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Magnetic properties of a 3U CubeSat with electric propulsion
... They can also be launched as secondary payloads, utilizing excess capacity on larger launch vehicles. Moreover, miniaturized satellites can be designed more affordably and are easier to mass-produce (Levchenko et al., 2018a;Wolfgang et al., 2023;Lian et al., 2023). Another primary reason for developing small satellites is the opportunity to perform tasks that large satellites cannot, such as constellations for low data rate communications, data collection from multiple points on the Earth's surface, in-orbit inspection of larger satellites, university research, testing or qualification of new https://doi.org/10.1016/j.asr.2024.05.039 0273-1177/Ó 2024 COSPAR. ...
Based on its successful CubeSat mission PEGASUS, the University of Applied Sciences Wiener Neustadt (FHWN) is preparing its new CubeSat mission called CLIMB. CLIMB is a 3U CubeSat that will be launched to a low, circular orbit of about 500 km. Using a Field Emission Electric Propulsion (FEEP) system commercialized by the company ENPULSION, the satellite will be lifted to an elliptical orbit with its apogee around 1000 km – well inside the inner Van Allen belt. During its 1.5 yearlong ascent and its operation in the Van Allen belt, the satellite will continuously monitor the space radiation with a RadFET dosimeter payload and the impact on CLIMB’s subsystems. Comparisons with radiation testing on ground will allow the assessment of the capability of ground tests to predict effects of space radiation on CubeSat subsystems. The operation of the propulsion system will raise the satellite’s apogee on average 16 times a day. A comprehensive analysis has been conducted to assess its collision probability throughout its mission time. Using various tools, provided by ESA (CROC, MASTER and the DRAMA ARES python package), the collision probability for the entire mission duration (~3 years) was calculated to be 3.38 × 10-5, i.e. a magnitude smaller than the requested probability of 10-4. The second payload of CLIMB is an anisotropic magnetoresistance (AMR) magnetometer with a, for CubeSats high, sensitivity of about 10 nT RMS. The first results of measurements with this COTS based magnetometer are presented as well as experimental assessments of the satellite’s magnetic cleanliness. The benign thermal conditions on CubeSats operating close to Earth are complicated by the relatively high-power propulsion system onboard CLIMB. Detailed numerical analysis (ANSYS, ESATAN) and experimental verifications resulted in the identification of possible methods to deal with up to 18 W of dissipated electric power. The main heat sources are the thruster and the battery unit, during thruster operation
Magnetometers are widely used on satellites for both attitude sensing and scientific observations. Spaceborne magnetometers have enabled the creation of accurate maps of Earth’s magnetic fields. However, these models have limited spatial and temporal resolution, and therefore are much less accurate in locations with fast or localized magnetic perturbations. Such perturbations can be particularly problematic near Earth’s poles where field aligned currents come close to the surface of the Earth and are concentrated near satellites in LEO. Science missions which need to know the local magnetic field in the polar regions need to bring their own high-fidelity magnetic sensors. The AERO-VISTA mission comprises a pair of 6U CubeSats which will determine the propagation modes and directions of high frequency (400 kHz–5 MHz) waves in Earth’s ionosphere in the presence of Earth’s aurorae. This mission science requires accurate in-situ magnetic sensing of auroral currents for RF measurement context. This thesis details the design, integration, and testing of the magnetic sensors in the AERO-VISTA Auxiliary Sensor Package (ASP). We discuss the estimation of spacecraft self-interference and implement an informal magnetic interference control process. We present some simple ground testing strategies for magnetic screening of components and measurement of spacecraft self-interference. We evaluate the performance and non-ideal effects of our selected anistropic magnetoresistive (AMR) 3-axis magnetometer. We create a measurement equation, which together with regression techniques, allows for calibration to better than 100 nT repeatability despite non-ideal effects, meeting AERO-VISTA’s requirements. This calibration strategy is extended to include current path and material interference effects. We describe the detailed design of the magnetic sensing system, including the electronics, mechanical design, and software of the ASP. Without self-interference effects, this design has a noise floor better than 10 nTrms.
Magnetic cleanliness and in-flight identification and rejection of residual magnetization effects for a magnetically controlled 3U CubeSat are considered. The routine we propose starts from accurate acquisition of all the components' magnetic fields at the stage when the engineering model of the spacecraft is ready. This allows making a map of the assembled CubeSat's internal magnetic fields using electromagnetic simulation and analysis software. Such maps can be produced for each operational regime of the spacecraft and compose a magnetic atlas in accordance with the concept of operations. By feeding the maps to the extended Kalman filter, which processes the magnetometer data and estimates the residual magnetic dipole and the magnetometer bias along with the state variables, we ensure a better initial guess for the disturbances, which is crucial for the filter's convergence. It is shown to be of importance, whenever the level of magnetic disturbances abruptly changes as the spacecraft switches between the regimes.
The Electron Loss and Fields Investigation with a Spatio-Temporal Ambiguity-Resolving option (ELFIN-STAR, or heretoforth simply: ELFIN) mission comprises two identical 3-Unit (3U) CubeSats on a polar (∼93∘ inclination), nearly circular, low-Earth (∼450 km altitude) orbit. Launched on September 15, 2018, ELFIN is expected to have a >2.5 year lifetime. Its primary science objective is to resolve the mechanism of storm-time relativistic electron precipitation, for which electromagnetic ion cyclotron (EMIC) waves are a prime candidate. From its ionospheric vantage point, ELFIN uses its unique pitch-angle-resolving capability to determine whether measured relativistic electron pitch-angle and energy spectra within the loss cone bear the characteristic signatures of scattering by EMIC waves or whether such scattering may be due to other processes. Pairing identical ELFIN satellites with slowly-variable along-track separation allows disambiguation of spatial and temporal evolution of the precipitation over minutes-to-tens-of-minutes timescales, faster than the orbit period of a single low-altitude satellite (Torbit ∼ 90 min). Each satellite carries an energetic particle detector for electrons (EPDE) that measures 50 keV to 5 MeV electrons with
Δ
E/E < 40% and a fluxgate magnetometer (FGM) on a ∼72 cm boom that measures magnetic field waves (e.g., EMIC waves) in the range from DC to 5 Hz Nyquist (nominally) with <0.3 nT/sqrt(Hz) noise at 1 Hz. The spinning satellites (Tspin
∼
3 s) are equipped with magnetorquers (air coils) that permit spin-up or -down and reorientation maneuvers. Using those, the spin axis is placed normal to the orbit plane (nominally), allowing full pitch-angle resolution twice per spin. An energetic particle detector for ions (EPDI) measures 250 keV - 5 MeV ions, addressing secondary science. Funded initially by CalSpace and the University Nanosat Program, ELFIN was selected for flight with joint support from NSF and NASA between 2014 and 2018 and launched by the ELaNa XVIII program on a Delta II rocket (with IceSatII as the primary). Mission operations are currently funded by NASA. Working under experienced UCLA mentors, with advice from The Aerospace Corporation and NASA personnel, more than 250 undergraduates have matured the ELFIN implementation strategy; developed the instruments, satellite, and ground systems and operate the two satellites. ELFIN's already high potential for cutting-edge science return is compounded by concurrent equatorial Heliophysics missions (THEMIS, Arase, Van Allen Probes, MMS) and ground stations. ELFIN's integrated data analysis approach, rapid dissemination strategies via the SPace Environment Data Analysis System (SPEDAS), and data coordination with the Heliophysics/Geospace System Observatory (H/GSO) optimize science yield, enabling the widest community benefits. Several storm-time events have already been captured and are presented herein to demonstrate ELFIN's data analysis methods and potential. These form the basis of on-going studies to resolve the primary mission science objective. Broad energy precipitation events, precipitation bands, and microbursts, clearly seen both at dawn and dusk, extend from tens of keV to >1 MeV. This broad energy range of precipitation indicates that multiple waves are providing scattering concurrently. Many observed events show significant backscattered fluxes, which in the past were hard to resolve by equatorial spacecraft or non-pitch-angle-resolving ionospheric missions. These observations suggest that the ionosphere plays a significant role in modifying magnetospheric electron fluxes and wave-particle interactions. Routine data captures starting in February 2020 and lasting for at least another year, approximately the remainder of the mission lifetime, are expected to provide a very rich dataset to address questions even beyond the primary mission science objective.
Orbital capability is a decisive step forward for nanosatellites in general and CubeSats in particular. Although trajectory maneuvers and their implementation have been thoroughly studied for classical satellites, the high level of constraints on CubeSats in terms of mass, volume and power, makes the transition delicate. Orbit, attitude and power control subsystems available for this format limit too optimistic performance available in literature. To verify this hypothesis, we simulate trajectory maneuvers in Earth orbit with representative CubeSat hardware and software. A low-thrust trajectory solver based on classical orbital elements from the literature is adapted to our context. A home-made attitude control simulation tool is coupled to include both control and perturbative dynamics. Increases in time and propellant consumption of more than 100 % are caused by thrust direction errors such as misalignments and attitude control limitations, sometimes leading to mission loss. These results highlight an important increase in complexity for the CubeSat format that is not covered by the usual approach. Such limitations should be considered from the very start of the design of a nanosatellite mission with trajectory modification requirements.
CubeSats are being increasingly utilized for demanding astronomical and Earth observation applications where precise pointing and stability are critical requirements. Such precision is difficult to achieve in the case of CubeSats, mainly because of their small moment of inertia, this means that even small disturbance torques, such as those due to a residual magnetic moment, can still be an issue and have a significant effect on the attitude of spacecraft, when a high degree of stability is required. Recently, a PhD research program has been undertaken at Surrey University, to investigate the magnetic characteristics of CubeSats, it has been found that the disturbances may be mitigated by good engineering practice, in terms of minimizing current-loop areas and reducing the use of permeable materials. This paper discusses the dominant source nanosatellites disturbances and presents a survey and a short description of magnetic cleanliness techniques to minimize the effect of the residual magnetic field. It is mainly intended to supply a guide for CubeSat community to design future CubeSats with improved attitude stability
The IFM Nano Thruster is a variable specific impulse electrostatic thruster based on Field Emission Electric Propulsion (FEEP), in which a liquid propellant is electrostatically extracted and accelerated to high exhaust velocity. The core element of this propulsion technology is a passively fed, porous ion emitter consisting of 28 sharp emitter tips. This emitter technology has been developed and qualified over decades at FOTEC and the Austrian Institute of Technology, and has recently been adapted for the use as main propulsion system in Nano-and Small-satellites. The resulting IFM Nano Thruster occupies approximately 0.8 U and can be operated between 10 and 40W, resulting in thrust of up to 0.35mN. The thruster can be operated at specific impulse levels between 2000s and 6000s, adapting to mission needs as well as power availability, allowing for significant throttling capability between a couple of µN and 0.5mN. Due to the high specific impulse and high propellant density, the thruster can produce total impulses between 5000Ns and beyond 12000Ns when operated at specific impulses at 2000s and 5000s respectively. The first IFM Nano Thruster has been successfully integrated into a commercial 3U CubeSat in 2017 after undergoing environmental testing, and was launched in January 2018 for a first in-orbit demonstration (IOD). This IOD represents the first instance of a liquid metal FEEP thruster to be operated in space as a primary propulsion system. This paper will present the FEEP thruster principle and experimental characterization, including the first in-orbit test results.
Magnetic effects of satellites should be carefully studied starting from the design phase by creating a stringent magnetic cleanliness program. This program is required to maintain DC magnetic compatibility between units and reduce the interaction of satellite magnetic field with external magnetic field resulting parasitic torques. DC magnetic compatibility is especially important for satellites aiming scientific missions requiring sensitive magnetic equipment. Magnetic behavior of satellites and its units can be understood by test and analysis. Although testing is the most reliable method to evaluate magnetic assessment quantitatively, analysis becomes the first step because of cost and time restrictions. In this study, common analysis methods such as centered dipole approximation, multiple dipole modeling, spherical harmonics methods are discussed for their powerful and weak points. © 2016 Wiley Periodicals, Inc. Int J RF and Microwave CAE, 2016.
Formations of small satellites offer promising perspectives due to improved temporal and spatial coverage and resolution at reasonable costs. The UWE-program addresses in-orbit demonstrations of key technologies to enable formations of cooperating distributed spacecraft at pico-satellite level. In this context, the CubeSat UWE-3 addresses experiments for evaluation of real-time attitude determination and control.
UWE-3 introduces also a modular and flexible pico-satellite bus as a robust and extensible base for future missions. Technical objective was a very low power consumption of the COTS-based system, nevertheless providing a robust performance of this miniature satellite by advanced microprocessor redundancy and fault detection, identification and recovery software. This contribution addresses the UWE-3 design and mission results with emphasis on the operational experiences of the attitude determination and control system.
A prototype of a Service Oriented Spacecraft Magnetometer (SOSMAG) is being developed for European Space Agency's Space Situational Awareness program, which shall serve as a ready-to-use space weather monitoring system to be mounted on a variety of different spacecraft built without a magnetic cleanliness program. The separation between the natural magnetic field in space and the artificial magnetic field generated by the spacecraft is one of the key issues for its successful performance. The SOSMAG design is based on two types of magnetic sensors. One or two low-noise fluxgate sensors for the required high measurement accuracy will be mounted along a boom of length 1 m, and two anisotropic magnetoresistance (AMR)-based sensors will be used within the spacecraft for detection and characterization of its magnetic disturbances. This paper presents the design and performance of the AMR magnetometer that contains an application-specified integrated circuit (ASIC) and a hybrid sensor as core elements. The magnetometer front-end ASIC, originally developed for National Aeronautics and Space Administration's Magnetospheric Multiscale mission, is based on a fourth-order sigma-delta A/D converter design. All the active electronics needed for the readout of the AMR sensor and its digitization, as well as for digitizing the magnetometer's housekeeping data are part of the ASIC. Each of the three-axis hybrid sensors consists of a ceramics circuit board with three HMC1021 sensor elements, one bypass capacitor, a temperature sensor, and a MOSFET driver mounted in chip-on-board technology. A novel approach is used to set and reset the AMR elements. It is done in a narrow-pulsed way at 32 kHz with a pulsewidth of only 238 ns, thus minimizing the required excitation power to 30 mW in total.
The paper describes a sophisticated and realistic control and prediction method for the magnetic cleanliness of spacecraft, covering all phases of a project till the final system test. From the first establishment of the so-called magnetic moment allocation list the necessary boom length can be determined. The list is then continuously updated by real unit test results with the goal to ensure that the magnetic cleanliness budget is not exceeded at a given probability level. A complete example is described. The synthetic spacecraft modeling which predicts only quite late the final magnetic state of the spacecraft is also described. Finally, the most important cleanliness verification, the spacecraft system test, is described shortly with an example. The emphasis of the paper is put on the magnetic dipole moment allocation method.
The rapid pace of replacing fossil fuel propelled transport by electric vehicles is critically dependent on high-performing, high energy density batteries. Optimal and safe use of existing battery cells and development of much-needed novel battery chemistries and geometries require a multitude of diagnostic and monitoring tools. While structural and chemical information is readily extracted through a host of imaging techniques, non-invasive functional detection of interior battery processes remains limited. Here we introduce sensitive magnetometry performed outside the battery, revealing internal current distribution. As a key application, we use a sensor array to image the internal current flow of a pouch cell cycling between charge states. We find good agreement between measured and modelled fields with sufficient resolution to detect percent-level deviations around high current density areas. This opens the path towards rapid and reliable assessment throughout the battery life cycle, from battery development and manufacturing quality assurance to optimised use.
Traditionally, the space industry produced large and sophisticated spacecraft handcrafted by large teams of engineers and budgets within the reach of only a few large government-backed institutions. However, over the last decade, the space industry experienced an increased interest towards smaller missions and recent advances in commercial-off-the-shelf (COTS) technology miniaturization spurred the development of small spacecraft missions based on the CubeSat standard. CubeSats were initially envisioned primarily as educational tools or low cost technology demonstration platforms that could be developed and launched within one or two years. Recently, however, more advanced CubeSat missions have been developed and proposed, indicating that CubeSats clearly started to transition from being solely educational and technology demonstration platforms to offer opportunities for low-cost real science missions with potential high value in terms of science return and commercial revenue. Despite the significant progress made in CubeSat research and development over the last decade, some fundamental questions still habitually arise about the CubeSat capabilities, limitations, and ultimately about their scientific and commercial value. The main objective of this review is to evaluate the state of the art CubeSat capabilities with a special focus on advanced scientific missions and a goal of assessing the potential of CubeSat platforms as capable spacecraft. A total of over 1200 launched and proposed missions have been analyzed from various sources including peer-reviewed journal publications, conference proceedings, mission webpages as well as other publicly available satellite databases and about 130 relatively high performance missions were downselected and categorized into six groups based on the primary mission objectives including “Earth Science and Spaceborne Applications”, “Deep Space Exploration”, “Heliophysics: Space Weather”, “Astrophysics”, “Spaceborne In Situ Laboratory”, and “Technology Demonstration” for in-detail analysis. Additionally, the evolution of CubeSat enabling technologies are surveyed for evaluating the current technology state of the art as well as identifying potential areas that will benefit the most from further technology developments for enabling high performance science missions based on CubeSat platforms.
Nano-satellites provide space access to broader range of satellite developers and attract interests as an application of the space developments. These days several new nano-satellite missions are proposed with sophisticated objectives such as remote-sensing and observation of astronomical objects. In these advanced missions, some nano-satellites must meet strict attitude requirements for obtaining scientific data or images. For LEO nano-satellite, a magnetic attitude disturbance dominates over other environmental disturbances as a result of small moment of inertia, and this effect should be cancelled for a precise attitude control. This research focuses on how to cancel the magnetic disturbance in orbit. This paper presents a unique method to estimate and compensate the residual magnetic moment, which interacts with the geomagnetic field and causes the magnetic disturbance. An extended Kalman filter is used to estimate the magnetic disturbance. For more practical considerations of the magnetic disturbance compensation, this method has been examined in the PRISM (Pico-satellite for Remote-sensing and Innovative Space Missions). This method will be also used for a nano-astrometry satellite mission. This paper concludes that use of the magnetic disturbance estimation and compensation are useful for nano-satellites missions which require a high accurate attitude control.
Ecss-e-hb-20-07a - electromagnetic compatibility handbook. ESA Requirements and Standards Division
- Esa
Auroral Reconstruction CubeSwarm - ARCS: a NASA Heliophysics mission concept for decoding the aurora
- Lynch
Small Satellite Design for High Sensitivity Magnetic Measurements. Master’s thesis Luleå
- N S Janes
Auroral Reconstruction CubeSwarm - ARCS: a NASA Heliophysics mission concept for decoding the aurora
- K A Lynch
- J W Labelle
- J Lyon