Thruster Performance Research Articles

Fabrication of controllable porous tungsten tips for indium FEEP by dynamic reciprocating electrochemical etching combined with ultrasonic cleaning method

The porous tungsten emitter is a core component of field emission electric propulsion (FEEP). It combines the advantages of sharp tip and internal propellant transport. The emission performance of a FEEP thruster is highly depended on the emitter's tip apex radius. Traditional processing methods for solid tungsten tips are unsuitable for porous tungsten due to its unique structure and electrochemical properties. This paper introduces a novel approach combining dynamic reciprocating electrochemical etching with ultrasonic cleaning to fabricate cone-shaped porous tungsten tips with controlled apex radii. By alternating between 5 cycles of etching and an ultrasonic cleaning, a 1.9 μm apex radius is achieved after 90 repetitions. The etching current waveform reveals the material dissolution mechanism. The electrochemical characteristics of porous tungsten are investigated, with an optimal processing voltage of 2 V identified. To refine the apex radius and prevent tip collapse, a 1.9 V etching voltage is used in the final stages, maintaining the apex within 2 μm. The effectiveness of the fabricated tip is confirmed through wetting in indium and ignition tests.

Experimental Study and Application Prospects of Coal Dust/Methane/Oxygen Pulse Detonation Performance: Coal Dust Equivalence Ratio and Types.

In order to achieve reliable application of coal in pulse detonation combustion chambers, this study explored the influence mechanism of equivalence ratio and coal type (lignite, anthracite) on the detonation characteristics of coal/methane/oxygen methane-solid two-phase fuel. A high-energy ignition device was used to ignite the preblast methane fuel, and the high-temperature and high-pressure energy generated directly initiated the detonation of coal/methane/oxygen methane-solid two-phase fuel. The experimental results are as follows: (1) The maximum detonation pressure of coal/methane/air fuel is positively correlated with the initial pressure. With the increase of the φglobal, the maximum detonation pressure of the two types of coal gradually increases. To be specific, when the φglobal rises from 0.5 to 1.6, the sensitivity of lignite and anthracite to the maximum detonation pressure with respect to the initial pressure increases from 4.56 to 6.11 and 5.48, respectively; (2) In terms of detonation flame and maximum pressure wave velocity, anthracite shows better dynamic performance than lignite. The maximum velocity of detonation flame and pressure wave is achieved when the φglobal is 1.4. (3) In terms of detonation thrust performance, as the φglobal increases from 0.5 to 1.6, the rate of increase in average thrust of the two types of coal first increases and then decreases. Compared with lignite, anthracite has a wider time domain and higher peak value in the average thrust curve and the generated detonation wave displays a greater impulse. The experimental data and conclusions obtained in this study are of practical significance for realizing the engineering application of pulverized coal detonator power generation.

Numerical Study on Auxiliary Propulsion Performance of Foldable Three-Element Wingsail Utilizing Wind Energy

Sail-assisted propulsion is an important energy-saving technology in the shipping industry, and the development of foldable wingsails has recently become a hot topic. This type of sail is usually composed of multiple elements, and its performance at different folding configurations is very sensitive to changes in incoming airflow, which result in practical operational challenges. Therefore, original and optimized three-element wingsails (bare and concave) are modeled and simulated using the unsteady RANS method with the k-ω SST turbulence model. Next, certain key design and structural parameters (such as angle of attack, apparent wind angle, and camber) are employed to characterize the auxiliary propulsion performance, and the differences are explained in combination with the flow field details. The results show that, in the unfolded state, the aerodynamic performance of the concave wingsail is better than that of the bare wingsail, exhibiting higher lift coefficients, lower drag coefficients, and a more stable surface flow. In the fully folded state, wherein both the nose and flap are rotated, the thrust performance of the concave wingsail remains superior. Specifically, at an angle of attack of 8 degrees, the thrust coefficient of the concave wingsail is approximately 23.5% higher than that of the bare wingsail, indicating improved wind energy utilization. The research results are of great significance for engineering applications and subsequent optimization design.

Kinetic Simulation of Nozzle Flow in a Micronewton-Class Cold Gas Thruster

Many scientific space missions need highly precise attitude and orbit control or ultrafine drag compensation, which relies on the variable-thrust propulsion technology operating at the micronewton level. Cold gas thruster (CGT) is a very promising solution, mainly because of its high reliability. One of the keys to the success of micronewton variable-thrust CGT is to understand the flow in its nozzle, whose configuration is much more complex than traditional CGT nozzles. This paper applies kinetic-based multiscale models to investigate the gas flow in the complex nozzle of a micronewton variable-thrust CGT. The simulations reveal that the flow simultaneously experiences all kinds of regimes from continuum to free molecular. The continuum breakdown is likely to occur near the throat region due to large gradients of flow variables and in the expander due to low gas density. Frictional choking is observed, and the nozzle length can be optimized to improve the thruster performance. Nozzle performance measures such as thrust, discharge coefficient, and thrust efficiency are found to change only with the throat Knudsen number Knt. The performance curves can be divided into two sections at Knt≃0.1, and thereby an empirical piecewise formula for thrust prediction is proposed.

Study on the spray characteristics of transverse jets with elliptical nozzles in supersonic crossflows using the volume of fluid–discrete phase model

The atomization characteristics of liquid jets injected transversely into a supersonic crossflow significantly affect the performance of scramjet engines. Existing research on this topic has mainly focused on circular nozzles, while the influence of nozzle geometry, particularly elliptical nozzles, has received relatively limited attention. Therefore, this study employs a numerical simulation method coupling the volume of fluid and discrete particle model to investigate the breakup and atomization characteristics of transverse liquid jets from elliptical nozzles with different aspect ratios under supersonic crossflow conditions, as well as the total pressure loss. The simulation model is validated against experimental data obtained from a pulse wind tunnel, including measurements of the liquid jet penetration depth and the Sauter mean diameter (SMD). The results indicate that for elliptical nozzles with an aspect ratio (AR) greater than 1, columnar breakup occurs earlier, and the columnar breakup length is shorter compared to circular nozzles. As the AR increases, the jet penetration depth decreases, while the spray expansion angle increases. Furthermore, the SMD of the atomized spray field from the circular nozzle is larger than that from the elliptical nozzles, and the SMD of the spray field is smallest for an elliptical nozzle with AR of 4. Finally, the elliptical nozzles exhibit a higher total pressure recovery coefficient, indicating reduced total pressure loss in the combustion chamber. This reduction in pressure loss is expected to improve the thrust performance of the scramjet engine.

Effects of differently shaped textures on the tribological properties of static and dynamic pressure thrust bearings and multiobjective optimization

PurposeThe purpose of this study is to evaluate three types of textures designed to enhance the tribological performance of static and dynamic pressure thrust bearings.Design/methodology/approachTo explore the effects of different types of textures on tribological performance, the Reynolds equation is modified using lubrication theory and computational fluid dynamics methods while considering the influence of cavitation and turbulence on the physical field. In addition, the tribological performance is optimized through an improved selection algorithm based on Pareto envelope (PESA).FindingsThe results indicate that textured thrust bearings exhibit superior tribological performance compared to untextured ones. The circular texture outperforms other textures in terms of load-bearing and friction performance, with improvements of approximately 28.8% and 18.9%, respectively. In addition, the triangular texture exhibits the most significant temperature improvement, with a reduction of approximately 1.93%.Originality/valueThe study proposes three types of textures and evaluates the friction performance of thrust bearings by modifying the Reynolds equation. In addition, the optimal texture design is determined using an improved selection algorithm based on PESA.

Propulsive performance of a heaving and pitching foil with large amplitudes in unsteady ground effect

To reveal the effect of ground on the force behavior, propulsion performance, and vortex dynamics of a flapping foil, we numerically study a flapping foil actuated by combined heaving and pitching motions with large amplitudes near a solid ground surface. The instantaneous forces in both streamwise and cross-stream directions undergo a sudden drop during upstroke in close ground effect. The time-averaged thrust is enhanced by the ground but varies non-monotonically with ground distance at high foil oscillating frequencies. Three force regimes are identified in lift production, giving rise to two equilibrium altitudes of zero lift, one stable and another unstable. Moreover, we propose a simple scaling law model by a path-length-based Strouhal number and ground effect for thrust and power performance of a large-amplitude flapping foil. The results of the model are consistent with numerical results and previous experimental studies. The vortex dynamics in the wake structure are largely affected by ground proximity; in particular, the formed vortex pairs and their orientations are sensitive to their initial relative positions. Different sources and mechanisms of satellite vortex formation are then discussed. The ground exerts little effect on the phase angles that optimize the propulsive efficiency. This study is expected to provide new insights into the development and maneuvering of flapping foil-based propulsive systems.

Time- and angular-distribution of ion current in pulse-laser ablation plume of aluminum

Thrust generation through pulse-laser ablation has been proposed for use in propulsion systems in space. However, to achieve a high thrust performance, the energy efficiency must be improved and it is important to determine the behavior of the ablation plume to identify where energy losses occur. In this study, the ion-current distribution of a plume with aluminum ablation was measured. A neodymium-doped yttrium aluminum garnet (Nd:YAG) laser with a wavelength and pulse width of 1064 nm and 9 ± 2 ns, respectively, was used for the ablation. Experiments were conducted at three different fluences: 10, 15, and 20 J/cm2. The fluence was varied in two ways: changing the beam-spot diameter while maintaining the single-pulse energy and changing the single-pulse energy while maintaining the beam-spot diameter. The ion current was measured at various angles using Faraday probes. The generated impulse and reduction mass were also measured. By combining and analyzing the results, the divergence angle of momentum, mean valence of ablated aluminum, and energy efficiencies of several processes were estimated.

Deposition characteristics and influence regularity of sputtered products in a low-power RF gridded ion thruster

RF ion thruster has a broad prospect of development via its simple structure, high specific impulse and electrode-less discharge. During the RF operating process, an important but ignored generally issue is that the high-energy metal particles generated by sputtering grid system can diffuse towards the ionization chamber and then deposit on the wall to form a conductive film. As a result, the RF power feed can be significantly inhibited which affects the thruster perform and reliability. In this work, the axial distribution of film thickness on the ionization chamber has been measured which has the characteristic of nonlinear distribution. And the film roughness can be significantly affected due to the gradient of film thickness and deposition temperature. Another corresponding measurement of plasma parameter evolutions in the ionization chamber has been carried out within 6000 min of thruster operation. Results shows that the presence of the deposited film can result in a ∼30 % decrease in the plasma density, which weakens the grid sputtering erosion and subsequently leads to a decrease in film growth rate. The combination of contributed to the degradation of the thruster performance, with the effect of deposited film accounting for about 41 %.

PIV And Schlieren Measurements Of A Linear Plug Nozzle Jet Interacting With External Flow

Plug nozzles represent a promising concept as part of a propulsion system for space launchers. The ability to adapt the nozzle jet to the ambient pressure level improves thrust performance in overexpanded operating conditions compared to conventional bell nozzles. In this study, the dynamics of a jet flow of a cold air plug nozzle model are investigated by means of PIV and high-speed schlieren measurements. For a fixed nozzle pressure ratio, the jet flow is studied in an externalouter flow environment with sub-, trans-, and supersonic Mach numbers. The effect of plug truncation is analyzed as well. It is found that the outer Mach number strongly influences the flow pattern, local velocity magnitudes and aerodynamic modes. The frequency and dominance of vortex shedding and jet screeching modes are found to be highly dependent on the Mach number of the surrounding flow. A truncated plug introduces local accelerations in its base wake region, causing an increased shock strength in the jet structure. In addition,it causes severe periodic fluctuations in the flow that are transmitted to the shear layer and induce acoustic wave emissions. The impact of the plug length on the flow dynamics has been seen to decrease with increasing outer Mach number.

Effect of Spherical Silver Particles Size of the Catalyst Bed on Hydrogen Peroxide Monopropellant Thruster Performance

Effect of Spherical Silver Particles Size of the Catalyst Bed on Hydrogen Peroxide Monopropellant Thruster Performance

Monodisperse Porous Emitter Materials for Ion Electrospray Propulsion

Electrospray spacecraft propulsion typically uses commercially available porous glass materials for their emitter components. The quality of this emitter material is crucial to the operation of the thrusters, as nonuniformities in the pore distribution can lead to differential propellant flow rates, variation in tip geometries, and less-than-optimal performance of the thruster. The focus of this work is on the development of new porous materials that are more uniform and can mitigate these issues caused by nonuniformities present in current commercially available porous materials. A silica nanoparticle photo-curable resin solution is developed to create high-uniformity porous substrates to improve tip geometry and overall thruster performance.

Characterization of water propellant in an electron cyclotron resonance thruster

A coaxial electron cyclotron resonance thruster operating on water vapor propellant is investigated to determine the influence of molecular propellant chemistry on thruster performance. The performance is characterized at different mass flow rates (0.1–0.4 mg/s) and powers (20–200 W) using a thrust stand, ion energy analyzer, and spectrometer. Experimental data are compared to the results from a theoretical model that includes non-equilibrium chemical kinetics. The thruster is observed to transition from a state where power deposition favors propellant dissociation and ionization into a state where it favors electron heating and ion acceleration. The results suggest that the majority of plasma heating occurs in a core region that surrounds the thruster antenna and that propellant flow outside this region is not efficiently utilized for propulsion.

Armature Reaction Analysis and Suppression of Voice Coil Actuator Based on an Improved Magnetic Equivalent Circuit Model

The high thrust density voice coil actuator (VCA) is a key component for the deployers to determine the CubeSats' separation velocity in orbit directly. To improve the VCA's thrust density and power consumption, it is essential to analyze the armature reaction under heavy current work conditions. This paper proposes a dynamic splitting method of the magnetic equivalent circuit (MEC) model for analyzing the armature reaction of the VCA. This MEC model can effectively reveal the interaction mechanism between the dynamic winding field and the permanent magnet (PM) field, which provides an approach to readily and quickly analyze the armature reaction. Based on the understanding of the mechanism, we found an effective way to suppress the armature reaction and propose a new VCA with double PM flux circuits. The electromagnetic performances of the new configuration are evaluated by comparing it with the conventional one. Both the theoretical and experimental results show that the new structure achieves lower magnetic saturation, lower armature reaction, and higher thrust density. A research prototype was constructed and experimentally tested to verify the theoretical results of the armature reaction, magnetic field, thrust, separation velocity, and other performances. The new VCA realizes a wide and accurate modulation of separation velocity for the CubeSats.

Design and Experiment on Heat Dissipation Structures of Ducted Fan Motor for Flying Electric Vehicle

Ducted fan motors play a crucial role in promoting various applications of flying electric vehicles. In ducted fan motor systems, motor performance affects the speed of the fan, the flow field of the fan affects the thermal field of the motor, and the thermal field influences the performance of the motor. The coupling model between fan static thrust, motor power, and motor temperature rise is established in this paper. After confirming the external dimensions of the motor, three cooling schemes of the motor casing are designed. The casing forms are as follows: model 1 with smooth surface, model 2 with circular fins, and model 3 with longitudinal fins. The optimization work was carried out on the geometric dimensions of two types of fins for model 2 and model 3, and the static thrust and heat transfer performance of the motors were calculated. This study proposes that the ratio of thrust-to-temperature rise is an indicator for future optimization design of ducted fan motors. Model 3 with longitudinal fins has a higher thrust-to-temperature rise ratio. The thrust temperature rise in model 3 has increased by 24.77% compared to model 1.

Influences of initial voltage and electrode size on propellant surface evolution in a coaxial pulsed plasma thruster

Coaxial pulsed plasma thrusters are well suited for microsatellites, but the discharge therein is radially nonuniform, resulting in uneven propellant ablation that ultimately degrades the thruster performance. This study investigates how the propellant surface evolves during 50 000 discharges in a coaxial pulsed plasma thruster with different outer electrode inner diameters and initial voltages, encompassing the propellant surface profile, concave depth of the propellant, and carbon deposition on the propellant surface. The results show that the propellant surface changes in real time throughout the 50 000 discharges. Increasing the outer electrode inner diameter leads to a proportional increase in the concave depth of the propellant, resulting in unstable ablation, and the increased inward concave depth increases the likelihood of discharge failure. Furthermore, increasing the initial voltage increases the energy density in a unit area of propellant, further contributing to the increase in the concave depth and discharge failure. The energy density distribution on the propellant surface is obtained via simulation, and by comparing the results with experimental data, the critical ablation threshold for the propellant is estimated as being ca. 0.7 J/mm2.

Research on the Combustion Mode and Thrust Performance of Rotating Detonation Scrarmjet Engines

Decades of research in hypersonic science and engineering have shown that scramjet engines are the preferred choice for powering hypersonic vehicles. In these engines, oblique shock combustion and rotating detonation combustion are two common combustion modes. This article primarily focuses on the numerical simulation and theoretical analysis of these two combustion modes. The research indicates that as the combustible mixture injected into the combustion chamber increases, the combustion mode transitions from deflagration combustion to self-sustaining detonation combustion and then to forced detonation combustion to maintain stable combustion in the chamber. Under self-sustaining detonation combustion mode, as the amount of combustible mixture injected into the chamber increases, the number of detonation waves gradually increases from one to multiple and the angle between the detonation waves and the inflow changes from acute to right angles. In the forced detonation combustion mode, the minimum shock wave intensity required to initiate detonation combustion can be obtained by drawing a tangent from the Rayleigh line. The structure generating oblique shock waves in the combustion chamber needs to match the shock wave intensity. In the detonation combustion mode, both the exit total temperature and specific impulse of the combustion chamber increase with the increase in Mach number at the inlet, while the pressure ratio decreases. This article provides a reference for the design of the combustion chamber in supersonic combustion ramjet engines.

Numerical study on ignition start-up process of an underwater solid rocket motor across a wide depth range

Solid rocket motors have important applications in the propulsion of trans-media vehicles and underwater launched rockets. In this paper, the ignition start-up process of an underwater solid rocket motor across a wide depth range has been numerically studied. A novel multi-domain integrated model has been developed by combining the solid propellant ignition and combustion model with the volume of fluid multiphase model. This integrated model enables the coupled simulation of the propellant combustion and gas flow inside the motor, along with the gas jet evolution in the external water environment. The detailed flow field developments in the combustion chamber, nozzle, and wake field are carefully analyzed. The variation rules of the internal ballistics and thrust performance are also obtained. The effects of environmental medium and operating depth on the ignition start-up process are systematically discussed. The results show that the influence of the operating environment on the internal ballistic characteristics is primarily reflected in the initial period after the nozzle closure opens. The development of the gas jet in water lags significantly compared with that in air. As the water depth increases, the ignition delay time of the motor is shortened, and the morphology evolution of the gas jet is significantly compressed and accelerated. Furthermore, the necking and bulging of the jet boundary near the nozzle outlet and the consequent shock oscillations are intensified, resulting in stronger fluctuations in the wake pressure field and motor thrust.

Trends in mass utilization of a magnetically shielded Hall thruster operating on xenon and krypton

The trends in mass utilization with increasing discharge voltage and current are investigated for a magnetically shielded Hall thruster operating on xenon and krypton. A 9 kW class shielded thruster is operated with discharge voltages from 300 to 600 V and discharge currents from 15 to 30 A on xenon and krypton. Experimental measurements of discharge current, thrust, anode efficiency, and ion velocity as a function of axial position are used to calibrate a multi-fluid 2D Hall thruster code at all operating conditions. The results of these calibrated simulations are employed to interrogate the plasma properties inside the thruster channel. A simplified 0D model for mass utilization evaluated on spatial averages of the simulated plasma parameters is employed to interpret the response of this efficiency mode with power for each propellant. It is found that with both higher voltage and current, mass utilization increases for both gases and their relative gap in this efficiency decreases. This can be attributed to the higher plasma densities and ionization rate coefficients at high voltage, and solely to higher plasma densities at high current. The driving factors for the increase in mass utilization are examined in the context of its nonlinear response to internal plasma properties. The behavior of mass utilization is also discussed in context of the gap in overall efficiency between the propellants. Finally, the implications of these results for improving the performance of high power Hall thrusters operating on krypton are examined.

Numerical modeling of the non-equilibrium condensation in multi-component flows through the high expansion ratio nozzles

Excessive flow expansion in the thruster nozzles reduces the flow temperature and pressure. This provides conditions for the condensation phenomena in the nozzle, which can affect the thrust performance of thrusters. In this paper, the condensation effect of water components in the decomposition products of hydrogen peroxide in a typical thruster nozzle is modeled. To describe the state and properties of the real gas mixture, the Peng-Robinson equation of state is used. Moreover, the effects of H2O2 concentration and nozzle profile on the condensation and nozzle performance parameters are studied. The results show that by more flow expansion and crossing the saturated state of condensable components, the thrust coefficient improves up to 6%. This improvement decreases by increasing the H2O2 concentration in the reaction chamber for a fixed nozzle expansion ratio. Due to the small growth of droplets during the free molecular regime of the droplet growth in the high-speed flow, the non-equilibrium condensation continues until the nozzle outflow. In addition, although all the profiles have the same expansion ratio for nucleation start, increasing the nozzle length results in larger nano-droplets, more liquid mass fraction formation up to 4%, and moving the Wilson point to the further upstream sections in the nozzle.