A micro Z-pinch pulsed plasma thruster (mu-ZPPT) was designed to address the low propulsion efficiency (<10 %) and low thrust power ratio (<20 mu N/W) common in traditional micro pulsed plasma thruster (mu PPT). Experimental and theoretical analyses demonstrate that within a capacitance range of 0.5-2.0 mu F, the discharge mode of mu-ZPPT transitions from multi-period oscillation to single period. Among these, 1.5 mu F is identified as the optimal capacitance, achieving the highest comprehensive performance. The plasma resistance exceeds 100 m Omega (compared to about 50 m Omega for traditional PPT), with an energy deposition efficiency of 95.1 %. Notably, at 1W power with an optimal capacitance of 1.5 mu F, the mu-ZPPT exhibits exceptional propulsion performance with a propulsion efficiency of 25.5 %, specific impulse of 613.3 s, and thrust ratio of 84.9 mu N/W-exceeding the efficiency of traditional micro pulsed plasma thrusters (mu PPTs) by over 150 % and thrust ratio by more than 8 times. The mechanism analysis attributes the superior performance of mu-ZPPT to electrothermal acceleration, which accounts for over 97 % of the impulse generation. The enclosed configuration of mu-ZPPT effectively accelerates the neutral gas produced by late-time ablation, reduces thermal radiation losses, enhances Joule thermal ablation effects, and ultimately boosts thrust power ratio. The mu-ZPPT in this study achieved higher propulsion efficiency (25.5 %) and higher thrust-to-power ratio (84.9 mu N/W) at lower power (1 W), bringing a new high-performance solution in the micro-propulsion field.

The very-low Earth orbit (VLEO) atmosphere-breathing electric propulsion (ABEP) system captures the rarefied atmosphere as the propellant in situ, which is expected to achieve long-term in-orbit residence of VLEO satellites. However, VLEO missions face critical challenges including intense atmospheric drag, severe surface erosion from oxygen-containing species, and strict constraints on onboard propellant storage and power supply. As an electrodeless thruster, the radio-frequency (RF) plasma thruster can effectively avoid the erosion of the electrode system by oxygen-containing atmospheric propellant, making it a highly promising atmosphere-breathing electric thruster solution. In this work, experimental research was conducted on RF plasma discharge characteristics using nitrogen–oxygen (1:1 to simulate the 210 km condition) mixed propellant, with the relative position of the inductive RF coil as a key reference for magnetic field configuration. Three inlet modes (axial, radial, tangential) and three magnetic field positions (upstream, midstream, downstream relative to the RF coil) were set as experimental variables, and the distribution laws of plasma parameters such as electron temperature and electron density in the discharge chamber were systematically analyzed via emission spectroscopy and RF-compensated Langmuir probe. The results show that under the same conditions, the axial inlet mode yields the highest electron temperature in the discharge chamber (up to 3.72 eV at 400 W/30 sccm), while the tangential inlet mode results in the most uniform electron temperature distribution (variation <0.3 eV in the [−3,3] cm axial range), which is more conducive to the uniform distribution and ionization of the nitrogen–oxygen propellant in the discharge chamber; when the magnetic field position is adjusted from upstream to downstream relative to the RF coil, the electron density exhibits a significant reverse response characteristic, that is, the peak region of electron density migrates upstream (from 2.0 × 1017 m⁻3 at 14 cm/300 W to 3.4 × 1017 m⁻3 at 12 cm/300 W). This study reveals the regulation mechanism of inlet mode and magnetic field position on nitrogen–oxygen RF plasma discharge, which is different from the research on single inert gas propellants in previous studies. The findings can provide a direct experimental basis for the coordinated optimization of the inlet structure and magnetic field configuration of RF plasma thrusters for ABEP systems.

In recent years, with the maturation of space technology, CubeSats have seen rapid development. Due to mass, volume, and power constraints, most CubeSats are not equipped with propulsion systems, and only very few are equipped with single propulsion systems. However, as the demands of space missions continue to increase, there is a growing need for CubeSats to gradually acquire capabilities such as autonomous deorbiting, rapid maneuvering, and precise attitude and orbit control. This paper proposes a novel dual-mode concept using HAN green ionic liquids as a shared propellant, combining cold gas and electric ignition. A DC electrolysis method was employed to investigate the effects of different mass fractions of ionic liquids and varying voltage conditions on the gas production phenomenon and electrical properties of HAN ionic liquids during electrolysis. Furthermore, the electric ignition method is applied to the gaseous mixture produced by electrolysis to investigate the ignition combustion characteristics and performance under different electric ignition power conditions. This paper adopts a TG-DSC-FTIR method to analyze the thermal decomposition behavior, energy release characteristics, and gas phase product composition of the most suitable ion liquid formulation for the dual-mode system, aiming to infer possible thermal decomposition mechanisms. The results show that the HAN ionic liquid propellant can achieve stable electrolysis gas production under specific voltage conditions (90 V) and HAN mass fraction (70 %). Under these conditions, the average power for electrolytic gas production was 65.8 W, with a decomposition time of 3.2 s. At an ignition power of 28 W, electric ignition for gas production could be achieved. Increasing the ignition power improved the ignition effect, resulting in a brighter flame. The novel cold gas-electric ignition dual-mode concept proposed in this paper is feasible through experimental validation of the principle. This system concept shares the propellant and has significant potential for application in future CubeSats and micro-nano satellites.

At room temperature, the mixture of hydroxylamine nitrate (HAN) solution and polyvinyl alcohol (PVA) powder solidifies to form an electrically controlled solid propellant. Investigating its curing mechanism can provide important guidance for formulation design. This study systematically examines the effects of different HAN solution concentrations, varying PVA contents, and molecular weights on the curing characteristics of the propellant. Multiple characterization techniques, including compression tests, thermogravimetric analysis, Fourier-transform infrared spectroscopy, and scanning electron microscopy, were employed to thoroughly investigate the curing mechanism of HAN-electrically controlled solid propellants (HAN-ECSP). The results indicate that HAN can form stronger hydrogen bonds with PVA, and high-concentration HAN solutions effectively promote the dissolution of PVA, leading to the formation of a uniform and dense three-dimensional network structure. This structure not only imparts excellent structural stability to the propellant but also provides good self-recovery properties. However, low PVA content and molecular weight result in insufficient crystalline regions within the propellant, with inadequate entanglement between PVA chains, which adversely affects the curing performance. To achieve optimal curing performance for HAN-ECSP, the HAN concentration should be maintained above 80%, the PVA content should be at least 15%, and the PVA molecular weight should exceed 166 000 g/mol.

The interaction between transient ablation plasmas and surfaces governs contamination and performance in pulsed plasma thrusters and related devices. Using nanosecond laser ablation as a controllable plasma source, this study investigates the deposition characteristics of aluminum-polytetrafluoroethylene propellant under high-vacuum conditions. Spatially resolved collection and multi-scale characterization (confocal microscopy, scanning electron microscope, energy-dispersive x-ray spectroscopy) reveal a strong correlation between plume dynamics and deposit properties. The results suggest pronounced spatial anisotropy in deposition, which appears to be associated with laser-induced momentum and asymmetric plume expansion. A systematic morphological evolution is observed: from a porous, ballistically aggregated network in the high-flux core, to dense, thermally sintered (and cracked) ceramic films, and finally to carbon-dominated spherical particulates via gas-phase nucleation and subsequent condensation in the periphery. Energy-dispersive x-ray spectroscopy unveils a sharp chemical stratification: central regions retain high F and Al concentrations, suggesting near-field reactive deposition and in-flight condensation of fluorine-bearing Al-F products (e.g. AlFx), where as peripheral deposits become carbon-rich and fluorine-deficient due to volatile species escape. This work proposes a unified deposition model that correlates the thermo-chemical evolution of a reactive ablation plasma source with its surface modification patterns. The findings provide fundamental insights into plasma-material interactions and offer a physics-based framework for predicting contamination and improving the design of ablation-based plasma sources.

Atmosphere-breathing electric propulsion (ABEP) systems can capture the rarefied atmosphere as propellant and are therefore promising for drag compensation in ultra-low Earth orbit (ULEO). In this study, a radio-frequency (RF) nitrogen-oxygen plasma thruster was tested under a ground nitrogen-oxygen propellant condition, and a calorimetric plume power measurement device was developed as a complementary diagnostic for long-duration plume-energy monitoring. The revised measurement interpretation distinguishes the RF input power, the calorimetrically inferred plume power, and the directed jet kinetic power. Under a representative direct-thrust measurement condition of 1000 W RF input power and 40.63 sccm N2 + 52.79 sccm O2, the pendulum thrust target measured a thrust of approximately 37 mN. The corresponding specific impulse, jet kinetic power, and RF-input-based total efficiency were approximately 178 s, 32 W, and 3.2%, respectively. During the approximately 209 h endurance test, the average cooling-water power was approximately 120 W, and the average calorimetrically inferred plume power was approximately 314.7 W after baseline subtraction and heat-transfer calibration. A linear trend analysis showed that the long-duration change of the inferred plume-power signal was much smaller than its short-period fluctuation, indicating no clear long-duration decrease in the intercepted plume-energy signal during the present test. The thruster also showed a stable active temperature-control response, with the average response deviation ranging from approximately -3.4% to 0.0%. Drag-compensation analysis at the 200 km reference altitude indicates that the measured performance level is relevant to ABEP-oriented ULEO operation under the assumed ground-test condition. The results should be interpreted as a controlled nitrogen-oxygen ground endurance test rather than a complete reproduction of the atomic-oxygen-dominated orbital environment.

Electrically controlled solid propellants (ECSPs) primarily convert electrical energy into thermal and electrochemical energy during operation. Understanding the role of these energy conversion pathways is essential for elucidating their ignition and combustion mechanisms. In this study, the ignition and combustion behaviors of the propellant were investigated by varying its intrinsic energy characteristics and energy application methods, and multiple characterization techniques and observation methods were employed. Adding boron powder enhances the heat release during high-temperature decomposition of the propellant, thereby improving its ignition and combustion performance. Under different power densities, the propellant exhibits similar ignition and combustion regulation capabilities under both thermal and electrical energy inputs. At a power density of 0.44 W/mm2, the propellant cannot be ignited by either method. When the power density increases from 0.88 W/mm2 to 2.20 W/mm2, the ignition delay time decreases from 4.58 s to 0.51 s under laser heating, and from 5.01 s to 0.55 s under electrical activation. As the power density increases, the proportion of electrical energy converted into electrochemical energy gradually decreases; at 2.20 W/mm2, only 1.5 % of the input electrical energy is converted into electrochemical energy, while the majority is dissipated as thermal energy. Most of the electrical energy is converted into thermal energy, which is concentrated on the anode propellant surface, leading to its ignition and combustion. However, the thermal feedback from the flame alone is insufficient to sustain the combustion. By adjusting the external energy input power, repeated ignition and combustion control can be achieved.

In recent years, the development of micro-electro-mechanical systems (MEMS) technology has driven the rapid advancement of CubeSats. The constraints of volume, power consumption, and cost have resulted in the majority of CubeSats not being equipped with propulsion systems, with very few exceptions. However, traditional monopropellant thrusters utilise expensive, high-concentration, precious metal catalysts, resulting in extremely high costs. This hinders the widespread application of chemical thrusters in the CubeSat domain. The paper proposes a novel approach for catalyzing the decomposition of green monopropellants using low-concentration catalysts. The catalytic decomposition of a promising green dual-mode propellant was studied by applying low-concentration platinum and iridium catalysts. Using drip experiments, the effects of different temperatures and catalyst conditions on the gas generation delay time of the green propellant were investigated. Furthermore, TG-DSC-FTIR analysis and SEM-EDS characterization were performed to examine the thermal and catalytic decomposition behavior, energy characteristics, functional group composition, and gas-phase product composition of the green propellant, along with catalyst characterization data, to infer the possible decomposition mechanisms. The results indicate that an increased preheating temperature leads to a shorter gas generation delay time for low-concentration Pt and Ir catalysts. At 200 degrees C, the average gas generation delay time was reduced to 0.41 similar to 0.79 s. This finding suggests that low-concentration Pt and Ir catalysts have the potential to meet the activation performance requirements of monopropellant thrusters. In comparison with thermal decomposition, the addition of low-concentration catalysts has been shown to reduce the initial reaction temperature, decrease residual mass, accelerate the decomposition process, and enable uniform energy release. Based on FTIR results, the composition and sequence of gaseous products differed without a catalyst and after the utilization of Pt and Ir catalysts. SEM-EDS characterization analysis indicates that under the reaction conditions of the dual-mode propellant, low-concentration Pt and Ir catalysts exhibit macroscopic structural stability. The results of the study indicate that low-concentration metal catalysts have the potential to catalyse the decomposition of green dual-mode propellants. This approach offers cost advantages and significant potential for future micro-nano satellite applications.

This study employs multiple diagnostic techniques, including high-speed photography, optical flow method for velocity field analysis, and B-dot probe magnetic field measurements, to systematically investigate the discharge characteristics, plasma plume evolution, and internal nozzle magnetic field distribution of a micro Z-pinch pulsed plasma thruster (mu-ZPPT). The key findings are as follows: (1) the mu-ZPPT demonstrates a high energy deposition efficiency of 90%, enabling it to accelerate restrike products (15-20 km s-1) and harness late-time neutral gas for thrust, thereby alleviating the detrimental impact of the late-time ablation effect. However, a significant 97.3% of this energy is deposited within the initial 0.64 mu s, and the subsequent restrike, characterized by negative power, warrants suppression. (2) The evolution of the plasma plume progresses through four distinct stages: initial, development, restrike, and late-time ablation, exhibiting a notable plasma grouping phenomenon. The initial stage is marked by high velocity (up to 65 km s-1) and low density, while the subsequent development stage is characterized by low velocity but high density, a clear indication of ablation lagging behind the discharge. Concurrently, the magnetic field evolves from a weak negative field (peak -0.021 T, magnetic pressure 260 Pa) to a strong positive field (peak 0.121 T, magnetic pressure 5600 Pa). (3) The plasma plume's dynamics are significantly governed by the magnetic field. Specifically, its characteristic downward tilt stems from a high magnetic pressure zone localized in the mid-to-lower region of the nozzle. This study elucidates the fundamental operational mechanisms of the mu-ZPPT, offering critical experimental guidance for optimizing the discharge circuit, mitigating the lag between ablation and discharge, and suppressing the restrike.

The thrust chamber is a critical yet vulnerable component of liquid rocket engines, with the "dog-house" failure mechanism on its inner wall constituting a primary constraint for enhancing engine life. The sandwich beam theory serves as a classical framework for predicting the service life of thrust chamber's inner wall, enabling the characterization of inner wall's damage and thinning as continuous temporal functions. Based on the concept of the sandwich beam model, this study independently establishes an improved sandwich beam model and presents a detailed derivation process. The improved model considers the existence of the cooling channel fillets, and is suitable for the case that the thermal strain difference of the cold and hot wall surfaces changes along the circumference, which better aligns with the physical and engineering practice. The calculation result shows that the introduction of the fillets, while improving the stress concentration, will lead to a reduction in the service life of the inner wall. Therefore, the fillet radius should not be too large. Comparative analyses between thinning models based on linear deflection distribution and arc-shaped deflection distribution were conducted, leading to the establishment of an optimized thinning model for inner wall with fillets. The model considers the influence of fillet radius and plastic compression effect on the thinning amount, and improves the reliability of life prediction. Finally, the improved model combined with creep constitutive relation was used to calculate the damage and life of the thrust chamber's inner wall, and the influencing factors of life were analyzed. This research provides valuable theoretical references for thrust chamber's life prediction and real-time damage assessment, while also offering conceptual foundations for offline damage mitigation control system design.

In the tandem dual-cavity scramjet combustor at the Mach 2.52 inflow condition, a detailed mechanism investigation of the plasma-assisted combustion process is conducted using high-speed flame chemiluminescence, schlieren imaging, and numerical simulations. Experiments show a weak flame state in the single rear cavity at a fuel injector ratio of approximately 3:1. In the presence of the gliding arc plasma, a stronger global flame and a stable state in the dual cavity are achieved. The flame intensity and heat release rate are significantly increased. OpenFOAM coupled with ZDPlasKin is applied in numerical simulations to calculate the detailed threedimensional plasma-kinetics mechanism of flow combustion and to reveal the combustion mechanism. Under the influence of the gliding-arc plasma, the kernel generates hot bubbles and modifies the flow-field characteristics of the shear layer above and within the front cavity. This facilitates the further transport of the kernel and flame. The kernel flows backward from the bottom of the cavity to the leading-edge corner recirculation zone, where it develops into a global flame connecting the two cavities. At the same time, the high turbulence shear-layer over the rear cavity is thickened and lifted. Detailed chemical reaction kinetics in the plasma region are investigated. The heat release process is dominated by the de-excitation reactions of N2* and O*, with a contribution rate two orders of magnitude higher than that of typical combustion elementary reaction heat release rates, such as the consumption of CH2 and OH. Novelty and significance statement: Through experiments and three-dimensional, detailed chemical-kinetics numerical simulations, this work innovatively investigated the mechanism and detailed chemical kinetics of the gliding-arc plasma-assisted supersonic combustion process in the tandem dual-cavity combustor. The plasma kinetics and main excited-state components contributing to heat release are revealed, along with the elementary reactions involved. The work on simulating 3-D detailed plasma dynamics for the tandem dual-cavity combustor configuration represents the most innovative contribution. This work guides the design and application of dual-cavity scramjet combustors. By applying gliding-arc plasma, supersonic combustion is enhanced in the weak-flame case. The verified comparison of experimental and computational flow-field characteristics and pressure distributions demonstrates the reliability of using OpenFOAM coupled with ZDPlasKin for three-dimensional simulations of plasma-kinetics reaction flows.

Radio-frequency plasma thrusters,as a type of electrode-less electric thruster,represent a preferred thruster solution for the long-duration operation of air-breathing electric propulsion systems in very low Earth orbits.However,ambient parameters significantly influence the radio-frequency discharge modes and the transition character-istics,which directly determine the thruster's operational stability under a wide range of conditions.We employed test diagnostic apparatus such as Langmuir probes and optical emission spectroscopy to measure the variations of plasma density and spectral intensity with radio-frequency power under different working conditions,including propellant type,mixture ratio,and flow rate.The influence of the propellant on the transition characteristics of radio-frequency dis-charge modes was subsequently analyzed.For argon,the discharge mode transition from capacitively coupled plasma to inductively coupled plasma occurs at 300 W,accompanied by consistent jumps in plasma density and spec-tral intensity.Due to its higher ionization energy,the required radio-frequency power threshold for discharge mode transition is significantly higher for nitrogen than for oxygen.In nitrogen-oxygen mixtures,an increased oxygen propor-tion reduces the transition threshold;however,a high oxygen fraction can facilitate the formation of negative ions due to its electronegativity,thereby suppressing the free electron density.The threshold of radio-frequency power for mode transition increases with the increase of gas flow rate,driven by the reduced gas residence time,the heightened electron-neutral collision frequency,and the consequent decline in ionization efficiency.The results provide important test evidence for elucidating the characteristics of radio-frequency plasma discharge mode transitions and for optimizing the operational parameters of radio-frequency plasma thrusters under wide-range conditions.

Micro flow piezoelectric proportional valve play an important role in the accurate control of thrust in electric thrusters such as electrospray thrusters and hall thrusters. However, the leakage and inability to open due to the micrometer level displacement nature remains a block in front of its reliable using.This study aims to explore the effects of operating temperature and pre-tightening force on the performance of micro flow piezoelectric proportional valves, and guide the pre-tightening force setting and assembly requirements of the valves through simulation results. The research results indicate that the performance of micro flow piezoelectric proportional valves is mainly affected by the operating temperature. When the working temperature is lower than the assembly temperature, liquid leakage may occur,while the working temperature is higher than the set value, it may not be able to open normally. At operating conditions below assembly temperature, excessive pre-tightening force required for valve sealing can cause the valve to remain normally closed at assembly temperature and above. Therefore, it is recommended to assemble at low temperatures. The main factor affecting valve performance is the working temperature, and ensuring the stability of the working temperature of the micro flow piezoelectric proportional valve is a key measure to ensure its optimal performance.

As the complexity and challenges of space missions continually increase, propulsion systems composed of clusters of engines, driven by the demand for high-thrust launch vehicles, are becoming mainstream. This paper constructs a dynamic simulation model of an engine cluster that includes multiple engines using oxygen-rich staged combustion cycles and their propellant supply systems. By setting different engine start time offsets and using asynchronous start time combinations generated by Sobol sequences, a systematic simulation analysis of the dynamic response of the engine cluster is conducted. The research results show that asynchronous startups cause severe nonlinear fluctuations and transient dynamics in the pressures and flow rates of the propellant tanks, as well as the inlet pressures and flow rates of the fuel and oxygen pumps. Additionally, the asynchronous startup of the engine cluster also results in significant fluctuations in the temperature and mixture ratio of the gas generator, with temperatures exceeding the design limit, potentially leading to component damage. In contrast, the temperature, mixture ratio, and pressure of the combustion chamber are also affected by the asynchronous startup, but the fluctuations are less severe. These findings not only reveal the complex dynamic behavior of clustered engines during asynchronous startups but also provide crucial theoretical support and practical guidance for the design and optimization of rocket propulsion systems.

Rolling element bearings are key components in rotating machinery, and their vibration responses provide essential information for fault diagnosis. However, traditional fault characteristic frequency models often neglect the combined effects of amplitude modulation (AM) and frequency modulation (FM) that occur under real operating conditions, leading to inaccurate spectral interpretation. This study proposes a novel analytical vibration signal model for rolling element bearings by analogizing their kinematic structure to a planetary gear system. Closed-form equations are derived to describe the spectral components caused by different bearing defects, and a simplified reconstruction method is introduced to calculate fault characteristic frequencies based on bearing geometry. The proposed model is validated using both the Case Western Reserve University (CWRU) dataset and experimental signals collected from a custom bearing test rig. Results demonstrate that the model effectively predicts fault frequencies and clarifies the spectral distribution of local and distributed defects, providing a new theoretical framework for fault feature interpretation in bearing diagnostics.

Laser ablation propulsion for microsatellites is hindered by the excessive mass loss and low coupling efficiency of polytetrafluoroethylene under repetitive pulsed laser irradiation. To address this limitation, we systematically investigate the ablation behavior of four propellant configurations: metal mesh-reinforced polytetrafluoroethylene, large-pore mesh/polytetrafluoroethylene, pure polytetrafluoroethylene, and an aluminum powder/polytetrafluoroethylene composite. Using a pulsed laser under vacuum conditions, we characterize the temporal evolution of surface morphology, ablation depth, and mass loss via 3D microscopy and scanning electron microscopy. Results reveal that the metal mesh-reinforced polytetrafluoroethylene exhibits the most stable and linear mass loss, with significantly suppressed ablation depth, attributed to a synergistic thermo-mechanical mechanism involving lateral heat spreading, structural reinforcement, and confinement of molten polytetrafluoroethylene. In contrast, large-pore mesh composites suffer from lamellar peeling due to anisotropic thermal stress, pure polytetrafluoroethylene develops deep porous craters, and the aluminum powder composite shows only moderate ablation resistance with compromised structural integrity. This study establishes that a continuous metal mesh skeleton effectively mitigates laser-induced ablation, providing a new design strategy for high-performance laser propulsion propellants.

Conventional parallel-plate pulsed plasma thrusters (PPTs) suffer from low propulsion efficiency (<10%), severely limiting their application in power-constrained micro- and nano-satellites. To address this, we propose a micro Z-pinch pulsed plasma thruster (mu-ZPPT) that utilizes a confined capillary structure and a divergent cathode nozzle, thereby enhancing energy conversion through the confinement of plasma and neutral gas. Performance characterization involved measuring impulse bits with a torsional pendulum, ablation mass with an analytical balance, and discharge parameters with probes to assess energy deposition efficiency. Plasma electrical parameters were derived using least-squares fitting based on an RLC circuit model, and experimental findings were compared with theoretical impulse bit calculations. Experiments conducted under capacitances ranging from 0.5 to 2.0 mu F reveal that the Z-pinch configuration enables single-pulse discharge (versus multi-oscillation in parallel plates), suppresses secondary discharge, and increases arc plasma resistance to above 120 m Omega (versus 50 m Omega in parallel plates). The mu-ZPPT mainly relies on the electrothermal acceleration mechanism to generate thrust, with the electrothermal impulse bit contributing >95%. At a capacitance of 1.5 mu F, the thruster achieves peak performance, with a 19.0% propulsion efficiency, a specific impulse of 531.9 s, and a thrust-to-power ratio of 72.8 mu N/W, doubling the efficiency of state-of-the-art micro-PPTs. Moreover, such a structure ensures an energy deposition efficiency of over 90% at lower discharge energies (<2 J), establishing the mu-ZPPT as a transformative solution for high-efficiency micropropulsion in next-generation nano-satellites.

Traditional electromechanical models typically assume a constant ablation mass, which deviates significantly from the actual operating conditions of thrusters. To more accurately describe the operating principles of micro-pulsed plasma thrusters (mu PPT), this paper establishes an ablation-electromechanical coupling model that considers the time-varying characteristics of propellant ablation mass. The model comprehensively accounts for the thermal decomposition of solid propellant, Fourier's law of heat conduction, and the ablation interface recession process, with plasma motion equivalently modeled using current sheet theory. The accuracy is validated by comparing it with mu PPT experimental data, showing prediction errors of less than 10 %, confirming the reliability of the coupling model. Based on the established model, a systematic parameter analysis is conducted, revealing the influence patterns of key design parameters on thruster performance. By studying ablation characteristics, the negative impact of the late-time ablation effect on the thruster is verified, confirming the dominant role of surface acceleration mechanisms in mu PPT. The first discharge period is the critical stage for thrust generation. This coupling model provides a reliable numerical tool for mu PPT design optimization and performance prediction.

As a promising candidate for micro/nano-satellite propulsion, the Micro-Pulsed Plasma Thruster (μPPT) requires a deep understanding of its acceleration mechanisms for effective optimization. This paper experimentally investigates these mechanisms. The μPPT’s low ionization efficiency and high neutral gas density result in the dominance of Electrothermal (ET) acceleration. Temporally, the first discharge cycle is dominated by Electromagnetic (EM) acceleration, while subsequent cycles are governed by ET effects. Owing to the circuit’s oscillatory nature, EM acceleration is prominent during the first and third quarter-cycles of the initial discharge. Spatially, the cathode side is dominated by EM acceleration, whereas the anode side is dominated by ET acceleration. This study elucidates the spatiotemporal variation of acceleration mechanisms within the μPPT and identifies a critical circuit limitation in small parallel-plate configurations: the significant energy loss induced by multi-cycle discharge oscillations.