Microsatellite Missions Research Articles

Earth-Venus Mission Analysis via Weak Capture and Nonlinear Orbit Control

Exploration of Venus is recently driven by the interest of the scientific community in understanding the evolution of Earth-size planets, and is leading the implementation of missions that can benefit from new design techniques and technology. In this work, we investigate the possibility to implement a microsatellite exploration mission to Venus, taking advantage of (i) weak capture, and (ii) nonlinear orbit control. This research considers the case of a microsatellite, equipped with a high-thrust and a low-thrust propulsion system, and placed in a highly elliptical Earth orbit, not specifically designed for the Earth-Venus mission of interest. In particular, to minimize the propellant mass, phase (i) of the mission was designed to inject the microsatellite into a low-energy capture around Venus, at the end of the interplanetary arc. The low-energy capture is designed in the dynamical framework of the circular restricted 3-body problem associated with the Sun-Venus system. Modeling the problem with the use of the Hamiltonian formalism, capture trajectories can be characterized based on their state while transiting in the equilibrium region about the collinear libration point L1. Low-energy capture orbits are identified that require the minimum velocity change to be established. These results are obtained using the General Mission Analysis Tool, which implements planetary ephemeris. After completing the ballistic capture, phase (ii) of the mission starts, and it is aimed at driving the microsatellite toward the operational orbit about Venus. The transfer maneuver is based on the use of low-thrust propulsion and nonlinear orbit control. Convergence toward the desired operational orbit is investigated and is proven analytically using the Lyapunov stability theory, in conjunction with the LaSalle invariance principle, under certain conditions related to the orbit perturbing accelerations and the low-thrust magnitude. The numerical results prove that the mission profile at hand, combining low-energy capture and low-thrust nonlinear orbit control, represents a viable and effective strategy for microsatellite missions to Venus.

Best practices on adopting open-source and commercial low-cost devices in small satellites missions

The NewSpace approach of utilizing CubeSats and smaller satellites for complex LEO missions that once were reserved only for bigger platform, paired with the requirement of enduring reliability and quality assurance on batches of tens of satellites, is paving the way for the use of commercial-off-the-shelf (COTS) components in micro-satellite missions. The recent global supply chain issues and extended lead times in the procurement of micro-controllers, memories and passive electronic components, caused by the pandemic and factories shutdowns, have resulted in a forced procurement diversification of several key elements, even in partially assembled products.GAUSS has been designing, manufacturing, and launching Small Satellites since early 2000.Its tenth satellite, UNISAT-7, was launched by a Soyuz-Fregat LV on March 22nd, 2021. Thanks to the vast experience gained on the application of commercial components in all the subsystems developed by the company, the typical overall costs for a complete platform are substantially lower compared to competitors, while retaining similar reliability.UNISAT-7, developed using lean-satellite methodologies, included more than 28 independent subsystems with their own telemetry, and most of them were developed in-house using ground COTS components; no space-grade products were employed, to save on overall mission costs. Some systems were devised from the start using open-source hardware, to expedite the prototyping process, while having the benefit of using readily available software, generally included with these platforms. Raspberry embedded devices were employed on the flight model for two remote sensing systems, to test how such systems, duly adapted to sustain space conditions, may perform and show reliability similar to commercial space systems worth thousand times the value.All systems have been activated and successfully verified for an entire year in LEO. They were hardened for vacuum, extreme temperatures, radiation, and launch stresses using common items found in general hardware stores.This work will delve into the open-source/COTS components used, the process employed to strengthen them, and the programmed flight firmware routines for maximizing embedded devices’ reliability in space.

Laboratory Study of Microsatellite Control Algorithms Performance for Active Space Debris Removal Using UAV Mock-Ups on a Planar Air-Bearing Test Bed

In this paper, a planar air-bearing test bed with unmanned aerial vehicles (UAV) was used to test a microsatellite motion control system. The UAV mock-ups were controlled by four ventilator actuators that imitated the satellite thrusters and provided the required acceleration vector in the horizontal plane, and torque along the vertical direction. The mock-ups moved almost without friction along the planar air-bearing test bed due to the air cushion between the test bed surface and the flat mock-up base. The motion of the mock-ups motion imitated the motion of satellites in the orbital plane. The problem of space debris can be solved using special microsatellite missions able to dock to space debris objects and change their orbit. In this paper, two control algorithms based on the virtual potentials approach and the State Dependent Ricatti Equation (SDRE) controller, were proposed for docking to a non-cooperative space debris object. The algorithms were tested in a laboratory facility, and the results are presented and analyzed, including their main features demonstrated during the laboratory study. It was shown that the SDRE-based control was faster, although the virtual potential-based control required less characteristic velocity.

An Earth-Mars microsatellite mission leveraging low-energy capture and low-thrust propulsion

Growing interest in deep-space exploration using microsatellites is leading the development of novel trajectory design techniques compatible with the constraints of this class of spacecraft. The mission design discussed in this paper examines the possibility of injecting a microsatellite into a desired orbit around Mars taking advantage of multibody dynamics, performing Earth and Moon flybys, which concludes in a low-energy capture at Mars, followed by nonlinear control for orbit injection maneuvers. If compared to a traditional Earth-Mars transfer, the proposed solution leads to relevant propellant savings and extended launch windows, but the capture orbit established has chaotic behavior and time-varying osculating orbital parameters which may not satisfy the mission requirements. Consequently, the microsatellite is capable of transferring from the original low-energy capture orbit to the desired operative one. For this phase, the microsatellite is assumed to operate using a low-thrust propulsion system, which is ignited with the intent of driving the microsatellite toward two operational orbits: (a) a quasi-synchronous repeating 2-SOL orbit and (b) a sun-synchronous orbit with altitude of 200 km. To do this, nonlinear orbit control is employed. Convergence toward the desired operational orbits is investigated, and can be guaranteed - using the Lyapunov stability theory, in conjunction with the LaSalle invariance principle - under certain conditions related to the orbit perturbing accelerations and the low-thrust magnitude. The numerical simulations prove that the combination of multiple-flyby transfer (simulated with real ephemeris, with the use of GMAT) and low-thrust nonlinear orbit control represents a viable and effective strategy for microsatellite missions to Mars.

AN OVERVIEW ON INNOVATIVE COMPOSITE MATERIALS EMPLOYED IN THE CONSTRUCTION OF MICROSATELLITES

The space industry has a growing interest in microsatellite missions, which are frequently employed in the defense, agriculture, business intelligence, in case of disasters, in communications. The paper presents recent findings from the literature regarding the structure and properties of composite materials, in order to establish the possibilities of their use in the development of microsatellites, in the context of a sustainable development. Innovative composites must have dimensional stability during exposure to thermal cycling in space, low degree of outgassing, high resistance to microcracking, UV radiation, atomic oxygen, proton irradiation and orbital debris.

Multidisciplinary design and optimization of expendable launch vehicle for microsatellite missions

PurposeThe capability to predict and evaluate various configurations’ performance during the conceptual design phase using multidisciplinary design analysis and optimization can significantly increase the preliminary design process’s efficiency and reduce design and development costs. This research paper aims to perform multidisciplinary design and optimization for an expendable microsatellite launch vehicle (MSLV) comprising three solid-propellant stages, capable of delivering micro-payloads in the low earth orbit. The methodology’s primary purpose is to increase the conceptual and preliminary design process’s efficiency by reducing both the design and development costs.Design/methodology/approachMultidiscipline feasible architecture is applied for the multidisciplinary design and optimization of an expendable MSLV at the conceptual level to accommodate interdisciplinary interactions during the optimization process. The multidisciplinary design and optimization framework developed and implemented in this research effort encompasses coupled analysis disciplines of vehicle geometry, mass calculations, aerodynamics, propulsion and trajectory. Nineteen design variables were selected to optimize expendable MSLV to launch a 100 kg satellite at an altitude of 600 km in the low earth orbit. Modern heuristic optimization methods such as genetic algorithm (GA), particle swarm optimization (PSO) and SA are applied and compared to obtain the optimal configurations. The initial population is created by passing the upper and lower bounds of design variables to the optimizer. The optimizer then searches for the best possible combination of design variables to obtain the objective function while satisfying the constraints.FindingsAll of the applied heuristic methods were able to optimize the design problem. Optimized design variables from these methods lie within the lower and upper bounds. This research successfully achieves the desired altitude and final injection velocity while satisfying all the constraints. In this research effort, multiple runs of heuristic algorithms reduce the fundamental stochastic error.Research limitations/implicationsThe use of multiple heuristics optimization methods such as GA, PSO and SA in the conceptual design phase owing to the exclusivity of their search approach provides a unique opportunity for exploration of the feasible design space and helps in obtaining alternative configurations capable of meeting the mission objectives, which is not possible when using any of the single optimization algorithm.Practical implicationsThe optimized configurations can be further used as baseline configurations in the microsatellite launch missions’ conceptual and preliminary design phases.Originality/valueSatellite launch vehicle design and optimization is a complex multidisciplinary problem, and it is dealt with effectively in the multidisciplinary design and optimization domain. It integrates several interlinked disciplines and gives the optimum result that satisfies these disciplines’ requirements. This research effort provides the multidisciplinary design and optimization-based simulation framework to predict and evaluate various expendable satellite launch vehicle configurations’ performance. This framework significantly increases the conceptual and preliminary design process’s efficiency by reducing design and development costs.

Design of fuel-saving lunar captures using finite thrust and gravity-braking

The emergence of microsatellite missions leaving low Earth orbits to pursue the exploration of the solar system is currently leading the development of new techniques of mission analysis. In fact, the small size of these satellites affects the performance of the propulsion systems, in particular because of the limited mass budget available for propellant stowage. This work focuses on the analysis of lunar microsatellite missions, proposing a method to design fuel-saving lunar captures. The solution is aimed at establishing the capture conditions taking advantage of the gravity-braking effect of the Moon, before injecting the spacecraft into a lunar capture orbit. The maneuvers are developed in the form of guidance laws, implemented based on primer vector theory and on the definition of a quasi-integral of motion for the elliptic restricted 4-body problem. The method is verified, by means of numerical analysis, on a 6U CubeSat equipped with a monopropellant thruster and considering deployment conditions compatible with microsatellite lunar missions. The achievement of lunar capture in a few days from the deployment, with propellant usage on the order of tenths of grams, corresponding to ΔV on the order of some meters per second, indicates the suitability of the method to expand the possible profiles of microsatellite missions to the Moon.

Numerical Analysis and Modelling of a 100 N Hypergolic Liquid Bipropellant Thruster

This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100 N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.

Assessment of Micropropulsion System Unifying Water Ion Thrusters and Water Resistojet Thrusters

This study proposes a micropropulsion system unifying ion thrusters and resistojet thrusters and assessing that propulsive capability. The remarkable features of the system are the usage of water propellant and unification of the two types of thrusters by the single propellant. Water has been regarded as an attractive propellant in the view points of safety, availability, handling ability, low molecular mass, and future procurement in space. A multimode propulsion system is an attractive solution for the increasing demand for nano-/microsatellite missions. The proposal is to use microwave discharge water ion thrusters, tolerant for oxidization by water, and low-temperature water resistojet thrusters, enabling reuse of the waste heat. As a result of the assessment, it was expected that the propulsion system would have 3U size () and 3.70 kg mass, which realize in total a 6U and 10 kg satellite with 3U and 6 kg satellite bus system. The ion thruster would provide the maximum of by 47 W system power and the resistojet thruster would have 3.80 mN thrust and 72 s specific impulse by 19.4 W. Additionally, reuse of the waste heat from ion-thruster power supplies would enable the simultaneous operations of the two thrusters even at 50 W, which is almost the same power as the single ion thruster operation.

Laser-Ignited Micromotor Using Multiple Stacked Solid Propellant Pellets

This study proposes the laser-ignited micromotor (LIMO), which is a new propulsion system for CubeSats to overcome the problems associated with conventional chemical propellants. The solid propellant pellets are stacked and stored in the micromotor. This concept makes it possible to design the total impulse of the micromotor in a flexible manner at a low cost and in a short time, depending on microsatellite missions. In addition, a laser-ignition system is integrated with the micromotor to address the problems of ignition delay and the difficulty in cutting the high-current feed lines of the ignition system. Firing tests are conducted in a vacuum chamber using four types of micromotors with different characteristics; that is, each type has a different number of solid propellant pellets and a different ignition system. Total impulses of 184 and are achieved by the two types of LIMOs with 98 and 140 stored pellets, respectively, with a burning rate of . Ignition delay is improved from 0.37 to 0.04 s by using the laser-ignition system. In addition, an engineering model of the LIMO is designed.

Simulation of Gamma-Ray and Neutron Spectrometers for Microsatellite Missions

Microsatellites have recently opened windows of frequent and low cost missions for planetary exploration. The performance of gamma-ray and neutron spectrometers on future microsatellite missions is simulated to assess the possibility of observation of hydrogen and major elements, given their concentration on the observation target. The measured elemental abundance will provide important geological constraints, and some of them may serve as space resources. Four different types of target bodies with various hydrogen concentrations in the range of 0 - 20,000 ppm are assumed as target compositions; Earth’s core, C-type, S-type and Martian meteorites. Gamma-ray and neutron emission rates show unique footprints that are related to the different elemental compositions. The starting point is the solid angle subtended between observation target and spectrometers that allow estimating the gamma-ray and neutron count rates emitted by the celestial bodies. In this work, three types of gamma-ray detectors; high-purity germanium (HPGe), CeBr3 and LaBr3(Ce), a neutron spectrometer combining a lithium glass scintillator with a boron loaded plastic scintillator and a dual mode spectrometer Cs2LiYCl6(Ce) (CLYC) are simulated, focusing on their observation backgrounds as a model case for microsatellite based measurements. The background count level of both gamma-ray (except for the LaBr3 detector) and neutron count rates was negligible under these particular conditions. The gamma-ray detectors were compared by the figure of merit, which was determined by their efficiency and energy resolution. It was found that each detector has unique advantages. The HPGe detector has the highest figure of merit due to its excellent energy resolution, whereas the CLYC detector is low in weight and power consumption due to its dual sensitivity to gamma-ray and neutron. The CeBr3 detector is an intermediate choice. The neutron count rates are calculated separately in three energy ranges, i.e. , thermal ( 500 keV), as a function of the hydrogen concentration in the 0 - 20,000 ppm range. The thermal and epithermal neutron count rates are found to decrease with hydrogen concentration, while the fast neutron count rate increases with the target average atomic mass. The optimal detector should be decided by the mission restraints on mass, power consumption, and heat thermal design.

Propulsion options for very low Earth orbit microsatellites

The growing competitiveness in the commercial space market has raised the interest in operating small spacecraft at very low altitudes. To make this feasible, the space industry has started developing propulsion options tailored specifically to these platforms. This paper presents a review of emerging micropropulsion technologies and evaluates their applicability to microsatellite missions in the altitude range 250–500km. The results of the proposed analysis are demonstrated on two different remote sensing applications.

Simulation of Fault-Tolerant Space Systems Based on COTS Devices With GPSS

In microsatellite missions devoted to remote sensing, the cost-reliability issue is an important constraint. This paper presents a low-cost methodology to simulate the reliability of a system for the development of fault-tolerant space systems. This methodology does not rely on special facilities or expensive equipment and is focused on the onboard data handling (OBDH) system for microsatellite's remote sensing payloads with commercial off-the-shelf components. The method presented uses the general-purpose simulation system language to simulate the system's reliability when fault-tolerance techniques (FTTs) are applied to field-programmable gate arrays. The objective of the proposed methodology is to validate the mathematical model corresponding to the system's reliability or to compute the reliability when there is no mathematical model. Therefore, system-level decisions can be made, in order to select the most suitable components and develop the systems for microsatellites. Specifically, the OBDH computers must have a high reliability as a mandatory requirement for microsatellites. Moreover, the simulation provides statistics to support decisions regarding the necessary FTT to be implemented.

Jitter reduction of a reaction wheel by management of angular momentum using magnetic torquers in nano- and micro-satellites

Nowadays, nano- and micro-satellites, which are smaller than conventional large satellites, provide access to space to many satellite developers, and they are attracting interest as an application of space development because development is possible over shorter time period at a lower cost. In most of these nano- and micro-satellite missions, the satellites generally must meet strict attitude requirements for obtaining scientific data under strict constraints of power consumption, space, and weight. In many satellite missions, the jitter of a reaction wheel degrades the performance of the mission detectors and attitude sensors; therefore, jitter should be controlled or isolated to reduce its effect on sensor devices. In conventional standard-sized satellites, tip-tilt mirrors (TTMs) and isolators are used for controlling or isolating the vibrations from reaction wheels; however, it is difficult to use these devices for nano- and micro-satellite missions under the strict power, space, and mass constraints. In this research, the jitter of reaction wheels is reduced by using accurate sensors, small reaction wheels, and slow rotation frequency reaction wheel instead of TTMs and isolators. The objective of a reaction wheel in many satellite missions is the management of the satellite’s angular momentum, which increases because of attitude disturbances. If the magnitude of the disturbance is reduced in orbit or on the ground, the magnitude of the angular momentum that the reaction wheels gain from attitude disturbances in orbit becomes smaller; therefore, satellites can stabilize their attitude using only smaller reaction wheels or slow rotation speed, which cause relatively smaller vibration. In nano- and micro-satellite missions, the dominant attitude disturbance is a magnetic torque, which can be cancelled by using magnetic actuators. With the magnetic compensation, the satellite reduces the angular momentum that the reaction wheels gain, and therefore, satellites do not require large reaction wheels and higher rotation speed, which cause jitter. As a result, the satellite can reduce the effect of jitter without using conventional isolators and TTMs. Hence, the satellites can achieve precise attitude control under low power, space, and mass constraints using this proposed method. Through the example of an astronomical observation mission using nano- and micro-satellites, it is demonstrated that the jitter reduction using small reaction wheels is feasible in nano- and micro-satellites.

Compensation of time-variable magnetic moments for a precise attitude control in nano- and micro-satellite missions

Theses days, many nano- and micro-satellites are applied to several astronomy and remote sensing missions. In order to achieve mission requirements, these satellites must control the attitude precisely. A magnetic disturbance is one of the dominant sources of attitude disturbances. Therefore, this disturbance should be canceled in-orbit or on the ground to achieve the attitude strict requirements. This paper presents the effect of the magnetic disturbance to the attitude in nano- and micro-satellite missions and the sources of the residual magnetic moment of the satellites, which causes the magnetic disturbance. Then, the paper proposes a method to compensate the residual magnetic moment both in-orbit and in the design phase of the satellites. The research also focused on a time-varying residual magnetic moment. Finally, the method is applied to a micro-astrometry satellite as an example.

The New Fault Tolerant Onboard Computer for Microsatellite Missions

This paper describes an onboard computer with dual processing modules. Each processing module is composed of 32 bit ARM reduced instruction set computer processor and other commercial-off-the-shelf devices. A set of fault handling mechanisms is implemented in the computer system, which enables the system to tolerate a single fault. The onboard software is organized around a set of processes that communicate among each other through a routing process. Meeting an extremely tight set of constraints that include mass, volume, power consumption and space environmental conditions, the fault-tolerant onboard computer has excellent data processing capability that can meet the erquirements of micro-satellite missions.

06/00941 Theoretical expression of rate of rise of recovery voltage across a circuit breaker connected with fault current limiter: Calixte, E. et al. Electric Power Systems Research, 2005, 75, (1), 1–8.

This paper describes an onboard computer with dual processing modules. Each processing module is composed of 32 bit ARM reduced instruction set computer processor and other commercial-off-the-shelf devices. A set of fault handling mechanisms is implemented in the computer system, which enables the system to tolerate a single fault. The onboard software is organized around a set of processes that communicate among each other through a routing process. Meeting an extremely tight set of constraints that include mass, volume, power consumption and space environmental conditions, the fault-tolerant onboard computer has excellent data processing capability that can meet the erquirements of micro-satellite missions.

Ground solar radius survey in view of microsatellite missions

For the last 25 years, ground time series of the solar radius have shown (different) apparent variations according to different instruments. The origin of these variations may search in the observer, the instrument, the atmosphere and up to the sun. Ground instruments are automated to reduce the “personnal equation” and place that origin in the atmosphere and/or in the sun. Astrometric satellites scheduled at the end of this decade will perform non ambiguous diameter measurements. A survey of the Solar radius has been initiated in 1975 by Francis Laclare, at the Calern site of the Observatoire de la Côte d’Azur, which have been chosen for hosting the ground segment of the Centre National d’Etudes Spatiales (CNES) Microsatellite PICARD mission, to be launched in 2008. This reference series was obtained by Visual observations of the Sun, with a Solar Astrolabe whose metrological character has to be stressed. Considering the Visual series results, we have compared the solar diameter variations with the solar activity cycle, and we found an opposite phase, for the whole series and at the different times of the cycles. Parallel to that series, CCD measurements were made with the same instrument and gave results which are perfectly blended together, within our quoted uncertainties. Located next to the Solar Astrolabe, DORAYSOL (Définition et Observation du Rayon Solaire) is a second generation instrument, which keeps the major features of the design of its predecessor and, which is designed to increase the number of CCD measurements and to be eventually automated. Since 1999, both series overlap correctly within our quoted uncertainties. Some information is added to explain the pattern of the PICARD mission ground segment, next to those instruments at Calern Observatory, as well as the international network intended to carry out the Sun’s Radius ground survey ( R2S3: Réseau de Suivi au Sol du Rayon Solaire).

Study on the on-Board Computer System Based on ARM Processor

The new fault-tolerant onboard computer(OBC) with dual processing modules is presented to improve the micro-satellite data handling. Each processing module is composed of 32-bit ARM processor. Using the fault tolerance method, The OBC's hardware structure is implemented based on commercial-off-the-shelf (COTS) devices. As well as, a detail analysis of the fault handling mechanism and the software architecture is given. Considering the demanding of the extremely tight constraints on mass, volume, power consumption and space environmental conditions, the new fault tolerant onboard computer's data handling is enough to meet the micro-satellite missions.

Microsatellite missions to conduct midlatitude studies of equatorial ionospheric plasma bubbles

Two missions presently under development by the United States Air Force Academy (USAFA), FalconSAT-2 and FalconSAT-3, include mission scientific objectives targeting the study of ionospheric F region plasma density depletions and topside bubbles associated with the so-called Equatorial Spread F (ESF) phenomena. The Miniature Electrostatic Analyzer (MESA), a USAFA-designed patch sensor that measures differential energy fluxes of electrons from 0.05 to 13 eV in six channels, is the primary experiment aboard FalconSAT-2, a 25-kg microsatellite intended for launch into an International Space Station (ISS) orbit via the Space Shuttle. Because the orbit will be approximately 360 km in altitude and of 52° inclination, FalconSAT-2 observations will complement those of low latitude missions (e.g., C/NOFS) and polar latitude, higher altitude missions (e.g., DMSP). Realistic internal magnetic field models demonstrate that field lines with apex heights of 1500 km (representing the upper altitude limit of equatorial plasma bubbles) may intersect the orbit plane at dip latitudes greater than 35°. Thus, FalconSAT-2 will be able to observe plasma depletions that have propagated poleward along the field lines and lower in altitude, depletions that may not be observed with the high altitude DMSP and the low latitude C/NOFS. Additionally, there may be opportunities for FalconSAT-2 to make simultaneous multipoint in situ measurements of large-scale plasma bubbles with other low altitude satellites, such as C/NOFS and DMSP. We will present a statistical analysis of the probability of making such measurements using nominal orbital parameters of the relevant spacecraft. Finally, a description of the FalconSAT-3 follow-on mission, including scientific objectives associated with seeking kinetic effects, is presented.