A method for on-orbit autonomous temperature control of satellite-mounted solar panels

Abstract

A method for autonomous on-orbit temperature control of satellite-mounted solar panels is proposed. This method monitors the temperature of the satellite's solar panels in real time and compares it with a temperature alarm threshold. When the measured temperature exceeds the alarm threshold but the difference is not greater than a temperature threshold, the method uses a solar panel temperature calculation model to determine the required current shunt to bring the solar panel temperature back below the alarm threshold. The shunt is then reduced by increasing the load current, thus lowering the solar panel temperature. Conversely, when the measured temperature exceeds the alarm threshold but the difference is greater than the temperature threshold, the method uses the same model to determine the required satellite offset angle to bring the solar panel temperature back below the alarm threshold. This offset method alters the angle of solar radiation incidence to further reduce the solar panel temperature. This method enables precise temperature control of the solar panels under energy balance conditions, ensuring normal on-orbit operations while effectively mitigating the high-temperature problem of the solar panels.

Description

Technical Field

[0001] This invention belongs to the field of satellite energy control and relates to a method for on-orbit autonomous temperature control of satellite solar panels. Background Technology

[0002] With the rapid development of aerospace technology, the application of body-mounted solar arrays in low-cost small satellites is becoming more and more widespread, especially in microsatellites. In order to achieve miniaturization, satellite structural plates are often used as solar array substrates, and the solar cells are attached to the satellite structure to form body-mounted solar arrays.

[0003] Because the body-mounted solar array is in close contact with the satellite body, heat dissipation is poor, resulting in a much higher temperature compared to the deployed solar array. Furthermore, since body-mounted solar arrays are typically in a sun-facing cruising attitude, there is a risk that the solar array temperature will exceed its upper limit when directly facing the sun. In the shadow region, the solar array temperature decreases, and extreme temperature fluctuations can affect the reliability of the solar array's operation. Therefore, effective methods for high-temperature control of the solar array are needed in orbit to address the problem of excessively high temperatures in body-mounted solar arrays.

[0004] Existing temperature control methods for deployable solar arrays are clearly unsuitable for solving the high-temperature problem of body-mounted solar arrays. Summary of the Invention

[0005] The technical problem solved by this invention is to overcome the shortcomings of the prior art and propose an autonomous temperature control method for satellite-mounted solar arrays in orbit, which can ensure that the temperature of the solar arrays in orbit is within the normal range under different operating conditions, thereby improving the reliability of the solar arrays in orbit.

[0006] The technical solution of this invention is: an on-orbit autonomous temperature control method for satellite-mounted solar panels, comprising:

[0007] A solar array temperature calculation model is established to calculate the theoretical value of the solar array temperature on the satellite body. At the same time, the measured value of the solar array temperature on the satellite body is acquired in real time. The measured value is used to correct the model so that the error between the two is within the allowable calculation error range.

[0008] The temperature of the satellite's solar array is monitored in real time and compared with the temperature alarm value. When the measured temperature exceeds the temperature alarm value and the difference between the two is not greater than the temperature threshold, the model is used to calculate the amount of current required for the satellite's solar array temperature to return to below the temperature alarm value. The current is then reduced by increasing the load current to lower the temperature of the satellite's solar array. When the measured temperature exceeds the temperature alarm value and the difference between the two is greater than the temperature threshold, the model is used to calculate the angle of satellite offset required for the satellite's solar array temperature to return to below the temperature alarm value. The satellite offset method is used to change the angle of the sun's incident rays to lower the temperature of the satellite's solar array.

[0009] Furthermore, the solar panel temperature calculation model is as follows:

[0010] T={{[(Scosθ+10)×α+66×ε]×A+P 分流} / (εAσ)} 0.25

[0011] Where T is the temperature of the solar array in Kelvin; S is the solar constant; θ is the angle between sunlight and the normal to the solar array; α is the absorptivity of the solar array; ε is the emissivity of the solar array; A is the area of ​​the solar array; and σ = 5.67 * 10⁻⁶. -8 P 分流 The power shunt to the solar array.

[0012] Furthermore, the shunt power P of the body-mounted solar array 分流 Obtained in the following way:

[0013] P 分流 =P 全分流 +P i分流 P 全分流 =SUM(I1, I i-1 )×V 母线 P i分流 =I i ×Duty cycle×V 母线

[0014] Where i represents the n square arrays constituting the solar array that satisfy the condition SUM(I) i+1 I n )≤I 负载 ≤SUM(I i I n ), the index of the square matrix i≥1, SUM(I1, I i-1 SUM(I) is the sum of the currents of each of the square matrices 1, 2, ..., i-1. i+1 I n SUM(I) is the sum of the currents in square matrices i+1, i+2, ..., n.i I n Let I be the sum of the currents in square matrices i, i+1, ..., n. 负载 For the entire satellite load current, V 母线 For satellite bus voltage, I i Let I be the current in matrix i, and the duty cycle be 1 - (Ii)2 负载 -SUM(I i+1 I n )) / I i .

[0015] Furthermore, the P mentioned above 分流 To ensure the energy balance of the entire satellite during its operation in orbit.

[0016] Furthermore, the required offset angle for the satellite is θ, satisfying the relationship P = P0 cosθ, where P is the output power of the satellite's solar panels after attitude offset, and P0 is the output power of the satellite's solar panels when the sun is shining directly on it. The P condition satisfies the requirement that the battery pack discharge depth not exceed 30% and that the entire satellite maintains energy balance during on-orbit operation. P is calculated from P. 分流 This allows us to determine the temperature of the satellite's solar panels.

[0017] Furthermore, the energy balance of the entire satellite in orbit is conditional upon the satellite completing its payload mission and exiting the shadow zone, while in the sunlit area, the satellite's battery pack can be fully charged.

[0018] Furthermore, the depth of discharge of the battery pack is obtained in the following way:

[0019] Maximum depth of discharge = (Discharge during shadow period + Replenished discharge during sunshine period) / Battery pack capacity

[0020] Discharge during the shadow period = Long-term power consumption × Shadow period duration / (Final discharge voltage of a single battery cell × Number of cells in series)

[0021] Supplemental discharge capacity during sunshine period = ((long-term power consumption - solar array power / 2) × maneuver time + (long-term power consumption + load power) × load operating time) / (final discharge voltage of individual battery cell × number of cells in series),

[0022] Rechargeable capacity during sunshine period = (Solar array power - Long-term power consumption) × (Sunlight period duration - Mobility duration - Load operating duration) / (Final charging voltage of individual battery cells × Number of cells in series × Recharge factor)

[0023] Energy surplus = Rechargeable energy during the sunshine period - Discharged energy during the shadow period - Supplemented discharge energy during the sunshine period;

[0024] The long-term power consumption refers to the power continuously consumed by the satellite during its operation in orbit.

[0025] Preferably, the calculation error range is 5℃. The difference is 5℃.

[0026] Preferably, the method of increasing the load current is: using the satellite module as the load to consume power, and the use of the module as the load to consume power does not affect the normal operation of other products of the satellite.

[0027] The advantages of this invention compared to the prior art are:

[0028] (1) In the design of this invention, the influence of various factors such as satellite working mode, long-term load current, and orbital characteristics on the temperature of the solar array is comprehensively considered. The invention proposes a strategy to control the temperature of the solar array by adjusting the magnitude of external heat flow and the magnitude of shunt current. The reduction of external heat flow is mainly achieved by increasing the satellite attitude offset. The adjustment of the shunt current is achieved by increasing the load current and reducing the magnitude of the array current (using the solar array offset). The determination of the solar array offset angle and the magnitude of the load current are determined under the constraints of satellite on-orbit energy balance. Therefore, the method of this invention is a refined temperature control of the solar array under the condition of satisfying satellite energy balance, which not only ensures the normal on-orbit operation of the satellite, but also effectively avoids the high temperature problem of the solar array.

[0029] (2) This invention constructs a refined analysis method for solar wing temperature, which incorporates parameters such as solar wing area, absorptivity, emissivity, sunlight angle, shunting, solar constant, and Earth infrared and albedo corresponding to the orbit into the thermal control calculation model, providing an accurate thermal control model for on-orbit body-mounted solar wing temperature control, which is conducive to precise quantitative control of body-mounted solar wing temperature.

[0030] (3) The present invention adopts an on-orbit closed-loop solar array temperature control strategy, which has the advantages of fast response time and high efficiency compared with the ground-based telemetry and control operation method. It reduces the risk of excessive temperature of the solar array, and also reduces the dependence on ground-based telemetry and control resources, thereby improving the efficiency of long-term on-orbit management of the satellite. Attached Figure Description

[0031] Figure 1 This is a flowchart illustrating the overall design of the method of the present invention;

[0032] Figure 2 This is a flowchart of the flow splitting size calculation for the present invention;

[0033] Figure 3 This is a flowchart of the calculation of the offset angle after the solar array is offset in this invention;

[0034] Figure 4 This is a flowchart of the solar panel temperature calculation process of the present invention;

[0035] Figure 5 This is a flowchart illustrating the design of the on-orbit autonomous closed-loop temperature control strategy of this invention. Detailed Implementation

[0036] Given a fixed thermal control design for the solar array, the main factors affecting its temperature include the angle of sunlight incidence and the shunt power distribution to the solar array. During satellite operation, in shadowed areas, the satellite receives no external power and relies on its battery bank for power. In sunlit areas, the satellite establishes a solar array orientation to the sun, satisfying its basic energy needs while charging the battery bank; any excess energy is shunted. With the same angle of sunlight incidence, a larger shunt results in a higher solar array temperature; conversely, a larger angle of sunlight incidence (the angle between sunlight and the solar array normal) results in a lower solar array temperature. The angle of sunlight incidence also affects the total output power of the solar array and the magnitude of the shunt power distribution. When the satellite's solar array is aligned with the sun, its temperature needs to be assessed. If it exceeds a safe threshold, temperature control measures are required. Therefore, this invention, considering the satellite's energy requirements during operation, focuses on these two key factors to adjust and control the high-temperature value of the solar array.

[0037] To explain the technical content, objectives, and effects of the present invention in detail, the following description is provided in conjunction with the embodiments and accompanying drawings.

[0038] like Figure 1 As shown, the temperature control of the solar array is mainly achieved through the adjustment of external heat flux and the magnitude of the shunt. External heat flux adjustment is primarily controlled by adjusting the satellite offset angle. The magnitude of the shunt can be controlled by altering the output power based on the satellite offset angle and the satellite load. When adjusting the satellite offset angle, energy balance, satellite thermal control, and load-related influence analyses are required to determine the maximum acceptable offset angle. The shunt is calculated based on the offset angle and load. The offset angle, shunt, solar emissivity, and reflectivity are then input into the solar array temperature calculation model to calculate the temperature range of the solar array. Based on the measured solar array temperature in orbit, the solar array temperature calculation model is revised, and an on-orbit autonomous solar array temperature adjustment strategy is designed to achieve autonomous on-orbit solar array temperature control.

[0039] like Figure 2 As shown, to assess the solar array temperature at different stages of satellite operation in orbit, the shunt current to the solar array is first determined, which can be calculated using the load size and the output current of the solar array. First, based on the power consumption of the satellite in different operating modes throughout its lifespan, the corresponding shunt current under different operating conditions needs to be traversed. In the initial stage of orbit insertion, due to the high energy conversion efficiency of the solar array and sufficient energy per revolution, a large shunt current occurs, leading to excessively high solar array temperatures. Therefore, in the initial stage of orbit insertion, the long-term power consumption of the satellite platform can be increased by adjusting the threshold values ​​of some heating circuits or activating some equipment. With the satellite's output power remaining constant, increasing power consumption reduces the shunt current and lowers the solar array temperature.

[0040] To calculate the current shunt magnitude of the solar array, taking a non-regulated bus as an example, under a sequential current shunt regulation power distribution system, the current shunt state of each solar array is determined by the primary bus voltage and the magnitude of the array and load current. As the battery pack charges, the primary bus voltage gradually increases, and the battery pack transitions from constant current to constant voltage charging, with each array sequentially shunting current. The shunt begins at the corresponding primary bus voltage value for each array stage; when the primary bus voltage rises to that threshold, that array stage begins shunting.

[0041] Assume the corresponding arrays on the solar array are array 1, array 2, ..., array n. The corresponding array currents are I1, I2, ..., In. n The current distribution sequence is based on the array order, with current distributed level by level. The calculation method for different array current distribution or power supply states and current duty cycles when the battery pack is fully charged is as follows:

[0042] (1) Determine the value of the index i of the square matrix that satisfies the following formula.

[0043] SUM(I i+1 I n )≤I 负载 ≤SUM(I i I n ), (i≥1)

[0044] Among them, I 负载 This represents the total satellite load current. SUM(I) i+1 I n ) represents the array current I i+1 …I n Summation, SUM(I) i I n ) represents the array current I i …I n Sum.

[0045] (2) Arrays i+1, i+2, ..., n are in full power supply mode, the i-th array is in pulse width modulation mode (i.e., constantly switching between shunt and non-shunt states), and arrays 1, 2, and i-1 are in full shunt state. The shunt current is then the sum of the shunt currents of arrays 1 to i-1. The shunt power P of arrays 1 to i-1 is calculated based on the shunt current. 全分流 =SUM(I1, I i-1 )×V 母线 Where SUM(I1, I i-1 ) is the summation of the shunt currents, V 母线 This is the bus voltage.

[0046] (3) The i-th level matrix is ​​in pulse width modulation state, and its shunt state duty cycle = 1 - (I 负载 -SUM(I i+1 I n)) / I i .

[0047] P i分流 =I i ×Duty cycle×V 母线

[0048] Among them, P i分流 Let I be the shunt power of array i. i Let be the shunt current of matrix i.

[0049] (4) Calculate the total shunt power of the solar array as P 分流 =P 全分流 +P i分流 .

[0050] Considering that the current shunting varies under different operating conditions of the satellite in orbit, we can analyze the corresponding current shunting magnitude under various combined operating conditions based on the different load current requirements corresponding to different load operating modes in orbit and the current output magnitude of the array in different seasons.

[0051] like Figure 3 As shown, when the solar array has sufficient energy leading to a high temperature, the satellite's offset angle is calculated based on the constraints of on-orbit energy balance. Combined with constraints such as the illumination angle of the satellite's thermal control and heat dissipation surfaces and the payload's illumination angle, the satellite is offset at a certain angle to change the solar array's incident angle, reducing the solar array's power generation and external heat flow, thereby lowering the solar array's temperature. The specific steps are as follows:

[0052] Calculate the power output of the solar array at a certain tilt angle to analyze the satellite's energy balance and battery discharge depth. The solar array output power is approximately linearly related to the angle of sunlight incidence. After tilting, the angle between sunlight and the normal to the front of the solar array is generally less than 45°. The theoretical and actual values ​​of the solar array output power deviate very little. The power output of the solar array after tilting is P = P0cosθ. Where P is the output power of the solar array after tilting, P0 is the output power of the solar array when sunlight is directly overhead, and θ is the angle between sunlight and the normal to the front of the solar array after tilting.

[0053] Calculate the energy balance and battery discharge depth, solar panel output power P, and recalculate the battery discharge depth and energy balance under different operating modes. This should ensure that, after biasing, the battery discharge depth (generally no more than 30%) and energy balance (meaning that the batteries in the sunlit area can be fully charged after the satellite completes its payload mission and is in the shadow area) meet the overall on-orbit operational requirements. Calculate using the following formula:

[0054] 1) Rechargeable capacity during sunshine period = (Solar array power - Long-term power consumption) × (Sunlight period duration - Mobility duration - Load working duration) / (Final charging voltage of individual battery cells × Number of cells in series × Recharge factor);

[0055] 2) Discharge capacity during the shadow period = Long-term power consumption × Shadow period duration / (Final discharge voltage of a single battery cell × Number of cells in series);

[0056] 3) Supplemental discharge capacity during sunshine period = ((long-term power consumption - solar array power / 2) × maneuver time + (long-term power consumption + load power) × load working time) / (final discharge voltage of individual battery cell × number of cells in series);

[0057] 4) Energy surplus = Rechargeable energy during the sunshine period - Discharged energy during the shadow period - Supplemented discharge energy during the sunshine period;

[0058] 5) Maximum depth of discharge = (discharge during shadow period + supplementary discharge during sunshine period) / battery pack capacity.

[0059] Long-term power consumption refers to the power continuously consumed by a satellite during its operation in orbit.

[0060] In the above calculation formula, the power of the solar cell array is closely related to the offset angle θ. The larger θ is, the smaller the power of the solar cell array. By analyzing the satellite energy balance under different offset angles θ, the size of the offset angle θ can be determined.

[0061] like Figure 4 As shown, during satellite operation, since the back of the body-mounted wing is the satellite body, to reduce the impact of the high temperature of the solar panels on the temperature of the internal equipment, multi-layer heat insulation components are used in some areas. A 15-unit double-sided aluminized polyester film is selected, and 5mm thick fiberglass or titanium alloy heat insulation pads are used at the joints with adjacent side panels and partitions for heat insulation connection. Due to poor heat dissipation, the temperature of the body-mounted solar wings becomes excessively high when exposed to sunlight in sunlit areas.

[0062] The calculation method for analyzing the impact of satellite offset angle and shunt magnitude on solar array temperature is as follows:

[0063] First, determine the known input information:

[0064] 1) The area of ​​the solar array is A;

[0065] 2) The absorptivity of the solar array is α (generally taken as 0.666 in the early stage and 0.8 to 0.85 in the later stage; it is recommended to take a smaller value for larger shunting).

[0066] 3) Emissivity ε (recommended value: 0.82);

[0067] 4) Figure 3 The calculated angle between sunlight and the normal to the solar wing is θ (0°~90°);

[0068] 5) Figure 2 Calculated split P 分流 ;

[0069] 6) Solar constant S (varies with the seasons, 1322~1414W / m) 2 )

[0070] Next, analyze the solar panel temperature using the following steps:

[0071] (1) Analyze the external heat flux of the solar array under different solar incidence angles. The external heat flux consists of three parts: absorbed direct solar heat flux, Earth infrared radiation, and Earth albedo. The direct solar heat flux is mainly related to the solar array area, solar constant, and solar incidence angle, and is calculated according to the formula S×cosθ×A. Earth infrared radiation and Earth albedo are related to the orbital altitude. Taking a 500km sun-synchronous orbit as an example, when the solar incidence angle is 0°, the Earth albedo is approximately 10W / m. 2 The Earth's infrared radiation is approximately 66 W / m. 2 .

[0072] Considering the effects of orbital altitude on infrared and albedo, the Earth's albedo is 10 × [6878 / (6378+h)] 2 The Earth's infrared value is 66 × [6878 / (6378+h)] 2 h represents the altitude of the satellite's orbit.

[0073] (2) Calculate the total absorbed external heat flow, which is related to the direct solar heat flow, Earth infrared, Earth albedo, solar wing absorptivity and emissivity. Specifically, calculate according to the following formula, where different absorptivity coefficients are used in the initial and final stages.

[0074] Total absorbed external heat flow = [(S×cosθ+10)×α+66×ε]×A

[0075] (3) Calculate the total absorbed heat, which is related to the solar array's absorptivity α and emissivity ε, and conforms to the following calculation formula.

[0076] Total heat absorbed = [(Scosθ + 10) × α + 66 × ε] × A + P 分流 =εAσT 4 Where σ = 5.67 * 10 -8 T is the temperature of the body-mounted solar array, measured in Kelvin (K). P 分流 This refers to the shunt power of the solar array.

[0077] Therefore, we can calculate T = {{[(Scosθ+10)×α+66×ε]×A+P} 分流} / (εAσ)} 0.25

[0078] If T is converted to °C as T1, then T1 = T - 273.15.

[0079] The temperature value is calculated based on the solar array offset angle, shunt power, inherent solar array design, and orbital influences. This temperature can be compared with the measured temperature in orbit to optimize parameters such as external heat flux, Earth infrared radiation, solar constant, absorptivity, and emissivity, thus completing the model revision.

[0080] like Figure 5 As shown, the satellite utilizes on-orbit computer software to execute an autonomous temperature control strategy for its solar panels. This strategy is designed with two states: enabled and disabled. In the enabled state, it can autonomously determine telemetry data such as solar panel temperature and shunt status, and execute the control strategy accordingly. In the disabled state, it does not autonomously execute the control strategy. The enabled or disabled state can be set via commands based on fluctuations in the satellite's operating temperature.

[0081] When enabled, the spaceborne computer continuously monitors the measured temperature of the solar array and, based on... Figure 4 The temperature values ​​obtained from the solar array temperature calculation method are compared. During the initial orbital period, it is determined whether the measured and calculated temperatures are consistent. If the deviation is significant, the solar array temperature calculation model is installed on the on-orbit correction module to ensure that the measured and calculated temperatures are basically consistent (≤5℃). Then, it is determined whether the solar array temperature is close to the required temperature threshold. If the temperature exceeds the threshold, it is first determined whether the array current is greater than the expected value (the current when the solar array is under normal irradiation), and then... Figure 2 The method used in the analysis determines the magnitude of the shunting. Based on these two conditions, it can be determined that the temperature rise of the solar array is due to the large external heat flow and shunting of the solar array, which are influenced by sufficient sunlight.

[0082] If the solar array temperature exceeds the required threshold by no more than 5°C, the priority is to increase the load current to reduce the shunt current and achieve the temperature reduction target. The increased load current employs a satellite panel power-consuming load design, which can be activated when needed without affecting the normal operation of other satellite components. The required shunt power reduction to lower the solar array temperature to the required threshold is determined based on the solar array temperature calculation model. Figure 2 The shunt power calculation method determines the required increase in load current. After determining the load current, it is then based on... Figure 3 The energy balance calculation method was used to determine that increasing the load current would still maintain energy balance. The increased load current requirement was then used to open the hatch, consuming power.

[0083] When the solar array temperature exceeds the required threshold by more than 5°C, a satellite offset method is used to reduce the solar array temperature. The minimum offset angle for reducing the solar array temperature to the required threshold is determined based on the solar array temperature calculation model. First, an initial offset angle is preset in the calculation model, based on... Figure 3 The energy balance calculation method determines that the energy balance can still be guaranteed at this bias angle; secondly, based on the array current after the bias angle, according to... Figure 2 Calculation method analysis of shunt power P 分流 Finally, based on the initial offset angle, it is analyzed whether the solar array temperature value meets the required threshold. If it does not meet the threshold, the offset angle is increased and the above analysis process is repeated until the offset angle corresponding to the solar array temperature meeting the required threshold value is determined.

[0084] The satellite's computer sets an autonomous offset enable flag and an autonomous offset flag bit, which defaults to "0 - no offset". When the enable flag is enabled and the offset flag bit is "0 - no offset", the satellite autonomously performs an attitude offset maneuver based on the determined attitude offset angle. After offsetting, it monitors the solar array temperature changes in a closed loop. If the solar array temperature meets the requirements, it continues to monitor the power supply and energy balance of the solar array. If energy shortages occur due to seasonal changes or load demand fluctuations, the solar array offset angle is recalculated, and the autonomous offset operation is performed again.

[0085] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for on-orbit autonomous temperature control of a satellite-mounted solar array, characterized in that... include: A solar array temperature calculation model is established to calculate the theoretical value of the solar array temperature on the satellite body. At the same time, the measured value of the solar array temperature on the satellite body is acquired in real time. The measured value is used to correct the model so that the error between the two is within the allowable calculation error range. The temperature of the satellite's solar array is monitored in real time and compared with the temperature alarm value. When the measured temperature exceeds the temperature alarm value and the difference between the two is not greater than the temperature threshold, the model is used to calculate the amount of current required for the satellite's solar array temperature to return to below the temperature alarm value. The current is then reduced by increasing the load current to lower the temperature of the satellite's solar array. When the measured temperature exceeds the temperature alarm value and the difference between the two is greater than the temperature threshold, the model is used to calculate the angle of satellite offset required for the satellite's solar array temperature to return to below the temperature alarm value. The satellite offset method is used to change the angle of the sun's incident rays to lower the temperature of the satellite's solar array.

2. The satellite-mounted solar array on-orbit autonomous temperature control method according to claim 1, characterized in that: The solar panel temperature calculation model is as follows: T={ {[(Scosθ+10) ×α+66×ε] ×A+P 分流 } / (εAσ)} 0.25 Where T is the temperature of the solar array in Kelvin; S is the solar constant; θ is the angle between sunlight and the normal to the solar array; α is the absorptivity of the solar array; ε is the emissivity of the solar array; A is the area of ​​the solar array; and σ = 5.

67. 10 -8 P 分流 The power shunt to the solar array.

3. The on-orbit autonomous temperature control method for satellite-mounted solar arrays according to claim 2, characterized in that: The shunt power P of the body-mounted solar array 分流 Obtained in the following way: P 分流 =P 全分流 +P i分流 P 全分流 =SUM(I1, I) i-1 )×V 母线 P i分流 =I i ×Duty cycle×V 母线 Where i represents the n square arrays constituting the solar array that satisfy the condition SUM(I) i+1 I n ) ≤ I 负载 ≤ SUM (I i I n SUM(i1, i2, ..., i3) represents the index of the square matrix i ≥ 1. i-1 SUM(I) is the sum of the currents in each of the square matrices 1, 2, ..., i-1. i+1 I n SUM(I) is the sum of the currents in square matrices i+1, i+2, ..., n. i I n Let I be the sum of the currents in square matrices i, i+1, ..., n. 负载 For the entire satellite load current, V 母线 For satellite bus voltage, I i Let i be the current in the matrix, and the duty cycle be 1 - (Ii). 负载 - SUM (I i+1 I n )) / I i .

4. The satellite-mounted solar array on-orbit autonomous temperature control method according to claim 3, characterized in that: The P mentioned 分流 To ensure the energy balance of the entire satellite during its operation in orbit.

5. A satellite-mounted solar array on-orbit autonomous temperature control method according to claim 2, characterized in that: The required offset angle for the satellite is θ, satisfying the relationship P = P0cosθ, where P is the output power of the satellite's solar panels after attitude offset, and P0 is the output power of the satellite's solar panels when the sun is shining directly on it. The P condition satisfies the requirement that the battery pack discharge depth not exceed 30% and that the entire satellite maintains energy balance during on-orbit operation. P is calculated from P. 分流 This allows us to determine the temperature of the satellite's solar panels.

6. A satellite-mounted solar array on-orbit autonomous temperature control method according to claim 4 or 5, characterized in that: The aforementioned energy balance for the entire satellite's on-orbit operation is conditional upon the satellite completing its payload mission and exiting the shadow zone, while its battery pack can be fully charged in the sunlit area.

7. A satellite-mounted solar array on-orbit autonomous temperature control method according to claim 5, characterized in that: The depth of discharge of the battery pack is obtained in the following way: Maximum depth of discharge = (Discharge during shadow period + Replenished discharge during sunshine period) / Battery pack capacity Discharge during the shadow period = Long-term power consumption × Shadow period duration / (Final discharge voltage of a single battery cell × Number of cells in series). Replenishment discharge capacity during sunshine period = ((long-term power consumption - solar cell array power / 2) × maneuver time + (long-term power consumption + load power) × load working time) / (final discharge voltage of individual battery cell × number of cells in series). Rechargeable capacity during sunshine period = (Solar array power - Long-term power consumption) × (Sunlight period duration - Mobility duration - Load working duration) / (Final charging voltage of individual battery cells × Number of cells in series × Recharge factor). Energy surplus = Rechargeable energy during the sunshine period - Discharged energy during the shadow period - Supplemented discharge energy during the sunshine period; The long-term power consumption refers to the power continuously consumed by the satellite during its operation in orbit.

8. The on-orbit autonomous temperature control method for satellite-mounted solar panels according to claim 1, characterized in that: The calculation error range is 5℃.

9. A satellite-mounted solar array on-orbit autonomous temperature control method according to claim 1, characterized in that: The temperature threshold is 5°C.

10. A satellite-mounted solar array on-orbit autonomous temperature control method according to claim 1, characterized in that: The method of increasing the load current is as follows: the satellite module is used as the load to consume power, and the use of the module as the load to consume power does not affect the normal operation of other products of the satellite.

Images