Dissertations / Theses on the topic 'Picosatellite'

The Poly-Picosatellite Orbital Deployer (P-POD) has undergone a series of revisions over the years. The latest revision, described in this Master’s Thesis, incorporates new capabilities like EMI shielding, an inert gas purge system, and an electrical interface to the CubeSats after they are integrated into the P-POD. Additionally, some mass reduction modifications are made to the P-POD, while its overall strength is increased. The P-POD inert gas purge system successfully flew, on a previous revision P-POD. The P-POD components are analyzed to a set of dynamic loads for qualification, and successfully undergoes random vibration qualification testing. The P-POD encounters some problems in thermal vacuum cycling qualification and EMI testing, but there is evidence that the issues can be mitigated. A path forward is laid out to complete both sets of testing.

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The military applications for miniature, low-cost satellites that could be quickly launched to provide ad-hoc tactical networks have risen in recent years. Currently, the smallest practical variant of these miniaturized satellites is known as the picosatellite. In order to evaluate the performance of the picosatellite constellation-based network, a model that can accurately simulate the orbital physics of the constellation as well as the satellite-to-ground communication links and data traffic is necessary. The focus of this thesis was to build such a model using commercially available software and assess the effects of orbital geometries on the performance of the picosatellite constellation-based network. The research revealed that orbital planes that were inclined near the latitude of the area of interest could provide better coverage. In addition, when the satellites were spaced farther apart in the orbital plane the constellation access times were also extended. This was at a cost, however, as the link quality could be compromised. The model that was created for this research could be integrated into the Naval Postgraduate School Tactical Network Topology testbed environment to study the extension of tactical networks to orbit and allow the modelling of picosatellite architectures applied to different maritime and inland missions.

This thesis outlines an orbit determination and 3-axis attitude determination system for use on orbit as applicable to 1U CubeSats and other picosatellites. The constraints imposed by the CubeSat form factor led to the need for a simple configuration and relaxed accuracy requirements. To design a system within the tight mass, volume, and power constraints inherent to CubeSats, a balance between hardware complexity, software complexity and accuracy is sought. The proposed solution consists of a simple orbit propagator, magnetometers with a magnetic field look-up table, Sun sensors with an analytic Sun direction model, and the TRIAD method to combine vector observations into attitude information. The orbit propagator is a simple model of a circular trajectory with several frequently updated parameters and can provide orbital position data with average and maximum errors—when compared to SGP4—of less than 3.7km and 10.7km for 14 days. The magnetic field look up table provides useful information from a small memory footprint; only 480 data points provide a mean error of approximately 0.2° and a maximum error of approximately 2°—when compared to the IGRF model. The Sun’s direction is modeled, and as expected, can be modeled simply and accurately. Combining the magnetic field and Sun direction models with inaccurate sensors and the TRIAD method results in useful attitude information from a very simple system. A system with Sun sensor error standard deviation of 1° and magnetometer error standard deviation of 5° yields results with average error of only 2.74°, and 99% of the errors in this case are less than approximately 13°. The system outlined provides crude attitude determination with software and hardware requirements that are well within the capabilities of current 1U CubeSats—something that many other systems, such as Kalman filters or star trackers, cannot do. It also provides an excellent starting point for future ADCS systems, which will significantly increase the ability of CubeSats.

International Telemetering Conference Proceedings / October 20-23, 2003 / Riviera Hotel and Convention Center, Las Vegas, Nevada
The NMSUSat is part of the AFRL/NASA University Nanosatellite program. The constellation will consist of a main microsatellite that will have a command link from ground and a telemetry link to ground while a picosatellite will act as a sensor reporting data to the microsatellite. Innovative command and data handling will be incorporated at low cost and greater accessibility. In this paper we present the necessary communications and control architecture for the space segment and the ground segment of the nanosatellite.

Small satellites, and CubeSats in particular, have quickly become a hot topic in the aerospace industry. Attitude determination is currently one of the most intense areas of development for these miniaturized systems and future Cal Poly satellite missions will depend heavily on magnetometers. In order to utilize magnetometers as a viable source of attitude knowledge, precise calibration is required to ensure the greatest accuracy achievable. This paper outlines a procedure for calibrating and testing magnetometers on the next generation of Cal Poly CubeSates, utilizing a Helmholtz cage to simulate any desired orbital magnetic field that would be experienced by a spacecraft around Earth, as well as investigation of magnetic interference as a result of on-board electrical activity.

Satellites are playing an increasingly important role in collecting scientific information, providing communication services, and revolutionizing navigation. Until recently satellites were large and very expensive, creating a high barrier to entry that only large corporations and government agencies could overcome. In the past few years the CubeSat project at California Polytechnic University in San Luis Obispo (Cal Poly) has worked to refine the design and launching of small, lightweight, and less expensive satellites called pico-satellites, opening space up to a wider audience. Now that Cal Poly has the launch logistics and hardware under control, a new problem has arisen. These pico-satellites are within communication range of a ground station only 40 minutes a day. This, combined with their 1200 bps communication speed, limits the usefulness of the satellite missions to those only transmitting small amounts of data back to Earth. This thesis proposes a novel protocol that allows a sparse network of pico-satellites to communicate among one another and to larger satellites called data mules, which relay the information back to the ground station at much higher speeds. The data mules are able to provide higher speeds because they are larger satellites with less power constraints. This protocol makes it possible for a pico-satellite to send more data over a given amount of time with less end-to-end delay. When every satellite has large amounts of data almost three times as much aggregate data can be sent through the network, and almost five times more data can be sent if only a single satellite has large amounts of data to send. The end-to-end delay is cut almost in half when sending 1 MB of data per day per satellite and is decreased by a factor of at least three when sending large amounts of data from only one satellite.

Small satellites are becoming increasingly important to the aerospace industry mainly due to their significantly reduced development and launch cost as well as shorter development time frames. In order to meet the requirements imposed by critically limited resources of very small satellites, e.g. picosatellites, innovative approaches have to be taken in the design of effective subsystem technologies. This thesis presents the design of an active attitude determination and control system for flight testing on-board the picosatellite 'Compass-1' of the University of Applied Sciences Aachen, Germany. The spacecraft of the CubeSat class with a net spacecraft mass of only 1kg uses magnetic coils as the only means of actuation in order to satisfy operational requirements imposed by its imagery payload placed on a circular and polar Low Earth Orbit. The control system is capable of autonomously dissipating the tumbling rates of the spacecraft after launch interface separ ation and aligning the boresight of the payload into the desired nadir direction within a pointing error of approximately 10°. This nadir-pointing control is achieved by a full-state feedback Linear Quadratic Regulator which drives the attitude quaternion and their respective rates of change into the desired reference. The state of the spacecraft is determined by a static statistical QUEST attitude estimator processing readings of a three-axis magnetometer and a set of five sun sensors. Linear Floquet theory is applied to quantify the stability of the controller and a non-linear dynamics simulation is used to confirm that the attitude asymptotically converges to the reference in the absence of environmental disturbances. In the presence of disturbances the system under control suffers from fundamental underactuaction typical for purely magnetic attitude control but maintains satisfactory alignment accuracies within operational boundaries.

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2011.
ENGLISH ABSTRACT: This thesis describes the development of attitude sensors required for the Attitude Determination and Control System (ADCS) for a Cubesat. The aim is to find the most suitable sensors for use on a small picosatellite by implementing miniaturised sensors with available commercial-off-the-shelf (COTS) technology. Specifically, the algorithms, hardware prototypes, software and filters required to create accurate sensors to determine the 3-axis orientation of a CubeSat are discussed.
AFRIKAANSE OPSOMMING: Hierdie tesis beskryf die ontwikkeling van oriëntasiesensors wat benodig word vir die oriëntasiebepaling en -beheerstelsel (Engels: ADCS) van ’n CubeSat. Die doelwit is om sensors te vind wat die geskikste is om in ’n klein picosatelliet te gebruik, deur miniatuursensors met kommersiële maklik verkrygbare tegnologie (Engels: COTS technology) te implementeer. Daar word in die bespreking veral aandag geskenk aan die algoritmes, hardewareprototipes, programmatuur en filters wat benodig word om akkurate sensors te skep wat op hul beurt 3-as oriëntasie van die CubeSat kan bepaal.

The CubeSat standard originated in 1999. It was a joint development led by Dr. Jordi Puig-Suari of California State Polytechnic University San Luis Obispo and Professor Robert Twiggs of Stanford University. The engineering challenges that have come from this picosatellite class have created incredible educational opportunities for engineering students throughout the world. Since the challenges of engineering a CubeSat abound the designers are always looking at novel and even revolutionary solutions to each one. One of those opportunities is in thermal subsystem design, implementation and characterization. A potential solution for CubeSats is adaptive component usage. This thesis is the written catalogue of my study of adaptive component utilization to solve the thermal management problem inherent in picosatellites. Inside the limited design space of a picosatellite’s electrical, mechanical and software subsystems active spacecraft thermal control often is a necessary forfeiture. This does not preclude CubeSat teams from addressing the thermal aspect of spacecraft design. To the contrary it forces them down a different route to ensuring the spacecraft is verified to meet appropriate environmental constraints. Most CubeSat teams, Cal Poly included, use punishing qualification testing, robust system design and a restricted spacecraft operational lifespan ensure their system will operate through all of the environments it will encounter during launch, separation, spacecraft activation and on until the end of operations. The testing, engineering and modeling I performed were to answer the hypothesis, can a standard* 1-U CubeSat utilize existing hardware and software to improve its thermal condition and operational lifetime? This hypothesis assumes thermal control or situational improvement would have to be gained without the addition of thermal control surfaces, active heaters, heat pipes or louvers and no additional flight software. Ground control software and operation alterations were explicitly not included in these assumptions. The thesis began with defining the many unknowns that existed in the material properties. This required: research into the methods required, specialized measurement hardware to be obtained and set-up, controlled measurements to be taken and thorough testing procedures to be developed. Once the unknowns were better defined the thesis required a detailed satellite thermal analysis by multiple methods along with both thermal vacuum chamber simulation trials and finally on-orbit testing. Based on the research, modeling and testing performed and results obtained through this study, yes, a standard* 1-U CubeSat utilizing existing hardware and software can improve its thermal condition and operational lifetime. As is shown in Section 3.0 and discussed in detail in Section 4.0, utilizing only the onboard electronics and existing flight software the orbital temperature delta that components are experiencing can be reduced by up to 35.8%. Further analyses in section 4.0 use the temperature data to show that by lowering the temperature deltas the satellite does in fact have the capability to both improve its lifetime and certain key subsystem performance parameters.

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.

碩士
國立中央大學
機械工程研究所
91
The research of picosatellite can let us to understand the requirements of space mission and help us to design a ‘smaller, cheaper, faster, and better’, satellite. We utilize a finite element analysis software to analyze the structure and thermal-control subsystem in a picosatelite. In structure subsystem, we have to consider the influence during the launching period. According to the results of center of mass, natural frequency vibration, stress analysis, we can estimate the safety in the structure design. In thermal-control subsystem, heat source and the method of thermal control should be considered. Software simulation is used to understand the picosatellite’s temperature variation when it operates in space. All subsystems are assured to be workable work in the appropriate temperature.

碩士
國立成功大學
電機工程學系碩博士班
91
PACE satellite is the first indigenous Picosat that attempts to achieve 3-axis stabilization. In this thesis, the design and simulation of the attitude control subsystem of the PACE satellite are presented. The requirement is to achieve 5 degree earth pointing accuracy at 600 km altitude, 98 degree inclination orbit. A momentum-biased control system is employed to control the pitch angle and stabilize the roll/yaw dynamics. In addition, three orthogonal magnetic coils are used to damp the angular rate after separation, eliminate nutation and precession phenomena, and prevent the momentum wheel from saturation. Through analysis and simulation, it is shown that a 5° attitude control requirement can be achieved.

This thesis consists in designing and testing the payload of a TubeSat low-cost satellite, called CTSAT-1. Instead of spend large quantities of money to space research, this picosatellite offers a very low-cost solution to provide answers for institutions investigations. Designing and assembling a picosatellite of this size represents a big challenge due to packing of a lot electronics and equipments inside TubeSat low dimensions, therefore providing a unique efforts in reaching a complete small spacecraft. To keep the budget of the project low as possible, maintaining TubeSat specifications and space flight requirements, for intensive testing will be used homemade techniques to achieve satisfactory results without spending too much on expensive equipment.

碩士
國立成功大學
航空太空工程學系碩博士班
92
In this thesis, according to a pico-satellite design specification of 600 Km altitude, 98 degree inclination, 3-axis stabilization, and the requirement of 5 degree earth pointing accuracy, the attitude control subsystem is considered based on the method of 3-axis magnetic active control combined with momentum-biased stabilization. In the initial acquisition mode, de-tumbling is accomplished by using three orthogonal magnetic coils with robust B-dot control law. In the normal mode, a momentum wheel is used with its spin axis aligned along the satellite Y axis to control pitch motion and provide roll/yaw stabilization. Together, the magnetic coils are used to perform pointing control and momentum dumping to prevent the momentum wheel from saturation. For attitude determination, measurement sensors used include a 3-axis magnetometer, a gyroscope, and a coarse sun sensor. The Extend Kalman Filter is selected as the estimation algorithm. To test the correctness and to evaluate the reliability of the designed attitude control flight software, a simulation platform using the xPC toolbox with an attitude control 8051 micro-processor is established to perform real-time hardware-in-the-loop experiment. Several cases with different environmental conditions were tested and the simulation results are included which indicate that the control and determination algorithm is able to satisfy the attitude control requirement.