Electrospray Propulsion Research Articles

Electrical propulsion systems for satellites: a review of current technologies and future prospects

Electrical propulsion systems have revolutionized satellite technology by offering greater efficiency, longer mission durations, and increased maneuverability compared to traditional chemical propulsion systems. This review explores current technologies and future prospects in the field of electrical propulsion systems for satellites. The discussion begins with an overview of existing technologies, including ion propulsion systems, Hall effect thrusters, and pulsed plasma thrusters. Each technology's principles of operation, advantages, limitations, and notable applications are examined. The review delves into the future prospects of electrical propulsion systems, exploring advanced concepts such as magnetoplasmadynamic thrusters, variable specific impulse magnetoplasma rockets, and electrospray propulsion systems. Additionally, miniaturization and efficiency improvements, as well as sustainable and green propulsion alternatives, are discussed. Challenges and opportunities facing the field are addressed, including technical hurdles like power generation and management, thruster lifespan, and regulatory and economic considerations such as policy frameworks and market dynamics. In conclusion, the review underscores the critical role of continued research and development in electrical propulsion systems for satellites. As the demand for more capable and sustainable satellite missions grows, advancements in propulsion technology will be essential in meeting these evolving needs and pushing the boundaries of space exploration.

Emission and performance characterization of ionic liquids for an externally wetted electrospray thruster

Ionic liquids are the base for electrospray propulsion, yet few have been explored. The performance of an externally wetted electrospray thruster is characterized for seven different ionic liquids: 1-ethyl-3-methylimidazolium tetrafluoroborate (EMI-BF4), dicyanamide (EMI-DCA), bistrifluoromethylsulfonylimide (EMI-Im), thiocyanate (EMI-SCN), ethylsulphate (EMI-EtOSO3), triflate (EMI-OTf), and trifluoromethanesulfonate (EMI-TFA). This research marks the first time that the latter four liquids are reported for externally wetted emitters. A complete thruster assembly has been tested with each of the propellants. Using time-of-flight mass spectrometry and a Faraday cup for beam divergence, the performance of these ionic liquids has been characterized, assessing thrust, specific impulse and thrust efficiency. EMI-BF4, EMI-DCA and EMI-SCN operated in pure ionic regime while EMI-Im, EMI-EtOSO3, EMI-OTf and EMI-TFA operated in mixed regime.

A novel annular slit-type emitter developed for multi-jet electrospray propulsion

Electrospray thrusters employ ionization in the liquid phase to produce and propel streams of molecular ions or highly charged droplets at significant velocities. In this study, we developed a novel annular slit-type emitter for electrospray and investigated its operational modes under varying applied potentials in both open atmosphere and vacuum conditions. To assess the performance of the annular slit-type emitter in comparison to the conventional capillary-type emitter, benchtop electrospray experiments were conducted using water and glycerin as working fluids for both emitter types. The study examined the formation of the Taylor cone, cone-to-jet transition, stable jet, whipping jet, and multi-jet, along with their dependence on fluid viscosity and electric potential for both emitter designs. Clear distinctions in hydrodynamic mode, drop-to-cone mode, and cone-to-jet transition mode were observed between the two emitters. As the electric potential increased, the capillary-type emitter exhibited a whipping and pulsating water jet, while glycerin displayed a steady tilted jet. In contrast, the annular slit-type emitter demonstrated a pulsating water jet followed by a distinctive dripping mode at higher electric potentials, while glycerin formed multiple steady jets around the annular slit. Notably, the annular slit-type emitter, when subjected to an 18.5 kV potential, produced seven electrospray jets for glycerin, a phenomenon attributed to the novel design of the emitter and the viscosity of glycerin enabling the generation of multiple cone-jets at a specific electrostatic potential around the slit peripheral meniscus. Vacuum chamber tests of the annular-type emitter using liquid indium as an ion source at 1 × 10−5 Torr revealed an ion-current density of 0.3 mA/mm, resulting in a thrust of 290 μN.

A simple retarding-potential time-of-flight mass spectrometer for electrospray propulsion diagnostics

The time-of-flight mass spectrometer (ToF-MS) is a useful tool for quantifying the performance of electrospray thrusters and characterizing their plumes. ToF-MS data can be used to calculate the mass-to-charge distribution in the plume, but the kinetic-energy-to-charge (i.e., the potential) distribution must be known first. Here we use a ToF-MS in tandem with a retarding potential (RP) analyzer. By sweeping the retarding potential through the range of potentials present in the plume, both the mass-to-charge distribution and the potential distribution can be measured independently. We demonstrate this technique in a case study using a capillary electrospray emitter and the ionic liquid propellant 1-Ethyl-3-methylimidazolium bis(trifluoromethylsulfonyl)imide, abbreviated EMI-Im. We report a linear correlation between retarding potential and mass-to-charge ratio that agrees with published data from more complex orthogonal RP/ToF-MS instruments. Calculated values for the jet velocity and jet breakup potential match within 2% and 12%, respectively. Using conventional ToF-MS, we estimated the propellant flow rate and compared those estimates to direct flow rate measurements. For flow rates between 233 pL/s and 565 pL/s, the error in ToF-based flow rate estimates ranged from -16% to -13% when the plume potential was assumed to be a function of mass-to-charge. Assuming a constant plume potential yielded mixed results. However, using the average stopping potential measured by a retarding potential analyzer resulted in higher errors, ranging from -26% to -30%. Data and MATLAB code are included as supplemental materials so that readers can easily apply the techniques described here.

On the cone-to-jet transition region and its significance in electrospray propulsion

A novel method for identifying the cone-to-jet transition region is proposed to characterize the electrospray dynamics, and thus the Taylor cone, a fundamental principle of ion extraction from the tip of the Taylor cone in an electric field to produce μN thrust for nanosatellites' application. A numerical model is developed and validated against the droplet's diameter with a proper characterization of the cone-to-jet transition region based on the hydrodynamic pressure gradient at the cone-jet axis. The results show that the cone-to-jet transition region is reduced when the electric potential increases and the flow rate decreases. Moreover, the current density is found to be higher for the lower flow rate and large electric potential. Interestingly, the current density at the jet axis increases significantly at the flow rate of 1 cc/hr. This implies that below a certain droplet diameter, the current density carried by the jet interface is completely migrated into the jet, and the substantial change in the current density occurs within the cone-to-jet transition region. When comparing the current density along the jet axis, the small flow rate and large electric potential carry the maximum current density at the jet interface. The identification of cone-to-jet transition region is central in modeling the ion beam trajectory via initial current density profile required for particle-in-cell calculation for assuring the performance of a field emission electric propulsion thruster.

Formation and fragmentation of 2-hydroxyethylhydrazinium nitrate (HEHN) cluster ions: a combined electrospray ionization mass spectrometry, molecular dynamics and reaction potential surface study.

The 2-hydroxyethylhydrazinium nitrate ([HOCH2CH2NH2NH2]+NO3-, HEHN) ionic liquid has the potential to power both electric and chemical thrusters and provide a wider range of specific impulse needs. To characterize its capabilities as an electrospray propellant, we report the formation of HEHN cluster ions in positive electrospray ionization (ESI) and their collision-induced dissociation. The experiment was carried out using ESI guided-ion beam mass spectrometry which mimics an electrospray thruster in terms of ion emission, injection into a vacuum and fragmentation in space. Measurements include compositions of primary ions in the electrospray plume and their individual dissociation product ion cross sections and threshold energies. The results were interpreted in light of theoretical modeling. To determine cluster structures that are comprised of [HE + H]+ and NO3- constituents, classical mechanics simulations were used to create initial guesses; and for clusters that are formed by reactions between ionic constituents, quasi-classical direct dynamics trajectory simulations were used to mimic covalent bond formation and structures. All candidate structures were subject to density functional theory optimization, from which global minimum structures were identified and used for construction of reaction potential energy surface. The comparison between experimental values and calculated dissociation thermodynamics was used to verify the structures for the emitted species [(HEHN)nHE + H]+, [(HEHN)n(HE)2 + H]+, [(HE)n+1 + H]+ and [(HE)nC2H4OH]+ (n = 0-2), of which [(HE)1-2 + H]+ dominates. Due to the protic nature of HEHN, cluster fragmentation can be rationalized by proton transfer-mediated elimination of HNO3, HE and HE·HNO3, and the latter two become dominant in larger clusters. [(HE)2 + H]+ and [(HE)nC2H4OH]+ contain H-bonded water and consequently are featured by water elimination in fragmentation. These findings help to evaluate ion formation and fragmentation efficiencies and their impacts on electrospray propulsion.

Fragmentation modeling of gas-phase ionic liquid clusters in high-voltage electric field

Cluster fragmentation subjected to a high-voltage electric field may play an important role in controlling the novel ionic liquid (IL) electrospray propulsion but has received little attention. In this work, the micro-fragmentation mechanisms of three kinds of IL clusters are numerically investigated using both static and dynamic approaches based on the quantum mechanical principles. Firstly, it reveals that the Coulomb interaction dominates the total inter-fragment energy for all three IL dimers, whereas the hydrogen-bonding interaction is identified as the main cause for the much slower break-up rate of the 1-ethyl-3-methylimidazolium tetrafluoroborate ([EMIM-BF4]) dimer. In addition to the electrostatic interaction, the van der Waals interaction is a secondary energy component to which only a small fraction of atoms contribute significantly, regardless of IL cluster type. We find, only when the applied electric field is larger than 1 V/nm, the total inter-fragment energies of all three IL clusters start to dramatically reduced, indicating the possible occurrence of massive break-up. Secondly, ab initio molecular dynamics simulation further confirms that the [EMIM-BF4] negatively charged dimer is the most stable of three IL clusters under the identical conditions. As far as the high-voltage electric field is concerned, the one-parameter exponential function of break-up ratio versus time, routinely used in the field-free region, is not valid. Finally, on the basis of simulated results, a compound exponential-decay model is proposed, and the derived differential-sum equations from constant electric fields can be further applied to the non-uniform electric field.

Experimental characterization of the electrospray propulsive performance for ionic liquid propellants [EMIm][DCA] and [BMIm][DCA]

A chemical-electrospray dual-mode propulsion system is significant in improving mission capability and flexibility of a spacecraft. Hypergolic ionic liquids are potential propellants of chemical and electrospray propulsion systems. In this work, the electrical propulsive performance of two hypergolic ionic liquids 1-ethyl-3-methylimidazolium dicyanamide ([EMIm][DCA]) and 1-butyl-3-methylimidazolium dicyanamide ([BMIm][DCA]) are experimentally characterized using two array-emitter electrospray thrusters. Compared with a commonly used ionic liquid propellant 1-ethyl-3-methylimidazolium tetrafluoroborate ([EMIm][BF4]), the current magnitude and current steadiness of the electrospray thruster with [EMIm][DCA] and [BMIm][DCA] are efficiently improved, especially with [EMIm][DCA] as propellant. The results show that the low viscosity is beneficial for steady and sufficient flow supply and thus favors a long-time steady operation of the thruster. For the propellant [EMIm][DCA], the maximum specific impulse is 1791 s, i.e., 27.8% higher than that of [EMIm][BF4], while the maximum thrust 32.5 μN is 128.9% greater than that of [EMIm][BF4]. For the propellant [BMIm][DCA], the thrust-to-power ratio of the thruster is approximately 82.8% higher than that of [EMIm][BF4] at a similar specific impulse of about 1120 s resulting from a much greater thruster efficiency. These results demonstrate that [EMIm][DCA] has a better performance than [EMIm][BF4] in both maximum thrust and specific impulse, and [BMIm][DCA] performs best in the aspect of thrust-to-power ratios. This work highlights the prospect of hypergolic ionic liquids and is valuable to promote their application in chemical and electrical propulsion communities.

Scalable Microfabrication of Multi-Emitter Arrays in Silicon for a Compact Microfluidic Electrospray Propulsion System.

The recent proliferation of SmallSats and their use in increasingly demanding applications require the development of onboard electric propulsion compatible with the power, mass, and volume constraints of these spacecraft. Electrospray propulsion is a promising technology for SmallSats due to its unique high efficiency and scalability across the wide power range of these platforms, for example, from a few watts available in a CubeSat to a few hundred watts in a MiniSat. The implementation of electrospray propulsion requires the use of microfabrication techniques to create compact arrays of thousands of electrospray emitters. This article demonstrates the microfabrication of multi-emitter electrospray sources of a scalable size for electrospray propulsion. In particular, a microfabrication and assembly process is developed and demonstrated by fabricating sources with arrays of 1, 64, and 256 emitters. The electrospray sources are tested in a relevant environment for space propulsion (inside a vacuum chamber), exhibiting excellent propulsive performance (e.g., absence of beam impingement in the extractor electrode, absence of hysteresis in the beam current versus propellant flow rate characteristic, proper operation in the cone-jet electrospraying mode, etc.) and nearly coincident output per emitter. Several design elements contribute to this performance: the even distribution of the propellant among all emitters made possible by the implementation of a network of microfluidic channels in the backside of the emitter array; the small dead volume of the network of microfluidic channels; the accurate alignment between the emitters and extractor orifices; and the use of a pipe-flow configuration to drive the propellant through closed conduits, which protects the propellant.

Electrospray propulsion: Modeling of the beams of droplets and ions of highly conducting propellants

This article presents a model for the beams of charged droplets and ions typical of electrospray thrusters. The model is applied to a typical highly conducting propellant operating in the mixed droplet-ion emission regime but can be extended to simulate other emission regimes. A key feature of the model is the separation of the beam in two zones: a small inner region including the droplet emission area, where the equations of motion of all droplets are integrated simultaneously while retaining direct droplet-on-droplet Coulomb interactions and a much larger outer region where the beam is treated as a continuum. In addition to the geometry of the electrodes, inputs to the model include the distribution of diameters and charge-to-mass ratios of the droplets and their initial positions and velocities. Most of this information is determined from experiments. The analysis of the numerical solution confirms that lateral oscillations of the slender jet producing the droplets is an important factor in the expansion of electrospray beams, and that droplets in flight are the main sources of ion emission in these propellants. In addition to improving the knowledge of the physics of electrospray beams, the model is a significant tool for designing the extractor and accelerator electrodes of electrospray thrusters and for minimizing the deposition of beam particles on the surfaces of spacecrafts.

Multi-scale modeling of ionic electrospray emission

The physics of ionic electrospray propulsion spans multiple length scales. This paper combines a molecular dynamics model, a particle–particle model, and a particle-in-cell model to investigate the physics of ionic electrospray propulsion over 9 orders of magnitude in length scale. The combined models are applied to simulate beam emission for an ionic electrospray propulsion system with porous emitter tips and 1-ethyl-3-methylimidazolium tetrafluoroborate ionic liquid propellant from the emission site to the downstream plume. Additionally, the impact of multiple emission sites from a single emitter tip is analyzed with regard to extractor grid interception and overall beam neutralization for bipolar thruster pairs. Results show that beams consisting of species of different masses (i.e., monomer and dimer species) are affected by particle–particle forces during acceleration and should not be treated as a superposition of independently accelerated species in macro-scale plume models. The activation of multiple emission sites also causes a noticeable increase in the beam’s spread, leading to increased intercepted current but relatively little adverse effects in the downstream plume region.

On the Cone-to-Jet Transition Region and its Significance in Electrospray Propulsion

On the Cone-to-Jet Transition Region and its Significance in Electrospray Propulsion

Performance measurement and evaluation of an ionic liquid electrospray thruster

As a novel micro-propulsion system for small satellites (from micro to nano), the ionic liquid electrospray propulsion system is a promising candidate. However, performance measurement and evaluation of the Ionic Liquid Electrospray Thruster (ILET) is one of the most challenging issues for practical application, due to the difficulties in the development of a prototype and direct measurements of micro-thrust and small flow rate. To address this issue, a Modular Ionic Liquid Electrospray Thruster (MILET) prototype is constructed, and a diagnostic system for thrust and mass flow rate is specially developed based on an analytical balance method. With the ionic liquid 1-ethyl-3-methylimidazolium tetrafluoroborate as the propellant, a series of experiments is carried out on the MILET prototype under a wide operating condition through changing the applied voltage to control the thrust. Under different applied voltages, the thrust and the mass flow rate of the propellant are directly measured. The propulsive performance parameters of the thruster, such as thrust, specific impulse, thrust-to-power ratio, thruster efficiency, etc., are comprehensively analyzed. Then, a performance comparison is made between the MILET and other representative ILETs. With a relatively low applied voltage ranging from 1550 V to 2000 V, the MILET achieves a quasi-constant specific impulse of 1263 s with the averaged thrust-to-power ratio of 65.2 μN/W and thruster efficiency of 40.7%. The performance of ILET is also compared with other typical electric propulsions. The results demonstrate that the ILET exhibits an excellent ability of minimalization with high specific impulse and thruster efficiency, which guarantees a great superiority in micro propulsions. Finally, the ways to further improve the performance of ILET are discussed, which further confirms the potential prospect of ILET. The present result helps to advance the development and application of ILET.

Creation of electrospray emitters from a flat polytetrafluoroethylene sheet through the use of laser micro-manufacturing

Electrospray propulsion is a form of electrostatic propulsion that shows potential for use as a thruster on small satellites. Electrospray emitters traditionally use arrays of capillaries, however, this paper investigates the suitability of manufacturing an emitter by laser drilling holes in sheets of polytetrafluoroethylene (PTFE), also known as Teflon. This is done by examining factors such as the machining time, hole shape and degree of taper in holes drilled through different methods. Experimental work was done to determine the shape of holes drilled in a flat sheet, and a computational model was used to simulate the performance of these drilled holes. These results were used to judge if this would be a valid method to manufacture electrospray thrusters, and how these thrusters would compare to more traditional emitters. Based on these results, it was concluded that a PTFE sheet with a width of 0.81 mm would be sufficient to achieve electrospray at voltages lower than those needed for traditional emitters.

Electrospray Propulsion Engineering Toolkit (ESPET)

We report on the development of a software tool, the Electrospray Propulsion Engineering Toolkit (ESPET), that is currently being shared as a web application with the purpose to accelerate the development of electrospray thruster arrays for space propulsion. ESPET can be regarded as a database of microfluidic properties and electrohydrodynamic scaling models that are combined into a performance estimation tool. The multiscale model integrates experimental high-level physics characterization of microfluidic components in a full-scale electrospray propulsion (ESP) microfluidic network performance solution. ESPET takes an engineering model approach that breaks the ESP system down into multiple microfluidic components or domains that can be described by either analytical microfluidic or reduced order numerical solutions. ESPET can be divided into three parts: a central database of critical microfluidic properties, a microfluidic domain modeler, and a microfluidic network solver. Two options exist for the network solution, a detailed multi-domain solver and a QuickSolver designed for rapid design and testing of simple three-domain reservoir-feed-emitter arrays. The multi-domain network solver exploits the Hagen–Poiseuille/Ohm’s law analogy by using the publicly available SPICE (Simulation Program with Integrated Circuit Emphasis) electric circuit simulation software to solve the flow properties of the microfluidic network. Both the multi-domain and QuickSolver solutions offer Monte Carlo analysis of arrays based on user supplied tolerances on design parameters. Benchmarking demonstration examples are provided for experimental work in the literature, as well as recent experimental work conducted at Busek Co. The demonstration examples include ionic liquid propelled systems using active and passive capillary emitters, externally wetted emitter needles, and porous glass emitters, as well as a liquid metal system based on an externally wetted emitter needle.

Orbital Maintenance of a Constellation of CubeSats for Internal Gravity Wave Tomography

The advent of radio occultation (RO) instruments aboard CubeSats leads to the possibility of a mission to sound atmospheric internal gravity waves if such satellites are deployed in close-flying constellation. The satellites in the constellation must have slightly perturbed orbital inclinations in order to spread the RO soundings within clusters in two horizontal dimensions, and consequently the satellites will disperse because they will experience different rates of regression of nodes. This dispersion must be countered by propulsive maneuvering in order to maintain the close formation of the constellation. Here, a theoretical approach to the necessary propulsive maneuvering is presented and simulations using comprehensive orbit propagators are performed to analyze four propulsive systems: two cold gas propulsion systems and two electrospray propulsion systems. Cold gas propulsion permits greater separations in inclination between satellites in a constellation by virtue of the greater thrust they can exert on a spacecraft: cold gas propulsion can permit inclination separations of 1 to 10 $^\circ$ while electrospray limits separations to less than 0.2 $^\circ$ . On the other hand, electrospray propulsion provides much longer mission lifetime by virtue of the greater total thrust it offers: cold gas propulsion expends all of its fuel in maintaining the constellation formation in less than approximately 100 days while electrospray propulsion can maintain formation for greater than 1000 days before expending all of its fuel. Mission lifetime is the most critical consideration for a mission, thus electrospray propulsion is recommended for the constellation-flying of CubeSats, but the accelerations that they offer must be greatly increased to enable spacecraft separations useful for tomography of internal gravity waves. Note that any close-flying constellation involving satellites with slightly perturbed inclinations will experience the same dispersing effect as the constellations described herein and, thus, require the same propulsive maneuvering to maintain formation.

Development of a microthrust balance and ion beam measurement system: Characterizing a dual-mode thruster for spacecraft

There is a rapid progression toward spacecraft miniaturization due to the prospect of reduced launch costs, the requirement for satellite constellations, and the integration of microelectromechanical systems (MEMS). Satellite propulsion is required for increased mission flexibility and orbital debris mitigation, but down-scaling of such devices has proven non-trivial. Some electrostatic thrusters are both challenging and promising given their increase in performance with small geometry, but the relatively low thrust produced through ion acceleration limits their mission applicability. For this reason, it may be desirable to produce thrusters that are able operate across a wide range of specific impulses, depending on maneuver time, thrust bit requirements, and several other factors that influence propulsion system selection. At RAPPEL, a bimodal thruster that can operate at high specific impulse with electrospray propulsion or high thrust using cold gas expansion has been developed. Inevitably, such a propulsion system comes with complications related to performance characterization. A novel and economical data acquisition system has been established to measure the performance of both modes in high vacuum. For the cold gas mode, a torsional thrust balance was built in-house to measure thrust levels between 1−15mN, and can be easily modified for lower ranges. The natural frequency of the thrust stand was measured to be 5.6Hz. The vacuum system sustained pressures below 5mbar while the cold gas nozzle operated, and 1.3×10−5mbar during electrospray operation. A 3-D printed linear cold gas nozzle was used to validate the thrust stand. Its thrust was close to 9mN at a mass flow rate of 20mg/s of nitrogen, producing a specific impulse of just under 45s. Due to the noise present in thrust stand measurements at the sub-mN force-level produced by electrosprays, the thrust stand was built to also accommodate a complementary ion beam current measurement system. The power processing unit, used to generate square-wave ±2500V signals using optodiodes, was built to operate a bipolar ionic liquid ion source (ILIS). The rise and fall times of the push/pull switch were close to 1ms and 30ms, respectively. The electrical power system used to operate the electrospray and measure its performance is examined in detail, followed by the results from validation of both data acquisition systems.

CubeSat constellation management using Ionic Liquid Electrospray Propulsion

The Space Propulsion Laboratory (SPL) of the Massachusetts Institute of Technology (MIT) is developing the Ion Electrospray Propulsion System (iEPS), designed to address a current need in CubeSat technology: miniaturized electric thrusters. These could be used for different applications, ranging from attitude control to interplanetary flights. In this work, performed together with the Space Systems Laboratory of the Costa Rica Institute of Technology (SETEC Lab), we explore a case study in which the iEPS is used for constellation management in Low Earth Orbit (LEO) when integrated in a 3U CubeSat. We analyze how a 180° separation in the Right Ascension of the Ascending Node (RAAN) between two CubeSats (SatA and SatB) starting in the same orbit can be achieved by modifying one of the spacecraft's orbital altitude, resulting in a difference in their rate of nodal precession (defined as the drift rate) due to the J2 effect, and therefore a difference in their relative RAAN. The method consists of SatB increasing its semi-major axis, drifting in a higher orbit with a lower drift rate, and returning to the original semi-major axis once the desired difference in RAAN in achieved relative to the other spacecraft. SatA will stay in its original orbit, using its thruster to compensate for orbital energy loss due to atmospheric drag, therefore demonstrating another application of iEPS for constellation management. Three different simulations were studied, defined as the minimum time trajectory, minimum propellant trajectory and a hybrid trajectory, consisting of reaching a higher altitude orbit, but actively changing the RAAN using the propulsion system instead of drifting. It was observed that the difference in this orbital element could be achieved using 85 g of propellant in as little as 164 days for the minimum time trajectory. The same difference could also be achieved using only 44 g of propellant in 245 days for the minimum propellant trajectory. Furthermore, the results of the hybrid trajectory showed that the goal could be achieved in 161 days, but using 158 g of propellant mass, demonstrating the benefit of using a drift orbit. The results proved the feasibility of implementing iEPS for constellation management using 3U CubeSats in LEO.

Interplanetary CubeSats for asteroid exploration: Mission analysis and design

Recent advances in CubeSats technology are leading the transition from purely education tools to actual scientific missions. The small volumes and masses, the versatile purpose, as well as the fast development time associated with a potential high return-to-cost ratio are at the origin of the increasing number of new mission proposals, also beyond low Earth orbit (LEO). The purpose of this study is to assess the CubeSats ability to complement an interplanetary scientific mission. The proposed AIDA (Asteroid Impact and Deflection Assessment) mission, an ESA/NASA joint effort to demonstrate the kinetic impact technique to change the motion of an asteroid in space, has been selected as case study, having a mission context particularly suitable in showing CubeSats supporting capabilities. The feasibility study of a mission involving the use of CubeSats as secondaries for technology demonstration and science purposes was performed. Mission objectives and requirements were defined, followed by the development of concepts of operations and mission architectures proposals. Eventually, multiple trade-off tools were adopted to define the proposed mission baseline, which involves the deployment of two 3U CubeSats performing a detachment to achieve the final configuration of one 2U and four 1U CubeSats. The scientific campaign conducted by the CubeSats includes gravitational and magnetic field mapping, on-surface chemical-physical measurements via multiple wide chip-size-sensor nets deployed from orbit, on-surface seismic measurements via landing of the 2U CubeSat, direct observation of the impact from multiple viewpoints and evaluation of the asteroid's orbit deflection due to the impact. The mission proposed involves also some important technological demonstrators. The S-iEPS (Scalable-ion Electrospray Propulsion System) has been considered as propulsion system, while a potential landing system has been proposed to achieve the soft touchdown of the 2U CubeSat. Finally, an inter-satellite communication link via laser has been included as main communication system. The proposed mission baseline has shown that CubeSats can be successfully integrated as multi-platform systems to provide useful support to interplanetary missions. This solution may enable the capability to acquire more detailed information with the possibility to combine them to obtain better results with respect to single-platform systems. The proposed mission concept represents a valuable low-cost piggyback solution adaptable to several mission contexts, with a potential high return and a remarkable attitude for the implementation and testing of new technologies and operations. In this context, this study provides a useful framework for the design and development of interplanetary CubeSats missions.

Molecular Dynamics Electrospray Simulations of Coarse-Grained Ethylammonium Nitrate (EAN) and 1-Ethyl-3-Methylimidazolium Tetrafluoroborate (EMIM-BF4)

In this work, the use of molecular dynamics as a predictive tool for modeling the atomistic behavior of electrospray propulsion is discussed. 1-ethyl-3-methylimidazolium tetrafluoroborate (EMIM-BF 4 ) and ethylammonium nitrate (EAN) were considered as two limits of ionic liquid (IL) propellants that tend to operate in an ion versus a droplet mode. The emission modes were found to depend on the electro-chemical properties of the IL propellant. The aprotic EMIM-BF 4 -based electrospray emitted primarily monomers and trimers as the dominant species and only small quantities of droplets. In contrast, trimers were the dominant emitted species in the protic EAN emissions with a significantly large contribution from droplets towards the total emission current, suggesting that EMIM-BF 4 -based colloid thrusters operate in ion mode and EAN-based devices operate in the droplet mode. Furthermore, the formation of the Taylor cone was found to depend on the mass flow rate and the external electric field strength. This paper provides a framework that can be extended for use to simulate any other ILs or their combinations.