Space Propulsion Research Articles

Confined Z-Pinch Magnetic Nozzle : A New Approach to Space Propulsion System that Utilizes Trapped Charged Particles

This paper provides a theoretical overview of a Z-pinch magnetic nozzle plasma propulsion system designed for Low Earth Orbit (LEO). The system utilizes trapped charged particles from the spacecraft's environment and magnetic confinement to generate thrust, reducing the need for onboard propellants. By employing a magnetic mirror configuration, the plasma is confined and accelerated to create propulsion. The proposed propulsion system is tailored to function in the dynamic space radiation environment of LEO, particularly in polar orbits, accounting for variations in solar activity that affect trapped protons and electrons. The mechanism involves converting magnetic plasma energy into directed kinetic energy, facilitating efficient plasma detachment and momentum transfer to the spacecraft. This study examines the feasibility, design overview, and constraints of implementing this propulsion technology. Key challenges include the variability of the space environment, technological complexities, radiation effects, plasma detachment efficiency, energy requirements, and material durability. Overcoming these challenges through innovative engineering and rigorous testing is essential for successful implementation. The Z-pinch magnetic nozzle plasma propulsion system promises to advance space propulsion technology, offering a more efficient and cost-effective method for satellite operations in LEO. With further research and development, it has the potential to support future satellite missions and space exploration initiatives.

Design of 2-DOF Thrust vector control system for Rockets and Missiles

This paper outlines the development and implementation of a Thrust Vector Control (TVC) system tailored specifically for solid propellant rocket engines, aimed at enhancing the portability of hybrid and liquid propellant rocket systems. The proposed system ensures trajectory stability and rapid response to flight emergencies by effectively addressing external perturbations like wind. Key components are Gyroscopic (GYRO) sensors, good micro-controller, and Servo motors, collectively referred to as Thrust vectoring components, which dynamically adjust thrust direction opposite to the rocket's trajectory to maintain stability. Drawing from an extensive literature review analyzing existing TVC system designs, with a focus on thrust characteristics and engine combustion timings, the design integrates considerations of geometric properties, kinematics, forces, energy requirements, safety, cost, and control methods, aligning with both mechanical and avionic design principles. The anticipated outcome is an advancement in rocket propulsion technology, characterized by improved angular speed control, trajectory linearity, and rapid response capabilities.

Verification of thermodynamic parameters of a mixture of generator gas on oxygen-hydrogen fuel with an excess of one of the fuel components

Liquid gas generators (LGG) are additional firing units in the power system of liquid rocket engines (LPRE). The LGG ensure the operation of the power units of the turbopump unit (TPU) of the engine by feeding combustion products (CP) to the turbine drive. The main criteria for the efficiency of the generator gas is the complex (RT)gg and the thermodynamic properties of the mixture, depending on temperature, pressure, the degree of excess of the oxidizer and the enthalpy of the fuel, attributed to the conditions of supply to the nozzles of the GG. Changing the parameters of the generator gas leads to a change in the turbine power parameters due to its effect on the adiabatic operation of the Lad turbine. Depending on the engine circuit under consideration, CP GG can perform work in other units and elements of the engine, as well as influence many parameters of the LPRE. Among the main ones can be noted: the power of the booster gas turbine of the booster turbopump unit (BTPU) in the case of the selection of the generator gas after the GG or turbogas after the main turbine; the temperature of heating the refrigerant in the heat exchanger introduced in the GG; specific impulse of a liquid rocket propulsion system (LRPS), depending on the quantity and properties of the turbogas entering the exhaust pipe of the engine (for the engine circuit without afterburning the generator gas); mixing in the combustion chamber (CC) due to afterburning of turbogas entering the engine chamber after the turbine (for the engine circuit with afterburning of generator gas); parameters of the firing wall of the engine in the case of using a high-temperature gas curtain by blowing generator gas into the supersonic part of the nozzle. For many pairs of fuel during combustion in GG, the nonequilibrium of combustion products is characteristic (especially in hydrocarbon fuels).Due to the fact that the combustion products (CP) during the combustion of an oxygen-hydrogen mixture, due to the simplicity of the reaction, have time to form while staying in the GG (i.e., the time of chemical equilibrium of the CP is less than or equal to the time of stay in the GG), their thermodynamic parameters can be reliably determined using programs that simulate chemical equilibrium reactions. In this article, the issue of obtaining reliable results of thermodynamic calculations of generator gas at low and high coefficients of oxidant excess is investigated. Verification of parameters obtained in the programs “Astra” and “Rocket Propulsion Analysis” with calculated values was carried out. The most suitable program for performing engineering calculations and modeling the thermodynamics of liquid gas generators has been determined.

Tetrazole-grafted-hydroxyl terminated polybutadiene: A novel energetic binder for solid rocket propulsion

The present study reports on the synthesis, characterisation, and energetic aspects of poly-nitrogen containing azole-grafted hydroxyl terminated polybutadiene (HTPB) based energetic binders (EBs). A comprehensive computational analysis at the B3LYP/6-311++G(d,p) level of theory reveals that azole-grafted HTPB variants exhibit significantly higher HoF (1061.6 to 1659.8 kJ/kg) and densities (1.26 to 1.61 g/cc) as compared to traditional non-energetic binders such as HTPB (specific heat of formation (HoF): 781.8 kJ/kg, density: 0.92 g/cc),. Their ballistic properties, including specific impulse (Isp: 266.2 to 283.8 s) and density specific impulse (ρIsp: 484.5 to 547.4 g/cc.s) are also superior as compared to HTPB (Isp: 283.8 s, ρIsp: 547.4 g/cc.s). Moreover, the synthetic methodology for grafting azole compounds onto HTPB, utilizing cost-effective, metal-free, and environmentally sustainable procedures with straightforward iodine and base mediation, underscores its potential for large-scale industrial applications. In addition to its advantages in synthesis and energetic properties, tetrazole-grafted HTPB exhibits favourable physicochemical properties: it has a low molecular weight ([Mn(VPO)] = 3023), high specific gravity (1.25 g/cc at 30 °C), moderate flowability (η = 623.6 Pa.s at 30 °C and 62.7 at 60 °C), a low glass transition temperature (−57 °C), mechanical insensitivity (low impact sensitivity and no ignition up to 400 °C), high thermal stability (onset temperature, Td,onset = 190 °C, TDSC = 246 °C), and a desirable hydroxyl value (47 mg/KOH). These properties make tetrazole-grafted HTPB an appealing choice for advancing novel propellant composites. Overall, these findings not only pave the way for advancements in HTPB-based propellants but also inspire further exploration in real-world applications, highlighting the transformative potential of azole-grafted HTPB in propellant technology.

A Space Propulsion System without Mass Ejection

Flight performance of long-range spacecraft is limited by the need to lift large amounts of fuel into orbit; thus, a propulsion system which will not depend on mass- ejection will be a most welcome development. Various past proposals which attempted to achieve this were rejected because it was thought that they contradicted basic laws of physics. An experimental mechanical device which operates without mass-ejection, (or interaction with the environment), and which can serve as a spacecraft propulsion system, was constructed and successfully tested. It is based on manipulating centrifugal forces, internally generated by rotating masses, and is not size or scale dependent. Full explanation, with photos and movies of its construction and operation, is presented. When fully developed, this technology will enable constant long-duration acceleration of spacecraft (decreasing flight times), permit novel physics research, or enable hovering with or without the presence of atmosphere.

Enhancement effect of Mo electrodes on the ignition characteristics of HAN-based gel propellant fuel and its combustion mechanisms

Hydroxylammonium nitrate (HAN) serves as a monopropellant fuel, renowned for its low toxicity and high-performance in space propulsion systems. HAN-based gel propellant may offer additional advantages, including long-term stability and excellent throttling capabilities. To better understand the ignition performance and combustion characteristics of this gel propellant, we prepared HAN-based gel propellant, and monitored its decomposition products using thermogravimetry (TG), Fourier Transform Infrared Spectroscopy (FTIR) and mass spectrometry (MS). Electrode materials of Mo, Cu, Ti and SUS 304 were fabricated to investigated their impact on ignition performance. It was found that the Mo electrode exhibited the shortest ignition delay time (ti) compared to Cu, Ti, and SUS 304. Meanwhile, the role of surface roughness of Mo electrode on the ignition delaytime (ti) and the ignition energy (Qi) were also studied. Our results revealed that, as the voltage increased from 100 V to 200 V, the fine-surfaced Mo electrode demonstrated a lower sensitivity to changes in ti compared to the rough-surfaced electrode. Additionally, the Qi of the rough-surfaced electrode decreased from 43.5 J to 16.7 J, while the Qi of the fine-surfaced electrode decreased from 52.5 J to 17.6 J. Combining a thermocouple, the combustion wave structure of the gel propellant was divided into pre-heating zone, condensed phase, electro chemical reaction zone, microexplosion-gases mixing zone and flame zone. Finally, the combustion model of HAN-based gel propellant was inferred to explain its combustion mechanism.

Exploring efficiency in next generation high temperature superconducting-enhanced applied field magnetoplasmadynamic thrusters: A combined numerical and experimental study

The integration of superconductors into space propulsion has emerged as a promising avenue. Superconductors have demonstrated significant effectiveness in space propulsion devices in terms of performance. This study introduces a low-power High-Temperature superconducting Applied Field MPD thruster, utilizing High-Temperature superconductors to propel plasma out of the thruster to generate thrust. Replacing traditional copper magnets with High-Temperature superconducting magnets has not only led to substantial improvements but has also addressed challenges such as the increased weight of traditional copper coils, the energy-intensive nature of copper magnets consuming high power, and the extremely low temperatures required by conventional superconductors, necessitating the use of helium cooling tanks, which can add to the overall weight of the thruster.The research combines numerical simulations with detailed experimental validation to evaluate performance. As predicted by the simulation model, experimental results demonstrate a thrust of 220 mN at 15 mg/s and 15.2 kW, and 180 mN at 5 mg/s and 14.8 kW, with a specific impulse of 2760s, while maintaining a discharge current of 200 A. Additionally, the discharge voltage in the experimental setup increases with magnetic field strength but decreases with mass flow rate. Plasma fall voltage also increases with both the applied magnet field strength set by the High-Temperature superconducting magnet and the mass flow rate. Meanwhile, the fractional electrode fall voltage decreases with the applied field. Discharge voltage, plasma fall voltage, and electrode fall voltage are analyzed as factors influencing the increase or decrease in thrust, specific impulse, and efficiency of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster.Furthermore, the study concludes with a comprehensive examination of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster, offering valuable insights for researchers. Notably, the achieved efficiencies for Applied field Magnetoplasmadynamic thruster at low input power surpass those of the same configuration with traditional copper magnet, highlighting the potential benefits of utilizing High Temperature superconducting magnets for enhanced performance.

Electrical ignition of ADN-based green liquid propellant in constant volume combustion chamber and continuous flow

Ammonium dinitramide (ADN) propellant for green space propulsion is perceived as a focal point in space propulsion research. The development of electrical ignition methods to realize the ignition for ADN-based liquid propellant can overcome the technical bottleneck of the catalytic ignition methods currently applied to ADN-based space thrusters. This work presents an experimental investigation of the electrical ignition and combustion characteristics of an ADN-based liquid propellant in a constant volume combustion chamber, considering ignition voltages and initial pressures. It is observed that increasing the ignition voltage and initial pressure reduces the flame retardation period, combustion duration, droplet lifetime, ignition energy, and total energy. Also, the combustion peak pressure decreases with an increase in the ignition voltage but increases with higher initial pressures. The energy consumption is mainly concentrated in the ignition stage, with the ignition energy accounting for approximately 76% to 96% of the total energy. The measurement of the concentrations of the keycombustion products are reported, highlighting that the mole fraction of CO decreases with an increase in ignition voltage and pressure. The ignition and combustion of the ADN-based liquid propellant in continuous flow were achieved for the first time through a combination of resistive ignition and high-temperature arc-assisted combustion. During the testing of the propellant ignition in continuous flow, there is a need for a continuous supply of electrical energy, and the combustion cannot be self-sustaining.

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.

Combustion behavior of NMP-001, a nitromethane-based green rocket propellant

Hydrazine’s carcinogenic and toxic nature makes its use in space propulsion expensive and necessitates careful launch campaign planning, which may cause delays. One already established alternative is to use energetic salt-based ionic solution propellants. However, the costs remain high because these salts are costly and highly regulated energetic materials. Additionally, propellants based on these energetic salts produce much higher temperatures during decomposition and combustion, as is the case for hydrazine. Consequently, expensive refractory metal-based alloys and high-temperature oxidation-resistant catalytic reactors must be used, which further drives up the costs. Alternatively, hydrogen peroxide may be used, though only in applications with a low specific impulse requirement. This study explores an alternative approach: using nitromethane, a commonly available laboratory solvent. NMP-001 is a nitromethane-based monopropellant that contains additives that allow for low-pressure combustion and reduce sensitivity to mechanical stimuli. Multiple combustion chamber configurations with varying characteristic lengths were tested in rocket combustion experiments. With the 11.5 m L∗-configuration, a 15.1 bar low-pressure limit could be determined. The 7.3 and 5.0 m L∗-configurations yielded a 16.2 and 28.3 bar low-pressure combustion limits, respectively. These outcomes are significantly below the 30 bar limit reported in the literature. Despite using only a simple heat-sink combustion chamber with a non-sophisticated injector element and a non-optimized propellant formulation, combustion was 83 to 95 % efficient with a C* ranged between 1133 to 1302 m/s. These figures are higher than for hydrogen peroxide and are comparable to the theoretical performance of state-of-the-art green liquid monopropellants.

PHM for Spacecraft Propulsion Systems: Developing Resilient Models for Real-World Challenges

PHM for Spacecraft Propulsion Systems: Developing Resilient Models for Real-World Challenges

Application of Lee–Tarver Model for Energetic Materials Safety Assessment Utilized in Aerospace Applications

Energetic materials, essential for rocket propulsion, play a crucial role in aerospace systems. These materials lack sufficient safety assessments and pose significant risks to life and infrastructure. The slow progress in this field is hindered by prohibitively high experimental costs and the inherently destructive nature of these experiments. This study aimed to evaluate modeling-based safety assessments, primarily utilizing the Lee–Tarver model. It specifically addressed the threat posed by shaped charges during transportation. Experimental analyses employed 50 mm JH-2 shaped charges, targeting RDX-based JH-2 formulation, since RDX is widely used in various rocket motor propellant formulations. TNT was also used to enhance the validation of the Lee–Tarver model findings. Various logistical configurations were explored, testing steel container walls ranging from 10 mm thickness to reinforced walls of 50 mm, 60 mm, and 65 mm steel. Comparative analysis with experiments reinforced the accuracy of the Lee–Tarver model’s numerical predictions. Both simulations and experiments affirmed that a 65 mm steel protection suffices for the safe transportation of RDX-based JH-2 formulations and TNT within a 10 mm storage box in logistical setups. Furthermore, this study emphasizes the Lee–Tarver model’s reliability for most energetic materials safety assessments, while acknowledging certain limitations.

Micro-flow properties of NEPE propellant in a rotating depressurization combustion environment of solid rocket motor

A profound understanding of microflow processes within the combustion chamber is paramount for optimizing engine performance and ensuring safe operation in the realm of solid rocket propulsion. Specifically, the scrutiny of Nitrate Ester Plasticized Polyether (NEPE) propellant under rotating depressurization combustion conditions has garnered substantial scholarly interest. The non-uniform distribution of propellant constituents and the ensuing heterogeneity in combustion rates convolve fluid flow dynamics within the chamber, ultimately engendering intricate flow patterns, turbulent phenomena, and rotational effects. Hence, conducting numerical simulation studies assumes an imperative role in comprehensively analyzing and apprehending the intricate microflow characteristics under such demanding circumstances. This paper endeavors to accomplish this overarching objective by introducing a novel kinetic mechanism in the gas phase that meticulously accounts for complicated reactions between distinct oxidizing powders and fuel binders, representing an uncharted territory in the existing literature. Additionally, this study takes into meticulous consideration the influence of radiation effects and non-planar moving surfaces, as well as other pivotal components. Upon establishing the combustion model, exponential depressurization is induced within the gas phase, while the computational domain is deflected to accurately simulate the rotational effects. A series of methodical investigations is subsequently undertaken to validate the proposed numerical framework, yielding commendable predictive accuracy. Ultimately, this research meticulously examines the minute intricacies of microflow characteristics under such complex transient combustion conditions.

Propulsion Technologies for CubeSats: Review

This paper explores the wide-ranging topography of micro-propulsion systems that have been flown in different small satellite missions. CubeSats, known for their compact size and affordability, have gained popularity in the realm of space exploration. However, their limited propulsion capabilities have often been a constraint in achieving certain mission objectives. In response to this challenge, space propulsion experts have developed a wide spectrum of miniaturized propulsion systems tailored to CubeSats, each offering distinct advantages. This literature review provides a comprehensive analysis of these micro-propulsion systems, categorizing them into distinct families based on their primary energy sources. The review provides informative graphs illustrating propulsion performance metrics, serving as beneficial resources for mission planners and satellite designers when selecting the most suitable propulsion system for a particular mission requirement.

Rocket science unveiled: A differential equation exploration of motion

Through the perspective of differential equations, the report "Rocket Science Unveiled" explores the amazing invention of rocket propulsion. In order to study, comprehend, and forecast the behavior of rocket engines, differential equations are essential. In order to better understand and analyze this intricate anomaly, the report aims to investigate the underlying mathematics of rocket propulsion and how differential equations work. We apply the differential equation to clarify the fuel consumption and thrust generation rates. In addition, we utilize Newton's rule of motion to explain the relationship among thrust, mass, and acceleration. Working on this study allowed us to discover the anticipated outcome for both position location and spacecraft position determination. For iterative operations, we used Euler's approach because the analytical calculation of differential equations is complicated, we used Euler's method for iterative operations. Knowing the rocket's initial or previous value allows us to locate or establish its placements with ease.

DEVELOPMENT OF THE DEVICE FOR CONTROLLING AND MEASURING THE SUPPLY OF WORKING GAS FOR LABORATORY TESTING OF THE ELECTRIC PROPULSION THRUSTERS

The object of the research is a device for measuring, controlling, and indicating the flow rate of working gas during laboratory testing of electric propulsion systems for space applications. The problem lies in the need to ensure high accuracy in controlling and measuring the flow rates of working gas for such devices. The obtained results include: analysis of existing industrial flow meters that could be used in development; proposed structural and functional diagrams of the device; experimental investigations into the accuracy of the developed device. Based on the analysis of parameters of the most common flow meters for development, the Bronkhorst F-201CV-100-AAD-22-V regulator was selected. Based on the selected regulator, a device was developed with a 32-bit ARM architecture. This architecture provides for the use of floating-point calculations, easy reprogramming of the device, and fast indication of values displayed on the indicator. The device features galvanic isolation between the control bus and the UART port of the microcontroller. The device allows for regulation, measurement, and indication of the flow rate of working gas into the anode block and hollow cathode during testing of Hall thrusters. Using the device allows for recommendations regarding the design of onboard systems for supplying working fluid to electric rocket propulsion systems. Based on the presented development and conducted research, laboratory prototypes of the devices were manufactured. The measurement error throughout the entire range of working gas flow rates does not exceed 0.7%. The overall error of the stand control and measurement system for working gas flow rates is up to 3%. The developed device provides measurement accuracy that satisfies the developers of electric propulsion systems.

Elucidating the challenges in the development and deployment of refractory complex concentrated alloys for additive manufacturing

The increasing demand for advanced materials with load-bearing capabilities in extremely high-temperature environments has brought considerable attention to the design and fabrication of novel refractory complex concentrated alloys. Despite the notable progress in design of customized alloys, there are still significant challenges regarding their processing and fabrication, highly limiting their exploitation. In this study, model non-equiatomic Nb-Ta based refractory complex concentrated alloys were created as potential materials with improved high-temperature stability and performance for space propulsion applications. Gas atomization was used for producing spherical pre-alloyed powders. Two different additive manufacturing techniques, i.e., laser powder bed fusion and cold spray were used to explore the manufacturing complexities of this novel alloy. The results highlight the importance of powder’s features on its processability by laser powder bed fusion and cold spray. High interstitial content, inhomogeneous microstructures, fractured and hollow particles, and the presence of microsegregation, led to embrittlement, and in case of cold spray, the wide particle size range of the feedstock resulted in undesired flowability, adversely affecting the processibility of the developed material. The obtained results provide valuable insights and recommendation for improving design, preparation, and fabrication of custom-designed refractory alloys for additive manufacturing.

Experimental and numerical study of stabilized flame in inverse coflow turbulent jet using nanosecond repetitively pulsed discharges

Methane has become increasingly popular in rocket propulsion, but low stability and limited flammability range have always been a concern about methane-powered systems. Many stabilization methods have been developed to change the geometrical or flow characteristics of the burner. However, most of these efforts have yet to be practically successful due to cost and compatibility issues. Alternatively, other methods such as microwave, dielectric barrier, and nanosecond repetitive pulse (NRP) discharges have been proven to be efficient by modifying the kinetic and transport pathways. NRP discharges have shown promising results as one of the most effective low-temperature plasma (LTP) methods. In this paper, chemiluminescence imaging is used to study the effect of NRP discharge on liftoff and blowout, as the important stabilization parameters, by recording the liftoff height and liftoff/blowout velocities under a wide range of discharge (f=0–10 kHz and V=11–19 kV) and jet velocity (ν=2–60 m/s). Depending on these parameters, four different discharge regimes of corona, diffuse, filamentary, and arc were observed. The results have shown that high-intensity plasma in a filamentary discharge regime can provide a significant advantage in delaying the liftoff conditions, but no improvements in the blowout were observed. It was also found that NRP discharge can reduce the liftoff height. To explore the cause of the increased stability, a parametric numerical study is conducted using detailed plasma-assisted methane kinetic modeling coupled to a 1D opposed diffusion flame simulation. Results show that the extinction limits of diffusion flames can be dramatically enhanced by LTP due to the local formation of high radical and excited species concentrations, with subsequent recombination leading to increased temperature and higher reactivity in the flame zone. In addition, a 1D laminar flame speed evaluation shows that the plasma-generated active species can dramatically increase the flame speed, which, in turn, reduces the lifted flame height above the burner surface.

Internal Nuclear Propulsion System for Future Space Applications

Internal nuclear pulse propulsion (INPP) represents a novel approach to space propulsion, where controlled nuclear detonations occur within the spacecraft itself, offering potential advantages in safety and propulsion control compared to conventional external detonation designs like Project Orion. INPP relies on the same fundamental principles of nuclear pulse propulsion, utilizing the momentum generated by nuclear explosions to propel the spacecraft forward. Proposed INPP designs involve strategically placing small nuclear devices within the spacecraft's structure, detonating them in sequence to produce controlled impulses for propulsion. However, significant engineering challenges, such as radiation shielding, precise detonation control, and safety considerations, must be addressed to realize the full potential of INPP. Despite these hurdles, ongoing research and development efforts aim to harness the transformative capabilities of INPP, promising to revolutionize space exploration by enabling efficient and rapid interstellar travel while ensuring the safety of spacecraft and crew.

Effects of Different Initial Conditions on Combustion Process of Ammonium Dinitramide-Based Energetic Propellant in Straight Nozzle

As a new type of green propellant, ammonium dinitramide (ADN)-based energetic propellants have wide application value and development potential in the field of space propulsion. This paper delves into the intricate impact of varying initial temperatures, pressures, and propellant component ratios on critical parameters, including temperature, combustion rate, and heat release, in the straight nozzle of an ADN-based propellant. The findings indicate that an elevation in both initial temperature and ADN ratio expedites the thermal decomposition rate of ADN, thereby elevating the average temperature in the nozzle. However, the elevation in initial temperature has a negative effect on the overall rise amplitude of average temperature. Furthermore, the initial pressure setting is crucial in determining whether the oxidation reaction of the fuel CH3OH occurs in ADN propellants. When the initial pressure is greater than 10 atm, CH3OH is completely consumed, and the final average temperature is about 2650 K, which increases by 558.89% compared with that at 1 atm. Our work aims to provide theoretical guidance and practical optimization strategies for enhancing propellant performance and optimizing thruster structure design.