Analysis and Design of Integrated Magnetorquer Coils for Attitude Control of Nanosatellites

The nanosatellites typically use either magnetic rods or coil to generate magnetic moment which consequently interacts with the earth magnetic field to generate torque. In this research, we present a novel design which integrates printed embedded coils, compact coils and magnetic rods in a single package which is also complaint with 1U CubeSat. These options provide maximum flexibility, redundancy and scalability in the design. The printed coils consume no extra space because the copper traces are printed in the internal layers of the printed circuit board (PCB). Moreover, they can be made reconfigurable by printing them into certain layers of the PCB, allowing the user to select any combination of series and parallel coils for optimized design. The compact coil is wound around the available space in a 1U complaint CubeSat panel and it can accommodate much more number of turns compared to printed coil; consequently generating more torque. The magnetic rod is made complaint with the existing available options and can easily be integrated in the panel. This design gives a lot of flexibility because one can choose to optimize power, optimized torque or rotation time by choosing among the available magnetorquer options. The proposed design approach occupies very low space, consume low power and is cost effective. The analysis in terms of generated torque with certain applied voltages, trace widths. The analysis results in terms of selection of optimized parameters including torque to power ratio will be presented

Smart attitude control system for small satellites

The attitude control system is one of the most important systems for satellites, which is essential for the satellite's detumbling, pointing, and orbital maneuver. The conventional attitude control system consists of magnetorquers, reaction wheels, and thrusters. Among these actuators, magnetorquers are widely used for satellite detumbling and attitude control, especially for small satellites and CubeSats. It consumes zero propellant compared with thrusters and has a high chance of survival compared with the reaction wheel as it does not contain any moving parts, which makes them last longer in harsh environments. Conventional magnetorquers utilize air or soft magnetic materials, e.g., iron and alloys, as core, and the magnetic field is generated by feeding the electric current to the wrapped solenoid. Due to the power limit of the small satellites, the magnetic field strength is strictly limited, and the continuous current supply results in massive energy consumption for detumbling and other attitude adjustment missions. The long copper wire of the solenoid will also result in high resistance and generate significant heat. To improve the current design and overcome the proposed drawbacks, a novel electro-permanent magnetorquer has been designed and developed in this thesis as one actuator of the attitude control system. Unlike conventional magnetorquers, the electro-permanent magnetorquer utilizes hard magnetic materials as the core, which can maintain the magnetization when the external magnetic field is removed, to generate the required magnetic field. A special driving circuit is designed to generate the desired dipole moment for the magnetorquer, and the components used for the circuit are carefully selected. The experiments show that the electro-permanent magnetorquer can generate 1.287 Am2 dipole moment in either direction. The magnetorquer works in pulse mode to adjust the dipole moment, and it requires around 0.75 J energy maximum per pulse. A single-axis detumbling experiment has been conducted using only one torque rod on the air-bearing table inside an in-house manufactured Helmholtz cage. The experiment results show that the magnetorquer can detumble the air bearing table with 0.0612 kgm2 moment of inertia from an initial speed of around 27°/s to zero within 800s, and total energy of 82.92 J was consumed for the detumbling experiment. A single torque rod single-axis pointing experiment has been conducted with a sliding mode controller on the same platform. The results show that a single torque rod can achieve the target angle and maintain the error discrepancy within the ±0.4° boundary under a specific system configuration. A micro air-fed magnetoplasmadynamic thruster has been designed and tested as another attitude control system actuator. The thruster is a miniaturized electric propulsion system based on the conventional full-scale magnetoplasmadynamic thruster that operates at hundreds of kilowatts. The thruster is designed and tested using normal air as the propellant under the pulse operation mode on a calibrated micro-force measurement thruster stand. The experiments revealed that the thruster could generate a 34.534 µNs impulse bit with an average power input of 1.857 ± 0.0679 W and thrust to power ratio of 8.266 µN/W. The specific impulse is calculated to be 2319 s with a thruster efficiency of 9.402%, which is quite competitive compared with other solid-state and liquid-fed pulse-mode thrusters. This paper presents the design and test results for the thruster under a low power level, as well as an analysis of its problems and limitations with corresponding future research and optimization directions noted at the end. The electro-permanent magnetorquer as a payload of the CUAVA-2 satellite mission has been introduced in this thesis. The design considerations and adjustment based on the requirement of the CUAVA-2 has been introduced in detail. A simple sliding mode controller has been developed to achieve three-axis attitude control using both electro-permanent magnetorquer and the micro air-fed magnetoplasmadynamic thruster. The controller's performance has been tested using MATLAB-based simulation with the experimentally obtained performance parameters and some assumptions. The results show that the smart attitude control system can achieve ±0.005° pointing error discrepancy with the help of both actuators

Study of the Business Model of three Earth Observation (EO) companies already present in the Very Low Earth Orbit market (VLEO)

The emergence of a new private spaceflight industry has taken the Earth Observation (EO) sector by surprise. NewSpace companies are challenging the traditional satellite sector by addressing their services to mass market requirements of high-quality and low-cost EO. As part of the DISCOVERER project, this study aims to determine the Key Success Factors to consider by a new EO company at Low Earth Orbit (LEO). Hence, three businesses fitting the description were analyzed with the Case Study Methodology to establish their Business Model Canvas (BMC), associated Patterns, and Key Success Factors. The investigation consolidated the newly proposed Democratizing Business Model Pattern and added new characteristics. Successful EO NewSpace firms are getting divided between integrated operators, integrated manufacturers, and end-user specialists. A new EO company should consider the Democratizing Pattern success factors and the Vertically Integrated Strategies (VIS), depending on its disruptive idea and resource capabilities. Further research is needed to identify new factors, strengthen the validity of the Pattern, and VIS tendencies

State of the Art: Small Spacecraft Technology

This report provides an overview of the current state-of-the-art of small spacecraft technology, with particular emphasis placed on the state-of-the-art of CubeSat-related technology. It was first commissioned by NASAs Small Spacecraft Technology Program (SSTP) in mid-2013 in response to the rapid growth in interest in using small spacecraft for many types of missions in Earth orbit and beyond, and was revised in mid-2015 and 2018. This work was funded by the Space Technology Mission Directorate (STMD). For the sake of this assessment, small spacecraft are defined to be spacecraft with a mass less than 180 kg. This report provides a summary of the state-of-the-art for each of the following small spacecraft technology domains: Complete Spacecraft, Power, Propulsion, Guidance Navigation and Control, Structures, Materials and Mechanisms, Thermal Control, Command and Data Handling, Communications, Integration, Launch and Deployment, Ground Data Systems and Operations, and Passive Deorbit Devices

Modular, reconfigurable approach for a commercial space spacecraft programme

This thesis presents the work performed in producing a system-level design for a modular, multipurpose small satellite platform. A multipurpose platform may be applied to a wide range of missions, and, to be commercially viable, the envelope of missions for which it is suitable should be as large as possible. The research therefore addresses the particular requirements that are specific to different mission types, and produces characteristic requirement sets for each. General design requirements are also derived, such as those for enabling modularity and allowing compatibility with different launch vehicles. The commercial requirements arising from the different market and customer sectors are also examined. Industry analysis allows identification of general market trends, and predictions are made regarding the likely size and characteristics of the market in which the proposed platform would compete. It is anticipated there could be a worldwide demand for more than twenty small satellites each year, for which a flexible small spacecraft platform could potentially compete. After derivation of the necessary requirements has been performed, a system-level design of the spacecraft platform is undertaken. The resulting design is based on a multi-module, reconfigurable concept, which can be adapted to fit the different launch envelopes of Pegasus-XL, Taurus, ASAP-5 and larger launchers, and also to accommodate a wide range of payloads. The subsystems are offered in different capability variants, which may be interchanged in response to different mission requirements. The platform equipment and structure forms a 'standard parts lisf', from which the appropriate configuration can be built up. Schedule reductions are obtained due to the modular design allowing more of the integration and testing of the platform to be performed in parallel. The proposed programme for development of the platform uses up-front investment to conduct much of the detailed design of the platform in advance of any actual project. This allows the design effort to be shared across many subsequent projects, and the design phase of each new project to be minimised. The key benefits of the proposed platform and programme are adaptability, ability to rapidly reconfigure to mission requirements, suitability for future upgrading, and reduction of the project schedule.EThOS - Electronic Theses Online ServiceGBUnited Kingdo

Modular, reconfigurable approach for a commercial space spacecraft programme

This thesis presents the work performed in producing a system-level design for a modular, multipurpose small satellite platform. A multipurpose platform may be applied to a wide range of missions, and, to be commercially viable, the envelope of missions for which it is suitable should be as large as possible. The research therefore addresses the particular requirements that are specific to different mission types, and produces characteristic requirement sets for each. General design requirements are also derived, such as those for enabling modularity and allowing compatibility with different launch vehicles. The commercial requirements arising from the different market and customer sectors are also examined. Industry analysis allows identification of general market trends, and predictions are made regarding the likely size and characteristics of the market in which the proposed platform would compete. It is anticipated there could be a worldwide demand for more than twenty small satellites each year, for which a flexible small spacecraft platform could potentially compete. After derivation of the necessary requirements has been performed, a system-level design of the spacecraft platform is undertaken. The resulting design is based on a multi-module, reconfigurable concept, which can be adapted to fit the different launch envelopes of Pegasus-XL, Taurus, ASAP-5 and larger launchers, and also to accommodate a wide range of payloads. The subsystems are offered in different capability variants, which may be interchanged in response to different mission requirements. The platform equipment and structure forms a 'standard parts lisf', from which the appropriate configuration can be built up. Schedule reductions are obtained due to the modular design allowing more of the integration and testing of the platform to be performed in parallel. The proposed programme for development of the platform uses up-front investment to conduct much of the detailed design of the platform in advance of any actual project. This allows the design effort to be shared across many subsequent projects, and the design phase of each new project to be minimised. The key benefits of the proposed platform and programme are adaptability, ability to rapidly reconfigure to mission requirements, suitability for future upgrading, and reduction of the project schedule.EThOS - Electronic Theses Online ServiceGBUnited Kingdo

Communication Networks in CubeSat Constellations: Analysis, Design and Implementation

CubeSat constellations are redefining the way we approach to space missions, from the particular impact on scientific mission possibilities, constellations potential is growing with the increasing accessibility in terms of low development and launch costs and higher performances of the available technologies for CubeSats. In this thesis we focus on communication networks in CubeSat constellations: the project consist of developing a clustering algorithm able to group small satellites in order to create an optimized communication network by considering problems related to mutual access time and communication capabilities we reduce the typical negative effects of clustering algorithms such as ripple effect of re-clustering and optimizing the cluster-head formation number. The network creation is exploited by our proposed hardware system, composed by a phased array with up to 10dB gain, managed by a beamforming algorithm, to increase the total data volume transferable from a CubeSat constellation to the ground station. The total data volume earned vary from 40% to a peak of 99% more, depending on the constellation topology analyzed

Power Management, Attitude Determination and COntrol Systems of Small Satellites

Satellites have always been considered to be extremely expensive and risky business, which not only requires extensive knowledge and expertise in this field but also huge budget. Primarily, this concept was based on initial development and launching cost. Secondly, it was also impossible to repair and substitute parts (this was true up to 1993: the first Hubble Space Telescope servicing mission), which makes design more tough because it requires advanced fault tolerance solutions and extreme reliability. But with the passage of time many space actors entered in this market. Low cost design techniques played an important role in the aerospace market growth in the past years, but they can still play a major part in future developments. At present, several private companies are also providing launch services which further lower the accumulative cost. Many universities and SMEs (Small Medium Enterprises) worldwide are also trying to reduce satellite costs. The Department of Electronics and Telecommunication (DET) at Politecnico di Torino has been working on NanoSatellites since 2002 and developed their first NanoSatellite called PiCPoT, which was intended to be launched together with other university satellites by a DNEPR LV rocket in July 2006. Unfortunately a problem in the first stage of the carrier led to the destruction of all satellites. After that DET started work on a comprehensive NanoSatellite project called AraMiS (Italian acronym for Modular Architecture of Satellites). The main idea of the AraMiS is modularity at mechanical, electronic and testing levels using Commercial-Off-The-Shelf (COTS) components. These modules can be assembled together to get the targeted mission, which allows an effective cost sharing between multiple missions. AraMiS satellites have mass up to 5kg with different shapes and dimensions. AraMiS-C1 is a CubeSat Standard satellite developed on the AraMiS approach. Four sides of the AraMiS-C1 are equipped with identical tiles called 1B8_CubePMT that mount solar panels on the exterior and a combined power management, attitude control and computing subsystem on the interior. The other two sides are devoted to the telecommunication tiles called 1B9_CubeTCT which carry a commercial deployable UHF antenna (one side) and a patch type SHF antenna (the other side). Thesis discusses in detail the design, implementation and testing of the 1B8_CubePMT module. It is developed on the design approach of AraMiS architecture with dimension 98×82.5×1.6 mm3. 1B8_CubePMT module contains electric power supply (EPS) and attitude determination & control subsystems (ADCS) of AraMiS-C1 satellite. The integration of such a large number of systems in a small area was not a trivial job. Several techniques were employed for reduction of size, weight and power consumption of the different subsystems while still achieving best performances. COTS components were selected for the EPS subsystems, on the basis of power loss analysis and minimum dimensions which helped in efficiency enhancement and also miniaturization of the subsystems. ADCS subsystems components were also selected on the basis of minimum dimensions and lower power consumptions while still achieving targeted performances. The most interesting feature of the 1B8_CubePMT module is the design and integration of a reconfigurable magnetorquer coil within four internal layers occupying no excess space. Coils in each layer are treated separately and can be attached/detached through straps. Changing the arrangement of these straps make the magnetorquer reconfigurable. Different housekeeping sensors have been employed at various points of the 1B8_CubePMT module. Thesis also discusses thermal modeling of CubeSat, AraMiS-C1 satellite and 1B8_CubePMT module. Thermal resistance and temperature differences between different sides of the satellites and individual tiles have been found. At the end, preliminary thermal and spin analysis of NanoSatellites have been presented. Chapter 1 gives an introduction to the problem and proposed solutions which will be discussed in this thesis. Chapter 2 presents an introduction to AraMiS project and AraMiS-C1 satellite. Chapter 3 discusses different satellite design flow configurations and their comparison. Chapter 4 discusses 1B8_CubePMT module which is a CubeSat standard power management tile, developed on the AraMiS concept, for AraMiS-C1 satellite. It has EPS and ADCS subsystems which are the most essential elements of any aerospace mission. Chapter 5 deals with the design and development of the EPS system of AraMiS-C1 satellite. This chapter discusses how to reduce the size, weight and power consumption of the EPS subsystems while achieving better efficiency and fulfilling satellite power requirements. The selection of COTS components on the basis of power loss analysis and minimum dimensions is discussed in detail. Housekeeping sensors such as current, voltage and temperature sensors which are employed at different points of the 1B8_CubePMT module to cope with anomalies, have been discussed in detail in this chapter. At the end of the chapter, the designed EPS is evaluated on the basis of AraMiS-C1 power budget. Chapter 6 discusses design and implementation of attitude determination sensors (ADS) of the AraMiS-C1 satellite. 1B8_CubePMT has three types of attitude determination sensors: sun sensor, magnetometer and gyroscope. This chapter discusses in detail the design and operation of these sensors. Chapter 7 discusses the attitude control (ADC) system of AraMiS-C1 satellite. The design and implementation of a reconfigurable magnetorquer coil which is embedded inside the 1B8_CubePMT module, is discussed in detail. The designed magnetorquer has been evaluated on different parameters and compared with the magnetic actuator already available in the market. In chapter 8 testing procedure and results of 1B8_CubePMT subsystems are discussed in detail. Chapter 9 presents thermal modelling of NanoSatellites. Detailed and simplified thermal models of CubeSat panel have been discussed. Thermal resistances measured through both models are compared. Generic thermal model of a CubeSat is presented. Utilizing the proposed models, thermal resistance of 1B8_CubePMT and AraMiS-C1 are measured. In order to verify the theoretical results, the thermal resistance of the AraMiS-C1 is measured through an experimental setup. Chapter 10 discusses preliminary thermal and spin analysis of NanoSatellites in space environment. All the heat sources and their effects on the satellite have been discussed. A thermal balance equation has been established and satellite temperature for different structures and various conditions has been found. At the end a satellite spin analysis on the basis of different absorption coefficient related with colors, has been discussed