Battery power requirements and related key components

Battery power requirements and related key components

For the battery requirements of satellites operating in different orbits, the focus will be on related components that require reliability and long-lasting electrical performance. The battery performance specifications of low earth orbit and geosynchronous earth orbit (GEO) satellites are very different. The typical orbital parameters of the first-generation rotating satellite are shown in Figure 1. The onboard battery is designed to meet the power load requirements of the mission parameter monitor and to provide power for the surplus system selected by the satellite design engineer. Both satellite-based solar cell arrays and airborne batteries must meet the electrical energy requirements of electronic sensors, stabilizing components, and control mechanisms. The typical power consumption requirements of early communication satellites launched from 1970 to 1980 ranged from 10 to 20 kW. When powerful communication satellites entered the communication field, power consumption later increased to more than 25 kW. The power requirements for the latest monitoring and reconnaissance spacecraft are at least 25kW or higher because of the deployment of high-resolution millimeter-wave side-looking radars, high-power tracking lasers, and precision photoelectric sensors. According to these power requirements, space solar arrays and batteries must be able to meet such power consumption requirements.

Battery power requirements and related key components
Figure 1 Design configuration of solar cell array (A) A 12 kw live TV broadcast; (b) a 5 kw TV satellite.

The design of satellite-based solar arrays must be able to meet the requirements of basic power consumption, as well as additional power requirements, to compensate for any electrical damage to the solar cells caused by the Van Allen radiation belt during the entire life cycle of the satellite. This additional power demand varies from 10% to 20% of the basic power level, depending on the life expectancy and the altitude of the satellite.

1. Performance requirements of solar cell arrays
The performance requirements of solar cell arrays strictly depend on the following issues:
●Satellite orbit;
●Spacecraft’s stability control and attitude control mechanism ;
● Power demand exceeding the operational life of the spacecraft or satellite;
●Mission requirements (communication or surveillance and reconnaissance);
●Available space for installing solar arrays on the launch vehicle;
●Specific design features, such as the extra power required by the dual-spin method or triaxial stabilization technology.

2. The power demand of solar arrays in the dark period
Depending on the duration, LEO satellites in low earth orbit may experience several dark periods every 12h. Satellite designers claim that communications satellites launched in equatorial orbit will experience the least amount of darkness. The designer also believes that due to its unique orbital characteristics, synchronous earth-orbiting satellites rarely encounter a dark period. Therefore, the solar array will receive unlimited solar energy to generate electricity for the needs of airborne sensors and tracking devices. The power consumption requirements of stability control and attitude control are relatively high. Therefore, the designer of the solar cell array must ensure that such a power level can be continuously provided. The deployment of solar arrays facing the sun provides almost uniform power output per track, which provides the best electrical energy with high reliability. The temperature of the solar array facing the sun varies from 60 to 80°C, depending on the atmospheric temperature at the altitude of the satellite’s operation.

3. Requirements for the orientation of the solar array to achieve the best power from the sun
In order to achieve the best power output from the panel, the orientation of the solar array is sometimes necessary. For surveillance and reconnaissance spacecraft, the orientation of solar arrays will be the most effective because more power is needed to supply multiple microwaves, high-resolution infrared cameras, high-power lasers for precise target tracking, attitude control mechanisms, and stability control sensors . In the equatorial orbit, the orientation of the battery array with one degree of freedom will maintain the normal solar impact plus or minus 23.5°C with seasonal changes. Other orbits may require two degrees of freedom. On these orbits, relative to the installation plane of the solar cell array, the spacecraft-solar line may be at any angle.

4. The most suitable solar array configuration for spacecraft or communication satellites
The electrical power output capacity of a solar cell array strictly depends on the number of solar cells used, the conversion efficiency of the cells, the array direction relative to the direction of solar impact, the panel installation plan, and the maximum size allowed by the cell array. The self-installed battery array shown in Figure 1 can provide hundreds of watts, while the fixed array configuration is most suitable for applications that require electrical power levels of several kilowatts.

5. Direct energy transfer system
When satellites are operating in the dark period (lack of sunlight) or flying in sunlight, the direct energy transfer (DET) system plays a vital role. Figure 2 clearly shows the key components of the system and its structural block diagram.

Battery power requirements and related key components
Figure 2 The key electrical components of the basic spacecraft power system

This special direct energy transfer system eliminates the insertion loss between a series of solar cell arrays and the load. The mode selection switch plays a key role in the operation of the system. This circuit ensures that the flow of load current and charging current in the system is appropriate regardless of the solar lighting conditions. The shunt regulator is always in working condition. When the bus voltage VR maintains the specified level, a small amount of current is connected to the ground through the branch line. The shunt current passes through one of the parallel resistors, generating a small threshold voltage. The direct energy transfer system plays a stabilizing role in the dark period and sun exposure period during flight.

During the sun exposure part of the satellite flight, if the output level of the battery array exceeds the requirements of the electrical load, the shunt regulator will remove the extra current and cause the small threshold voltage (ΔV) to exceed its predetermined threshold. When the overvoltage is sensed by the mode selection circuit, the circuit will turn on the charging regulator. Excessive voltage will allow excess current to be supplied to the battery.

If the communication satellite goes through a dark period, the threshold voltage (ΔV) will often drop below the designed threshold voltage level, allowing the mode selection circuit to turn on to turn on the boost regulator and the charging regulator at the same time. This will allow the necessary minimum battery discharge to meet the dark period load. The function of the shunt regulator can be implemented in several ways, namely as a dissipative full shunt circuit, or as a pulse width modulation (PWM) shunt regulator, or as a partial shunt circuit that operates continuously or sequentially.

The nodal current equations describing the direct energy transfer system in the dark part of the orbit and the bright part of the orbit are as follows:
eBR(VR/VBRO)IBD=(IL+ISH)
IA=(IBC+IL+ISD)

In the formula: VR is to regulate the load voltage; IA is the input current of the solar cell array; IBD is the discharge current of the battery; ISH ​​is the shunt current; eBR is the efficiency of the shunt regulator (see Figure 3); VBRO boost regulator Output voltage; IL is the load current; IBC is the charging current of the battery; IBD and ISD are the current of the shunt regulator.

Battery power requirements and related key components
Figure 3 Block diagram of the direct energy transfer power supply system on the spacecraft

The efficiency of a typical boost regulator is around 90%. If the optimal value of the circuit parameters is used, the efficiency can be increased to 95%.

Direct energy transfer power system designers believe that for the best system reliability, the minimum size of the solar cell array should be 110 series in parallel, with 63 cells in series in each string, including at least 6,930 solar cells. These solar cells will limit the battery charging current to approximately 5.52 A.