Attitude Determination and Control System (ADCS)

Introduction

This section contains general ADCS design information. For mission-specific ADCS documentation, click the links below.

The ADCS subsystem serves two primary purposes: a) to orient the spacecraft’s solar panels towards the sun and b) to illuminate the payload samples with sunlight.

Simple three-axis control will be obtained using electromagnetic torque rods (developed by York University) for actuation. Dual-redundant three-axis chip-style magnetometers and gyros will supply magnetic field and attitude rate information.

The design incorporates two fine sun sensors and solar arrays. Two sun-sensors from York University will provide fine-pointing sun direction telemetry, while solar panel currents will provide a rough estimate of the sun direction in off-nominal attitudes where the sun is not pointed to the front face of the spacecraft.

The ADCS subsystem hardware interface will consist of the main microcontroller with an SPI interface to the CDH computer. This microcontroller is used for controlling and powering the sun sensors, sensors (magnetometers and gyros connected to microcontrollers via SPI interface) and magnetorquer. The following figures illustrate the functional block and system block diagrams for the ADCS subsystem.

ADCS attitude control strategy will be conducted mostly in a Matlab / Simulink simulation environment.

Deliverable Items

A deliverable is anything that the team has agreed to submit to the mission integrator as part of the legal obligations under the CSA requirements. These deliverables will be used to verify that the CubeSat meets the requirements set in the mission ICD. Here we describe the ADCS deliverables required.

Deliverables.xlsx

ADCS Hardware Block Diagram

Figure 1: ADCS Hardware Architecture including interfaces

Functional Block Diagram

The following depict the ADCS functional block diagram. The sensors are a part of attitude determination and the ADCS controller collect data from sensors and process it to get reliable positioning information. ADCS should communicate with both PCU in detumbling. CDH will collect the ADCS data after commissioning.

The ADCS interface board will be breadboarded and then built into an engineering model prior to the flight model.

ADCS Software Design

ADCS software architecture

A solar ephemeris model will relate the measured sun vectors to expected values (based on our orbital position) for validation.

A PD controller, along with a model of the Earth’s magnetic field will compute the required control torques and associated magnetorquer moments for maintaining adequate sun pointing. Note that at any given time in the mission, the spacecraft will have only two-axis control, corresponding to the two axes that are perpendicular to the earth’s magnetic field lines. Fortunately, this uncontrollable axis will change throughout the orbit, eventually enabling the ability to correct for the small disturbance torques we anticipate throughout the mission.

Detumbling Mode

Upon separation from the launch vehicle (and following the mandatory 30-minute hold period), the spacecraft will autonomously enter a detumble phase to remove residual rotation rates resulting from the separation springs. A standard B-Dot control law will apply magnetic dipoles proportional to the rate of change of the measured magnetic field. Since the magnetorquers and magnetometers must detumble the spacecraft prior to commissioning, these components will undergo special testing and characterization prior to launch.

PD Attitude Control Strategy

This section presents the PD controller designed for the Manitoba Sat-1 which uses magnetic actuators.

adcs_d_phil.pdf

Simulation Results-CDR PF comments.pdf

Pointing Budget

The sun sensors' accuracy should be 10 times better than the pointing requirement. The maximum satellite attitude tolerance is +-24 degrees and therefore the sun sensors pointing error should be less than +-2.4 degrees. In Iris the sun sensors provide an accuracy of 0.5 degrees which fulfill the pointing budget.

ADCS Hardware Design

The maximum height of the board including all components is 9.4mm +/- 0.1mm .

Parts List and Bill of Material

In the following the parts list is shown for both LOTAD(UMS-0054) and MOLTRESS(UMS-0380). Each part contains the component along with the part number, the manufacturer and bill of material.

ACS Project Control and Sun Sensor BoM Updated 10_07_2021.xlsx

Sun Sensors

York University is developing a novel low-cost sun sensor applicable for Nano-satellites with 0.1-degree accuracy and reliability. The current design consists of two orthogonal arrays of photodiodes with a pattern of slits (shown below) to allow greater precision in sensing and calibration. Micro-fabricated sun sensor will be first demonstrated on the Iris mission to showcase not only the novel concept of low-cost micro-sun sensor but also the micro-fabrication approach in developing satellite attitude sensors. The sun sensors run on an SPI connection and they are digital. Depending on the output you require, 8,4,2,1.5 and 1-bit output modes are available. The sun-sensors would be tied to the common ground bus. The specifications of the sun-sensors as suggested by York is mentioned on the right of the document. The survival and operating temperature ranges of the low-cost sun-sensors are [-50,+150 C] and [-40,+105 C].

Sun sensor Mask

Redundant Sensors

There are redundant sensors used to prevent failure from radiation and external torques. Double redundant sensors for the gyros and magnetometers are used. Similarly, two fine sun-sensors are utilized.

Torque Rods

Torque rods, provided by the York University team, are based on the rod concept described in the design page. The proposed design will be demonstrated on the SIGMA mission. The magnetorquers were controlled by PWM signals. For the PWM signals, we included h-bridge circuits. The torque rods will be tied to the common ground bus. The torque rods shall be mounted orthogonally. The maximum operational temperature ranges for the torque rods are 100 C. During the last mission done by York University they had mentioned about the same specification they would use for designing the rods for Iris mission which has a magnetic dipole of 0.1 Am Sq. During torque rod sizing we estimated the power consumption to be around 0.6 W. For each torque rod, a one to one mapping for Dipole vs. Power and Dipole vs. Current are shown in the following images.

Dipole Moment vs. Power

Dipole Moment vs. Current

Magnetometers

The MEMSIC, high performance and a low cost 3-axis magnetic sensors which has a field range of +/- 8 Gauss with 16 bits operation. This was chosen because it requires very less external components and has a dynamic range with a better accuracy. The RMS noise is 0.4 mG giving a heading accuracy of +/- 1°. Dimensions are 3.0×3.0×1.0 mm which is easy to fit in 3U form factor of the satellite. The main purpose for using this is that it clears the residual magnetization coming from the strong fields and it also has a built temperature sensor for monitoring the temperature throughout the mission. The power requirements are pretty low which suits for the CubeSat mission. It has a 1 µA of power down current and 2.5 V single low power supply. It has a voltage of 1.8 V with I2C interface.

Gyroscopes

The gyros used for the Iris mission are COTS having the model number A3G4250D from STMicroelectronics. They also include an adaptor board with them. These gyros are 3-axis, chip style and surface mounted on the CubeSat with the voltage range of 2.4 V to 3.6 V , eventually having a low power consumption. They work on the SPI/I2C interface and work on 16 bit rate data output.

Detailed Schematics

The LOTAD and MOLTRESS are designed as the ADCS hardware. LOTAD is referred to ADCS PCB(UMS-0054) and MOLTRESS(UMS-0380) points to sun sensor boards.

LOTAD Schematics.pdf

LOTAD(PCB) schematic

MOLTRESS Schematics.pdf

MOLTRES(sun sensor) schematic

ADCS Budget

Assembly Plan

The following document is the assembly plan including the board assembly, harnessing and mounting the board to the shell.

ADCS harnessing assemble.docx

Risk Registry

ADCS Risk registry.docx

Single point Failure Analysis

The single point failure(SPF) for ADCS is the failure of torque rods. In order to eliminate the single point failure, the torque rods shall be properly calibrated and tested through out Phase C and D according to the test plans. Regarding the high reliability of magnetorquers, we do not have any redundant actuators.

References

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