1. Introduction

As part of a multidisciplinary spacecraft design project, our team developed the conceptual design of a 2U CubeSat mission intended to measure thermospheric CO₂ concentration, temperature, and atmospheric density. The spacecraft also served as a technology demonstrator for deployable solar arrays used for passive attitude stabilisation, a 3D-printed primary structure, and shape-memory-alloy deployment mechanisms.

The mission was motivated by climate-change predictions suggesting that rising CO₂ concentrations should cool the thermosphere and reduce its density (a trend NASA’s TIMED mission had previously attempted to verify by comparing atmospheric data from 2002 and 2008). The CubeSat concept we designed was intended to contribute further data toward correlating these climatic models.

The project followed a concurrent engineering approach, requiring the close integration of payload, power, thermal, communications, orbital analysis, and attitude control considerations throughout the design process.

My responsibility was the complete design of the Attitude Determination and Control System (ADCS), including pointing requirements definition, disturbance analysis, sensor and actuator selection, and the evaluation of a distinctive passive aerodynamic stabilisation concept.

2. The Challenge: Requirements

The scientific payload required the spacecraft to maintain its longitudinal axis aligned with the orbital velocity vector in order to obtain meaningful atmospheric measurements. At the same time, the spacecraft had to maximise power generation from its solar arrays while operating within the severe mass, volume, and power constraints of a 2U CubeSat.

This translated into two distinct pointing requirements: precise alignment of the roll axis with the velocity vector for the payload, and alignment of the spacecraft’s Z+ axis with the Earth vector so that only the nadir-facing panel (the one face without solar cells) pointed towards the planet, maximising the sunlit solar array area.

One of the first tasks was defining realistic pointing requirements together with the payload team, establishing the general orientation and error margin needed for the mission. Separately, after consulting a specialist in mass spectrometry to clarify the payload’s operating principles, the original 5° pointing accuracy requirement was tightened to 1°, which became one of the primary design drivers for the entire ADCS architecture.

3. Attitude Determination

Accurate attitude knowledge was essential not only for spacecraft pointing but also for calculating atmospheric density from orbital decay measurements. The proposed sensor suite was selected for both accuracy and redundancy and included:

• A Fine Sun Sensor for high-precision Sun vector measurements.
• An Earth Horizon Sensor for Earth vector determination.
• GPS position data for absolute attitude reconstruction.
• A three-axis magnetometer as an independent, redundant attitude reference.
• MEMS gyroscopes and accelerometers for attitude propagation during orbital eclipse periods.
• A coarse Sun vector estimation method using solar array power output measurements.

The Fine Sun Sensor and Earth Horizon Sensor combination allowed attitude determination with a resolution of approximately 0.2°, comfortably exceeding the mission’s pointing requirements. The magnetometer, combined with GPS data through the International Geomagnetic Reference Field model, gave a coarser cross-check with an accuracy on the order of a few degrees.

3.1. Disturbance Analysis

A critical part of the design process was determining the external torques acting on the spacecraft and quantifying their relative importance. This analysis directly influenced the sizing of the control actuators.

I evaluated the four major disturbance sources and estimated their order of magnitude at the CubeSat’s operating altitudes:

• Aerodynamic drag: ranging from approximately 1,000 nNm at 400 km (beginning of life) up to 100,000 nNm at 180 km (end of life).
• Gravity gradient effects: on the order of 10 nNm.
• Solar radiation pressure: approximately 65 nNm.
• Residual magnetic torques: approximately 24 nNm.

The results showed that aerodynamic disturbances were by far the dominant effect throughout the mission lifetime, two to four orders of magnitude larger than the remaining disturbances depending on altitude. This maximum combined disturbance torque, around 0.3 mNm, drove the sizing of the actuators.

This finding ultimately motivated the investigation of aerodynamic stabilisation as a passive attitude control mechanism.

3.2. Passive Aerodynamic Stabilisation

A particularly distinctive aspect of the project was the study of a passive stabilisation concept inspired by the behaviour of a shuttlecock.

The spacecraft incorporated four deployable fins that acted as aerodynamic control surfaces; three of the four also doubled as solar arrays, with the fourth (the nadir-facing fin) left without solar cells so that it could satisfy the Z+ Earth-pointing requirement described earlier. By deploying the panels at an angle, the centre of pressure was shifted away from the centre of mass, generating a restoring aerodynamic moment whenever the spacecraft deviated from its nominal velocity-vector orientation.

This concept allowed the spacecraft to naturally align itself with the velocity vector, reducing the burden on active control elements and minimising power consumption.

Several deployment angles were analysed to weigh the trade-off between:

• Aerodynamic stability.
• Pointing authority.
• Orbital lifetime.

A deployment angle of 45° was selected as the final configuration. It delivered the strongest restoring moment while keeping orbital lifetime within acceptable limits. Simulations performed by the orbital analysis team, using the STK HPOP propagator, estimated a mission lifetime of approximately 152 days (0.42 years) at this deployment angle. This result was cross-checked using an independent lifetime tool, which produced a comparable estimate of 0.53 years, giving confidence in the model.

3.3. Control Actuation

While passive stabilisation handled most of the low-frequency disturbance rejection, active actuators were still required for fine pointing and momentum management.

The final architecture combined:

• Three orthogonal reaction wheels, each storing up to 1.77 mNms of momentum before saturating (around 7.4 seconds of continuous actuation at nominal torque), for precise attitude control.
• A three-axis magnetorquer system for detumbling after deployment and reaction wheel momentum dumping.

The reaction wheels provided sufficient torque and momentum storage to perform attitude manoeuvres and maintain the required pointing precision, while the magnetorquers offered a reliable and low-complexity mechanism to unload accumulated momentum using Earth’s magnetic field.

4. Supporting Analysis and Simulation

Beyond the ADCS design itself, I also developed the spacecraft geometry used for solar power generation simulations.

A simplified CAD model was prepared and exported to STK through a Blender-based workflow. Solar cell groups were defined individually so that power generation from each spacecraft face could be analysed throughout the orbit. The resulting model was used by the power subsystem team to generate detailed power-production estimates for the deployable solar array configuration.

5. Key Takeaways

Looking back, the most valuable aspect of the project was not any individual ADCS calculation, but the exposure to true systems engineering.

Every design decision had consequences across multiple subsystems. Increasing solar array deployment angles improved aerodynamic stability but reduced orbital lifetime. Tightening pointing requirements affected sensor selection, power consumption, and actuator sizing. Mass distribution influenced both structural design and attitude dynamics.

This was my first experience working on a genuinely multidisciplinary spacecraft system where performance emerged from the interaction of many interconnected subsystems.