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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.

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 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.

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.

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%.

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.

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.

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)

A solar-pumped laser (SPL) that converts sunlight directly into a coherent and intense laser beam generally requires a large concentrating lens and precise solar tracking, thereby limiting its potential utility. Here, we demonstrate a fully-planar SPL without a lens or solar tracking. A Nd3+-doped silica fiber is coiled into a cylindrical chamber filled with a sensitizer solution, which acts as a luminescent solar collector. The body of the chamber is highly reflective while the top window is a dichroic mirror that transmits incoming sunlight and traps the fluorescence emitted by the sensitizer. The laser-oscillation threshold was reached at a natural sunlight illumination of 60% on the top window. Calculations indicated that a solar-to-laser power-conversion efficiency could eventually reach 8%. Such an SPL has potential applications in long-term renewable-energy storage or decentralised power supplies for electric vehicles and Internet-of-Things devices. Solar-pumped laser systems are attractive for applications including hydrogen generation and space propulsion, but current technologies are cumbersome and rely on accurate tracking of the sun’s light. Here, lasing is achieved using a planar, luminescent solar collector removing the need for lenses or tracking.

Satelliten und andere Raumfahrzeuge werden häufig mit chemischen Kleintriebwerken ausgestattet, die zur Anhebung oder Absenkung des Orbits, zur Lageregung und gegebenfalls auch zum Manövrieren bei Andockmanövern oder der Landung auf Himmelskörpern verwendet werden. Die Verwendung von mehreren kleineren Triebwerken in einer Clus- teranordnung, auch Triebwerksbündel genannt, ist dabei nicht unüblich. Die unter Weltraumvakuumbedingungen frei expandierenden Einzeltreibstrahlen des Clusters treten bei genügend kleinem Abstand zueinander in Wechselwirkung, was zur Ausbildung eines Sekundärtreibstrahls und zu einer erhöhten Beaufschlagung des Raumfahrzeugs stromauf der Düsenaustrittsebene führt. Bei einem Lande- oder Andockmanöver kommt es zudem zu einer Wechselwirkung zwischen den Treibstrahlen und der stromab gelegenen Oberfläche. Diese Wechselwirkungsphänomene wurden in der Literatur bereits untersucht, häufig im Rahmen von analytischen Be- trachtungen oder numerischen Simulationen. Experimentelle Untersuchungen eines frei expandierenden Treibstrahls unter Weltraumvakuumbedingungen sind nur aus der STG-CT des DLR Göttingen bekannt. Dort wurden in den letzten Jahren experimentelle und numerische Untersuchungen zur Expansion von zwei benachbarten Düsen angefertigt. Anknüpfend an diese Arbeiten wird in der vorliegenden Arbeit die Treibstrahl-Treibstrahl- und Treibstrahl-Oberflächen-Wechselwirkung innerhalb eines Clusters aus vier identischen Kaltgastriebwerken untersucht. Ziel der vorliegenden, neuen Arbeiten ist die experimentelle Charakterisierung der Wechselwirkungen, ihre Interpreta- tion auf Basis vorhandener Literaturdaten sowie die Gegenüberstellung neuer mit bereits bekannten Konfigurationen. Insbesondere werden die Zusammenhänge zwischen den Strukturen von Einzel- und Mehrfachtreibstrahlen, sowie zwischen diesen Freistrahlen und der Beaufschlagung auf stromauf und -ab gelegenen Oberflächen betrachtet. Im Rahmen dieser Arbeit wird dabei vordergründig die axiale Expansion und angulare Molekülstromdichtenverteilung der Treibstrahlen sowie die räumliche Verteilung der Beaufschlagungsdrücke auf stromauf und -ab positionierten Platten untersucht. Dazu wird die neue experimentelle Konfiguration mittels numerischer Berechnungen und experimenteller Daten auf ihre Übertragbarkeit zu bereits vorliegenden Arbeiten untersucht. Das Fernfeld des Einzel- und Mehrfachtreibstrahls wird vermessen und beschrieben, auch im Hinblick auf die Beeinflussung durch Ruhedruck im Triebwerk und Um- gebungsdruck in der Kammer. Die stromauf gelegene Beaufschlagung (Rückströmung) wird vermessen und ihre Beeinflussung durch die Anzahl der verwendeten Triebwerke, den Ruhe- und Umgebungsdruck sowie die Nähe zu einer stromab gelegenen Oberfläche untersucht. Zuletzt wird das Beaufschlagungsdruckprofil auf einer Oberfläche stromab der Düsenaustrittsebene des einzelnen Triebwerks und des Triebwerksclusters vermessen und dem Fall des frei expandierenden Treibstrahls gegenübergestellt. Es konnten konsistente experimentelle Daten gewonnen werden, die in späteren Arbeiten mit DSMC-Simulationen kombiniert werden können, um letztere zu validieren und zugleich die im Experiment gewonnenen Erkenntnisse zu er- weitern. DSMC steht für Direct Simulation Monte-Carlo, es ist ein numerisches Verfahren zur Lösung der Boltzmann- Gleichung in Nicht-Kontinnummsströmungen mittels stochastischer Methoden. Der wechselwirkende Treibstrahl des 4-Düsenclusters zeigt deutliche Unterschiede zum Einzeltreibstrahl, aber auch zum wechselwirkenden Treibstrahl aus zwei Düsen. Im in dieser Arbeit betrachteten 4-Düsencluster schließen deren Einzeltreibstrahlen den Sekundär- treibstrahl, also den Strömungsbereich stromab des ersten Wechselwirkungsstoßes, nach allen Seiten ab, wodurch die Rückströmung innerhalb des Clusterzentrums überproportional verstärkt wird. Das Gas strömt frei aus und isoliert das Clusterinnere von Änderungen der Umgebung. Selbst unter den hier betrachteten frei-molekularen Bedingungen werden auf stromab gelegenen Beaufschlagungsplatten Kontinuumsbedingungen und damit die Ausbildung von Bo- denstößen beobachtet. Bemerkenswert ist die komplexe Interaktion des Mehrfachtreibstrahls mit dem Bodenstoß. Es können verschiedene, ineinander übergehende Phasen dieser Wechselwirkung in Abhängigkeit der Flughöhe experimen- tell nachgewiesen und gasdynamisch beschrieben werden. Bei großen Flughöhen dominiert der Sekundärtreibstrahl, bei sehr kleinen Flughöhen die einzelnen Düsen des Clusters die Beaufschlagung. Dazwischen, im Entstehungsbereich des Sekundärtreibstrahls und bevor die Strömung so stark expandiert ist, dass er von aufgeweiteten statt diskreten Stößen begrenzt ist, wird eine komplexe Wechselwirkung zwischen den Treibstrahlen und dem stromab sowie stromauf gelegenen beaufschlagten Oberflächen, also allen Komponenten des betrachteten Systems, beobachtet.

In this paper, the laser-accelerated plasma–propulsion system (LAPPS) for a spacecraft is revisited. Starting from the general properties of relativistic propellants, the relations between specific impulse, engine thrust and rocket dynamics have been obtained. The specific impulse is defined in terms of the relativistic velocity of the propellant using the Walter’s parameterization, which is a suitable and general formalism for closed–cycle engines. Finally, the laser-driven acceleration of light ions via Target Normal Sheath Acceleration (TNSA) is discussed as a thruster. We find that LAPPS is capable of an impressive specific impulse Isp in the 105 s range for a laser intensity I0≃1021W/cm2. The limit of Isp≲104 s, which characterizes most of the other plasma-based space electric propulsion systems, can be obtained with a relatively low laser intensity of I0≳1019W/cm2. Finally, at fixed laser energy, the engine thrust can be larger by a factor 102 with respect to previous estimates, making the LAPPS potentially capable of thrust-power ratios in the N/MW range.

A cikkben a szerzők bemutatják az Airbus H145M típusú könnyű harci helikopter fegyverzetét. Rövid fejlesztési történetet követően a helikopteren alkalmazható különböző fegyverrendszerekkel foglalkozunk, többek között tűzfegyverekkel, nemirányítható és irányítható rakétákkal, a típushoz rendszeresített rakéták rakétahajtóművével, valamint a helikopterfedélzeti rakéták irányításával.

Neutronic analysis of a conceptual heat pipe-cooled micronuclear reactor with 70 % enriched uranium nitride fuel is carried out by modeling the core and peripheral control drum movement to estimate the power distribution. The core configuration results in non-uniformities and hotspots. For the heat removal, empirical formulae have been used in the case of sodium, lithium, and potassium working fluids. The neutronic simulation was carried out by the OpenMC code. It has been found that the radial flux peaking as high as ~20 % can occur at various stages of the drum movement. The novelty of this research is the investigation of the effect of variable enrichment on the overall system multiplication, which can form the basis for optimal fuel distribution. It has been found that non-uniform fuel distribution can mitigate peaking factors, and thus reduce the hotspots. This analysis is useful for the design optimization of compact micro nuclear reactors for underwater, portable and space propulsion systems.

This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100 N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.

The electric sail is an innovative concept for spacecraft propulsion, which can generate continuous thrust without propellant by reflecting solar wind ions. In previous studies, the thrust of an electric sail is described by a classical model that neglects the effects of the electric sail attitude on the propulsive thrust modulus and direction. This paper reappraised the performance of the electric sail in the Vesta and Ceres exploration mission with an advanced thrust model that considers the effect of the spacecraft attitude on both the thrust modulus and direction. By using a hybrid optimization method, the trajectory optimization of the electric-sail-based spacecraft from earth to Vesta and Ceres is implemented in an optimization framework. Numerical results show that the minimal flight time with the advanced thrust model is longer than that with the classical model. The difference in performance between the classical and advanced models is attributable to overestimation of the maximum thrust cone angle and the thrust modulus by the classical model.

3D Carbon/Carbon composites have an important use in space propulsion and atmospheric re-entry of space objects. We propose a modeling approach for its non-linear mechanical and thermomechanical behavior based on the introduction of internal interfaces, under the form of cohesive and sliding zones located between the macro-constituents (bundles, matrix pockets). The interface model parameters have been identified from bundle push-out experiments at temperatures from ambient to 1000∘C. The model allowed reproducing successfully a 45∘ off-axis tensile test, only with initial damage and interface sliding. The sole incorporation of interface sliding as the only nonlinear phenomenon already allows to successfully reproduce the off-axis behavior. It is also correctly found that the material has a higher yield limit at high temperature, because its interfaces are closing due to thermal expansion of the bundles. The validity of the model does not encompass yet cases where progressive damage of the bundles occurs, as e.g. in torsion tests.

One of the issues of using a normal graphite nozzle for the hybrid rocket propulsion is the serious throat erosion due to combustion with a higher O/F ratio. This may undermine the throttling capability of the hybrid rocket engine. In this study, we would like to address how the propulsion performance changes under the conditions of different nozzle throat diameters and O/F ratios. We designed and tested a 40-kgf class single-port hybrid rocket motor with the swirling injection of oxidizer, which utilizes the 90 wt% hydrogen peroxide and polypropylene (PP) as oxidizer and fuel, respectively. Three different diameters of graphite nozzle throat (10, 11, and 12 mm) were used to approximate various conditions of nozzle erosion, while the mass flow rates of the injected oxidizer were kept the same. The hot-fire test results indicated that the thrust was nearly the same even the nozzle throat diameters and O/F ratios were different. Especially, the discrepancies among the measured thrusts and oxidizer ISP under various test conditions were found to be less than 1%, which is highly beneficial for the purpose of the thrust control using mass flow rate control of the oxidizer for a hybrid rocket engine with swirling injection of high-concentration hydrogen peroxide.

Evaluating the climbing performance of rocket-based combined cycle (RBCC) powered aircraft has a guiding role for single-stage to orbit. This study establishes a new modularized thrust model of the RBCC engine for low speeds (Ma∞ = 0.8–3) based on independent ramjet stream and pressure matching. The model can quickly estimate the quasi one-dimensional flow characteristics of secondary flow and primary flow at given runner geometry and operating conditions. Comparison with experimental results shows that the thrust model is correct. In this model, the influence of altitude, Mach number, and rocket flow on thrust and specific impulse is considered. The model was used to optimize the climbing trajectory of RBCC powered aircraft with a typical lifting-body. The comparison of RBCC propulsion and pure rocket propulsion shows that the former offers propellant saving advantages. During the climb, the RBCC powered aircraft needs to strictly control the air flow rate and ensure the pressure matching in the flow path to improve propulsion efficiency.

For the human species and its unique culture to survive the death of the sun, a bridge must be built to other solar systems with earthlike planets. The Kepler space telescope has discovered a large number of extrasolar planets, but only a few with earthlike conditions, and those are many light years away. Assuming that no new fundamental laws of physics which would greatly facilitate interstellar spaceflight are awaiting discovery, one can only see two avenues: First, at 10% of the speed of light via deuterium fusion bomb propulsion, harvesting the deuterium in the comets of the Oort clouds surrounding our and other suns, and by hopping from comet to comet. Second, with relativistic velocities by matter-antimatter generated GeV laser beams released from relativistically stabilized hydrogen-antihydrogen super-pinch discharges, transmitting the recoil of the laser beam by the Mossbauer effect to the spacecraft. The production of the anti-hydrogen can be done with solar energy in robotic factories on the planet Mercury. For the first, but much more for the second possibility, very large masses must be lifted in one stage into low Earth orbit. This can conceivably be done by chemically ignited pulsed pure fusion bomb propulsion.

Advances in miniaturization using micromachining processes have led to propulsion systems small enough to consider the feasibility of carrying a large number of thrusters even on small spacecrafts such as Cubesats. Electrospray thrusters developed at the Space Propulsion Laboratory of the Massachusetts Institute of Technology are composed of a highly miniaturized emitter array attached to a tank structure. In terms of volume and mass, the thrusters are small compared to the tank and overall system. This feature makes it possible to envision a staging concept, in which multiple propulsion units are powered in succession, with staging of those that have depleted their propellant. As the satellite mission advances, such a staging operation reduces the spacecraft structural mass, leading to an increased total Δ v capability compared to traditional mission designs. This work examines the impact of such an operational concept in the context of a Cubesat that is capable of raising its orbit from a geosynchronous orbit (GEO) to interplanetary space. While this concept can be applicable to missions of different interplanetary objects such as Near-earth-objects, a mission from GEO to lunar space is investigated in this work as an example. To fully assess the benefit of the staging concept, comparison to a more traditional mission approach is made. When comparing the staging method to a fixed structural spacecraft for a 3U Cubesat, the reduction of propellant mass and time of flight is 12.7 % for a transfer from GEO to the Moon's Sphere of Influence.