Characteristics Of Thrusters Research Articles

Investigating the vibration characteristics of vector thrusters under complex inclination angle variations

ABSTRACT As autonomous underwater vehicles advance, the study of design and dynamic characteristics of vector thrusters is of significant engineering importance. In this paper, the mechanical model based on a vector thruster with a universal joint is established, followed by the analysis of its vertical torsional coupled vibration across different modes of inclination angle variations. Then, a simulated calculation of dynamic characteristics is carried out by the Newmark method. Finally, following the introduction of several structure parameters, the characteristics of vertical and torsional vibrations of the driven shaft under impact load are studied. The results show that the increase in the stiffness ratio can reduce not only the amplitude of angular displacement but also the stability time of vibration. The combinations of inclination angles should be selected preferentially, and the action value of actuators must be reasonably adjusted to ensure the inclination angle ratio remains as small as possible in the thruster control.

Research on numerical simulation of tandem thrusters flow-induced noise

A tandem propeller is a kind of propulsion device that is installed on the outer axis and the inner axis with the same axis respectively and rotates positively and negatively. Taking DTMB tandem thrusters as an example, which consists of propeller3689 and propeller3849, the hydrodynamic characteristic simulation of the tandem thrusters is completed, using sound analogy theory. The simulation results are consistent with the experimental results. The speed distribution of the tandem thrusters and the hydrodynamic wake of the tandem thrusters are analyzed. Through the analysis of acoustic radiation characteristics, it can be concluded that the noise characteristics of tandem thrusters are more complex than those of ordinary propellers, and there are more line spectrum characteristics in them.

Modeling of the Flight Performance of a Plasma-Propelled Drone: Limitations and Prospects

The resurgence in interest in aircraft electro-aerodynamic (EAD) propulsion has been sparked due to recent advancements in EAD thrusters, which generate thrust by employing a plasma generated through electrical discharge. With potentially quieter propulsion that could contribute to the generation of lift or the control of attitude, it is important to determine the feasibility of an EAD-propelled airplane. First, the main propulsive characteristics (thrust generation and power consumption) of EAD thrusters were drawn from the literature and compared with existing technologies. Second, an algorithm was developed to couple standard equations of flight with EAD propulsion performance and treat the first-order interactions. It fairly replicated the performance of the only available autonomous EAD-propelled drone. A test case based on an existing commercial UAV of 10 kg equipped with current-generation EAD thrusters anticipated a flight of less than 10 min, lower than 30 m in height, and below 8 m · s −1 in velocity. Achieving over 2 h of flight at 30 m of height at 10 m · s −1 requires the current EAD thrust to be doubled without altering the power consumption. For the same flight performance as the baseline UAV, the prediction asked for a tenfold increase in the thrust at the same power consumption.

The interelectrode breakdown mechanism and discharge characteristics of the electrospray thruster

In the actual working process of electric thrusters based on high-voltage electric fields, the discharge breakdown phenomenon is universal and complex, and such phenomena will have a significant impact on the thruster structure, working state and spacecraft system. In order to study the interpolar discharge breakdown characteristics of ionic liquid electrospray thrusters, a basic electrospray model and test system were constructed, and the change curves of discharge characteristic parameters such as breakdown voltage, threshold current, breakdown voltage frequency and so on in the range of 7×10−3~105 Pa with air pressure and transmitter inner diameter were obtained, and the air pressure range that the electrospray model could work in normally was calibrated. The results show that the breakdown voltage characteristic curve of the electrospray model has typical minimum characteristics, and the minimum values all appear around 80 Pa. Lowering the air pressure below 10−2 Pa can effectively increase the breakdown threshold between the poles and the emission current, thereby obtaining a larger voltage regulation range, and when the air pressure is reduced to 7×10−3 Pa, the breakdown can reach more than 3200 V. The 60-μm inner diameter emitter performed better in the discharge experiment, and the breakdown threshold, emission current and operating area range were better than the slightly larger inner diameter emitter under the same working conditions.

Effect of radial scaling down on the performance of low-power external discharge plasma thrusters

The external discharge plasma thruster removes the annular channel of the traditional Hall thruster, allowing the ionization region to appear outside the thruster, which is expected to reduce the plasma-wall interaction and extend the thruster lifetime. Due to its simpler structure and lighter weight, the external discharge plasma thruster is expected to be scaled down to lower power levels. In this work, the effect of radial scaling down on the performance of low-power external discharge plasma thrusters was experimentally investigated by comparing the discharge characteristics, thrust performance, beam characteristics, and spectral characteristics of thrusters with two different anode diameters of 15 and 30 mm. The 30 mm thruster has thrust of 78–235 mN, specific impulse of 566–1071 s, and anode efficiency of 11.8–23.4% at 78–235 W anode power. Compared with the 30 mm thruster, the anode power and thrust of the 15 mm thruster were decreased by approximately half. Plume diagnostic results show that the 15 mm thruster had a higher current utilization, higher ion peak energy and more active ionization due to better electron confinement. However, the propellant utilization of the 15 mm thruster was lower because of the shortened ionization region length.

Fault Diagnosis Method for an Underwater Thruster, Based on Load Feature Extraction

Targeting the problem of fault diagnosis in magnetic coupling underwater thrusters, a fault pattern classification method based on load feature extraction is proposed in this paper. By analyzing the output load characteristics of thrusters under typical fault patterns, the load torque model of the thrusters is established, and two characteristic parameters are constructed to describe the different fault patterns of thrusters. Then, a thruster load torque reconstruction method, based on the sliding mode observer (SMO), and the fault characteristic parameter identification method, based on the least square method (LSM), are proposed. According to the identified fault characteristic parameters, a thruster fault pattern classification method based on a support vector machine (SVM) is proposed. Finally, the feasibility and superiority of the proposed aspects are verified, through comparative simulation experiments. The results show that the diagnostic accuracy of this method is higher than 95% within 5 seconds of the thruster fault. The lowest diagnostic accuracy of thrusters with a single failure state is 96.75%, and the average diagnostic accuracy of thrusters with five fault states is 98.65%.

Investigation of the ion-optical system of an ion thruster with a microwave plasma generator with a power of up to 10 W

A method for improving the tactical and technical characteristics of ion thrusters while maintaining the specified electrical characteristics of the power source is proposed. This method is considered on the example of increasing the accelerating voltage on the electrodes of the ion-optical system of the prototype ion thruster. The method is based on the creation of a dielectric layer on the electrodes of an ion-optical system. Experimental and theoretical studies of the operation of the ion-optical system as part of the ion thruster have shown that the presence of a layer of dielectric coating on the electrodes allows you to achieve a steady-state voltage of up to 740 V, and in the absence of this layer-no more than 570 V. The calculated values of the thrust for the prototypes of ion thrusters were 14 μN with an ion-optical system having a dielectric layer and 8.4 μN a dielectric on the electrodes. The presence of dielectric coatings on the electrodes of the ion-optical system makes it possible to improve the characteristics of ion thrusters without significant design changes.

Time-frequency-domain method for thrust noise characteristics of electric thrusters

Electric thrusters are considered to be the most promising propulsion system of choice for space gravitational wave detection. However, the requirement of high accuracy and drag-free control results in significant challenges in thrust measurements. To better understand the thrust noise characteristics, this work proposes a time-frequency-domain method for thrust noise analysis, i.e., a scalogram method attained through wavelet analysis. Unlike the conventional Fourier transform method, it can provide 3-D relationships for the time-frequency and thrust noise. A free-state thrust measurement experiment identified that the environmental temperature produces the dominant impact in the thrust noise. During a 70,000 s measurement, even without thruster operation, the thrust noise increased from 0.3 μN/Hz1/2 to 19 μN/Hz1/2 as the environmental temperature increased from −1 °C to 3 °C. The corresponding frequency of peak thrust noise for the entire period tends to be the same of 0.1 Hz. Additionally, a Hall thruster thrust measurement experiment was conducted to verify the method's time-varying analysis capability for thrust noise. The complete time-frequency thrust noise variation throughout the entire operating process of a Hall thruster was observed. The results at different stages of thruster operation indicate that thrust noise has the greatest influence 150–300 s after ignition. This is because the thrust stand will take hundreds of seconds to reach thermal equilibrium after ignition. Another finding during the Hall thruster operation is that the 0.1 Hz frequency corresponding to the peak thrust noise is highly consistent with the frequency of peak thrust noise in the free-state thrust experiment. This reveals the fact that the 0.1 Hz frequency should be the dominant frequency for thrust noise caused by temperature drift in our thrust measurement system.

Numerical study of plasma-fluid characteristics and thrust performance of a low-power argon inductively coupled plasma electrothermal thruster

Low-power (from the sub-kilowatt range up to a few kilowatts) inductively coupled plasma (ICP) electrothermal thrusters for space propulsion are potential alternatives to low-power arcjet thrusters, which are often implemented on geostationary satellites for north–south station keeping. We develop an axisymmetric two-dimensional magnetohydrodynamic (MHD) numerical simulation technique using a two-temperature plasma model as a computer-aided engineering tool for low-power argon ICP electrothermal thrusters. Furthermore, to the best of our knowledge, this is the first study to provide a basic understanding of the plasma-fluid characteristics of low-power ICP electrothermal thrusters. To this end, we perform the MHD numerical simulation for a low-power argon ICP electrothermal thruster model that was developed for thrust measurement experiments to validate the proposed numerical simulation technique. The numerical results indicate that the plasma flow produced in the low-power argon ICP electrothermal thruster model is basically in a strongly thermal and ionizing nonequililbrium state. In addition, the numerical results show that the experimentally measured thrust forces can be reproduced with an underestimation of 10% or less over the RF input power range of 0–0.6 kW considered in the experiment. Moreover, the numerical results suggest that the main reason why the performance of the low-power argon ICP electrothermal thruster model is significantly lower than the target performance is the considerable heat loss of more than 90% of the input power through the flow-channel wall.

Modeling working processes of the marine thruster of the PMM-2M ferry-bridge machine

Objective. The article deals with aspects of modeling the working processes occurring in marine thrusters of amphibious vehicles, taking into account the specifics of their operation. Methods. The methods of 3D modeling of propellers in CAD and CAE packages are applied, which can determine and optimize the parameters of ongoing work processes with reliable accuracy. Results. A mathematical construct is proposed that allows calculating the characteristics of marine thrusters of amphibious vehicles. The propeller is designed to provide more thrust compared to the original design, making it possible to increase the speed of movement on the water and reduce the radius of circulation when moving through the water. The calculated version of the propeller provides an increase in thrust by 36%, allows developing a high speed on the water, and significantly reduces the radius of circulation of the ferry-bridge machine when maneuvering on the water. Conclusion. The proposed option for increasing the speed and maneuverability of ferry vehicles on the water is the most effective and least expensive; a promising direction for further research to achieve maximum efficiency is the creation and verification of software, hardware, and methodological complexes for modeling the joint operation of the "marine thruster - hull - power plant" system.

Correction to the paper “a simple model to determine the interrelation between the integral characteristics of hall thrusters” [Plasma Physics Reports 40, 229 (2014)

The paper is devoted to comparison of experimental data with theoretical predictions concerning the dependence of the current of accelerated ions on the operating voltage of a Hall thruster with an anode layer. The error made in the paper published by the authors in Plasma Phys. Rep. 40, 229 (2014) occurred because of a misprint in the Encyclopedia of Low-Temperature Plasma. In the present paper, this error is corrected. It is shown that the simple model proposed in the above-mentioned paper is in qualitative and quantitative agreement with experimental results.

A simple model to determine the interrelation between the integral characteristics of hall thrusters

A simple model to determine the interrelation between the integral characteristics of Hall thrusters with an anode layer is proposed. The model includes the equation describing the balance of forces acting on the Hall current region, as well as the relationship between the current of accelerated ions and the rate of working gas consumption. For fixed geometrical characteristics of a specific Hall thruster, this model allows one to interrelate the main integral characteristics of the thruster, such as the accelerating voltage, the magnetic field, the propellant flow rate, and the current of accelerated ions. The results of calculations for TAL-WSF/D-55—one of the most widely used Hall thrusters with an anode layer—are presented. The domain of existence of the discharge in the channel of the Hall thruster is analyzed. It is shown that taking into account kinetic effects of neutral particle expansion and working gas preheating lead to a decrease in the current of accelerated ions and the engine thrust.

Influence of the erosion of the discharge channel wall on the efficiency of a stationary plasma thruster

The integral characteristics of stationary plasma thrusters are studied experimentally during their long-term operation. It is shown that the monotonic fall of the thrust efficiency and specific impulse within the initial 500–1000 h of operation is explained, in particular, by a decrease in the propellant utilization efficiency. A physical mechanism underlying this phenomenon is suggested, and expressions predicting the variation of the output characteristics of the thrusters during the operating life are derived. Satisfactory agreement is observed between experimental and predicted results on the anode specific impulse variation in the course of testing several thrusters with different powers for endurance. An experimentally found correlation between the magnetic field configuration and the position of the erosion zone boundary on the discharge chamber wall is accounted for.

Performance characteristics according to the channel length and magnetic fields of cylindrical Hall thrusters

Performance characteristics of low power cylindrical Hall thrusters are investigated in terms of the length of the discharge channel. Thrust, efficiency, discharge current, and propellant utilization are evaluated for different channel lengths of 19, 22, and 25 mm. It is found that the propellant utilization and ion energy distribution function are strongly associated with the channel length. Increase of thrust and efficiency are also found with increasing channel lengths. These characteristics of the thruster are interpreted with possible generation of multi-charged ions due to increased residing time within the extended space inside the channel.

Transient Behavior of H2O2 Thruster: Effect of Injector Type and Ullage Volume

ROCKET-GRADE hydrogen peroxide has been used as a monopropellant and a storable oxidizer. However, because of the demand for a higher specific impulse, hydrazine andN2O4 are being used as the monopropellant and storable oxidizer, respectively. Recently, due to concerns regarding propellant toxicity, there has been a renewed interest [1] in the use ofH2O2 in propulsion systems [2–10]. A monopropellant thruster is operated in either the continuous or pulse mode. The thrust force and pressure instability are important issues in the continuous mode. For generating the desired thrust, the catalytic reactor size required for completely decomposing the propellant must be determined [8]. However, in the pulse mode (the main operationmode for attitude control systems), the response characteristics of the thruster are important. The catalyst activity, thruster component design (including the injector design), manifold volume, ullage volume in the reactor, and operating pressure influence the thruster response time. Tian et al. investigated the response time when using a combination of PbO and MnO2 catalysts [11]; they found that Ir=Al2O3 is unsuitable for use as a catalyst in a H2O2 monopropellant thruster [12]. Xu et al. studied the activities of various catalysts during H2O2 decomposition [13]. El-Aiashy et al. reported that the catalyst activity ofMnO2 increased when ZnO was added [14]. Hasan et al. reported that the activity ofMnO2 increased when promoters such as Ni, Cu, Bi, and Ce were added [15]. None of the aforementioned studies have addressed the effect of thruster design parameters on response times, although a few researchers have measured the thruster response time. Optimization of the thruster design (determination of the appropriate injector and ullage volume in the reaction chamber) can also influence the response characteristics. Therefore, we investigate the response characteristics of H2O2 monopropellant thrusters for three different thruster designs andmeasure the response times by varying the injector type, reactor volume, and catalyst grain size. AMnO2=Al2O3 catalyst is used for the decomposition of concentrated H2O2 (90 wt%).

Effect of laser ablation on the mechanical impulse formation in capillary discharge plasma

Within studies on the creation of new types of engines for small-sized space vehicles, we have experimentally studied the effect of the laser ablation of an electrode (insulator) material on the mechanical impulse of capillary discharge plasma. It is demonstrated that laser pulses can be used to reliably initiate of lowvoltage capillary discharge at low pressures, that is, under conditions where the usual discharge initiation is complicated. The proposed combined scheme allows the weight and size characteristics of plasma thrusters operating in pulsed and pulse-train regimes to be reduced.

Effects of channel wall material on thrust performance and plasma characteristics of Hall-effect thrusters

The effects of channel wall material on Hall thruster performance and on plasma characteristics were investigated. A laboratory-model Hall thruster THT-III was operated with three channel wall materials of BN, BNSiN and BNAlN. Both the discharge current and the thrust were affected by the nature of the channel wall materials. The measured axial distributions of wall and plasma potentials, radial and axial electron temperatures, and electron number density near the channel walls showed that the wall material affected ionization region and ion wall loss in the channel, resulting from secondary electron emission, although ion acceleration region was determined by the axial distribution of radial magnetic field. The difference in discharge current between channel wall materials was attributed to the difference in axial current density near the inner channel wall, depending on secondary electron emission.

Operational Characteristics of Colloid Thrusters

The operational characteristics of colloid thrusters were determined using experimental testing of several annular emitters and an analytical study based on empirical data. The analytical study was initiated following the exploration of the previously unknown linear region of operation at low-mass-flow rates where the current and thrust vary linearly with extraction voltage and mass-flow rate. This analytical description of colloid operation has reduced the extent of performance-mapping data required to characterize the performance of an emitter-propellant combination. Typical performance mapping includes determinations of thrust, specific impulse, emitter current, and charge-to-mass ratio as functions of emitter voltage and mass flow rate. Emitter performance was upgraded by obtaining stable operation with very low extractor drain currents and with no glow discharge. The relationship between performance level with beam divergence and emitter temperature was also examined. The performance indices and operational curves of a thruster designed to satisfy specific performance goals are derived using the analytical techniques. Methods of achieving improved performance are discussed.

Attitude-Translation Coupling in Drag-Free Satellites

The orbit of a drag-free satellite is determined entirely by gravity because it follows an unsupported proof mass which is shielded from all external surface forces by the satellite. A control system in the satellite senses the relative position of the satellite with respect to the proof mass and actuates reaction jets forcing the satellite to follow the proof mass. A drag-free control system may be added to any type of satellite, and in principle is independent of the satellite's attitude control system. Coupling, however, can occur because both the pickoff, which senses the relative position of the satellite, and the proof mass may be located away from the satellite mass center and the thrusters may be misaligned with respect to the mass center. In passive attitude control systems, misaligned translational control thrusters may produce significant disturbances to the attitude. These attitude motions can in turn be coupled back into the translation control system through an offset of the pickoff with respect to the vehicle mass center. This paper describes the mechanism of this type of coupling and provides design guides for gravity gradient satellites which are made drag free. The discussion begins with a simplified model to illustrate the nature of the coupling and develops the full six-degree-of-freedom equations to examine thoroughly all paths of coupling. It is shown that an important dynamic path for this type of coupling is through the interaction of along-track and radial translations (conventional orbital mechanics perturbation theory). Thrust which is produced in-track due to relative displacement along the orbit produces radial displacement which in turn activates thrusters in the radial direction. The vertical thrust, if misaligned, produces torques which change the vehicle's orientation and thereby produce an error signal in the horizontal direction if the sensor is not located at the satellite mass center. This error signal in turn requests additional horizontal thrust, and may thereby close the loop unstably. Stability boundaries as a function of thruster misalignment, pickoff offset, and attitude-motion damping ratio, are established and the alignment requirements are discussed. The analytical discussion is substantiated with results from a simulation which includes the nonlinear characteristics of thrusters and their modulators.