The greatest single cause of satellite failure is a malfunction in the power system. A number of satellites have been lost when the solar array failed and those on the ground had to watch helplessly as the batteries slowly waned away and gave up the last of their life. It is therefore quite imperative to have a robust power management system that is capable of supplying all the power needs but is also simple enough to work without the possibility of failure.

The main component to the satellite’s power is the solar array. Even with fully charged batteries, the satellite will only be able to operate a maximum of 2 hours without an operational array.

Several criteria were used to judge the selection of a solar cell brand:

• Dimensions of Individual Cells (constrained by specific array dimensions)
• Efficiency (desire maximum efficiency and GaAs/Ge type)
• Voltage Output (desire voltage output appropriate for battery charging)
• Ease of Implementation (desire coverglass, metal interconnects, i.e. CIC form)
• Cost (desire minimum cost)

The solar cell we chose to fly was one made by a company called Spectrolab (Fig. 5). This company makes space-grade cells that have as high as 25% efficiency. The 1.23" x 2.72"cells that we will be flying have an average efficiency of 19% (which is more than enough for our uses) and will produce a current of 0.273 Amps per cell. We will be putting 48 cells on the satellite, these cells are arranged in eight strings of six cells each. This configuration will give us a total voltage of 12.24 V as well as a power output of 3.34 W. 52 watt hours is the maximum power produced per orbit if the satellite is under eclipse orbit or 136 watt hours maximum per orbit if the satellite is in constant sun orbit. These numbers mean that we will have a high enough voltage to charge our batteries as well as a sufficient amount of power to run all the electrical components on the satellite. On top of this, these GaAs/Ge type cells also come in Cell Interconnect Cover (CIC) form. This means that the cells themselves already have metal interconnects to attach to each other as well as a special protective coverglass that will slow down the degradation process. All we have to do is attach them to the satellite using SN62 solder (62% tin, 36% lead, 2% silver) to solder cell connections and wiring and a two part GE silicon RTV566 to adhere the cells to the backing. We will also be putting a diode on each cell as well as each string to prevent non-working cells from adversely affecting those that are operating properly.

The backing we are using for the solar cells is a 0.012" thick copper clad PC board. This board is not only lightweight, but will act as a very convenient wiring system for the cells. In addition, the boards themselves give some diagonal stability to the frame itself.

Fig. 5—A sample of one of the GaAs/Ge solar cells we will be flying.

Batteries comprise a very important component to the functionality of our satellite as a whole. If we are in a 50% eclipse orbit, then we will be relying on batteries for half of the operation time. Also, initial start up and signal beaconing is going to be run solely on battery power until the solar panels are up and running. Therefore, choosing the correct battery to fit not only our size and weight requirements, but our power usage requirements as well becomes imperative. The criteria we used to judge the selection of a battery cell brand were:

• Dimensions of Individual Cells (constrained by specific volume)
• Longevity (desire maximum cycle life)
• Voltage Output (desire voltage output appropriate for battery charging)
• Capacity (desire maximum capacity)
• Mass (desire minimum mass)

We chose a Lithium Ion battery (Panasonic CGP345010). Each of these 1.34" x 1.96" x 0.41" cells hold 3.7 V and can deliver up to 1400 milliamp hours. Our design incorporated using two battery packs of two cells each. Each battery pack has the two cells in series, which increases the total voltage up to 7.4 V and keeps the current the same. This gives us enough power to operate all the electrical components at the same time for a little over an hour. By having two battery packs we have a multifunctional power supply which will give us a wide range of power capabilities. For instance, each battery pack can be used together in parallel if a burst of current is needed, one battery pack can be in use while the other is being charged, or we could toggle the use of each pack depending on performance or amount of energy stored in each. Our pack arrangement meets any electrical need that could arise on the satellite including emergencies. Moreover, the StenSat program has performed extensive testing of the Lithium Ion battery and demonstrated its abilities for space use. They claim to have run the batteries for 4000 cycles so far, although the data indicates this was for a ~5 hour constant current charge, ~2 hour constant discharge, with a 20 minute break between sessions. This is comforting for similar cycling would occur during the operation of our satellite. The charge/discharge period becomes more balanced with respect to time and more variable with respect to current if the satellite is in the eclipsed orbit. However, on average it will be about 50 minutes on and 40 minutes off.

Charge and discharge protection circuits will be needed in addition to the protective circuits installed in the battery. The pack has an upper charging limit of 8.2 V and 900 milliamp hours. If the pack voltage rises above this, power must immediately be cut-off. If the voltage were to drop below ~7.2 V, the satellite should go into emergency mode. Shortly after this point, battery output begins to drop quickly. If the pack drops to six volts, all power usage must be terminated because the battery is about to die.

To lengthen the life of the batteries and enhance the function of the satellite, the power distribution electronics have maximum flexibility (Fig. 6). This flexibility will help prevent critical failure of the mission if any problems arise. For example, power produced from the solar array passes into a digital multiplexing switch. From there, it can be diverted to charge either of the two battery packs or power the satellite directly. Furthermore, power from either or both of the battery packs can be used to supplement power to the satellite electronics during high current draws. If either battery pack exhibits erratic behavior, that pack can be isolated from the rest of the system without compromising the mission. Furthermore, the satellite can operate solely upon the solar arrays, although functions would be limited to daylight use only.

High efficiency regulators adjust the system voltage of eight volts to the various needs of the spacecraft: seven volts for the plasma probe and TNC, six volts for both of the transceivers, and five volts for all other digital electronics. The microcontrollers receive data from a series of amperage and voltage probes. They are programmed to automatically adjust the power distribution during emergencies and limit power usage if the batteries are insufficiently charged.