attitude prediction using gyro output integral；(2) carry out whether satellite is motor-driven judges in place；(3) if judging that satellite is not motor-driven in place in step (2), step (6) are transferred to；Otherwise judge whether to have carried out the quick posture resetting of star, if being transferred to step (4), be transferred to step (5) if not；(4) calculating of posture filter correction is carried out, the attitude quaternion after obtaining filter correction is transferred to step (6)；(5) it carries out determining the resetting calculating of appearance posture, the attitude quaternion after being reset is transferred to step (6)；(6) according to scheduled targeted attitude and target angular velocity, control calculating is carried out using current Attitude rate estimator value and Attitude estimation value, obtains control moment, and be applied on celestial body and carry out stability contorting.
Description
Technical field
The present invention relates to spacecraft attitude determination and control field, especially a kind of boat using the quick information posture resetting of star
Its device attitude maneuver antihunt means.
Background technique
Continuous promotion with satellite platform to control ability demand controls the fast reserve fast and stable of satellite and proposes
Increasingly higher demands.To realize fast reserve function, angular speed of satellite when motor-driven is larger, this makes fast reserve
Posture determination in journey is faced with and situation very different when stability contorting.Star sensor is difficult to mention in big angular speed first
For accurate attitude measurement data, or even attitude measurement data can not be provided.Secondly gyro in big angular speed output bias with
Estimated constant value drift has differences when stability contorting, on the one hand this species diversity is derived from the little deviation of gyro during installation,
On the other hand derived from output stability and constant multiplier of the gyro in different angular speed, there are deviations.In view of these feelings
Condition, satellite generally carries out posture without using star sensor during fast reserve and determines, and gyro output integral is used to carry out
Attitude rate estimator and attitude prediction.But in this way when satellite it is motor-driven in place when, influenced by gyro output bias, estimating posture will be with
There are certain deviations for true posture.Due to the presence of attitude misalignment, when satellite is transferred to stability contorting after motor-driven, need
Longer time eliminates posture determination deviation, is unable to complete the target of fast and stable.To solve the problems, such as this kind, need for boat
Its device attitude maneuver fast and stable proposes new solution.
Summary of the invention
A kind of the quick information of star is utilized to carry out technical problem solved by the present invention is having overcome the deficiencies of the prior art and provide
The method of posture resetting, by star sensor in small angular speed the higher advantage of attitude measurement accuracy and gyro moves when big angular speed
State responds faster advantage and combines, on the basis of original fast reserve process is integrated using gyro, it is motor-driven in place after
Satellite uses star sensor exact posture measurement data to top in the case where being in small angular speed state before being transferred to stability contorting
Spiral shell integral posture is reset, and the purpose for quickly eliminating gyro integral deviation is realized, after substantially reducing posture fast reserve
The stabilization time, realize fast and stable.
The technical solution of the invention is as follows: a kind of Spacecraft During Attitude Maneuver stabilization side using the quick information posture resetting of star
Method includes the following steps:
(1) Attitude rate estimator and attitude prediction are carried out using gyro output integral；
(2) carry out whether satellite is motor-driven judges in place；
(3) if judging that satellite is not motor-driven in place in step (2), step (6) are transferred to；Otherwise judge whether to have carried out star quick
Posture resetting, if being transferred to step (4), is transferred to step (5) if not；
(4) calculating of posture filter correction is carried out, the attitude quaternion after obtaining filter correction is transferred to step (6)；
(5) it carries out determining the resetting calculating of appearance posture, the attitude quaternion after being reset is transferred to step (6)；
(6) according to scheduled targeted attitude and target angular velocity, current Attitude rate estimator value and Attitude estimation value are used
Control calculating is carried out, control moment is obtained, and is applied on celestial body and carries out stability contorting.
Attitude prediction in the step (1) uses Eulerian angles attitude prediction or quaternary number attitude prediction.
Attitude rate estimator and attitude prediction are carried out using gyro output integral in step 1) method particularly includes:
Δ g=[Δ g is exported according to gyro1,Δg2,...,ΔgN]T, N >=3 control period Δ t, and the upper period determines appearance and provide
Quaternary number q=[q of the star ontology relative to inertial system1,q2,q3,q4]T, gyro often floats correction value b=[bx,by,bz]T, carry out
Attitude rate estimator and attitude prediction calculate, and obtain star ontology inertia angular velocity omega=[ωx,ωy,ωz]TWith the quaternary number of update
Discreet value q '；Wherein N is the gyro number determining appearance and using, Δ gi, i=1,2 ... N is the angle step of each gyro output
Value, subscript x, y, z are illustrated respectively in the component of star ontology x-axis, y-axis, z-axis；Specifically:
Star ontology inertia angular velocity omega=R ' Δ g/ Δ t-b is calculated, wherein R ' is related with gyro installation defeated from gyro
Output to magnitude of angular velocity transition matrix, its calculation formula isO in formulag1=[Ogix,
Ogiy,Ogiz], i=1,2 ... N be representation in components of each gyro to measure axis direction vector in star body coordinate system, subscript x, y,
Z is illustrated respectively in the component of star ontology x-axis, y-axis, z-axis；
Calculate the quaternary number discreet value updatedWherein Eq () is transfer function, is defined as
The whether motor-driven criterion in place of satellite is carried out in the step 2) specifically: whether arrived the intentional maneuver time,
Whether whether attitude prediction value and maneuvering target attitude error meet specified threshold, angular speed discreet value and target angular velocity error
Meet specified threshold.
According to the quick output q of stars=[qs1,qs2,qs3,qs4]TWith quaternary number discreet value q ' updated in step (1), carry out
Posture filter correction calculates, the attitude quaternion q " after obtaining filter correction.
The method for carrying out the calculating of posture filter correction uses Kalman filter algorithm.
Attitude quaternion after being reset in the step 5) method particularly includes: according to the quick output q of stars=[qs1,
qs2,qs3,qs4]TIt carries out determining the resetting calculating of appearance posture, attitude quaternion q " '=Qa (C after being resetsb TAq(qs))；Wherein
CsbIt is the transition matrix of the quick coordinate system from star body coordinate system to star, Aq () is the conversion letter that Direct cosine matrix is counted to from quaternary
Number, Qa () is the transfer function from Direct cosine matrix to quaternary number.
Current angular velocity estimated value in the step 6) is step 1) culminant star ontology inertia angular velocity omega；Current pose is estimated
Evaluation is determined according to the mode for being transferred to step 6), is the quaternary number discreet value q ' updated in step 1) when being transferred to from step 3), from
It is the attitude quaternion q " after filter correction in step 4) when step 4) is transferred to, after being reset when being transferred to from step 5) for step 5) is middle
Attitude quaternion q " '.
The specific acquisition methods of control moment restrain method using PID control in the step 6).
The present invention compared with the existing technology the advantages of be:
(1) present invention is by distinguishing attitude of satellite mobile process and stable control process and determining appearance using different sensors
The shortcomings that mode, precision is poor when having evaded star sensor wide-angle when motor-driven or even can not effectively export, it is defeated to play gyro
The high advantage of responsive bandwidth out, attitude measurement accuracy higher advantage when playing the small angular speed of star sensor in stability contorting,
It makes up gyro output stability and the shortcomings that there are long term drifts, accomplishes to maximize favourable factors and minimize unfavourable ones in the use of star sensor and gyro,
Have complementary advantages；
(2) present invention by it is motor-driven be transferred to stability contorting afterwards in place before satellite be in small angular speed state in the case where
Gyro integral posture is reset using star sensor exact posture measurement data, realizes and quickly eliminates gyro integral deviation
Purpose, the stabilization time after substantially reducing posture fast reserve, realize fast and stable.
Detailed description of the invention
Fig. 1 is a kind of Spacecraft During Attitude Maneuver antihunt means flow chart using the quick information posture resetting of star of the present invention；
Fig. 2 is practical posture (solid line) He Dingzi posture (dotted line) curve in the embodiment of the present invention；
Fig. 3 is to determine appearance error (solid line) curve in the embodiment of the present invention.
Specific embodiment
The present invention proposes a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star, as shown in Figure 1,
The method of the present invention includes that specific implementation process is as follows:
(1) Δ g=[Δ g is exported according to gyro1,Δg2,...,ΔgN]T, N >=3, period Δ t was controlled, the upper period determines appearance
Quaternary number q=[q of the star ontology provided relative to inertial system1,q2,q3,q4]T, gyro often floats correction value b=[bx,by,bz]T,
It carries out Attitude rate estimator and attitude prediction calculates, obtain star ontology inertia angular velocity omega=[ωx,ωy,ωz]TWith the four of update
First number discreet value q '.Wherein N is the gyro number determining appearance and using, Δ gi, i=1,2 ... N is that the angle of each gyro output increases
Magnitude, subscript x, y, z are illustrated respectively in the component of star ontology x-axis, y-axis, z-axis.Specifically:
Star ontology inertia angular velocity omega=R ' Δ g/ Δ t-b is calculated, wherein R ' is related with gyro installation defeated from gyro
Output to magnitude of angular velocity transition matrix, its calculation formula isO in formulag1=[Ogix,
Ogiy,Ogiz], i=1,2 ... N be representation in components of each gyro to measure axis direction vector in star body coordinate system, subscript x, y,
Z is illustrated respectively in the component of star ontology x-axis, y-axis, z-axis；
Calculate the quaternary number discreet value updatedWherein Eq () is transfer function, is defined as
(2) then carry out whether satellite motor-driven judges in place, judgment method can according to whether arrived the intentional maneuver time,
Whether whether attitude prediction value and maneuvering target attitude error meet specified threshold, angular speed discreet value and target angular velocity error
Meet the condition settings such as specified threshold；
(3) if judging that satellite is not motor-driven in place in step (2), step (6) are transferred to；Otherwise according to whether to have carried out star quick
Posture resetting, if being transferred to step (4), is transferred to step (5) if not；
(4) according to the quick output q of stars=[qs1,qs2,qs3,qs4]TWith the quaternary number discreet value q ' provided in step (1), into
Row posture filter correction calculates, the attitude quaternion q " after obtaining filter correction.Modification method uses Kalman filter algorithm.So
After be transferred to step (6)；
(5) according to the quick output q of stars=[qs1,qs2,qs3,qs4]TIt carries out determining the resetting calculating of appearance posture, the appearance after being reset
State quaternary number q " '=Qa (Csb TAq(qs)).Wherein CsbIt is the transition matrix of the quick coordinate system from star body coordinate system to star, Aq ()
It is the transfer function that Direct cosine matrix is counted to from quaternary, Qa () is the transfer function from Direct cosine matrix to quaternary number；Then turn
Enter step (6)；
(6) according to scheduled targeted attitude and target angular velocity, current Attitude rate estimator value and Attitude estimation value are used
Control calculating is carried out, calculation method is restrained using PID control, obtains control moment, and be applied on celestial body and carry out stability contorting.Its
Middle current angular velocity estimated value is step 1) culminant star ontology inertia angular velocity omega, and current pose estimated value is according to being transferred to step 6)
Mode determines, is the quaternary number discreet value q ' updated in step 1) when being transferred to from step 3), is step 4) when being transferred to from step 4)
Attitude quaternion q " after middle filter correction is the attitude quaternion q " ' after resetting in step 5) when being transferred to from step 5)
The method of the present invention is described in detail below with reference to embodiment.
Embodiment 1: certain satellite carries out motor-driven, utilization according to the invention from absolute orientation zero attitude to 45 degree of postures of rolling
The Spacecraft During Attitude Maneuver fast and stable method of the quick information posture resetting of star, the simulation result of mobile process and stable control process
As shown in Figure 2 to 3.Wherein Fig. 2 is practical posture (solid line) He Dingzi posture (dotted line) curve, and Fig. 3 is to determine appearance error (in fact
Line) curve.200s~228s satellite is in motor-driven stage, mobile process in the case where estimating using only gyro when star as shown in the figure
In determine appearance error and be gradually increased, it is motor-driven in place when roll axis error and have been approached 0.5 degree, appearance posture determine at the end of 228s is motor-driven and is realized
Posture resetting, three axis determine appearance error and decrease below 0.02 degree, avoid that due to determining appearance error is larger to be caused to stablize adjustment
Time longer problem, realizes fast and stable.
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 (9)
1. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star, it is characterised in that including walking as follows
It is rapid:
(1) Attitude rate estimator and attitude prediction are carried out using gyro output integral；
(2) carry out whether satellite is motor-driven judges in place；
(3) if judging that satellite is not motor-driven in place in step (2), step (6) are transferred to；Otherwise judge whether to have carried out the quick posture of star
Resetting, if being transferred to step (4), is transferred to step (5) if not；
(4) calculating of posture filter correction is carried out, the attitude quaternion after obtaining filter correction is transferred to step (6)；
(5) it carries out determining the resetting calculating of appearance posture, the attitude quaternion after being reset is transferred to step (6)；
(6) it according to scheduled targeted attitude and target angular velocity, is carried out using current Attitude rate estimator value and Attitude estimation value
Control calculates, and obtains control moment, and be applied on celestial body and carry out stability contorting.
2. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 1,
Be characterized in that: the attitude prediction in the step (1) uses Eulerian angles attitude prediction or quaternary number attitude prediction.
3. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 2,
It is characterized in that: carrying out Attitude rate estimator and attitude prediction using gyro output integral in step 1) method particularly includes:
Δ g=[Δ g is exported according to gyro1,Δg2,...,ΔgN]T, N >=3 control period Δ t, and the upper period determines the star that appearance provides
Quaternary number q=[q of the ontology relative to inertial system1,q2,q3,q4]T, gyro often floats correction value b=[bx,by,bz]T, carry out angle speed
Degree estimation and attitude prediction calculate, and obtain star ontology inertia angular velocity omega=[ωx,ωy,ωz]TIt is estimated with the quaternary number of update
Value q '；Wherein N is the gyro number determining appearance and using, Δ gi, i=1,2 ... N is the angle step value of each gyro output, under
X is marked, y, z are illustrated respectively in the component of star ontology x-axis, y-axis, z-axis；Specifically:
Star ontology inertia angular velocity omega=R ' Δ g/ Δ t-b is calculated, wherein R ' is related with gyro installation from gyro output quantity
To the transition matrix of magnitude of angular velocity, its calculation formula isR '=inv (RTR)RT, O in formulag1=[Ogix,Ogiy,
Ogiz], i=1,2 ... N is representation in components of each gyro to measure axis direction vector in star body coordinate system, subscript x, y, z point
It Biao Shi not be in the component of star ontology x-axis, y-axis, z-axis；
Calculate the quaternary number discreet value updatedWherein Eq () is transfer function, is defined as
4. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 1,
Be characterized in that: carrying out the whether motor-driven criterion in place of satellite in the step 2) specifically: whether arrived the intentional maneuver time,
Whether whether attitude prediction value and maneuvering target attitude error meet specified threshold, angular speed discreet value and target angular velocity error
Meet specified threshold.
5. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 3,
It is characterized in that: according to the quick output q of stars=[qs1,qs2,qs3,qs4]TWith quaternary number discreet value q ' updated in step (1), into
Row posture filter correction calculates, the attitude quaternion q " after obtaining filter correction.
6. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 5,
Be characterized in that: the method for carrying out the calculating of posture filter correction uses Kalman filter algorithm.
7. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 1,
It is characterized in that: attitude quaternion after being reset in the step 5) method particularly includes: according to the quick output q of stars=[qs1,
qs2,qs3,qs4]TIt carries out determining the resetting calculating of appearance posture, attitude quaternion q " '=Qa (C after being resetsb TAq(qs))；Wherein
CsbIt is the transition matrix of the quick coordinate system from star body coordinate system to star, Aq () is the conversion letter that Direct cosine matrix is counted to from quaternary
Number, Qa () is the transfer function from Direct cosine matrix to quaternary number.
8. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 1,
Be characterized in that: the current angular velocity estimated value in the step 6) is step 1) culminant star ontology inertia angular velocity omega；Current pose
Estimated value is determined according to the mode for being transferred to step 6), is the quaternary number discreet value q ' updated in step 1) when being transferred to from step 3),
It is the attitude quaternion q " after filter correction in step 4) when being transferred to from step 4), is resetting in step 5) when being transferred to from step 5)
Attitude quaternion q " ' afterwards.
9. a kind of Spacecraft During Attitude Maneuver antihunt means using the quick information posture resetting of star according to claim 1,
Be characterized in that: the specific acquisition methods of control moment restrain method using PID control in the step 6).
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