Thruster Performance Research Articles

A comparison of Fourier and POD mode decomposition methods for high-speed Hall thruster Video

Hall thrusters are susceptible to large-amplitude plasma oscillations that impact thruster performance and lifetime and are also difficult to model. High-speed cameras are a popular tool to study these dynamics due to their spatial resolution and are a popular, nonintrusive complement to in-situ probes. High-speed video of thruster oscillations can be isolated (decomposed) into coherent structures (modes) with algorithms that help us better understand the evolution and interactions of each. This work provides an introduction, comparison, and step-by-step tutorial on established Fourier and newer Proper Orthogonal Decomposition (POD) algorithms as applied to high-speed video of the unshielded H6 6-kW laboratory model Hall thruster. From this dataset, both sets of algorithms identify and characterize $m=0$ and $m>0$ modes in the discharge channel and cathode regions of the thruster plume, as well as mode hopping between the $m=3$ and $m=4$ rotating spokes in the channel. The Fourier methods are ideal for characterizing linear modal structures and also provide intuitive dispersion relationships. By contrast, the POD method tailors a basis set using energy minimization techniques that better captures the nonlinear nature of these structures and with a simpler implementation. Together, the Fourier and POD methods provide a more complete toolkit for studying Hall thruster plasma instabilities and mode dynamics. Specifically, we recommend first applying POD first to quickly identify the nature and location of global dynamics and then using Fourier methods to isolate dispersion plots and other wave-based physics.

Investigation of variable aperture on the performance and lifetime of ion thruster

Beam flatness is an important parameter that determines the performance and the lifetime of a gridded ion thruster. To improve the beam flatness of the 30 cm (LIPS-300) ion thruster, variable aperture ion optics that adapts to the decreasing ion density as the radius increases is proposed. It is the ion optics that the screen grid surface is divided into several zones, where the aperture diameter in each zone is determined by the ion density and the electron temperature upstream of the screen grid. The beam current density in the central area is artificially reduced. A particle in cell-Monte Carlo collision model is applied in this work to investigating the effect of variable aperture on the perveance and the maximum beam current per aperture by simulating the extraction, focusing and acceleration processes of ions. Taking into account the engineering implementability, the screen grid surface is divided into four zones. The hole diameter in each zone is decreased from 1.95 mm to 1.8 mm, 1.9 mm, 1.8 mm and 1.7 mm, respectively. The simulation results show that the maximum ion density in the center area of grid is decreased by 10.6% and 6.99%, while it is increased by 6.49% and 22.3% in the edge region, respectively. The beam flatness of the variable aperture ion optics is improved from 0.69 to 0.88. The erosion rate is decreased by 31.9%, but the total beam current is also decreased by 7.15%. The simulation results can provide a valuable reference of the development of the ion thruster.

On Cathode Position and Inclination Angle in Magnetically Shielded Hall Thrusters

This study systematically investigated the influence of the position and inclination angle of an externally mounted hollow cathode in 1.35 kW magnetically shielded Hall thrusters to establish the relationship between thruster performance, plume state, and cathode placement. For this purpose, a combined motor driver with a three-degree-of-freedom movement was designed to move the cathode over a wide range of positions and inclination angles. In an experiment performed on a thruster with outer magnetic columns, placing the cathode inside the magnetic separatrix resulted in a considerable amount of thrust. However, an extremely small distance between the cathode and thruster can cause a sharp increase in the discharge current; hence, the optimal cathode position should be determined. It was also found that a small inclination angle was favorable for obtaining high performance. In addition, a strong correlation was found between the plume image and thruster performance. Under improved performance conditions, the visible light region of the plume had a shorter axial length and a broader pink halo. Finally, a comparative experiment was conducted on a thruster with an outer magnetic shell. The thruster of this structure was observed to be insensitive to the cathode position; hence, to protect the cathode from ion sputtering, the distance between the cathode and thruster can be appropriately increased.

Direct Formulation of Bipropellant Thruster Performance for Quantitative Cold-Flow Diagnostic

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination, and nozzle expansion ratio. The theoretical formulation deduces a framework for a quantitative noncombustion test or cold-flow test to predict the performance indices. We simply employ water and dyed water as simulant liquids and then measure the local ratios of mixture and flow rate using an absorbance spectrometer. Density ratio mismatch between hypergolic propellants and water can be reasonably convertible. Combined with chemical equilibrium analysis, we obtain characteristic velocity and specific impulse across wide injection mixture ratios. The validity of the quantitative water-flow diagnostic is evidenced by comparing the results with those of corresponding combustion tests using nitrogen tetroxide and monomethylhydrazine as the propellants under several injector unlike-doublet and triplet configurations, showing that the mixing states of bipropellant thrusters under combustion can be reproduced using the water-flow diagnostic.

Three-dimensional neutralizer effects on a Hall-effect thruster near plume

The near plume of an SPT100-like Hall thruster is simulated with a hybrid particle-in-cell/fluid code to assess the three-dimensional effects of the neutralizer for two different asymmetric positions. The electron fluid model accounts for the effects of both magnetization and collisions, and closes the momentum equation with a simple polytropic closure for the electron temperature. The observed asymmetries are small in the majority of the plasma plume properties and in the achieved magnetic thrust and torque, which also show consistent trends with the available experimental characterizations. The only exception is represented by the 3D electric and electron current densities, which show persistent asymmetries that tend to diminish as the neutralizer is placed both farther downstream and radially outward, outside of the magnetic separatrix. These findings permit to justify the use of axisymmetric codes for Hall-effect thruster performance estimation. The results on the neutralizer asymmetry effects are found robust to different polytropic coefficients (i.e. electron temperature gradients). Finally, the comparison of the adopted polytropic closure model against an advanced one including the energy balance equation shows relevant differences in the electron temperature but still a good agreement in the electron current density map.

Space Electroosmotic Thrusters in Ion Partitioning Soft Nanochannels.

Space electroosmotic thrusters (EOTs) are theoretically investigated in a soft charged nanochannel with a dense polyelectrolyte layer (PEL), which is considered to be more realistic than a low-density PEL. When the PEL is dense, its permittivity is smaller than the one of the electrolyte solution layer, leading to rearrangement of ions in the channel, which is denoted as the ion partitioning effect. It is noted that fluid viscosity becomes high within the PEL owing to the hydration effect. An analytical solution for electroosmotic velocity through the channel is obtained by utilizing the Debye–Hückel linearization assumption. Based on the fluid motion, thruster performances, including thrust, specific impulse, thrust-to-power ratio, and efficiency, are calculated. The ion partitioning effect leads to enhancement of the thruster velocity, while increase of the dynamic viscosity inside the PEL reduces the flow rate of the fluid. Therefore, these performances are further impacted by the dense soft material, which are discussed in detail. Moreover, changes or improvements of the thruster performances from the dense PEL to the weak PEL are presented and compared, and distributions of various energy items are also provided in this study. There is a good result whereby the increase in electric double layer thickness promotes the development of thruster performances. Ultimately, the simulated EOTs produce thrust of about 0 to 20 μN and achieve thruster efficiency of 90.40%, while maintaining an appropriate thrust–power ratio of about 1.53 mN/W by optimizing all design parameters.

Real-Time On-Orbit Estimation Method for Microthruster Thrust Based on High-Precision Orbit Determination

In order to enable the micro-nanosatellites equipped with microthrusters to better complete various space applications, it is necessary to estimate the thrust performance of the microthrusters in real-time on orbit. This paper proposes a real-time on-orbit estimation method for microthrust based on high-precision orbit determination. By establishing a high-precision orbit dynamic model, the microthrust generated by a microthruster is modeled as a first-order Markov model, combined with a high-precision GNSS measuring device, and the satellite position is obtained through the cubature Kalman filter algorithm, velocity, and thrust real-time on-orbit estimates. For a thrust of 100 μN, the error accuracy of the on-orbit estimation is 3.98%; for a thrust of 500 μN, the error accuracy is 1.79%; for a thrust of 5 mN, the error accuracy can be reduced to 1.43%; and when the thrust is 500 μN, the accuracy of orbit determination is 16 cm. This method solves the problem that the traditional on-orbit thrust estimation method cannot perform real-time on-orbit estimation of microthrust on the order of hundreds of μN. The real-time on-orbit estimation of microthrust of micro-nanosatellites equipped with microthrusters of the order of hundreds of micronewtons or even several mN to tens of mN has certain reference value.

Impulse and electric charge characteristics of chemical propellant under pulsed laser irradiation

To enhance the thrust performance of laser-pulsed plasma thruster (LPPT), we proposed to use chemical propellant as substitute for ordinary PTFE propellant. The discharge characteristics and thrust performance of chemical propellant were analyzed. Meanwhile, the contribution of different energy sources and the physiochemical process between laser and chemical propellant were investigated. The results showed that the chemical propellant had higher plasma peak current and shorter discharge delay time compared to PTFE propellant. For chemical propellant, the contribution of laser energy was about 3%, while the electric energy and chemical energy was less than 17% and more than 80%, respectively. Based on the self-made torsion pendulum micro-impulse test bench, we found the chemical propellant had the highest impulse bit of 27.9 μN·s at 0 V and 69.9 μN·s at 1500 V. Meanwhile, the PTFE propellant had the lowest laser ablation mass, which is 3.70 μg/shot. The chemical reaction between laser and chemical propellant enhances the impulse bit, and increases the mass consumption of the propellant. The chemical propellant had the highest specific impulse, which is 295.1 s at 0 V and 738.5 s at 1500 V. These findings can indicate that the chemical propellant is a better choice than PTFE propellant.

Pulse Performance Analysis of a 45 Newton Additively Manufactured Bipropellant Thruster

This paper describes the testing and pulse performance analysis of a 45 N thruster, designated A45, using a 19.78% monomethylhydrazine–80.22% hydrazine fuel blend with mixed oxides of nitrogen containing 3% nitric oxide as the oxidizer. The thruster components are fabricated using additive manufacturing (AM), which enabled fine design features on injectors to achieve fast response times. The main focus of this work is to demonstrate precise and repeatable pulse performance and combustion stability and to compare thruster performance with theoretical predictions from NASA CEA across the chosen test matrix. As part of the pulse performance analysis, impulse produced by each pulse of a given 50 Hz pulse train is characterized, thereby illustrating that true pulse repeatability on such thrusters only starts with the third pulse. This is caused by the priming of the injector manifold at the start of a pulse train. Specific impulse as functions of pulse ON and OFF times are also characterized for a 50 Hz duty cycle and dry injector manifold operation. This method of per-pulse performance characterization aids guidance, navigation, and control teams in thrust/impulse control for attitude control. Overall, the specific impulse, characteristic velocity efficiency, and response times demonstrated by this additively manufactured thruster are similar to those fabricated by traditional methods and support wider use of AM in the aerospace industry.

Modeling of the non-uniform combustion in a scramjet engine

Due to the high-speed of the air stream, the scramjet combustion is neither uniform nor complete. An empirical fuel distribution model is developed to describe non-uniform combustion for scramjet engines. The combustor is subdivided into different regions by radius with different local equivalence ratios. The conservation equations of the regions for the combustion and exhaust expansion are computed independently. The results indicate that scramjet thrust is more related to the fuel equivalence ratio and combustion efficiency. If the combustion efficiency is 100% and the fuel equivalence ratio is constant, there is no obvious effect of different fuel distribution to the engine except for the exhaust parameter distribution. It is also revealed that the sum of isolator shock loss and combustor Rayleigh loss is nearly constant under the same isolate cross-sectional area. Lower isolator inlet Mach is benefit to the thrust performance and the best thrust performance is at the thermal choking boundary. When the isolator inlet Mach increases, the thrust decreases.

Improving the performance of a water ion thruster using biased electrodes

This paper investigates whether biased electrodes can improve the performance of a microwave-discharge water ion thruster, which has been proposed as a propulsion system for CubeSats. Previously, for several microwave ion sources, the beam currents have been increased by mounting a biased electrode in the discharge chamber. The effect of a biased electrode on a water ion thruster was unclear because its magnetic field profile and characteristic length compared with the microwave wavelength are different from those used in previous studies. To investigate the effect, a thruster was divided into three parts: magnets, a microwave antenna, and a discharge chamber, which is comprised of an electrically conductive sidewall and a screen grid. The parts were isolated electrically. Bias voltages from −44 to ＋ 44V were applied to the microwave antenna or magnets by an additional power supply. These bias voltages were applied with respect to the screen grid and sidewall voltage. A maximum screen current of 12.7 mA was obtained for an antenna bias voltage of 38.2V. This is 2.3 times larger than the current without a bias voltage. This increase could be caused by a combination of an increase of plasma density and ion extraction efficiency. The plasma density could increase with the gain of the total absorbed power. The ion extraction efficiency could increase due to the change in the plasma potential distribution. Overall, the mass utilization efficiency of the water ion thruster increased from 3.3% to 7.9% and the discharge loss decreased from 280 to 150W/A at the antenna bias of 38V. The performance with water electron source was evaluated. The thrust force, specific impulse, thrust-to-power ratio, thrust efficiency were 190μN, 350s, 5.9μN/W, and 1.0%, respectively at the antenna bias of 25V.

Performance model of vacuum arc thruster with inductive energy storage circuit

A vacuum arc thruster is a type of micro-thruster based on pulsed ablative vacuum arc discharge. A simple inductive energy storage circuit in a vacuum arc thruster is particularly suitable for CubeSats because of its compact size and low cost. In practice, it is necessary to predict the thruster performance with the given design parameters. However, unlike the pulsed plasma thruster, there is a lack of integrated performance models for the vacuum arc thruster that cover all three processes of the circuit, ablation, and acceleration. In this study, a circuit model that accounts for the real characteristics of the components, such as the inductor, was developed. Discharge experiments were conducted to validate the proposed circuit model. The current-time curve and discharge energy were calculated and had relative errors of less than 15%. The performance model was obtained by combining the circuit model with the existing ablation and acceleration models. The input power and thrust-to-power ratio were predicted, and the results were compared to a flight test of the thruster prototype. The error in the input power was small, whereas the error in the thrust-to-power ratio was relatively large. This large error can be explained by the differences in the surface conditions.

Design and Analysis of a High Efficient Multi-Segment Redundancy Linear Motor

Background: The electromagnetic aircraft launch system helps the carrier aircraft accelerate take-off on a short distance on the flight deck. Linear motors are required to have high efficiency and thrust / volume ratio, as the actuator of the EMALS. In addition, linear motors were required to have the fluctuation of the thrust is as small as possible to prevent impact and overstress on the airborne equipment. The thrust performance and thrust fluctuation suppression of linear motors are the current research hotspots in the field of EMALS. Objective: This paper proposes a multi-segment LPMSM to improve the efficiency of the system. Methods: We design a decentralized primary structure based on the Autonomous Decentralized System (ADS) concept. Only the coils near the moving secondary are motivated during the launch process. The others could be released from being driven after the secondary slide over; this control will save energy. Results: A multi-segment linear motor is proposed to improve the efficiency of the EMALS. This motor consists of two sections - the one with iron core for the low speed process and the one without iron core for the high speed process. Conclusion: The normal direction component of the air gap magnetic flux density is much larger than that of the longitudinal component, and their values are 80-90% and 10-20% respectively. For the normal direction component, it is mainly affected by the third, seventh, and ninth harmonics. The accelerating and energy test results show that, in the accelerating process, the efficiency of the linear motor can be more than 90%, and the acceleration is stable.

Plume Composition Measurements of a High-Emission-Density Electrospray Thruster

An electrospray thruster designed for nanosatellites is introduced in this paper. The thruster has a compact size featuring a passive propellant transport method through the use of porous materials. Porous electrospray emitters with 1, 25, and 100 emission tips were fabricated using conventional machining techniques. The geometrical features of the emission tips were measured using a high-precision laser profilometer. Thruster prototypes fitted with different emitters were tested in a bipolar operation mode. The emission currents and voltage characteristics were studied, which demonstrated a wide range of operational voltage with relatively high emission currents. The maximum overall emission currents were and achieved using a 100-tip emitter at , and the maximum per-tip emission currents were and achieved using a 25-tip emitter at , respectively. Time-of-flight systems were used to characterize the mass and velocity distribution of the charged particles in the thruster plume, and the results suggest that the electrospray emission were highly ionic, containing mostly monomer ions and dimer ions. The thruster performance was conservatively estimated, with the thrust ranging from approximately 0.92 to , and the specific impulse ranging from 2733 to 5550 s. The maximum thrust per emission tip was approximately in a purely ionic emission regime, which is considerably high in this type of technology.

A Review of Electric Aircraft Drivetrain Motor Technology

This article reviews the power density and torque density of permanent magnet (PM) motor technology that is currently commercially developed for use in electric aircraft. The review shows that in order to meet the ambitious electrical motor power density goals recently set forth, the power density of a motor needs to more than double. The requirements imposed by the maximum propeller rotor tip speed, as well as the diameter constraints set by the aerodynamic requirements, will further limit the practical electric motor design sizing options. This article advocates for considering using magnetic gearing (MG) and using an integrated contra-rotating motor propeller co-design approach as a way of both increasing propulsive thrust performance and thereby increasing the aircraft’s overall power density.

Electric thruster mode-pruning strategies for trajectory-propulsion co-optimization

Application of idealized constant-specific-impulse, constant-thrust electric thruster performance models or curve-fitted polynomials is quite common for spacecraft trajectory design. However, incorporation of realistic performance models of multi-mode electric thrusters leads to notable challenges, and at the same time, offers unprecedented system-level optimization opportunities. In this paper, a framework is developed for co-optimization of 1) spacecraft trajectory, 2) operation modes of multi-mode propulsion systems, and 3) solar array size. The selection of the most optimal operation modes is in accordance to Pontryagin's minimum principle for solving a payload-mass-maximization problem. The novelty of the work lies in solving a mixed-integer trajectory optimization problem featuring user-defined constraints on the maximum number of operating modes used along the trajectory. Utility of the framework is demonstrated through optimization of a multi-year trajectory from Earth to comet 67P/Churyumov–Gerasimenko using an SPT-140 Hall thruster with 21 operating modes; the results are interesting and of significant practical utility.

Numerical simulation and experimental research of LRIT-30 radio frequency ion thruster

This study aimed at discussing the laws of the design parameters of a radio frequency (RF) ion thruster that influence the thruster performance, guide the performance optimization-oriented design of the thruster, and realize the high-accuracy continuous adjustment of the thrust performance. The key influencing parameters of the thrust performance were analyzed by the numerical simulation method. The influencing laws of RF parameters on key plasma parameters, as well as the thermal characteristics of the thruster under the rated parameters, were explored. Moreover, a LRIT-30 RF ion thruster was developed and subjected to the performance adjustment test. The research results demonstrated that the simulation model can describe the plasma parameter distribution of the discharge chamber and the thermal distribution of key components. The 3 cm RF ion thruster performs well under 2 MHz operational frequency. The RF power and gas flow rate are the key influencing factors of beam extraction and are appropriate for the accurate adjustment of parameters. The wide-range adjustment of thrust (0.5–2.3 mN) and specific impulse (869–2564 s) can be realized when the screen-grid voltage, decelerating screen voltage, RF power, and gas flow rate are 1500 V, −200 V, 40–65 W, and 0.4–0.8 SCCM, respectively. The performance indices of the proposed RF ion thruster are close to the international advanced level.

Studies of a modulated Hall thruster

A typical Hall thruster is powered from a DC power supply and operates with a constant discharge voltage. In operation, the discharge current is subject to strong low frequency oscillations (so-called breathing oscillations). Recent studies have shown that not only can these breathing oscillations be correlated with improved performance, but these oscillations can be induced and controlled by modulating the anode voltage. In this work, a systematic experimental study of the plasma flow in a modulated cylindrical Hall thruster was performed to characterize the effect of natural and modulated breathing oscillations on thruster performance. Measurements suggest that modulating the anode voltage in resonance with the natural breathing frequency does increase the thrust, but a corresponding phase alignment of discharge current and discharge voltage causes the efficiency gains to be insignificant. In addition, the outward shift of the acceleration region causes the plasma plume divergence to increase at the resonance condition and thereby, limit the thrust increase. Mechanisms underlying the relative phase between discharge current, ion current, and discharge voltage are investigated experimentally and corroborated with one-dimensional hybrid simulations of the thruster discharge.

Performance and plume characteristics of a Hall-effect thruster with asymmetrical gas supply

An analysis of the discharge behavior of a symmetrically designed Hall-effect thruster (HET) under an asymmetrical condition will aid the understanding of the working characteristics of HETs from different perspectives. Therefore, in this paper, with experimental methods, the asymmetrical gas distribution inside the channel of a HET was realized by supplying the gas asymmetrically, and changes in both the performance and plume parameters with the degree of asymmetry were measured. Both the discharge current and thrust were found to vary nonlinearly with the degree of asymmetry; moreover, the discharge mode also shifted, as observed from the variations in the discharge current and oscillation. Furthermore, the relationships between the plume ion current and energy distributions and the degree of asymmetry were determined and further used to explain the nonlinear change in the thrust. Overall, the findings of this study strengthen the understanding of the physics of HET discharge.

SU2 REAL GAS MODELS’ PERFORMANCE PREDICTIONS ON A COLD GAS THRUSTER

Cold gas propulsion systems are preferred, especially for the altitude and trajectory control for satellites, since the 1960s. Both depressurizing the propellant in the propellant tank throughout the mission and the expansion occurring in the divergent part of the thruster nozzle are the reasons for observing very low temperatures and pressure at the outlet section. We have decided to use an open-source compressible CFD tool, SU2, in order to predict the performance of the propulsion system in the early design phase. The scope of this study is the comparison of the results of the vacuum chamber performance tests and the outcomes of unsteady simulations with SU2 using different gas models, including ideal gas, van-der Waals, and Peng-Robinson models. Then, the accuracy and performance of these models are evaluated for the extremely low temperature and pressure conditions. Thruster performance tests have been conducted in the thermal vacuum chamber test campaign at the specially built testing facilities. In order to simulate nominal and low-temperature operation conditions, a propellant tank is thermally conditioned to the predetermined values, and performance test parameters are used as the input for the CFD simulations. The obtained results showed that the van der Waals gas model is the most appropriate gas model for both cases and provides the most realistic results in terms of performance parameters.