This thesis presents the design of a working and testable satellite with particular emphasis on the electrical, mechanical, and thermal modelling and performance issues for the ECOSat project in the framework of the Canadian Satellite Design Competition. In order of importance, based on the design challenges for the satellite structure were the dynamics modelling and analysis, thermal modeling and analysis, and assembly and integration modeling. Both the dynamics and thermal modeling of the satellite were completed using Finite Element Analysis (FEA) in NX with the NASTRAN solver. The dynamic analysis study was performed first since it has the primary design driver for the structure. These frequencies are of concern due to the 90 Hz or greater fundamental frequency requirement for each axis. The dynamic modes of the satellite structure had the largest influence not only on the design of the structure but also its interface to the electronic systems as these had to meet the required testing qualification levels. It was found that the first fundamental frequency appeared near 200 Hz in the XY plane of the structure. The second study performed was on the thermal modeling of the satellite both for extreme operating conditions in “Hot” and “Cold” cases. Operational limiting cases were identified for the batteries in the cold and hot case study, and the power amplifier for the transmitter was identified for the hot case study. For the batteries to perform satisfactorily for the cold and hot case problem, a metal bracket with an electric heater was added to the design. The heaters were added to the design as a resistive heating element, the additional thermal coupling from the bracket improved heat transfer during the hot case. A trade study analysis was conducted for the power amplifier. Here, a bi directional heat spreader made of pyrolytic graphite attached to a frame member with high thermal inertia was chosen as the optimal solution. Finally, the third study performed tested the interface and clearance requirements of the satellite. The synergistic integration of the electrical and mechanical systems required significant attention in order to ensure the successful assembly, integration, and testing of the two systems. The investigation focused on the cabling assemblies of the satellite. Several design iterations were required for the power regulation, transmitter, receiver, modem, and onboard computer systems. Detailed assembly drawings were created for the cabling assembly fabrication prior to the final integration of the electrical and mechanical systems. The performance simulations show that the satellite systems meet or exceed the required launch qualification tests as well as the thermal cycling requirements for all systems and their components to operate within the manufacturer specified values. Once completely assembled and launched into orbit, the satellite should be able to perform and within its operational and mission requirements in both a sun synchronous or polar orbit at a range of altitudes. / Graduate / 0538 / 0544 / 0548 / jtcurran@uvic.ca

International Telemetering Conference Proceedings / October 17-20, 1994 / Town & Country Hotel and Conference Center, San Diego, California / There has been much interest recently in the possibility of using the NASA Tracking and Data Relay Satellite System (TDRSS) instead of proprietary ground stations in supporting small space payload communications. These payloads operate on fairly low power and do not use the sophisticated tracking equipment standard on more complex user spacecraft. This paper is part of a feasibility study for such use of TDRSS, and focuses on the effect of the method of providing the Grade-2 data link layer services specified in Consultative Committee for Space Data Systems (CCSDS) Advanced Orbiting Systems (AOS) recommendations upon a hypothetical spacecraft using a MIL-STD-1553B polled data bus as the backbone of its onboard LAN. In particular, one case in which the 1553B bus controller, assumed to be some Intel 80X86 microprocessor, provides all CCSDS services will be contrasted with another where these services are split between the bus controller and a device which interfaces between the spacecraft LAN and the TDRSS Return Service spacelink. The comparison will be made for a 15 orbit/day scenario using a small helical antenna with a comparatively wide beamwidth. The main performance criteria considered here are end-to-end data throughput and expected delays, along with required buffer sizes for the LAN. Also, it may be noted that the data rate of the TDRSS return link and the size of the sliding window used for flow and error control will have a large impact on the required values for the chosen criteria, and so choices for these parameters significantly affect the outcome of any system service comparison. The two LAN types will be modeled and analyzed using NETWORK II.5. This simulator allows tracking of the number of packets read into LAN memories and sent down from the payload to ground via TDRSS, as well as accurately modeling the delays involved with data processing and transmission over the link.

Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.

Risk management plans improve the likelihood of mission success by identifying potential failures early and planning mitigation methods to circumvent any issues. However, in the aerospace industry to date, risk management plans have typically only been used for larger and more expensive satellites, and have rarely been applied to satellites in the shape of 10 x 10 x 10 centimeter cubes, called CubeSats. Furthermore, existing risk management plans typically require experienced personnel and significant time to run the analysis. The purpose of this research was to develop two risk management software tools, the CubeSat Risk Analysis tool and the CubeSat Decision Advisor tool, which could be used by anyone with any level of experience. Moreover, the tools simply require the user to enter their mission-specific data; the software tools calculate the required analysis. The CubeSat Risk Analysis tool was developed for the purpose of reducing the subjectivity associated with estimating the likelihood and consequence of spacecraft mission risks. The tool estimates mission risk in terms of input characteristics, such as satellite form factor, mass, and development cycle. Using a historical database of small satellite missions, which was gathered in the course of this research, the software determines the mission risk root causes which are of the highest concern for the given mission. The CubeSat Decision Advisor tool uses components of decision theory such as decision trees, multi-attribute utility theory, and utility elicitation methods to determine the expected utility of a mitigation technique alternative. Based on the user’s value preference system, assessment of success probabilities, and resources required for a given mitigation technique, the tool suggests the course of action which will normatively yield the most value for the cost, personnel, and time resources required. The goals of this research were met in the development of two easily-accessible and free risk management software tools to assist in university satellite mission development. But more importantly, these tools will reach beyond the academic setting and allow small satellites to continue to evolve as a platform to accomplish educational, scientific, and military objectives. / text

Scheduling of aperiodic and independent tasks in hard real-time symmetric multiprocessing systems is an NP-complete problem that is often solved using heuristics like particle swarm optimization (PSO). The performance of these class of heuristics, known as evolutionary algorithms, are often evaluated based on the number of iterations it takes to find a solution. Such metrics provide limited information on how the algorithm reaches a solution and how the process could be accelerated. This thesis presents a methodology to analyze the trajectory formed by candidate solutions in order to analyze them in both the time and frequency domains at a single scale. The analysis entails (i) the impact of different parameters for the PSO algorithm, and (ii) the evolutionary processes in the swarm. The work reveals that particles have a directed movement towards a solution during a transient phase, and then enter a steady state where they perform an unguided local search. The scheduling algorithm presented in this thesis uses a variation of the minimum total tardiness with cumulative penalties cost function, that can be extended to suit different system needs. The experimental results show that the scheduler is able to distribute tasks to meet the real-time deadlines over 1, 2, and 4 processors and up to 30 tasks with overall system loads of up to 50\% in fewer than 1,000 iterations. When scheduling greater loads, the scheduler reaches local solutions with 1 to 2 missed deadlines, while larger tasks sets take longer to converge. The trajectories of the particles during the scheduling algorithm are examined as a means to emphasize the impact of the behaviour on the application performance and give insight into ways to improve the algorithm for both space and terrestrial applications.

This thesis describes the design, implementation and testing of a spacecraft bus implemented on KySat-1, a picosatellite scheduled to launch late 2009 to early 2010. The spacecraft bus is designed to be a robust reusable bus architecture using commercially available off the shelf components and subsystems. The bus designed and implemented for the KySat-1 spacecraft will serve as the basis for a series of future Kentucky Space Consortium missions. The spacecraft bus consists of attitude determination and control subsystem, communications subsystem, command and data handling subsystem, thermal subsystem, power subsystem, and structures and mechanisms. The spacecraft bus design is described and the implementation and testing and experimental results of the integrated spacecraft engineering model. Lessons learned with the integration, implementation, and testing using commercial off the shelf components are also included. This thesis is concluded with future spacecraft bus improvements and launch opportunity of the implemented spacecraft, KySat-1.

Scheduling of aperiodic and independent tasks in hard real-time symmetric multiprocessing systems is an NP-complete problem that is often solved using heuristics like particle swarm optimization (PSO). The performance of these class of heuristics, known as evolutionary algorithms, are often evaluated based on the number of iterations it takes to find a solution. Such metrics provide limited information on how the algorithm reaches a solution and how the process could be accelerated. This thesis presents a methodology to analyze the trajectory formed by candidate solutions in order to analyze them in both the time and frequency domains at a single scale. The analysis entails (i) the impact of different parameters for the PSO algorithm, and (ii) the evolutionary processes in the swarm. The work reveals that particles have a directed movement towards a solution during a transient phase, and then enter a steady state where they perform an unguided local search. The scheduling algorithm presented in this thesis uses a variation of the minimum total tardiness with cumulative penalties cost function, that can be extended to suit different system needs. The experimental results show that the scheduler is able to distribute tasks to meet the real-time deadlines over 1, 2, and 4 processors and up to 30 tasks with overall system loads of up to 50\% in fewer than 1,000 iterations. When scheduling greater loads, the scheduler reaches local solutions with 1 to 2 missed deadlines, while larger tasks sets take longer to converge. The trajectories of the particles during the scheduling algorithm are examined as a means to emphasize the impact of the behaviour on the application performance and give insight into ways to improve the algorithm for both space and terrestrial applications.

The magnetic cleanliness of a spacecraft during magnetic measurements is an important aspect in the design of many space science missions. The adequate reduction or removal of the spacecraft's magnetic disturbance plays a vital role in allowing the ambient magnetic field to be measured with the required accuracy. There are three main approaches to reduce the impact of the spacecraft's magnetic disturbance on the final magnetic measurement, with each approach imposing its own set of changes and constraints on the spacecraft. In turn these changes and constraints introduce additional complexity and cost to the system design. The required changes in the spacecraft's mission profile and configuration also need to be factored in during the design phase of a spacecraft, but cannot be avoided if high quality measurements are desired.  One of these approaches is the use of a magnetic cleanliness programme, and such programmes have a long history of successful use on large satellite missions.CubeSats have become increasingly technically capable and have in recent years begun to undertake scientific missions with challenging sensitivity requirements, including for magnetic measurements. The small size of the CubeSat form factor poses some unique challenges to the implementation of magnetic cleanliness techniques, but are also in increased need of limiting the residual magnetic moment when compared to large satellites. This thesis details the early phases of the magnetic cleanliness programme on the FORESAIL-2 science mission. Nine magnetic cleanliness requirements on the FORESAIL-2 satellite platform were derived from the FORESAIL-2 measurement and instrument requirements. A simple magnetic model was established, the results of which were used to propose a configuration of spacecraft subsystems. The resulting preliminary configuration of subsystems reduced the effective magnetic field of the REPE payload at the sensor by 352pT, 30.1%, when compared to the worst-performing configuration. Subsequently an improved model, utilising RSS analysis, was created. Combined with updated location information for each subsystem, defined using the proposed configuration, this second model yielded an estimated magnetic field of 2710pT at the reference point. The results of the second model were also used to identify the CDE payload and the TT&C subsystem as potentially problematic from a magnetic cleanliness perspective. A list of ferromagnetic materials was compiled, and a total mass of 453.72g of ferromagnetic materials was estimated. The work presented in this thesis is expected to form the basis of FORESAIL-2's continued magnetic cleanliness programme throughout the design and integration phases of the project.  For example, the improved model could be extended to include the estimated magnetic dipole moment's of each subsystem. Additionally, the inventories of materials, currents, and frequencies established as part of this thesis can be maintained throughout the FORESAIL-2's development cycle, and used to update the estimated total magnetic field of the spacecraft platform.

Lunar constellations providing the capabilities of situational awareness and surveillance for future mission operators are analyzed in this study. The use of specialty orbits, such as sun-synchronous and frozen orbits, are analyzed to determine the applicability of these unique orbits. Additionally, altitude and inclination trades are performed to determine the degree to which mission objectives are achieved through ranges of these orbital parameters. Using the analyzed orbits, constellations of varying patterns are developed and surface coverage figures of merit are used to evaluate them. The research concludes with calculation of the yearly cross-track and in-track stationkeeping costs of a representative constellation. This stationkeeping is necessary for preservation of the designed coverage statistics.

Once a satellite has completed its operational period, it must be removed responsibly in order to reduce the risk of impacting other missions. Geostationary Transfer Orbits (GTOs) offer unique challenges when considering disposal of spacecraft, as high eccentricity and orbital energy give rise to unique challenges for spacecraft designers. By leveraging small satellite research and integration techniques, a deployable drag sail module was analyzed that can shorten the expected orbit time of launch vehicle stages in GTO. A tool was developed to efficiently model spacecraft trajectories over long periods of time, which allowed for analysis of an object’s expected lifetime after its operational period had concluded. Material limitations on drag sail sizing and performance were also analyzed in order to conclude whether or not a system with the required orbital performance is feasible. It was determined that the sail materials and configuration is capable of surviving the expected GTO environment, and that a 49 m2 drag sail is capable of sufficiently shortening the amount of time that the space vehicles will remain in space.