Cubesats

In this paper, a printed Yagi antenna with an integrated balun is proposed for CubeSat communications. The printed antenna is mechanically adjustable to realize three functional states at different operating frequencies in the L-band and S-band respectively. Three different angle deployments are proposed at 10°, 50° and 90°, so that the antenna operates at three different operating frequencies, namely 1.3 GHz (L-band), 2.4 GHz (S-band) and 3 GHz (S-band). The measured results of the fabricated antenna are well matched with the simulation, having frequencies of 2.82–3.07 GHz, 1.3–1.4 GHz and 2.38–2.57 GHz, with similar radiation patterns. The measured gain of the antenna is 8.167 dBi at 2.4 GHz, 5.278 dBi at 1.3 GHz and 6.120 dBi at 3 GHz. Keeping within the general theme of cheap off the shelf components for CubeSats, this antenna design allows the CubeSat designers to choose from three popular frequencies, through a simple angle configuration. The main contribution of this work lie...

CubeSats are miniature satellites that have been instrumental in low Earth orbit research for the past 15 years. But recently the tiny CubeSat has been called upon to go where no CubeSat has gone before – interplanetary space. Learn more in an Introduction to CubeSats course by Tonex.

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly, Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnológico de Aeronáutica, in São José dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnológico de Aeronáutica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.

The Satellite Applications Catapult (the Catapult) In-Orbit Demonstration (IOD) Programme aims to stimulate and support the growth of UK businesses. Specifically, the IOD programme provides a series of flight opportunities for businesses to demonstrate new commercial satellite services. Applicants are required to provide a comprehensive business plan, utilising a mission concept based on a 3U or 6U CubeSat platform. The proposal must also include plans for implementing any bespoke payloads or user segment equipment required to demonstrate the complete service. The Catapult manages manufacture of the CubeSat platform and delivery to orbit, in addition to provision of operations, data processing and dissemination infrastructure services, typically in under 12 months. IOD-1 GEMS, the first IOD programme mission, will be delivered in partnership with Orbital Micro Systems to provide global weather data services. The objective of IOD missions is to reduce commercial risk and improves time to market for novel satellite services. It has also identified a growing interest in spacecraft larger than the 3U CubeSat platform. The IOD Programme continues to offer flight opportunities, however, it is identified that the commercial use of small satellites has evolved during the duration of the IOD Programme, which will need to be revised accordingly.

Cubesats have become quite popular especially among university students as these miniaturized satellites help them in carrying out space exploration feasibly. Till date, cubesats have never flown in space with any well defined propulsion system. But in order to increase mission capabilities like orbit raising and formation flying, or to perform proximity operations, fine attitude control, drag make up and de-orbit without much risks, it is important to use a propulsion system. Here in this report we have focused on the use of cold gas thrusters in a cubesat.Cold gas thrusters consist of a pressurized tank containing gaseous propellant, such as nitrogen, and a solenoid actuated valve system leading to exit nozzles. Since the propellant is unheated and relies solely on the enthalpy of the stored gas, the velocity at the nozzle exit is relatively low resulting in a low specific impulse, typically around 60 sec, useful for small attitude adjustments and low ΔV maneuvers. However, the power level, pressure, and weight required in the vessel are though less, but still higher than the Cubesat Specifications provided by the CalPoly. Rather than relying on high-pressure chambers, we propose the use of solid gas cartridges, which upon ignition by lasers can be released into a chamber in the gaseous form, from which nitrogen can be drawn for propulsive use. Moreover since only a fraction of the propellant is used for propulsion use, it helps in reducing the tank size and minimizes the risk of leakage and pressure limitations, with reduction in mass by around 20%. This way a suitable propulsion system for a cubesat can be developed which renders increased performance with respect to mission capabilities without increasing the overall mass of the satellite. The idea would also prove to be of much benefit for industries which are planning future cubesat missions.

After a decade of experimental nanosatellite projects, the past few years have seen a growing trend of utilizing nanosatellites in more complex scientific and commercial missions. More complex nanosatellite missions typically involve longer target lifetimes, sophisticated instruments and complex ground operations. These missions can achieve lower costs than traditional satellites by using commercial off-the-shelf (COTS) components where possible. However, COTS components may not function reliably in space for extended periods of time, and more ambitious missions with longer lifetimes require more reliability than experimental nanosatellites. Components developed by the traditional space industry are known to be reliable and robust, but they are usually too expensive and bulky compared to the requirements of nanosatellites. Using components designed for terrestrial safety-critical applications, such as automotive use, could provide cost savings while maintaining reliability compared to space-grade components. This paper considers the Texas Instruments Hercules TMS570 microcontroller family for the on-board data handling (OBDH) system of a nanosatellite mission in development at Aalto University. Several parties, including NASA, already plan to use Hercules in space applications. The proposed on-board computer (OBC) design is based on a Texas Instruments TMS570LS3137 Hercules Safety Microcontroller with built-in fault tolerance, including two processors running in lockstep and error correction codes (ECC) implemented in internal memories. Other components in the OBC are selected from automotive and other high-grade catalogs. The proposed OBC design is compared to LEON-based traditional OBCs and some CubeSat designs, and the built-in fault-tolerant lockstep architecture is compared to some previous fault-tolerant methods used in nanosatellites. Using microcontrollers with built-in fault tolerance for nanosatellite OBDH provides more reliability while speeding up mission development.

3 | P a g e CERTIFICATE I hereby certify that the work which is being presented in the project report entitled "Trajectory Analysis of UPESSAT" in partial fulfilment of the requirements for the satisfactory performance for Major Project, final year, submitted in the Department of Aerospace Engineering, University of Petroleum and Energy Studies, Dehradun is an authentic record of these students" own work carried out during a period from 4 | P a g e DECLARATION I hereby declare that the major project work titled "Trajectory Analysis of UPESSAT" is unique to the best of my knowledge and submitted to UPES, Dehradun. It is the bonafide record of the original work carried out by us under the exemplary mentorship of Dr. Ugur

With the surge of interest in nano-satellites, there is a concomitant need for high quality, yet affordable simulation and testing environments. It is particularly challenging to experimentally evaluate nano-satellite attitude control systems in a test environment. This article investigates the technical feasibility of fabricating a low-cost air-bearing platform with three degrees of freedom of angular motion using desktop 3-D printing technology with limited printing resolution. An open source air-bearing attitude simulator for complete 1U CubeSat is proposed , manufactured, and characterized. The platform is equipped with directional air nozzles that enable external torque generation in order to cancel out i) parasitic moments of inertia from the satellite's enclosure and ii) error torque produced by imperfections. It is also capable of simulating disturbances in a space environment. The results show that the torques produced by the nozzles can reach beyond 0.001 Nm and are sufficient to remove error torques and provide torque compensation of the orders of 0.0004 Nm. Removing the effects of gravity torque with the nozzles proved to be unachievable with the current design, requiring precise positioning of the CubeSat within the enclosure. Future work has been identified for a number of improvements to the design and details for the further development of the platform.

AbstrAct: Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly, Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnológico de Aeronáutica, in São José dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnológico de Aeronáutica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel de-saturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.

A compact hydrazine/hydrogen peroxide pressure-fed bi-propellant propulsion system for long term linear 4U CubeSat missions has been designed. The requirements of providing a delta v of 600m/s, containing a 1U payload of 1.33kg, and to not exceed an acceleration of 3g have been satisfied. The specific impulse was determined to be 332 s with a given constant thrust of 10N.
The design decisions for propellant choice has been justified via calculations using a NASA CEA software package, and operating principles and key design considerations for each component are outlined. A complete analysis of the thrust chamber was completed with material selected to maintain simplicity and increase reliability, and injector plate sizing performed. COTS components have been reviewed to aid in calculations and potential inclusion into the design.
CAD drawings and schematics have been developed and failures modes considered and analysed leading to potential off design conditions.

Co-Authors: Azzimonti Nicolò, Bassissi Enrico, Bologna Andrea, Colombo Alessandro, De Blasi Francesca, De Luca Maria Alessandra, di Trocchio Marco, Fiore Walter, Mucci Marco, Parisi Adrian Burton

SwissCube: the first entire Swiss satellite is now on orbit; planned to operate for at least twelve months, it will have four years in the 23th of September; four years of data and results from the space and from its payload and operations. As cubesat made by students it is a good result: more than 200 students worked on it before and after the launch in 2009, designing, testing and validating the data from the instruments on board. In this paper it is presented a general description of the cheap technology used for SwissCube, giving a particular attention to the tests done before and after flight, describing the payload mounted and the results achieved. This study wants to show the good results and the errors made, the good operations and the malfunctions during the flight in order to spread the technology used for future CubeSats. As example, results obtained by a cheap and an easy B-dot controller using magentotorquers are here presented; camera's pictures, gyros, magneto-meters, sun-sensors and thermometers results are also shown, giving particular relevance to the performances and to the difficulties had during the years. The goal of SwissCube was mainly educational, teaching and showing to students how to built a Nano/Pico satellite with available just 10x10x10cm and 1kg and this study wants to show the results achieved during the design and the test according to the flight data, in order to increase the technology and the “know-how" experience of CubeSats platforms.

The aim of this research is to investigate how to track the QB50 cubesats when they are
deployed from the rocket launchers. Also the use of the ground station in Von Karman
Institute to track, identify and distinguish the signals for each of the 50 cubesats of the
QB50 constellation is examined in this research.
Cubesats, due to their small size and weight, are low cost platforms to place in orbit. In
parallel to this, cubesats are becoming more powerful everyday with very promising future
from earth observation to telecommunications.
Cubesat tracking and signal analysis is a crucial milestone in the cubesats constellation.
The deployment of multiple cubesats from a rocket at an interval of about 30-60 seconds,
creates a cloud of multiple cubesats orbiting close to each other. The process of tracking
an individual cubesat and distinguish each signal is a complex process. Thus a research
in the antennas, tracking system, orbital mechanics, electromagnetics, software and
general ground station assembly is needed.
Furthermore to track the cubesats exact orbital elements for a satellite are needed. Orbital
elements or two line elements are also very hard to be generated at the early stages of
deployment. NORAD the organization who is generating the two line elements for all
space objects, is able to generate accurate two line elements for the cubesats after one
month of deployment.
As a conclusion to this research two methods are presented to track the QB50
constellation. The first is aiming to use the existing antennas and devices at VKI and the
second is presenting a more sophisticated mini radar like system with a parabolic antenna
which can be used alone or in a combination with the existing system. Also for future
cubesats constellations a proposed system of time flags and time measuring is proposed.

The Norwegian Student satellite program was established in 2007 and is now coming to an end. After running the program for several years, the tangible output from the program is one CubeSat in orbit and two still under construction. This paper gives an overview of the achievements during this period. The student satellite program had several goals, not limited only to the launch of 3 CubeSats. Several students have attended conferences and workshops; some presented their work at international conferences or even published their work in scientific papers. There are multiple spin-off projects and lessons learned from this program. Some examples of these, in addition to the main results from the satellite program will be presented. An analysis of which of the objectives have been accomplished and why others have not been reached will be presented. During the project period, there has been an increasing interest in continuing work with student satellites at university level. Two participating universities now have ongoing activities on scientific payloads for CubeSats for various applications. Andøya Space Center has, together with partners, investigated the possibility of establishing a launch facility for nano-and micro satellites. If this becomes a reality, it will open many opportunities for universities and the space industry in Norway. Space technology is often a strong field of interest for young students, and can be an important recruitment factor for STEM studies.

The production of a ferrofluid container, intended for use in the KRAKsat (CubeSat type) satellite in space conditions, is presented. Mechanized laser beam welding for AISI 316L stainless steel test joint and container prototype was developed and tested. The welded test joints were examined by non-destructive visual, penetration and radiographic testing and destructive testing by macro-and microscopic examination, static tensile test, static bending test, and hardness measurements. The welded container prototype was examined by leak test, temperature-vacuum test and vibration test. Test joints' evaluation showed a proper selection of welding parameters and expected quality of joints. Austenitic microstructure with small δ-ferrite content in base materials, heat-affected zones, and welds guarantees sufficient mechanical properties for this part geometry. The tensile strength range of test joints was 687-729 MPa, hardness range was 140-200 HV3, and the bending angle was 180°. Welding of the prototype container and testing of tightness, resistance to temperature changes, and vibration were successful. Compliance with flywheel design and manufacturing requirements will enable the launch of a research satellite into orbit with such a wheel.

In this paper, we will describe the Finnish Aalto-1 CubeSat satellite mission, the satellite status and the first in-orbit observations. The Aalto-1 satellite is designed by students is built by a consortium of universities and institutes. It is the first Finnish nanosatellite project and also the first satellite registered to Finland. The mission has primary goals in education, technology demonstration and science. The mission has initiated several New Space companies and initiatives in Finland and brought the nation to the circle of space faring nations. The satellite platform was mainly developed by Aalto University students while the payloads were developed by consortium partners. The satellite was successfully launched in June 2017. By the end of July, all instruments on-board of the satellite were commissioned and first scientific measurements were performed with spectral camera and radiation measurement instrument. Also, data was collected about all other sensors and subsystems, such as solar cells with conductive cover, novel sun sensors and redundant Linux-based onboard computer.

In less than a decade, Cubesats have evolved from purely educational tools to a standard platform for technology demonstration and scientific instrumentation. The use of COTS (Commercial-Off-The-Shelf) components and the ongoing miniaturization of several technologies have already led to scattered instances of missions with promising scientific value. Furthermore, advantages in terms of development cost and development time with respect to larger satellites, as well as the possibility of launching several dozens of Cubesats with a single rocket launch, have brought forth the potential for radically new mission architectures consisting of very large constellations or clusters of Cubesats. These architectures promise to combine the temporal resolution of GEO missions with the spatial resolution of LEO missions, thus breaking a traditional tradeoff in Earth observation mission design. This paper assesses the current capabilities of Cubesats with respect to potential employment in Earth observation missions. A thorough review of Cubesat bus technology capabilities is performed, identifying potential limitations and their implications on 17 different Earth observation payload technologies. These results are matched to an exhaustive review of scientific requirements in the field of Earth observation, assessing the possibilities of Cubesats to cope with the requirements set for each one of 21 measurement categories. Based on this review, several Earth observation measurements are identified that can potentially be compatible with the current state-of-the-art of Cubesat technology although some of them have actually never been addressed by any Cubesat mission. Simultaneously, other measurements are identified which are unlikely to be performed by Cubesats in the next few years due to insuperable constraints. Ultimately, this paper is intended to supply a box of ideas for universities to design future Cubesat missions with high scientific payoff.

In order to support the interoperability between heterogeneous missions, Federated Satellite Systems (FSS) should include flexible communications technologies supporting protocol translation among different federated spacecrafts. A spacecraft equipped with such technology is what we refer to as a federated negotiator spacecraft. A federated negotiator node acts as a middle man between spacecraft missions that would be otherwise unable to interoperate, acting as a protocol translator and enabling the opportunistic, ad-hoc exchange of resources envisioned in the FSS paradigm. This paper presents the ongoing efforts in the development of a negotiator spacecraft for federated satellite systems. These efforts include the implementation of Software-defined Radio (SDR) systems as an enabling technology for negotiator spacecraft. In particular, this work evaluates the use of commercial-off-the-shelf SDRs and GNU Radio as the technical solution to implement negotiating functions on federated spacecraft. The live signal received from ESTCube-1 is used for the demonstration of the protocol translation between SDRs, emulating different spacecraft operating at different protocols.

CubeSats have a wet mass ranging from 1.3 (1U) to 6 kg (3U) and employ commercial off-the-shelf electronics [1]. A key advantage of CubeSats is that they are small, lightweight and have the ability to form a constellation of cube satellites that communicate directly with one another [2]. To this end, CubeSats require a cross-link communication antenna [3]. This, however, poses some challenges. One of which is designing a low profile antenna that occupies a small area, provides high gain and achieves wide directivity. Another is establishing reliable cross-link communications between cube satellites in a swarm without a priori knowledge of their position.  Therefore, it is critical CubeSats employ an antenna system that provides wide directivity and high gain. The most suitable antennas that address most of the aforementioned challenges are planar antennas. They have a low-cost, low profile, light weight and easily integrated with electronic devices. Having said that, a key consideration is the configuration required to enable cross link communications between CubeSats operating in a swarm. To date there has been limited work that uses multiple planar antennas for CubeSats. To the best of our knowledge, only the authors of [4] have proposed a similar set up to us.  However, their design suffers from back radiation that may cause interference with electronic components inside the CubeSat. Another limitation is their robustness during launch because the antennas require soldering connections.  Lastly, their antenna is designed for a 3U CubeSat while ours is targeted at 1U CubeSat.
In this paper, we propose a unique configuration of six small ACM antennas.  Each individual ACM patch antenna is placed at the top corner of each face of a 1U CubeSat.  The placed antennas thus provide communication in all six directions; see Fig.1.  Moreover, they only occupy 11.5% of the total surface area on each CubeSat’s face.  Hence, the remaining real estate, i.e., 88.5%, is sufficient to mount solar cells. Fig. 2 shows a single ACM patch antenna. It has a small volume, i.e., 50×23×3.05 mm3, and is easy to fabricate. Its main advantage is the use of an aperture coupling technique that does not require a soldering connection. This is important as it makes the antenna robust during the launch stage and in space. As shown in Fig. 3, each individual ACM patch antenna achieves a total gain of 5.72 dB, directivity of 6.24 dB and a return loss of -26 dB at a resonance frequency of 2.45 GHz.

​ Mountain glaciers are important because they work as precipitation buffers and water sources for downstream populations. In addition, they have been a source of data for climate dynamics analysis. Measuring glaciers meteorological variables, estimating the amount of black carbon, as well as recording mass slides on a periodic basis becomes mandatory to understand and drive effective environmental policies. While these measurements are traditionally executed via expensive field campaigns, they tend to be scarce and non-periodic. Thus, remote sensing alternatives have been explored by the community, they are based in monolithic satellite architectures which besides the long development time, are bounded in mission lifetime. In this paper we propose a Glacier Segmented Observation System (SOS), based on a number of interconnected cubesats under the segmented architecture paradigm to provide frequent revisit times and extendable lifetime for critical observation missions. We evaluate how SOS can support a worldwide glacier remote sensing mission to then discuss relevant mission requirements, formation flying disposition of satellites and communication architecture challenges.

This paper investigates the verification, validation, and testing (VVT) processes that MarCO underwent. When available, actual design, integration, and VVT steps undertaken for the MarCO mission are used. When MarCO specifics are not available, general NASA standards are referenced. After a review of the MarCO mission background, this paper discusses MarCO for each of the six SYSE 640 course goals.

TURKSAT-3USAT is a three unit CubeSat developed for voice communication at low Earth orbit. The payload is a VHFIUHF transponder operating at amateur bands. The transponder and all other subsystems, except the stabilization, are doubled for redundancy. Where possible, both COTS systems and in-house development is employed. The energy is provided using lithium polymer batteries together with super capacitors. Satellite stabilization is accomplished using passive magnetic attitude control system with hysteresis rods. The TURKSAT-3USAT is launched from JSLC of China, on April 26, 2013.

GPS receivers are increasingly used in nanosatellite applications to provide the means for autonomous and accurate navigation in orbit. However, many of the available nanosatellite GPS receivers consume a large amount of power compared to the available power budget. This can be a problem when the main science instruments have high power requirements and GPS is only used to support the science mission. This paper presents the design and proposed integration and qualification of a GPS subsystem that uses commercial, low-power GPS components to support the science mission of the 3U CubeSat-based Aalto-1 nanosatellite. Aalto-1 has three science payloads which require accurate positioning: a Fabry-Pérot imaging spectrometer used for remote sensing, a radiation monitor used for studying the near-Earth radiation environment and an electrostatic plasma brake used for deorbiting the satellite at the end of the mission. Ground-based tracking with only one ground station or NORAD-provided two-line element sets cannot position the satellite accurately enough for the science payloads, and it was concluded that GPS navigation is the best option. The GPS receiver and antenna have a combined power budget of 160 mW. Navigation accuracy of less than 100 meters is needed by the scientific instruments of Aalto-1. Effects of antenna placement on GPS navigation performance are considered, as the antenna on Aalto-1 will have a limited view of the sky. The preliminary functional tests and simulations with a GPS signal simulator indicate that the subsystem is capable of navigating in orbit and providing the Aalto-1 mission with the required navigation capability.

Works on pico-satellites have gained momentum recently, especially those that consider pico-satellites as part of a much larger constellation or swarm. This feature allows pico-satellites to provide high temporal resolution of observational data and redundancy. In particular, it reduces the need for satellite-to-ground communications and, hence, helps save energy and allows the execution of distributed processing algorithms on the satellites themselves. Consequently, satellite-to-satellite or cross-link communication is critical. To realize these advantages, the cross-link antenna employed on pico-satellites must meet many criteria, namely, small size, lightweight, low-power consumption, high gain, wide bandwidth, circular polarization, and beam steerability. To date, no works have examined the suitability of existing planar antenna designs for the use on pico-satellites. To this end, this paper contributes to the literature by focusing on microstrip patch and slot antennas that have the ability to achieve high gain, beam steering, and wide bandwidth. This paper reviews 66 planar antenna designs, which includes 38-patch and 28-slot antennas. In addition, we provide an extensive qualitative comparison of these antennas in terms of their mass, size, gain, beam steerability, type of polarization, operating frequency band, and return loss. In addition, we have evaluated three antenna designs that best address the pico-satellite challenges on a common platform. We find that the asymmetric E-shaped patch antenna design is the most suitable for the use on 2U CubeSats. This is because of its small size (34 × 13 mm 2) and high gain (7.3 dB). In addition, the E-shaped patch antenna yields a wide −10-dB bandwidth of 2300 MHz and a small return loss of −15.2 dB. INDEX TERMS Circular polarization, reconfigurable antennas, antenna arrays, beam steering, CubeSat, pico-satellites, planar antennas, microstrip patch antennas, sequential phase-rotation, photonic band-gap (PBG).

Quantum Key Distribution (QKD) using entangled photon pairs enables highly-secure communications but requires that individual photons be exchanged between the two communicating parties. Practical limitations prevent the distance over which this is performed being more than a few hundred kilometres on the ground, but from space much longer distances can be achieved. A first step towards enabling space-to-ground QKD is to develop and demonstrate a space-compatible entangled photon source. To pursue this goal CQT are developing miniaturised ruggedised sources that are compatible with the CubeSat nanosatellite platform. Although CubeSats impose significant size, weight and power restrictions, their large developer base and widespread adoption enables low-cost raising of technology readiness level with rapid development cycles. This paper gives an overview of the SPEQS-1 devices that are being built to prove the new principles of our miniaturised entangled photon sources, SPEQS-2 which is being developed to be powerful enough for space-to-ground QKD and SpooQySats, the CubeSat platforms we are developing to test SPEQS-2 prior to full space-to-ground QKD demonstrations. This abstract has been modified from the one originally submitted to the IAC conference proceedings, for original see appendix.

—In recent years, the interest for the Arctic area has been increasing. Harvesting of scientific data and environmental monitoring are key activities. The Arctic has poor communication infrastructure, both by terrestrial and satellite systems. Launching a free-flying constellation of small Cube-Sats is one proposal to help mitigate this service gap. CubeSats are traditionally built with industrial-grade components, which reduces the development time and hardware cost. Since the cost of one satellite is low, it is possible to launch several together. The CubeSats are generally launched without any station keeping capabilities, as this increases cost and complexity in both the production and operational phase. Without station keeping, the swarm satellites will drift relatively to each other. The governing parameter of the drift is the velocity difference of the satellites at the time of deployment. This paper shows how a freely drifting swarm can improve the coverage for sensor networks in the Arctic, when an effort is made to optimize this velocity difference at deployment. E.g., a free-flying constellation of three satellites will have better coverage properties than a fixed two-satellite constellation for more than 80% of the time.

The production of a ferrofluid container, intended for use in the KRAKsat (CubeSat type) satellite in space conditions, is presented. Mechanized laser beam welding for AISI 316L stainless steel test joint and container prototype was developed and tested. The welded test joints were examined by non-destructive visual, penetration and radiographic testing and destructive testing by macro-and microscopic examination, static tensile test, static bending test, and hardness measurements. The welded container prototype was examined by leak test, temperature-vacuum test and vibration test. Test joints' evaluation showed a proper selection of welding parameters and expected quality of joints. Austenitic microstructure with small δ-ferrite content in base materials, heat-affected zones, and welds guarantees sufficient mechanical properties for this part geometry. The tensile strength range of test joints was 687-729 MPa, hardness range was 140-200 HV3, and the bending angle was 180°. Welding of the prototype container and testing of tightness, resistance to temperature changes, and vibration were successful. Compliance with flywheel design and manufacturing requirements will enable the launch of a research satellite into orbit with such a wheel.

TURKSAT-3USAT is a three unit CubeSat developed for voice communication at low Earth orbit. The payload is a VHFIUHF transponder operating at amateur bands. The transponder and all other subsystems, except the stabilization, are doubled for redundancy. Where possible, both COTS systems and in-house development is employed. The energy is provided using lithium polymer batteries together with super capacitors. Satellite stabilization is accomplished using passive magnetic attitude control system with hysteresis rods. The TURKSAT-3USAT is launched from JSLC of China, on April 26, 2013.

This paper considers controlling the orientation of a rigid body using noncommutative rotation sequences. In particular, we consider a cube that is fixed at its center of mass and has two internal masses that can vibrate within the cube. The mass vibrations cause small angular oscillations about two orthogonal body-fixed axes. When the internal masses vibrate out of phase and at the same frequency, net rotation about the remaining body-fixed axis is achieved. The dynamic equations of motion for the system are derived and open-loop numerical simulations demonstrate the effects of amplitude, frequency, and phase difference of the mass vibrations. We then present two control techniques for single-axis orientation control. Closed-loop numerical simulations demonstrate the effectiveness of these control techniques.

GPS receivers are increasingly used in nanosatellite applications to provide the means for autonomous and accurate navigation in orbit. However, many of the available nanosatellite GPS receivers consume a large amount of power compared to the available power budget. This can be a problem when the main science instruments have high power requirements and GPS is only used to support the science mission. This paper presents the design and proposed integration and qualification of a GPS subsystem that uses commercial, low-power GPS components to support the science mission of the 3U CubeSat-based Aalto-1 nanosatellite. Aalto-1 has three science payloads which require accurate positioning: a Fabry-Pérot imaging spectrometer used for remote sensing, a radiation monitor used for studying the near-Earth radiation environment and an electrostatic plasma brake used for deorbiting the satellite at the end of the mission. Ground-based tracking with only one ground station or NORAD-provided two-line element sets cannot position the satellite accurately enough for the science payloads, and it was concluded that GPS navigation is the best option. The GPS receiver and antenna have a combined power budget of 160 mW. Navigation accuracy of less than 100 meters is needed by the scientific instruments of Aalto-1. Effects of antenna placement on GPS navigation performance are considered, as the antenna on Aalto-1 will have a limited view of the sky. The preliminary functional tests and simulations with a GPS signal simulator indicate that the subsystem is capable of navigating in orbit and providing the Aalto-1 mission with the required navigation capability.

Future space telescopes with diameter over 20 m will require new approaches: either high-precision formation flying or in-orbit assembly. We believe the latter holds promise as a potentially lower cost and more practical solution in the near term, provided much of the assembly can be carried out autonomously. To gain experience, and to provide risk reduction, we propose a combined mico/nano-satellite demonstration mission that will focus on the required optical technology (adaptive mirrors, phase-sensitive detectors) and autonomous rendezvous and docking technology (inter-satellite links, relative position sensing, automated docking mechanisms). The mission will involve two "3U" Cubesat-like nanosatellites ("MirrorSats") each carrying an electrically actuated adaptive mirror, and each capable of autonomous un-docking and redocking with a small central "15U" class micro/nano-satellite core, which houses two fixed mirrors and a boom-deployed focal plane assembly. All three spacecraft will be launched as a single 40kg microsatellite package. The spacecraft busses are based on heritage from Surrey's SNAP-1 and STRaND-1 missions (launched in 2000 and 2013 respectively), whilst the optics, imaging sensors and shape adjusting adaptive mirrors (with their associated adjustment mechanisms) are provided by CalTech/JPL. The spacecraft busses provide precise orbit and attitude control, with inter-satellite links and optical navigation to mediate the docking process. The docking system itself is based on the electromagnetic docking system being developed at the Surrey Space Centre (SSC), together with rendezvous sensing technology developed for STRaND-2. On orbit, the mission profile will firstly establish the imaging capability of the compound spacecraft before undocking, and then autonomously re-docking a single MirrorSat. This will test the docking system, autonomous navigation and system identification technology. If successful, the next stage will see the two MirrorSat spacecraft undock and re-dock to the core spacecraft in a linear formation to represent a large (but sparse) aperture for high resolution imaging. The imaging of stars is the primary objective, but other celestial and terrestrial targets are being considered. Teams at CalTech and SSC are currently working on the mission planning and development of space hardware. The autonomous rendezvous and docking system is currently under test on a 2D air-bearing table at SSC, and the propulsion and precision attitude control system is currently in development. Launch is planned for 2015. This paper details the mission concept, technology involved and progress to date.

TURKSAT-3USAT is a three unit CubeSat developed for voice communication at low Earth orbit. The payload is a VHFIUHF transponder operating at amateur bands. The transponder and all other subsystems, except the stabilization, are doubled for redundancy. Where possible, both COTS systems and in-house development is employed. The energy is provided using lithium polymer batteries together with super capacitors. Satellite stabilization is accomplished using passive magnetic attitude control system with hysteresis rods. The TURKSAT-3USAT is launched from JSLC of China, on April 26, 2013.

One of the objectives of the RANGE mission is to perform relative navigation using differential aerodynamics. The aerodynamic force coefficients, moment coefficients, and heating can be computed numerically, though this process is too computationally expensive to integrate directly into an orbit propagator. A surrogate model is developed to improve the modeling fidelity beyond a simple sphere or plate model without significantly increasing computational cost. The training points for this model come from an industry standard code for Direct Simulation Monte Carlo analysis. Detailed discussions of model development, validation, and results are included.

In this paper we present CubedOS, a lightweight application framework for CubeSat flight software. CubedOS is written in SPARKand proved free of certain classes of runtime errors. It consists of a collection of interacting , concurrent modules that communicate via message passing over a microkernel based on Ada's Ravenscar tasking model. It provides core services such as, for example, communication protocol processing and pub-lish/subscribe message handling. Application-specific modules can be added to provide both high level functions such as navigation and power management, as well as low level device drivers for mission-specific hardware.

TURKSAT-3USAT is a three unit CubeSat developed
for voice communication at low Earth orbit. The payload is a
VHFIUHF transponder operating at amateur bands. The
transponder and all other subsystems, except the stabilization,
are doubled for redundancy. Where possible, both COTS systems
and in-house development is employed. The energy is provided
using lithium polymer batteries together with super capacitors.
Satellite stabilization is accomplished using passive magnetic
attitude control system with hysteresis rods. The TURKSAT-
3USAT is launched from JSLC of China, on April 26, 2013.