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Electrospray thrusters require significant improvements in operational lifetime for use inmulti-year spacecraft propulsion missions. The primary thruster lifetime-limiting mechanismis propellant overspray, in which wide-angle particles impinge on and saturate downstreamelectrodes instead of exiting through the electrode aperture and contributing to producedthrust. Electrospray particles are emitted within a small radial range, but diverge as theymove downstream from emission to form a 3D plume, the edges of which contribute tooverspray. In order to improve electrospray thruster designs towards minimizing oversprayand optimizing operational lifetime, we need to understand what causes electrospray plumedivergence.This dissertation investigates electrospray plume divergence using the Discrete ElectrosprayLagrangian Interaction (DELI) Model to simulate electrospray particle dynamics. Thegoverning equation for particle propagation includes the applied electrostatic force from thepotential difference between the emitter and downstream electrode, the Coulomb forcesbetween particles (including image charges), and the drag force. Each of these forces is investigatedtheoretically and computationally to determine its influence on plume divergence.None of the forces introduce radial divergence into a set of particles emitted straight down the axis of emission with no range in radial coordinate. However, electrospray particles are always emitted with some small range in radial coordinate due to hydrodynamic instabilitiesand minute asymmetries in the emitter. All three forces exacerbate existing radial divergenceamong a set of particles: the applied electric field has a radial component due to jetcurvature and the electrode aperture; there is a radial component to Coulomb forces betweenparticles with a difference in radial coordinate; and drag counters particle motion, keepingparticles in a clustered state in which Coulomb forces are magnified.Simulations compare the radial divergence of groups of particles with equal velocities andwith an upstream velocity gradient, in which upstream particles are moving faster than theirforward neighbors. In the upstream velocity gradient case, faster particles catch up to theirforward neighbors, magnifying the Coulomb interaction between the two in response to theirincreased proximity. We term this interaction a ‘traffic jam’ and correlate it with increasedplume divergence through Coulomb interactions. We present two novel means of characterizingplume divergence: 1) a metric for positional divergence based on three standards of aGaussian or Super-Gaussian fit to particle mass density distribution as a function of radialcoordinate, and 2) emittance as a metric for positional and velocity divergence. We furtherdescribe how emittance can be used to identify when an electrospray plume has reached thesteady state.Machine learning is applied for the first time to electrospray particle dynamics data,produced by the DELI Model. Results demonstrate predictive abilities for downstreamparticle dynamic properties given particle properties at emission. Furthermore, a novelmethod is proposed for combining experimental electrospray particle data, computationalplume evolution models, and machine learning algorithms to optimize diagnostic design.In summary, this dissertation presents a comprehensive consideration of electrosprayplume divergence using computational and analytical models supported by experimentaldata. The origins and sources of growth of electrospray plume divergence are identified, new metrics for electrospray plume divergence are presented, and machine learning algorithmsare developed to predict electrospray plume divergence.In summary, this dissertation presents a comprehensive consideration of electrosprayplume divergence using computational and analytical models supported by experimentaldata. The origins and sources of growth of electrospray plume divergence are identified, new metrics for electrospray plume divergence are presented, and machine learning algorithmsare developed to predict electrospray plume divergence.

The static relation between business and engineering design hinders the pace of innovation. While program managers often evaluate innovation in terms of financial value generated over a number of business scenarios, engineering design teams base their activities on improving product functionality and meeting technical requirements. This results in an insufficient common understanding during gate meetings about the business implications of alternative technological tradeoffs, thus negatively impacting the pace of innovation. This article presents the results from the introduction of a methodology–based on value and functional modeling–into the practice of design teams working with next-generation electric propulsion systems for satellite applications. The introduction of the methodology was evaluated via interviews, workshops, and observations with nine industrial partners. The results indicate business stakeholders and technology-focused design teams’ bidirectional interest in the methodology. In particular, the results highlight the benefits of the methodology in creating cross-boundary representations that can be used by stakeholders to share knowledge and find common ground in gate meetings. The dynamic interaction with such representations enables a faster decision-making pace during the management of innovation initiatives.

This study demonstrated the performance of flight control systems of a hybrid rocket in a hovering flight test by developing a rocket designated HTTP-3AT powered by High Test Peroxide (HTP, a term used for concentrated hydrogen peroxide, H 2 O2). Hybrid rocket excels in system simplicity, operational safety, oxidizer storability, cost, and throttling capability compared to current solid and liquid rocket engine systems. Although issues such as the severe oxidizer-to-fuel (O/F) ratio shift during combustion and difficulty in gimbaled thrust vector control (TVC) caused by the lengthy chambers need to be solved, hybrid rocket propulsion is nevertheless a promising propulsion technology for future space exploration. To achieve accurate orbit insertion, thrust magnitude control and TVC of the rocket engines are necessary. However, no organizations have successfully implemented this technology on a practical hybrid rocket, not even using this technology for hovering flight tests. On September 8th, 2020, a hovering flight test of HTTP-3AT was conducted, achieving a steady hover 3 m above ground for 25 s utilizing both attitude and position controls. This test showed that a hybrid rocket could achieve a stable hovering flight with the capability of vertical takeoff and vertical landing (VTVL), demonstrating excellent throttling control and TVC capabilities of hybrid rocket propulsion.

Hybrid rocket propulsion can contribute to reduce launch costs by simplifying engine design and operation. Hypergolic propellants, i.e. igniting spontaneously and immediately upon contact between fuel and oxidizer, further simplify system integration by removing the need for an ignition system. Such hybrid engines could also replace currently popular hypergolic propulsion approaches based on extremely toxic and carcinogenic hydrazines. Here we present the first demonstration for the use of hypergolic metal-organic frameworks (HMOFs) as additives to trigger hypergolic ignition in conventional paraffin-based hybrid engine fuels. HMOFS are a recently introduced class of stable and safe hypergolic materials, used here as a platform to bring readily tunable ignition and combustion properties to hydrocarbon fuels. We present an experimental investigation of the ignition delay (ID, the time from first contact with an oxidizer to ignition) of blends of HMOFs with paraffin, using White Fuming Nitric Acid (WFNA) as the oxidizer. The majority of measured IDs are under 10 ms, significantly below the upper limit of 50 ms required for functional hypergolic propellant, and within the ultrafast ignition range. A theoretical analysis of the performance of HMOFs-containing fuels in a hybrid launcher engine scenario also reveals the effect of the HMOF mass fraction on the specific impulse (Isp) and density impulse (ρIsp). The use of HMOFs to produce paraffin-based hypergolic fuels results in a slight decrease of the Isp and ρIsp compared to that of pure paraffin, similar to the effect observed with Ammonia Borane (AB), a popular hypergolic additive. HMOFs however have a much higher thermal stability, allowing for convenient mixing with hot liquid paraffin, making the manufacturing processes simpler and safer compared to other hypergolic additives such as AB.

This article, the last in a series of three articles, attempts to unravel the underlying physics of recent experiments regarding the contradictory properties of the neutron lifetime that has been a complete riddle for quite some time. So far, none of the advanced theories beyond the Standard Models (SMs) of particle physics and cosmology have shown sufficient potential to resolve this mystery. We also try to explain the blatant contradiction between the predictions of particle physics and experiments concerning the nature and properties of the (so far undetected) dark matter and dark energy particles. To this end the novel concepts of both negative and hypercomplex matter (giving rise to the concept of matter flavor) are introduced, replacing the field of real numbers by hypercomplex numbers. This extension of the number system in physics leads to both novel internal symmetries requiring new elementary particles – as outlined in Part I and II, and to novel types of matter. Hypercomplex numbers are employed in place of the widely accepted (but never observed) concept of extra space dimensions – and, hence, also to question the corresponding concept of supersymmetry. To corroborate this claim, we report on the latest experimental searches for novel and supersymmetric elementary particles by direct searches at the Large Hadron Collider (LHC) and other colliders as well as numerous other dedicated experiments that all have come up empty handed. The same holds true for the dark matter search at European Council for Nuclear Research (CERN) [CERN Courier Team, “Funky physics at KIT,” in CERN Courier, 2020, p. 11]. In addition, new experiments looking for dark or hidden photons (e.g., FUNK at Karlsruhe Institute of Technology, CAST at CERN, and ALPS at Desy, Hamburg) are discussed that all produced negative results for the existence of the hitherto unseen but nevertheless gravitationally noticeably dark matter. In view of this contradicting outcome, we suggest a four-dimensional Minkowski spacetime, assumed to be a quasi de Sitter space, dS 1,3, complemented by a dual spacetime, denoted by DdS 1,3, in which the dark matter particles that are supposed to be of negative mass reside. This space is endowed with an imaginary time coordinate, −it and an imaginary speed of light, ic. This means that time is considered a complex quantity, but energy m(ic)2 > 0. With this construction visible and dark matter both represent positive energies, and hence gravitation makes no distinction between these two types of matter. As dark matter is supposed to reside in dual space DdS 1,3, it is principally undetectable in our spacetime. That this is evident has been confirmed by numerous astrophysical observations. As the concept of matter flavor may possibly resolve the contradictory experimental results concerning the lifetime of the neutron [J. T. Wilson, “Space based measurement of the neutron lifetime using data from the neutron spectrometer on NASA’s messenger mission,” Phys. Rev. Res., vol. 2, p. 023216, 2020] this fact could be considered as a first experimental hint for the actual existence of hypercomplex matter. In canonical gravity the conversion of electromagnetic into gravity-like fields (as surmised by Faraday and Einstein) should be possible, but not in cosmological gravity (hence these attempts did not succeed), and thus these conversion fields are outside general relativity. In addition, the concept of hypercomplex mass in conjunction with magnetic monopoles emerging from spin ice materials is discussed that may provide the enabling technology for long sought propellantless space propulsion.

A computational tool able to perform a fast analysis of hybrid rocket engines is presented, describing briefly the mathematical and physical models used. Validation of the code is also shown: 16 different static firing tests available in the open literature are used to compare measured operational parameters such as chamber pressure, thrust, and specific impulse with the code’s output. The purpose of the program is to perform rapid evaluation and assessment on a possible first design of hybrid rockets, without relying on computationally expensive simulations or onerous experimental tests. The validated program considers as benchmark and study case the design of a liquid-oxygen/paraffin hybrid rocket engine to be used as the upper stage of a small launcher derived from VEGA building blocks. A full-factorial parametric analysis is performed for both pressure-fed and pump-fed systems to find a configuration that delivers the equivalent total impulse of a VEGA-like launcher third and fourth stage as a first evaluation. This parametric analysis is also useful to highlight how the oxidizer injection system, the fuel grain design, and the nozzle features affect the performance of the rocket.

The high density of iodine, and its storage in solid form with no pressurization requirements, have driven the increase on the use of the halogen for space propulsion systems. This paper presents the design and development process of a miniaturized propellant management system which can supply iodine at a stable mass flow rate to the thrusters. This system approach has been adapted to produce the flow control units of the NPT30-I2 gridded ion propulsion system and the I2T5 cold gas thruster, both of which have performed their first in-orbit demonstration flights. The paper is dedicated to the description of the design process, with a focus on the particularities of the use of iodine for space propulsion systems, mainly related to chemical interaction and deposition. The numerical analysis of rarefied flows serves as a base for the design of the flow path elements of the system. A performance characterization of the system obtained on ground and the challenges tackled during the development process are evaluated with the in-orbit data from the first flight missions of both propulsion systems.

The performance (specific impulse, Isp) of a modular, 150-lbf-class rotating detonation rocket engine (RDRE) was measured with three gaseous fuels (methane, ethane, and ethylene) and gaseous oxygen...

Since modern propulsion systems are insufficient for large-scale space exploration, a breakthrough in propulsion physics is required. Amongst different concepts, the EMDrive is a proposed device claiming to be more efficient in converting energy into propulsive forces than classical photon momentum exchange. It is based on a microwave resonator inside a tapered cavity. Recently, Taylor suggested using a laser instead of microwaves to boost thrust by many orders of magnitude due to the higher quality factor of optical resonators. His analysis was based on the theory of quantised inertia by McCulloch, who predicted that an asymmetry in mass surrounding the device and/or geometry is responsible for EMDrive-like forces. We put this concept to the test in a number of different configurations using various asymmetrical laser resonators, reflective cavities of different materials and size as well as fiber-optic loops, which were symmetrically and asymmetrically shaped. A dedicated high precision thrust balance was developed to test all these concepts with a sensitivity better than pure photon thrust, which is the force equivalent to the radiation pressure of a laser for the same power that is used to operate each individual devices. In summary, all devices showed no net thrust within our resolution at the Nanonewton range, meaning that any anomalous thrust must be below state-of-the-art propellantless propulsion. This puts strong limits on all proposed theories like quantised inertia by at least 4 orders of magnitude for the laboratory-scale geometries and power levels used with worst case assumptions for the theoretical predictions.

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

State-of-the-art electrodynamic tethers reach a steady electric current by using a bare segment to capture electrons passively from the ambient plasma (anodic contact) and an active electron emitter or a tether segment coated with a low-work-function material (cathodic contact) to emit electrons back and close the electrical circuit. This work proposes to take advantage of recent developments on thin-film solar cells and insert a photovoltaic (pv) tether segment in between the anodic and the cathodic contacts. Since thin-film solar cells can be folded and manufactured with any desired length and the same cross-section dimensions as the bare segment, i.e. width and thickness around few centimeters and tens of microns, the resulting device is compact and preserves bare tether simplicity. Detailed analysis of the current and voltage profiles throughout the tether shows that the electrical power introduced by the pv-segment into the tether-plasma circuit improves the performance and makes them less dependent on ambient conditions. The pv-segment decreases considerably the tether-to-plasma bias at the cathodic contact, thus opening the possibility to emit substantial current while using consumable-less electron emitters like thermionic and electron field emitters. The pv-segment also favors the current collection by increasing the tether-to-plasma bias at the bare segment. Propulsion and power generation applications and alternative architectures of bare-pv tethers are briefly discussed. This work received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 828902 (E.T.PACK project). GSA work is supported by the Ministerio de Ciencia, Innovación y Universidades of Spain under the Grant RYC-2014-15357.

The focus of this review is to collect and compare research studies on hybrid rocket fuels and components produced via additive manufacturing. The use of hybrid rocket propulsion for launching applications is limited due to its inherently low regression rate and actual specific impulse. To address these limitations, various additive manufacturing methods have been employed so far to fabricate fuel grains with complex port geometries or composite fuel grains possessing intricate shaped, printed scaffolds into which a second fuel material is incorporated. In addition to the fuel grains, additive manufacturing technology has also proven to be beneficial for fabricating a considerable number of hybrid rocket components, thereby reducing the part count, production time and cost. In this review, the rationale, types of the materials used, methods and comparison of performance data are presented. Future directions that use additive manufacturing to enhance hybrid rocket propulsion are also provided.

This paper considers the combination of an analytical reference trajectory with linear state feedback control to allow for autonomous guidance and control of a spacecraft for coplanar circle-to-circle transfers with a stage-based propulsion system. Staging of electric propulsion components, such as tanks and thrusters, could allow small spacecraft to achieve high- Δ V capabilities with current propulsion technology. In order to utilize these propulsion systems, further developments in guidance and control of such spacecraft are required due to limitations in computational power and communications. Analytical approximations for low-thrust trajectories could allow for computationally simple guidance and control of autonomous spacecraft for circle-to-circle transfers around large central bodies. Many trajectories have been developed for conventional propulsion systems, based either on their shape or input thrust, and applied for preliminary mission design. A previously developed analytical trajectory is extended to account for the effects of staging propulsion system components. In order to stabilize the trajectory in the presence of disturbances, a linear state feedback control law is designed with linear quadratic regulator methods. Finally, a methodology for determining the correct phasing between the spacecraft and a target object is developed and is practical to implement on power-limited computers. The use of the analytical reference trajectory is simulated on an orbit transfer from low-Earth orbit to geostationary orbit.

A promising aspect of inductive plasma generators for space propulsion applications is their electrodeless nature that leads to a high propellant compatibility and opens the door to in-situ resource utilisation propellants. To further develop inductive plasma generators as a space propulsion technology it is important to understand the relationship between system inputs (e.g. input power and propellant material) and system outputs (e.g. power efficiencies) to aid design and control. In this work, methodologies are developed for non-intrusively assessing the inductive coupling from antenna current measurements in terms of the ‘instantaneous ignition power’, the ‘inductive duty cycle’ and the ‘inductive frequency shift’. Experimental results are presented for IPG7, a high-power (up to 180 kW) 5.5-turn helical antenna inductive plasma generator. Potential in-situ resource utilisation materials oxygen and carbon-dioxide, and their respective mixtures with argon, are assessed in terms of their inductive coupling characteristics and resulting power efficiencies. Oxygen is shown to be an attractive propellant material, especially when supplemented with argon. In particular, high thermal efficiency ( ∼ 84 % ) can be achieved at high input powers. A basis is provided for understanding the power efficiencies of an inductive plasma generator in terms of non-intrusive antenna current measurements, to aid power supply sizing and to develop monitoring and control techniques for thruster applications.

3D printing is an advanced manufacturing technology, enabling production of complex shapes by adding material layer-upon-layer. A fuel injection system component designed with Inconel-718 nickel-based superalloy for an advanced rocket propulsion system was realized through Laser Powder Bed Fusion method. 3D printing allowed realizing the complex component having interconnected channels and having minute openings, which are impossible to manufacture by conventional manufacturing methods such as CNC machining. Detailed metallurgical characterization and mechanical property evaluation were performed to qualify the component for space application. The process helped in avoiding multiple piece assemblies for the fuel injection system and developing an integrated component within a short turnaround time. Microstructure evaluation revealed absence of any defects such as porosity, lack of fusion and microcracks. The mechanical properties were evaluated on 3D-printed test coupons in four different orientations along with the component. The fractographs in STA condition in four different orientations revealed a typical ductile fracture with the presence of dimples. Computed tomography of component was carried out and it was concluded that defects > 139 μm were not present, which is the limitation for the thickness of Inconel-718 component scanned. The formation of fine, complex channels and absence of defects > 139 μm as observed in CT, makes the component suitable for complex fuel injection operation for high-temperature space application.

Direct thrust measurements by means of a thrust balance (TB) are the golden standard for measuring thrust and, concurrently, the specific impulse in electric thrusters. To measure these properties in the novel class of electrodeless plasma thrusters, a new TB based on the Variable Amplitude Hanging Pendulum with Extended Range (VAHPER) concept has been developed. The TB has a mechanical amplification mechanism with an angular magnification of $31^\circ /^\circ $ . Using Lagrangian mechanics, we show that the TB loaded with a 5.2-kg thruster prototype behaves like a damped harmonic oscillator with a natural frequency of 0.37 Hz. A variable damping system provides damping with an optimal damping ratio of 0.78, which corresponds to a settling time of only 1.8 s. Both the model and the damping and calibration system have been validated. To accommodate the particularities of medium power electrodeless plasma thrusters, the TB design includes the following features: an optical displacement sensor, water cooling, liquid metal connectors, and dedicated vacuum-rated electronics for autoleveling, remote (in-vacuum) calibration, and temperature monitoring. To test the TB, measurements were performed on a 500-W Helicon Plasma Thruster breadboard model. When loaded with this thruster, the measured stiffness of the system was 12.67 ± 0.01 mN/mm. For this stiffness, the thrust range is 150 mN with a 0.1-mN resolution. The relative uncertainty on the thrust measurements is found to be on the order of 2%.

Paraffin-based fuels incorporating solid amine boranes are investigated to identify formulations suitable for use as hypergolic hybrid rocket propellants. Their ignition delays are measured followi...

The paper presents the results of a research project carried out at the Military University of Technology aimed at designing a technology demonstrator of an active protection system – a smart counter-projectile for combating anti-tank missiles at a fixed distance from the protected object. Since the design of the counter-projectile head includes electronic components sensitive to high loads, a solid propellant rocket motor was used as the propulsion system. Based on the specification and requirements for the propulsion system, the propellant charge and nozzle dimensions were determined, and the performance properties of the designed system (chamber pressure, thrust with time and total thrust pulse), calculated. The tests and analyses were carried out using the known properties of homogenous solid rocket propellants manufactured in Poland. To verify the results of the theoretical analysis, experimental studies were carried out in collaboration with “GAMRAT” Sp. z o.o. Special Production Plant (Jasło, Poland) to validate the selected solid propellant and the initial assumptions made on the operation of the propulsion system of the designed counter-projectile.

Electrodynamic tethers are proposed as propulsion and energy harvesters for space probes orbiting planets with a magnetic field and ionosphere, however there are no descriptions in the technical literature of the design of an electrical power system for such an application at subsystem and circuit detail. This paper presents a proposal for such a power system that extracts energy from the kinetic energy of a spacecraft using a bare electrodynamic tether, i.e. an unsheathed conductive wire or band, in low-Earth orbit. The application of the system is the powering of the spacecraft while in its final de-orbiting maneuvers at end-of-life with no reliance on the main spacecraft bus.

A pulsed electric thruster named Pulsed Electrogasdynamic (PEG) thruster was investigated. In this thruster, the propellant gas is accelerated gasdynamically to high exhaust velocities (10–15 km/s...

Although several investigations have been performed regarding the fundamental physics of axial-injection, end-burning hybrid rocket motors, investigations into the mission performance of a propulsi...

With the development of electric space propulsion and all-electric propulsion technologies, the application range of the Hall effect thruster (HET) has gradually expanded. As an important part of the HET, the magnetic field plays a decisive role in its performance. In the development of versatile HETs, it is necessary to comprehensively and thoroughly understand the effects of the magnetic field on the HET discharge characteristics. In this paper, a HEP-100X thruster with multiple degrees of freedom in terms of magnetic field control is used to study the effects of the peak magnetic field position on the thruster discharge characteristics. The results indicate that the peak magnetic field position can effectively control the main ionization zone position and affect the ionization rate. The thruster has the best comprehensive discharge performance when the peak magnetic field position is located near the channel exit. Further analysis reveals that the performance improvement is mainly due to the reduced ion loss on the wall surface and plume divergence half-angle, and the increased current utilization efficiency and voltage utilization efficiency. The HEP-100X shows the above trends under various flow conditions, and the anode efficiency is the highest when the anode flow is 50 sccm. In this condition, the propellant is sufficiently ionized and the ion loss on the wall surface is small. This research provides the necessary support for a comprehensive understanding of the role and significance of the magnetic field in the thruster discharge process, and has important significance for the design optimization of multifunctional Hall thrusters and subsequent research.

The paper considers non-spherical objects with low gravitational attraction, such as asteroids, satellites of the planet and comets. We considered possibility of a mission to small bodies of the solar system of irregular shape on the example of the asteroid Apophis. The authors of the article suggest using a nanoclass spacecraft with an electric rocket propulsion system for a long mission to study Apophis. The purpose of this work is to determine the necessary costs of the working body for all stages of the mission, which includes reaching the asteroid, forming and maintaining a given orbit relative to it. The gravity of the Earth, Sun, and asteroid is taken into account when modeling the controlled movement of the spacecraft. When a spacecraft is moving relative to an asteroid, its gravitational field is described as a superposition of the gravitational fields of two rotating massive points. In this paper, it is proposed to divide the mission into two sections for preliminary ballistic design. The first optimal speed heliocentric flight Earth-asteroid Apophis with the alignment of the speed of the spacecraft and the asteroid. The second is the movement in the vicinity of the asteroid, which includes the optimal speed maneuver for forming the working orbit and maintaining the working orbit for a given time.

A method has been developed for the combined de-orbiting of large-size objects of space debris from low-Earth orbits using an electro-rocket propulsion system as an active de-orbiting means. A principal de-orbiting technique has been devised, which takes into consideration the patterns of using an electric rocket propulsion system in comparison with the sustainer rocket propulsion system. A procedure for determining the parameters of the de-orbiting scheme has been worked out, such as the minimum total speed and the time of the start of the de-orbiting process, which ensures its achievement. The proposed procedure takes into consideration the impact exerted on the process of the de-orbiting by the ballistic factor of the object, the height of the initial orbit, and the phase of solar activity at the time of the de-orbiting onset. The actual time constraints on battery discharge have been accounted for, as well as on battery charge duration, and active operation of the control system. The process of de-orbiting a large-size object of space debris has been simulated by using the combined method involving an electro-rocket propulsion system. The impact of the initial orbital altitude, ballistic coefficient, and the phase of solar activity on the energy costs of the de-orbiting process have been investigated. The dependences have been determined of the optimal values of a solar activity phase, in terms of energy costs, at the moment of the de-orbiting onset, and the total velocity, required to ensure the de-orbiting, on the altitude of the initial orbit and ballistic factor. These dependences are of practical interest in the tasks of designing the means of the combined de-orbiting involving an electric rocket propulsion system. The dependences of particular derivatives from the increment of a velocity pulse to the gain in the ballistic factor on the altitude of the initial orbit have been established. The use of these derivatives is also of practical interest to assess the effect of unfolding an aerodynamic sailing unit

A cathode-less radio frequency (RF) thruster is a plasma-based system for space propulsion. It consists on a source for plasma generation driven by an RF antenna and a magnetic nozzle for the acceleration of the exhausted plasma stream. A parametric analysis has been conducted in order to evaluate the influence of some key design parameters on the performances of such a thruster. Specifically, the intensity and the topology of the magneto-static field, along with the antenna geometry have been varied. Their influence has been assessed on the plasma parameter profiles within the source (i.e., electron density, electron temperature, and power deposition), the antenna properties (i.e., impedance and current distribution), and the propulsive performances (i.e., thrust and specific impulse). The results presented in this work have been obtained with the validated code 3D-VIRTUS. The latter solves self-consistently both the electro-magnetic wave propagation and the plasma transport within a source driven by a RF antenna. An analytical model has been subsequently employed in order to solve the plasma dynamics in the plume and, in turn, to estimate the propulsive performances.

The use of multimodal space propulsion has the benefit of increasing satellite mission flexibility with the ultimate goal of decreasing mission development costs. The present paper explores the dev...

Staging of space propulsion systems would allow lifetime limitations inherent to small propulsion systems to be bypassed in order to enable high-ΔV capabilities for small spacecraft, in particular mass and volume constrained CubeSats. In addition, staging can be used to provide redundancy in the propulsion system, counteract thruster degradation, or open up new avenues of mission optimization. Analytical approximations are developed in order to provide a computationally simple approach to the design, analysis, definition of propulsion technology requirements, and online autonomous decision making for systems that make use of staging of propulsion elements. In addition, the analytical approximations provide insight into the dependencies of performance metrics on system parameters. Equations are developed for any mission, defined by its ΔV, and then specialized for escape trajectories., NASA (Grants 80NSSC18M0045, 80NSSC18K1186)