AODCS

Mission and Systems Level Requirements Impact Analysis

LI Bus builds on the Attitude Determination and Control System (ADCS) from Iris. In addition to the sensors on Iris, LI Bus will include a star tracker as an additional attitude determination tool. The only attitude actuators on Iris were three magnetorquers, these are supplemented with a reaction wheel cluster to meet the finer pointing requirements for LI Bus. LI Bus must also know its position in orbit and be able to make changes to its orbit. For this, LI Bus will also have a GNSS module for position and velocity determination, and a cold gas thruster to perform orbital maneuvers. The addition of these components will result in increased power draw, an increased volume required within the satellite structure, and more software to be written to implement these sensors and actuators.

The more major changes from Iris are that LI Bus will be a 6U CubeSat rather than a 3U, which could ease the implementation of additional AODCS components and would change its mass properties meaning the actuators must be configured differently to Iris. 

Lastly, the LI Bus mission will operate in an orbit different from Iris. Where Iris operated in an ISS orbit which sees an eclipse every orbit, LI Bus will be in a higher altitude sun-synchronous orbit. This introduces nearly consistent sun exposure throughout the year. This risks putting thermal strain on the structure and other subsystems, but reduces strain on the batteries while operating in eclipse.

See the Valispace registry for more requirement information about LI Bus.

AODC System

The AODCS subsystem determines the position, velocity, and orientation of the spacecraft using various sensors. The spacecraft then turns that data into commands for actuators to control the orientation of the spacecraft. With the addition of a thruster to LI Bus, controlling the orbit will also be possible. A block diagram depicting the interactions between AODCS and intersubsystem components is shown in the figure below. This diagram shows the flow of power from the power subsystem, and data generated and received by different components.

For detailed information about the subsystem and all components, see the Interface Control Document.

EKF  for Attitude Estimation

The AODCS will run an Extended Kalman Filter (EKF) onboard the CDH to provide attitude estimation using updates from all of the attitude determination components. The purpose of this estimator is to combine telemetry from all attitude and position sensors, and output a single pointing quaternion as the spacecraft's current orientation. Using this, the spacecraft can regularly estimate its orientation even if some attitude sensors are not available. For example, if the spacecraft performs a maneuver that causes the star tracker to point at the sun for a time, the spacecraft can continue to update its orientation based on other sensors like the gyroscope, with the last viable star tracker measurement acting as a high accuracy reference point. This also provides some resilience to bias and drift among sensors like the gyroscope and magnetometer. More info below.

SGP4 for Orbital Position Estimation

LI Bus will run an SGP4 orbit propagator alongside the attitude estimator for orbital position estimation to aid in pointing. This allows for orbital information to be polled at any time internally if GNSS updates are unavailable. SGP4 remains a popular orbit propagator due to its purely analytical method. Research and code for the propagator is also widely available.[1] SGP4 has been found to be accurate to about a half-kilometer in the tangential direction during propagation.[2] This is likely not accurate enough for most imaging purposes, but is sufficient for helping align with ram and the sun while in science mode. 

Another advantage of this method is that the propagator epoch can be updated using two different sources. A TLE can be uploaded from the ground, or the CDH can convert the last GNSS update into a TLE with an updated epoch.  All satellites are tracked by the United States Space Force and regular updates can be found on Space-Track.org and CelesTrak. 

RSO-based Navigation

LI Bus will also have the ability to test RSO-based navigation as an additional navigation tool. This would utilize the star tracker to use other objects in orbit as reference points to determine orbital position and orientation. More detailed information can be found below.

Estimation Results

Operating Mode Philosophy

This section describes the high level AODCS philosophy for LI Bus's operational modes. LI Bus will include three operating modes for nominal operations:

• Detumbling

• Sun Pointing

• Orbital Maneuver

• Science

Detumbling Mode

The purpose for this mode is to detumble the spacecraft following deployment in orbit, but can also be commanded at any time in the mission to reduce excessive body rates. This mode uses a simple control law (figure below) to calculate the magnetic dipole needed to reduce the rate of change of the Earth's magnetic field. The IMU will be the primary sensor used. The IMU's magnetometer will be used to calculate the rate of change of the Earth's magnetic field in the spacecraft's body frame. The IMU's gyroscope will be used to monitor the spacecraft's rotational rate as detumbling progresses. The spacecraft will exit detumbling mode automatically once acceptably low, or if a maximum time of 8 hours (~5 orbits) is elapsed. 

Sun Pointing

In sun pointing mode, the primary attitude determination sensor will be the sun sensors to measure the angle of the sun directly. This way, as long as the spacecraft is not in eclipse, the spacecraft can orient to point its solar panels to the sun without the need for accurate orbital position knowledge. This way, pointing accuracy to the sun can be maintained if the GNSS and star tracker modules are turned off, as is the case in low power mode.

Orbital Maneuver Mode

Orbital maneuver mode is intended for using the thruster. This mode will point the thruster nozzle in the intended direction, and negate the body torque caused by the thruster to maintain pointing in the correct direction.

Science Mode

Science mode will be the main operating mode for LI Bus in orbit. Nominal pointing will point the payload in either the ram or anti-ram direction. The spacecraft will rotate in the payload imager axis to align the solar panels with the sun as best as it can in this pointing orientation. Updates to the pointing direction will come with each update of the attitude estimator, which will help provide smooth, accurate pointing for payload imaging. The spacecraft can also be commanded to point elsewhere given an inertial pointing quaternion.

 The high level operating philosophy is shown in the flowchart below.

Components and Requirements Compliance

The ADCS subsystem for LI Bus includes many components in order to control the spacecraft and meet mission requirements. This section will detail all of the components included in the subsystem, and their requirements compliance. Please find the high level subsystem compliance below.

AODC Control Board

The control board acts mainly as an interface board for most AODC components. It will also serve as a mechanical mounting point for the inertial measurement unit and the magnetorquers. The parts list for the board can be found below.

The AODC control board uses automotive grade components wherever possible in its design. However, many of the passive components on the board are commercial or industrial grade parts. Critical components like the STM32 ARM processor, the BNO055 IMU, the 3.3 V linear voltage regulator, and the DC/DC buck converter are also commercial or industrial grade.

Temperature:

• All parts on the AODC board comply with the given temperature requirements.

Vibration:

• Many of these parts were included on the AODC board for MBSAT-1, which passed vibration testing. Similar commercial grade components like the gyroscope and magnetometers also passed, so we do not expect any different from the components that were replaced in this iteration of the board.

Radiation:

• STM32 Processor: Test results reported in literature shows that the commercial grade STM32F1 family microcontrollers have a failure total ionizing dose of around 107 krad which is higher than the expected 20 krad(Si) dose in 10 years based on the SPENVIS analysis. Since it is in the same family, it is believed that the STM32F0 series will behave similarly.

• BNO055: An investigation on another COTS MEMS IMU found that although sensor drift was present during irradiation from a proton beam, the sensors suffered no permanent effects even after the maximum TID of 50 krad(Si) was reached.  Given that the TID from the SPENVIS analysis was found to be around 20 krad(Si) over 10 years, the IMU should perform adequately.

• 3.3V Linear Voltage Regulator: One study on the TID of COTS linear voltage regulators showed that many ceased functioning after 10 krad(Si). Since this is the 5 year dose in the SPENVIS analysis, the team is currently investigating whether this part would be suitable as a flight model. We have reached out to the CSA for further assistance.

• TPS82140SILR DC/DC Converter: A study using a similar Texas Instruments buck converter found that it could sustain about 40 krad before experiencing damage. Given that the TID from the SPENVIS analysis was found to be around 20 krad(Si) over 10 years, this should not be a problem.

Phase C requirements compliance can be found in the table below.

Reaction Wheel Cluster

LI Bus will use a NanoAvionics 4RW0 reaction wheel cluster to provide fine pointing for mission operations. A figure of the cluster and the requirements compliance for this component can be found below.

Magnetorquers

Magnetorquers will be made by STAR Lab, and are the same design supplied by York University for the Iris mission. LI Bus will use 3 orthogonally oriented magnetorquers. The primary functions of the magnetorquers will be for magnetic detumbling and reaction wheel desaturation.

Thruster

STAR Lab is developing its own cold gas thruster to be used by LI Bus. The purpose of the thruster is to demonstrate LI Bus's ability to make adjustments to its orbit during flight. For more information on the thruster, see its dedicated wiki page.

Gyroscope

LI Bus will use a STMicroelectronics A3G4250D gyroscope. It is an AEC-Q100 qualified sensor mounted to the AODC Control Board. It features a full-scale range of ±245 dps, with a typical sensitivity of 8.75 mdps, and a 32 slot FIFO buffer to store measurements.

Magnetometer

LI Bus will use a Memsic MMC5983 high performance magnetometer. It uses anisotropic magnetoresistive technology for high sensitivity, and features AEC-Q100 qualification for reliability and a SET/RESET function to reduce sensor drift.

Sun Sensor

The sun sensors that LI Bus will use are based on the ManitobaSat Onboard LEO-Tested Raised External Sun Sensor (MOLTRES) from York University, the same model used on the Iris mission. The sun sensor consists of a PCB with two Melexis MLX75306 linear optical arrays mounted perpendicularly, and an aluminum mask with two thin slits to let light hit the sensors. 

The sun sensor's total field of view is 60 degrees, and the minimum angular resolution is 0.0242 degrees.  For the Iris mission, York University was able to calibrate to <1 degree accuracy. STAR Lab will recreate this setup and procedure with provided documentation from York University.

GNSS Receiver

LI Bus will use a NovaTel OEM719H-WFN-LNN-TMN-H GNSS receiver. Its features include GPS+Galileo+QZSS, L1/L2/L5/L6/E1/E5a/E5b/AltBOC/E6, SBAS L1/L5 Single Point+DGPS PNT, 20 Hz Data Output Rate, High Speed Includes GLIDE & RAIM. The flight model is COCOM removed and considered Controlled Goods, the engineering model is a COCOM enabled OEM719-WFN-LNN-TMN.

Development resources for the receiver can be found below.

• https://docs.novatel.com/OEM7/Content/Appendix/PDFDocuments.htm?tocpath=Support%7CPDF%20Documents%7C_____0

• https://docs.novatel.com/OEM7/Content/Core_Installation/OEM7_Dev_Kit.htm

GNSS Antenna

LI Bus will use an Antcom 1.2G-mini-GNSS-AS4 GNSS Antenna. This model has flight heritage, and has a small footprint (37x32 mm) to mount to the CubeSat structure.

Star Tracker

LI Bus will use a Rocket Lab ST-16HV star tracker. Requirements compliance for this component can be found below.