Primary Propulsion System Research Articles

Trajectory & maneuver design of the NEA Scout solar sail mission

The Near-Earth Asteroid Scout (NEA Scout) mission aimed to deliver a 6U CubeSat to a slow flyby of a Near Earth Asteroid using a solar sail as the primary propulsion system. NEA Scout was launched as a secondary payload aboard Artemis I, on November 16, 2022, but communication with the spacecraft was not achieved. This paper describes the target asteroid selection and the innovative trajectory and maneuver design for the NEA Scout mission, including cold gas maneuvers, multiple lunar flybys, solar sail thrusting, and design margins for solar sailing. The resulting trajectory design tools, solutions, and analyses are potentially applicable to future solar sail missions.

Three-Dimensional Guidance Laws for Spacecraft Propelled by a SWIFT Propulsion System

This paper discusses the optimal control law, in a three-dimensional (3D) heliocentric orbit transfer, of a spacecraft whose primary propulsion system is a Solar Wind Ion Focusing Thruster (SWIFT). A SWIFT is an interesting concept of a propellantless thruster, proposed ten years ago by Gemmer and Mazzoleni, which deflects, collects, and accelerates the charged particles of solar wind to generate thrust in the interplanetary space. To this end, the SWIFT uses a large conical structure made of thin metallic wires, which is positively charged with the aid of an electron gun. In this sense, a SWIFT can be considered as a sort of evolution of the Janhunen’s E-Sail, which also uses a (nominally flat) mesh of electrically charged tethers to deflect the solar wind stream. In the recent literature, the optimal performance of a SWIFT-based vehicle has been studied by assuming a coplanar orbit transfer and a two-dimensional scenario. The mathematical model proposed in this paper extends that result by discussing the optimal guidance laws in the general context of a 3D heliocentric transfer. In this regard, a number of different forms of the spacecraft state vectors are considered. The validity of the obtained optimal control law is tested in a simplified Earth–Venus and Earth–Mars transfer by comparing the simulation results with the literature data in terms of minimum flight time.

Material Research for Hypersonic Travel

The goal of travelling faster than five times the speed ofsound, or hypersonic travel, has enormous potential to transform global transportation, defence capabilities, and space access. However, materials and structures face severe obstacles because to harsh aerothermal conditions inherent in high Mach number trajectories. This work focuses on new approaches to do research for ceramics, composites, and refractory alloys that are essential for building hypersonic vehicles. Essential design concepts for primary structures, heat shielding, and propulsion systems are covered, highlighting the critical role that theory and computation play in comprehending the links between structure, property, and processing. The remarkable high-temperature capabilities, stiffness, strength, and corrosion resistance of ceramic materials are highlighted in the study as reasons for theirincreasing importance in aircraft applications. Based on their distinct features, titanium alloys, nickel aluminides, metal- matrix composites, carbon-carbon, and ceramic-matrix composites stand out as top choices for a high temperature. In pursuit of lightweight, high- performance materials for hypersonic travel, this paper summarises ongoing researchefforts, evaluates the state of the art, offers insights into technological hurdles, and identifies areas that require future improvement. To explore the complex field of hypersonicmaterials design and selection in this research, this paper hope to shed light on the critical role that materials science plays in pushing the boundaries of aerospace technology by investigating the special difficulties presented by hypersonic flight as well as the characteristics and potential of various material classes.

Materials design for hypersonics.

Hypersonic vehicles must withstand extreme conditions during flights that exceed five times the speed of sound. These systems have the potential to facilitate rapid access to space, bolster defense capabilities, and create a new paradigm for transcontinental earth-to-earth travel. However, extreme aerothermal environments create significant challenges for vehicle materials and structures. This work addresses the critical need to develop resilient refractory alloys, composites, and ceramics. We will highlight key design principles for critical vehicle areas such as primary structures, thermal protection, and propulsion systems; the role of theory and computation; and strategies for advancing laboratory-scale materials tomanufacturable flight-ready components.

Optimal Trajectories of Diffractive Sail to Highly Inclined Heliocentric Orbits

Recent literature indicates that the diffractive sail concept is an interesting alternative to the more conventional reflective solar sail, which converts solar radiation pressure into a (deep space) thrust using a thin, lightweight highly reflective membrane, usually metalized. In particular, a diffractive sail, which uses a metamaterial-based membrane to diffract incoming solar rays, is able to generate a steerable thrust vector even when the sail nominal plane is perpendicular to the Sun–spacecraft line. This paper analyzes the optimal transfer performance of a diffractive-sail-based spacecraft in a challenging heliocentric scenario that is consistent with the proposed Solar Polar Imager mission concept. In this case, the spacecraft must reach a near-circular (heliocentric) orbit with a high orbital inclination with respect to the Ecliptic in order to observe and monitor the Sun’s polar regions. Such a specific heliocentric scenario, because of the high velocity change it requires, is a mission application particularly suited for a propellantless propulsion system such as the classical solar sail. However, as shown in this work, the same transfer can be accomplished using a diffractive sail as the primary propulsion system. The main contribution of this paper is the analysis of the spacecraft transfer trajectory using a near-optimal strategy by dividing the entire flight into an approach phase to a circular orbit of the same radius as the desired final orbit but with a smaller inclination, and a subsequent cranking phase until the desired (orbital) inclination is reached. The numerical simulations show that the proposed strategy is sufficiently simple to implement and can provide solutions that differ by only a few percentage points from the optimal results obtainable with a classical indirect approach.

Diffractive Sail-Based Displaced Orbits for High-Latitude Environment Monitoring

This paper analyzes the possibility of maintaining a circular displaced non-Keplerian orbit around the Sun by means of a Sun-facing diffractive sail. With the goal of monitoring the Earth’s high-latitude regions, the spacecraft is required to track its displaced orbit at an angular velocity equal to the mean motion of the planet. In doing so, the spacecraft keeps a constant average phase shift with respect to Earth’s angular position along its orbit, allowing the objectives of the scientific mission to be achieved. The diffractive sail, recently proposed by Swartzlander and chosen in this paper as the spacecraft’s primary propulsion system, is a special photonic solar sail in which the membrane film is covered by an advanced diffractive metamaterial. In particular, a Sun-facing diffractive sail with a grating at normal incidence generates radial and transverse thrust components of equal magnitude; that is, the thrust vector is tilted 45 degrees from the Sun-spacecraft line. This peculiarity enables the diffractive sail to maintain a family of circular displaced non-Keplerian orbits, each of which is characterized by unique values of radius and a lightness number for an assigned value of spacecraft displacement relative to the Ecliptic. A comparison with the ideal reflecting sail shows that the diffractive sail performs better because for the same overall spacecraft mass, the latter needs about 30% less surface area exposed to the Sun. Finally, this paper discusses the classical stability problem, assuming an error in orbit insertion of the diffractive sail-based spacecraft. In this context, extensive numerical simulations show that such displaced orbits are marginally stable.

Analysis on increasing energy efficiency in regional railways by the introduction of permanent magnet synchronous motors

Enhancing efficiency has become a fundamental requirement across contemporary technological endeavors aimed at energy conversion. The imperative lies in achieving global energy resource preservation objectives. Novel methods of energy utilization with heightened efficiency are progressively gaining traction, accompanied by the intricacies of refining existing technologies. This research undertakes an examination of the amplified energy efficiency within regional rail systems, achieved through the adoption of permanent magnet synchronous motors (PMSMs) as primary electric propulsion systems, particularly when juxtaposed with induction motors. Delving into the theoretical engineering attributes of PMSMs, coupled with the establishment of dynamic motion characteristics and comprehensive simulations, offers a pathway to amassing empirical insights pertaining to efficiency studies.

Solar Sail Orbit Raising with Electro-Optically Controlled Diffractive Film

The aim of this paper is to analyze the transfer performance of a spacecraft whose primary propulsion system is a diffractive solar sail with active, switchable panels. The spacecraft uses a propellantless thruster that converts the solar radiation pressure into propulsive acceleration by taking advantage of the diffractive property of an electro-optically controlled (binary) metamaterial. The proposed analysis considers a heliocentric mission scenario where the spacecraft is required to perform a two-dimensional transfer between two concentric and coplanar circular orbits. The sail attitude is assumed to be Sun-facing, that is, with its sail nominal plane perpendicular to the incoming sunlight. This is possible since, unlike a more conventional solar sail concept that uses metalized highly reflective thin films to reflect the photons, a diffractive sail is theoretically able to generate a component of the thrust vector along the sail nominal plane also in a Sun-facing configuration. The electro-optically controlled sail film is used to change the in-plane component of the thrust vector to accomplish the transfer by minimizing the total flight time without changing the sail attitude with respect to an orbital reference frame. This work extends the mathematical model recently proposed by the authors by including the potential offered by an active control of the diffractive sail film. The paper also thoroughly analyzes the diffractive sail-based spacecraft performance in a set of classical circle-to-circle heliocentric trajectories that model transfers from Earth to Mars, Venus and Jupiter.

Cluster of electric thrusters for astronautic and robotic INPPS flagship space flights to Mars and Europa moon

This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship.Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

E-Sail Optimal Trajectories to Heliostationary Points

The aim of this paper is to investigate the performance of a robotic spacecraft, whose primary propulsion system is an electric solar wind sail (E-sail), in a mission to a heliostationary point (HP)—that is, a static equilibrium point in a heliocentric and inertial reference frame. A spacecraft placed at a given HP with zero inertial velocity maintains that heliocentric position provided the on-board thrust is able to counterbalance the Sun’s gravitational force. Due to the finite amount of storable propellant mass, a prolonged mission toward an HP may be considered as a typical application of a propellantless propulsion system. In this respect, previous research has been concentrated on the capability of high-performance (photonic) solar sails to reach and maintain such a static equilibrium condition. However, in the case of a solar-sail-based spacecraft, an HP mission requires a sail design with propulsive characteristics that are well beyond the capability of current or near-future technology. This paper shows that a medium-performance E-sail is able to offer a viable alternative to the use of photonic solar sails. To that end, we discuss a typical HP mission from an optimal viewpoint, by looking for the minimum time trajectory necessary for a spacecraft to reach a given HP. In particular, both two- and three-dimensional scenarios are considered, and the time-optimal mission performance is analyzed parametrically as a function of the HP heliocentric position. The paper also illustrates a potential mission application involving the observation of the Sun’s poles from such a static inertial position.

Solenoid Engine

Given the impact on the environment as well as financial considerations like the gradually rising cost of liquid fossil fuels, maintenance, and other costs, electric vehicles are becoming a more and more appealing option to vehicles with combustion engines. Due to the fact that these cars are well renowned for having zero emissions and being fueled by renewable energy. The project's goal is to replace the current electric motor with an alternate electric vehicle primary mover. Generally speaking, EVs are propelled and controlled by the fusion of electrical, electronic, and mechanical components, but their primary propulsion system is its electric motor. By transforming electrical energy into kinetic energy, an electric motor operates according to the electromagnetic induction theory. An electric motor's primary function is energy conversion, and this actuator is widely used in most EV designs. Therefore, the electric motor will be replaced with a solenoid as the prime mover. A solenoid prototype is created, constructed, and tested for this. A kicking mechanism will be made of the solenoid. A solenoid has been studied as a shooting mechanism in prior investigations. One study looks into the solenoid as the best kicking mechanism. The solenoid was created and optimised in the other investigation. A prototype solenoid is created and tested in this study.

Optimal Circle-to-Ellipse Orbit Transfer for Sun-Facing E-Sail

The transfer between two coplanar Keplerian orbits of a spacecraft with a continuous-thrust propulsion system is a classical problem of astrodynamics, in which a numerical procedure is usually employed to find the transfer trajectory that optimizes (i.e., maximizes or minimizes) a given performance index such as, for example, the delivered payload mass, the propellant mass, the total flight time, or a suitable combination of them. In the last decade, this class of problem has been thoroughly analyzed in the context of heliocentric mission scenarios of a spacecraft equipped with an Electric Solar Wind Sail as primary propulsion system. The aim of this paper is to further extend the existing related literature by analyzing the optimal transfer of an Electric Solar Wind Sail-based spacecraft with a Sun-facing attitude, a particular configuration in which the sail nominal plane is perpendicular to the Sun-spacecraft (i.e., radial) direction, so that the propulsion system is able to produce its maximum propulsive acceleration magnitude. The problem consists in transferring the spacecraft, which initially traces a heliocentric circular orbit, into an elliptic coplanar orbit of given eccentricity with a minimum-time trajectory. Using a classical indirect approach for trajectory optimization, the paper shows that a simplified version of the optimal control problem can be obtained by enforcing the typical transfer constraints. The numerical simulations show that the proposed approach is able to quantify the transfer performance in a parametric and general form, with a simple and efficient algorithm.

Station-keeping for a solar sail during lander/probe deployment using feedback control

Due to its propellantless nature, a solar sail can provide the primary propulsion system for a high energy mission, such as that of a multiple asteroid rendezvous. Upon arrival at an asteroid, it is often desirable to interact with the surface of the body, such as for sample extraction. The deployment of a lander from a solar sail carries the difficulty of an instantaneous, and sometimes considerable, change to the system dynamics at the point of separation. This paper investigates the effects of changing sail performance during the release of multiple “ChipSat” probes as well as a large MASCOT-type lander and the control of the sail into a positional hold at an equilibrium point or periodic orbit. In one scenario, 20 ChipSat probes are released, with one-hour spacing between each release. The sail is then controlled to maintain the sailcraft close to the initial deployment point. The performance of a Linear Quadratic Regulator (LQR) is compared with maintaining a fixed sail attitude after deployment. In a second scenario, at the point of separation of the larger MASCOT-type lander, there will be a considerable instantaneous change in the sail characteristic acceleration, as opposed to the gradual small change for the staggered deployment of the small ChipSat probes. It is shown that the Time-Delayed Feedback Control (TDFC) method is effective in controlling the orbit of the sail after this deployment. The sail converges to an orbit in the same region of phase space when deployment is made from both a lower and higher inclination orbit.

Enhancing one-dimensional particle-in-cell simulations to self-consistently resolve instability-induced electron transport in Hall thrusters

The advent of high-power Hall thrusters and the increasing interest toward their use as a primary propulsion system for various missions have given a new boost to the efforts aiming at self-consistent predictive modeling of this thruster technology. In this article, we present a novel approach, which allows enhancing the predictive capability of one-dimensional particle-in-cell (PIC) simulations to self-consistently capture the wave-induced electron transport due to the azimuthal instabilities in Hall thrusters. The so-called “pseudo-2D” PIC scheme resulting from this approach is extensively tested in several operating conditions. The results are compared against a well-established 2D3V axial–azimuthal reference case in terms of the axial profiles of the time-averaged plasma properties, the azimuthal electric field fluctuations and their dispersion features, and the contributions of the force terms in the electron azimuthal momentum equation to the cross-field mobility. We have demonstrated that the pseudo-2D PIC provides a prediction of the above aspects that compares very closely in almost all conditions with those from the full-2D simulation. In addition, the sensitivity of the pseudo-2D simulation results to the numerical parameters associated with our approach is assessed in detail. The outcomes of these analyses have casted light on the next steps to further improve the approach.

Approximate solution of spacecraft motion with an inward constant radial acceleration

This Short Communication analyzes the motion of a spacecraft whose primary propulsion system generates an inward, radial propulsive acceleration of constant magnitude. Starting from the classical literature results regarding the more common outward radial-thrust case, the proposed mathematical model uses a set of modified non-singular orbital elements to describe the spacecraft motion around a generic primary body. Assuming a circular parking orbit and a small propulsive acceleration magnitude, a perturbation method, based on an asymptotic series expansion truncated at second order, is used to obtain analytical expressions of the spacecraft propelled trajectory. The obtained results are applied to an orbital targeting problem and, in this context, the resulting spacecraft trajectory is accurately approximated through analytical functions. The proposed procedure is also able to well estimate the thrust level and the swept angle required to reach the target state.

Extended Lifetime of CubeSats in the Lower Thermosphere with Active Attitude Control

A wide variety of missions could be enabled by extended orbital flight in extreme low Earth orbit, defined as an altitude range of 150–250 km. This study investigates the feasibility of a nanosatellite (mass ) using a propulsive, attitude control system in conjunction with a primary propulsion system to extend mission life. The primary propulsion system consists of a pair of electrospray thrusters providing a combined thrust of 0.12 mN at 1 W. Pulsed plasma thrusters are used for attitude control. The mission consists of two phases. In Phase I, a 4U CubeSat is deployed from a 414 km orbit and uses the primary propulsion system to deorbit to an initial altitude within the targeted range of . Phase I lasts 12.73 days, with the propulsion system consuming 5.6 g of propellant to deliver a of 28.12 m/s. In Phase II the mission is maintained until the remaining 25.2 g of propellant is consumed. Phase II lasts for 30.27 days, corresponding to a of 57.22 m/s with a mean altitude of 244 km. Using this approach, a primary mission life of 30.27 days could be achieved, compared with 3.1 days without primary propulsion.

Accessibility assessment and trajectory design for multiple Near-Earth-asteroids exploration using stand-alone CubeSats

Continuous breakthrough of related technologies enables stand-alone interplanetary CubeSats, requiring primary propulsion systems for orbit maneuvering and trajectory control, to be applied in deep space missions. The application of stand-alone interplanetary CubeSats to explore multiple Near-Earth asteroids (NEAs) is investigated in this paper. Indirect and direct transfer trajectories to rendezvous with NEAs are designed assuming CubeSats are firstly parked on the Sun-Earth Halo orbits. To select candidate target asteroids that CubeSats could potentially reach, two assessment methods of the accessibility of NEAs are proposed based on the characteristic of the Sun-Earth Circular Restricted Three-Body Problem (CRTBP). Compared to the traditional methods, the two methods derive analytical expressions of the velocity increments of the transfer trajectories, which can simplify the accessibility assessments. By applying the assessment methods, the list of asteroids, ranked with respect to the accessibility of NEAs between the years 2020 and 2031, is produced. Then the transfer trajectories from the Sun-Earth Halo orbits to asteroids are calculated by numerical optimization method in the CRTBP model. It can be found that the two methods could serve as the effective tools to assess the accessibility of asteroids and the accuracy of the assessment methods of indirect transfer and direct transfer are ±0.02 km/s and ±0.1 km/s, respectively.

On the Expansion of On-Board Diagnostics (OBD) to Electric Propulsion Systems in Battery Electric Vehicles

<div class="section abstract"><div class="htmlview paragraph">Currently the On-Board Diagnostics (OBD) requirements enforced by government agencies do not cover electric vehicles. Although the California Air Resources Board (CARB) mandates all light and medium duty vehicles and heavy duty engine dynamometer certified engines equipped with fossil fuel-powered engines, including all hybrid vehicles, must follow the OBD requirements in California Code of Regulation (CCR) 1968.2 and 1971.1, Battery Electric vehicles (BEVs), are exempted from OBD requirements. The legislators, such as CARB, have started to make proposals for on-board systems to monitor electric propulsion system health. In addition, there may be customer needs to obtain standard vehicle service information and the Original Equipment Manufacturers (OEMs) may also have the desire for common diagnostic strategies across different vehicle applications to lower the development costs. This paper, different from other published technical papers or proposals from OEMs or government agencies, analyzes the feasibility and compatibility of the existing OBD system design requirements, and aligns them with the need of industrial-common OBD system architecture and existing/upcoming SAE standards to propose the monitoring and reporting requirements on the identified primary propulsion system components/systems from technical point of view. This study makes some key technical proposals/solutions for the possible use of OBD-like monitoring and reporting requirements for Electric Propulsion Systems, including, what are the monitoring requirements, what is to be monitored, what are the malfunction indicator light (MIL) illumination requirements, and what are the generic scan tool (GST) interface requirements that are based on J1979-2 (i.e., Unified Diagnostic Services -UDS) to facilitate the repair service by technicians, including from independent service shops.</div></div>

Review of State-of-the-Art Green Monopropellants: For Propulsion Systems Analysts and Designers

Current research trends have advanced the use of “green propellants” on a wide scale for spacecraft in various space missions; mainly for environmental sustainability and safety concerns. Small satellites, particularly micro and nanosatellites, evolved from passive planetary-orbiting to being able to perform active orbital operations that may require high-thrust impulsive capabilities. Thus, onboard primary and auxiliary propulsion systems capable of performing such orbital operations are required. Novelty in primary propulsion systems design calls for specific attention to miniaturization, which can be achieved, along the above-mentioned orbital transfer capabilities, by utilizing green monopropellants due to their relative high performance together with simplicity, and better storability when compared to gaseous and bi-propellants, especially for miniaturized systems. Owing to the ongoing rapid research activities in the green-propulsion field, it was necessary to extensively study and collect various data of green monopropellants properties and performance that would further assist analysts and designers in the research and development of liquid propulsion systems. This review traces the history and origins of green monopropellants and after intensive study of physicochemical properties of such propellants it was possible to classify green monopropellants to three main classes: Energetic Ionic Liquids (EILs), Liquid NOx Monopropellants, and Hydrogen Peroxide Aqueous Solutions (HPAS). Further, the tabulated data and performance comparisons will provide substantial assistance in using analysis tools—such as: Rocket Propulsion Analysis (RPA) and NASA CEA—for engineers and scientists dealing with chemical propulsion systems analysis and design. Some applications of green monopropellants were discussed through different propulsion systems configurations such as: multi-mode, dual mode, and combined chemical–electric propulsion. Although the in-space demonstrated EILs (i.e., AF-M315E and LMP-103S) are widely proposed and utilized in many space applications, the investigation transpired that NOx fuel blends possess the highest performance, while HPAS yield the lowest performance even compared to hydrazine.

Solar sails for perturbation relief: Application to asteroids

A major cause of spacecraft orbital variation comes from natural perturbations, which, in close proximity of a body, are dominated by its non-spherical nature. For small bodies, such as asteroids, these effects can be considerable, given their uneven (and uncertain) mass distribution. Solar sail technology is proposed to reduce or eliminate the net secular effects of the irregular gravity field on the orbit. Initially, a sensitivity analysis will be carried out on the system which will show high sensitivity to changes in initial conditions. This presents a challenge for optimisation methods which require an initial guess of the solution. As such, the Genetic Algorithm (GA) is proposed as the preferred optimisation method as this requires no initial guess from the user. A multi-objective optimisation is performed which aims to achieve a periodic orbit whilst also minimising the effort required by the sail to do so. Given the system sensitivity, the control law for one orbit is not necessarily applicable for any subsequent orbit. Therefore, a new method of updating the control law for subsequent orbits is presented, based on linearisation and use of a Control Transition Matrix (CTM). The techniques will later find application in a multiple asteroid rendezvous mission with a solar sail as the primary propulsion system.