Power (EPS)

Subsystem Top Level Requirements
-- The system shall provide a nominal power output of 0.5 Watts
-- The system shall provide 8 Watts of peak power for communications over Boulder, Colorado
-- The system shall be capable of continuing operations through all parts of the satellite orbit

Subsystem Functional Requirements
The system shall not exceed a mass of 200 grams and a volume of 150 cm^3
-- The system shall have two parallel separation switches in the power path to ensure the circuit is open prior to ejection from the CalPoly P-POD
-- The system shall have a remove before flight pin that will disable all power to the spacecraft prior to integration with the launch vehicle
-- The system shall automatically cycle power to the entire satellite without any input from the Command and Data Handling (CDH) subsystem at a specified interval
-- The system shall provide a control method for the power distribution to each subsystem
-- The system shall provide health and status data of current power usage and battery status to CDH
-- The system shall automatically protect all hardware in a low power situation
-- The system shall have an efficiency of greater than 70% from collection to distribution
-- The system shall be scalable for larger future missions
-- The system shall be capable of launching with the batteries at any charge state


The Power subsystem for the Hermes CubeSat will provide an average power of about .5W and a peak power of 15W, which will come from solar cells and Lithium-ion battery packs.  The source of power will be controlled by a Smart Battery Manager/charger chip will automatically change the power path as the satellite passes from the sunlit portion of its orbit to the eclipse portion and vice versa.

Solar Panels
Each of the six sides of the cube satellite will have a two-solar cell panel built on a PCB substrate. Each solar cell has a nominal voltage of 2V and an efficiency rating of 25.1%. This yields a maximum power output of 2W per cell. The two cells will be connected in series, yielding a voltage of 4V and a maximum power output of 4W. As the satellite orbits the earth, the solar cells will degrade due to radiation damage and other environmental effects such as the presence of atomic oxygen. De-rating the solar cells appropriately to take into account the degradation, the expected end of life maximum power output will be roughly 3.5W after two years in orbit.

Current Sensors
Each of the solar panels will have a MAX4070 current sensor in series between the panel and the boost converter. This will allow CDH to gather data about incoming voltages from the solar panels to be used for power system management as well as attitude determination.

Boost Converters
The voltage from the solar panels is not high enough to charge the batteries or to operate some of the subsystems such as PCOM or ADCS. To rectify this issue two MAX1771 boost converters, connected in parallel for redundancy, are located between the solar panels and the battery charger/manager to convert the voltage to 10V. This stage will be between 80-92% efficient, depending upon the current load.

Charger Manager
The heart of the Power subsystem is the LTC1760 battery system charger and manager. This component will control which of the power sources provides power to the satellite at any given time: solar panels, one or both of the batteries, or a combination of sources. The charger manager interfaces to CDH via the two-line SM Bus protocol. A third interrupt line will also connect power and CDH for the charger manager to interrupt the main CDH processor when it needs attention.

Battery Protection Circuitry
The two Lithium-ion batteries will provide 100% of the power when the satellite is in eclipse and will augment solar cell power during high current draw operating times such as communication passes. There are two battery packs, connected in parallel, each constructed of two cells in series. The packs have a nominal voltage of 7.4V and a capacity of 600mAh. Each battery pack has fuel gauging and protection circuitry which interfaces to the LTC1760 charger manager chip.

Buck Converters
Since each of the subsystems require a different operating voltage, four buck converters are used to convert 10V to the appropriate voltage. The 3.3V buck converter for CDH and the 5V buck converter for HSCOM each have an expected efficiency of greater than 95%. The two 7.4V buck converters for PCOM and ADCS have an efficiency of greater than 97%. Each voltage line will be regulated to 5-10% ripple and will have high frequency components removed.

CDH has the ability, via Field Effect Transistors (FETs), to turn on and off the buck converters for various subsystems. Only the 3.3V buck converter does not have FET control as it powers the satellite controller and should always be on whenever there is power.

While the Terminal Node Controller (TNC) is part of the PCOM system, it actually uses 3.3V rather than 7.4V. Rather than powering it from the 7.4 PCOM buck converter and being forced to regulate the voltage down, it is powered off of the 3.3V CDH buck converter. A FET is located post-buck converter, but pre-TNC so that CDH can turn the TNC on and off.

Power Predictor Model
A power predictor was created in Matlab to model the expected power usage throughout the lifetime of the mission. The model was also run with End of Life (EOL) efficiencies inserted to determine the available power as the mission nears its end. The Power system was designed so that even at EOL, the spacecraft will have enough available power to fulfill all of its needs.

If you have questions or desire more information about the Hermes Power subsystem, please email the Project Manager.


Links & Downloads

MAX4070 Current Sensor Datasheet

MAX1771 Boost Converter Datasheet

LTC1760 Dual Smart Battery System Manager & Charger Datasheet

TI-bq20z90 Fuel Gauge Datasheet

TI- bq29330 Lithium-Ion Protection Datasheet

Lithium-Ion Battery Datasheet

LTC1148 Buck Converter Datasheet