University of LeicesterPLUMERef: PLM-SYS-PSUInterSDD-604-1

Date: 25/03/2009

PSU Interface and SDD

D.S.W.Gray

Date / Updated Reference Number / change
25/03/2009 / PLM-SYS-PSUInterSDD-604-1 / first draft issued

Definition

The role of the PSU is to provide electrical power for all the subsystems onboard the PLUME cubesat. This is will be achieved by the use of photovoltaic (PV) panels when the satellite is in direct sunlight and by two Lithium Polymer batteries when the Sun is eclipsed by the Earth’s shadow. Both of these will provide direct current (DC) power. The batteries are charged by the solar panels through three battery charge regulators (BCRs) which optimize the solar arrays’ voltages independently for maximum power transfer. Power is then transferred from the battery or directly from the solar panels through a number of power buses, rated at 5V, 3.3V and unregulated, to which the other subsystems are attached. Both the regulated buses have over-current protection, to prevent damage from any short circuit, and the battery has under-voltage protection to prevent a complete discharge.

Telemetry and telecommand is provided to the PSU via an I2C bus which connects through the central OBDH/MCU through standard PC/104 connectors. Commands can be sent though this system and information returned on the battery’s current, voltage and temperature as well as the Solar panels’ current, voltage and temperature and the buses’ currents.

Below is the current up to date block diagram for the whole PSU system.

Fig 1. PSU block diagram

Now follows a more detailed description of each of the main components of the PSU.

Components

Solar Panels

The solar panels have currently not been finalised but will consist of Clyde Space’s 1U CubeSat panels, customised with holes cut for the payload and camera in the arrangement shown in figure 2.

Fig 2. PV cell arrangement: 3x fully covered 100cm2, 2x with 16cm2 free for the payload, 1x with 4cm2 free for the camera.

The cells themselves will be Spectrolab’s Ultra Triple Junction (UTJ). It’s a triple junction Germanium/Gallium Arsenide cell with efficiency of 28%. For full details see the Spectrolab datasheet.

As the spacecraft’s orientation with respect to the sun cannot be controlled the panel area receiving sunlight is effectively random. Based on calculations of an averaged projected area the panels should produce an average of 4.60W when in direct sunlight. Over an entire orbit the average power will be lower than this, dependent on the exact time spent in the Earth’s shadow. If the satellite spends half its time in eclipse, for example, it will generate an average of 2.30W over an orbit.

For general information on PV cells see Solar Cells Overview and Solar Cell Thermal Properties.

For detailed calculations on predicted power output see Power Production and Power Production Calculations.

See also Mechanical Interface.

Battery

The battery to be used consists of twoClyde Space Lithium Polymer cells, mounted together in series, side by side on top of the power generation board. It contains an integrated thermostatically controlled heater, battery telemetry and cell over and under-voltage protection and over-current protection. The battery boardhas the following characteristics:

  • Size: 58.5mm x 37mm x 5mm.
  • Mass (per cell):22.5g
  • Mass (total for two cells + PCB)62g
  • Capacity: 1250mAh.
  • Maximum charge voltage 8.2V.
  • Minimum discharge voltage 6.4V.
  • End of Charge Limit (EoC): 4.1V.
  • End of Discharge Limit: 3.2V.

and is rated for the following conditions:

  • Radiation up to 500krad.
  • Discharge between -20°C and 60°C.
  • Charge between 0°C and 45°C.

The battery heater is an independent analogue circuit which automatically comes on to keep the batteries above 0°C. It can be overridden by a command from the I2C bus if necessary (i.e. to conserve power or if a fault is detected).

A battery board has been purchased from Clyde Space and will be used in the lab (see Testing below) to test the capacity, charge and discharge times etc. The actual flight battery will be identical but has not yet been purchased.

For more information on the battery board see theClyde Space manual.

Power Generation Board

The battery board is mounted on top of the Clyde Space power generation board to create the Electronic Power Supply (EPS) component of the PSU. The total EPS is contained on a single PC/104 board, occupying one level of PLUME’s stack of boards. This has been purchased and measured in the lab to have the following characteristics (total including battery):

  • Size: 95mm(l) x 90mm(w) x 15mm(d)
  • Mass: 152.96g
  • Power Consumption:<0.1W

The board contains three BCRs (one for each axis: x, y, z, see figure 1). These automatically match the solar panel voltage to their input voltage in order to transfer maximum power. Each is capable of delivering up to 3W with an input voltage of 3-10V and a maximum output of 10V and all three are rated as over 90% efficient. The BCRs are also self-sustaining and do not require power from the battery, meaning they can effectively charge the battery at any time. They do this using a taper charge method whereby they act as a current source to the battery when it is below its pre-set end of charge (EoC) voltage. Once this point is reached the BCR output voltage and current are kept at this level to charge the battery and their input from the solar panels is allowed to drift.

BCR 1 can also interface with the main connector to the 5V USB input, allowing the batteries to be charged via USB before launch.

The BCRs connect to the battery through a pull-pin and to the power buses through a separation-switch, as shown in figure 1.

Also contained on the board are the regulators for the power buses which actually supply DC electrical power to the other cubesat subsystems. These are the 5V bus which is rated for a full load current of 1.2A and the 3.3V bus, rated for 1A. There is also a third unregulated bus. All three buses have over-current protection switches which prevent damage to the PSU in the event of a short circuit by switching the affected power bus off for a few milliseconds. They also contain telemetry units that can be called on to display the currents drawn from each bus.

Finally the board also contains the TTC node, a detailed view of which can be seen in figure 3. This is made up of a Microchip PIC16F690 8-bit micro-controller as well as an analogue-to-digital converter (ADC), FLASH based program memory, built in RAM, 8,000MHz crystal oscillator clock and I2C slave interface. Together this system receives analogue inputs from the battery and power bus telemetry units and digital inputs from the I2C bus and can produce digital outputs for the I2C bus and an override command for the heater control. The outputs to the I2C bus are then sent to the MCU and back to mission control and can then be decoded with the calibration equations (see interfacing below) to give telemetry on the current status of the battery and power buses.

Fig 3. TTC node

For more information about the EPS see the Clyde Space Manual.

Interfacing

Telemetry and telecommands interface to the EPS is provided by an I2C digital interface. This is physically connected through the CubeSat standard 4 by 26 pin PC/104 push-to-fit connector mounted on the board which matches pin to pin with the MCU board. The pin configuration is shown below in figure 4.

Name / Type / Bus Pin No.
Alternative I2C Clock / 3.3V logic / H1.21
Alternative I2C Data / 3.3V logic / H1.23
ON_I2C / 3.3V logic / H1.24
I2C Data / 3.3V logic / H1.41
I2C Clock / 3.3V logic / H1.43
+5V_BUS / +5V Power / H2.25
+5V_BUS / +5V Power / H2.26
+3.3V_BUS / +3.3V power / H2.27
+3.3V_BUS / +3.3V power / H2.28
GND / Digital Ground / H2.29
GND / Digital Ground / H2.30
AGND / Analogue Ground / H2.31
GND / Digital Ground / H2.31
VBAT(PP_NC) / Launch switch connections / H2.33
VBAT(PP_NC) / Launch switch connections / H2.34
VBAT(SS_NC) / Launch switch connections / H2.35
VBAT(SS_NC) / Launch switch connections / H2.36
VBAT(PP_SS_C) / Launch switch connections / H2.41
VBAT(PP_SS_C) / Launch switch connections / H2.42
VBAT(PP_SS_C) / Launch switch connections / H2.43
VBAT(PP_SS_C) / Launch switch connections / H2.44
BAT_BUS / Battery Voltage Unregulated / H2.45
BAT_BUS / Battery Voltage Unregulated / H2.46

Fig 4. Pin configuration for PSU interfacing. For full description of pin functions see Electronics Interface Document.

On Board EPS to Solar Array connectors (called SA1-3) are type DF13-6P-1.25DSA.

Harnesses should use connector type DF13-6S-1.25C.

Solar panel connectors should be type DF13-6P-1.25H.

The TTC node is configured as a slave through the I2C interface to the MCU which acts as the master. Read and write commands can be sent through this interface and the TTC node will respond with telemetry data. For a full list of these commands see the Software Interface Document. PSU telemetry can be stored on the SD card by the MCU for transmission to the surface. Data includes panel voltage, current and temperature, and battery voltage, current, current direction and temperature.

Mechanically the battery board slots onto the Power Generation Board and together they take up one single slot in the stack. The order of the stack depends on the masses of each board, and will be changed to ensure that the centre of mass is no more than 2cm from the geometric centre of the cube. Each PCB is connected to the others via four main bolts (approximately 2mm diameter and 93mm long) with aluminium spacers and washers that fit between and separate the PCBs. These spacers are either 15 or 25mm long depending on the required space between boards. The EPS and batteries together are 14mm high, so require the 15mm spacers (one in each corner).

For more information on interfacing see the Software Interface Document, Electronics Interface Document and Mechanical Interface Document.

Testing

The full test plan has not yet been finalised at the time of documentation.

All testing of delicate, static sensitive equipment is undertaken on a laminar flow bench in the SRC lab while testers are required to wear gloves and a grounding wrist strap.

The tests to be performed include measuring the charging and discharging cycle of the battery. This will be done by using an external power source to simulate the solar cells and timing how long the test battery takes to reach its maximum charge level. The battery is then timed as it is discharged to its minimum level through a battery discharge circuit, designed to emulate the maximum load the PSU can expect from the other subsystems. See Charging/Discharging Test Document for further details.

The I2C bus commands also need to be tested to make sure telemetry and telecommand function correctly. For the purposes of this test the solar panels are again simulated by connecting an external power supply to the SA connectors and the battery is replaced with a battery simulation circuit, designed by Clyde Space. The MCU is replaced by a laptop with the appropriate software, connected through the USB connector and functioning as the I2C master. Command are inputted in hexadecimal through the laptop and the resulting returns recorded and decoded as data on the voltage, current, temperature etc. These are then verified using multimeters. This testing was in progress at the time of documentation. For further information see PSU Test Document.

Once the flight battery has been purchase this will need to be tested in order to establish whether it works correctly. This will likely involve a single charge/discharge cycle in order not to reduce its capacity.

The solar cells will also need to be tested once they are purchased to make sure that they are function and are able to charge the battery correctly.

Once all of the separate sections have been purchased an assembled together the PSU as a whole will need to be tested to determine that it works as described in the above document.

For More information on testing see the Test Plan.

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