Power

Overview

This page contains power information specific to LI Bus. For more general power information, see the general power page.

Power Design Changes Impact Assessment

The LI Bus project will borrow heavily from the power architecture of previous projects. At the same time, some elements of its design will require significant changes. The main areas where this will happen are the Li Bus payload interfacing and orbital control.
Little is known about the power requirements of the LI Bus payload. Currently, a conservative preliminary estimate for continuous power supply is all the team has. As the payload design develops further, the power team will need to stay up to speed on the exact power requirements of the payload so they can ensure that they are met.
Current plans are for LI Bus to include some type of thrusters for orbital control. Again, power requirements for these are unknown, and the team must acquire this information before completing the main level power architecture. Depending on the power requirements for the thrusters, the team may need to incorporate more switches to control them, which might mean redesigning the power control boards or even designing new ones.

Power Design

The preliminary design of the Power system is derived from the Iris Power. The Power subsystem provides means for generation, storage, and distribution of power among all of the satellite's subsystems. A summary of design considerations for the power subsystem is listed below. 

• All of the energy storage devices are enclosed within the internal structure of the satellite bus. 

• The power design avoids using any sort of pressure vessels in forms of electrolytic capacitors or any other electronic components. If any is used, a fully detailed report shall be provided by PWR lead. Additionally, appropriate records and procedures will be provided to assure the system is safe and meets NASA requirements. 

• Electrical bonding of the power components including soldering of the components to the board, crimping the connectors, and using screw terminals will follow the guidelines highlighted in SSP-30245. 

• No beryllium, cadmium, mercury, silver and other prohibited materials listed in SSP-30233 is used. (if silver is used in solar panel assembly, an RFD shall be submitted)

• Imported products and services such as solar cells, batteries, electronic boards, test equipment, etc. will comply with the import conditions imposed by the federal and provincial governments. 

• The batteries will go through standard assessment and screening as per highlighted by NanoRacks [launch provider].  

• Additional assessment and screening will be planned if the capacity of the battery design exceeds 80.00 W.hr. 

• Power design does not contain any stress corrosion susceptible materials. If any of the materials listed in "Table II of MSFC- SPEC-522B" is used, those materials will be documented in the ICA. None of the materials listed in "Table III of MSFC- SPEC-522B" can be used.

• if any capacitors are used as energy storage devices, these capacitors shall be treated and reviewed like batteries. 

• ArcticSat's power system does not use any "wet capacitors". if any is used, it shall be reported to NASA with all the details of the capacitor. 

• Battery charging equipment will provide at least two levels of protection (e.g. overvoltage and overcurrent protection features). The design has OC, OV, SC, UV features.

• The ground charge equipment will not energize the satellite systems.  

• A timer circuit inhibits the power system from turning on other systems until hazardous conditions after deployment are eliminated. Pressing the inhibits will reset this timer to increase safety. 

• The wiring within the EPS shall use appropriate gauge and lengths of wire. For the batteries, no longer that 6 inches of 26 AWG wire is allowed. 

• All energy generated by the power system is autonomously and sustainably made using solar cells. 

• Power system is able to communicate with satellite's main processor "CDH", allowing it to send health telemetry to ground through CDH and COMS subsystems. 

• Communication with other subsystems happens through the primary CAN bus protocol.

• Auxiliary CAN bus is implemented to communicate with the load switches on each subsystem. 

• Temperature of the batteries and other electronics will be controlled and maintained within the proper range using a heater system. 

• Power will provide all the other systems with their required power for the mission duration of at least 1.00 year under orbital conditions. 

• The power subsystem is able to withstand airlock de-pressurization rate of 6.9 kPa/S. 

•  The power system is able withstand an airlock pressure of 0 to 104.80 kPa. 

• The Power design achieves reliability of at least 80% by using redundant systems. Two analog and one digital timer, three inhibit switches, and two battery thermistors are examples of redundant designs. The majority of components are chosen to be automotive grade to allow power to function in the orbit radiations environment.

• Health telemetry containing the power draw of each subsystem, temperature readings, power inputs from each solar panel, and battery SoC is collected regularly and is stored for transmission to ground during the next pass.

• The power design follows the guidelines highlighted in NASA's hazardous materials guideline documents. 

• All non-metallic components in power subsystem shall be compliant with NASA's guidelines for outgassing properties. 

• The power system shall comply with NASA space debris mitigation guidelines as per NASA-STD-8719.14A 

• A remove before flight (RBF) feature is implemented in power design to keep the satellite in an unpowered state throughout the ground handling and integration process. This RBF feature is accessible via the access panel on the +y face of the dispenser. 

• The batteries and load switches are equipped with protection circuitry to avoid internal and external short circuits.

•  To comply with outgassing requirements, the boards shall be conformally coated. This ensures that CVCM value is less than 0.10 percent and TML is less than 1.00 percent. 

• The deployable solar arrays are designed to be independent of the dispenser's mechanisms. 

The power system will provide power to the satellite's subsystems based on the operation mode. The operation mode is based on the battery's state of charge and is based on the conditions below: 

• If SOC >= 50% --> NORMAL POWER MODE

• If SOC < 50% --> LOW POWER MODE

• if SOC < 40% --> SURVIVAL MODE 

• if SOC < 30% --> CRITICAL HOLD MODE 

*The power modes decrement autonomously by the power control board based on the battery state of charge. Increment of the power modes is only allowed through commands from ground station which are relayed to the power subsystem through the COMS subsystem. The only exception is transitioning from critical hold mode to survival mode. This transition happens autonomously by the power subsystem when the battery state of charge goes above 50%. 

Table below shows the power modes and how the power subsystem performs load shedding depending on the power modes. After the satellite is released from the deployer, the power subsystem starts harvesting the solar energy until the battery state of charge is at least 60% before commencing the normal operations. 

The logic for mode switching is presented in more details in the power state flow chart below. 

State of Charge Estimation

The Power subsystem has three methods of estimating the state of charge of the batteries. 

1. Using the LTC4150 IC that is placed on the board and connects directly to the battery pack. 

2. Using a software algorithm loosely based on the enhanced Coulomb counting algorithm. This method measures the current going into the battery by subtracting the current drawn by the subsystems from the current generated by the solar panels. The flow chart for this method is shown below where Cmax is the battery's rated capacity in Coulombs and Vmax is the maximum battery voltage.  

3. Using the battery pack's voltage as a backup method and referring to a lookup table 

Power EM Design

The Power module handles the power generation, storage and distribution on the satellite. The main components on the Power subsystem include a inhibit switches, 30-minute timer, batteries, solar panels, control board, and CAN/SPI transceivers. All hardware is mounted inside a 1U shell, forming the power subsystem assembly.  

The current Power design is in its preliminary phase and further details pertaining to the Power, such as specific components and battery sizes, will be defined later. The Power system block diagram is shown in Figure below. 

The main design for the power subsystem is based on the Iris power design. The new changes are mainly on the energy distribution section. The new power design distributes the load switches across all the other subsystems that require power by using CAN Controlled Load Switch and Monitor (CCLSM) devices. These devices provide a smart power control system for the subsystems, achieving the power distribution goals while simplifying the power subsystems electronics and simplifying the required harnessing.

The block diagram for the CCLSMs is shown below. In addition to switching loads, the CCLSMs can provide power telemetry for each of the satellite systems to the Power subsystems. 

Power Model Philosophy 

We plan on manufacturing multiple models for the power subsystem to ensure that the functionality of the system is evaluated in multiple stages. Two possible designs for the power subsystem are going to be made. One is exactly the same design as our previous satellite (Iris) power subsystem and another design based on the new distributed load switch design. A breakdown of models for each phase is presented here which is derived from the LI Bus Model Philosophy spreadsheet. For the Legacy Power design (Iris) we have: 

• Phase A: Rebuild the engineering model based on the design of Iris and test it.

• Phase B: No action needed for this phase. Continue development using the Engineering model.

• Phase C: Use the Engineering model in FlatSat and integrate it with the rest of the satellite subsystems and evaluate the performance of the boards when operating with the other subsystems. In this phase, build the Flight model as well.

• Phase D: test the flight model in flatSat.

 For the new Power design (Power MkII) we have: 

• Phase A: make a power simulation model and evaluate the preliminary design parameters of the power subsystem. Also develop a Breadboard model to test the distributed load switches design.  

• Phase B: Update the simulation model and update the breadboard model. Build an Engineering model in this phase. 

• Phase C: Update the simulation model. Update the engineering model. Build an Engineering Qualification Model to test the subsystem before making the final flight model.

• Phase D: Update the simulation model.  After making the final revisions on the Engineering model, a final flight-ready version of the power subsystem electronics is made. This model is the "Flight Model."

Power Analysis

The power subsystem is to be analyzed in multiple stages to ensure adequate power is available for all of the satellite's functions. 

In order to do a power analysis, the following power requirements is estimated for each subsystem. 

• Each subsystem's power consumption in Watts  

• Solar panels power generation 

• Battery voltage and capacity 

• Satellite's orbit and the duration of sun light/ eclipse moments

The analysis values are updated in each phase as more accurate estimates are provided from each subsystem. More details are presented on the Power Analyses page.

Energy Generation - Photovoltaic cells

Photovoltaic (PV) cells are the primary source of power on the spacecraft. The PV cells are triple junction n-on-p cells (RocketLab's ZTJ-Ω). The solar cell's datasheet is shown below. Detailed information about the cells and array design can be found under solar array sizing. The power system autonomously handles power generation by harvesting the electrical energy generated by these panels and storing them in the battery pack. 

Figure below shows the electrical connection of the solar strings. The specified cells include integrated bypass diodes, which allow current to flow safely in event of partial shading or string current mismatch. The electrical specifications detailed in the table are based on the 3S string configuration. Power system will collect current and temperature telemetry from solar panels for satellite health monitoring purposes. 

Energy Storage - Lithium Iron phosphate (LFP) battery pack

Energy is stored onboard the spacecraft in a battery pack constructed from Lithium iron Phosphate cells. The specified cells are obtained from AA Portable Power corp. (IFR-18650EC-1.5Ah). Each cell has a rated nominal capacity of 1500 mAh and nominal potential difference of 3.2 V. The main reason for selecting LFP cells is to take advantage of its relative flat voltage profile, because the battery pack is the main source of capacitance on the unregulated MPB. Consequently, the operating voltage of the solar strings and spacecraft loads is equal to the battery pack voltage, the relatively flat voltage profile is therefore desirable.

Figure below shows the battery pack in a 2S3P configuration. Based on the given battery pack specifications, operating temperature control is crucial. As a result, two battery heaters are used from the Thermal subsystem to condition the batteries. 

Power subsystem will collect temperature, current, and voltage telemetry from the battery pack for health monitoring purposes. Additionally, battery's state of charge will be measured in regular intervals every 1 minute. 

Separation switches and RBF

There are a total of 3 separation switches connected in series. The separation switches are placed at the bottom of the power shell, making contact with the deployer and when compressed, they shut off the power of the entire satellite. The RBF pin is located on the side of the shell and is depressed by the RBF pin at all times during integration and testing and keeps the spacecraft at an unpowered state. These switches are wired according to the diagram below and are connected to the main power board. 

The location of these switches will be on the outer shell of the satellite close to the power board. An example illustration of the switch locations is shown below.

Mission and Systems Level Requirements Impact Analysis

The LI Bus project will borrow heavily from the power architecture of previous projects. At the same time, some elements of its design will require significant changes. The main areas where this will happen are the LI Bus payload interfacing and orbital control.

Little is known about the power requirements of the LI Bus payload. Currently, a conservative preliminary estimate for continuous power supply is all the team has. As the payload design develops further, the power team will need to stay up to speed on the exact power requirements of the payload so they can ensure that they are met.

Current plans are for LI Bus to include some type of thrusters for orbital control. Again, power requirements for these are unknown, and the team must acquire this information before completing the main level power architecture. Depending on the power requirements for the thrusters, the team may need to incorporate more switches to control them, which might mean redesigning the power control boards or even designing new ones.

The driving mission and systems level requirements are shown below:

• R-LIB-MIS-001: The LI Bus shall accommodate a generic imaging/remote sensing payload.

  • R-LIB-BUS-078: The Bus shall be capable of providing 10 W of continuous power to payload.

• R-LIB-MIS-007: The LI Bus shall be capable of interfacing with the Rocket Lab 6U canisterized satellite dispenser (2002367F).

See the Valispace registry for more requirement information.

Schematics 

The electrical schematics for the power subsystem are shown below. The critical parts of the design such as the inhibit mechanisms and the startup timer are designed with redundant backup mechanisms to improve reliability of the design. 

Parts List 

The majority of sensitive electronic parts used for the design of the power avionics were chosen to be of automotive grade or space grade as shown in the parts list below. There are some commercial grade parts in the parts list that are mainly passive components and connectors. The commercial ICs will be closely monitored and will be swapped for automotive grade components in the FM version if direct replacement parts are found that do not alter the design much, do not cost significantly more, and do not result in an extended lead time. The parts are all procured from global suppliers like Digi-Key and Mouser which are compliant with all the import laws imposed by the federal and provincial governments. 

Note that where possible, the Power uses automotive grade parts to ensure high reliability. However, the majority of the passive components on the Power subsystem are commercial or industrial grade parts. Critical components such as the STM32 ARM processor, G3VM-31HR1 Solid State Relay, INA250 Instrumentational amplifiers and the LTC4150 Coulomb counter are also industrial or commercial grade. 

Temperature

• All parts on the Power meet the thermal operation and environment minimum and maximum temperature requirements.

Vibration

• The base Power design for LiBus is the same as the Iris Power design which underwent random vibration testing prior to launch. The Iris Power was tested before and after the full satellite random vibration test and was fully functional after the vibration test and in space after deployment from the ISS. Given that the LiBus Power design is identical in its core and manufactured from the same manufacturer, it is expected that the LiBus Power will also be reliable under launch vibrations.

Radiation

• STM32 Processor: Test results reported in literature [reference] show that the commercial grade STM32F1 family microcontrollers have a failure total ionizing dose of around 107 krad which is higher than the expected 10 krad(Si) dose for LiBus based on the SPENVIS analysis. Given that the power system uses STM32 microcontroller from the same family as the one tested in the literature, we do not expect it to fail during LiBus mission. 

• G3VM-31HR1 Solid State Relay: According to the previous success of this part in a cubesat mission highlighted in their lessons learned [reference] which used the same component in their power system, we do not expect our SSRs to fail during LiBus mission. 

• INA250 Instrumentational amplifiers and LTC4150 Coulomb counter: According to the previous success of this part in a cubesat power systems design highlighted in [reference] which used the same component in their power system, we do not expect our SSRs to fail during LiBus mission. 

PCB Layout EM

Manufactured Hardware and Testing

On February 8th, 2024, we received the boards for POWER-EM hardware. These boards were manufactured by Servetronics company.  Coding and testing of these boards are underway and will be documented here. The components are properly soldered and bonded which is compliant with the guidelines outlined in the SSP-30245. 

FM PCB Changes

After finishing the testing on the EM board and finalizing the design details of other subsystems (structure and data interfaces and power switches), we performed the following changes on the EM board to make our Flight Model board for the Power subsystem. 

1. Increase deployment timer to 5 minutes (300 seconds)

2. Add a second harness connector on the bottom side of the board

3. Add a power switch for the thermal heater with 2A current limit 

4. Move components that are in the keep-out zone as highlighted by structural drawings

5. Remove the J22 connector -- this connector provided pinouts for SPI pins of the power control board which is not used.