Application of Medium-Voltage Fully-Regulated High-Power EPS on the BDS-3 Satellites

LI Xuping,LEI Hu,XU Huidong,LUO Guo

Shanghai Institute of Space Power-sources,Shanghai 200245

Abstract:China's BeiDou Navigation Satellite System (BDS) construction has been completed and the system has been formally commissioned.Most of the Electric Power Systems (EPSs) for MEO satellites were developed by the Shanghai Institute of Space Power-sources.The 42 V medium-voltage fully-regulated high-power EPS has been adopted for the first time in medium Earth orbit,with an output power reaching about 3 kW.Compared with the 42 V medium-voltage semi-regulated bus power system used in the Regional Navigation BDS-2 satellite,the EPS of the BDS-3 MEO satellites has increased power by about 80%,adopting many newly developed products such as high-efficient triple junction GaAs solar cells,high-energy-density lithium ion batteries and a high-efficient autonomous power control unit (PCU).Based on the studies on the medium-voltage fully-regulated and high-power EPS technical principles,and the adaptability and reliability of various working modes,the test verifications for the EPS were conducted both on the ground and in orbit.Compared with other global navigation satellite systems such as GPS,Galileo and GLONASS,the EPS of the BDS-3 MEO satellite has a long design life time which is equivalent to that of the GPS and Galileo,but with a larger power supply capability and power ratio,distinguishing its advancement in the field of satellite power technology.

Key words:BDS-3 satellite,EPS,medium-voltage,fully-regulated bus

1 INTRODUCTION

In June 2020,the construction of China's BeiDou 3 Navigation Satellite System (BDS-3) was completed.BDS-3 is one of the four major global navigation satellite systems in the world.China's Global Navigation Satellites System (GNSS) consists of 24 MEO satellites,3 IGSO satellites and 3 GEO satellites[1].The Electric Power Systems (EPSs) of MEO satellites were mostly developed by the Shanghai Institute of Space Power-sources(SISP).

As a new-generation navigation satellite,the BDS-3 MEO satellite has a higher demand for power from the EPS compared with that of BDS-2,due to higher output power and longer design life.A medium-voltage fully-regulated bus and an upper stage are used for entering orbit.Meanwhile,high power ratio devices,such as high-efficient triple junction GaAs solar cells,high-energy-density lithium ion batteries and a high-efficient autonomous power control unit (PCU) are adopted to meet the higher power demand,enabling the satellite to have the ability to operate autonomously for more than 60 days[1-2].See Table 1 for the characteristics of the EPS of the BeiDou MEO satellite.

2 TECHNICAL SOLUTION

2.1 EPS Composition and Functions

The long-term power load of a BDS-3 MEO satellite is 2.8 kW,which is more than 80% higher than that of a BDS-2 MEO satellite.The EPS of a BDS-3 MEO satellite adopts a single bus system at a 42 V medium fully-regulated voltage with sequence switch shunt regulation(S3R) control topology,to maximize its efficiency and improve quality.

The EPS consists of a solar array (SA),lithium ion batteries,a PCU,a battery management unit (BMU) and a battery connection relay box (BCRB),as shown in Figure 1.

The functions of each product are as follows[4]:

1) Solar array

The function of the solar array is to convert solar energy into electrical power when the satellite is in sunlight.Through PCU regulation,part of the electricity is provided to satisfy the load of a satellite,and the other part is provided to the lithium ion batteries for charging and storing energy.The remaining part is lost in the form of heat in space.

Two symmetrical solar wings are adopted.Each solar wing is composed of 3 solar panels (rigid substrate).Average triple junction GaAs solar cells are selected with a photoelectric conversion efficiency of not less than 28%.The solar array is modular in design,with a total of 24 subarrays,12 subarrays for each wing,and the same series number cells for each array.The maximum output power at the initial life of the solar array in orbit is more than 4 kW (on the 42 V bus),and the minimumoutput power at the end of life will be more than 3.3 kW (on the 42 V bus).

Figure 1 EPS composition

Table 1 The characteristics of the EPS of BDS MEO satellites

2) Lithium ion battery

Lithium ion batteries are used to supply power to satisfy the electrical load of the satellite via the regulation function of the PCU to ensure sufficient power for the upper stage and for operation during its orbit while in eclipse .

2 sets of lithium ion batteries with a rated total capacity of 150 A·h are adopted.Each set of batteries is equipped with bypass switches and interfaces for power,ignition,voltage sampling,cell balance,bypass driver,heater,etc.

3) Power control unit

The PCU is used to regulate the power of the solar array during the sunlight period and charge the lithium ion batteries,as well as to regulate the power discharge of the batteries during the eclipse period to ensure a stable bus voltage,and provide remote measurement,remote control along with a 1553B bus interface for the EPS.The PCU contains power management software,and a hardware reinforced design for anti-single event effects along with an independent watchdog and software functions including check code,three-mode redundancy,a software trap and other features.

The EPS uses 1 PCU,including 24 hot redundancy S3R circuits with frequency limited,4 hot redundancy charging control circuits of Superbuck type,6 hot redundancy discharge regulation circuits of Weinberg Boost type,a triple redundant Main Error Amplifying (MEA) circuit,TM/TC circuit with primary backup redundancy and other functional circuits.The output capacity of the PCU bus is greater than 3.6 kW.The PCU adopts a modular structure,which is made of 12 structural blocks connected in series.The two wings are symmetrically arranged on the left and right sides,which is good for heat dissipation.The capacity of each SA shunt circuit is greater than 7 A,with shunt efficiency greater than 98%,and the transient discharge current caused by the junction capacitance of solar cells is limited within 15 A.Battery charging capacity is greater than 32 A,and charging conversion efficiency is greater than 94%.The output power of each battery discharge circuit is not less than 700 W,and the discharge conversion efficiency is more than 94%.The difference of discharge current sharing between 6 circuits is not more than 1%,and the bus over voltage protection is limited to not more than 45 V to avoid damaging other equipments.

4) Battery management unit

The function of the BMU is to realize the voltage balance control of each lithium-ion battery cell to ensure that the voltage difference between cells is within the allowed range.

One BMU is adopted in the EPS,which includes measurement and control circuits for 2 sets of battery cell voltage and block voltage sampling and transformation,balance control,bypass control,TM/TC control and serial bus control,etc.Each functional circuit is set with a cold backup redundancy.The voltage sampling accuracy of the BMU is less than 10 mV within the full temperature range,which ensures the uniformity of the battery voltage.

5) Battery connection relay box

The function of BCRB is to disconnect the battery charging and discharging power path directly during AIT.

One BCRB is adopted in the EPS to control the on-off of the battery power path through a high-power relay.The BCRB has an anti-surge current-limiting circuit when the battery is connected and has a charging function when the battery discharge path is disconnected,which ensures the safety of the satellite.

2.2 EPS Design [4]

The principle of the EPS is shown in Figure 2.In sunlight,the solar array is regulated by 24 S3R circuits,which are controlled by MEA.The S3R circuits realize the shunt of excess solar power,and the main bus is controlled within the range of 42.2 V±0.2 V through the switch regulation of one shunt circuit any time.At the same time,at the beginning of the sunlight period,the bus power supply will charge the batteries via the battery charging regulator (BCR) circuit with constant current control first,and then convert to constant voltage control when batteries reach the end of voltage (EOV).The constant current and constant voltage charging value is set by the power management software.The BCR module can work in hot backup or cool backup.In sunlight,if during the charging process the satellite power load increases,such that the bus power cannot meet the requirements,the battery error amplifying (BEA) control circuit will automatically reduce the charging current and stabilize the main bus voltage,so as to give priority to the satellite load power requirements.

In eclipse,the main bus voltage will be regulated within the range of 42.2 V±0.2 V by the 6 battery discharge regulator(BDR) modules from two batteries.In the process of eclipse entry,when the solar array power gradually decreases and cannot meet the power load requirement of the satellite,the MEA signal will be turned off for charging first,and the discharge regulation will be turned on for joint combined power supply until the eclipse is entered.As the satellite comes out of the eclipse,the solar array power increases gradually,and all S3R circuits develop full output,the output current of the BDR circuits gradually decrease,and the combined power supply is continued until it enters the sunlight.The BDRs will be turned off and the battery will be charged automatically under the MEA signal.No matter how the satellite state and load characteristics change within the normal range in orbit,the MEA signal is used for unified control of shunt,charge and discharge functions,and the EPS will maintain a stable output of 42.2 V±0.2 V as the bus power supply.

The large-capacity lithium ion battery for a MEO satellite has characteristics such as more charge-discharge cycles,longer shelf life,high state of charge (SOC),and long storage life.For the first time,in-orbit autonomous control and management strategies have been applied in the EPS design,which include long storage control in sunlight,treatment before entering into eclipse,charge-discharge cycle control in eclipse,bypass judgment and processing,batteries or cell overcharge and over discharge protection control,including the function of automatic open balance control to ensure the cell voltage uniformity in maintained within 10 mV.In addition,EPS selects different constant charging voltages to keep the battery voltage within the appropriate SOC.

All functions of the EPS are autonomously controlled,with the ability to operate autonomously for more than 60 days.

2.2.1 System stability design

According to the principles of circuit stability,the output filter capacitance of the main bus is a decisive factor for the frequency band response width and output impedance in the system control circuit.The transfer function block diagram of the EPS is shown in Figure 3.

Figure 2 Block diagram of working principle of the EPS

Figure 3 Transfer function block diagram of the EPS

Mathematical expression of frequency band response:

Mathematical expression of bus output impedance:

K.Feedback coefficient

VrefReference signal

A(S) MEA gain

G(S) Power regulating system

CbusFilter capacitor

Under the conditions of charging and discharging,the phase margin of the EPS is greater than 60°,and gain margin of the EPS is more than 10 dB that meets the requirements for system stability.At the same time,the output impedance in the case of charging and discharging operations can be less than 50 mΩ thus guarantee the EPS output characteristic.

2.2.2 MEA control design

In the EPS,operations of S3R,BCR and BDR circuits are uniformly controlled by the MEA.MEA has a set of three domains of S3R,BCR and BDR according to the voltage,which is shown in Figure 4.The MEA circuit is set with S3R-MEA and BCDR-MEA.S3R-MEA is further amplified S3R domain on the basis of BCDR-MEA,that is beneficial to the 24 S3R circuits control.All MEA circuits adopt a three-by-two voting to ensure reliability.

The MEA circuit works on the principle that the DC openloop output voltage gain of the operational amplifier is infinite,and the AC output gain varies as a logarithmic form according to the frequency domain of the AC input.The mathematical expression of the relation of the MEA output voltage VMEAand bus voltageVbusis generally as follows[5]:

Figure 4 Diagram of three-domain control voltage

KPMEA voltage amplification proportional coefficient

KiMEA voltage amplification integration coefficient

kBus voltage feedback partial voltage proportional coefficient

VbusBus voltage

VrefReference voltage

The bus voltage and reference voltage samples of the MEA circuit are obtained by proportional voltage dividers from the bus voltage,thus minimizing the in-domain error caused by the drift of temperature coefficient.Meanwhile,to ensure reliable operation of the EPS,the MEA voltage values of S3R,BCR and BDR for different operations remain constant.The gap between S3R&BCR and BCR&BDR shall be more than 0.5 V.

2.2.3 BEA control design

The unified battery error amplification (BEA) circuit is used for charging constant voltage,constant current and bus voltage stability control.Each battery group corresponds to a BEA circuit,which can realize the functions of constant voltage,constant current and pricing a stable bus under the control of the BCRs together.Each BEA circuit also adopts the redundancy mode of three-by-two voting to ensure the reliability of the functional circuit.The principle of BEA circuit is shown in Figure 5.

Figure 5 Block diagram of working principle of BEA

2.3 Operational Mode of EPS

The complete mission of the EPS includes the launch active period,the upper stage entry period and the normal operating period in orbit.The operational modes of the EPS mainly include non-power supply,joint combined power supply,power supply charging in sunlight,power supply in eclipse and battery cell balancing.The conversion relationship and conversion conditions among all the operational modes are shown in Figure 6.

2.3.1 Upper stage orbit insertion period

The BDS-3 MEO satellite adopts the mode of orbit insertion directly from the upper stage with dual satellites,which is different from that of other satellites.During orbit insertion of the upper stage,the dual satellites are powered by the upper stage.The satellites are powered by their EPSs automatically up to separation from the upper stage.

1) Power supply by the upper stage:during the orbit injection,the upper stage spins around the axis of the paralleled satellites (Zaxis) at a certain angular velocity,as shown in Figure 7.The solar wings are folded up,where only two outside solar panels of each satellite can be irradiated with sunlight to generate power (and only one panel of each satellite can be irradiated at any time),the cosine change rule for the incidence angle of sunlight is cosθ=cosωsinβ(ω=0 °-360 °,βis the angle between incident direction of sunlight and flight direction of the upper stage),the solar array output current is indicated in Figure 8.As the solar wings are not spread out and hence cannot dissipate heat,the upper stage power supply causes the solar array output current into a full shunt state,so the solar cells have to withstand high temperature of up to 125 ℃.

Figure 6 The conversion relation and conversion conditions of EPS operational modes

Figure 7 The state of the satellite in the upper stage

2) Power supply transition from the upper stage to EPS.During the orbit insertion of the upper stage,the upper stage provides a voltage of about 45 V to the two satellites,which is higher than the satellite main bus of 42 V supplied by EPS.At the moment of separation,the EPS of each satellite will automatically and quickly supply power under the necessary condition so that all loads are not affected.The instantaneous descent recovery time of main bus voltage is less than 5 ms,the lowest main bus voltage will be higher than 39 V,and the main bus power supply will normalize,as shown in Figure 9.

Figure 9 Power supply conversion characteristics (blue line for bus voltage)

2.3.2 In-orbit management of lithium ion batteries

Each year,MEO satellites go through about 100 days of eclipse period,with a maximum discharge time approaching 1 hour.In eclipse season,the battery is charged and discharged twice a day,and for the rest time of a year kept in long-term storage.Therefore,the life of the lithium ion batteries is directly determined by the storage life and cycle life,thus special control strategies need to be adopted in orbit.

1) Long-term sunlight storage period:the batteries are stored in a 60%-80% low state of charge (SOC),and the storage temperature is controlled between-5-10℃.In orbit,due to the electrical loss of the circuits connecting with the batteries and cells,the reduction of the battery SOC from 80% to 60% will last for several days.When the battery SOC drops to 60%,the EPS will automatically start charging to 80% SOC state.In the long-term sunlight storage period,if the voltage difference between individual cells exceeds 60 mV,the balancing function will be automatically turned on until the voltage difference of the corresponding individual cell is less than 10 mV.

2) Charging-discharging cycle control in eclipse:the charging and discharging temperature of batteries is controlled within the range of 10-30℃.

3) Real-time charge and discharge protection control:when the voltage of batteries is more than 35 V,the overcharge protection function will be initiated,and the overcharge protection threshold is set to increase byΔV based on the current voltage set,and the battery cell overcharge protection threshold is set.As overdischarge will cause permanent failure of the batteries,when the voltage of batteries is lower than 22.5 V and the voltage of more than 3 individual cells is lower than 2.5 V (failure is not considered),the EPS will automatically start overdischarge protection and cut off the discharge path.When the battery voltage returns to a certain voltage,such as 32 V,the discharge path will be automatically switched on to continue to supply power for the satellite.All protection values can be modified in orbit.

4) Cell health check:considering the attenuation failure of individual cell performance in batteries,if the voltage difference between the fault cell and the average cell voltage exceeds more than the set value (the value can be modified in orbit),the EPS will isolate and interrupt it.If an individual cell voltage is higher than 4.3 V for 3 times in the charging process or lower than 3 V for 3 times in the discharging process,and the voltage difference between it and other cells reaches the set value,the corresponding battery cell's bypass shall be activated under the premise of excluding the fault of the cell sampling circuit.After the bypass is activated,some battery related values should be adjusted,such as the charging voltage and storage voltage.

2.4 System Reliability Design and Verification [6]

The EPS is designed to last for 12 years.According to the calculation of the reliability model of the EPS,the reliability of the EPS is expected to be 0.9987 before the normal operating period in orbit and 0.9885 in normal operational orbit.

2.4.1 Reliability verification at sub-system level

In order to ensure the reliability of the satellite,the EPS uses actual devices and equipment to carry out the sub-system level reliability verification,simulating battery cell bypass operation and isolation,and fuse fusing faults due to load short-circuits on the satellite,as well as bus response to a pulsating load,etc.The EPS is proved reliable and could meet the reliability design requirements.

1) Verification of the battery cell bypass operation and isolation

In the case of battery minimum SOC (about 2.8 V/cell),a domestically-made bypass is derived and operated successfully.Performances such as operation time and fuse time all meet the requirements,and there is no influence on the main bus and the batteries during the operation,as shown in Figure 10.

2) Verification a fuse fusing fault in load short-circuit

The EPS was set to full output power condition to the satellite in eclipse,which reduced the EPS output power margin to the lowest state,with the maximum the load 24.5 A fuse was short circuited on main bus in an instant,the dynamic response recovery time of both the bus voltage and current was within 300 μs,and minimum bus voltage still exceeded 34 V,and bus current was higher than 300 A at that instant.It did not affect the normal operation of the bus power supply or the load operation,as shown in Figure 11.

Figure 10 Verification of bypass operation

Figure 11 Waveform of bus voltage and current response

3) Verification of bus response with pulsating load

When the satellite is in orbit,it may have a bus load of about 800 W which is effected by a 50 Hz frequency pulse.In the worst state,the operational mode of the EPS is continuously transformed between the shunt domain and the discharge domain.The bus responded timely and operated steadily,meeting the requirements,as shown in Figure 12.

Figure 12 Response waveform of bus under 800 W pulsating load

2.4.2 Reliability verification of key devices

Large capacity lithium ion batteries and the high efficient autonomous PCUs are the key products of BDS-3 MEO satellites.In order to verify the design of long life and high reliability,many verification tests on the ground were specially carried out during the development process.

The PCU was subjected to five tests and passed the requirements on the ground,including backup redundancy design function verification test,system-level performance verification test,safety protection function verification test,12-year accelerated life test,and bias and diagnostic test,which verified the power supply reliability and safety of the PCU.

Similarly,the Lithium ion battery was subjected to and passed seven tests on the ground,including long term shelf survival verification test,charge-discharge cycle verification test,charging ratio test,open circuit fault verification test,short circuit fault verification test,influence of the cell difference on life test,cell and battery accelerated life tests,and verification of the lithium ion battery 12-year life design and the effectiveness of the management strategy in orbit .

2.4.3 Reliability verification in orbit[7]

Two BDS-3 MEO test satellites were launched in July 2015 and have been in operation for more than 5 years.The EPSs operate stably in orbit,and the output power of solar array,battery capacity and battery cell voltage difference all meet the design expectations,as shown in Figure 13 and Figure 14.

Figure 13 Solar array current curves in orbit

3 COMPARISON OF GLOBAL NAVIGATION SATELLITE POWER SYSTEMS

BDS-3 MEO satellites were compared with GPS satellites(GPS BLOCK IIF)[8,9,13,14],Galileo satellites,Glonass satellites(Glonass-M and Glonass-K)[16,17],while the EPSs of BDS-3 MEO satellites have the same design life as those of GPS and Galileo,and the key power products onboard BDS-3 MEO satellites are the same as the Galileo's,the output power supply of the BDS-3 satellite is the largest of the four satellite systems,as shown in Table 2.

4 CONCLUSION

The EPS of the BDS-3 MEO satellite uses many newly developed devices such as high efficient triple junction GaAs solar cells,high-energy-density and large capacity lithium ion battery and high efficient autonomous PCU.Medium-voltage fully regulated high-power EPSs have been applied successfully on BDS-3 MEO satellites.Compared with other global navigation satellite systems,the EPSs of BDS-3 MEO satellites provide the maximum output power.

Figure 14 Battery voltage curve in sunlight and eclipse

Table 2 Comparison of global navigation satellite power systems