RU2354592C2 - Method of determining spacecraft solar battery maximum output - Google Patents

Abstract

FIELD: transport, space engineering.

SUBSTANCE: propose method designed to determine spacecraft solar battery maximum output comprises measuring the spacecraft orbit altitude for the angular half-width of the Earth disk visible from the spacecraft (Q_z) and the angle of elevation of atmosphere top boundary (ε) above the Earth horizon visible from the spacecraft, determining the angular half-width of the Sun disk visible from the spacecraft (Q_s) and measuring spacecraft orbital angular velocity (ω). It includes also measuring the angle between direction to the Sun and spacecraft orbit plane (β), measuring the angle of elevation of direction to the Sun above the Earth horizon visible from spacecraft (g) and maximum solar battery output at the batteries minimum temperature. The system, to this end, comprises the unit to measure the angle between direction to the Sun and spacecraft orbit plane, unit to measure the angle of elevation of direction to the Sun above the Earth horizon visible from spacecraft, the unit to measure spacecraft orbital angular velocity, the unit to measure the spacecraft orbit altitude variation, the unit to determine the angle of elevation of atmosphere top boundary (ε) above the Earth horizon visible from the spacecraft, the unit to determine the angular half-width of the Sun disk visible from the spacecraft. It incorporates also the unit to determine the orbit whereon solar battery maximum output is determined, the unit to generate identification parameters of solar batteries temperature conditions and key.

EFFECT: higher accuracy of determining solar battery maximum output.

2 cl, 4 dwg

Description

The invention relates to the field of space technology, namely, power supply systems (SES) of spacecraft (SC), and can be used in the operation of solar panels (SB) SES SC.

The main electrical characteristic of the SB is the maximum output power of the SB (this power differs from the current actual output power, which depends on the load and on the influence of the environment). At the design and manufacturing stage of the SB (before the SC launch), a theoretical calculation of the SB performance is carried out, which is also called the calculation of the SB output parameters (see [1]; see [2], p. 49). A theoretical calculation of the operational characteristics of the SB and prediction of the process of their degradation under the influence of space flight factors is carried out using a computer, while the calculation of the output parameters of the SB is based on the method of moving the current-voltage characteristics, taking into account various environmental influences and load parameters on the characteristics of the SB (see [ 2], p. 54).

The disadvantage of this method for determining the maximum output power of the SB is that the models of space flight factors used in the calculations have limited accuracy, which does not allow obtaining reliable data on the actual characteristics of the SB in flight.

To control the actual characteristics of the SB during flight, special flight operations are carried out - sessions of evaluating the effectiveness of the SB, in which the actual maximum output power of the SB is measured. Based on the results of determining the maximum output power of the SB, the current effectiveness of the SB is estimated as the ratio of the measured maximum output power of the SB to its nominal value - the design or some initial value (for example, at the time the spacecraft began to function).

To determine the SB output power, the system presented in [2], p. 48, which implements a model of a solar cell connected to a load and containing a solar cell connected to an ideal wattmeter unit connected to a load, can be used. In this case, the unit of an ideal wattmeter contains current sensors (ammeter) and voltage (voltmeter) and a calculator that determines the output power P of the solar cell by the formula:

where I is the measured current value from the solar cell;

V is the measured voltage.

The closest of the analogues adopted for the prototype is a method for determining the maximum output power of the SB SC, described in [3], p.17-18. The essence of the method is as follows. To determine the maximum output power of the SB, use the measured value of the maximum current from the SB I _m - the current generated when the illuminated working surface of the SB panels is oriented perpendicular to the sun's rays. For this, the SB panels are deployed to a position corresponding to the normal to the working surface of the SB panels with the direction to the Sun. This orientation of the SB provides the maximum supply of electricity from the SB.

The system for implementing the method adopted for the prototype is described in [3], p. 6, and contains an SB, on a rigid substrate of which there is a block of photovoltaic batteries (BFB), a rotation device of the SB (UPSB), an amplifier-conversion device (UPU) , a control unit for the orientation of the SB in the direction to the Sun (БУСБС), a block of current regulators (BRT), a current sensor (ДТ), a power determination unit (БОМ), a power supply system control unit (BUSES), and a power supply bus (ШЭ). In this case, the output of the BSE located on the SB is connected to the BRT, the output of which is connected to the ST, and the BOM is connected by its first and second inputs to, respectively, the ST and DT, which is connected, in turn, to the ST, and the output of the BUSES is connected to the input BUOSBS, whose output is connected to the input of the UPU, the output of which is connected to the input of the UPSB, the output of which is connected to the second input of the UPS, and the UPSB is mechanically connected to the SB. Moreover, in the description of the prototype system and its block diagram (see [3], pp. 6-8), the BOM function is realized when the SES charger unit is functioning.

The system operates as follows.

On command from BUSES (9), the BUOSBS unit (5) controls the orientation of the SB (1) to the Sun. Input information for the SB control algorithm (1) are: the position of the unit direction vector on the Sun relative to the coordinate axes associated with the spacecraft; the position of the SB relative to the spacecraft body, obtained in the form of the current measured values of the angle α between the current position of the normal to the working surface of the SB and the direction to the Sun from angle sensors (ДУ) installed on UPSB (3). When the SB is oriented to the Sun, α≈0. The output of the control algorithm are commands for rotating the SB relative to the axis of the output shaft of the UPSB (3) and commands for stopping rotation, and the remote control of the UPSB (3) give signals about the current position of the SB (1).

The UPU (4) plays the role of an interface between the BSECS (5) and UPSB (3).

Electricity from the BSE (2) installed on the SB (1), through the BRT (6) is supplied to the power supply bus of the SES SHE (10). DT (7) measures the current current on the BW (10) and the measured current value is supplied to the BOM (8), in which the voltage on the BW (10) is measured and the value of the output power of the SB is calculated by relation (1).

It is known (see [2], pp. 10-15) that the efficiency of solar cells depends on their temperature. Moreover, the analysis of the output parameters of the SB (see [2], pp. 49-51) includes an analysis of the SB operation in the entire temperature range, both at the operating temperature and taking into account the temperature during darkening and unlit SB.

The method and system adopted for the prototype have a significant drawback - they do not allow to take into account the temperature regime of the SB when determining the maximum output power of the SB and evaluating their effectiveness.

The challenge facing the invention is to increase the accuracy of determining the maximum output power of the SB by taking into account the temperature regime of the SB, namely, to determine the maximum output power of the SB both at the maximum steady-state operating temperature and at the minimum temperature of the SB.

The technical result is achieved in that in a method for determining the maximum output power of solar panels of a spacecraft, including turning solar panels into a working position, corresponding to combining the normal to their illuminated work surface with the direction to the Sun, measuring the current value from solar panels, measuring voltage and determining the maximum output power of the solar panels of the spacecraft as a product of the measured values of voltage and current from solar batteries, additional They carefully measure the orbit height of the spacecraft and determine from it the values of the angular half-solution of the Earth’s disk visible from the spacecraft (Q _z ) and the elevation angle of the upper atmosphere above the Earth’s horizon visible from the spacecraft (ε), determine the value of the angular half-solution of the Sun’s disk visible from the spacecraft (Q _s ), measure the angular velocity of the orbital motion of the spacecraft (ω), measure the angle between the direction to the Sun and the plane of the orbit of the spacecraft (β), on turns in which the value of the measured angle β is less than or equal to the value of β °, determined by the formula

β ° = arccos (cos Q _z / cos (Δt ° ω / 2)},

where Δt ° is the required minimum allowable duration of the shadow portion of the orbit,

they measure the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft (g) and the maximum output power of solar panels at their minimum temperature is determined as the product of the voltage and current from solar panels measured at the moment the solar disk touches the upper boundary of the atmosphere visible from the spacecraft Earth at sunrise, determined from the condition that the measured angle g is equal to the sum of the angles ε and Q _s with increasing angle g, and the maximum output power solar cells at the maximum steady-state operating temperature is defined as the product of the voltage and current from the solar cells measured at the moment the solar disk visible from the spacecraft touches the upper boundary of the Earth’s atmosphere at sunset, determined from the condition that the measured angle g is equal to the sum of the angles ε and Q _s as the angle g decreases.

In this case, the task is solved in that in the system for determining the maximum output power of the solar panels of the spacecraft, including a solar battery with a block of photovoltaic batteries installed on it, a device for turning solar panels, an amplifying-converting device, a control unit for orienting the solar panels in the direction to the Sun , a block of current regulators, a current sensor, a power determination unit, a power supply system control unit and a power supply bus, while the output of the photoelectric unit batteries is connected to the input of the current regulator unit, the output of which is connected to the power supply bus, to which the current sensor and the power determination unit are connected, and the output of the power supply control unit is connected to the first input of the control unit for orientation of the solar batteries in the direction to the Sun, the output of which is connected to the input of the amplifier-converting device, the output of which is connected to the input of the solar battery rotation device, the output of which is connected to the second input of the orientation control unit with solar batteries in the direction to the Sun, and the solar battery rotation device is mechanically connected to the solar battery, an angle measurement unit between the direction to the Sun and the orbit plane of the spacecraft, a unit for measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft, a measurement unit the angular velocity of the orbital motion of the spacecraft, the unit for measuring the height of the orbit of the spacecraft, the unit for determining the elevation angle of the upper boundary of the atmosphere the Earth’s spheres above the Earth’s horizon visible from the spacecraft, the half-angle block for determining the solar disk visible from the spacecraft, the coil determination unit for determining the maximum output power of solar batteries, the unit for determining the moments for determining the maximum output power of solar batteries, the unit for generating identification parameters of solar temperature batteries and a key, while the second input of the power determination unit is connected to the output of the key, information and control the inputs of which are connected to the outputs, respectively, of the current sensor and the unit for determining the moments of determining the maximum output power of solar batteries, the first to fourth inputs of which are connected to the outputs, respectively, of the unit for measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft, block determining the half-angle angle of the solar disk visible from the spacecraft, the block determining the angle of elevation of the upper boundary of the Earth's atmosphere above the horizontal horizon visible from the spacecraft ntom of the Earth and the unit for determining the turn of the operation of determining the maximum output power of solar batteries, the first, second and third inputs of which are connected to the outputs, respectively, of the unit for measuring the angle between the direction to the Sun and the orbital plane of the spacecraft, the unit for measuring the angular velocity of the orbital motion of the spacecraft and a unit for measuring the height of the orbit of a spacecraft, the output of which is also connected to the input of the unit for determining the angle of elevation of the upper boundary of the Earth’s atmosphere above visible m from the spacecraft to the Earth’s horizon, and the output of the block measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft is also connected to the input of the unit for identifying the temperature parameters of solar batteries, the output of which is connected to the third input of the power determination unit.

The essence of the invention is illustrated by drawings, which show: in Fig.1 - a diagram of the lighting of the SB at sunset and sunrise; figure 2 - lighting scheme of the orbit of the spacecraft orbit by the Sun; figure 3 is a block diagram of a system for implementing the proposed method; figure 4 is a graph of the arrival of electricity from the SB of the Russian segment (PC) of the international space station (ISS).

Let us explain the proposed method of action.

The equilibrium operating temperature of the SB (T _slave ) and the temperature of the unlit SB (T _t ) are determined by the thermomechanical and electrical properties of the SB and, for example, can be calculated from the ratios presented in [2], p.90. The SB temperature when it enters the shadow depends on the dimming time, but if the shadow period lasts more than a certain time (in [2], p. 96, an estimate of this time is given for 30 min), then the temperature of the unlit SB is practically independent of T _slave , t .e. at the end of the shadow, the minimum SB temperature in flight is reached. In this case, the equilibrium operating temperature in the steady state operating mode of the SB is the maximum temperature of the SB in flight.

In spacecraft flight, these minimum and maximum temperatures are guaranteed to be reached, respectively, by the time the shadow portion of the orbit ends (provided that its duration is not less than the required value) and by the time the SB reaches the established temperature mode of operation in the illuminated portion of the orbit. This mode is established after a certain time after the spacecraft comes to light (for example, 15-20 minutes for the ISS PC PC), of course, under the obvious condition that the satellites are oriented to the Sun and are not obscured from the Sun by the spacecraft design elements. Thus, it is assumed that by the end of the illuminated portion of the orbit of the spacecraft, satellites in the working position — oriented toward the sun and illuminated by the sun (not obscured by the spacecraft’s structural elements), are guaranteed to reach a steady-state temperature regime.

Along with this, it is known (see [2], p. 272) that solar radiation entering the Earth is reflected from its surface, from clouds, and is scattered by the atmosphere. The energy of reflected radiation, concentrated in the spectral range of the sensitivity range of solar cells SB, is perceived by SB and increases their output power. Thus, in addition to direct solar radiation, the SB receives a stream of radiation reflected from the Earth, which introduces uncertainty in determining the maximum output power of the SB and in the subsequent solution of the problem of evaluating their effectiveness. Uncertainty in this case lies in the presence of not predicted and not subject to overestimation of the obtained values of the output power of the SB.

However, at the beginning and end of the illuminated portion of the spacecraft’s orbit, the direction of the flow of solar radiation entering the spacecraft passes tangentially to the Earth’s surface — in this case, the radiation reflected from the Earth entering the spacecraft’s SB does not arise except from radiation from the limb formed by the atmosphere illuminated by the Sun The Earth, whose influence on the power generation of the SB is negligible in comparison with the direct radiation coming from the Sun to the SB. This, in particular, is illustrated by the graph shown in figure 4, - at the time of sunset (when the SB works in a steady temperature mode), the increase in energy supply to the SB does not occur.

Thus, at the end of the illuminated portion of the spacecraft’s orbit, the SS reaches the steady-state temperature mode of operation and, at the same time, there is no unpredictable and not accountable overestimation of the SB output power values from the possible impact of radiation reflected from the Earth on the SB (i.e., there is no negative effect of the reflected radiation from the Earth to the operation of determining the maximum output power of the SB).

In the proposed technical solution, when determining the maximum output power of the SB, the output power of the SB is determined in the maximum possible temperature range of the SB - at the minimum temperature at the beginning of the illuminated portion of the orbit and at the maximum steady-state operating temperature at the end of the illuminated portion of the orbit.

In Fig. 1, which explains the lighting scheme of the SB at sunset and sunrise, the construction is performed in the plane formed by the radius vector of the spacecraft and the direction to the Sun, and the notation is introduced:

R - SB KA;

N is the normal to the working surface of the SB panels;

C is the center of the disk of the Sun visible from the spacecraft;

S, PC — direction vector from the spacecraft to the center of the solar disk;

Z - Earth;

O is the center of the earth;

D is the surface of the earth;

G is the upper boundary of the Earth's atmosphere;

H _atm - the height of the upper boundary of the Earth’s atmosphere from the Earth’s surface;

Q _z is the half-angle of the Earth’s disk visible from the spacecraft;

R _z is the radius of the Earth;

H _orb - the height of the orbit of the spacecraft;

g is the angle of elevation of the direction to the Sun above the Earth's horizon visible with the SC;

ε is the elevation angle of the upper boundary of the Earth’s atmosphere above the Earth’s horizon visible with the SC;

Q _s is the half-angle of the solar disk visible from the spacecraft.

We consider the current orientation of the SB, in which the normal to the working surface of the SB panels N is aligned with the direction to the Sun S.

After the spacecraft comes to light, the earliest point in time (t ₁ ), when the spacecraft is illuminated by the full solar disk (without “shading” part of the solar disk visible from the spacecraft by the Earth’s atmosphere), corresponds to the following condition (see figure 1):

The same condition determines the latest time (t ₂ ) on the illuminated part of the orbit, when the spacecraft is illuminated by radiation from the full solar disk. When the angle g exceeds this value (between time instants t ₁ and t ₂ ), the effect of the additional arrival of radiation reflected from the Earth to the SC SB will already begin to negatively manifest.

The layer of the Earth’s atmosphere, which scatters the radiation arriving at the spacecraft from the Sun, is set by the height of its upper boundary from the Earth’s surface N _atm (see [7], pp. 19-20, pp. 128-160). The determination of the angle ε is carried out according to the following formula:

Where

The determination of the angle Q _{s is} carried out according to the technique used in calculating the tables of the apparent radius of the Sun of the Astronomical Yearbook (see [8], p. 624).

Figure 2, representing the lighting scheme of the orbit of the spacecraft orbit by the Sun, is additionally indicated:

N _orb is the normal to the orbital plane of the spacecraft;

V is the spacecraft velocity vector;

β is the angle between the direction to the Sun and the orbital plane of the spacecraft;

T _in , T _s - the position of the spacecraft at the moments, respectively, of sunrise and sunset t _in and t _s (at the moments of the beginning and end of the illuminated portion of the orbit);

T _t and T _s - the position of the spacecraft at the moments of the middle of the shadow portion and the middle of the illuminated portion of the orbit;

T _in K - the projection of the direction to the Sun on the plane of the orbit of the spacecraft;

γ is the angle between the directions of T _in O and T _in K;

T ₁ , T ₂ - the position of the spacecraft at time, respectively, t ₁ and t ₂ .

The duration of the shadow portion of the orbit Δt:

is determined from the equality of the angles T _t OT _in , T _t OT _s , OT _in K and from the ratio obtained as a solution of a rectangular spherical triangle formed by the directions T _in C, T _in O, T _in K, in which the angle ST _in K = β , the angle CT _in O = Q _z , the angle OT _in K = γ:

where ω is the angular velocity of the orbital motion of the spacecraft.

The condition for the duration of the shadow is not less than the specified value Δt °:

determined by the fulfillment of the ratio:

Thus, condition (10) is satisfied only on the turns at which the value of the angle β is bounded above by the maximum allowable value β °:

In this case, the maximum possible value Δt ° realized for a particular spacecraft is given by the relation

To implement the method, a system is proposed, presented in figure 3 and containing the following blocks:

1 - SB; 2 - BSE; 3 - UPSB; 4 - UPU; 5 - BUOSBS; 6 - BRT; 7 - DT; 8 - BOM; 9 - BUSES; 10 - SE;

11 is a block measuring the angle between the direction of the Sun and the plane of the orbit of the spacecraft (BIUSPOKA),

12 - block determining the turn of the operation of determining the maximum output power of solar batteries (BOVOMVMSB);

13 - unit for measuring the angle of elevation of the direction to the Sun above the Earth's horizon visible from the spacecraft (BIUVSGZ);

14 - unit for determining the moments of determining the maximum output power of solar batteries (BOMOMVMSB);

15 - unit for measuring the angular velocity of the orbital motion of the spacecraft (BIUSODKA);

16 - unit for measuring the height of the orbit of the spacecraft (BIVOKA);

17 is a block for determining the elevation angle of the upper boundary of the Earth’s atmosphere above the Earth’s horizon visible with the SC (BOUVVGAGZ).

18 - block determining the half-angle of the solar disk visible from the spacecraft (BOPSA);

19 - unit for the formation of identification parameters of the temperature conditions of solar batteries (BFIPTRSB);

20 is the key

the output of the BFB (2) located on the SB (1) is connected to the BRT (6), the output of which is connected to the BW (10), to which the DT (7) and the BOM (8) are connected, the second input of which is connected to the output key (20). The information and control inputs of the key (20) are connected to the outputs, respectively, DT (7) and BOMOMVMSB (14). From the first to the fourth inputs of the BOMOMVMSB (14) are connected to the outputs, respectively, BIUVSGZ (13), BOUPDS (18), BOUVVGAGZ (17) and BOVOMVMSB (12). The first, second and third inputs of BOVOMVMSB (12) are connected to the outputs, respectively, BIUSPOKA (11), BIUSODKA (15) and BIVOKA (16). The output BIVOKA (16) is also connected to the input BOUVVGAGZ (17). The output of BIUVSGZ (13) is also connected to the input of BFIPTRSB (19), the output of which is connected to the third input of the BOM (8). The output of the BUSES (9) is connected to the input of the UPS (5), the output of which is connected to the input of the UPU (4), the output of which is connected to the input of UPSB (3), the output of which is connected to the second input of the UPS (5). UPSB (3) is mechanically connected to SB (1).

In Fig. 3, the dotted line shows the mechanical connection of the UPSB (3) with the SB body (1) through the output shaft of the SB drive.

The system operates as follows.

On command from BUSES (9), the BUOSBS unit (5) controls the orientation of the SB (1) to the Sun. Input information for the SB control algorithm (1) are: the position of the unit direction vector on the Sun relative to the coordinate axes associated with the spacecraft; the position of the SB relative to the spacecraft body, obtained in the form of the current measured values of the angle α between the current position of the normal to the working surface of the SB and the direction to the Sun from angle sensors (ДУ) installed on UPSB (3). When the SB is oriented to the Sun, α≈0. The output of the control algorithm are commands for rotating the SB relative to the axis of the output shaft of the UPSB (3) and commands for stopping rotation, and the remote control of the UPSB (3) give signals about the current position of the SB (1).

The UPU (4) plays the role of an interface between the BSECS (5) and UPSB (3).

Electricity from the BSE (2) installed on the SB (1), through the BRT (6) is supplied to the power supply bus of the SES SHE (10). DT (7) measures the current current on the ST (10) and the measured current value is supplied to the information input of the key (20). In the initial state, the key (20) is “closed”.

In BIUSPOK (11), the angle between the direction to the Sun and the orbit plane of the spacecraft β is determined. In BIUSODKA (15), the angular velocity of the spacecraft orbital motion ω is measured. In BIVOKA (16), the measurement of the orbit height of the spacecraft N _orb is performed. The obtained values are transmitted to BOVOMVMSB (12).

In BOVOMVMSB (12), the angle Q _z is determined by relation (4), β ° is determined by relation (16), and condition (15) is checked to determine whether the maximum output power of the SB can be determined at the current orbit of the spacecraft. When condition (15) is met, BOMOMVMSB (12) generates a signal supplied to the input of BOMOMVMSB (14).

At the same time, in the BIUVSGZ (13), the current elevation angle of the Sun above the Earth's horizon g, visible with the SC, is determined. The measured value of the orbit height from BIVOKA (16) is also transmitted to BOUVVAGAGZ (17), in which the value of the angle ε is determined from relations (3), (4). In BOPSDS (18), the half-angle angle of the solar disk Q _s visible from the spacecraft is determined. These values are received by the BOMOMVMSB (14).

Upon receipt of the control signal from BOVOMVMSB (12), BOMOMVMSB (14) by checking conditions (2) determines the moments t ₁ and t ₂ at which current current measurements from the SB will be performed to determine the maximum output power of the SB at different temperature conditions of the SB functioning. When condition (2) is fulfilled, the BOMOMVMSB (14) generates a signal supplied to the control input of the key (20).

In BFIPTRSB (19), the identification parameters of the temperature conditions of the SB are formed. For example, an identification parameter is formed equal to “1” when the angle g increases (this value corresponds to the case of the minimum SB temperature at sunrise) and equal to “2” when the angle g decreases (this value corresponds to the case of the maximum steady-state working temperature of the SB at sunset). The values of the generated parameters are transferred to the BOM (8).

Upon receipt of the signal at the control input of the key (20), the key (20) is “open” and through it the measured current value from the DT (7) is supplied to the BOM (8), in which the voltage is measured at the ST (10) and according to the relation (1) ) the value of the output power of the SB is calculated. In this case, the identification of the temperature regimes of the SB operation, at which the output power of the SB were determined, is carried out by the corresponding identification parameters received from the BFIPTRSB (19).

Implementation of BIUSPOK (11), BIUVSGZ (13), BIUSODKA (15), BIVOKA (16) can be performed on the basis of sensors and equipment of the Control and Navigation System (SUDN) and the On-Board Digital Computing System (BTsVS) of the spacecraft (see [4] , [5]). Implementation of BOMOMVMSB (12), BOMOMVMSB (14), BOUVVGAGZ (17), BOUPDS (18) and BFIPTRSB (19) can be performed on the basis of the BCVS KA. The key (20) can be made in the form of an elementary analog circuit. SB (1), BSE (2), UPSB (3), UPU (4), BUOSBS (5), BRT (6), DT (7), BOM (8), BUSES (9) can be made on the basis of elements SES (see [1]).

Thus, an example of the implementation of the fundamental blocks of the system is considered, according to the results of which a decision is made and the proposed operations are implemented.

We describe the technical effect of the proposed inventions.

The proposed technical solution allows to increase the accuracy of determining the maximum output power of the SB by taking into account the temperature regime of the SB, namely, ensuring the determination of the maximum output power of the SB both at the maximum steady-state operating temperature and at the minimum temperature of the SB. At the same time, an increase in the accuracy of determining the output power of the SB is also achieved by minimizing (eliminating) the use of radiation reflected from the Earth to generate electricity during the operation to determine the maximum output power of the SB, which eliminates the unpredictable and not subject to overstating current values of the output power of the SB.

For illustration, Fig. 4 shows a graph of the arrival of electricity from SB SB MKC I (A) versus time t (sec) during an orbit while maintaining the SB orientation to the Sun: 02.02.2004, turn 1704 (telemetric number), time 17.35 -19.06 DWM, orientation of ISK (see [6]). The time points t ₁ and t ₂ corresponding to the fulfillment of condition (2) at sunrise (with increasing angle g) and at approach (with decreasing angle g) are marked on the graph; for MKC orbit, these time moments are very close to the times of sunrise and sunset . Horizontal dashed lines indicate the levels of electric power arrival at time t ₁ and t ₂ .

Information sources

1. The power supply system of the spacecraft. Technical description. 300GK.20Yu. 0000-ATO. RSC Energia, 1998.

2. Rauschenbach G. Handbook for the design of solar panels. Moscow, Energoatomizdat, 1983.

3. Kovtun B.C., Soloviev S.V., Zaikin S.V., Gorodetsky A.A. A method for controlling the position of solar panels of a spacecraft and a system for its implementation. Description of the invention to the patent of the Russian Federation No. 2242408 according to the application 2003108114/11 of March 24, 2003.

4. Engineering reference for space technology. Publishing House of the Ministry of Defense of the USSR, M., 1969.

5. The motion control and navigation system of the spacecraft. Technical description. 300GK.12YU. 0000-ATO. RSC Energia, 1998.

6. Rulev D.N., Stazhkov V.M., Korneev A.P., Panteleimonov V.N., Melnik I.V. Evaluation of the performance of solar panels in the Russian segment of the international space station according to telemetric information // Proceedings of the XXXIX Readings on the development of the scientific heritage and development of ideas of K.E. Tsiolkovsky (Kaluga, September 14-16, 2004). Section "Problems of rocket and space technology." - Kazan: Kazan State University. V.I. Ulyanov-Lenin. 2005.

7. Kroshkin M.G. Physical and technical foundations of space research. - M.: Mechanical Engineering. 1969.

8. Astronomical Yearbook of the USSR for 1993. Leningrad, "Science", 1991.

Claims (2)

1. A method for determining the maximum output power of solar panels in a spacecraft, including turning solar panels into a working position corresponding to combining the normal to their illuminated work surface with the direction to the Sun, measuring the current value from solar panels, measuring voltage and determining the maximum output power of solar panels spacecraft as a product of the measured values of voltage and current from solar panels, characterized in that they additionally measure the height from the orbit of the spacecraft and determine from it the values of the angular half-solution of the Earth's disk visible from the spacecraft (Q _z ) and the elevation angle of the upper boundary of the atmosphere above the Earth's horizon visible from the spacecraft (ε), determine the value of the angular half-solution of the solar disk visible from the spacecraft (Q _s ), measure the angular velocity of the orbital motion of the spacecraft (ω), measure the angle between the direction to the Sun and the plane of the orbit of the spacecraft (β), in turns on which the value of the measured angle β is less than or equal to β ° determined by the formula
β ° = arcos {cosQ _z / cos (Δt ° ω) / 2)},
where Δt ° is the required minimum allowable duration of the shadow portion of the orbit, the angle of elevation of the direction to the Sun above the Earth's horizon (g) visible from the spacecraft is measured and the maximum output power of solar batteries at their minimum temperature is determined as the product of the voltage and current from solar batteries, measured at the moment of contact with the disc of the Sun visible from the spacecraft on the upper boundary of the Earth’s atmosphere at sunrise, determined from the condition of equality of the measured angle g is the sum of the angles ε and Q _s as the angle g increases, and the maximum output power of the solar panels at the maximum steady-state working temperature is determined as the product of the voltage and current from the solar cells measured at the moment the solar disk touches the upper boundary of the Earth’s atmosphere at sunset, determined from the condition that the measured angle g is equal to the sum of the angles ε and Q _s when the angle g decreases. 2. A system for determining the maximum output power of solar cells in a spacecraft, including a solar battery with a block of photovoltaic batteries installed on it, a solar rotation device, an amplifying-converting device, a solar orientation control unit in the direction to the Sun, a block of current regulators, a current sensor, power determination unit, power supply system control unit and power supply bus, wherein the output of the photovoltaic battery unit is connected to the input of the regulator current generators, the output of which is connected to the power supply bus, to which a current sensor and a power determination unit are connected, and the output of the power supply control unit is connected to the first input of the solar orientation control unit in the direction to the Sun, the output of which is connected to the input of the amplifying-converting device, the output of which is connected to the input of the solar rotator, the output of which is connected to the second input of the solar orientation control unit in the direction to the Sun, m the device for turning the solar panels is mechanically connected to the solar battery, characterized in that it additionally introduces a unit for measuring the angle between the direction to the Sun and the plane of the orbit of the spacecraft, a unit for measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft, and a unit for measuring the angular velocity of the orbital the motion of the spacecraft, the unit for measuring the height of the orbit of the spacecraft, the unit for determining the angle of elevation of the upper boundary of the Earth’s atmosphere above the visible of the osmotic apparatus with the Earth’s horizon, a block for determining the half-angle of the solar disk visible from the spacecraft, a block for determining the turn of the operation to determine the maximum output power of solar batteries, a unit for determining moments of determining the maximum output power of solar batteries, a unit for generating identification parameters of the temperature regimes of solar batteries and a key, when the second input of the power determination unit is connected to the output of the key, the information and control inputs of which are connected are connected with the outputs, respectively, of the current sensor and the unit for determining the moments of determining the maximum output power of solar batteries, the first to fourth inputs of which are connected to the outputs, respectively, of the unit for measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft, and the unit for determining the half-angle the disk of the Sun visible from the spacecraft, the block for determining the elevation angle of the upper boundary of the Earth’s atmosphere above the Earth’s horizon visible from the spacecraft, and the block a turn of the operation for determining the maximum output power of solar cells, the first, second and third inputs of which are connected to the outputs, respectively, of the angle measuring unit between the direction to the Sun and the orbit plane of the spacecraft, the unit for measuring the angular velocity of the orbital motion of the spacecraft and the unit for measuring the orbit height spacecraft, the output of which is also connected to the input of the unit for determining the angle of elevation of the upper boundary of the Earth’s atmosphere above the visible from the spacecraft the Earth’s horizon, and the output of the block measuring the angle of elevation of the direction to the Sun above the Earth’s horizon visible from the spacecraft is also connected to the input of the unit for identifying the temperature parameters of the solar batteries, the output of which is connected to the third input of the power determination unit.

Images