Propellant Thruster Research Articles

Optimization investigation of vacuum air-intake for atmosphere-breathing electric propulsion system

An atmosphere-breathing electric propulsion system uses the rarefied atmospheric molecules as the propellant for the electric thruster. In the best case, it can allow spacecraft complete a long-time mission in the lower Earth orbit without carrying any propellant. In this article, the intake geometry is designed, analysed and optimized to improve the performance of atmospheric particles capture, including collection efficiency and compression ratio. The orthogonal method is used in the simulation test to analyse the sensitivities of main parameters, including the configuration of grid ducts, the configuration of tapered chamber, the length-to-diameter ratio of tapered chamber and the diameter of tube. The results show that the performance of air-intake can be optimized with different parameter combinations. Compared with different intake designs of previous studies, the optimal design in this article shows the better particle capture performance under the same boundary conditions. The particles compression ratio is over 100, and the collection efficiency can reach 81.08%.

Intake design for an Atmosphere-Breathing Electric Propulsion System (ABEP)

Challenging space missions include those at very low altitudes, where the atmosphere is source of aerodynamic drag on the spacecraft. To extend the lifetime of such missions, an efficient propulsion system is required. One solution is Atmosphere-Breathing Electric Propulsion (ABEP) that collects atmospheric particles to be used as propellant for an electric thruster. The system would minimize the requirement of limited propellant availability and can also be applied to any planetary body with atmosphere, enabling new missions at low altitude ranges for longer times. IRS is developing, within the H2020 DISCOVERER project, an intake and a thruster for an ABEP system. The article describes the design and simulation of the intake, optimized to feed the radio frequency (RF) Helicon-based plasma thruster developed at IRS. The article deals in particular with the design of intakes based on diffuse and specular reflecting materials, which are analysed by the PICLas DSMC-PIC tool. Orbital altitudes h=150−250km and the respective species based on the NRLMSISE-00 model (O, N2, O2, He, Ar, H, N) are investigated for several concepts based on fully diffuse and specular scattering, including hybrid designs. The major focus has been on the intake efficiency defined as ηc=Ṅout∕Ṅin, with Ṅin the incoming particle flux, and Ṅout the one collected by the intake. Finally, two concepts are selected and presented providing the best expected performance for the operation with the selected thruster. The first one is based on fully diffuse accommodation yielding to ηc<0.46 and the second one based on fully specular accommodation yielding to ηc<0.94. Finally, also the influence of misalignment with the flow is analysed, highlighting a strong dependence of ηc in the diffuse-based intake while, for the specular-based intake, this is much lower finally leading to a more resilient design while also relaxing requirements of pointing accuracy for the spacecraft.

Applicability of Iodine as a propellant for hall Effect thrusters: Performance parameters prediction

In deep space missions with smaller payloads, electric propulsion is one of the most preferred propulsion techniques. Electric propulsion possesses the advantages of prolonged mission span and a very high specific impulse (Isp). However, the complexity lies in determining the suitable propellant for the electric propulsion-based systems viz., Hall effect thrusters (HET). The present paper reports the approach suitable for 6U CubeSat to utilize Iodine as a potential alternate fuel. The study focuses on computational simulation using the commercial package ANSYS FLUENT 18. A Hall effect thruster simulation is carried out using MHD module in fluent, wherein effect of power supplies were studied by varying the voltage supply of 200 V and 300 V. It is observed that Iodine propellant with the HET configuration delivers Isp of 2650 s and thrust ∼ 37.5 (mN) in the STP at 600 W, and Isp of 4565 s followed by thrust ∼ 27.5 (mN) in STP at 400 W. The simulation is carried out by assuming the plasma fluid model and serves as the base for the earlier stage development of experimental techniques.

Аналіз переваг використання аеродинамічного компенсатора при безконтактному видаленні космічного сміття

A modified scheme of the known technology for contactless space debris removal, which is called Ion Beam Shepherd, is considered. This scheme uses an aerodynamic compensator in order to reduce the propellant consumption of the additional electrojet thruster of the shepherd spacecraft. The thruster serves to compensate the spacecraft motion caused by the action of the main electrojet thruster, whose ion plume “brakes” the space debris object. The aerodynamic compensator significantly increases the spacecraft cross-sectional area compared to the space debris object one. This fact, together with the aerodynamic perturbations acting in the direction perpendicular to the orbital plane, calls for estimating the propellant consumption of the control system thruster to maintain the required position of the spacecraft relative to the space debris object in that direction. The goal of this article is to identify the advantages of using the aerodynamic compensator in space debris removal from low Earth orbits using the Ion Beam Shepherd technology. The tasks of the study are to estimate the reduction in the cost of the momentum of the additional electrojet thruster during contactless space debris object de-orbiting due to the use of the aerodynamic compensator and the additional cost of the momentum of the thruster of the spacecraft – space debris object relative position control system to correct deviations perpendicular to the orbital plane. Using a number of simplifying assumptions, integral estimates of these costs are obtained. Using these cost estimates, it is shown that the use of an aerodynamic compensator is advantageous in terms of the cost of the saved electrojet thruster propellant (xenon) regardless of the type of the spacecraft control system thruster.

Naphthalene as a Cubesat Cold Gas Thruster Propellant

The `cubesat' form factor (multiples of 10x10x10 cm^3 volume and 1.33 kg mass) has been adopted as the defacto standard for a cost effective and modular, nano-satellite platform. Many commercial options exist for nearly all components required to build such satellite; however, there is a limited range of thruster options that suit the power and size restrictions of a cubesat. This work presents the design, implementation and direct thrust measurements of a proof of concept cubesat cold gas thruster system using naphthalene (C10H8) as the propellant. The proposed design is optimised for simplicity to match the requirement of entry level cubesat missions, yet, due to the properties of naphthalene, it can achieve a total impulse in the order of tens of newton-seconds.

RF Helicon-based Inductive Plasma Thruster (IPT) Design for an Atmosphere-Breathing Electric Propulsion system (ABEP)

Challenging space missions include those at very low altitudes, where the atmosphere is source of aerodynamic drag on the spacecraft. To extend such missions lifetime, an efficient propulsion system is required. One solution is Atmosphere-Breathing Electric Propulsion (ABEP). It collects atmospheric particles to be used as propellant for an electric thruster. The system would minimize the requirement of limited propellant availability and can also be applied to any planet with atmosphere, enabling new mission at low altitude ranges for longer times. Challenging is also the presence of reactive chemical species, such as atomic oxygen in Earth orbit. Such species cause erosion of (not only) propulsion system components, i.e. acceleration grids, electrodes, and discharge channels of conventional EP systems. IRS is developing within the DISCOVERER project, an intake and a thruster for an ABEP system. The paper describes the design and implementation of the RF helicon-based inductive plasma thruster (IPT). This paper deals in particular with the design and implementation of a novel antenna called the birdcage antenna, a device well known in magnetic resonance imaging (MRI), and also lately employed for helicon-wave based plasma sources in fusion research. This is aided by the numerical tool XFdtd®. The IPT is based on RF electrodeless operation aided by an externally applied static magnetic field. The IPT is composed by an antenna, a discharge channel, a movable injector, and a solenoid. By changing the operational parameters along with the novel antenna design, the aim is to minimize losses in the RF circuit, and accelerate a quasi-neutral plasma plume. This is also to be aided by the formation of helicon waves within the plasma that are to improve the overall efficiency and achieve higher exhaust velocities. Finally, the designed IPT with a particular focus on the birdcage antenna design procedure is presented.

Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene (PTFE). Impulse bit measurements for PTFE indicate 100 ± 20 µN-s at an initial energy level of 5 J, which increases linearly with energy by approximately 30 µN-s/J. Within the error of the experiment, measurements of the impulse bit for the electric solid propellant are identical to PTFE. Specific impulse when operating on PTFE is calculated to be about 450 s. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which decreases the average specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s.

Near-Infrared-Emitting Probes for Detection of Nanomolar Hydrazine in a Complete Aqueous Medium with Real-Time Application in Bioimaging and Vapor-Phase Hydrazine Detection

Hydrazine is a highly reactive inorganic compound that is commonly used as a thruster propellant in rocket fuels. In this report, the synthesis of a naphthofluorescein-based simple, sensitive, and selective colorimetric and fluorescent probe for the qualitative and quantitative determination of hydrazine is described. An ultrasensitive NIR probe is employed for the nanomolar detection of hydrazine with real-time applicability. The probe is capable of detecting hydrazine in a completely aqueous medium without the interference of other analytes. The probe reacts selectively with hydrazine and exhibits turn-on NIR emission with quantum yields ranging from Φ = 0.0011 to Φ = 0.5338. This enables the probe to selectively detect extremely low concentrations of hydrazine (∼47 nM). As a consequence, practical utility of the probe is greatly augmented. Eco-friendly test strips were prepared to monitor the hydrazine under aqueous and vapor phases. The potential of the probe was verified by monitoring the presence of hydrazine in a living system by fluorescence bioimaging studies of living cells.

A highly stable and efficient spherical underwater robot with hybrid propulsion devices

Underwater robots have been promoted a significant interest in monitoring the marine environment. In some complex situation, robots sometimes need to keep moving fast, sometimes need to keep low speed and low noise. To address this issue, a novel spherical underwater robot (SUR IV) with hybrid propulsion devices including vectored water-jet and propeller thrusters is proposed in this paper. The diversity of the movement modes is also proposed for the different targets as remote or hover and general or silent. To analyze the hydrodynamic characteristics of the hybrid thruster, the computational fluid dynamics simulation is calculated in ANSYS CFX by using the multi-reference frame method. The simulation results show the interaction between the propeller and water-jet thruster. The thrust experiment to evaluate the performance of the improved hybrid thruster is also conducted. The maximum thrust of the hybrid thruster is increased 2.27 times than before. In addition, a noise comparison experiment is conducted to verify the low noise of the water-jet thruster. Finally, the 3 DoF motions which include the surge, heave and yaw for the SUR IV were carried out in the swimming pool. The improvement of the overall robot is assessed by the experimental results.

A micro-force measurement system based on high-temperature superconducting magnetic levitation

A thrust stand based on high-temperature superconducting (HTS) magnetic levitation technology capable of measuring the micro-thrust in two methods is presented in this paper. The stand mainly consists of a superconducting array, a permanent magnetic levitation platform, refrigerating devices, optical measuring devices, calibration devices and a data acquisition system. HTS magnetic levitation technology was used to provide a platform for micro-force measurements due to its low-friction and high-load-bearing characteristics. The measurements are divided into a direct method and an indirect method according to the principles. The direct method is required to measure the angular displacement of the platform and convert it into thrust with proper calibration. The indirect method, on the other hand, is required to measure the angular velocity to calculate the thrust according to the angular momentum theorem. Experiments indicate that, although the hysteresis loss of the HTS levitation system is unavoidable and may affect the accuracy of the system, it can be reduced effectively by multiple calibrations. The calibration results illustrate excellent repeatability and linearity. A propeller thruster is used to verify the performance of the stand. Four sets of micro-force with different propeller speeds were tested. The relative error between results from both methods at each speed was 11.48%, 18.11%, 22.73% and 21.00%. The direct method is more accurate and has fewer sources of uncertainty. It presents a new approach for measuring micro-thrust.

Fabrication and Testing of the Cold Gas Propulsion System Flight Unit for the Adelis-SAMSON Nano-Satellites

Adelis-SAMSON is a nano-satellite mission aimed at performing geo-location of target signals on Earth using a tight three-satellite formation in space. To maintain formation, each nano-satellite is equipped with a cold gas propulsion system. The design, qualification, and integration of the Adelis-SAMSON nano-satellite propulsion system is presented in this paper. The cold gas propulsion system mass is approximately 2 kg, takes a volume of 2U, and generates a thrust of 80 mN from four thrusters using krypton as a propellant. We first present the propulsion system requirements and corresponding system configuration conceived to meet the mission requirements. Subsequently, we present the system architecture while listing all the components. We overview the particular role and qualification process of four of the propulsion system’s components: the propellant tank, thruster assembly, pressure regulators, and fill and vent valve. We detail the tests performed on each component, such as proof pressure tests, vibration tests, and external leak tests. Finally, we present the propulsion system level tests before delivery for satellite integration and discuss the propulsion system’s concept of operations.

Molecular propellants for ion thrusters

There is no ideal atomic propellant for ion thrusters. Xenon commonly used as propellant becomes resource-critical in light of electric propulsion commercialization. Combining these considerations leads to seeking alternatives to xenon as propellant. In this review, we summarize the current literature on molecular propellants. We define two classes of molecules, group I and II, comprising diatomic molecules and more complex molecules, respectively. We identify basic properties which a candidate molecule belonging to either group, I or II, should possess in order to be suitable as molecular propellant. We discuss the pits and traps in testing such candidate molecules inside a thruster on the basis of our experiences with iodine (a member of group I) and adamantane (a member of group II). The thruster system needs to be individually adopted for each propellant candidate in order to enable a thorough testing inside the thruster. The same holds for optimizing the thruster’s performance when fed with a new propellant because the microscopic processes occurring inside the plasma will differ from molecule to molecule. These circumstances make such testing time-consuming and costly. To accelerate systematic screening of the vast number of molecular species in terms of suitability as propellant, we propose a screening and evolution procedure based on combining chemical engineering and fundamental physical measurements.

Experimental studies of iodine stationary plasma thruster

The paper demonstrates the feasibility of using iodine as propellant for thrusters with closed electron drift and its economic viability. It describes a test setup for running experiments. It provides the results of experimental studies of the stationary plasma thruster using iodine as its propellant with xenon gas-passage hollow cathode, as well as of the operational mode of the thruster where a mixture of xenon and iodine is used. During tests gas dynamic and electrical properties of the thruster were analyzed. Thermal conditions in the iodine storage and supply system were studied. Conclusions were drawn on how the test object could be improved and upgraded. The paper describes the option to use a thermionic non-flow cathode as the compensator cathode for the operation of the iodine thruster. The paper provides the results of an experimental study of the prototype non-flow compensator cathode in diode mode. Based on the results of the studies an experimental facility was built for testing a thruster with non-flow compensator cathode. Key words: cathode, compensator cathode, thruster with closed electron drift, stationary plasma thruster, iodine.

Argon ionization improvement in a plasma thruster induced by few percent of xenon

In spite of its high cost, xenon gas is known as both the most efficient and commonly used propellant for plasma thrusters in space technologies. Argon, a gas by far less costly, is widely used in other technologies, but a much lower efficiency of ionization, as obtained for example in closed electron drift thrusters, prevents its use in R&amp;D programs and development of space thrusters. This paper shows that a drastic increase in argon ionization can be obtained in a low power thruster when only a few percent of xenon are added in the argon flow. Besides the increase in the ion beam current in the plume generated by the thruster, a net increase in the ion kinetic energy is observed. These two features are of interest in terms of thrust efficiency. These results, obtained for a small size closed electron drift thruster, could be even more spectacular for higher power devices, suggesting further investigations for space propulsion and/or ion source applications.

Experimental and numerical study on capillary flow along deflectors in plate surface tension tanks in microgravity environment

Surface tension tanks are the most widely used satellite propellant tanks, which use liquid surface tension for liquid transportation and gas-liquid separation to provide pure propellant for thruster. As the second generation surface tension tank, the plate type tank has characteristics of simple structure, easy processing and high reliability, and represents development direction of surface tension tanks. In this paper, drop tower experiments are used to investigate transfer speed of deflectors in plate type tanks, and corresponding numerical simulation is carried out in Fluent with Volume of Fluid method.

On the Possibility of Using Water As a Propellant for Powerful Plasma Thrusters

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Zeeman splitting measurements of magnetic fields in iodine plasma.

Iodine is an attractive propellant for next generation ion thrusters. Laser induced fluorescence (LIF) is widely used with other propellant species as a non-perturbative technique for measuring flow for thruster prediction models. We apply LIF methods recently demonstrated for singly-ionized iodine to a magnetized plasma environment similar to those found in ion thrusters and in magnetically confined laboratory plasmas. We demonstrate the feasibility of remotely determining the local magnetic field from the Zeeman effect-split spectrum of I+.

System analysis and test-bed for an atmosphere-breathing electric propulsion system using an inductive plasma thruster

Challenging space mission scenarios include those in low altitude orbits, where the atmosphere creates significant drag to the S/C and forces their orbit to an early decay. For drag compensation, propulsion systems are needed, requiring propellant to be carried on-board. An atmosphere-breathing electric propulsion system (ABEP) ingests the residual atmosphere particles through an intake and uses them as propellant for an electric thruster. Theoretically applicable to any planet with atmosphere, the system might allow to orbit for unlimited time without carrying propellant. A new range of altitudes for continuous operation would become accessible, enabling new scientific missions while reducing costs. Preliminary studies have shown that the collectible propellant flow for an ion thruster (in LEO) might not be enough, and that electrode erosion due to aggressive gases, such as atomic oxygen, will limit the thruster lifetime. In this paper an inductive plasma thruster (IPT) is considered for the ABEP system. The starting point is a small scale inductively heated plasma generator IPG6-S. These devices are electrodeless and have already shown high electric-to-thermal coupling efficiencies using O2 and CO2. The system analysis is integrated with IPG6-S tests to assess mean mass-specific energies of the plasma plume and estimate exhaust velocities.

Theoretical and Experimental Insights into the Dissociation of 2-Hydroxyethylhydrazinium Nitrate Clusters Formed via Electrospray.

Ionic liquids are used for myriad applications, including as catalysts, solvents, and propellants. Specifically, 2-hydroxyethylhydrazinium nitrate (HEHN) has been developed as a chemical propellant for space applications. The gas-phase behavior of HEHN ions and clusters is important in understanding its potential as an electrospray thruster propellant. Here, the unimolecular dissociation pathways of two clusters are experimentally observed, and theoretical modeling of hydrogen bonding and dissociation pathways is used to help rationalize those observations. The cation/deprotonated cation cluster [HEH2 - H]+, which is observed from electrospray ionization, is calculated to be considerably more stable than the complementary cation/protonated anion adduct, [HEH + HNO3]+, which is not observed experimentally. Upon collisional activation, a larger cluster [(HEHN)2HEH]+ undergoes dissociation via loss of nitric acid at lower collision energies, as predicted theoretically. At higher collision energies, additional primary and secondary loss pathways open, including deprotonated cation loss, ion-pair loss, and double-nitric-acid loss. Taken together, these experimental and theoretical results contribute to a foundational understanding of the dissociation of protic ionic liquid clusters in the gas phase.

Study on Endurance and Performance of Impregnated Ruthenium Catalyst for Thruster System.

Performance and endurance of the Ru catalyst were studied for nitrous oxide monopropellant thruster system. The thermal decomposition of N2O requires a considerably high temperature, which make it difficult to be utilized as a thruster propellant, while the propellant decomposition temperature can be reduced by using the catalyst through the decomposition reaction with the propellant. However, the catalyst used for the thruster was frequently exposed to high temperature and high-pressure environment. Therefore, the state change of the catalyst according to the thruster operation was analyzed. Characterization of catalyst used in the operation condition of the thruster was performed using FE-SEM and EDS. As a result, performance degradation was occurred due to the volatilization of Ru catalyst and reduction of the specific surface area according to the phase change of Al2O3.