Thruster Performance Research Articles

Investigation on plasma enhanced decomposition of ammonium dinitramide (ADN) based propellant with optical diagnosis

Plasma-enhanced decomposition is a promising method to achieve reliable ignition and better thruster performance with an ADN-based propellant. High-speed imaging and optical emission spectroscopy are used to characterize the discharge plasma produced from ADN-based propellant vapor and argon carrier gas within parallel planes. Discharge regimes are independently controlled by varying working parameters. The transition from abnormal glow discharge (AGD) to filamentary discharge (FD) is identified from voltage-current characteristics and sudden changes in excitation electron temperature (Te-exc) and electron density (Ne) under a pressure of 0.2–10 kPa. Instabilities develop because there are more freedom degrees and higher collision probabilities after ADN-based propellant vapor is added. The product of Te-exc and Ne shows a higher energy transfer efficiency from input power to vapor in the FD regime. Water molecules increase the net dissociative attachment rate and effectively quench Ne. Although preheated vapor (as the discharge medium) has a much lower Ne, more radicals appear such as OH, NH, CH, CN, N2, N2+, and C2, which promote chemical interactions in an electric field and increase the probability of successful ignition. This can be attributed to thermalization due to an increase in the translational kinetic energy and rotational excitation, which has an important impact on chemical reaction kinetics. Therefore, preheating is verified as critical for improving the ignition and performance of a plasma-assisted ADN-based thruster.

Precision Feedback Control Design of Miniature Microwave Discharge Ion Thruster for Space Gravitational Wave Detection

The space gravitational wave detection mission has put forward multiple performance requirements for micro-newton thrusters, such as “fast response, high resolution, and low noise”. In order to solve the problem that the traditional open-loop propulsion system can hardly meet the high-precision performance, the microwave ion thruster with high accuracy and controllability of thrust estimation is adopted as the research object, and the precision feedback control method combining analog circuit and digital control technology is adopted to achieve a resolution of less than 0.1 μN, fast response of 20 ms, and 10−3~1 Hz frequency band noise less than 0.05 μN/Hz1/2 of the thrust performance. Experimental results show that, compared with open-loop regulation, the feedback control can effectively suppress various noises and disturbances caused by device temperature drift and complex operating characteristics of the thruster, and the anti-aliasing filter scheme adopted in this paper can suppress the folding of high-frequency data in the low- and medium-frequency bands during the digital control process, while it can effectively improve the system performance and reduce the impact of thrust noise. The feedback control strategy proposed in this paper verifies the possibility of a microwave ion thruster as an actuator for super “static and precise” space experiment satellites and provides a feasible solution for the application of a microwave ion thruster in deep space science experiment missions such as gravitational wave detection.

Simulation Study of the Swirl Spray Atomization of a Bipropellant Thruster under Low Temperature Conditions

The spray atomization of an injector significantly influences the performance and working life span of a bipropellant thruster of a spacecraft. Deep space exploration requires the thruster to be able to operate reliably at a low temperature range from −40 °C to 0 °C, so the effect of low temperature conditions on the atomization characteristics of injector spray is motivated to be comprehensively investigated. To study the swirl atomization characteristics of MMH (methylhydrazine), which is more difficult to atomize than NTO (nitrogen tetroxide), numerical simulations were conducted, employing the methods of VOF (volume of fluid) and LES (large eddy simulation) under low temperature conditions. The physical model with a nozzle size of 0.5 mm and boundary conditions with a velocity inlet of 3.89 m/s both follow the actual operation of thrusters. The development of spray atomization at low temperatures was observed through parametric comparisons, such as spray velocity, liquid total surface area, droplet particle size distribution, spray cone angle and breakup distance. When the temperature decreased from 20 °C to −40 °C at the same condition of flowrate inlet, those atomization characteristics of MMH propellant vary following these rules: the spray ejection velocity of MMH is significantly reduced by 7.7%, and gas-liquid disturbance sequentially decreases; the liquid film development is more stable, with a negative influence on atomization quality, causing difficulties for primary and secondary breakup, so the total surface area of droplets also decreases by 6.4%; the spatial distribution characteristics, spray cone angle and breakup distance vary less than 5%. Therefore, the low temperature condition can directly lower the combustion efficiency of thrusters with obvious performance degradation, but there are no significant changes in the propellant mixing and liquid film cooling. It is concluded that the bipropellant thruster can reliably work at low temperatures around −40 °C for deep space probe operation.

Editorial: Numerical simulations of plasma thrusters and/or related technologies

In the last few decades, electric propulsion has become in high demand on modern spacecrafts, from 8 very small nano-satellites to large communications orbital platforms. The difficulty of experimentally 9 characterizing the plasma discharge channels and the operative in-space conditions of plasma thrusters make 10 simulations a real game changer in this advanced technology/research field. In this respect, the rapid growth 11 of high-performance computing and esa-scale flops of modern architecture supercomputers has enabled 12 tackling more and more complex and computationally costly scenarios. With the use of different numerical 13 approaches, typically based on kinetic, hybrid or fluid plasma models, a large number of applications 14 can be covered: thruster physics and performance estimation, plasma thruster plume interaction with the 15 spacecraft, synthetic simulations of plasma diagnostic tools, etc.This research topic aimed to investigate, from a numerical point of view, plasma propulsion applications 17 or physical phenomena. When possible, in order to validate and/or tune the numerical tools, simulation

Micro-grooved hybrid spherical thrust bearing with Non-Newtonian lubricant behaviour

Recent tribology research has primarily focused on leveraging engineered grooved surfaces to enhance the lubrication performance of tribo-contacts. The performance of fluid film bearings may be improved by using a properly designed micro-groove on the surface of bearing. This paper deals with the performance of micro-grooved hole-entry hybrid spherical thrust bearing (HSTB) considering Non-Newtonian behavior of the lubricant. To analyse the Non-Newtonian behaviour of lubricant the cubic law model has been used. The solution of the governing Reynold equation in the fluid film domain of the hole-entry hybrid spherical thrust bearing is obtained by the FEM. In this study, the influence of rectangular and circular shaped micro-groves has been analysed. The parametric analysis has also been performed to obtain the optimum attributes (i.e. w¯g,d¯g,l¯gr) of the micro-grooves for both the shapes. The present investigation shows that the performance of the hole-entry hybrid spherical thrust bearing gets enhanced by considering the micro-grooves on the surface of bearing. Furthermore, the bearings performance is affected by the geometry of the micro-groove used. The value of fluid film reaction (F¯y) and stiffness coefficient (S¯) is observed to be increased by 23.33% and 343.98% at εy = 0 and κ¯=1 for the case of rectangular-shaped micro-grooved hole-entry HSTB configuration.

Control law synthetizing for an innovative indirect precooled airbreathing engine under off-design operation conditions

Indirect precooled engine (IPE) is a promising power for single-stage-to-orbit (SSTO) and the booster stage of a two-stage-to-orbit (TSTO) vehicles. Thrust of IPE is realized the most decisive factor that affects vehicle mission performance; therefore it is particularly critical to ensure that the engine can fully exert its thrust performance through active control since the flight parameters will change substantially alongside the trajectory. On account of this, control law of maximizing thrust for a single branch IPE is developed in current study. To this end, a high-fidelity off-design model of IPE is constructed to evaluate how engine performance is regulated by the various changeable parameters. The results indicate that engine thrust and specific impulse are functions of total fuel mass flow, nozzle throat area and the temperatures of preburner and main combustor, which can be conveniently controlled in real engines by actuators or changing the related fuel mass flow through fuel valves. Based on the influence characteristics revealed, the maximum thrust control law of the engine is synthetized, with the range of engine thrust can be regulated along a representative flying trajectory is evaluated. The results may provide some useful guidance about performance control of IPE with complex multi-branch intermediate cycles.

Thirty percent conversion efficiency from radiofrequency power to thrust energy in a magnetic nozzle plasma thruster

Innovations for terrestrial transportation technologies, e.g., cars, aircraft, and so on, have driven historical industries so far, and a similar breakthrough is now occurring in space owing to the successful development of electric propulsion devices such as gridded ion and Hall effect thrusters, where solar power is converted into the momentum of the propellant via acceleration of the ionized gases, resulting in a high specific impulse. A magnetic nozzle (MN) radiofrequency (rf) plasma thruster consisting of a low-pressure rf plasma source and a MN is an attractive candidate for a high-power electric propulsion device for spacecraft, as it will provide a long lifetime operation at a high-power level due to the absence of an electrode exposed to the plasma and a high thrust density. The high-density plasma produced in the source is transported along the magnetic field lines toward the open-source exit and the plasma is then spontaneously accelerated in the MN. By ejecting the plasma flow from the system, the reaction forces are exerted to the thruster structure including the source and the MN, and the spacecraft is resultantly propelled. The thruster will open the next door for space technologies, while the performance of the MN rf plasma thruster has been lower than those of the mature electric propulsion devices due to the energy loss to the physical walls. Here the thruster efficiency of about 30%, being the highest to date in this type of thruster, is successfully obtained in the MN rf plasma thruster by locating a cusp magnetic field inside the source, which acts as a virtual magnetic wall isolating the plasma from the source wall. The increase in the thrust by the cusp can be explained by considering the reductions of the loss area and the plasma volume in a thrust analysis combining a global source model and a one-dimensional MN model.

Maximizing the thrust performance of flexible caudal fin panels via experimental optimization

This paper discusses the experimental optimization of an artificial caudal fin to maximize thrust production. A combination of direct experimental measurement and Gaussian Process modeling was applied to optimize the time-averaged net thrust force produced by the fin panel. Specifically, this paper investigates the effect of the fin panel’s stiffness and flapping frequency on the thrust force. The Gaussian Process regression is primarily used to model the relationship between the flapping frequency and thrust force, with variations in the fin panel’s stiffness. Results show that the optimum flapping frequency that yields maximum force tends to decrease as the stiffness decreases. Furthermore, increasing the lateral amplitude displacement of the fin base leads to a higher time-averaged net thrust force. Further investigation shows that the phase difference between the trailing and leading edge of the oscillating fin should be about 90°to yield optimum time-averaged net thrust force.

Unsteady aerodynamic performance of a tandem flapping–fixed airfoil configuration at low Reynolds number

In nature, insects with their forewings and hindwings undergoing small-gap flapping motion experience strong aerodynamic interaction. Conventional studies mainly focus on the propulsion performance of tandem flapping wings, while the interaction between a flapping wing and a fixed wing in the tandem configuration at low Reynolds numbers (Re) is unclear. In this paper, we numerically studied the aerodynamic performance and vortex structure of this tandem flapping–fixed airfoil configuration. The effects of horizontal distance (LX), vertical distance (LY), and geometric angle of attack (α) of the fixed wing on the thrust and lift performance are investigated. The results show that LX dominates the propulsion performance, while LY and α control the lift performance. The thrust enhancement of the flapping airfoil is effective only within a small range of LX, and the thrust is mainly determined by the changing rate of the impulse of the vortices directly connected to the airfoils. The lift reaches its peak when LY approaches the plunging amplitude. Compared with a fixed airfoil, the flapping–fixed configuration shows a larger lift-to-drag ratio, indicating a lift enhancement led by the interaction with the upstream flapping airfoil. Moreover, increasing LY and α simultaneously can lead to additional advantages in lift generation. Further analysis shows that changes of LY and α both manifest in a variation of the effective angle of attack of the fixed airfoil, thereby manipulating its lift generation. This paper provides an aerodynamic database and guidance for the design of micro air vehicles using tandem flapping–fixed wings.

Stochastic learning and extremal-field map based autonomous guidance of low-thrust spacecraft

A supervised stochastic learning method called the Gaussian Process Regression (GPR) is used to design an autonomous guidance law for low-thrust spacecraft. The problems considered are both of the time- and fuel-optimal regimes and a methodology based on “perturbed back-propagation” approach is presented to generate optimal control along neighboring optimal trajectories which form the extremal bundle constituting the training data-set. The use of this methodology coupled with a GPR approximation of the spacecraft control via prediction of the costate n-tuple or the primer vector respectively for time- and fuel-optimal trajectories at discrete time-steps is demonstrated to be effective in designing an autonomous guidance law using the open-loop bundle of trajectories to-go. The methodology is applied to the Earth-3671 Dionysus time-optimal interplanetary transfer of a low-thrust spacecraft with off-nominal thruster performance and the resulting guidance law is evaluated under different design parameters using case-studies. The results highlight the utility and applicability of the proposed framework with scope for further improvements.

Longitudinal Vibration Transmission Control of Marine Propulsion Shafting with Friction Damper Integrated into the Thrust Bearing

Propeller-induced longitudinal vibration resonance in marine propulsion shafting systems causes great harm to the hull structure and is the primary source of shipboard noise. Integrating a friction damper with designed parameters into thrust bearings can prevent these issues. To investigate the performance of the damper-integrated thrust bearing in longitudinal vibration transmission control, an experimental and theoretical study is carried out in a laboratory-assembled test rig, which consists of components similar to the existing marine propulsion system. We developed a prototype of a thrust bearing designed with a friction-damping generation that allows switching from two supporting states, i.e., damper-connected and damper-disconnected states. Furthermore, a nonlinear analysis method for friction dampers is proposed. By this method, the way in which the friction damper changes the dynamic characteristics of the shafting system is analyzed. Based on the test rig, the acceleration frequency response function (AFRF) of the thrust bearing with and without a friction damper is measured. By comparison, the effectiveness of the friction damper is proved. The experimental results show that the friction damper suppresses the shafting longitudinal vibration response in a broadband frequency range and also confirms the stability of the damping effect, which does not change with the shafting rotational speed or static thrust from the propeller.

Coupling plasma physics and chemistry in the PIC model of electric propulsion: Application to an air-breathing, low-power Hall thruster

This work represents a first attempt to include the complex variety of electron-molecule processes in a full kinetic particle-in-cell/test particle Monte Carlo model for the plasma and neutral gas phase in a Hall thruster. Particular emphasis has been placed on Earth’s atmosphere species for the air-breathing concept. The coupling between the plasma and the gas phase is self-consistently captured by assuming the cold gas approximation and considering gas-wall and gas recycling from the walls due to ion neutralization. The results showed that, with air molecular propellants, all the most relevant thruster performance figures degraded relative to the nominal case using Xe propellant. The main reasons can be ascribed to a reduced ionization cross-section, a larger gas ionization mean free path due to lighter mass air species, and additional electron collisional power losses. While vibrational excitations power losses are negligible, dissociation and electronic excitations compete with the ionization channel. In addition, for molecular oxygen, the large dissociation leads to even faster atoms, further reducing their transit time inside the discharge channel. Future studies are needed to investigate the role of non-equilibrium vibrational kinetics and metastable states for stepwise ionization.

Effects of the emitters density on the performance of an atmospheric ionic thruster

A laboratory study on the performance of corona thrusters as a function of the spatial arrangement of the emitting electrodes is presented. The setup consists in arrays of emitters and collectors with independently adjustable spacings. This thruster is tested in different benches, for output velocity and direct thrust measurements. The measured velocity is also used to determine thrust by momentum balance, comparing it with direct readings. Electric measurements complete the dataset. The parameter space consists of two different collector spacings and several emitter spacings, with fixed interelectrode gap and voltage. Within this space lies an optimal collector–emitter spacing ratio 2:1.

Electrical characteristics of a Hall effect thruster body in a vacuum facility testing environment

The operational characteristics of Hall effect thrusters are altered by conductive surfaces in vacuum facilities. Conductive surfaces alter charge distribution in the plume by providing pathways for electron-ion recombination that do not exist in the spaceflight environment. Charge recombination pathways impact thruster performance and plume behavior through mechanisms that are not entirely understood. The incomplete understanding of the relationship between charge recombination pathways and thruster behavior limits the ability to characterize thruster performance through ground testing. This paper quantifies the effect of polarity and magnitude of body-to-cathode voltage on coupling between the thruster body and the local plasma environment. The effort operates the T-140 Hall thruster at a single, fixed operating condition of 300 V, 3.5 kW, with anode and cathode xenon flow rates of 11.6 ± 0.03 mg/s and 1.61 ± 0.12 mg/s, respectively. During data collection, the chamber was maintained at a pressure of 8.7 × 10–6 Torr-Xe. The thruster body-to-ground voltage is manipulated by varying body-to-ground resistance. Results show the thruster pole face and body circumference couple to the local plasma environment through distinct sheaths. The polarity of the body-to-cathode voltage determines the characteristics of these sheaths. Therefore, the body-to-cathode voltage controls the interaction between the thruster body recombination pathway and the local plasma environment.

Effect of combustion mode on thrust performance in a symmetrical tandem-cavity scramjet combustor

To investigate the relationship between the thrust performance and combustion mode in a scramjet combustor, the acceleration process of a symmetrical tandem-cavity combustor is numerically investigated based on previous experiments. The numerical results are essentially in agreement with the experimental results through high-speed photography and wall pressure comparisons, which verifies the reliability of the numerical method. Then, three groups of simulations with different equivalent ratios are designed under the condition of inlet flow Mach numbers 2.4, 2.5, 2.6, 2.7, and 2.9. Complete combustion efficiency and specific section thrust are defined to measure the combustion efficiency of kerosene and thrust performance. Four combustion modes are defined and observed in simulations, i.e. strong ram mode, weak ram mode, dual-mode ram mode, and scram mode. The simulation results show that the combustion modes in the former and the later cavities for this tandem-cavity scramjet combustor are related to each other. Combustion modes have a significant impact on the thrust performance of the combustor and the specific section thrust mutation is caused by mode transition. Under different inlet Mach numbers, appropriate equivalent ratio settings can improve thrust performance for this tandem-cavity supersonic combustor. Two ideal combustion modes realized by different injection schemes are proposed for better thrust performance of symmetrical tandem-cavity supersonic combustor.

The effect of magnet core size on the performance of a cylindrical Hall thruster with the near-anode cusp magnetic field

Effect of magnet core size on the performance of a cylindrical Hall thruster with the near-anode magnetic field has been investigated. Experiments found that extending magnetic core and rising discharge voltage show similar effects on the ionization process. When the magnetic core ends flush with the anode and discharge voltage is fixed at 240 V, the upstream ionization related with the electron E × B drift is hardly established under an approximately axial magnetic field and the ionization region is located at the near-axial magnetic mirror field, the thruster manifests the highest propellant utilization efficiency (as high as 1.43), the lowest current utilization efficiency (0.54) and beam divergence efficiency (0.52). When the magnet core extends towards the thruster exhaust, the enhancement of upstream ionization results in significant enhancements in the current utilization efficiency (by 14.6%) and the beam divergence efficiency (by 27.1%). However, the reduction of magnetic mirror ionization also led to a reduction of propellant utilization efficiency (by 32.1%). Therefore, the anode efficiency varies non-monotonically (climbing up and then declining) and has an optimal solution: shorter magnet core performs higher anode efficiency in low voltage condition and longer magnet core performs higher anode efficiency in high voltage condition.

Experimental investigation of an external discharge very low anode power (< 20 W) hall thruster

A sub 20 W wall-less Hall Effect Thruster (HET) was developed at the Asher Space Research Institute (ASRI), Technion. In this work, an initial study of the thruster performance and underlying physics was conducted. It was found that the anode efficiency of the thruster was low (~ 1%), mainly due to the low mass utilization efficiency. Typical performance figures are 90 mu N of thrust, specific impulse of 90 s and anode efficiency of ~ 1% at 3–4 W anode power. The thruster far-field plume was analyzed using a retarding potential analyzer. It was found that the beam divergence was relatively low at 57.7^circ (for 95% of the beam current) compared to other wall-less HETs. The voltage utilization efficiency was 38% for a discharge voltage of 1 kV and a mass flow rate of 1 sccm xenon. We speculate that the leading driver to the low mass utilization efficiency is the small ionization fraction associated with these very low power wall-less devices. It was found that the beam efficiency can be over 90% at discharge power levels < 3 W, and decreases with power down to less than 50%.

Thrust characteristics of a truncated Laval nozzle with a bell-shaped tip

This paper presents the results of a thrust performance study of an unconventionally shaped supersonic nozzle in the form of a truncated Laval nozzle with a bell-shaped tip. This nozzle shape may be used in the development of compact layouts of multistage rockets. The study was carried out using the ANSYS software package in a 3D formulation. The methodological approaches to the numerical calculation of a complex separated gas flow used in this study were verified in a previous study of the flow pattern in similar nozzle. Some results of exact calculations were compared with the results of experimental studies carried out at the Institute of Technical mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine for a model of a similar truncated nozzle with a bell-shaped tip blown with a cold air. This study detailed the features of the separated gas flow in a spherical tip connected (at the corner point) to a truncated supersonic Laval nozzle of conical shape. It was found that the pattern of the separated flow in the tip depends on the nozzle flow expansion degree (nozzle inlet pressure). At a relatively low nozzle inlet pressure, a developed separation zone is observed in the nozzle tip (between the jet boundary and the nozzle wall) with a subsonic flow from the external environment, which forms an almost constant static pressure from the tip inlet cross-section to the tip exit. At a nozzle inlet pressure at which the free boundary of the jet flowing from the truncated nozzle adjoins the nozzle wall, the static pressure in the tip varies almost linearly along the tip length from the corner point with the minimum pressure to the tip exit. The dependence of the thrust of a tipped nozzle on the nozzle inlet pressure is nonlinear. As the pressure upstream of the nozzle increases (or the ambient pressure decreases), the effect of the external environment on the tipped-nozzle thrust diminishes. It is shown that under "terrestrial conditions" the thrust of a truncated nozzle with a tip exceeds the thrust of a profiled nozzle with the same geometric expansion degree (due to the atmosphere “entering” the tip). Under "vacuum” conditions, the former is 8% less than the latter.

Mechanism and characteristics of thrust generated by a submerged detonation tube for underwater propulsion

The detonation engine, which can produce high specific impulse during the underwater detonation process (UDP), has become the forefront of underwater propulsion. In this paper, the thrust mechanism conducted in UDP and the propagation characteristics of the complex pressure waves are numerically studied, and the correlation between those two features is analyzed. The thrust from UDP is generated in a submerged detonation tube (SDT) and driven by the stoichiometric methane-oxygen mixture. The results show that detonation of the pre-filled combustible gas mixture gives rise to complex pressure waves and delivers several force impulses to the SDT. The impulses present different effects on the thrust performance, which is divided into two stages. In the first stage, before the detonation wave collides with the exterior water, the thrust is provided by the persistent back pressure effect of the detonation product. When the detonation wave propagates through the SDT exit and strikes the gas–water interface, a transmitted shock wave and a reflected shock wave are formed, which produce the impulses dominating the second stage. The reflected shock wave eventually impinges on the inner wall, imposing a force impulse on it. The pressure disturbance on the annular wall caused by the transmitted shock wave and subsequent detonation gas jet leads to another two thrust impulses. Finally, a comparison between the thrust of the SDT and its counterpart in the air is conducted to characterize the influence of UDP, and the effects of dimensional parameters of the SDT are also investigated.

Thrust and Levitation Coupling and Fluctuation Analysis of Linear Doubly-Fed Induction Motors With Variable Air-Gap

Within variable air-gap linear motors, such as those used in maglev trains, levitation and thrust forces are both realized through the common air-gap magnetic field, and thus they are intrinsically coupled. Additionally, nonidealities such as cogging, end effect, and saturation influence the thrust and levitation performance and result in force fluctuation. In this article, an accurate modeling method for a variable air-gap linear doubly-fed induction motor based on the magnetic coenergy reconstruction is proposed, where the effects of magnetic saturation, variable air-gap, and spatial harmonics are taken into consideration. This method uses a unified dataset to describe thrust, levitation, and their fluctuations. Furthermore, an adjustable air-gap linear motor prototype is designed and constructed. Bench tests are conducted to verify the proposed model.