Navigation satellite orientation system

FIELD: physics.

SUBSTANCE: invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite.

EFFECT: high accuracy of determining navigation-time data on navigation artificial earth satellites by consumers.

4 cl, 12 dwg

The invention relates to the field of space technology, and more specifically to method the orientation of the navigation satellites.

During normal operation of the satellite in orbit is its spatial orientation to the Earth, the Sun in the orbital plane: at the same time on two or three of the above [1].

The spatial location associated with the center of mass of the satellite coordinate system X_cY_cZ_cdepending on the angular motion of the satellite along the orbit (angle γ) in the terrestrial orbital coordinate system X₀Y₀Z₀is determined from the solution of the spherical triangle (see Fig.1):

$$\cos \beta =-\cos \eta \cdot \cos \gamma ,\text{(1)}$$

$$\sin \alpha =\frac{\sin \eta }{\sin \beta }=\frac{\sin \eta }{\sqrt{1-{\cos }^2\eta \cdot {\cos }^2\gamma }},\text{(2)}$$

where β is the angle Sun - Earth - Satellite calculated by the formula: β=180-R_POPSwhere R_POPSangle of the Sun - Object (Satellite) - Ground (hereinafter POPS) in satellite the orbital coordinate system; γ is the angle position of the satellite in orbit; η is the angle of declination of the Sun above the plane of the orbit; α is the angle between the orbit plane and the plane of the POPS.

For one revolution of the satellite in orbit (0≤γ≤360°) of angle tracking is limited to the following ranges: η≤β≤180°-η, η≤α≤90°.

Thus, while the orientation of the antennas of the satellite to the Earth and solar panels (PRP) in the Sun, you must enter the kinematic relationship which trace the angles β and α by using single or two-stage actuators.

The angular velocity tracking is determined by differentiation of equations (1) and (2):

$$\stackrel{}{\beta }=K_{\beta }\cdot \stackrel{}{\gamma }$$,$$K_{\beta }=\frac{\cos \eta \cdot \sin \gamma }{\sqrt{1-{\cos }^2\eta \cdot {\cos }^2\gamma }},\text{(3)}$$

$$\stackrel{}{\alpha }=K_{\alpha }\cdot \stackrel{}{\gamma }\text{,}{K}_{\alpha }=\frac{\sin \eta \cdot \cos \eta \cdot \cos \gamma }{1-{\cos }^2\eta \cdot {\cos }^2\gamma },\text{(4)}$$

where K_β, K_α- the transformation ratio of the speed tracking;$$\stackrel{}{\gamma }$$- the angular velocity of the motion of the satellite, for a circular orbit with a constant value of$$\stackrel{}{\gamma }=\frac{{360}^o}{T}$$(here T is the orbital period of a satellite in orbit).

The known method the orientation of the satellite [1], providing a continuous three-axis orientation of the housing of the satellite together with the rigidly installed antennas and engines correction (DC) in the orbital coordinate system (on the Ground and in the plane of the orbit and the orientation of the solar panels in the Sun with a drive kinematically associated with the body of the satellite (see Fig.2). This scheme orientation was used on communication satellites, requiring continuous tracking targeted charts antennas on the selected area of the Earth's surface, maintaining the orbits of the pulse of the correction for the simultaneous operation of the satellite for the intended purpose with the organization of nepreryvnogo tracking PSB in the Sun turn with a drive around the binormal to the orbit with angular velocity $$\stackrel{}{\beta }$$(for geostationary satellites with a single drive) and additionally around the orthogonal direction to the binormal with angular velocity$$\stackrel{}{\alpha }$$(for satellites with any inclination of the orbit in the presence of a two-stage actuator).

The navigation satellites on an antenna with a wide beam pattern covering the whole Earth (global coverage), and in the process the operation does not permit the extradition of correction pulses. For these satellites is more than acceptable solar-terrestrial scheme orientation [1], which is closest to the claimed technical solution to the technical essence and the achieved technical result (see Fig.3).

The known method the orientation of the navigation satellite includes the orientation of the first axis of the satellite with the antenna on the Earth (radius-vector of the orbit and the orientation of the solar panels in the Sun turn the satellite with solar panels relative to the first axis of the satellite to align normal to the solar panels to the plane of the Sun - Satellite - Earth and spread of solar panels around the second axis of rotation perpendicular is th first to align normal to the solar panels pointed at the Sun. The described method is taken as a prototype of the invention.

During the operation of the navigation satellites in orbit during the year, the angle of declination of the Sun (angle η) varies in the range +90°, resulting in situations with uncertainty in the orientation of the satellite, due to the presence of shadow orbits (the shadow from the Earth crosses the orbit of the satellite in the area of small values of the angle POPS), as well as the emergence in the area of large values of the angle POPS high angular velocity tracking angle α exceeding the capabilities of Executive bodies.

The conditions for the occurrence of the shadow period of the orbits are determined by the inequality:

$$0\le \eta \le {\beta }_T,\sin {\beta }_T=\frac{R}{R+H},\text{(5)}$$

where β_Tthe angular size of the shadow area of the satellite's orbit from the Earth (ACE); R is the Earth's radius; H is the altitude of the circular orbit of the satellite.

The duration of the ACE for circular orbits used by the navigation satellites, is determined from equation (1):

$$t_T=\frac{2{\gamma }_T}{\stackrel{}{\gamma }},{\gamma }_T=\arccos \left(\frac{\cos {\beta }_T}{\cos \eta }\right),\text{(6)}$$

where γ_Tthe angular area of the shadow area in the plane of the orbit.

Thus, for a circular orbit navigation satellite height H≈20000 km angular velocity$$\stackrel{}{\gamma }$$≈0,5°/min, the angle β_T≈14,5° and the period of time of the existence of the shadow orbits can reach 25% for each six months. The maximum duration of an ACE is about 8% of the duration of the period of satellite revolution T [1].

At the time of passing the ACE tracking PSB on the angles β and α is not performed due to the lack of a reference point (the Sun), while the orientation of the first axis X_Csatellite and, accordingly, the electrical axis of the antenna on the Ground is supported according to the device orientation on the Ground. Therefore, after passing the ACE starts the restore staff orientation PSB in the Sun by rotation by the angle α around the first axis X_Csatellite search speed W_P1to align normal to the PSB with the plane of the POPS and spread PSB angle β around the second axis Z_Csatellite search near the STU W _PSBto align normal to the PSB with the direction to the Sun$$\stackrel{\to }{s}$$.

The duration of the recovery orientation depends on the magnitude of the error angle tracking α_Pat the time of exit from the ACE and includes two operations: alignment with the plane of the POPS (duration t₁and durasport PSB in the plane of the POPS (duration t₂) to align normal to the PSB with the direction to the Sun:$$\begin{array}{c}{t}_P=t_1+t_2=\frac{{\alpha }_P}{W_{P1}}\cdot \left(1+\frac{\stackrel{}{\gamma }}{W_{P\mathrm{With\; a}B}}\right)\text{,}{t}_1=\frac{{\alpha }_P}{W_{P1}},\\ t_2=\frac{\stackrel{}{\gamma }}{W_{P\mathrm{With\; a}B}},\end{array}\text{(7)}$$

where W_PSB- angular velocity of rotation of the PSB.

The initial value of the angle α_Pthat starts the restore orientation, depends on the reversal of the satellite in the shadow of the Earth around the first axis due to the presence of residual (random) angular velocity and is the predicted value being in the range of 0°...180°. Therefore, the duration of rotation about the first axis t_Pis a random value.

In addition, the shadow orbits depending on the position of the Sun relative to the plane of the orbit angular velocity tracking changes in a wide range and reach maximum values at the following points of the orbit:

$$\begin{array}{c}\gamma ={90}^o{\mathrm{,270}}^o,K_{\beta }^{\max }=\left|\cos \eta \right|,\gamma =0^o{\mathrm{,180}}^o,\\ K_{\alpha }^{\max }=\left|ctg\eta \right|.\end{array}\text{(8)}$$

So, for shadow orbit navigation satellites (H≈20000 km), in accordance with formulas (8) and (9), this range is equal to 0.97≤$$K_{\beta }^{\max }$$≤1,0, 3,9≤$$K_{\alpha }^{\max }$$<∞. Ie shadow orbits that satisfy the condition$$0\le \left|\eta \right|\le arctg\left(\frac{\stackrel{}{\gamma }}{W_{P1}}\right)<{\beta }_T$$in the areas of small and large angles POPS may encounter situations when the maximum speed tracking around the first axis of the satellite (the angle α) can exceed the capacity of the Executive bodies of tracking$$K_{\alpha }^0=\frac{W_{P1}}{\stackrel{}{\gamma }}$$(see Fig.4).

Interval plot of the satellite's orbit, where the angular velocity tracking exceeds the speed tracking of the Executive bodies, are determined from the solutions of the quadratic equation (4) with respect to cos γ:

$$\cos {\gamma }_{P\mathrm{With\; a}}=\frac{tg\eta }{2K_{\alpha }^0}-\frac{1}{\cos \eta },\text{(9)}$$

where γ_PS- the value of the angle γ to the showing $$K_{\alpha }\ge K_{\alpha }^0$$;$$K_{\alpha }^0$$- realized the value of the transformation coefficient,$$K_{\alpha }^0=\frac{W_{P1}}{\stackrel{}{\gamma }}$$.

In the area of the satellite's orbit with small angles POPS, where$$W_{P1}\le \stackrel{}{\alpha }$$is inconsistent program combining the second axis of the satellite with the actual plane of the POPS, which leads to the increase of error orientation PSB in the Sun during the following time:

$$t_{P\mathrm{With\; a}}=\frac{{\alpha }_P}{W_{P1}},\Delta {\gamma }_{\text{PS}}=\stackrel{}{\gamma }\cdot t_{P\mathrm{With\; a}}\text{(10)}$$

where α_P- the angle of rotation about the first axis in the recovery process ori is ncacii, which is a random variable, distributed in the range 0<α_P≤2(90°-η);

γ_PSthe area of the corners of uncertainty tracking.

In a period of uncertainty orientation PSB during the passage of the satellite shadow of the orbit, the angle between the normal to the PSB and the direction to the Sun is defined by the following relationship (see Fig.5):

$$\begin{array}{c}\text{cos}\phi =\text{-cos}\Delta \beta \cdot \text{cos}\beta +\text{sin}\Delta \beta \cdot \text{sin}\beta \cdot \text{cos}{\alpha }_P\text{,}\\ \cos \Delta \beta =\cos \eta \cdot \cos \Delta {\gamma }_H,\end{array}\text{(11)}$$

where Δβ is the angle the normal to the PSB from the axis X_Withat the time of entry of the satellite in ACE; Δγ_Harea of uncertainty orientation, γ_Tor Δγ_PS.

Due to the unpredictable occurrence of the provisions of the normal to the PSB relative to the direction of the Sun at intervals of uncertainty (near the minimum and maximum values of the angles POPS) increase the unpredictable components of the acceleration from the force of radiation pressure acting on the satellite, which leads to deterioration of the prediction accuracy of the motion parameters of navigatio the aqueous satellite shadow orbits and, as a consequence, increases the accuracy of observation of a user navigation satellite.

The calculation of the forces of radiation pressure on a single platform PSB is carried out by known formulas [1, 2], taking into account the components absorbed from (S_P), mirror (S_GSand diffuse-reflected (S_TOthe total solar flux (see Fig.6) and is presented in two ways:

a) in the form of two vectors that are deployed on the angle φ:

f_τ- component light pressure parallel to the incident light stream;

f_n- component light pressure parallel to the normal to the ground$${\stackrel{\to }{n}}_{P\mathrm{With\; a}B}$$;

b) in the form of two orthogonal vectors:

- parallel to the normal to the ground$$f_H=f_P+f_t\cdot \cos \phi \text{(12)}$$

in the lateral direction (in the plane of the platform)$$f_{\delta }=f_t\cdot \sin \phi \text{(13)}$$

In the intervals of uncertainty lateral component may take arbitrary the e position relative to the velocity vector of the satellite (the angle α _P), which introduces error in the calculation of these forces ≤2f_δ.

Calculations by formulas (12...13) forces of radiation pressure for GLONASS satellites in uncertainty interval t_n=15 min, t_PS=40 min showed that they differ from the forecast values up to 10% of the radius-vector and up to 30% of the velocity vector.

This leads to increased inaccuracy of predicting the position of a satellite in orbit on the daily interval up to 10% (confirmed by results of field tests of the satellite system GLONASS).

Furthermore, the presence of unpredictable reversals of the satellite around the first axis (angle α) in the area of small and large values of the POPS introduces additional error in the measurement range from the satellite to consumers (∆D) in the event of displacement of the phase center of the navigation antenna relative to the center of mass of the satellite (see Fig.7):

$$\begin{array}{c}\Delta D=D_0-D=\Delta D_r+\Delta D_f,\Delta {\text{D}}_{\text{r}}=l_r\cdot \frac{r-R\cdot \cos \theta }{D_0},\\ \Delta D_f=l_f\cdot \sin \theta \frac{R}{D_0},{\text{D}}_{\text{0}}^{\text{2}}=r^2+R^2-2rR\cdot \cos \theta ,\\ l_f=l_0\cdot \cos \alpha ,\end{array}\text{(14)}$$

where D₀- distance to the center of mass of the satellite; D is the distance from the phase center; ∆ D_r, ∆ D_f- the components of the uncertainty range from offset of the phase center of the antenna relative to the center of mass of the satellite; l_rlinear displacement of the phase center of the antenna along the first axis of the satellite; R is the radius of the Earth, θ is the angular position of the user relative to the radius vector of the orbit of the satellite r; l₀linear displacement of the phase center of the antenna relative to the first axis of the satellite.

It should be noted that the offset of the phase center of the antenna along the first axis of the satellite (l_r) leads to the constant component of the measurement error range that does not depend on spreads companion (∆D_r=const).

Calculations by the formula (14) maximum measurement uncertainty range for GLONASS satellites for the second component when l₀≈0.5 m and$${\theta }_{\max }\approx \arccos \frac{R}{r}$$give the value of 0≤ ∆ D_r≤0,13 m, which makes a significant contribution in determining the location of the consumer and in the calculations of care Board time of the satellite [1, 3].

Thus, the shadow orbits of the regular orientation of the satellite is carried out at all angles tracking except for the uncertainty intervals near the maximum and minimum values of the corners of the POPS, which is a disadvantage of this method.

The technical purpose of this invention is to improve the accuracy of navigation and time definitions consumer navigation satellites.

This technical problem is solved due to the fact that in the way that the orientation of the navigation satellites, including the orientation of the first axis of the satellite with the antenna on the Ground and the orientation of the solar panels in the Sun turn the satellite with solar panels relative to the first axis of the satellite to align normal to the solar panels to the plane of the Sun - Satellite - Earth and reversal of solar panels around the second axis of rotation perpendicular to the first, to align normal to the solar panels pointed at the Sun, carried out at predetermined intervals orbit, covering the intervals of uncertainty orientation of a satellite shadow orbits, the independence of the s proactive software turns around first and second axes of the satellite on the estimated size of the intermediate holding a given orientation.

Independent proactive software reversals can be implemented in various ways.

The way the orientation of a navigation satellite in the intervals of uncertainty, namely the orientation of the PSB in the Sun, achieved by proactive software pivot around the second axis of the satellite to align normal to the solar panels with the direction parallel to the first axis of the satellite, hold in this position and subsequent alignment normal to the solar panels pointed at the Sun at predetermined intervals orbit, covering the intervals of uncertainty satellite orientation and located symmetrically with respect to the maximum and minimum values of the angles "Sun - Satellite - Earth".

The combination of the normal to the PSB with the direction of the first axis X_Withsatellite (Δβ=0) and hold in this position (orthogonal position PSB) leads to the fact that at the turn of the satellite around the first axis (angle α_P) the angle φ does not change and is equal to the angle β (see equation (11)), and therefore, the value of the acceleration from the force of radiation pressure also does not change (see formula (12)...(13)). That is, for a given position of the PSB at predetermined values of the intervals, covering the intervals of uncertainty, the value of the acceleration from the force of radiation pressure is the predicted value, defined C is achenium angle β and the angle γ for a given value of the angle of declination η.

It should be noted that the maximum displacement of the satellite under the influence of light pressure forces created its lateral component, which changes its sign at the transition of the satellite through the points of the orbit γ=0° and γ=180°. Therefore, the organization intervals orthogonal position PSB symmetrically with respect to the maximum and minimum values of the angles POPS will lead to mutual compensation components of light pressure forces to the velocity vector (making a maximum contribution to the offset of the satellite along the orbit) and would eliminate the forecasting errors due to errors of orientation in the knowledge of the optical characteristics of solar cells due to their degradation, and to simplify the calculations, since there is no need to calculate the forces of radiation pressure to the velocity vector on the interval of uncertainty.

Implementation of the proposed method on the navigation satellite can be carried out by the following method.

Turn and hold the PSB can be performed using the native schema of the reversal of the bcop, is extended to align normal to the PSB with the direction parallel to the first axis of the satellite, retention PSB in this position and move to the standard orientation.

The calculation of the positions of the satellites in orbit, covering the intervals of uncertainty orientation PSB and hosted balanced is a rule regarding the maximum and minimum values of the angles POPS can be carried out using the following relationships (see Fig.8-10):

where t_I, t_O- times of entry and exit from the shadow of the Earth or from the zone of uncertainty orientation at small angles SES (high angles POPS); t₁- when the command was issued for the installation of PSB in the orthogonal position and lock the standard schema tracking PSB for the Sun angle β₁; t₂- the moment of fixation PSB in the orthogonal position; t₄torque lock tracking PSB behind the Sun; t₅- the beginning of the regular tracking PSB for the Sun.

Team upravleniyami operation of the satellite, issued at time t₁, t₂, t₄, t₅can be formed from a temporary program of the satellite and offline.

The values of t_I, t_O,$$\stackrel{}{\gamma }$$, η are determined by the well-known equations, based on the orbital parameters of the satellite at the beginning of each subsequent round, and the position of the Sun relative to the plane of the orbit.

The value of W_PSBis determined from the system parameters orientation of the satellite.

The presence of the intervals of the transition from regular tracking PSB orthogonal to its position (t₂-t₁, t₅-t₄) does not introduce errors in the calculations. Components of light pressure forces on the velocity vector are mutually excluded (due to symmetry), and the radius-vector are calculated by the formulas (12)...(13) under the following conditions: φ=180°-β when cosγ≥0 and φ=when β cosγ<0, i.e., these values are projected.

The way the orientation of the satellite in the intervals of uncertainty, namely the orientation of the antennas of the satellite to the Earth, can be implemented by two schemes depending on the structural design of the satellite in terms of placing the emissivity of the radiator thermal control system.

Under the first scheme (see Fig.11) proactive software turn around the first of sisutemu is to align the second axis of the satellite orbit plane hold in this position and subsequent combination of the second axis of the satellite normal to the plane of the Sun - Satellite - to-Ground at predetermined intervals orbit, covering the intervals of uncertainty orientation of solar panels and located symmetrically with respect to the maximum and minimum values of the angles "Sun - Satellite - Earth". In this scheme, the value of the turn angle in the interval τ₁-τ₂and τ₃-τ₄is α=90°|α₁|, where α₁- the value of the angle α (formula (2)) at time τ₁. Due to the chosen scheme spreads the Sun illuminates the surface of the satellite is outside the uncertainty intervals on one side only, coinciding with the positive direction of the Y axis_C. This allows you to arrange the radiation surface of the satellite side, coinciding with the negative direction of the Y axis_Ci.e. not illuminated by the Sun.

In the second scheme (see Fig.12) proactive software rotation about the first axis of the satellite is to align the second axis of the satellite and the normal to the orbit plane hold in this position and subsequent combination of the second axis of the satellite normal to the plane of the Sun - Satellite - Earth at predetermined intervals orbit, covering the intervals of uncertainty orientation of solar panels and located symmetrically with respect to the maximum and minimum values of the angles "Sun - Satellite - Earth". In this scheme, the amount is and the turn angle in the interval τ ₁-τ₂and τ₃-τ₄is α_TIMES=|α₁|.Due to the chosen scheme spreads the Sun illuminates the surface of the satellite is outside the uncertainty intervals of the first half revolution from the positive direction of Y_Cand the second half of the round from the side of the negative direction of axis Y_Cthat makes it inappropriate organization of the radiation surfaces of the satellite on those sides of the satellite, because they are illuminated by the Sun.

The calculation of the positions of the satellites in orbit, covering the uncertainty intervals of the positions of the phase centers of the antennas and placed symmetrically with respect to the maximum and minimum values of the angles POPS can be carried out using the following relationships (Fig.11, 12):

τ₁=τ₂-Δτ₁; τ₂=τ₃-Δτ₂; τ₄=τ₁+Δτ₂; τ₅=τ₄+Δτ₁

$$\Delta t_1\ge |\begin{array}{c}\frac{{90}^o-{\alpha }_1}{W_{P1}}\text{(scheme 1)}\\ \frac{{\alpha }_1}{W_{P1}}\text{(scheme 2)}\end{array}\text{(16)}$$

where α₁- the value of the angle α at time τ₁calculated by the formula (2), Δτ₂- duration fixed position, asked the technical capabilities of the control loop, Δτ₂≥0.

Command modes control of the satellite, issued at time τ₁τ₂τ₄τ₅can be formed from a temporary program of the satellite and offline.

The values of τ₃,$$\stackrel{}{\gamma }$$, η, θ are determined by the well-known equations based on the orbital parameters of the satellite at the beginning of each subsequent round, and the position of the Sun relative to the plane of the orbit, the position of the consumer in a geographic coordinate system.

The value of W_P1, l₀is determined from the design parameters of the satellite. The presence of the intervals of the transition from regular satellite orientation orthogonal to the position of its axis relative to the orbital plane (τ₂-τ₁τ₅-τ₄) does not introduce errors in the calculations. The adjustments range on these intervals is carried out by consumers using the formulas (14) at a known position of the user relative to the satellite (the angle θ) and the known (or predicted) by the law of changing the angle of the software reversal of the satellite around-ear, closed the g first axis (angle α).

Thus, the technical result of the claimed method is:

- increase the accuracy of predicting the motion of a satellite, shadow orbits due to the reduction of the unpredictable components of the acceleration from the forces of radiation pressure;

- increase the accuracy of the measurement range due to the reduction predicted values of the angles of a reversal of the satellite around the first axis.

Sources of information

1. Chebotarev C. E. fundamentals of spacecraft information support: textbook. manual/C. E. Chebotarev, V. Kosenko E.; Sib. state. aerocosmic. University, Krasnoyarsk, 2011. - 488 C.[24] with silt.

2. Eliasberg P. E. introduction to theory of flight of the satellite. - 2nd ed. - M.: Librokom, 2011. - 544 S.

3. w.w.w.elsevier.com/locate/asr. The GLONASS - M satellite yaw-attitude model/. F. Dilssner, T. Springer, G. Gienger, I. Dow. ESOC, 2010.

1. The way the orientation of the navigation satellites, including the orientation of the first axis of the satellite antenna on the Ground and the orientation of the solar panels in the Sun turn the satellite with solar panels relative to the first axis of the satellite to align normal to the solar panels to the plane of the Sun-satellite-Earth and spread of solar panels around the second axis of rotation perpendicular to the first, to align normal to the solar panels pointed at the Sun, characterized in that the given interval is x orbit, covering the intervals of uncertainty orientation of a satellite shadow orbits are independently proactive software turns around first and second axes of the satellite on the estimated size of the intermediate holding a given orientation.

2. The way the orientation of the navigation satellite under item 1, characterized in that the proactive software pivot around the second axis of the satellite is to align normal to the solar panels with the direction parallel to the first axis of the satellite and hold in this position and subsequent combination of the normal to the solar panels pointed at the Sun at predetermined intervals orbit, covering the intervals of uncertainty satellite orientation and located symmetrically with respect to the maximum and minimum values of the angles "Sun-satellite-Earth".

3. The way the orientation of the navigation satellite under item 1, characterized in that the proactive software rotation about the first axis of the satellite is to align the second axis of the satellite orbit plane hold in this position and subsequent combination of the second axis of the satellite normal to the plane of the Sun-satellite-Earth at predetermined intervals orbit, covering the intervals of uncertainty orientation of solar panels and located symmetrical relative to the positive maximum and minimum values of the angles "Sun-Satellite-Earth".

4. The way the orientation of the navigation satellite under item 1, characterized in that the proactive software rotation about the first axis of the satellite is to align the second axis of the satellite and the normal to the orbit plane hold in this position and subsequent combination of the second axis of the satellite normal to the plane of the Sun-satellite-Earth at predetermined intervals orbit, covering the intervals of uncertainty orientation of solar panels and located symmetrically with respect to the maximum and minimum values of the angles "Sun-satellite-Earth".