SPACECRAFT POWER SUPPLY SYSTEM CONTROL METHOD Russian patent published in 2019 - IPC H02J7/35 

FIELD: electrical engineering.

SUBSTANCE: method of controlling the power supply system of a spacecraft containing regulation and control equipment (RCE) connected in parallel between themselves, photoelectric (PE) battery sectioned into m parts, n (n ≥ 1) rechargeable batteries (RB), where m = k ⋅ n (k ≥ 2), consists in cycling of RB in the charge-discharge mode, controlling of the EPS parameters, for example, output voltage, PE current, RB charge and discharge currents, current capacity of each battery; forming a control signal in the onboard control complex of the spacecraft to turn off part of the onboard equipment (OE) during RB emergency discharge to the minimum charge level; disconnection of all RB from the load while reducing the EPS output voltage to the specified minimum value; switching on all RBs to the load after charging all RBs to the specified level of charge. As the EPS the regulator of excess power (REP) PE is used. Resetting PE excess power, depending on the current values of the EPS parameters is performed by introducing/deactivating with the REP short circuit mode of the required number of PE sections, while in the shadow portion of the spacecraft orbit, the output voltage of the EPS is maintained within the range of variation corresponding to the range of variation of voltages n RB in the discharge mode; and in the EPS composition RB with flat volt-ampere characteristics are used, for example, lithium-ion RB. On the light section of the spacecraft orbit, the output EPS voltage is stabilized relatively to a constant reference voltage, value of the reference voltage is taken equal to the maximum allowable value of the EPS output voltage and close to the maximum charging RB voltage. Signal proportional to the difference between the output EPS voltage and the reference voltage, equal to AU, used to force changes and obtain the necessary ratio between the PE power, used to power the OE and the RB charge, and the PE power, dissipated by the REP, so that AU tends to zero; moreover, rechargeable batteries are charged with a current proportional to the voltage difference on the EPS output bus and this RB, while the amount (n) and type of RB are selected based on the PE parameters and the EPS energy balance on the following criteria: k⋅(I_PE)_max. ≤ (I_ch)_max., (I_heat)_max. ≤ n⋅(I_time)_max., where (I_PE)_max. is the maximum current of one PE section, (I_ch)_max. = c₁⋅Q_{No. RB}, (I_time)_max. = c₂ Q_{No. RB}, c₁ and c₂ – coefficients, characterizing the load capacity of each RB by charge and discharge currents, respectively, Q_{No. RB} is the rated electrical capacity of one RB, (I_heat)_max. is the maximum current consumed by the OE from the RB, k is the number of PE sections connected to one RB.

EFFECT: technical result is higher survivability and reliability of the autonomous power supply system (EPS) of spacecraft (SC).

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