CN107839903B - Method for estimating air bleeding time of sailboard in transfer orbit section of single-wing satellite - Google Patents
Method for estimating air bleeding time of sailboard in transfer orbit section of single-wing satellite Download PDFInfo
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- CN107839903B CN107839903B CN201710876495.8A CN201710876495A CN107839903B CN 107839903 B CN107839903 B CN 107839903B CN 201710876495 A CN201710876495 A CN 201710876495A CN 107839903 B CN107839903 B CN 107839903B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
- B64G1/443—Photovoltaic cell arrays
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
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Abstract
The invention provides a deflation time estimation method for sailboards in a transfer orbit section of a single-wing satellite, which comprises the following steps of: acquiring timing data of angular velocity and a thruster; selecting iteration time; step three, summing moment integrals; step four, resolving the disturbance torque; and step five, estimating the air bleeding duration of the sailboard. The method can utilize the satellite on-orbit attitude angular velocity and the working time length data of the thruster, and calculate the interference torque generated by the deflation of the solar sailboard in real time based on the basic principle of satellite attitude dynamics, so as to identify the deflation duration of the solar sailboard.
Description
Technical Field
The invention relates to an air bleeding time estimation method, in particular to an air bleeding time estimation method for a sailboard of a transfer orbit section of a single-wing satellite.
Background
With the development of aerospace technology, the types and the number of loads configured by the satellite are more and more, and particularly, the radiation refrigeration surfaces of the loads of some remote sensing satellites cannot be shielded, so that the satellite can only select a single-wing sailboard configuration. One surface of the solar sailboard is a triple-junction gallium arsenide or silicon solar cell, and the surface of the solar cell faces the sun when the solar sailboard is oriented to the sun; the other side is an aluminum honeycomb sandwich structure substrate. The aluminum honeycomb structure has the advantages of light weight, high strength, high rigidity and the like, is widely applied to spacecrafts such as satellites and the like, and is provided with aluminum hexagonal honeycomb at the center and a middle interlayer containing a large amount of air. In order to adapt to the space vacuum environment, a plurality of air vents are arranged on the aluminum honeycomb structure of the substrate during production, so that the internal and external air pressures of the honeycomb structure are gradually balanced after the satellite enters the space. Due to the vacuum environment in space, the difference in internal and external gas pressure causes the gas to be exhausted from the aluminum honeycomb panel, generating a counter acting force acting on the solar sailboard. The satellite just enters a flight orbit and after the solar sailboard is unfolded, the acting force generated by the deflation of the solar sailboard is large, the influence on the attitude of the satellite is large, and particularly the influence on the change of the attitude of the satellite is large due to the deflation torque of the solar sailboard in the asymmetric configuration of the single-wing satellite.
The gas discharged from the solar sailboard honeycomb panel is mainly air entering during ground storage, and the air is continuously discharged after entering the space, so that adverse effects are generated on optical instruments such as a satellite payload, an earth sensor, a sun sensor and a star sensor, and the functions and the performance of the solar sailboard honeycomb panel are affected during starting. Therefore, it is of great practical significance to be able to give the end time of the deflation of the solar panels.
When the satellite operates in orbit, the deflation duration of the solar panel is difficult to directly measure, and needs to be indirectly estimated by measuring other physical parameters. When the elliptic orbit and static orbit satellite is launched, the elliptic orbit and static orbit satellite can reach a preset orbit (after 3-7 days) only through multiple orbital transfer flights of a transfer orbit section, the satellite is oriented to the sun for a long time, and the solar cell slice always faces the sun direction and keeps static and does not rotate. In the process of satellite in-orbit flight, the attitude is influenced by sunlight pressure moment, gravity gradient moment and aerodynamic moment, the sun orientation process of the satellite is influenced by the sunlight pressure moment with the fixed direction magnitude of 0.2 mNm-0.4 mNm, one orbit period of the transfer orbit section influences the attitude by about 0.2 degrees, and the attitude angular speed and the attitude angular change of the satellite are small. The characteristics of the orbit of the elliptic orbit satellite and the orbit of the geostationary orbit satellite determine that the directions of the aerodynamic moment and the gravity gradient moment are changed continuously, and the duration is short. Therefore, the windage moment of the windsurfing board is estimated, and the influence of the sunlight pressure moment, the gravity gradient moment and the aerodynamic moment can be ignored.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a method for estimating the deflation time of a sailboard in a transfer orbit section of a single-wing satellite, which can calculate the interference torque generated by deflation of the solar sailboard in real time by utilizing the on-orbit attitude angular velocity of the satellite and the working time length data of a thruster based on the basic principle of satellite attitude dynamics, and further identify the deflation duration time of the solar sailboard.
According to one aspect of the invention, a deflation time estimation method for sailboards in a transfer orbit segment of a monowing satellite is provided, and is characterized by comprising the following steps:
acquiring timing data of angular velocity and a thruster;
selecting iteration time;
step three, summing moment integrals;
step four, resolving the disturbance torque;
and step five, estimating the air bleeding duration of the sailboard.
Preferably, the third step comprises the following steps:
thirty, integrating the change of the angular momentum;
thirty-one, calculating a control moment integral of the attitude control thruster;
step thirty-two, calculating the integral of the air release moment;
and step thirty-three, calculating the coupling moment action integral generated by the inertia product term in the satellite rotational inertia.
Preferably, the step four is performed to judge whether the air bleeding of the sailboard is finished according to the real-time calculated air bleeding disturbance moment value, judge whether the air bleeding disturbance moment Md is greater than 0, and if the air bleeding disturbance moment Md is greater than 0, it is determined that the air bleeding of the sailboard is not finished, and continue to perform the disturbance moment estimation of the step four; if the bleed disturbance moment Md is equal to 0, the end of the bleed of the sailboard is indicated, and the bleed duration can be calculated.
Preferably, the fourth step takes the influence of the coupling term into account in the disturbance torque solution process.
Preferably, the method for estimating the air bleeding time of the sailboard at the transfer orbit section of the monowing satellite calculates the interference torque generated by air bleeding of the sailboard in real time by utilizing the angular velocity of the in-orbit attitude of the satellite and the working time length data of the thruster according to the acting force and the reacting force in the third law of Newton mechanics and the dynamic principle of the attitude of the satellite, and further identifies the air bleeding duration time of the solar sailboard.
Compared with the prior art, the invention has the following beneficial effects: the method can estimate the interference moment value generated by the deflation of the solar sailboard in real time by utilizing the data of the real-time angular velocity of the satellite, the working timing of the thruster and the like, and further estimate the deflation ending time of the solar sailboard. The method provided by the invention is simple, can be used in high and low orbit satellites, can provide a real and accurate deflation torque value particularly for the deflation process of the single wing sailboard of the satellite, and the analysis result can be used in the fields of spacecraft space environment torque estimation, control system design, on-orbit use of optical instruments and the like.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a flow chart of a method for estimating the deflation time of a sailboard in a transfer orbit segment of a monowing satellite according to the invention.
FIG. 2 is a graph of an estimate of the moment of deflation disturbance after deployment for a certain single wing sailboard satellite.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that variations and modifications can be made by persons skilled in the art without departing from the spirit of the invention. All falling within the scope of the present invention.
As shown in FIG. 1, the method for estimating the deflation time of the sailboard of the transfer orbit segment of the monowing satellite comprises the following steps:
the method comprises the steps of firstly, collecting timing data of angular velocity and a thruster, reflecting the attitude change of a satellite by the angular velocity on the satellite, indirectly reflecting the action results of interference torque and control torque on each axis of the satellite, and generally realizing real-time data acquisitionThe sexual requirement is high, the acquisition frequency is high, and is generally 0.5s (second); the timing of the thruster and the air injection quantity can be slowly transmitted, and are generally 8s or 16 s. Suppose that the measurement data update period of the angular velocity on the satellite is T1(ii) a The timing updating period of the thruster is T2(T2Is T1Integer multiple of) such that the satellite attitude angular velocity data is at T2M times are collected within the time, and the calculation formula of m is shown as the following formula (1);
m＝T2/T1……(1)
step two, selecting iteration time and iteration step length TsThe time interval of two integral summations is represented, the data output frequency in the calculation of the disturbance torque is determined, and is not less than the updating period T of the angular velocity measurement data1Taken as nT1N is more than or equal to 1 and less than or equal to m, n is a constant integer, and the timing updating period of the thruster is T2Integration duration T to ensure data validityrNot less than the timing update period T of the thruster2；
Step three, summing moment integrals;
step four, resolving the disturbance moment, integrating the terms on the two sides of the equation of the satellite attitude dynamic equation respectively, decomposing to obtain the following formula (10),
using each integral result obtained by the above formula (10) to reversely solve to obtain a triaxial interference moment generated by the air discharge of the solar panel; determining and obtaining iteration step length T according to data requirementssEvery T, everysOne point is taken as t of the integral0At any moment, repeating the third step and the fourth step to obtain continuous disturbance torque output;
estimating the deflation duration of the sailboard, judging whether the deflation of the sailboard is finished according to the deflation disturbance moment value obtained by real-time calculation in the fourth step, judging whether the deflation disturbance moment Md is greater than 0, if the deflation disturbance moment Md is greater than 0, indicating that the deflation of the sailboard is not finished, and continuing to estimate the disturbance moment in the fourth step; if the bleed disturbance moment Md is equal to 0, the end of the bleed of the sailboard is indicated, and the bleed duration can be calculated.
The third step comprises the following steps:
thirty, integration of the change of angular momentum, i.e. in the equation of attitude dynamics
Integrating; integral end moment angular momentum
Subtracting integral initial moment angular momentum
That is, the amount of change in angular momentum, the formula for calculating the amount of change in angular momentum is shown in the following equation (2),
wherein J is a whole-satellite inertia matrix, the ground can obtain an accurate value through testing, omega is the angular velocity of the satellite inertial angular velocity in the system, t0Which represents the initial moment of the integration,
respectively, the differential of the three-axis angular velocity, i.e. the three-axis angular acceleration,
is t0+TrThe angular velocity of the x-axis at the moment,
is t0Angular velocity of axis x, omega at time xx(i+m)To adoptThe i + m x-axis angular velocities, ω, in the set of dataxiSetting the first value or 0 of the data for the ith x-axis angular velocity in the collected data and the integral initial value of the angular velocity, wherein the other two directions are similar;
thirty-one, calculating the integral of the control moment Mc of the attitude control thruster, only having the control moment when the thruster works, and viewing the control moment Mc as a constant value step pulse, obtaining an accurate thruster value through calibration of the in-orbit thruster, needing to respectively obtain the working time length of the thruster generating a certain axis of positive and negative moment from satellite telemetering data, superposing the products of the control moment generated by each thrust and the respective working time length to realize the working control moment integral of the thruster, wherein the calculation formula is shown as the following formula (3),
wherein M iscxX-axis control moment, M, generated for thrusterscxPFor positive moment, M, produced by the thruster in the x-axiscxNNegative moment, T, generated by the thruster on the x-axisxP(i+m)、TxPiRespectively (i + m) th and ith jet timing data, T, of the thruster generating the positive moment of the x axisxN(i+m)、TxNiRespectively generating the (i + m) th and the ith jet timing data of the thruster generating negative moment of the x axis, wherein the other two axes are similar;
step thirty-two, the integral of the deflation torque Md is obtained, the deflation disturbance torque exists all the time in the whole process, but the generated deflation reaction torque can also gradually become smaller along with the gradual reduction of the continuous discharge pressure of the gas, the deflation disturbance torque integral obtains the angular momentum change generated by deflation by the product of the disturbance torque and the integral time length, and the calculation formula is shown as the following formula (4)
Wherein, the air-bleeding disturbing moment of the solar sailboard in the upper formulas Mdx, Mdy and Mdz is continuously reduced along with the time;
step thirty-three, the coupling moment action integral generated by the inertia integral term in the satellite rotational inertia is calculated, and for the X direction, the coupling term is- (J)yy-Jzz)ωyωzThe integral is shown in the following formula (5), and the coupling term is- (J) for the Y directionzz-Jxx)ωzωxThe integral is shown in the following formula (6), and the coupling term is- (J) for the Z directionxx-Jyy)ωxωyThe integral is shown in the following formula (7).
And step four, the influence of the coupling term is considered in the interference torque resolving process, and the accuracy is improved.
The invention is based on the basic principle of acting force and reaction of the third law of Newton mechanics and the dynamic equation of satellite attitude
(wherein,j is a satellite rotational inertia matrix, omega is a satellite body angular velocity, Mc is a satellite control moment, Md is an interference moment received by the satellite, which is specially referred to as a solar wing deflation interference moment), the satellite body angular velocity is changed according to the reaction moment when the solar sailboard is deflated, the solar sailboard deflation action moment is estimated by measuring the satellite attitude angular velocity and the working condition of a thruster, the deflation is gradually weakened along with the passage of time, the deflation action moment is gradually reduced, and the deflation duration of the solar sailboard can be estimated; in the method, in order to estimate the on-orbit air release time of the satellite solar sailboard, the air release moment of the solar sailboard needs to be estimated in real time, and the advantage of solving the interference moment by adopting an integral method is that the angular velocity data is filtered so as to reduce the noise error caused by the angular velocity difference; the method solves the disturbance moment reversely through the integral of each item in the satellite attitude kinetic equation, rather than directly solving the disturbance moment reversely by using the original equation with angular velocity difference; performing numerical integration on each item in the dynamics according to respective characteristics; the working control moment of the thruster only has the working time of the thruster, so that only a period of time has effect in the integral duration; the interference moment exists all the time, and the interference moment acts in the whole integral duration; the integral of the coupling term is obtained by superposing the data at each moment by the step size.
After a certain single wing sailboard is unfolded, the method adopts an example of estimating the moment of disturbance of the deflation of the solar wing; after a satellite single wing sailboard is unfolded, the inertia matrix of the satellite is as the following formula (20) (unit: kg. m2),
calculating to obtain the x-axis control moment M generated by the attitude control thrustercxP＝McxN18.9Nm, the generated y-axis control moment McyP＝McyN19.75Nm, resulting in a z-axis control moment MczP＝MczN15.4 Nm; angular velocity update period T1At 0.5s, the working time of the thruster is updated by a period T2Is 16 s; to ensure moderate data volume and effective data, the iteration step length TsAnd integration duration TrAll for 16 s. According to a single-wing satellite sailboard configurationIt can be known that the sailboard deflation disturbance moment acts in the + X axis direction, so the body + X axis disturbance moment of the satellite solar sailboard deflation process estimated according to the above method is shown in fig. 2. As can be seen from the disturbance torque estimation curve, the maximum disturbance torque generated by the windage of the windage panel reaches 0.037Nm, and the windage time of the windage panel lasts for nearly 16 hours.
In conclusion, according to the acting force and the reacting force in the third law of Newton mechanics and the dynamic principle of the satellite attitude, the interference moment generated by the air release of the solar sailboard is calculated in real time by using the angular velocity of the satellite in-orbit attitude and the working time length data of the thruster, and the air release duration of the solar sailboard is further identified. According to the relation between the acting force and the reacting force in the third law of Newton mechanics, the thrust generated by gas emission is used on the solar sailboard in a reaction mode, and further attitude angular velocity and attitude angular change around the center of mass of the satellite are generated.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes and modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention.
Claims (4)
1. A deflation time estimation method for sailboards in a transfer orbit section of a monowing satellite is characterized by comprising the following steps:
acquiring real-time angular velocity measurement data of a satellite and working timing data of a thruster;
selecting iteration time;
step three, summing moment integrals;
step four, resolving the interference torque to obtain an air bleeding interference torque value;
estimating the air bleeding duration of the sailboard;
the third step comprises the following steps:
thirty, integration of the change of angular momentum, i.e. in the equation of attitude dynamics
Integrating; integral end moment angular momentum
Subtracting integral initial moment angular momentum
That is, the amount of change in angular momentum, the formula for calculating the amount of change in angular momentum is shown in the following equation (2),
wherein J is a whole-satellite inertia matrix, the ground obtains an accurate value through testing, omega is the angular velocity of the satellite inertial angular velocity in the system, t0Which represents the initial moment of the integration,
respectively, the differential of the three-axis angular velocity, i.e. the three-axis angular acceleration,
is t0+TrThe angular velocity of the x-axis at the moment,
is t0Angular velocity of axis x, omega at time xx(i+m)For the i + m x-axis angular velocity, omega, in the collected dataxiSetting a first value or 0 of data for the ith x-axis angular velocity and the angular velocity integral initial value in the collected data, wherein the other two directions are similar;
thirty-one, calculating the integral of the control moment of the attitude control thruster, wherein the control moment only has the function of the control moment when the thruster works, the control moment Mc is regarded as a constant value step pulse, an accurate thruster value is obtained through calibration of an on-orbit thruster, the working time of the thruster generating positive and negative moments of a certain axis needs to be obtained from satellite telemetering data respectively, and the products of the control moment generated by each thrust and the respective working time are superposed to realize the integral of the working control moment of the thruster;
the calculation formula is shown in the following formula (3),
wherein M iscxX-axis control moment, M, generated for thrusterscxPFor positive moment, M, produced by the thruster in the x-axiscxNNegative moment, T, generated by the thruster on the x-axisxP(i+m)、TxPiRespectively (i + m) th and ith jet timing data, T, of the thruster generating the positive moment of the x axisxN(i+m)、TxNiRespectively generating the (i + m) th and the ith jet timing data of the thruster generating negative moment of the x axis, wherein the other two axes are similar;
step thirty-two, the integral of the deflation torque is obtained, the deflation disturbance torque exists all the time in the whole process, but the generated deflation reaction torque can also gradually become smaller along with the gradual reduction of the continuous discharge pressure of the gas, the deflation disturbance torque integral obtains the angular momentum change generated by deflation by the product of the disturbance torque and the integral time length, and the calculation formula is shown as the following formula (4)
Wherein M isdx，Mdy，MdzThe air-bleed interference torque of the solar sailboard in the above formula is continuously reduced along with the time;
step thirty-three, the coupling moment action integral generated by the inertia integral term in the satellite rotational inertia is calculated, and for the X direction, the coupling term is- (J)yy-Jzz)ωyωzFor the Y direction, the coupling term is- (J)zz-Jxx)ωzωxFor the Z direction, the coupling term is- (J)xx-Jyy)ωxωy；
The fourth step comprises: the deflation disturbance moment is resolved by integrating the terms on both sides of the equation of the satellite attitude dynamic equation and decomposing to obtain the following formula (5),
j is a satellite rotational inertia matrix, omega is a satellite star angular velocity, Mc is a satellite control moment, and Md is an interference moment on the satellite, and is specially referred to as a solar wing deflation interference moment.
2. The method for estimating the air bleeding time of the sailboard at the transfer orbit segment of the monowing satellite according to claim 1, wherein the method comprises the steps of judging whether the air bleeding of the sailboard is finished according to the air bleeding disturbance torque value obtained by real-time calculation in the fourth step, judging whether the air bleeding disturbance torque Md is greater than 0, if the air bleeding disturbance torque Md is greater than 0, indicating that the air bleeding of the sailboard is not finished, repeatedly executing the fourth step to estimate the disturbance torque until the air bleeding disturbance torque Md is equal to 0, indicating that the air bleeding of the sailboard is finished, and further calculating the air bleeding duration time.
3. The method for estimating time to bleed a sailboard for a transfer orbit segment of a monowing satellite according to claim 1, characterized in that said step four takes into account the influence of a coupling term in the process of disturbance torque solution.
4. The method for estimating the air bleeding time of the sailboard at the transfer orbit segment of the monowing satellite as claimed in claim 1, wherein the method for estimating the air bleeding time of the sailboard at the transfer orbit segment of the monowing satellite is characterized by calculating the disturbance moment generated by air bleeding of the sailboard in real time by utilizing the angular velocity of the in-orbit attitude of the satellite and the working time length data of a thruster according to the acting force and the reacting force in the third law of Newton mechanics and the dynamic principle of the attitude of the satellite, and further identifying the air bleeding duration time of the solar sailboard.
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| CN103303495A (en) * | 2013-04-11 | 2013-09-18 | 北京控制工程研究所 | Method for estimating disturbance moment in power decreasing process |
| CN104590588A (en) * | 2014-12-04 | 2015-05-06 | 哈尔滨工业大学 | Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy |
| CN105819004A (en) * | 2016-04-21 | 2016-08-03 | 上海微小卫星工程中心 | Solar array control method and system of satellite and satellite |
| CN105905317A (en) * | 2016-06-07 | 2016-08-31 | 湖北航天技术研究院总体设计所 | Sun-pointing control system for satellite and control method of sun-pointing control system |
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| CN103303495A (en) * | 2013-04-11 | 2013-09-18 | 北京控制工程研究所 | Method for estimating disturbance moment in power decreasing process |
| CN104590588A (en) * | 2014-12-04 | 2015-05-06 | 哈尔滨工业大学 | Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy |
| CN105819004A (en) * | 2016-04-21 | 2016-08-03 | 上海微小卫星工程中心 | Solar array control method and system of satellite and satellite |
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