Thesis (MEng (Mechanical Engineering))--Cape Peninsula University of Technology, 2019 / The accumulating space debris has been a developing problem for many years. Technological advances led to the creation of nanosatellites, which allows more affordable access to space. As a result, the number of satellite launches is rapidly increasing, which, translates into an increase in debris in the low earth orbit (LEO) and geostationary orbit (GEO). To comply with the Inter-Agency Space Debris Coordination Committee (IADC) requirement of a 25-year maximum orbital lifetime, nanosatellites must have an end of life strategy. Failure to meet these guidelines may not only cause catastrophic collisions but may make future space travel even more challenging. Consequently, orbital lifetime predictions must be completed for nanosatellites. Considering this, the aim of this thesis is to investigate the orbital lifetime predictions for the nanosatellite ZACube-2, and the effects on the orbital lifetime if ZACube-2 is fitted with deorbiting technology, specifically a drag argumentation device. An in-depth literature review regarding the current state of technology pertaining to nanosatellite de-orbiting was conducted. This was followed by studies regarding orbital dynamics and perturbation forces. Four case studies were simulated in NASA’s Debris assessment software (DAS 2.0) using orbital parameters extracted from the two-line element (TLE) file. General information such as launch date and final mass was provided by F’SATI. The Baseline case study presented the orbital lifetime of ZACube-2, without any drag enhancement device. This was followed by case study 1,2 and 3 which represented ZACube-2 when fitted with three different drag enhancement devices. A comparison study indicated a reduction in all three cases. A new inflatable cube de-orbiting device (ICDD) concept was also presented, and the effects it has on the orbital lifetime predictions are showcased in case study three. Two deployment concepts were considered and evaluated against design requirements. Solidworks software was used to model the most suitable concept as well as perform finite element analysis on the structure. Static analysis was followed by natural frequency analysis in which the natural frequencies of the components and assembled structure were extracted. The Soyuz launch vehicle’s sinusoidal testing requirements were used to evaluate the structures survivability under dynamic loading. Based on the finite element , and harmonic analysis it was concluded that the structures will survive the launch conditions of the Soyuz launch vehicle. Furthermore, individual parameters affecting orbital lifetime predictions are also identified, in the form of a mass and cross-sectional sensitivity study and a ballistic coefficient versus orbital time study.

CubeSats have become popular among universities, research organizations, and government agencies due to their low cost, small size, and light weight. Their standardized configurations further reduce the development time and ensure more frequent launch opportunities. Early cubesat missions focused on hardware validation and simple communication missions, with little requirement for pointing accuracy. Most of these used magnetic torque rods or coils for attitude stabilization. However, the intrinsic problems associated with magnetic torque systems, such as the lack of three-axis control and low pointing accuracy, make them unsuitable for more advanced missions such as detailed imaging and on-orbit inspection. Threeaxis control in a cubesat can be achieved by combining magnetic torque coils with other devices such as thrusters, but the lifetime is limited by the fuel source onboard. To maximize the mission lifetime, a fast attitude control management algorithm that could optimally manage the usage of the magnetic and thruster torques is desirable. Therefore, a recently developed method, the BSpline-augmented virtual motion camouflage, is presented in this defense to solve the problem. This approach provides results which are very close to those obtained through other popular nonlinear constrained optimal control methods with a significantly reduced computational time. Simulation results are presented to validate the capabilities of the method in this application

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2009. / The nanosatellite is build-up of subsystems and payloads (defined as satellite nodes) connected together into the OBC using CAN bus as the main communication protocol. The flight software application is required to run within the eCos environment on the OBC to monitor and control satellite nodes. The ground station must generate commands and send them to the satellite in space. The application is developed to validate, schedule and dispatch the commands to the satellite nodes at appropriate times. Each node manager, in the flight software, is required to execute the response messages from its respective satellite node. The housekeeping and error recovery data files are defined to convey useful information about satellite status to the user and can be downloaded to the ground station. The flight software is developed using POSIX functions supported by eCos. Although it is not yet ready for real operation in space, the algorithm that can be used for full development is examined and approved.

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2006. / This thesis relates to the design of a camera system for a nanosatellite based on a CMOS image sensor. The design specifications and constraints are considered followed by the proposal of a versatile design with all the required functions imple- mented on a single FPGA. These functions include bad block management, data routing, an EDAC, a soft-core processor, glue logic to external devices, and com- munication busses. The Altera Nios II soft-core processor is implemented in this design, which en- ables simple changes to be made in software. A good mixture of intellectual prop- erty soft-cores, open-source cores, and user created logic are utilised in this broad base design, containing a combination of hardware, digital logic, and software. Low power and compact devices are selected for this design to minimize the power usage and the physical size of the camera system. The system's peak power consumption is 952mW which is below the required maximum consumption of 1W. This design's performance is therefore ideal for a subsystem onboard a nanosatel- lite.

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2007. / Next-generation nanosatellites are becoming a very cost effective solution to gain access to space. Modern manufacturing technology together with low power low cost devices makes the development of nanosatellites, using standard industrial components, very attractive. A typical nanosatellite will have only one microprocessor, capable of performing all the computing tasks onboard the satellite - housekeeping, AODC (Attitude and Orbit Control) and instructing the different payloads aboard the satellite. One of the major requirements was to choose a processor from a dominant manufacturer in the market that will still be available for future satellite missions. Just as the 8051 dominated the 8-bit market, the ARM7 processor is fast becoming a market leader in the segment for 16-bit applications. ARM processors has also been used much in handheld devices in recent years - which emphasize the low power requirements and stability of these processors in embedded applications. This thesis investigates the different processors that are currently available. A complete system design is done, taking into account all the different modules needed onboard a very small Low Earth Orbit (LEO) satellite. Finally, some test results are given showing how this system can be reliably used onboard a nanosatellite in future.

Thesis (Master of Engineering in Electrical Engineering)--Cape Peninsula University of Technology, 2017. / The concept of forward error correction (FEC) coding introduced the capability of achieving near Shannon limit digital transmission with bit error rates (BER) approaching 10-9 for signal to noise power (Eb/No) values as low as 0.7. This brought about the ability to transmit large amounts of data at fast rates on bad/noisy communication channels. In nano-satellites, however, the constraints on power that limit the energy that can be allocated for data transmission result in significantly reduced communication system performance. One of the effects of these constraints is the limitation on the type of channel coding technique that can be implemented in these communication systems. Another limiting factor on nano-satellite communication systems is the limited space available due to the compact nature of these satellites, where numerous complex systems are tightly packed into a space as small as 10x10x10cm. With the miniaturisation of Integrated-Circuit (IC) technology and the affordability of Field-Programmable-Gate-Arrays (FPGAs) with reduced power consumption, complex circuits can now be implemented within small form factors and at low cost. This thesis describes the design, implementation and cost evaluation of a ½-rate convolutional encoder and the corresponding Viterbi decoder on an FPGA for nano-satellites applications. The code for the FPGA implementation is described in VHDL and implemented on devices from the Artix7 (Xilinx), Cyclone V (Intel-fpga), and Igloo2 (Microsemi) families. The implemented channel code has a coding gain of ~3dB at a BER of 10-3. It can be noted that the implementation of the encoder is quite straightforward and that the main challenge is in the implementation of the decoder.

Future spacecraft missions are trending towards the use of distributed systems or fractionated spacecraft. Initiatives such as DARPA's System F6 are encouraging the satellite community to explore the realm of collaborative spacecraft teams in order to achieve lower cost, lower risk, and greater data value over the conventional monoliths in LEO today. Extensive research has been and is being conducted indicating the advantages of distributed spacecraft systems in terms of both capability and cost. Enabling collaborative behaviors among teams or formations of pico-satellites requires technology development in several subsystem areas including attitude determination and control subsystems, orbit determination and maintenance capabilities, as well as a means to maintain accurate knowledge of team members' position and attitude. All of these technology developments desire improvements (more specifically, decreases) in mass and power requirements in order to fit on pico-satellite platforms such as the CubeSat. In this thesis a solution for the last technology development area aforementioned is presented. Accurate knowledge of each spacecraft's state in a formation, beyond improving collision avoidance, provides a means to best schedule sensor data gathering, thereby increasing power budget efficiency. Our solution is composed of multiple software and hardware components. First, finely-tuned flight system software for the maintaining of state knowledge through equations of motion propagation is developed. Additional software, including an extended Kalman filter implementation, and commercially available hardware components provide a means for on-board determination of both orbit and attitude. Lastly, an inter-satellite communication message structure and protocol enable the updating of position and attitude, as required, among team members. This messaging structure additionally provides a means for payload sensor and telemetry data sharing. In order to satisfy the needs of many different missions, the software has the flexibility to vary the limits of accuracy on the knowledge of team member position, velocity, and attitude. Such flexibility provides power savings for simpler applications while still enabling missions with the need of finer accuracy knowledge of the distributed team's state. Simulation results are presented indicating the accuracy and efficiency of formation structure knowledge through incorporation of the described solution. More importantly, results indicate the collaborative module's ability to maintain formation knowledge within bounds prescribed by a user. Simulation has included hardware-in-the-loop setups utilizing an S-band transceiver. Two "satellites" (computers setup with S-band transceivers and running the software components of the collaborative module) are provided GPS inputs comparable to the outputs provided from commercial hardware; this partial hardware-in-the-loop setup demonstrates the overall capabilities of the collaborative module. Details on each component of the module are provided. Although the module is designed with the 3U CubeSat framework as the initial demonstration platform, it is easily extendable onto other small satellite platforms. By using this collaborative module as a base, future work can build upon it with attitude control, orbit or formation control, and additional capabilities with the end goal of achieving autonomous clusters of small spacecraft.

M. Tech. Electrical Engineering / Deals with nonlinear methods for magnetic attitude control and reaction wheel attitude control. The work is divided into a number of parts. The first part, deals with the satellite attitude control basic information and development of a mathematical model of a low Earth orbit satellite. The second part introduces the controllers used in this dissertation. The third part deals with the dimension between the output of controller and input of reaction wheel. The fourth part solves the problem of the magnetic torque calculation. The last part carries out the simulation tests of those controllers for small satellite and cube satellite.

M. Tech. Electrical Engineering / The topic of this dissertation is the attitude control of a CubeSat in an elliptic orbit using nonlinear control. The attitude control system (ACS) is a subsystem of a CubeSat. Its principal goal is to stabilise the orientation of the satellite after launch and during the orbital motion of the satellite. Although several methods have been applied to achieve this objective, this still remains a challenging objective and hence plays an integral role in many modern technologies. CubeSat "Cube Satellite" is a miniaturised satellite which, due to its low cost and application potential is often used by academic institutions for research purposes. However, due to its physical size and weight of 1 kilogram, CubeSat have comparatively limited power supply and computational resources; hence the need for an uncomplicated and reliable control system is critical.

The CanX-4 and CanX-5 mission currently under development at The University of Toronto Institute for Aerospace Studies Space Flight Laboratory UTIAS/SFL is a challenging formation flying technology demonstration. Its requirements of sub-metre control accuracy have yet to be realized with nanosatellites. Many large technical challenges must be addressed in order to ensure the success of the CanX-4/5 mission. This includes the development of software for an intersatellite communication system, integration and optimization of key formation flying algorithms onto the Payload On-Board Computer as well as the development of a Hardware-In-The-Loop simulator for full on-orbit mission simulations. This thesis will provide background knowledge of the Space Flight Laboratory and its activities, the CanX-4/5 mission, and nally highlight the authors contributions to overcoming each of these technical challenges and ensuring the success of the CanX-4 and CanX-5 mission.