Electric Propulsion Concepts Research Articles

Distributed Electric Propulsion and Flight Control Concept to Meet EASA SC-VTOL-01 10‐9 Catastrophic Failure Criteria

The objective of the current work was to develop an all-electric distributed propulsion and flight control (DPFC) architecture that will have no more than 10‐9 catastrophic failures per flight hour (pfh). The DPFC architecture was broken into four system design teams: (i) electrical power and distribution system, (ii) drive and power system, (iii) thermal management system, and (iv) flight control system. System designs were updated and iterated upon, working with reliability and safety analysis teams to develop compliant designs. The firm designs were reflected in a preliminary system safety assessment (PSSA) for initial verification of compliance. Additionally, design and analysis excursions are presented in which aircraft attributes were modified to investigate sensitivities to propulsion type, number of rotors, and control schemes. Excursion results found that all aircraft evaluated likely have paths to comply with the stringent, probabilistic catastrophic failure criteria. However, stability and control models showed large power transients that must be addressed, and PSSA results show that future work is needed in single load path structures, high voltage power storage and distribution, and motor/rotor overspeed protection.

Concept Evaluation of Radical Short–Medium-Range Aircraft with Turbo-Electric Propulsion

Ambitious targets for the coming decades have been set for further reductions in aviation greenhouse gas emissions. Hybrid electric propulsion (HEP) concepts offer potential for the mitigation of these aviation emissions. To investigate this potential in an adequate level of detail, the European research project IMOTHEP (Investigation and Maturation of Technologies for Hybrid Electric Propulsion) explores key technologies for HEP in close relation with developments of aircraft missions and configuration. This paper presents conceptual-level design investigations on radical HEP aircraft configurations for short–medium-range (SMR) missions. In particular, a blended-wing-body (BWB) configuration with a turbo-electric powertrain and distributed electric propulsion is investigated using NLR’s aircraft evaluation tool MASS. For the aircraft and powertrain design, representative top-level aircraft requirements have been defined in IMOTHEP, and the reference aircraft for the assessment of potential benefits is based on the Airbus A320neo aircraft. The models and data developed in IMOTHEP and presented in this paper show that the turbo-electric BWB configuration has potential for reduced fuel consumption in comparison to the reference aircraft. But in comparison to advanced turbofan-powered BWB configurations, which have the same benefits of the BWB airframe and advanced technology assumptions, this potential is limited.

Influence of applied magnetic field in an air-breathing microwave plasma cathode

The air-breathing electric propulsion concept refers to a spacecraft in very-low Earth orbit (VLEO) ingesting upper atmospheric air as propellant for an electric thruster. This compensates atmospheric drag and allows the spacecraft to maintain its orbital altitude, removing the need for on-board propellant storage and allowing an extended mission duration which is not limited by propellant exhaustion. There is a need for development of a robust, high current density and long life cathode (or neutralizer) for air-breathing electrostatic thrusters as conventional thermionic hollow cathodes are susceptible to oxygen poisoning. An Air-breathing Microwave Plasma CAThode is proposed to overcome this issue through the use of a microwave plasma discharge, producing an extracted current in the order of 1 A with 0.1 mg s−1 of air. In this paper, the effect of varying magnetic-field strength and topology is investigated by using an electromagnet coil, which reveals a significantly different behaviour for air compared to xenon. The extracted current with xenon increases by 3.9 times from the zero-field value up to a peak around 150 mT magnetic-field strength at the antenna, whereas an applied field does not increase the extracted current with air at nominal conditions. A non-zero magnetic-field with air is however beneficial for current extraction at reduced neutral densities. A distinct increase in extracted current is identified at low bias voltages with air for a field strength of around 50 mT at the internal microwave antenna, consistent across varying field topologies. The effect of a lowered magnetic-field strength in the orifice region is investigated through the use of a secondary coil, resulting in an extracted current increase of 25% for a relaxation from 6 mT to 1 mT, and demonstrating the beneficial impact of a locally reduced field strength on electron extraction.

Advanced Electric Propulsion Concepts for Fast Missions to the Outer Solar System and Beyond

Electric Propulsion (EP) comprises all types of space propulsion in which a certain amount of propellant is ionized and then accelerated by electric or magnetic fields, or both. This propulsion technology allows for much higher specific impulses (Isp > 2,000 s) than conventional chemical propulsion (Isp < 500 s), resulting in a major reduction of the propellant mass or a considerably higher final speed for a certain space mission. Hence, EP can enable very challenging space missions as DAWN has clearly shown. Furthermore, an EP system coupled with an advanced nuclear reactor could enable fast manned missions to Mars (one-way travel times less than 4 months). This propulsion technology can be scaled up to even higher specific impulses (Isp > 5,000s). However, the power needed for the same thrust is also increasing. An Oberth maneuver performed very close to the Sun could provide the additional power to a high-Isp Solar Electric Propulsion (SEP) system in order to reach the needed delta-v for challenging interstellar precursor missions. A breakthrough in power source specific mass is needed in order to enable missions with ultra-high specific impulses (> 10,000 s); this breakthrough could be realized having the power source not on board, as with Laser-powered Electric Propulsion (LEP), where the needed power is beamed to the spacecraft from an external laser source. In this case the on-board power source is limited to a light-weight photovoltaic receiver/converter. The development of ultra-high Isp ion thrusters powered by an external laser source could enable the most challenging interstellar precursor missions up to the Oort Cloud and beyond. This paper gives an update on the status of these advanced propulsion concepts, and provides some examples of interstellar precursor missions enabled by advanced EP systems which could be launched before 2040. Keywords: Interstellar, Electric Propulsion, Oberth, Laser

Development of a lanthanum hexaboride hollow cathode for a magnetic octupole thruster

Electric propulsion has undergone significant advancements, leading to increased usage in satellite constellation management and orbital transfer vehicles. This study describes the design and development of a lanthanum hexaboride (LaB6) hollow cathode for the Magnetic Octupole Plasma Thruster (MOPT), a novel electric propulsion concept for future space missions. Its operational characteristics are also provided. To gain insights into the coupling behavior between the cathode plume and the octupole magnetic field, the hollow cathode was operated with and without an octupole magnet over a range of discharge currents and flow rates. The results indicate that the discharge voltage was anti-correlated with the discharge current and flow rate. Notably, the octupole magnetic field observed reduced discharge voltage oscillations in both high- and low-discharge-current regimes. Nonetheless, the existence of the magnetic field was associated with the amplification of ionization-like instabilities in the high-flow-rate regime. Finally, the paper also delves into the plasma behavior under the influence of the octupole magnetic field, signifying an advancement in developing LaB6 hollow cathodes for their potential use in space-based MOPT applications.

Certification-Constrained Vertical Tail Sizing and Power Split Optimization for Distributed Electric Propulsion Aircraft

Distributed electric propulsion (DEP) concepts leverage the synergistic benefits of aeropropulsive coupling and improve the overall flight efficiency by placing propulsors at optimal locations, such as the wingtip. However, placing propulsors far away from the aircraft centerline may bring aircraft difficulties in complying with critical-engine-inoperative flight certification rules. To investigate the impact of certification requirements on the DEP aircraft design, this paper takes NASA’s PEGASUS concept as a use case and presents a study of certification-constrained vertical tail sizing and propulsive power split optimization. For comparison, a vertical tail retrofit study is also conducted on the ATR 42-500 aircraft using the same design space, to gain insights into the different impacts of certification constraints on the design process between conventional and unconventional aircraft. In each study, a design space exploration is performed, including a feasibility test and a multi-objective optimization. The feasibility test shows that the certification constraints yield a much smaller feasible design space for the DEP concept when compared to the conventional aircraft. Despite a much smaller feasible design space, the constrained Pareto optima of the PEGASUS still exhibit lower cruise drag coefficient, fuel burn, and operational energy cost when compared to those of the ATR 42.

On the prediction of noise generated by urban air mobility (UAM) vehicles. II. Implementation of the Farassat F1A formulation into a modern surface-vorticity panel solver

This study focuses on the integration of established acoustic prediction techniques directly into a surface-vorticity solver. The main objective is to enhance an aircraft designer's ability to characterize the acoustic signatures generated by urban air mobility (UAM) vehicles, in general, and distributed electric propulsion (DEP) concepts, in particular, including unmanned aerial vehicles. Our solver consists of a reliable, surface-vorticity panel code that incorporates viscous boundary-layer corrections. Thus, it offers a computationally efficient commercial tool for conceptual design and preliminary aerodynamic analysis. By implementing the Farassat F1A acoustics formulation directly into the solver, a new intuitive capability is achieved, which is both conversive with modern engineering tools and efficient in setup and speed of execution. In addition to its application to the X-57 high-lift propeller and the Revolutionary Vertical Lift Technology Tiltwing electric Vertical Take-Off and Landing (eVTOL) vehicle by the National Aeronautics and Space Administration, this capability is systematically demonstrated using three particular case studies. These consist of both single- and six-propeller Joby S4 eVTOL as well as a small eight-propeller Kittyhawk KH-H1 DEP vehicle. Although the details of this tool and underlying equations are showcased in this article, the acoustic metrics that can be effectively used to characterize the noise level generated by a UAM in flight are described in a companion article. By embedding this assortment of insightful metrics into a simple and user-friendly flow solver, a much improved flow-acoustic analysis capability is thereby provided to support the design of future aircraft.

Effects of magnetic field strength on the microwave discharge cusped field thruster

The microwave discharge cusped field thruster is a novel electric propulsion concept developed for the space-borne Gravitational Wave detection mission. It is an effective combination of the cusped magnetic field and the ECR microwave plasma source, which provide the thruster strong confinement for the plasma and the ability to operate on low mass flow rate conditions. To meet the requirements on the high accuracy, low noise thrust over a wide range, a series of studies are carried out on the microwave discharge cusped field thruster. Since the magnetic field strength is one of the critical factors with a significant impact on the thruster, a plume plasma diagnosis is performed to analyze the discharge performance and plasma characteristics with different magnetic field strength conditions. A Faraday probe, a retarding potential analyzer, and an emissive probe are employed to characterize the plume. The results show that the variation of the magnetic field strength performs strong effects on the plume structure, the ion acceleration, and the electron conduction process. As the magnetic field strength weakens, the single-peaked plume profile transforms to a hollow cone structure, which is attributed to the conversion of acceleration mechanism. The thruster prototype performs a continuous thrust of 2–26.9 μN and a highest specific impulse of 248s with superior reproducibility in a low mass flow rate of 0.1sccm.

Experimental investigation of electric propulsion systems using C12A7 electride hollow cathodes

The development and experimental investigation of two low-power electric propulsion concepts using compact heaterless C12A7 electride (C12A7:e-) hollow cathodes is presented. The first concept represents an electrothermal thruster, in which a cathode discharge is used to heat a gas that is subsequently accelerated in a nozzle-shaped anode. The second propulsion system is an attempt to develop a sub-500 W magnetoplasmadynamic thruster (MPDT) that uses a rectangular discharge channel that allows to increase the applied magnetic field and thus lower the necessary discharge current. Extensive parameter studies with both concepts were conducted, and the thrust and discharge properties of different geometric and operational configurations were determined. This work is a follow-up publication of a previous paper (Gondol and Tajmar in CEAS Space J 14:65–77, 2021).

Simulation-guided engineering of an air-breathing electric propulsion concept

Simulation-guided engineering of an air-breathing electric propulsion concept

Global plasma modeling of a magnetized high-frequency plasma source in low-pressure nitrogen and oxygen for air-breathing electric propulsion applications

This work presents a global plasma model of a gridded air-breathing electric propulsion concept based on magnetized high-frequency plasma operating in the pressure range of 10−3 Pa to 1 Pa. We illustrate that the global plasma model reproduces the experimental measurements of the extracted current over two orders of magnitude in pressure. Consequently, we use the model to investigate the theoretical scalability of the plasma source, finding that the plasma source performance scales reasonably well with the average absorbed power per molecule, even though this scaling factor has its limits. The global model presented in this work is a model of a specific laboratory device and, in future, it can be adapted to very low Earth orbit conditions by adjusting the boundary conditions. The model was implemented using PlasmaSolve p3s-globalmodel software and the configuration file containing all the equations is provided to the community as supplementary material.

Design of a hollow cathode thruster: concepts, parameter study and initial test results

Two electric propulsion concepts have been developed at Technische Universität Dresden as spin-off devices of regular hollow cathodes and initial testing has been conducted. Both devices represent millinewton thrusters that take advantage of thermionic electron emission using the low work function materials C12A7, LaB6, and thoriated tungsten in different design configurations. The first concept represents an electrothermal thruster which generates thrust by expanding and accelerating a heated propellant in a nozzle. Initial thrust measurement tests were carried out which showed thrust levels well above cold gas thrust, but low thrust efficiencies. The influence of different geometric parameters on the discharge properties and the performance is investigated and presented. The second thruster concept is a novel electromagnetic device in which charge carriers in a plasma discharge are accelerated by an applied magnetic field that is orthogonally oriented to the discharge current. Initial tests with C12A7 were not successful, but the functionality of the concept was shown by thrust measurements using a thoriated tungsten wire as an electron emitter.

Numerical analysis of pressure distribution in a low aspect ratio wing with distributed electric propulsion

The distributed electric propulsion concept expands the boundaries of aircraft designs, therefore, the study of aerodynamic effects on the propulsion/wing interaction is of great importance. This paper focuses on analyzing the pressure distribution in a low aspect ratio wing with distributed electric propulsion concept, considering the propulsion upstream of the wing. Using computational fluid dynamics, the wing pressure was numerically analyzed for different combinations of propeller rotation direction. Numerical methods were based on RANS equations, including the turbulence model, and actuator disk model based on “blade element theory,” which was used to represent the propeller. Numerical validation presented good representation of the pressure distributions obtained by the numerical method when compared with experimental data, especially in wing regions affected by the induced flow through the actuator disk. Numerical analysis has shown significant interplay between propeller rotation direction and pressure distribution on the wing, which in turn can influence stall initiation regions due to blade-induced flow.

Study on the large-range thrust throttling ability in a multi-cusped field thruster

The multi-cusped field thruster represents a type of grid-less electric propulsion concept with a multi-stage magnetic field topology. In this study, a two-stage multi-cusped field thruster is designed. It has a continuous thrust throttle range from 0.37 mN to 19.2 mN with the change of anode flow rate (from 3 to 10 sccm) and anode voltage (from 300 to 1100 V). The performance characteristics show that the thruster is an excellent candidate to fulfill the requirements for space vehicle drag compensation. The large-range thrust throttle-ability is further studied by plasma diagnostic methods. The test results show that the change of anode flow rate mainly affects the propellant ionization process in discharge channel, its influence on acceleration process is less obvious. Besides, the change of anode voltage mainly affects the ion acceleration process near exit, its influence on propellant ionization process is small. These phenomena indicate that there is a weak coupling relationship between ionization region and acceleration region, which becomes an inner decisive factor for its large-range thrust throttle-ability.

Simulation of a High Fidelity Turboshaft Engine-Alternator Model for Turboelectric Propulsion System Design and Applications

Abstract The term sustainability became a popular subject both in the automotive industries and in the aerospace industries. Increasing and threatening environmental pollution problems and reduction in limited fuel sources are motivating either industries and academicians to develop alternative power systems to sustain more healthier and economical life in the long term. One innovation that has been researched in the automotive industry is all electric and hybrid electric propulsion concepts. These concepts have also been proposed as alternative solutions for aviation. These novel propulsion technologies are composed of a gas turbine/internal combustion engine structure (necessary for hybrid electric and turboelectric propulsion systems) and/or energy storage components (battery, fuel cell and so on.) with multiple electric motors respectively. In this paper, simulation of a high fidelity turboshaft engine-alternator model for turboelectric propulsion system is derived. To develop an aero-thermal engine model, GE T700 turboshaft engine data is used and constructed model is connected to an alternator model on MATLAB/Simulink environment. Open-loop simulations are carried out and satisfactory results are obtained.Simulation results are compared to the real engine design point data. Results show that there are acceptable differences between the simulation results and the real engine data. The power balances between compressor - high pressure turbine and power turbine – alternator are proven in the mathematical model. It is expected that the proposed model can be easily extended to power system design and power management studies in turboelectric propulsion systems and also in other suitable novel propulsion systems.

Benefit evaluation of hybrid electric propulsion concepts for CS-23 aircraft

The importance of small passenger aircraft design fitting up to nine passengers and the evaluation of associated technologies have been recently increasing. This is related to the upcoming interest in on demand air mobility concepts not only for intra- or inter-urban transport but also for thin-haul routes ranging from 150 km to approximately 500–800 km. Such concepts seem feasible as key enabling technologies (e.g., flight automation and battery technology) are likely to be mature enough to enable these concepts in the foreseeable future. However, until battery specific energy in particular surpasses the threshold of approximately 400 Wh/kg, hybrid electric propulsion concepts could serve as an interim solution. Therefore, this paper deals with the question of how hybrid electric concepts score compared to conventional piston engine aircraft and which concept promises the most benefits. This includes consideration of propulsion-airframe integration benefits of electric engines through distributed electric propulsion (DEP). Results show that the series hybrid electric concept is superior to a parallel setup if at least a 15% higher cruise lift-to-drag ratio can be achieved due to DEP (30–50% increase likely). Despite higher weight, variable operating costs can be reduced by 15–35% with application of series hybrid electric propulsion concepts.

Electric Propulsion Concepts for an Inverted Joined Wing Airplane Demonstrator

One of the airplane design concepts that potentially allows for significantly increased efficiency, but has not yet been investigated thoroughly, is the inverted joined wing configuration, where the upper wing is positioned in front of the lower one. We performed wind tunnel and flight testing of a demonstrator of this concept, first by applying electrical propulsion to simplify wind tunnel testing, and then the same electrical-propulsion demonstrator performed several flights. As the chosen propulsion method proved to be too cumbersome for an intensive flight campaign and significant loss of battery performance was also observed, the electrical propulsion was then replaced by internal combustion propulsion in the second phase, involving longer-duration flight testing. Next we identified and analyzed two potentially beneficial modifications to the design tested: one involved shifting the center of gravity towards the aft, the other involved modifying the thrust vector position, both with the assumption that electric motors can be applied for propulsion. On this basis, the paper finishes with some conclusions concerning a new concept of electrical propulsion for an inverted joined wing design, combining two ideas: hybridization and distribution along the aft wing leading edge.

Modelling of a radio frequency plasma bridge neutralizer (RFPBN)

A performance model of a radio frequency plasma bridge neutralizer was developed to calculate the electrical parameters and optimize the neutralizer design. Minimization of power losses and gas consumption, and a maximization of the neutralizer lifetime and the reliability of the system are requirements of all electric propulsion concepts and strongly determine their future application. The requirements of the neutralizer depend on mission profiles.

Methods for simulation and analysis of hybrid electric propulsion systems

Today, research performed for new aircraft propulsion concepts is driven by the requirement of achieving significant emission reductions to meet the environmental objectives of future air traffic. A current trend visible in the aviation industry shows the attempt to reduce inflight emissions as well as overall energy consumption of conventional combustion engines through electrification via electrical energy sources. With the growing interest in novel hybrid electric propulsion concepts, in the same way, a demand for conceptual design and performance simulation methods in combination with analysis capabilities rises. Based on primarily introduced and discussed hybrid electric propulsion systems, this paper presents a set of aircraft propulsion system simulation (APSS) methods which allow for an integrated simulation and consistent analysis. Particularly, methods for the modelling of electric motor and battery systems as implemented in APSS are described in detail.

Study on Helicon Plasma Lissajous Acceleration for Electrodeless Electric Propulsion

(Received June 27th, 2011) In order to realize a long lifetime of an electric propulsion system, we have been investigating various electrodeless electric propulsion concepts utilizing a helicon plasma source. In one of our concepts, helicon plasma is electromagnetically accelerated using a rotating electric field in the presence of a diverging static magnetic field. This acceleration concept is called the Lissajous acceleration. Plasma acceleration experiments have been conducted and plasma acceleration was evaluated using a Mach probe. Although the experiments showed some features of the electromagnetic acceleration, most increment of the plasma velocity is caused by the increment of the electron temperature. The thrust (4.95 μN) did not reach feasible values for real applications, and therefore, it is important to find a better operational condition with the aid of a theoretical thrust model. We have developed a theoretical thrust model which consists of a trajectory analysis and an electric field penetration model in the electrostatic approximation. The model shows that experimental parameters are off from an optimum operational condition which provides the maximum thrust.