Feasibility of a New Ironless-stator Axial Flux Permanent Magnet Machine for Aircraft Electric Propulsion Application

With the development of aviation electrification,higher demands for electrical machines are put forward in aircraft electric propulsion systems.The aircraft electric propulsion requirements and propulsion motor features are analyzed in this paper.Comparing with conventional PM machines,ironless stator axial flux permanent magnet(AFPM)machine topologies with Litz wire windings allow designs with higher compactness,lightness and efficiency,which are suitable for high-frequency and high-power density applications.Based on the motor requirements and constraints of aircraft electric propulsion systems,this paper investigates a high-power 1 MW multi-stack ironless stator AFPM machine,which is composed of four 250kW modular motors by stacking in axial.The design guidelines and special attentions are presented,in term of electromagnetic,thermal,and mechanical performance for the high-frequency coils and Halbach-array PM rotor.Finally,an ironless stator AFPM motor is manufactured,tested and evaluated with the consideration of cost and processing cycle.The results show that the output power is up to 53.8kW with 95%efficiency at 9000r/min at this stage.The proposed ironless stator AFPM machine with oil immersed forced cooling proves to be a favorable candidate for application in electric aircraft as propulsion motors.

Kinetic Analysis of Vectored Electric Propulsion System for Stratosphere Airship

To enhance the controllability of stratosphere airship,a vectored electric propulsion system is used.By using the Lagrangian method,a kinetic model of the vectored electric propulsion system is established and validated through ground tests.The fake gyroscopic torque is first proposed,which the vector mechanism should overcome besides the inertial torque and the gravitational torque.The fake gyroscopic torque is caused by the difference between inertial moments about two principal inertial axes of the propeller in the rotating plane,appears only when the propeller is rotating and is proportional with the rotation speed.It is a sinusoidal pulse,with a frequency that is twice of the rotation speed.Considering the fake gyroscope torque pulse and aerodynamic efficiency,three blade propeller is recommended for the vectored propulsion system used for stratosphere airship.

Electric Thrusters Redundancy Configuration Strategy Study for All Electric Propulsion Platform Station-Keeping

Electric propulsion is used for all electric propulsion satellites to perform the orbit transfer,attitude control and station-keeping tasks. Generally electric propulsion subsystem contains 4 thrusters. But if one thruster fails in the beginning of satellite lifetime,other thrusters will undertake all the firing tasks. The firing time will be 2 to 3 times of thrusters without failure. Thus it may go beyond the allow ed lifetime of thruster. This paper puts forward two thruster redundancy configuration solutions with 6 thrusters to solve this problem. Two layout configurations and their corresponding station-keeping strategies are simulated and compared. The results show that the maximum firing time of both layout configurations can meet the lifetime limitation. This solution is a good reference for all electric propulsion satellites design.

Mechanism of capture section affecting an intake for atmosphere-breathing electric propulsion

Atmosphere-Breathing Electric Propulsion(ABEP)can compensate for lost momentum of spacecraft operating in Very Low Earth Orbit(VLEO)which has been widely concerned due to its excellent commercial potential.It is a key technology to improve the capture efficiency of intakes,which collect and compress the atmosphere for ABEP.In this paper,the mechanism of the capture section affecting capture efficiency is investigated by Test Particle Monte Carlo(TPMC)simulations with 3D intake models.The inner surface smoothness and average collision number are determined to be key factors affecting capture efficiency,and a negative effect growth model is accordingly established.When the inner surface smoothness is less than 0.2,the highest capture efficiency and its corresponding average collision number interval are independent of the capture section’s geometry and its mesh size.When the inner surface smoothness is higher than 0.2,the capture efficiency will decrease by installing any capture section.Based on the present results,the manufacturing process and material selection are suggested to be prioritized during the intake geometry design in engineering projects.Then,the highest capture efficiency can be achieved by adjusting the length and mesh size of the capture section.

Experimental research on aero-propulsion coupling characteristics of a distributed electric propulsion aircraft

Distributed Electric Propulsion(DEP)aircraft use multiple electric motors to drive the propulsors,which gives potential benefits to aerodynamic-propulsion interaction.To investigate and quantify the aerodynamic-propulsion interaction effect of the wing section,we built a DEP demonstrator with 24"high-lift"Electric Ducted Fans(EDFs)distributed along the wing’s trailing edge.This paper explores and compares the aero-propulsion coupling characteristics under various upstream speed,throttle,and EDF mounting surface deflection angles using a series of wind tunnel tests.We compare various lift-augmentation power conditions to the clean configuration without propulsion unit under the experiment condition of 15-25 m/s freestream flow and angles of attack from-4°to 16°.The comparison of computational results to the experimental results verifies the effectiveness of the computational fluid dynamic analysis method and the modeling method for the DEP configuration.The results show that the EDFs can produce significant lift increment and drag reduction simultaneously,which is accordant with the potential benefit of Boundary Layer Ingestion(BLI)at low airspeed.

Aerodynamic design of tractor propeller for high-performance distributed electric propulsion aircraft

Aiming to maximize the aerodynamic performance of the Distributed Electric Propulsion(DEP)aircraft,a hybrid design framework which focuses on the aerodynamic performance of the propeller/wing integration has been developed and validated numerically.Variable-fidelity modelling for propeller aerodynamics has been used to achieve computational efficiency with reasonable accuracy.By optimizing the aerodynamic loading distributions on the tractor propeller disk,the induced slipstream is redistributed into a form that is beneficial for the wing downstream,based on which the propeller blade geometry is generated through a rapid inversed design procedure.As compared with the Minimum Induced Loss(MIL)propeller at a specified thrust level,significant improvements of both the lift-to-drag ratio of the wing and the propeller/wing integrated aerodynamic efficiency is achieved,which shows great promise to deliver aerodynamic benefits for the wing within the propeller slipstream without any additional devices.

Aerodynamic performance of distributed electric propulsion with wing interaction

Distributed electric propulsion(DEP)uses multiple propellers driven by motors distributed along the leading edge of the wing to produce beneficial aerodynamic interactions.However,the wing will be in the sliding flow of the propeller and the lift and drag characteristics of the wing will change accordingly.The performance of the propeller will also be affected by the wing in its rear.In this paper,combined with wind tunnel tests,the low Reynolds aerodynamic properties of multiple DEP structures are numerically simulated by solving the Reynolds averaged Navier-Stokes(RANS)equation of multiple reference frames(MRF)or slip grid technology.The results demonstrate that the lift and drag of DEP increase in all cases,with the magnitude depending on the angle of attack(AOA)and the relative positions of propellers and wing.When the AOA is less than 16°(stall AOA),the change of lift is not affected by it.By contrast,when the AOA is greater than 16°the L/D(lift-to-drag ratio)of the DEP system increases significantly.This is because the propeller slipstream delays laminar flow separation and increases the stall AOA.At the same time,the inflow and the downwash effect,which is generated on both sides of the rotating shaft,result in the actual AOA of the wing being greater than the free flow AOA with a fluctuation distribution of the lift coefficient along the span.Also,for the propeller in the DEP,the blocking effect of the wing and the vortex of the trailing edge of the wing result in a significant increase in thrust.

Dynamic neighborhood genetic learning particle swarm optimization for high-power-density electric propulsion motor

To maximize the power density of the electric propulsion motor in aerospace application,this paper proposes a novel Dynamic Neighborhood Genetic Learning Particle Swarm Optimization(DNGL-PSO)for the motor design,which can deal with the insufficient population diversity and non-global optimal solution issues.The DNGL-PSO framework is composed of the dynamic neighborhood module and the particle update module.To improve the population diversity,the dynamic neighborhood strategy is first proposed,which combines the local neighborhood exemplar generation mechanism and the shuffling mechanism.The local neighborhood exemplar generation mechanism enlarges the search range of the algorithm in the solution space,thus obtaining highquality exemplars.Meanwhile,when the global optimal solution cannot update its fitness value,the shuffling mechanism module is triggered to dynamically change the local neighborhood members.The roulette wheel selection operator is introduced into the shuffling mechanism to ensure that particles with larger fitness value are selected with a higher probability and remain in the local neighborhood.Then,the global learning based particle update approach is proposed,which can achieve a good balance between the expansion of the search range in the early stage and the acceleration of local convergence in the later stage.Finally,the optimization design of the electric propulsion motor is conducted to verify the effectiveness of the proposed DNGL-PSO.The simulation results show that the proposed DNGL-PSO has excellent adaptability,optimization efficiency and global optimization capability,while the optimized electric propulsion motor has a high power density of 5.207 kW/kg with the efficiency of 96.12%.

Relative position fitting based control strategy for thruster point assembly mechanism of electric propulsion

Purpose-Thruster point assembly mechanism(TPAM)of the electric propulsion system allows to adjust the thrust vector,sothat the thrust vectorisdirectedtothesatellitecenterof gravity(COG)during the satellite on-orbit working period.In this way the impact of disturbance torque caused by deviation of the thrust vector from the satellite COG during thruster ignition can be decreased.Therefore,the control accuracy of satellite is influenced directly by the control accuracy of TPAM.On the other hand,the on-orbit application of TPAM is restricted to the on-orbit computer resource.Therefore,the purpose of this paper is to design a control strategy for TPAM,and the strategy should not only be able to control the TPAM precisely but also be easily implemented by the on-board computer.Design/methodology/approach-First,the structure and work principle of TPAM are discussed,and the mathematical model based on D-H coordinate system is built for it.Then the fitting methods are utilized to design the control strategy of TPAM.Absolute position fitting-based control strategy and relative position fitting-based control strategy are designed,and the least squares algorithm is introduced for parameter selection.Findings-Simulations and tests are provided for the TPAM.Compared with the state-of-the-art PD controller,the proposed control strategy shows smaller overshoot and more simple realization.The experiment results are matched with the simulation results and both the experiment and simulation results show the validity of the proposed control strategies.Practical implications-The designed control strategies can be used for the TPAM of some satellite’s electric propulsion system.Originality/value-The mathematical model of the TPAM based on D-H coordinate system is given.The absolute position fitting-based control strategy and relative position fitting-based control strategy are proposed.Compared with existing methods,the two control strategies have more simple structure and smaller amount of computations.Furthermore,the relative position fitting-based control strategy achieves high precision with simple structure.

A three-dimensional numerical study on the effect of geometric asymmetry on arcjet thruster performance

In an arcjet thruster,the cathode and constrictor degrade with time,and the electrical arc discharge may become unsymmetrical.In this work,a three-dimensional numerical model of a hydrogen plasma arcjet is developed and validated to study the effect of unsymmetrical electric arc discharge on thruster performance.The unsymmetrical arc discharge is realized by introducing a radial shift of the cathode so that the cathode tip offset is 80μm(25%of the constrictor radius).Simulations are conducted for both axially centered cathode(coaxial)and off-centered cathode(non-coaxial)configurations with identical propellant flow rates and input current.Simulations show asymmetrical arc discharge in the non-coaxial cathode configuration,resulting in azimuthally asymmetric Joule heating,species concentrations,and velocity field.This asymmetry continues as the plasma expands in the divergent section of the nozzle.Temperature,species concentrations,and axial velocity exhibit asymmetric radial distribution at the nozzle exit.The computed Joule heating was found to reduce with cathode shift,and consequently,the thrust and specific impulse of the thruster was decreased by about 6.6%.In the case of the non-coaxial cathode,geometric asymmetry also induces a small side thrust.

Presheath formation and area design limit satellite-based Langmuir probes

In this article,the effect of the finite conductive surface area of a satellite on the use of satellite-based Langmuir probes is reviewed in light of the basic theory of asymmetric double Langmuir probes(ADLPs).Recent theoretical and experimental studies have discussed electron sheath/presheath formation and the electron Bohm criterion along with their implications for satellite-based Langmuir probes.The effects predicted by the latest theory of the electron Bohm criterion were not experimentally observed and the experimental results remain supportive of a critical area ratio(A_(L)/A_(S))_(crit)=(m_(i)/(2.3m_(e)))^(1/2)between the probe area A_(S)and the satellite area A_L as conventionally believed.A satellite-based Langmuir probe must satisfy this criterion to physically act as a single Langmuir probe.However,experimental investigations also found that high-energy electrons adversely affect(A_(L)/A_(S))_(crit)and a Langmuir probe's signal quality by giving additional electron current to A_(L).Based on these results,a number of limitations of the maximum probe area are derived when designing satellite-based Langmuir probes,with consideration of both the aim of the satellite and the plasma where the satellite-based probe works.These proposed measures are expected to only partially alleviate the effect of the inadequate satellite surface area on the application of satellite-based Langmuir probes.Using a larger satellite to carry a Langmuir probe remains the most viable means to obtain precise space plasma parameters.

Performance of a 4 cm iodine-fueled radio frequency ion thruster

The performance of an iodine radio ion thruster with a 4 cm diameter(IRIT4)was studied experimentally in this paper.Regulation of the mass flow rates of the iodine propellant is achieved by using a temperature control method of the iodine reservoir.Performance of the thruster using iodine as propellants is obtained at different total thruster powers of 40.6–128.8 W,different grid voltages of 1000–1800 V and the iodine flow rate of 100μgs^(-1).Results show that thrust and specific impulse increase approximately linearly with the increasing total thruster power and the screen grid voltage.The thrust of 2.32 mN and the specific impulse of 2361 s are obtained at the nominal total thruster power of 95.8 W and the screen grid voltage of 1800 V.It is also indicated that performance of the iodine propellant is comparable to that of the xenon propellant;and a difference between them is that the iodine thrust is slightly higher than xenon when the total thruster power is more than 62 W.At the nominal 95.8 W total thruster power,the thrust values of them are 2.32 m N and 2.15 mN respectively,and the thrust-to-power ratios of them are 24.2 mN kW^(-1)and 23.5 mN kW^(-1),respectively.

Cathode erosion site distributions in an applied-field magnetoplasmadynamic thruster

Erosion can influence cathode life,and is thus considered to be one of the main factors limiting the application of applied-field magnetoplasmadynamic thrusters.In this paper,erosion sites on graphite cathodes are studied so as to identify the influence of applied magnetic field and the ratio of propellant mass flow rate supplied from cathode and anode.The experiment results show that the application of applied magnetic field can significantly reduce the erosion rate of the cathode compared to that without magnetic field.The erosion sites on the cathode vary with the relative position of the convergent-divergent magnetic field,and are mainly distributed in the divergent part of the field.The erosion sites on the cathodes are found to be related to the propellant supply.The decreasing anode mass flow rate enlarges the range of erosion.These results are much helpful for the analysis of cathode erosion site location since they provide evidences of erosion mechanisms and point out the directions for further research.

The plume diagnostics of 30 cm ion thruster with an advanced plasma diagnostics system

In order to achieve a better understanding of plume characteristics of LIPS-300 ion thruster,the beam current density,ion energy and electron number density of LIPS-300 ion thruster plume are studied with an Advanced Plasma Diagnostics System(APDS)which allows for simultaneous in situ measurements of various properties characterizing ion thruster,such as plasma density,plasma potential,plasma temperature and ion beam current densities,ion energy distribution and so on.The results show that the beam current density distribution has a double‘wing'shape.The high energy ions were found in small scan angle,while low energy ions were found in greater scan angle.Electron number density has a similar shape with the beam current density distribution.

大功率霍尔电推进研究进展

High power Hall electric propulsion technology is a very competitive electric propulsion technology for future large space missions such as large GEO satellites,manned space programs,deep space explorations,cargo ships,space tugs.Based on the experience of more than 20 years in research and development of Hall electric propulsion,the Shanghai Institute of Space Propulsion(SISP)has developed 3 high power Hall thrusters,i.e.,the 10 k W class HET-500,20 k W class HET-1000,and 50 k W class HET-3000.This paper presents the development status of the high power(≥10 k W)Hall electric propulsion at SISP,including tests of 3 high power Hall thrusters in the power range from 10 k W to 50 k W,the qualification of a single string of a 10 k W Hall electric propulsion system,and the study of a cluster of two 1.35 k W HET-80 Hall thrusters to understand the technical issues related to multi-thruster high power electric propulsion systems.

An atmosphere-breathing propulsion system using inductively coupled plasma source

CubeSats have attracted more research interest recently due to their lower cost and shorter production time.A promising technology for CubeSat application is atmosphere-breathing electric propulsion,which can capture the atmospheric particles as propulsion propellant to maintain longterm mission at very low Earth orbit.This paper designs an atmosphere-breathing electric propulsion system for a 3 U CubeSat,which consists of an intake device and an electric thruster based on the inductively coupled plasma.The capture performance of intake device is optimized considering both particles capture efficiency and compression ratio.The plasma source is also analyzed by experiment and simulation.Then,the thrust performance is also estimated when taking into account the intake performance.The results show that it is feasible to use atmosphere-breathing electric propulsion technology for CubeSats to compensate for aerodynamic drag at lower Earth orbit.

Machine learning-based method to adjust electron anomalous conductivity profile to experimentally measured operating parameters of Hall thruster

The problem of determining the electron anomalous conductivity profile in a Hall thruster,when its operating parameters are known from the experiment,is considered.To solve the problem,we propose varying the parametrically set anomalous conductivity profile until the calculated operating parameters match the experimentally measured ones in the best way.The axial 1D3V hybrid model was used to calculate the operating parameters with parametrically set conductivity.Variation of the conductivity profile was performed using Bayesian optimization with a Gaussian process(machine learning method),which can resolve all local minima,even for noisy functions.The calculated solution corresponding to the measured operating parameters of a Hall thruster in the best way proved to be unique for the studied operating modes of KM-88.The local plasma parameters were calculated and compared to the measured ones for four different operating modes.The results show the qualitative agreement.An agreement between calculated and measured local parameters can be improved with a more accurate model of plasma-wall interaction.

Numerical simulation of hydrogen arcjet thruster with coupled sheath model

In the present work,a complete 2D chemical and thermal non-equilibrium numerical model coupled with a relatively simple sheath model is developed for hydrogen arcjet thruster.Conduction heat transfer in the anode wall is also included in the model.The operating voltages predicted by the model are compared with those in the literature and are found to be in close agreement.Power distributions for the various operating conditions are obtained,anode radiation loss primarily determines the thruster efficiency.Higher thruster efficiency was found to be associated with longer arc length.At cathode ion diffusion contribution dominates except at low input current where thermo-field electron current is dominant.

Simulation of liquid meniscus formation in the ionic liquid electrospray process

Though tremendous efforts have been made to investigate electrospray,some aspects,such as the evolution of the menisci on the micropores of porous emitter tips and the transient response of the meniscus during the polarity alternation,need to be further understood.This paper presents a computation fluid dynamics(CFD)model to describe the meniscus formation in the ionic liquid electrospray process.The CFD model,based on the Taylor–Melcher leaky dielectric fluid theory and the volume of fluid(VOF)method,is validated by experiments.The evolution of the meniscus on the basis of a micropore is presented using two typical ionic liquids,EMI-BF_(4) and EMI-Im.The influences of the pore size,flow rate and applied voltage on the formation of the meniscus have been studied.Results show that a larger pore is more likely to start emission,and the time consumed for liquid meniscus formation decreases with increasing applied voltage and flow rate.Further,it is found that alternation of polarity does not destroy the structure of the meniscus but retards the formation process,and a faster polarity alteration leads to a shorter delay in meniscus formation time.

Investigation on plasma structure evolution and discharge characteristics of a single-stage planar-pulsed-inductive accelerator under ambient fill condition

The physical process of a single-stage planar-pulsed-inductive accelerator is investigated.Measurements include the waveforms of circuit current,capacitor voltage,plasma radiation intensity,and temporal plasma structure photos captured by a high-speed camera.Experiments are conducted under static ambient fill condition using argon as propellant.Varied values of capacitor voltage and gas pressure are compared.Further discussions quantify the EM interaction between circuit and plasma,as well as their energy deposition and current sheet acceleration.Based on the results of experiments,physical mechanisms of the initial ionization phase and the following acceleration phase are analyzed theoretically.