Recent Advances in Space Propulsion Technology

This article aims to conduct an in-depth investigation into the environmental impact of Hall thruster plumes on spacecraft surface charging. The non-uniform plasma plume generated by Hall thrusters may trigger charging and discharging effects, making the assessment of surface charging risks crucial. Through numerical simulations using SPIS system, this study evaluates the surface charging characteristics of a complex spacecraft in orbit, simulating the effects of turning on and off the thrusters, as well as varying distances between the thrusters and the spacecraft. The simulation demonstrates that turning on the thrusters significantly affects spacecraft charging, reducing the potential difference between spacecraft surfaces from 3740 V to 19.2 V, effectively alleviating electrostatic discharge on the spacecraft surface. The closer the thruster is to the spacecraft, the more CEX ions are collected on the surface, influenced by the beam ions, resulting in a surface potential change of 1.3 V, with minor effects on surface potential but contributing to increased deposition contamination on the spacecraft surface. Full article

The transfer between two circular, coplanar Keplerian orbits of a spacecraft equipped with a continuous thrust propulsion system is usually studied in an optimal framework by maximizing a given performance index. Using an indirect approach, the optimal trajectory and the maximum value of the performance index are obtained by numerically solving a two-point boundary value problem (TPBVP). In this context, the computation time required by the numerical solution of the TPBVP depends on the guess of unknown initial costates. The aim of this paper is to describe an analytical procedure to accurately approximate the initial costate variables in a coplanar, circle-to-circle, minimum-time transfer. In particular, this method considers a freely steerable propulsive acceleration vector, whose magnitude varies over a finite range with a sufficiently low maximum value. The effectiveness of the analytical method is tested in a set of both geocentric and heliocentric (simplified) mission scenarios, which model the classical LEO-GEO or interplanetary transfers toward Venus, Mars, Jupiter, and comet 29P/Schwassmann–Wachmann 1. Full article

The problem of noise reduction in supersonic aircraft design is one of the key problems, the solution of which largely determines the speed of development of supersonic aviation as a whole. The present study examines the noise generation during flights of supersonic civil aircraft. The effect of a thermally stratified energy source (TSS) used to control the supersonic flow past a pointed cylinder aerodynamic model on the near-field and ground pressure signatures, as well as on the perceived loudness in decibels (PLdB) on the ground, is evaluated. The complex conservative difference schemes, Tomas’ waveform parameter method, and Stevens’ algorithm Mark VII are used for near-field modeling, obtaining the ground pressure signature, and the evaluation of the PLdB on the ground, accordingly. The fields of flow parameters and the dynamics of a drag force are researched at the variation of temperatures in layers of TSS and for different numbers of layers. Simulations showed that changing the surface pressure due to drag reduction does not necessarily imply a change in the PLdB on the ground. In particular, it has been shown that when performing the flow control at freestream Mach numbers 1.5–2 using TSSs with the number of layers from 2.5 to 7.5 and rarefaction parameters in the layers from 0.15 to 0.3, some weakening of the bow shock wave in the near-field pressure signature due to the effect of TSS occurs, and no additional noise impact on the ground is introduced. Full article

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. Full article

Gridded ion thrusters (GITs) are an established technology that, by covering a wide range of power class, allows one to accomplish a lot of space mission types. Many analysis tools and analytical models describing the physics of GITs are present in the open literature, while there is a lack of tools for preliminary design, considering the mission requirements (i.e., thrust or power). Thus, in this work, a tool that takes as input thrust or power and that combines analytical formulas, describing GITs’ physics; a curve-fitting approach, exploiting data from different ion thrusters present in the open literature; and an FEMM (finite element method magnetics) simulation has been developed and validated against known medium-low-power (<5 kW) gridded ion thrusters (e.g., NSTAR, XIPS, ETS-8). Some of the main outputs of the developed tool are its specific impulse, efficiencies, voltages, and propellant flow rate. The results obtained by the tool have been in good agreement with the real performance and working parameters of the thrusters selected for the validation, obtaining an average error of less than 5–10%. The tool has been also compared with a tool proposed in the literature as a possible design tool, which makes use of a simple macroscopic plasma-source simulation (SMPS) code with a genetic algorithm (GA) and obtains slightly more accurate results on average. Finally, the tool has been exploited for the design of a very low-power GIT (100 W) that is able to produce 2 mN of thrust, as the interest of the scientific community in miniaturizing electric engines has recently grown because they could enable new space missions. Full article

In recent years, there has been a considerable growth in the demand for low-orbit satellites, leading to a need for more flexible and cost-effective launch systems. This study presents a low-cost “carrier-launcher” configuration designed for space missions in low earth orbit. The carrier is a remote-controlled unmanned flying wing that can fulfil the role of the first stage of a multi-stage earth-to-orbit launcher rocket. Making the carrier a flying wing increases its effectiveness and efficiency compared to other state-of-the-art options. The flying wing architecture allows for a significantly lighter carrier compared to the traditional aircraft. The launcher is carried on the wing’s upper surface and is released during a high-altitude almost “zero g” parabolic manoeuvre. A state-of-the-art analysis has been conducted to initialize and develop the carrier’s conceptual configuration. The aerodynamics and flight mechanics of the flying wing carrier were studied using the potential aerodynamic code Athena Vortex Lattice. The high-altitude launcher’s release manoeuvre has been investigated to properly assess the required installed thrust. Finite element analyses were also performed using NASTRAN to preliminarily evaluate the aeroelastic behaviour of the proposed “carrier-launcher” configuration. The overall results show the conceptual feasibility of the flying wing carrier for launching small satellites in low earth orbit. This study provides valuable insights into the development of cost-effective launch systems for the growing demand in the low-orbit satellite sector. Our proposed design has a maximum take-off mass of 122,000 kg, uses 4 Rolls-Royce UltraFan model engines, has a wingspan of 54 m, and can carry a 10,000 kg launcher to put a 460 kg payload in LEO. As it is an initial conceptual study, this investigation establishes an initial benchmark for forthcoming inquiries, hence providing a starting point of a breakthrough concept to foster its future development. Full article

The Solar Wind Ion Focusing Thruster (SWIFT) is a highly-innovative propellantless propulsion concept, recently proposed by Gemmer and Mazzoleni. In its nominal configuration, a SWIFT consists of a conically-shaped mesh of positively-charged conducting tethers, with its vertex linked to the spacecraft and its axis oriented towards the Sun. The SWIFT collects and filters the solar wind plasma and suitably directs the positive ions, which are then accelerated by an ion thruster. Such a device is theoretically able to generate a deep-space propulsive acceleration that comes, in part, from the solar wind dynamic pressure impinging on the conical grid and, in part, from the positive ion beam. In particular, the orientation of the ion beam may be chosen in such a way as to set the resultant propulsive acceleration and steer the spacecraft. The aim of this paper is to analyze the performance of a SWIFT-propelled spacecraft in an orbit-to-orbit two-dimensional interplanetary transfer. To that end, some mission scenarios are studied, in an optimal framework, by minimizing the total flight time necessary for the spacecraft to complete the transfer as a function of the propulsion system performance parameters. Numerical simulations are used to compare the optimal flight times calculated in simplified Earth–Venus and Earth–Mars transfers with those obtained by considering other propellantless propulsion systems. Full article

In order to support increasingly detailed simulation models in the field of plasma generation and expansion, more experimental data are needed to prove and verify these models. In addition to existing invasive probe methods and spectroscopy, optical diagnostics methods can be employed for this purpose. These can be used to obtain information about the transient behavior of the plasma plume itself. For this purpose, a pseudo stereo imaging system was assembled, which consisted of a biprism and one camera. The presented setup was used to observe the plasma plume expansion of a vacuum arc thruster and to detect how a magnetic nozzle changes the plume expansion behavior of the used thruster. For this, areas were calculated by means of contour detection from both view directions, and a 3D model was reconstructed by computing cross-sectional images. The results show that the magnetic nozzle provides a more uniform formation of the plasma plume, which can be seen by comparing the calculated areas from both directions of observation and is confirmed by the reconstructed 3D models. Thus, the used setup is suitable as an additional diagnostic tool in the future. Despite its simple design, it provides information about the spatial development of a plasma plume. It can serve as a fast and simple verification tool for simulation results, where otherwise complex tomographic setups and reconstructions would be necessary to obtain spatial information. Further measurements are recommended to improve and validate the recording and evaluation process. Full article