Thruster Performance Research Articles

Argon admixture-driven enhanced ionization and performance of a 5 kW Hall thruster on krypton

Utilizing alternative propellants has been recognized as a strategy to reduce the total cost of propellants in electric propulsion-based missions. The aim of this study is to quantify the Hall effect thruster (HET) performance and operating characteristics using a Kr–Ar mixture to enable mission designers to evaluate the impact on mission and spacecraft design. We present the performance and plume plasma properties of the P5 5 kW-class HET operated with a Kr–Ar mixture with Ar volumetric flow rate fractions from 0 to 100%. The thruster is characterized at discharge power levels of 2.6 kW and 4.1 kW at constant discharge current and voltage over the range of Ar fractions. Despite higher ionization energy and lower mass of Ar, the thruster exhibited a similar level of thrust within 2% when comparing the pure Kr and 26%-Ar mixture cases. The derived ion energy distribution functions and analytical modeling suggest that the characteristic length for the ionization region is extended as the Ar fraction increases. The increased residence time of Kr at the extended ionization region and background energetic electrons from the ionized Ar neutrals are considered to cause this enhanced ionization of the injected Kr neutrals. This leads to a 6% higher Kr ion density at the 26%-Ar case even at the 16% less injected Kr neutral density than in the pure Kr case. The enhanced Kr ionization and generated Ar ions in the 26%-Ar case consequently led to a comparable thrust with that of the pure Kr case. The study indicates that mixing Ar with approximately 26% volumetrically with Kr can provide a similar or even higher thrust performance at the same discharge power. This will be particularly advantageous for various space missions that require high impulses by reducing the total cost of the propellant.

Magnetron sputtering thruster operated with a compact gas feeding system using double pulsed valves

A continuous injection of gaseous water propellant into a magnetron-type sputtering propulsion device is performed in a pseudo manner by using a combination of two high-speed solenoid-type pulsed valves and a low-pressure gas reservoir . The gas injection system provides a temporally averaged gas flow rate of about 5–10 sccm with adjustment of the repetition frequencies of the pulsed valves, while the previously used single valve system attached to a liquid water tank has caused excessive gas injection. A discharge is sustained by a high voltage dc power supply; the discharge current has pulsation synchronized with the pulsed valve repetition. A temporally averaged force induced by the magnetron sputtering thruster and exerted on a pendulum plate downstream of the thruster is measured; the results are in good agreement with the previously reported force measurement under precisely controlled gas flow rate using a commercial mass flow controller. Therefore, the precise control of the gaseous water propellant with the compact gas feeding system is achieved, while maintaining a similar thruster performance to that with the mass flow controller.

Radial ablation uniformity of cathode and design of double anode micro-cathode arc thruster

Micro-cathode arc thruster offers promising applications in micro–nanosatellite space missions owing to its simple structure, high specific impulse, and low power consumption. However, the non-uniform ablation of the cathode poses a significant limitation on the total impulse of micro-cathode arc thruster. This non-uniformity can be categorized into circumferential and radial types, with previous research primarily focusing on the influence of circumferential ablation non-uniformity. Nevertheless, it has been observed that radial ablation non-uniformity also affects the stability of the thruster, despite the lack of studies in this area. To investigate the mechanism behind the cathode radial ablation, a slice model has been designed to study the motion process of the cathode spots during a single discharge using a high-speed camera. From this analysis aims to identify the root cause of uneven cathode radial ablation. Building upon these findings, a double anode micro-cathode arc thruster has been devised to enhance the uniformity of cathode ablation. Experimental verification has been conducted to assess the performance of the double anode micro-cathode arc thruster. The results indicate that in the traditional micro-cathode arc thruster discharge process, the cathode spots do not migrate towards the outer layer covered by the shell. Instead, they extend to the surface of the shell along the contact area between the shell and cathode. This leads to unequal ablation between the inner and outer layers of the cathode. The double anode micro-cathode arc thruster effectively addresses these issues by providing two discharge channels that maintain the dynamic balance of electrode resistance and increase the ablation amount on the outer layer of the cathode. Consequently, the total impulse of the thruster improved significantly while ensuring the desired performance index.

Performance investigation of a low-power Hall thruster fed on iodine propellant

The common propellants used for electric thrusters, such as xenon and krypton, are rare, expensive, and difficult to acquire. Solid iodine attracts much attention with the advantages of low cost, extensive availability, low vapor pressure, and ionization potential. The performance of a low-power iodine-fed Hall thruster matched with a xenon-fed cathode is investigated across a broad range of operation conditions. Regulation of the iodine vapor’s mass flow rates is stably achieved by using a temperature control method of the iodine reservoir. The thrust measurements are finished utilizing a thrust target during the tests. Results show that thrust and anode-specific impulse increase approximately linearly with the increasing iodine mass flow rate. At the nominal power of 200 W class, iodine mass flow rates are 0.62 and 0.93 mg/s, thrusts are 7.19 and 7.58 mN, anode specific impulses are 1184 and 826 s, anode efficiencies are 20.8% and 14.5%, and thrust to power ratios are 35.9 and 37.9 mN/kW under the conditions of 250 V, 0.8 A and 200 V, 1.0 A, respectively. The operating characteristics of iodine-fed Hall thruster are analyzed in different states. Further work on the measurements of plasma characteristics and experimental optimization will be carried out.

Enhancement of the electrostatic–magnetic hybrid thrust performance using a stagnant ring

Electrostatic–magnetic hybrid ion acceleration has been researched to realize high-thrust density and efficient thrust operations. In this study, a newly designed stagnant ring (SR) was installed in an electrostatic–magnetic hybrid thruster with the nominal anode inner radius of 40 mm to improve the thrust performance, and the effects of SR on the thruster operation of xenon propellant were analyzed. During constant-discharge-voltage operations, the discharge current increased with both electrically conductive (copper) and non-conductive (ceramic) SR materials, yet the thrust increased only with the copper SR, which was equipotential to the anode. At lower values of the radial width (H) of the copper SR, the thrust and discharge current were almost dominated by the propellant flow rate from the anode surface. However, the contribution of the propellant flow rate from the cathode tip on the thrust and discharge current became significant when H ≥ 9 mm. The thrust coefficient in electromagnetic acceleration form exhibited a maximum value at H = 9 mm, whereas that in electrostatic acceleration form monotonically increased when H ranged from 0 to 20 mm. The highest thrust efficiency was obtained at H = 12 mm. The study findings indicate that installing a conductive SR can effectively enhance the thrust performance in electrostatic–magnetic hybrid thrusters.

Numerical study on altitude-compensating mechanism of a permeable nozzle

A conceptual adaptive altitude compensation nozzle with a permeable wall (permeable nozzle for short) is proposed in the present study. The typical operating states of the nozzle are demonstrated using numerical methods and the underlying mechanism of its altitude compensation is interpreted. The numerical results show that, at low nozzle pressure ratios, the external atmosphere enters the nozzle through its permeable wall section, which inhibits the backflow at the nozzle wall, weakens the strength of the separated oblique shock, thus improves the low-altitude thrust performance of the nozzle. However, at high nozzle pressure ratios, the exhaust gas leaves the nozzle through its permeable wall, which weakens the expansion of the exhaust gas and reduces the thrust performance. The permeable nozzle is considered to possess altitude compensation ability when the enhancement in its low-altitude thrust performance surpasses the deterioration in high-altitude thrust performance. The thrust composition analysis of the permeable nozzle reveals that its altitude compensation ability derives from the internal pressure rise due to the atmospheric air penetration into the nozzle at low nozzle pressure ratios.

Aerodynamics of a flapping wing with stroke deviation in forward flight

In this paper, we numerically studied the effect of stroke deviation on the aerodynamic performance of the three-dimensional flapping wing in forward flight at a low Reynolds number. Six deviation motion patterns with different stroke deviation amplitudes were investigated. The results show that the distinct patterns exert a substantial influence on the aerodynamic forces of the flapping wing, with a more pronounced effect at higher values of deviation amplitude. For most patterns, stroke deviation enhances either lift or thrust performance unilaterally. The maximum lift and thrust of the wing with deviation motion can be 37% and 35% larger than that of the wing without deviation motion. A detailed analysis of typical flow characteristics underscores the pivotal role of deviation motion in aerodynamic force generation. Finally, two artificially created innovative deviation motion patterns are proposed, which exhibit an exceptional capacity to augment thrust by up to 123% or enhance comprehensive aerodynamic performance significantly. These findings establish a theoretical foundation for designing high-performance flapping-wing micro-air vehicles.

Nitrous Oxide/Ethanol Cylindrical Rotating Detonation Engine for Sounding Rocket Space Flight

There are few experimental studies on rotating detonation engines (RDEs) with liquid propellants. This study reveals the static thrust performance of a cylindrical RDE with ethanol and liquid nitrous oxide as propellants under atmospheric pressure. This RDE had an inner diameter of 40 mm, a maximum combustor length of 230 mm, a nozzle contraction ratio of 1.7, and a nozzle expansion ratio of 9.1. Nineteen experiments were conducted at total mass flow rates of 184–210 g/s, mixture ratios of 3.6–5.9, and combustion pressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of 1840 m/s (approximately 80% of the theoretical detonation velocity, 2251 m/s), maximum thrust at sea level of 294 N, and maximum specific impulse at sea level of 148 s. In addition, the maximum characteristic exhaust velocity, c*, was 1716 m/s, which was 99% of the theoretical value. The characteristic length of the combustion chamber at this time was 0.15 m. Since conventional rocket combustion requires 1.57 m to achieve the same c* efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.

Uji Performa Waterjet Thruster Produk 3D Printing Material Super Tought Polylactic Acid Terhadap Gaya Dorong

In the world of transportation, there are many things that must be considered, ranging from safety to efficiency. Propulsion system planning is very important, this is because the system is a thrust generator that will determine the performance of the ship. The thrust generated must be able to overcome the resistance or friction of the ship with water. In recent times, many studies have been conducted to improve the performance and efficiency of the waterjet thruster. In this research, the performance of the waterjet thruster will be tested with variations in the number of inlet blades and the number of turbo outlet blades. The levels used are 8 blades and 15 blades. The manufacturing process on waterjet thruster components such as inlet, impeller, and turbo outlet uses Fused Deposition Modeling (FDM) technology to obtain high dimensional accuracy. From the research conducted, it can be concluded that the number of blades at the turbo inlet and outlet has a major effect on the thrust generated. The turbo outlet has a greater role than the inlet in influencing the thrust force. From the research conducted by Nori and this study, it can be concluded that the turbo outlet with 8 blades produces the highest thrust force compared to 6 blades and 15 blades.

Research on the Support Performance of Internal Feedback Hydrostatic Thrust and Journal Bearing Considering Load Effect

This study aims to analyze the impact of uniform and eccentric load conditions on the performance of internal feedback hydrostatic thrust and journal bearing. Two distinct models are established: a three-degrees-of-freedom uniform load model and a five-degrees-of-freedom eccentric load model. The support stiffness, overturning stiffness, and flow rate for both thrust and journal bearings are calculated. Additionally, numerical analysis is conducted to examine the influence of oil film thickness, inlet pressure, and restrictor size on the operational characteristics of the bearings, revealing the interplay between an eccentric load and journal bearing speed. The validity of the theoretical algorithm is verified through finite element simulation. The research outcomes hold significant guiding implications for the design and application of internal feedback hydrostatic bearings.

Influence of acceleration stage electrode voltage on the performance of double-stage Hall effect thruster with adjustable zero magnetic point

The configuration of electrode voltage and zero magnetic point position has a significant impact on the performance of the double-stage Hall effect thruster. A 2D-3V model is established based on the two-magnetic peak type double-stage Hall thruster configuration, and a particle-in-cell simulation is carried out to investigate the influences of both acceleration electrode voltage value and zero magnetic point position on the thruster discharge characteristics and performances. The results indicate that increasing the acceleration voltage leads to a larger potential drop in the acceleration stage, allowing ions to gain higher energy, while electrons are easily absorbed by the intermediate electrode, resulting in a decrease in the anode current and ionization rate. When the acceleration voltage reaches 500 V, the thrust and efficiency are maximized, resulting in a 15% increase in efficiency. After the acceleration voltage exceeds 500 V, a potential barrier forms within the channel, leading to a decrease in thruster efficiency. Further study shows that as the second zero magnetic point moves towards the outlet of the channel, more electrons easily traverse the zero magnetic field region, participating in the ionization. The increase in the ionization rate leads to a gradual enhancement in both thrust and efficiency.

Numerical investigation on the effects of heel on the aerodynamic performance of wing sails

The utilization of wind-assisted propulsion technology could reduce ship fuel consumption and emission of pollutants. In the study, the effect of steady heel angles on the aerodynamic characteristics of wing sails was investigated. Full-scale three-dimensional simulations with the ship's superstructure were conducted at heel angles of 0°–15° in all apparent wind directions. The correctness of the numerical computations is verified by comparing them with the wind tunnel experimental data. Overall, the heel is detrimental to the capture of wind energy by the wing sails. The aerodynamic forces of the sail system decrease with increasing heel angles. Specifically, the average thrust performance is attenuated by 4.7%, 13.5% and 20.5%, respectively, at heel angles of 5°, 10°, and 15° compared to the upright condition. Due to changes in the airflow distribution around the sails, the aerodynamic force of sails in heeled to windward and heeled to leeward decreases by the reduced area of the low-pressure zone on the suction side and the high-pressure zone on the pressure side, respectively. Finally, a prediction scheme for assessing the effect of rolling motion on sail performance is proposed by steady heel results.

Research on tribological performance of textured thrust bearing using gray relational degree and improved multi-objective water circulation algorithm

To enhance the loading capacity of the oil film, reduce the lubricating oil temperature, and minimize the deformation of the friction pairs, three types of concave textures were proposed. The tribological properties of the thrust bearing underwent multi-objective optimization analysis using the two different methods. By comparing the two methods, it was found that the improved algorithm (MOWCA) demonstrated better optimization performance. Under this algorithm, when the texture shape was triangular, the number of textures was 18n, the size was approximately 0.135 mm, and the spacing between textures was approximately 4.6 mm, the loading capacity of the oil film increased by approximately 26.8%, and the temperature decreased by approximately 0.46%.

Numerical analysis of the effect of the duct geometry on the hydrodynamic performance of rim-driven thruster

This paper studies the effect of the duct geometry on the hydrodynamic performance of rim-driven thruster (RDT) based on Computational Fluid Dynamics (CFD) method. The effect of the thickness of the duct, the radius of the leading edge of the duct and the geometry of the trailing edge of the duct on the thrust coefficient, torque coefficient and efficiency are investigated and analyzed. The conclusion shows the thrust coefficient increases with the thickness of the duct decreasing, however the torque coefficient increases with the increase of the thickness of the duct, so the efficiency increasing with the decrease of the thickness of the duct. On the other hand, the radius of the leading edge of the duct has no significant effect on the thrust coefficient of the RDT. However, the torque coefficient of RDT decreases with the increase of the radius of the leading edge of the duct, so the efficiency of RDT increases with the increase of the radius of the leading edge of the duct. A comprehensive comparison shows that the RDT 15-0.5 (the duct thickness is 15 mm, the ratio of the radius of the leading edge of the duct and the duct thickness is 0.5) has the best hydrodynamic performance among the investigated RDT models in this work.

Research on the effect of shallow draft on the performance of azimuth thruster under bollard conditions

For the azimuth thruster installed on shallow draft ships, the shallow water effect leads to a decrease in propulsive performance, especially under the bollard condition. In order to accurately predict the bollard propulsion performance of the azimuth thruster under shallow draft conditions, this paper firstly obtains the bollard thrust performance in several waters through the shallow water test, and then analyzes the effect of hull wake and cavitation on the bollard thrust performance of the azimuth thruster by using the CFD method. The results are consistent with the test results under the open water. When the water depth is less than 1.0D, the thrust coefficient increases with the increase of water depth, when the water depth is greater than 1.0D, the thrust coefficient basically does not change; Considering the influence of hull wake, the axial inlet velocity of the azimuth thruster is 7.27% higher than that of the open water condition; meanwhile considering the influence of the hull wake and cavitation, the leading edge of the bottom of the azimuth thruster conduit and the back of the paddle blade are susceptible to cavitation, which leads to average decrease of the thrust coefficient and torque coefficient by 7.03% and 6.12%, respectively.

Performance of pulsed plasma thruster at low discharge energy

As the size of satellites scales down, low-power and compact propulsion systems such as the pulsed plasma thruster (PPT) are needed for stabilizing these miniature satellites in orbit. Most PPT systems are operated at 2 J or more of discharge energy. In this work, the performance of a PPT with a side-fed, tongue-flared electrode configuration operated within a lower discharge energy range of 0.5‒2.5 J has been investigated. Ablation and charring of the polytetrafluoroethylene propellant surface were analyzed through field-effect scanning electron microscopy imaging and energy-dispersive X-ray spectroscopy. When the discharge energy fell below 2 J, inconsistencies occurred in the specific impulse and the thrust efficiency due to the measurement of the low mass bit. At energy 2 J, the performance parameters are compared with other PPT systems of similar configuration and discussed in depth.

On the study of the pitch angular offset effects at various flapping frequencies for a two-dimensional asymmetric flapping airfoil in forward flight

This article explores the aerodynamic performance of a two-dimensional elliptical airfoil undergoing sinusoidal heaving and asymmetric pitching motions in forward-flight conditions for the Strouhal number (St) range of 0.1–0.6. The study employs numerical simulations and water tunnel experiments to investigate the effects of non-zero pitch angular offset angles (θoffset) while maintaining a fixed Reynolds number of 5000 and an effective angle of attack amplitude of 15° at the pivot location. The θoffset is varied from −15° to +15° at 5° intervals. The present research shows that these parameters significantly impact the leading-edge effective angle of attack, flow velocity, and the formation of high-pressure regions, which are crucial factors in thrust and lift generation throughout the flapping cycle. Moreover, the pitch angle determines whether the resultant force favors thrust or lift. It is observed that the cyclic time-averaged lift consistently increases with θoffset, surpassing symmetric cases (θoffset = 0°). Conversely, the cyclic time-averaged thrust is lower for non-zero θoffset values. Increasing St enhances both cyclic time-averaged thrust and lift up to the respective critical Sts, after which their performance declines. Notably, the critical St of cyclic time-averaged lift exceeds that of cyclic time-averaged thrust; interestingly, their values are invariant with θoffset. Moreover, in the conditions where thrust efficiency maximizes, lift efficiency attains a minimum value and vice versa. So, depending upon the application, one needs to suitably select the pitch angular offset and flapping frequency to maximize thrust or lift performance.

Experimental demonstration of a water electrolysis Hall Effect Thruster (WET-HET) operating with a hydrogen cathode

This study presents the first experimental demonstration of a Hall Effect Thruster operating with water electrolysis-derived constituents: oxygen for the anode (thruster) and hydrogen for the cathode, both from pressurised bottles. In contrast to prior research where the thruster was operated with oxygen to the anode and xenon/krypton to the cathode, this investigation uses an advanced set-up that safely allowed the combined use of pressurised oxygen and hydrogen. Firstly, a comparative assessment between the established plasma filament cathode (utilised in previous research) and a novel LaB6 cathode (referred to as Hydrocat), both krypton-fuelled, is conducted. Results show a different IV curve for these cathodes, but negligible disparities in thruster performance for fixed discharge power and anode mass flow. Thrust, anode specific impulse, thrust-to-power ratio, and both, anode and total efficiency of the system using oxygen (anode) and hydrogen (cathode) are then presented and compared against oxygen (anode) and krypton (cathode). The cathode successfully operated on hydrogen, with a reduction in thrust performance of under 10% compared to when krypton is used as the cathode propellant. This performance difference can be attributed to different factors, such as the reduced efficiency of the cathode running with hydrogen due to a higher ionisation potential and the cathode propellant’s contribution to thrust, more significant when using heavier propellants such as xenon/krypton compared to hydrogen. The impact of changing the hydrogen-to-oxygen ratio is also assessed and shown to have no impact on the thruster’s performance. The most optimal performance was obtained at the highest discharge power tested (Pd=1953±1 W), using a hydrogen:oxygen ratio of Φ = 1:25. This operating point corresponded to a thrust of 20.9±0.2 mN, anode specific impulse of 2130±20 s, thrust-to-power ratio of 10.7±0.5 mN/kW, anode thrust efficiency of 11.2±0.3 %, and total efficiency of 10.1±0.3 %.

Ground Effect on the Thrust Performance of Staggered Rotor System

In this study, the thrust performance of a staggered rotor system in-ground effect (IGE) and out-of-ground effect (OGE) while considering the interaction on wake characteristics were investigated experimentally. A thorough comprehension of their performance holds significant importance for trajectory planning, aircraft design, landing safety, and energy-efficient landings. The complex interactions within staggered rotor systems and the impact of ground effects make rotor distance and ground interactions critical factors influencing near-ground flight performance. The study investigated the influence and enhancements of rotor thrust performance through an examination of rotor speed, lateral distance, and altitude. Experimental tests were conducted on two small-scale rotor models to assess the effects of these parameters. These experiments compared the performance of staggered rotor systems with isolated rotors, analyzing the competition mechanism between the thrust loss caused by interference and the thrust gain of rotors IGE. Furthermore, emphasis was placed on analyzing the thrust gain issues exhibited by staggered rotor systems under the condition of H = 2R. Additionally, the analysis was focused on identifying prominent relative positions for thrust performance and parameters for improving thrust performance in ground effects in staggered rotor systems.

Temperature field analysis and cooling structure optimization of permanent magnet linear synchronous motor

A water-cooled permanent magnet linear synchronous motor was designed, and the investigation focused on the temperature distribution of motor and the optimization of the cooling structure to address its electromagnetic thermal effects. Firstly, the electromagnetic model of the motor was employed to calculate losses using finite element analysis. Subsequently, the motor's temperature distribution under various operating conditions was determined through electromagnetic-thermal coupling analysis, with prototype measurements validating the predicted outcomes. Secondly, the study placed emphasis on the impact of different water-cooled structure parameters on the motor's cooling performance, resulting in a 69 % reduction in the maximum temperature of the motor module after the cooling system was applied. Building on this, cooling structure parameters underwent optimization and design via the response surface optimization method, which achieved a reduction of 0.617 K in the motor's maximum temperature. The analysis findings revealed a substantial improvement in the motor's thrust performance by a factor of 3.08 when the water flow rate reached 1 m/s. This research can serve as a fundamental reference for the design of water cooling structures for permanent magnet linear synchronous motors.