Drag Compensation Research Articles

Cathode-less RF plasma thruster design and optimisation for an atmosphere-breathing electric propulsion (ABEP) system

Atmosphere-breathing electric propulsion (ABEP) is a concept that ingests residual atmospheric gases as a source of propellant for an electric thruster, removing the need for onboard propellant storage. This would enable continuous low-thrust drag compensation, extending the lifetime of spacecraft in Very-Low Earth Orbit (VLEO); <250 km. VLEO is an appealing region for spacecraft operations, enabling new remote sensing missions with improved radiometric performance and spatial resolution, whilst reducing size, mass and power requirements, as well as mission cost. A preliminary design review and optimisation is therefore conducted for an ABEP system that uses the cathode-less radio frequency (RF) plasma thruster from Technology for Innovation & Propulsion (T4i) S.p.A. This removes the issue of thruster erosion by means of magnetic confinement and offers reduced susceptibility to varying atmospheric composition. A semi-empirical oxygen-nitrogen global source model (GSM) has been developed which considers the volume-averaged flux, momentum, and energy balance of the RF discharge. This includes a detailed chemistry model for the complex electron-molecular reactions and energy-loss channels of air plasma in the ionisation chamber. The GSM is coupled to an analytical model of flux balance for an air intake, verified by Direct Simulation Monte-Carlo (DSMC) simulation, to consider its design for maximum collection efficiency. This is then utilised in a robust multi-objective optimisation of the ABEP system, accounting also for spacecraft aerodynamics and power requirements.

Kinetic Simulation of Nozzle Flow in a Micronewton-Class Cold Gas Thruster

Many scientific space missions need highly precise attitude and orbit control or ultrafine drag compensation, which relies on the variable-thrust propulsion technology operating at the micronewton level. Cold gas thruster (CGT) is a very promising solution, mainly because of its high reliability. One of the keys to the success of micronewton variable-thrust CGT is to understand the flow in its nozzle, whose configuration is much more complex than traditional CGT nozzles. This paper applies kinetic-based multiscale models to investigate the gas flow in the complex nozzle of a micronewton variable-thrust CGT. The simulations reveal that the flow simultaneously experiences all kinds of regimes from continuum to free molecular. The continuum breakdown is likely to occur near the throat region due to large gradients of flow variables and in the expander due to low gas density. Frictional choking is observed, and the nozzle length can be optimized to improve the thruster performance. Nozzle performance measures such as thrust, discharge coefficient, and thrust efficiency are found to change only with the throat Knudsen number Knt. The performance curves can be divided into two sections at Knt≃0.1, and thereby an empirical piecewise formula for thrust prediction is proposed.

On the critical parameters for feasibility and advantage of air-breathing electric propulsion systems

In recent years, air-breathing electric propulsion emerged as a potential enabling technology for long-duration space missions in Very Low Earth Orbit (VLEO). In this work, we show how the complex relation between mission environment, spacecraft configuration, and propulsive performance can be associated to a single requirement for full air-breathing drag compensation, which becomes less stringent for higher VLEO orbits. The impact of the spacecraft shape and size on the performed analysis is then evaluated and the results compared with conventional electric propulsion solutions with stored propellant, showing how the adoption of air-breathing propulsion becomes more advantageous, from a volume and mass fraction perspective, when the spacecraft scale is reduced.

Deflagration thruster for air-breathing electric propulsion in very low Earth orbit

This study establishes the performance requirements for an air-breathing electric propulsion device to operate in very low Earth orbit (<450 km). We demonstrate that existing electric propulsion architectures, including electrostatic thrusters, fail to generate enough thrust while operating on flow rates of air that can be harvested in orbit. Instead, we develop and characterize a type of electromagnetic pulsed thruster that operates in a magneto-deflagration mode (i.e., forms an expansion wave during the acceleration process) that becomes more efficient as less propellant is used. We show this mode of operation can generate specific impulses up to 104 s, thrust per power >2 mN/kW, thrust densities >100 mN/m2, and thrust efficiencies up to 10% while consuming <100μg of air per discharge. We map these performance metrics on a candidate spacecraft to show the deflagration thruster can enable fully air-breathing drag compensation at altitudes ranging from ∼200 km to 350 km depending on its geometry and is extendable to other altitudes if stored propellant is utilized.

Optimal use of electric propulsion for drag compensation in very low earth orbit on satellites with deployable solar panels

In recent years, thanks to advancements in commercial hardware technology miniaturization, small satellites have become more and more important and effective in the Space ecosystem. In previous papers it has been demonstrated that future smallest observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost, lower risk, shorter schedules, lower up-front development cost, more sustainable business model, mitigating the problem of orbital debris. Flying in very low Earth orbit requires the addition of a propulsion system capable of providing drag compensation. Thanks to its high specific impulse, electric propulsion is potentially well suited for this task. However, the use of an electric thruster often requires significant power that can be provided only with deployable solar panels. To minimize the drag, the solar panels should be aligned with the satellite orbital velocity. At the same time, to maximize the power input, the solar panel surfaces should point toward the Sun. In the majority of cases these two conditions are not met simultaneously, so an optimal trade-off is needed. This topic is investigated in this paper by the use of an orbital model of a generic satellite with deployable solar panels. The model is able to simulate any circular orbit of the satellite in any day of the year and calculate both the power input and the drag from the solar panels based on their inclination. An equivalent specific impulse is also defined taking into account the thrust that the electric thruster has to provide only to push the area of the solar panels required to power the propulsion system. The results show, as expected, that the sun-synchronous Dawn-Dusk (or Dusk-Dawn) orbit is the preferred one for the electric propulsion as the velocity vector and the Sun vector are almost orthogonal. In the other cases, at very low Earth orbit, even relatively small inclinations of the solar panels respect to the velocity vector have a significant impact on the drag produced, increasing the propellant consumption and the required power. Nevertheless, it is shown also that if the solar panels are aligned with the velocity, the power lost can be kept reasonable. The same analysis could also apply to a satellite in low Earth orbit that does not employ an electric thruster but still needs deployable solar panels because of other sources of high power consumption (e.g. the payload).

Parametric study on the flight envelope of a radio-frequency ion thruster based atmosphere-breathing electric propulsion system

The atmosphere-breathing electric propulsion (ABEP) system utilizes atmospheric species as propellants to generate thrust for drag compensation of a satellite in very-low-Earth-orbit (VLEO). A parametric approach is used to assess the impact of different parameters on the flight envelope of a sample VLEO satellite with an ABEP system based on a radio-frequency ion thruster (RIT). The considered parameters include capture efficiency, maximum input power, solar activity, and atomic oxygen recombination factor. The NRLMSIS 2.0 atmosphere model is used to determine the flow conditions at VLEO, at a target altitude of 150–300 km with high, moderate, and low solar activity levels. DSMC method is employed to calculate the drag of the sample satellite with the ABEP system. The 0-D model of the RIT discharge chamber is used to predict the thrust of the RIT-based ABEP system. The flight envelope of a sample RIT-based ABEP system shows that drag can be compensated at altitudes between 196 km and 248 km. Increasing capture efficiency and maximum input power expands the feasible range for drag compensation to higher and lower altitudes, respectively. Also, the flight envelope shifts to higher altitudes with increasing solar activity levels. However, the atomic oxygen recombination factor of the intake device has minimal effects on the flight envelope.

Feasibility of autonomous orbit keeping with optimal fuel consumption using drag-free CubeSat

In this paper, the composition of the drag-free CubeSat platform was introduced first. Then the control strategy of drag-free CubeSat orbit keeping was proposed. Finally, the simulation with the control scheme of the Homan transfer orbit was compared. By comparing and analyzing the fuel consumption of the drag-free track and Homan transfer track, it is demonstrated that the continuous drag compensation of the drag-free control scheme can significantly reduce the consumption of thruster mass, and the drag-free technology has good advantages in track stability.

Comparison between Different Re-Entry Technologies for Debris Mitigation in LEO

The population of satellites in Low Earth Orbit is predicted to growth exponentially in the next decade due to the proliferation of small-sat constellations. Consequently, the probability of collision is expected to increase dramatically, possibly leading to a potential Kessler syndrome situation. It is therefore necessary to strengthen all the technologies required for collision avoidance and end-of-life disposal of new satellites, together with active debris removal of current and potential future dead satellites. Both situations require the lowering of the altitude of a satellite up to re-entry. In this paper several de-orbiting technologies are evaluated: natural decay, chemical propulsion (solid and liquid), electric propulsion, drag sail, electrodynamic tether, and combinations of the previous ones. The comparison considers the initial altitude, system mass, de-orbiting time, collision probability during descent, reliability, and technological limits. Differences between active debris removal and satellite end-of-life self-disposal are taken into account. Moreover, the different types of re-entry, controlled vs. non-controlled, expendable vs. reusable system, demisable vs. non-demisable system are also discussed. Finally, the possibility to operate the satellite in Very Low Earth Orbits with a propulsion system for drag compensation and passive re-entry at end of life is investigated.

Design and optimisation of a passive Atmosphere-Breathing Electric Propulsion (ABEP) intake

In order to extend the orbital lifetime of spacecraft operating in Very-Low Earth Orbit (VLEO), Atmosphere-Breathing Electric Propulsion (ABEP) can be employed for atmospheric drag compensation. The concept is based on the ingestion of rarefied atmospheric particles to be used as propellant for an electric thruster, thereby removing the need for onboard propellant. The present paper aims to design and analyse a passive ABEP intake optimised for the RF Helicon-based Plasma Thruster (IPT), which is selected as the most promising system due to its electrodeless design, as it removes the issue of thruster corrosion, and reduced susceptibility to atmospheric variations. This is achieved by implementing the analytical model, referred to as the Balancing Model (BM), along with its main performance parameters. The Direct Simulation Monte Carlo (DSMC) solver dsmcFoam+ is thus employed to simulate the transmission probability of cylinders and hexagonal prisms with non-zero entrance velocity, yielding a mean percentage error of 2% when compared to independent DSMC results. In addition, a novel DSMC transmittance investigation of square prisms is performed. Various design configurations are optimised for circular and hexagonal intakes based on the functional requirements of the thruster, leading to a maximum collection efficiency of 45% for a cylindrical intake with a hexagonal honeycomb. The optimal intake is hence simulated via DSMC, showing a good accuracy with the BM with or without intermolecular collisions, yielding a percentage error of 0.22% in the former case and 5.53% in the latter one. The variation of incidence flow angle, accommodation coefficient and thruster transmission probability is finally investigated, and the results are validated by the agreement shown with values retrieved from the literature.

0-D composition and performance analysis of an air-breathing radiofrequency ion thruster

AbstractThis paper presents a physics-based 0-D steady-state model of a radiofrequency ion thruster (RIT) that is valid for any propellant blend, and hence, any plasma mixture, and the model is used to predict the drag compensation capability and plasma composition of an air-breathing RIT operating in very-low Earth orbit (VLEO). This study fills a gap in the modeling of air-breathing electric propulsion with an air-breathing RIT model that is validated with experimental data of a RIT operating with an $${\mathrm{N}}_{2}$$ N 2 /$${\mathrm{O}}_{2}$$ O 2 mixture. The model expands upon current 0-D RIT models by accounting for the dissociation of neutral molecules and molecular ions and by applying particle conservation for each ion and neutral species. Atomic oxygen in the atmosphere is assumed to completely recombine into $${\mathrm{O}}_{2}$$ O 2 in the collector, and the model is used to calculate an air-breathing RIT’s thrust, discharge efficiency, mass utilization efficiency, and discharge plasma composition at altitudes between 80 and 150 km. At 1 kW of total input power and an optimal flow rate of 0.17 mg/s, the net thrust was limited to 3 mN at all altitudes considered. At these same operating conditions, the mass utilization efficiency and discharge efficiency were approximately 0.15 and 870 eV/ion, respectively. To increase the performance of the thruster, the magnetic field should be increased in order to increase the beam current. The mole fraction of atomic oxygen in the discharge chamber was found to be between 0.19 and 0.27 depending on the altitude, indicating that grid erosion should be a focus of future studies.

Position Controller for a Flapping-Wing Drone Using UWB

This paper proposes an integral approach for accurate ultra-wideband indoor position control of flapping-wing micro-air vehicles. Three aspects are considered to achieve a reliable and accurate position controller. The first aspect is a velocity/attitude flapping-wing model for drag compensation. The model is compared with real flight data and shown to be applicable for more than one type of flapping-wing drone. The second improvement regards a voltage-dependent thrust control. Lastly, a characterisation of ground effects in flapping-wing flight is obtained from hovering experiments. The proposed controller improves position control by a factor [Formula: see text], reaching a mean absolute error of 10[Formula: see text]cm for the position in [Formula: see text] and [Formula: see text], and 4.9[Formula: see text]cm for the position in [Formula: see text].

Air-breathing electric propulsion: Flight envelope identification and development of control for long-term orbital stability

Air-breathing electric propulsion (ABEP) enables long duration missions at very low orbital altitudes through the use of drag compensation. A system-level spacecraft model is developed, using the interaction between thruster, intake and solar arrays, and coupled to a calculation of the drag. A quadratic solution is found for specific impulse and evaluated to identify the thruster performance required for drag-compensation at varying altitudes. An upper altitude limit around 190 km is based on a minimum thruster propellant density, resulting in required thruster performance values of Isp>3000s and T/P>8mN/kW for a realistic ABEP spacecraft. The orbit of an air-breathing spacecraft is propagated with time, which highlights the prescribed orbit eccentricity due to non-spherical gravity and therefore an increased variability in the atmospheric conditions. A thruster control law is introduced which avoids a divergent altitude behaviour by preventing thruster firings around the orbit periapsis, as well as adding robustness against atmospheric changes due to season and solar activity. Through the use of an initial frozen orbit, thruster control and an augmented T/P, a stable long-term profile is demonstrated based on the performance data of a gridded-ion thruster tested with atmospheric propellants. An initial mean semi-major axis altitude of 200 km relative to the equatorial Earth radius, a spacecraft mass of 200 kg, Isp=5455s and T/P=23mN/kW, results in an altitude range of around 10 km at altitudes of 160–183km during a period of medium to high solar activity.

Dual mode operation of a hydromagnetic plasma thruster to achieve tunable thrust and specific impulse

We report here on initial studies of a pulsed hydromagnetic plasma gun that can operate in either a pre-filled or a gas-puff mode on demand. These modes enable agile and responsive performance through tunable thrust and specific impulse. Operation with a molecular nitrogen propellant is demonstrated to show that the hydromagnetic thruster is a candidate technology for air-harvesting and drag compensation in the very low Earth orbit. A dual mode operation is achieved by leveraging propellant gasdynamics to change the fill fraction and flow collisionality within the thruster. This results in the formation of distinct modes that are characterized by the current-driven hydromagnetic waves that they allow, namely, magneto-deflagration and magneto-detonation, respectively. These modes form the basis of using gasdynamics to enable responsive thruster performance. Using time-of-flight emission diagnostics to characterize near-field flow velocities, we find that a relatively dramatic transition occurs between modes as gas is allowed to expand in the thruster, with exhaust velocities ranging from 10 to 55 km/s in the deflagration and detonation regimes, respectively. Simulations of the processed mass bit offer the first glimpse into possible thruster performance and trade-offs between specific impulse and thrust. An impulse bit tunability of ∼22% is predicted, with differing propellant fill fractions when operating in a burst mode.

Exponentially Stable Motion Control for Multirotor UAVs with Rotor Drag and Disturbance Compensation

Exponentially Stable Motion Control for Multirotor UAVs with Rotor Drag and Disturbance Compensation

Extended channel Hall thruster for air-breathing electric propulsion

Air-breathing electric propulsion is a promising concept that may allow sustained access to very low earth orbit by guaranteeing continuous drag compensation through in situ harvesting of propellant. For earth observation, such low orbits benefit image resolution and reduce signal latency and transmission power requirements, easing link budgets. Moreover, not having to carry typical electrostatic thruster propellants such as xenon and tanks positively impact overall mission costs. However, the low molecular mass and high first ionization energies of principal atmospheric constituents hinder efficient thruster operation and result in strongly degraded performance. In this study, we examine the performance of an extended channel Hall thruster design for air-breathing operations. An extended channel compensates for the reduced residence times of lighter neutrals. The longer channel, coupled with an extended radial magnetic field region, allows for an increased length of the ionization zone, enhancing molecular ionization probability. Here, we present direct thrust measurements of this thruster operating with a pure nitrogen flow. Analysis of the results show performance in the 500–800 W anode power range, with thrust, specific impulse, and total thrust efficiency ranging from 17 to 22 mN, 1000 to 1100 s, and 14% to 18%, respectively, with a constant 2 mg/s pure molecular nitrogen propellant flow. Performance tends to degrade with increased voltage, suggesting increased contributions of electron-molecule kinetics to performance losses, contrary to what is typically seen in the Hall thrusters operating with propellants such as xenon.

System modelling of very low Earth orbit satellites for Earth observation

The operation of satellites in very low Earth orbit (VLEO) has been linked to a variety of benefits to both the spacecraft platform and mission design. Critically, for Earth observation (EO) missions a reduction in altitude can enable smaller and less powerful payloads to achieve the same performance as larger instruments or sensors at higher altitude, with significant benefits to the spacecraft design. As a result, renewed interest in the exploitation of these orbits has spurred the development of new technologies that have the potential to enable sustainable operations in this lower altitude range. In this paper, system models are developed for (i) novel materials that improve aerodynamic performance enabling reduced drag or increased lift production and resistance to atomic oxygen erosion and (ii) atmosphere-breathing electric propulsion (ABEP) for sustained drag compensation or mitigation in VLEO. Attitude and orbit control methods that can take advantage of the aerodynamic forces and torques in VLEO are also discussed. These system models are integrated into a framework for concept-level satellite design and this approach is used to explore the system-level trade-offs for future EO spacecraft enabled by these new technologies. A case-study presented for an optical very-high resolution spacecraft demonstrates the significant potential of reducing orbital altitude using these technologies and indicates possible savings of up to 75% in system mass and over 50% in development and manufacturing costs in comparison to current state-of-the-art missions. For a synthetic aperture radar (SAR) satellite, the reduction in mass and cost with altitude were shown to be smaller, though it was noted that currently available cost models do not capture recent commercial advancements in this segment. These results account for the additional propulsive and power requirements needed to sustain operations in VLEO and indicate that future EO missions could benefit significantly by operating in this altitude range. Furthermore, it is shown that only modest advancements in technologies already under development may begin to enable exploitation of this lower altitude range. In addition to the upstream benefits of reduced capital expense and a faster return on investment, lower costs and increased access to high quality observational data may also be passed to the downstream EO industry, with impact across a wide range of commercial, societal, and environmental application areas.

Low-Earth-Orbit Constellation Phasing Using Miniaturized Low-Thrust Propulsion Systems

Many low-Earth-orbit constellations consist of multiple small satellites that are launched together in batches, and which make use of differential atmospheric drag for subsequent deployment and phasing. Depending on the initial altitude, this process can potentially take many months, or even years, to complete, and satellite altitudes can only ever be decreased, potentially affecting mission lifetimes. Miniaturized low-thrust propulsion systems are one option to achieve faster phasing and can additionally increase the constellation lifetime through drag compensation. Here, an analytical and numerical study comparing differential drag and low-thrust propulsion for constellation phasing is performed. Satellite electrical constraints associated with power generation and battery cycling life represent important factors that can decrease the performance of an onboard propulsion system. Despite these constraints, low-thrust propulsion can reduce total phasing times by an order of magnitude, or more, for altitudes in the range 300–700 km, while also increasing satellite flexibility and introducing a number of operational advantages. Remaining propellant can be used for drag compensation and collision avoidance maneuvers, while also reducing deorbiting times by many years.

Impulse Bit Measurements from Metal Plasma Thruster

A torsional thrust stand, calibrated for impulse bits in the range of , was used to measure impulse bits from a metal plasma thruster. Impulse data were obtained on roughly 1400 shots, with metal plasma thruster targets of molybdenum, niobium, palladium, aluminum, and carbon. Model predictions (based on a simple circuit model and published plasma parameters) were validated by data from the calibrated torsional thrust stand. Over a typical 50-shot firing sequence, the impulse bits measured showed a coefficient of variation of approximately 5%. This implies that individual impulse bits of about can be imparted to a satellite with variation about the mean in a total impulse burst of . For a 20 kg satellite, this would result in a velocity increment of just , enabling very precise attitude control and fine positioning control of nano- and microsatellites. The metal plasma thruster uses solid metal propellant and hence requires no liquids, gases, flow valves, or flow controls and has no moving parts. Total impulse approximately provides orbit raising and drag compensation capability.

SITAEL HC1 Low-Current Hollow Cathode

SITAEL is active in the field of electric propulsion and is involved in the development of different thruster technologies—mainly Hall thrusters (HTs)—of power levels ranging from 100 W up to 20 kW. Low-power HTs are the most effective choice to perform orbit transfer, drag compensation, and de-orbiting maneuvers for small satellites. This paper is dedicated to the activities regarding HC1, the hollow cathode conceived for the 100-W-class Hall thruster under development at SITAEL. Successful test campaigns were performed and are described, with emphasis on the improvements in the cathode design after an extensive research and development phase. The results are presented and discussed, along with future developments of the ongoing activities.

Impact of propulsion system characteristics on the potential for cost reduction of earth observation missions at very low altitudes

Earth observation is one of the most important satellites’ applications. Past earth observation systems have used traditional space technology to achieve the best possible performance, but have been very expensive. Recently, thanks to advancements in technology and modern microelectronics, small satellites have become more and more useful at much lower costs, even if with reduced performance. The resolution of the optical payload improves as the altitude is reduced. Space system mass is proportional to the cube of the linear dimensions. This means that by flying at lower altitudes, satellites can reduce their payload size and therefore the entire mass of the satellite, thus reducing the cost of the system dramatically. However, almost all the earth observation missions fly at the minimum altitude that provides a sufficient orbital life. The addition of a propulsion system capable of providing drag compensation for the entire satellite operative life provides the possibility to fly at very low earth orbit. In this way, the same performance can be obtained with a smaller and cheaper system. To obtain the same coverage more units are needed to replace a larger unit at higher altitude. In this paper it is confirmed that future smallsat observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost (by a significant factor), lower risk (both implementation and operations), shorter schedules, lower up-front development cost, more sustainable business model, to be more flexible and resilient, more responsive to both new technologies and changing needs, and to mitigate the problem of orbital debris. This paper focus in particular on the effect of the propulsion system parameters (performance and costs) on the cost model as a function of the altitude. It is demonstrated that new affordable chemical propulsion systems provide already significant benefits with limited constraints, allowing a useful reduction of altitude and, consequently, costs. Electric propulsion systems have the potential to allow even lower altitudes or longer lifetimes; however, they have a stronger impact on the satellite design related to their power consumption, generally requiring deployable solar panels, which can limit the flexibility in the orbit selection or the added weight and cost of batteries. The development of electric thrusters that have good performance and limited impact on the satellite architecture (particularly at small scales) is fundamental to exploit their potential for reduced mission costs through very low altitude flight.