RU2569999C2 - Navigation satellite orientation method - Google Patents

Abstract

FIELD: radio engineering, communication.

SUBSTANCE: invention relates to controlling orientation of navigation satellites with antennae and solar panels. The method includes orientation of the electrical axis of the antenna (first axis of the satellite) towards the Earth and orientation of solar panels towards the Sun. The latter is achieved by turning the satellite along with the solar panels around said first axis and turning the solar panels around a second axis, perpendicular to the first. Predicted satellite motion is facilitated when passing through special parts of an orbit, which include eclipse portions and portions with large Sun-satellite-Earth angles (greater than 175°). This is achieved through preemptive programmed turns around the first axis of the satellite, which are symmetrical relative to points of the orbit responsible for maximum and minimum Sun-satellite-Earth angles.

EFFECT: reduced error when predicting the movement of the centre of mass of a satellite and the error in knowing the position of the phase centre of an antenna.

4 cl, 5 dwg, 1 tbl

Description

The invention relates to the field of space technology and can be used in navigation spacecraft to reduce the error in predicting the motion of the center of mass of the navigation satellite and the error in knowing the position of the phase center of the antenna when passing specific points of the orbit.

A known method of orienting a navigation satellite, including orienting the electric axis of the antenna parallel to the first axis of the satellite, to the Earth and the orientation of solar panels on the Sun by turning the satellite together with the solar panels (BS) relative to the first axis of the satellite to align the normal to the solar panels with the plane Sun-satellite-Earth ”and the rotation of solar panels around the second axis perpendicular to the first, until the normal to the solar panels coincides with the direction to the Sun [Fundamentals of oektirovaniya spacecraft information provision: Proc. allowance / Chebotarev V.E., Kosenko V.E; Sib. state aerospace un-t - Krasnoyarsk, 2011, p. 275-282].

When the satellite is oriented, the normals to the solar panels should be oriented in the plane Sun-satellite (Object) - Earth (POPs). The rotation speed of the plane of POPs (w _cos ) is determined by the formula (1) using FIG. one.

In FIG. 1 marked: POPs - the angle of the Sun-satellite-Earth; α is the angle between the plane of the orbit and the direction to the Sun; E is the angle from the current position of the satellite in orbit to the point of the orbit at which the POP angle is minimal (maximum); E ₀ is the angle from the point in orbit at which the anticipatory turn is turned on to the point at which the POP angle is maximum (minimum); Ψ - yaw angle (current angle between the axis OY ₀ and the projection of the direction on the Sun onto the plane OZ ₀ Y ₀ ); ψ _DEVELOPMENT - the angle of rotation around the axis OX in the process of moving in orbit from the current position of the satellite to the point of passage of the maximum (minimum) angle of POPs; OX _O Y _O Z _O - the right orbital coordinate system (OX _O - is directed along the current radius-vector of the product from the Earth, OZ _O - is directed along the normal to the plane of the orbit); OX _Z Y _Z Z _Z - the right solar-terrestrial coordinate system (OX _Z - coincides with OX _O , OY _Z - lies in the plane X _O OX _S ); OXYZ is the coordinate system associated with the satellite (OX is the first axis of the satellite, OY is the third axis of the satellite, OZ is the second axis of the satellite).

where w ₀ is the orbital angular velocity.

When shadow areas pass from the Earth, the POP angle takes a minimum value. When passing the sections of the orbit in which the satellite is between the Earth and the Sun, the angle of POPs takes a maximum value. From the spherical triangle CSS ₁ in FIG. 1 we have:

и

, получим:differentiating (2) and assuming that

and

we get:

From the spherical triangle CSS ₁ in FIG. 1 we have:

Substituting (5) in (3) we obtain the formula (1).

From formula (1) it follows that at an angle α between the orbital plane and the direction to the Sun close to zero, the rotation speed of the POP plane is close to zero. However, at the point E = 0, the rotation speed of the POP plane tends to infinity.

Thus, tracking the plane of POPs when passing small (close to 0 °) and large (close to 180 °) angles of POPs using the executive organs of the orientation system is impossible without error. In FIG. Figures 2 and 3 show the processes of tracking by the XOY plane the POP plane without lead and with lead at an angle α = 0.

From consideration of FIG. 2, 3 it can be seen that, without preempting, the orientation error of the panels on the Sun is much larger than with preempting.

The following should also be noted. Due to the fact that the satellite is navigation, the error in the orientation of the BS panels on the Sun leads to unpredictable forces from solar pressure, which affect the movement of the center of mass of the spacecraft (SC).

An error in the orientation of the BS panels on the Sun during satellite rotation occurs in the XOZ plane, since the solar battery drive cannot eliminate the error in this plane (a single-stage solar battery drive eliminates the orientation error only in the XOY plane).

The projections of the solar pressure forces on the OY ₀ axis of the orbital coordinate system, which coincides with the satellite linear velocity vector, are compensated for with a symmetrical anticipatory turn relative to the point of the maximum POP angle. The projections of forces from solar pressure on the OZ axis _{0 of the} orbital coordinate system, which coincides with the normal to the orbit plane, are added.

The influence of forces from solar pressure directed along the satellite’s linear velocity vector on the motion of the satellite’s center of mass is an order of magnitude greater than from the forces directed from the OZ ₀ axis (normal to the orbit). Therefore, with a symmetric anticipatory U-turn, the influence of the anticipatory U-turn on the motion of the center of mass of the spacecraft is minimized.

Thus, we can state the following:

1. The movement during the passage of the special points of the orbit using the anticipatory U-turn is predictable, since it is a programmatic U-turn.

2. The anticipatory turn should be symmetrical with respect to the point of the maximum (minimum) angle of the POPs, while the influence of the anticipatory turn on the motion of the center of mass of the spacecraft is minimized.

Closest to the claimed technical solution in terms of technical nature and the technical result achieved is a method for orienting a navigation satellite, including orienting the antenna’s electric axis parallel to the first axis of the satellite to Earth and orienting solar panels on the Sun by turning the satellite along with solar panels relative to the first axis of the satellite until the normal to the solar panels coincides with the Sun-satellite-Earth plane and the solar panels turn around the second s of rotation perpendicular to the first, until the normal to the solar panels coincides with the direction to the Sun, in predetermined orbit intervals, covering the intervals of the satellite orientation uncertainty in shadow orbits, where independent predictive turns around the first and second axis of the satellite are carried out by an estimated value with intermediate retention of a given orientation [Application No. 2012152127 of 04/04/2112, the Method of orientation of the navigation satellite]. The described method is adopted as a prototype of the invention.

The disadvantages of the prototype:

We determine the magnitude of the force from solar pressure acting on the plane of the aircraft panel using FIG. 4, on which is indicated:

F _P - pressure force from the incident stream;

F _OTP - pressure force of the mirror-reflected flow;

F _D - pressure force from a diffusely reflected stream;

θ is the angle between the normal to the panel and the direction of the light flux.

The force of solar pressure (F _SD ) is determined by:

The value of the force F, equal to the sum of the forces from the incident stream and the specularly reflected stream, is equal to:

where S is the area of solar panels; h ₀ - specific value of solar pressure (h ₀ = 4,63 · 10 ^-6 N / m ² ); C _S - specular reflection coefficient (the ratio of the energy contained in a unit volume of the mirror-reflected flow to the energy value in a unit volume of the incident stream);

The angle θ ₁ between the normal to the panel and the force F ₁ is determined by the formula:

The magnitude of the force F _D from the diffusely reflected flow is determined by:

where C _D is the diffuse reflection coefficient (the ratio of the energy contained in a unit volume of a diffuse reflected stream to the energy value in a unit volume of the incident stream).

With the standard orientation, the angle θ is almost zero. When θ = 0, we have:

Forming a preemptive turn of the BS panel along the second axis by the angle θ, we obtain that when passing through the same sections of the orbit, we will have different values of the force acting on the spacecraft and the angle of the force relative to the standard orientation. Thus, the motion of the center of mass of the spacecraft will be carried out under the influence of other forces from solar pressure in comparison with the standard mode, which leads to an error in predicting the motion of the center of mass of the spacecraft. Instant adaptation of the prediction of the motion of the center of mass of the spacecraft under the influence of altered forces from solar pressure will not occur.

Moreover, to predict the motion of the center of mass, accurate knowledge of the coefficients of specular C _S and diffuse reflection C _{D is} necessary. The coefficients C _S and C _D depend on the time spent by the satellite in space and cannot be measured. Even if algorithms for operative forecasting the motion of the center of mass are created, knowledge of the exact current values of the coefficients C _S and C _{D is} necessary.

Thus, the apparent simplicity of reducing the error in predicting the motion of the satellite's center of mass when passing the Sun-satellite-Earth angles close to 180 ° (large POP angles) by forming a proactive rotation of the solar panel around the second axis leads to a shift in the value of the nominal force from the solar pressure on an unknown value, due to which a large error in predicting the motion of the satellite's center of mass will be obtained. Thus, it is advisable to go through large angles of POPs by anticipating a programmatic turn around the first axis in a portion of the orbit that is symmetrical about the point of the maximum angle of the POPs. In this case, the error in predicting the motion of the center of mass of the navigation satellite will be minimal.

The basis of the present invention is the creation of a method for orienting a spacecraft during the passage of specific points of the orbit, which allows to reduce the error in predicting the motion of the center of mass of the navigation satellite and the error in knowing the position of the antenna phase center.

The problem is solved as follows.

A method for orienting a navigation satellite is claimed, including orienting the antenna’s electrical axis parallel to the satellite’s first axis to Earth and orienting solar panels on the Sun by turning the satellite along with the solar panels relative to the first satellite axis until the normal to the solar panels is aligned with the Sun-satellite plane -Earth "and the rotation of solar panels around the second axis of the satellite, perpendicular to the first, until the normal to the solar panels coincides with the direction to the Sun, distinct In that, during the passage of special sections of the orbit, including the shadow sections of the orbit and the sections of large angles of the Sun-satellite-Earth (greater than 175 °), the predicted satellite motion is organized by anticipating programmed turns around the first axis of the satellite in the orbit sections symmetrical with respect to the points of maximum and the minimum angle of the Sun-satellite-Earth.

SUMMARY OF THE INVENTION

The sequence of operations when passing large (small) angles of POPs.

When passing large (small) angles of POPs when the modulus of the angle between the orbital plane and the direction to the Sun is less than the set value (2 °) and the angle between the OX axis and the direction to the Sun is less than the set value (5 °) on each control cycle, the following actions are performed:

1) the angle E is determined in the orbit plane between the current position of the spacecraft and the position of the spacecraft at the time of passage of the maximum (minimum) angle of POPs;

2) the rotation angle ψ _DEVELOPMENT relative to the OX axis is determined in the process of orbiting from the current spacecraft position to the point of passage of the maximum (minimum) POP angle, in which the OY axis should be perpendicular to the orbit plane;

3) the time T _{1 of the} turn along the axis OX to the above angle is determined at a given speed of turn around the axis OX;

4) the time T ₂ for orbiting the angle from the current position to the point of passage of the maximum (minimum) angle of the POPs is determined.

As soon as the time T ₁ required for a turn around the OX axis to the calculated turn angle becomes longer than the T ₂ orbit time from the current position of the spacecraft to the point of passage of the maximum (minimum) angle of the POPs, a sign is formed for proactive programmed turn of the spacecraft around the OX axis .

The pivot speed and pivot angle in a forward pivot are controlled by information from the speed measurement unit (LSI). Turning off the anticipatory program turn is carried out after 2T ₁ , and the orientation is carried out according to the algorithm of the orientation to the Earth (ROS).

With this sequence of operations, a pre-emptive program turn is formed on the orbit portion symmetrical with respect to the point of the orbit at which the Sun-satellite-Earth angle is maximum (minimum). If a pre-emptive program turn is formed under the condition T ₂ > T ₁ , then a pre-emptive program turn will be formed on a section of the orbit that is asymmetric with respect to the point of the orbit at which the Sun-satellite-Earth angle is maximum (minimum). Consequently, the error in predicting the motion of the center of mass will increase.

In order to reduce the error of the actual orientation relative to the reference along the yaw channel during the anticipatory software turn in the spacecraft software (ON), the following operations are provided:

1) the angle of rotation and the angular velocity of the rotation are calculated using a reference model in which the control moments and moments of inertia of the spacecraft relative to the axis OX are nominal;

2) at the same time, the actual angle and angular velocity are calculated according to the information from the LSI during control using the nominal value of the control code (moment) supplied to the flywheel engine;

3) there are differences between the actual and reference value of the angle and between the actual and reference value of the angular velocity, and an amendment to the control moment relative to the axis OX is generated from these differences;

4) the reference angular velocity and the reference turning angle are monitored using the correction to the nominal value of the control torque.

If the phase center of the navigation antenna is not aligned with the center of mass of the spacecraft, then when conducting forward-looking turns both when passing large and small angles of POPs, an error will arise in determining the position of the center of mass of the spacecraft. This error is due to the fact that consumer equipment measures the position of the phase center of the navigation antenna, and the movement is determined by the movement of the center of mass.

Exchange on the navigation channel with consumer equipment in the current time mode is not possible due to the low bandwidth of the navigation channel. Therefore, to calculate the corrections of the position of the center of mass, the source data should be sent to the consumer equipment to calculate the forward turn, namely:

- the time of the start of the anticipatory reversal;

- the end time of the anticipatory reversal;

- the initial velocity of the satellite relative to the first axis;

- angle of anticipatory reversal;

- angular acceleration of the satellite during a turn;

- the angle between the position of the SPACECRAFT in orbit at the moment of turning on the anticipatory turn and the point of the maximum (minimum) angle of the POPs.

Calculation of the angle during a forward turn in the consumer equipment is made according to the model of a symmetric turn relative to the point of the maximum (minimum) angle of the POPs.

The models of the satellite’s forward turn on board and in the consumer’s equipment are the same, therefore, the forward turn angle calculated in the consumer’s equipment is equal to the turn angle calculated on board and worked out by the orientation system.

The calculation of ballistics is carried out in the current time on each control cycle of the on-board digital computer (BCM). To ensure the issuance of anticipatory reversal parameters 10-15 minutes before the anticipation of the anticipatory reversal, parallel calculation of ballistic parameters and conditions for the inclusion of the anticipatory reversal is necessary for the orientation and stabilization system (OS) software 10-15 minutes ahead, which leads to additional loading of the digital computer due to the increase , almost doubled, the work of the SOS software and the software complex of ballistic tasks.

The kinematic parameters of the anticipatory reversal depend only on the angle between the orbital plane and the direction to the Sun. At the same angle between the orbital plane and the direction to the Sun, realized at different times, the parameters of the anticipatory reversal will be the same, except for the time of the beginning and end of the anticipatory reversal.

Therefore, if we calculate the tables of the dependences of the forward turn on the angle between the plane of the orbit and the direction to the Sun and put these tables on board the satellite, then, knowing the angle α at the time of turn on the forward turn, it is possible to obtain the forward turn parameters described almost without time on the computer in table 1.

Thus, to obtain data for the calculation of anticipatory reversal, you must know:

- the angle between the plane of the orbit and the direction to the Sun at the moment of turning on the anticipatory turn;

- the time of the beginning and end of the anticipatory reversal.

Below is a list of operations to solve the problem.

To predict the magnitude of the angle between the direction to the Sun and the plane of the orbit, it is necessary to know the rate of change of the angle between the plane of the orbit and the direction to the Sun. To do this, for the time ΔT until the specified moment of calculating the parameters of the anticipatory reversal, the current angle α between the orbit plane and the direction to the Sun is remembered. At a given point in time, the rate of change of the angle α is determined:

where α _Z is the stored angle; α is the current angle.

To predict the angle α at the start of a forward turn, it is necessary to know the time T _{P of} the satellite in orbit from the current point of the orbit to the point of the orbit at which the lead turn is turned on.

where Е _T is the orbital angular distance from the current point to the point at which the Sun-satellite-Earth angle is maximum (minimum);

E ₀ is the angular distance in orbit from the point at which the anticipatory turn is turned on to the point at which the angle of the Sun-object-Earth is maximum (minimum);

w ₀ is the orbital angular velocity.

The angle E _{0 is} determined approximately using the tables of dependence E ₀ (α).

Angle E _T using FIG. 1 is determined by the formula:

where POP _T is the current angle of the Sun-satellite-Earth; α _T is the current angle α.

The angle α _H between the plane of the orbit and the direction to the Sun is determined by the formula:

The angle α _Н from the tabular dependencies determines the parameters of the anticipatory reversal:

- the initial velocity of the satellite relative to the first axis;

- angle of anticipatory reversal;

- the angle between the position of the satellite in orbit at the moment of turning on the anticipatory program turn and the point of the maximum (minimum) angle of the Sun-satellite-Earth.

The start and end time of a proactive software reversal is determined by the formula:

The parameters of the anticipatory program turn, including the start and end times of the anticipatory program turn, are recorded in the navigation frame.

When the consumer equipment is operating at time T _N , the cyclogram for calculating the heading angle begins to unfold.

The block diagram of the formation of control moments relative to the first axis of the satellite during proactive software turns is shown in FIG. 5.

In FIG. 5 is indicated:

ψ _A is the reference angle;

ψ _F is the actual angle;

ω _N is the initial angular velocity;

ψ _DEVELOPMENT - the angle of anticipatory reversal;

ψ _ACCELERATION is the angle traveled in the acceleration section;

M _X - control moment;

M ₀ - rated control torque;

To ₁ , To ₂ - the coefficients of the control channel of the engine, flywheel;

T _RISE - time of satellite acceleration relative to the first axis to the search speed;

T _N - the time of the beginning of the anticipatory reversal;

T _To - the time of the end of the anticipatory reversal;

T _P - turn time at the search speed;

T _TEK - current time;

- эталонное угловое ускорение.

- reference angular acceleration.

The formation of the heading angle in the consumer equipment during the passage of singular points of the orbit is similar to the formation of the angle ψ _A described in FIG. 5.

Thus, the claimed invention allows to reduce the error of predicting the motion of the center of mass and the error of knowing the position of the phase center of the antenna of the navigation satellite when passing special points of the orbit.

Table 1 METHOD OF ORIENTATION OF NAVIGATION SATELLITE Initial data for the implementation of proactive reversal No. α _N , ° ψ _dev , ° ω _start , ° / s E ₀ , ° one 0 180 0 3.842131 2 0.016667 179.500994 5.77514e-4 3.830233 3 0.033334 178.998902 1.162176e-3 3.818335 four 0.050001 178.49368 1.7541e-3 3.806349 5 0.066668 177.985234 2.353512e-3 3.794184 6 0.083335 177.473603 2.960335e-3 3.78202 7 0.100002 176.958704 3.57477e-3 3.769767 8 0.116669 176.440492 4.196935e-3 3.757425 9 0.133336 175.918822 4.82718e-3 3.744906 10 0.150003 175.393828 5.465199e-3 3.732386 eleven 0.16667 174.865377 6.111313e-3 3.719778 12 0.183337 174.333285 6.76597e-3 3.706992 13 0.200004 173.797613 7.429036e-3 3.694118 fourteen 0.216671 173.258471 8.100249e-3 3.681243 fifteen 0.233338 172.7155 8.780484e-3 3.668191 16 0.250005 172.168796 9.469514e-3 3.65505 17 0.266672 171.618308 0.010167 3.64182 eighteen 0.283339 171.063984 0.010874 3.628502 19 0.300006 170.505538 0.011591 3.615006 twenty 0.316673 169.943121 0.012317 3.601421 21 0.33334 169.376939 0.013052 3.587836 22 0.350007 168.806155 0.013798 3.573984 23 0.366674 168.231495 0.014554 3.560133 24 0.383341 167.652337 0.01532 3.546104 25 0.400008 167.0689 0.016096 3.531987 26 0.416675 166.481126 0.016882 3.51778 27 0.433342 165.888604 0.01768 3.503396 28 0.450009 165.291597 0.018488 3.488923 29th 0.466676 164.690044 0.019307 3.474362 thirty 0.483343 164.083483 0.020138 3.459623 31 0.50001 163.472223 0.020981 3.444795

Claims (4)

1. A method for orienting a navigation satellite, including orienting the electric axis of the antenna parallel to the first axis of the satellite to Earth and the orientation of solar panels on the Sun by turning the satellite along with the solar panels relative to the first axis of the satellite until the normal to the solar panels is aligned with the Sun-satellite plane -Earth and rotation of solar panels around the second axis of the satellite, perpendicular to the first, until the normal to the solar panels is aligned with the direction to the Sun, different the fact that during the passage of special sections of the orbit, including the shadow sections of the orbit and the sections of large angles of the Sun-satellite-Earth (greater than 175 °), pre-emptive programmatic turns are made around the first axis of the satellite in parts of the orbit that are symmetrical with respect to the points of the orbit at which the angle of the Sun satellite-Earth is maximum or minimum, thereby organizing the predicted satellite motion. 2. A method for orienting a navigation satellite according to claim 1, characterized in that, at Sun-satellite-Earth angles smaller (large) than a predetermined value, an angle E in the orbit plane from the current position of the satellite to the passage point of the minimum (maximum) angle is determined on each control cycle ) angle of the sun-satellite-Earth, determine the angle of rotation ψ _developed about the first axis of the satellite from the current position to the point of passage of the minimum (maximum), the angle of the sun-satellite-Earth, determine the time T ₁ turn around a first axis by an angle ψ _{to develop} for at Anna reversal rate, determine the time _T2 passage angle E by controlling the condition T _1> T _2, when the condition T _1> T ₂ include predictive software turn around a first axis calculated parameters preemptive software turn, turn off predictive software reversal after time 2T ₁ while ensuring the symmetry of the portion of the orbit on which the anticipatory programmatic turn is carried out, relative to the points of the orbit at which the angle of the Sun-satellite-Earth is maximum or minimum. 3. A method for orienting a navigation satellite according to claim 1, characterized in that the parameters of the anticipatory program turn around the first axis of the satellite are determined, including the time of the beginning and end of the anticipatory program turn, for a predetermined time before the start of the anticipatory program turn using the calculated dependencies of the anticipatory program turn program turn from the angle between the plane of the orbit and the direction to the Sun at the moment of turning on the anticipatory program turn and transmit to the frame of navigation information, include in the estimated time a proactive software turn, form the reference model of the proactive software turn around the first axis of the satellite, form the tracking mode of actual movement around the first axis of the satellite reference model of the turn. 4. The orientation method of the navigation satellite according to claim 3, characterized in that when determining the parameters of the anticipatory program turn, including the start and end time of the anticipatory program turn, for a given time before the start of the anticipatory program turn

changes in the angle between the orbit plane and the direction to the Sun, determine the angle E ₀ between the position of the satellite at the moment of turning on the anticipatory programmed turn and the point of the orbit at which the Sun-satellite-Earth angle is maximum (minimum) according to the current angle between the orbit plane and the direction to the Sun using tabular dependencies determine the angle E _T in the orbit between the current position of the satellite and the point of the orbit at which the angle of the Sun-satellite-Earth is maximum (minimum), determine the angle in orbit from the current point to point VK of anticipating programmatic reversal using angles E ₀ and E _T , determine the time T _P , during which the satellite will orbit the angle from the current point to the point of inclusion of anticipatory programmatic reversal, determine the angle between the plane of the orbit and the direction to the Sun at the moment of switching on anticipatory programmatic reversal , determine the parameters of the anticipatory program turn using tabular dependencies of the parameters on the angle between the plane of the orbit and the direction to the Sun, determine the start and end time of the lead waiting for a software reversal.

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