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Electrification of aircraft is a very challenging task as the demand for energy and power is high. While the storage and generation of electrical energy are widely studied due to the limited specific energy and specific power of batteries and fuel cells, electric machines (power electronics and motors) which have years of experience in many industrial fields must be improved when applied to aviation: they generally have a high efficiency but the increase in power levels determines significant thermal loads which, unlike internal combustion engines (ICE), cannot be rejected with the exhaust. There is therefore a need for thermal management systems (TMSs) with the main objective of maintaining operating temperatures below the maximum level required by electric machines. Turboprop aircraft, such as the ATR 72 or the Dash 8-Q400, are commonly used for regional transport and are equipped with two gas turbine engines whose combined power is in the order of 4 MW. Electric and hybrid propulsion systems for these aircraft are being studied by several leading commercial aviation industries and start-ups, and the 1MW motor size seems to be the main option as it could be used in different aircraft configurations, particularly those that exploit distributed electric propulsion. With reference to the topics mentioned above, the present work presents the design of a TMS for a high-power motor/generator whose electrical architecture is known. Once integrated with the electrical part, the TMS must allow a weight/power ratio of 14 kW/kg (or 20 kW/kg at peak power) while maintaining the temperature below the limit temperature with reasonable safety margins. Submerged jet oil is the cooling technique here applied with a focus on diathermic oil. Parameters affecting cooling, like rotor speed and filling factor, are analysed with advanced CFD. [ABSTRACT FROM AUTHOR]

The aircraft power supply system is the power source of all electrical equipment on board，thus its safety and reliability are pretty important. Against the backdrop of environmental protection and high-efficiency development needs，research and application of more/all electric aircraft technology with electric power as the core is advanced. The widespread use of electric drive devices and power electronic devices has led to the complexity of aircraft power supply system structure，which puts forward higher requirements for aircraft reliability，safety，testability and maintainability，so that researches on fault diagnosis technology for aircraft power supply systems are of great significance. In this paper，firstly the composition structure and respective functions of aircraft power supply system are introduced. Secondly，the development process of aircraft power supply system is outlined. Thirdly，the characteristics of typical domestic and foreign power supply systems are compared. Fourthly，the main failure modes，fault characteristics and failure causes of aircraft power supply system are summarized，and a design architecture of aircraft power health management system is proposed. Fifthly，the research progress of fault diagnosis methods based on model and data are reviewed，then，the characteristics of various diagnostic methods are evaluated from aspects such as accuracy，data demand，applicability and implementation difficulty. Finally，the challenges and development trends of fault diagnosis technology for aircraft power supply system are pointed out.

This paper proposes a comprehensive power system designed for the use of a more electric aircraft power distribution system. Instead of traditional Nicad battery solutions as the energy source of the aircraft power system, lithium battery structures, which are a recent and promising solution in the field of aviation power systems, are modeled and analyzed. In this study, a WPWM-based, single-phase, multi-level pure sine wave static aircraft-inverter system is designed and integrated to improve the performance of conventional aircraft power systems. In the designed power system, a boost converter structure is proposed that boosts 28 VDC-to-270 VDC voltage coming from the lithium–ion battery pack and can reach a steady state in 0.032 s. The performance of the modeled WPWM-based aircraft-inverter system, compared to SPWM Bipolar and Unipolar switching techniques commonly used in single-phase inverter designs, reveals a THD reduction of approximately 27% with WPWM, resulting in a THD value below 2% for both load current and load voltage. As a result of the study, a power system that will enable the aircraft avionics, ventilation, and navigation systems to perform better than conventional power systems and comply with aircraft electric-power characteristic standards has been designed and detailed. [ABSTRACT FROM AUTHOR]

An important reflection of the overall efficiency, reliability, and passenger safety of a more electric aircraft (MEA) is the output performance of its electrical power system (EPS) controller. This output performance encompasses the rise time, setting time, and percent undershoot of the voltage across the capacitor bank. Therefore, this article presents an optimal controller design using an artificial intelligence method called the adaptive tabu search (ATS) algorithm. The state-variables-averaging model is applied with the ATS algorithm to reduce computational time. Moreover, stability analysis based on the eigenvalue theorem is used as the penalty condition during the searching process to avoid unstable operation. The output performance of the proposed controller design is superior to that of the conventional controller design. All design results are verified by good agreement with MATLAB and hardware-in-the-loop (HIL) simulations.

Electrification of aircraft is a very challenging task as the demand for energy and power is high. While the storage and generation of electrical energy are widely studied due to the limited specific energy and specific power of batteries and fuel cells, electric machines (power electronics and motors) which have years of experience in many industrial fields must be improved when applied to aviation: they generally have a high efficiency but the increase in power levels determines significant thermal loads which, unlike internal combustion engines (ICE), cannot be rejected with the exhaust. There is therefore a need for thermal management systems (TMSs) with the main objective of maintaining operating temperatures below the maximum level required by electric machines. Turboprop aircraft, such as the ATR 72 or the Dash 8-Q400, are commonly used for regional transport and are equipped with two gas turbine engines whose combined power is in the order of 4 MW. Electric and hybrid propulsion systems for these aircraft are being studied by several leading commercial aviation industries and start-ups, and the 1MW motor size seems to be the main option as it could be used in different aircraft configurations, particularly those that exploit distributed electric propulsion. With reference to the topics mentioned above, the present work presents the design of a TMS for a high-power motor/generator whose electrical architecture is known. Once integrated with the electrical part, the TMS must allow a weight/power ratio of 14 kW/kg (or 20 kW/kg at peak power) while maintaining the temperature below the limit temperature with reasonable safety margins. Submerged jet oil is the cooling technique here applied with a focus on diathermic oil. Parameters affecting cooling, like rotor speed and filling factor, are analysed with advanced CFD.

The electrification of the aircraft engine and the concept of more electric aircraft is a trend of recent decades that has attracted the attention of global aviation corporations. Electrification of traditional nonelectric systems increases efficiency and reliability and reduces emissions of harmful substances into the atmosphere. The attention of the scientific and technical literature is directed toward more electric aircraft, without separating the aircraft engine electrification technology. A review of related technologies and processes in aircraft engine electrification that have not received much disclosure in the publicly available reviews, but without which the full implementation of more electric engines in the aviation industry is impossible, is presented. This article is focused on the control of more electric engines, power transmission with simultaneous integration of starter-generators and electric generators, integration into aircraft engines, heat recovery, and training of highly qualified personnel for the design and operation of more electric engines. [ABSTRACT FROM AUTHOR]

This review proposes an overview of hybrid electric and full electric powertrains dedicated to greener aircraft in the "sky decarbonization" context. After having situated the state of the art and context of energy hybridization in the aviation sector, we propose the visit of several architectures for powertrain electrification, situating the potential benefits but also the main challenges to be faced to takeoff these new solutions. Then, as a first example, we consider the EU project "HASTECS" (Hybrid Aircraft: reSearch on Thermal and Electric Components and Systems) in the framework of Clean Sky 2. It relates to a series hybrid chain integrated into a regional aircraft. This energy system integrates especially power electronics and electric machines with a high degree of integration, which raises the "thermal challenge" and the need to integrate cooling devices. Through the snowball effects typical of the aviation sector, this example emphasizes how important it is to "hunt for kilos", an alternative solution consisting of eliminating the power electronics within the powertrain. This is why we propose a second example, which concerns an AC power channel without power electronics that only integrates synchronous magnet machines (generator and motor) directly coupled on an AC bus. This last architecture nevertheless raises questions in terms of stability, with one solution being to insert an auxiliary hybridization branch via battery storage. Theoretical analyses and experiments at a reduced power scale show the viability of this concept. Finally, some recommendations for future research with potential technological breakthroughs complete that review. [ABSTRACT FROM AUTHOR]

Advances in the areas of electrical and power electronic technologies play a key role in more electric aircraft (MEA), especially in power converters. Therefore, most electrical power loads on MEA are tightly controlled power converters that behave as constant power loads (CPLs). These CPLs look like a small-signal negative impedance that can significantly degrade the overall system stability. The destabilizing effect may result in poor performance until the DC bus voltage of the system does not adhere to the MIL-STD-704F standard. Thus, a stability study is very important to avoid unstable operations. A proposed MEA model which can be derived using the DQ approach is comprehensive in the introduced stability analysis methods in this article. These methods, i.e., a small-signal stability analysis using the eigenvalue theorem, a modal analysis technique called participation factor analysis, and a large-signal stability analysis via phase-plane analysis, are performed to investigate the stability margin. Moreover, the impact of key parameter variations on MEA stability is also taken into account to deliver the ways of parameter selection at the early design stages for an engineer. The MATLAB topology model and processor-in-loop (PIL) simulations validated the analytical results. The results indicate that a good agreement between theoretical, simulation, and PIL results can be achieved. [ABSTRACT FROM AUTHOR]

This paper proposes a comprehensive power system designed for the use of a more electric aircraft power distribution system. Instead of traditional Nicad battery solutions as the energy source of the aircraft power system, lithium battery structures, which are a recent and promising solution in the field of aviation power systems, are modeled and analyzed. In this study, a WPWM-based, single-phase, multi-level pure sine wave static aircraft-inverter system is designed and integrated to improve the performance of conventional aircraft power systems. In the designed power system, a boost converter structure is proposed that boosts 28 VDC-to-270 VDC voltage coming from the lithium–ion battery pack and can reach a steady state in 0.032 s. The performance of the modeled WPWM-based aircraft-inverter system, compared to SPWM Bipolar and Unipolar switching techniques commonly used in single-phase inverter designs, reveals a THD reduction of approximately 27% with WPWM, resulting in a THD value below 2% for both load current and load voltage. As a result of the study, a power system that will enable the aircraft avionics, ventilation, and navigation systems to perform better than conventional power systems and comply with aircraft electric-power characteristic standards has been designed and detailed.

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In order to enhance the transient performance of aircraft high voltage DC (HVDC) generation system with wound rotor synchronous machine (WRSM) under a wide speed range, the nonlinear PI multi-loop control strategy is proposed in this paper. Traditional voltage control method is hard to achieve the dynamic performance requirements of the HVDC generation system under a wide speed range, so the nonlinear PI parameter adjustment, load current feedback and speed feedback are added to the voltage and excitation current double loop control. The transfer function of the HVDC generation system is derived, and the relationship between speed, load current and PI parameters is obtained. The PI parameters corresponding to the load at certain speed are used to shorten the adjusting time when the load suddenly changes. The dynamic responses in transient processes are analyzed by experiment. The results illustrate that the WRSM HVDC generator system with this method has better dynamic performance.

Research on electric power systems (EPSs) for the aviation industry has recently grown significantly due to the need to reduce global CO2 emissions from transportation. To fulfill the power requirements of a more electric aircraft (MEA), DC power distribution has emerged as a potential solution. However, the progress of DC distribution faces significant difficulties related to system protection. Solid-state power controllers (SSPCs) are being considered in these applications due to their ability to provide fast-tripping mechanisms for system protection. Although SSPCs have been successfully implemented in low voltage DC 28V aircraft networks, their application in high voltage systems (270 V, ±270 V, or higher) presents challenges, such as over-voltage and excessive power loss, particularly for high-power applications. This paper focuses on the development of SSPCs for a 270 V DC system with a current rating of 125 A / 250 A. The paper presents designs for over-voltage suppression and thermal management of the SSPCs. The study also includes a comparative analysis of using a different number of SiC MOSFET modules connected in series and parallel and their effect on the cooling requirements and circuit temperature to assess power losses, power density, weight, and cooling requirements for the SSPCs. A prototype of the proposed SSPC has been built for experimental validation. Results show effective over-voltage suppression to 480 V and quick interruption capabilities with trip currents of 250 A and 375A within time intervals of $160 ~\mu \text{s}$ and $300 ~\mu \text{s}$ , respectively, for line inductance of $105 ~\mu \text{H}$ . The circuit withstands energy up to 22.5 J for a breaking current of 375 A.

This paper presents the current scenario related to the application of power electronics converters in aviation industries, major challenges and their reliability aspects. Recent developments in power electronics have given a major breakthrough in the aviation industry. These developments are not only limited to the more electric aircraft (MEA) but also in the field of electric propulsion, popularised as electric propulsion aircraft (EPA). The aircraft requires a well‐established protection scheme of these converter topologies. The overall reliability of aircraft solely depends on its protection and fast mitigation of any fault if occurs in the cruise time. This paper aims to provide a comprehensive analysis of power converters applications in MEA and EPA. Further, a review of various topologies of Power Converters and their reliability assessment is also described in detail. The physics of Failure of switches, their lifetime estimation and thermal cycling have also been discussed. Finally, the reliability assessment of different topologies and their comparison has been discussed for overall performance enhancement of the MEA and EPA. [ABSTRACT FROM AUTHOR]

This paper presents a novel direct power control (DPC) strategy based on an optimized sector division for a three-phase coupled inductor-based bipolar output active rectifier (TCIBAR) applied in more electric aircraft (MEA). First, based on the instantaneous power theory, the power model of the TCIBAR is built in the synchronous rotating coordinate system. Second, to implement the hysteresis power control of TCIBAR without causing the runaway of the zero-sequence current in the three-phase coupled inductor (TCI), a set of new voltage vectors that have the same zero-sequence voltage (ZSV) component are synthesized and adopted in the proposed DPC strategy. Third, by quantitatively analyzing the effect of the new synthesized voltage vectors on the power variation of TCIBAR, an optimized sector division method is proposed to improve the accuracy of power control and reduce the phase current harmonics in TCIBAR. Finally, to maintain the voltage balance of the bipolar dc ports in TCIBAR, voltage balance control is studied in the proposed DPC strategy. The proposed DPC strategy is researched on an experimental platform of TCIBAR, and the results show that the proposed DPC strategy is feasible and has good static and dynamic performance. [ABSTRACT FROM AUTHOR]

In present decades, the need for a more efficient air transport system is driving towards more electric aircraft subsystems. Electrified subsystems offer the opportunity to optimize the operational performance of systems by reducing the power required by the propulsion system therefore, reducing the fuel burnt. A considerable advantage can be obtained by electrifying the environmental control system which is the most power demanding aircraft subsystem. The paper presents a simplified model to estimate the main performances of the conventional and electrified environmental control system during the aircraft conceptual and preliminary design phases. Different air cycle machine architectures can be designed. The model is divided in two main modules. The first is dedicated to the estimation of the aircraft thermal loads including the effect of solar radiation, the conduction of external air, the presence of passengers and the avionic systems. With the main result of estimating the cabin airflow required to control the temperature and the air quality within the aircraft's mission profile. The second module of the model designs all major components of the air conditioning pack including the dedicated compressors of the electrified environmental control system. The model requires basic input data that can be easily estimated during the early stages of aircraft design. The model is calibrated using the available data of a conventional system and then applied to the electrified one. The results show the increased efficiency of the electrified system that results from optimizing both pneumatic power generation and the air cycle machine. [ABSTRACT FROM AUTHOR]

With the development of aircraft design toward more electricity,the development of electric environmental control system in more electric aircraft has received wide attention.This paper introduces the operation principle of Boeing B787 electric environmental control system,and reviews the current research status on the architectural design and trade-offs of electric environmental control system and the design of key components.It is proposed that since the adoption of electric environmental control system design will aggravate the complexity of the aircraft,integrated tools are needed to support the architectural design of electric environmental control system,and to develop the agent modeling techniques that can replace CFD calculations to shorten the design cycle of key components.This paper will provide a reference for the research and development of domestic electric environmental control systems.

The aircraft environmental control system (ECS) is the second-highest fuel consumer system, behind the propulsion system. To reduce fuel consumption, one research direction intends to replace conventional aircraft with more electric aircraft. Thus, new electric architectures have to be designed for each system, such as for the ECS. In this paper, an electric ECS is modeled and then sized and optimized for different sizing scenarios with the aim of minimizing fuel consumption at the aircraft level. For the system and for each component, such as air inlets and heat exchangers, parametric models are developed to allow the prediction of relevant characteristics. These models, developed in order to be adapted to aircraft design issues, are of different types, such as scaling laws and surrogate models. They are then assembled to build a preliminary sizing procedure for the ECS by using a multidisciplinary design analysis and optimization (MDAO) formulation. Results show that the ECS design is highly dependent on the sizing scenario considered. An approach to size the ECS globally with respect to all the sizing scenarios leads to an ECS that accounts for around 200 N of drag, 190 kW of electric power, and 1500 kg of mass for the CeRAS aircraft. [ABSTRACT FROM AUTHOR]

This paper proposes a novel multi-phase single-turn permanent magnet synchronous motor (MSPMSM) to meet the stringent requirements of electric aircraft, such as high power density, high torque, and low mass and volume. The MSPMSM features a unique single-turn winding configuration, which can reduce the coil overhang, increase the power density and torque, and reduce the weight of the motor in conventional multi-turn motors. The proposed MSPMSM is studied by the finite element method (FEM) and compared to a conventional PMSM with the same geometry and PM usage. The performance comparison results under different load conditions show that the proposed MSPMSM has higher torque and power density than the conventional motor. [ABSTRACT FROM AUTHOR]

Abstract This paper presents the current scenario related to the application of power electronics converters in aviation industries, major challenges and their reliability aspects. Recent developments in power electronics have given a major breakthrough in the aviation industry. These developments are not only limited to the more electric aircraft (MEA) but also in the field of electric propulsion, popularised as electric propulsion aircraft (EPA). The aircraft requires a well‐established protection scheme of these converter topologies. The overall reliability of aircraft solely depends on its protection and fast mitigation of any fault if occurs in the cruise time. This paper aims to provide a comprehensive analysis of power converters applications in MEA and EPA. Further, a review of various topologies of Power Converters and their reliability assessment is also described in detail. The physics of Failure of switches, their lifetime estimation and thermal cycling have also been discussed. Finally, the reliability assessment of different topologies and their comparison has been discussed for overall performance enhancement of the MEA and EPA.

The increased cost of fuel and maintenance in aircraft system lead to the concept of more electric aircraft, moreover this concept increase the use of power electronic converters in aircraft power system. Since in this application, the reliability is a crucial feature. Therefore, the use of more efficient, reliable and robust power converter with health management capability will be a big challenge. Multicellular topology of power converters has the required performance in terms of efficiency and robustness. However, the increased complexity of control and more power components (power switches and capacitors) goes along with an increase in possibility of failure in multicellular topology. Therefore, the main contribution of this paper is the use of multicellular topology advantageous with fault diagnosis and fault tolerant control in order to increase the robustness reliability. The health management using a fault detection with Fuzzy Pattern Matching (FPM) algorithm when a failure in power switches or flying capacitors of multicellular converter and a Fault Tolerant Control (FTC) with sliding mode of second parallel three cells multicellular converters. Simulation results with Matlab show the increased efficiency and the continuity of work during failure mode in aircraft power system.

In the literature, different DC/DC power electronic converters (PECs) have been found to interconnect high-voltage DC and low-voltage DC grids in the electric power distribution networks of aircraft. In this scenario, the dual-active-bridge (DAB) converter has been shown to be one of the most promising topologies. The main disadvantages of this PEC are the large output capacitance required to satisfy more electric aircraft (MEA) requirements and the high conduction losses produced in low-voltage power devices of (LV). Therefore, this paper proposes analytical models to determine the voltage ripple and root-mean-square (RMS) current in DC bus capacitors of DABs considering different modulation strategies. Moreover, an analysis of the design space in an MEA case study is performed to evaluate the influence of the design variables in power losses of power devices and peak-to-peak voltage ripple in DC bus capacitors. These models are useful for the design stage of this PEC, as well as to enable multi-objective optimization procedures by reducing the computational cost of these methodologies. Furthermore, the exploration of the switching frequency and limit of the modulation angle aid in reducing the resulting volume of the low-voltage DC capacitor. [ABSTRACT FROM AUTHOR]

The requirements for improvement of the efficiency, reliability, and environmental safety, as well as for reducing the weight and size for aircraft devices and systems, are decisive in choosing the direction of development of the aviation industry. The development of the concept of more electric aircraft and the concept of an more electric engine, which involves replacing hydraulic and pneumatic systems by electric equivalents, is gaining ever more popularity. Both individual research groups and global aircraft engineering and aviation corporations are paying attention to the development of these promising technologies. The concept of more electric aircraft is widely covered in the scientific and technical literature, whereas the concept of an more electric engine included therein has not received such attention. This paper presents an overview of publications devoted to the use of electric machines for the more electric engine. The paper is focused on electric machines embedding into aircraft engine shafts (embedded starter-generators for a high-pressure shaft cascade and embed electric generators for a low-pressure shaft cascade, and electric drives of fuel-and-oil systems, as well as the use of electric machines in air conditioning systems onboard an aircraft). [ABSTRACT FROM AUTHOR]

This paper presents a novel active power control scheme for SRMs that are used as S/G units at MEAs; the proposed method uses a vector-controlled SAF, facilitating the acceleration of the SRM during starter mode and the efficient power control of the SRG at various operational points during generator mode. Moreover, an energy storage unit is incorporated in the SAF, aiming at enhancing the SRM's starting torque production and taking full advantage of the SRG's energy generation capability, through the appropriate phases’ magnetization regulation that is imposed by the developed adaptive SAF vector control. In addition, the proposed scheme provides enhanced FRTC during both modes of operation. The mathematical analysis is verified through MATLAB/Simulink simulations, while the performance of the S/G concept under study is evaluated via real-time CHIL tests, with the use of the dSPACE 1202 (MicroLabBox) platform and the 16-bit dsPIC30F4011 microcontroller. [ABSTRACT FROM AUTHOR]

The secondary energy on board has been gradually replaced by electrical power in more electric aircraft, effectively improving the fuel economy, reliability and maintainability of the aircraft, and has become an important direction of the aviation technology. As the core of the secondary energy system on board, the power system plays a key supporting role in the development of more electric aircraft. The innovative development of the power system and related technologies are the necessary basis to realize the improvement and global optimization of the aircraft comprehensive performance. Based on the concepts and characteristics of more electric aircraft, this paper analyzes and compares the electric power system architecture of three typical more electric aircraft, summarizes the key technologies supporting the development of more electric aircraft electric power system, and discusses the development trend of high voltage, direct current and intelligent more electric aircraft electric power system in the future. [ABSTRACT FROM AUTHOR]

The aeronautical industry is one of the last groups which has decided to join the distributed architecture trend in order to increase both the safety and efficiency of future aircraft. However, the actual power system used in aircrafts has certain limitations that makes it difficult to achieve this objective due to its high centralisation. This paper proposes a novel methodology for the design of a new modular power electronic device that is based on semiconductor technology, which provides the equipment and functionality that makes the full decentralisation of the aircraft DC power system possible in the near future. The proposed and developed device can be adapted to the requirements of the system where it is going to be implanted, as it is possible to add as many devices as the system needs and have as many modules as needed. All the modules are bidirectional, which makes the system more redundant and fault tolerant, as the number of possible paths to feed the loads is increased. Tests were carried out with 5 devices (4 modules per device), 3 power supplies and 2 loads, where the correct operation of the system (path search, load feed and failure management) was proved.

This review proposes an overview of hybrid electric and full electric powertrains dedicated to greener aircraft in the “sky decarbonization” context. After having situated the state of the art and context of energy hybridization in the aviation sector, we propose the visit of several architectures for powertrain electrification, situating the potential benefits but also the main challenges to be faced to takeoff these new solutions. Then, as a first example, we consider the EU project “HASTECS” (Hybrid Aircraft: reSearch on Thermal and Electric Components and Systems) in the framework of Clean Sky 2. It relates to a series hybrid chain integrated into a regional aircraft. This energy system integrates especially power electronics and electric machines with a high degree of integration, which raises the “thermal challenge” and the need to integrate cooling devices. Through the snowball effects typical of the aviation sector, this example emphasizes how important it is to “hunt for kilos”, an alternative solution consisting of eliminating the power electronics within the powertrain. This is why we propose a second example, which concerns an AC power channel without power electronics that only integrates synchronous magnet machines (generator and motor) directly coupled on an AC bus. This last architecture nevertheless raises questions in terms of stability, with one solution being to insert an auxiliary hybridization branch via battery storage. Theoretical analyses and experiments at a reduced power scale show the viability of this concept. Finally, some recommendations for future research with potential technological breakthroughs complete that review.

This paper presents the integrated optimal design of the powertrain of a hybrid regional aircraft using multidisciplinary design optimization (MDO). The sizing of the main components of the propulsion chain is performed over the flight mission under various scenarios regarding energy management strategies and technological assessments. For that purpose, a complete set of multidisciplinary surrogate models are integrated into the MDO process, taking account not only of the main electrical, thermal and mechanical aspects but also of environmental constraints such as partial discharges in electric motors regarding flight conditions. Several MDO formulations are investigated comparing local (i.e., motor mass minimization) and global optimizations (i.e., powertrain mass then fuel burn minimization at aircraft level). Results emphasize main systemic couplings showing that despite future technological progress, the series hybrid architecture is heavier than a conventional thermal aircraft. Nevertheless, thanks to the whole aircraft optimization, potential gains related to kerosene consumption can be reached, reducing the environmental footprint. The "energy gains" focused on in this paper may be added with aerodynamic gains potentially involved in more electric powertrain. This work has been carried out in the frame of the HASTECS project under the Clean Sky II program which aims at reducing CO2 emissions and environmental impacts of the aviation sector. [ABSTRACT FROM AUTHOR]

Flight control surfaces guarantee a safe and precise control of the aircraft. As a result, hinge moments are generated. These moments need to be estimated in order to properly size the aircraft actuators. Control surfaces include the ailerons, rudder, elevator, flaps, slats, and spoilers, and they are moved by electric or hydraulic actuators. Actuator sizing is the key when comparing different flight control system architectures. This fact becomes even more important when developing more-electric aircraft. Hinge moments need to be estimated so that the actuators can be properly sized and their effects on the overall aircraft design are measured. Hinge moments are difficult to estimate on the early stages of the design process due to the large number of required input. Detailed information about the airfoil, wing surfaces, control surfaces, and actuators is needed but yet not known on early design phases. The objective of this paper is to propose a new methodology for flight control system sizing, including mass and power estimation. A surrogate model for the hinge moment estimation is also proposed and used. The main advantage of this new methodology is that all the components and actuators can be properly sized instead of just having overall system results. The whole system can now be sized more in detail during the preliminary design process, which allows to have a more reliable estimation and to perform systems installation analysis. Results show a reliable system mass estimation similar to the results obtained with other known methods and also providing the weight for each component individually. [ABSTRACT FROM AUTHOR]

The increasing number and functional complexity of power electronics in more electric aircraft (MEA) power systems have led to a high degree of complexity in modelling and computation, making real-time energy management a formidable challenge, and the discrete-continuous action space of the MEA system under consideration also poses a challenge to existing DRL algorithms. Therefore, this paper proposes an optimisation strategy for real-time energy management based on hybrid deep reinforcement learning (HDRL). An energy management model of the MEA power system is constructed for the analysis of generators, buses, loads and energy storage system (ESS) characteristics, and the problem is described as a multi-objective optimisation problem with integer and continuous variables. The problem is solved by combining a duelling double deep Q network (D3QN) algorithm with a deep deterministic policy gradient (DDPG) algorithm, where the D3QN algorithm deals with the discrete action space and the DDPG algorithm with the continuous action space. These two algorithms are alternately trained and interact with each other to maximize the long-term payoff of MEA. Finally, the simulation results show that the effectiveness of the method is verified under different generator operating conditions. For different time lengths T, the method always obtains smaller objective function values compared to previous DRL algorithms, is several orders of magnitude faster than commercial solvers, and is always less than 0.2 s, despite a slight shortfall in solution accuracy. In addition, the method has been validated on a hardware-in-the-loop simulation platform. [ABSTRACT FROM AUTHOR]

This article discusses a high-speed brushless doubly-fed machine (BDFM) for aviation turboelectric distributed propulsion application. After a brief overview of the machine fundamentals and propulsion architecture, the electromagnetic properties of the newly designed BDFM are analyzed. An innovative rotor mechanical design is introduced, which allows for a significant increase of speed compared to legacy rotor designs. Experiments were conducted to verify the 12 000-r/min high-speed operation. The findings in this article will give prominence to the feasibility of applying BDFMs for aviation applications. [ABSTRACT FROM AUTHOR]

Accurate identification of electrical load working status can provide information support to the remote electrical distribution system (EDS) of more electric aircraft (MEA), which could use it to realize redundant switching and protection. This paper presents a method to automatically identify the load status on the remote power distribution unit (RPDU) of MEA by using an intelligent algorithm. The experimental platform is built in an aircraft Electrical Power System (EPS) distribution large-scale test cabin. Four pieces of typical aviation equipment are installed in the test cabin and powered from RPDU. Voltage and current values under 15 working combinations on the RPDU are measured to extract the steady-state V-I trajectory. In total, 750 group samples were collected in the feature parameter database. A generalized regression neural network (GRNN) identification model was established, and the smoothing factor was calculated by using a conventional cross-validation method to train and reach an optimal value. However, the identification results are not ideal. In order to improve the accuracy, the parameter of GRNN was optimized by genetic algorithms. The proposed model shows great performance as accuracy of all 15 classifications reached 100%. The proposed model has advantages of flexible network structure, high fault tolerance, and robustness. It can realize global approximation optimization, avoid local optimization, effectively improve GRNN fitting accuracy, improve model generalization ability, and reduce model training calculation. [ABSTRACT FROM AUTHOR]

The operating conditions can have uncontrolled effects on the voltage of a High-Temperature Proton Exchange Membrane Fuel Cell (HT-PEMFC). For instance, the HT-PEMFC can be used at ambient pressure, i.e., without having a back pressure regulator. In this case, the variation in the atmospheric pressure directly affects pressures inside the fuel cell, which induces voltage variation. Moreover, in transient phases, several coupled phenomena can have an uncontrolled effect on the voltage. For example, following a change in the current operating point, thermal conditions in the fuel cell can vary, and the temperature stabilization then leads to a voltage variation. This article introduces a readjustment method for the fuel cell voltage to compensate for the effects of the pressure and temperature variations that are undergone and to decouple their effects. This methodology is based on the realization of a design of experiments to characterize the voltage sensitivity to pressure ([1; 1.5 bar]) and temperature ([120; 180 °C]) between 0.2 and 1 A/cm2 of an Advent PBI MEA (formerly BASF Celtec®-P 1100 W). The data obtained allowed identifying an empirical model that takes into account the aging caused by the experiment. Finally, the methodology is criticized before proposing an alternative method. [ABSTRACT FROM AUTHOR]

Brushless synchronous generators, which have the structure of three-stage generators to realize brushless excitation, are the most commonly used aero-generators. With the development of more electric aircraft technology, the output frequency of aero-generators is changing from constant to variable, and the characteristics of generators are also changed. To make sure that the generators can be adapted to the variable frequency operation, this paper analyzes the effects of different operating conditions on generator performance and proposes corresponding design methods. The design and verification of the electromagnetic field, fluid field, temperature field and stress field are carried out for the variable frequency generator. A prototype generator is manufactured according to the design results. An experimental platform is set up to test the performances of the designed generator. The final experimental results show that the generator can work well in variable frequency generation systems, which proves the design method proposed in this paper is effective and feasible. [ABSTRACT FROM AUTHOR]

The increased cost of fuel and maintenance in aircraft system lead to the concept of more electric aircraft, moreover this concept increase the use of power electronic converters in aircraft power system. Since in this application, the reliability is a crucial feature. Therefore, the use of more efficient, reliable and robust power converter with health management capability will be a big challenge. Multicellular topology of power converters has the required performance in terms of efficiency and robustness. However, the increased complexity of control and more power components (power switches and capacitors) goes along with an increase in possibility of failure in multicellular topology. Therefore, the main contribution of this paper is the use of multicellular topology advantageous with fault diagnosis and fault tolerant control in order to increase the robustness reliability. The health management using a fault detection with Fuzzy Pattern Matching (FPM) algorithm when a failure in power switches or flying capacitors of multicellular converter and a Fault Tolerant Control (FTC) with sliding mode of second parallel three cells multicellular converters. Simulation results with Matlab show the increased efficiency and the continuity of work during failure mode in aircraft power system. [ABSTRACT FROM AUTHOR]

With the aim towards lighter and more efficient electrical systems in future aircraft, design of DC/DC converters with high efficiency, power density and improved thermal management becomes necessary. This paper investigates the detailed design of isolated bidirectional DC/DC converters for more electric aircraft (MEA). Use of wide bandgap (WBG) devices to enhance system efficiency is considered. A transformer optimization technique based on switching frequency for different converter topologies is investigated. The control strategy of the discussed configurations are verified in the PLECS simulation environment. Dual active bridge (DAB), input-series output-parallel (ISOP), neutral point clamped (NPC) and active neutral point clamped (ANPC) converters are considered to exploit benefits offered by WBG devices for MEA. A comparison is performed in terms of efficiency, thermal management, power density and electromagnetic interference (EMI). The results indicate that the ANPC converter is better for efficiency, thermal management and power density. The DAB converter showed similar results to the ISOP with the DAB yielding higher efficiency. The NPC and ANPC configurations resulted in the best EMI performance. The NPC converter proved to be the least effective solution with regard to efficiency, thermal management and power density.

Aircraft electrification has been a major trend in aviation industry for past 20 years. Given the increasing electrical power requirement for future more electric aircraft and hybrid electric aircraft, research efforts has been ongoing in high power electrical conversion for air-borne systems. Safety critical nature of aviation systems places reliability of aerospace power converters as a critical design concern. In this paper, power electronic system reliability is studied with emphasis on lifetime limiting factors of critical sub components. Reliability of voltage source power converters at different system voltage levels are modelled for a starter generator drive converter. A key observation is that Si IGBT devices are sufficient with respect to reliability requirements in low and medium voltage systems (up to 540 V). At higher system voltages (above 540 V), multi level topologies are necessary for designing with Si IGBTs. In constant power profile drive, system reliability is minimally affected by wear-out failure of film capacitors in converter DC links. In multi level topologies without enhanced voltage derating, system reliability is dominated by cosmic ray induced random failures. Simulation results demonstrate that at high system voltages (810 V), 2 L topology with SiC mosfets outperform Si IGBT based 3 L topology with respect to reliability.

The arrival on the market of new power devices based on wide bandgap semiconductors has raised a relevant interest due to their superior properties compared to conventional technologies. On the other hand, these devices are inherently characterized by high rates of voltage changes over time, which may result in reliability challenges in electric drives adopting them. In fact, dangerous voltage overshoots at the motor terminals and uneven voltage distributions within the machine windings may occur. These phenomena can trigger a high insulation stress and partial discharges and, as a consequence, they may concur to the premature failure of the dielectric materials. This paper proposes a flexible and comprehensive modelling approach for the accurate analysis and estimation of both voltage overshoots and voltage distributions in a typical converter-cable-motor system intended for more electric aircraft applications. The modelling results are validated against experimental measurements carried out on a physical prototype comprising a wide bandgap-based converter, a connecting cable and an electrical machine stator. The findings are then used in the companion papers (part II and part III) to investigate the dependence of partial discharge phenomena on these voltage waveforms, highlight reliability challenges in modern ±270 V DC bus voltage drives for the more electric aircraft and discuss solutions.

In this paper, a comprehensive formulation of reliability models for active and passive power components (i.e., capacitors and power semiconductor devices (MOSFETS)) of a regulated transformer rectifier unit (RTRU) for more electric aircraft application is presented. Additionally, the work focuses on the reliability analysis and evaluation at the system level for a non-modular discrete MOSFET based auxiliary power supply unit for more-electric-aircraft. A rigorous quantitative assessment of reliability, volume and power loss is carried out at the component level as well as system level, verifying how a particular design relates to the specific performance parameters. Keeping this as a motivation, this paper also focuses on studying and analyzing the inter-relation among reliability and other performance metrics such as power density and efficiency taking a 3kW RTRU as an example design. The major metrics under consideration for this study are the failure in time (FIT), mean time to failure (MTTF), power density, efficiency while keeping the eye on most failure prone elements of the power converter - capacitors and active power semiconductors. A comprehensive quantitative evaluation based on thermal management size, system reliability is done on choice of the switching semiconductors where the candidate pool includes Si and SiC MOSFET technologies. From the analysis on three different types of widely used capacitors - film, ceramic, aluminum electrolytic, it is concluded that the optimum choices of capacitors for high voltage DC link and low-voltage output belong to Aluminum and Ceramic types, respectively. Finally, the optimum selection of power stage components is derived from the extensive evaluation of the overall RTRU system in terms of system size and reliability.

Drive system availability is critical in aerospace applications. Multi three phase machines are a known solution to provide redundancy and fault tolerance capability. Control of multi three phase machines is typically implemented in DQ domain or VSD domain. An alternate control approach is per phase current control using proportional resonant controllers in ABC domain. In this work, ABC domain current controllers are utilized to demonstrate key advantages in high reliability applications - fault reconfigurability with minimal controller modifications and ability to control multiple harmonic components. ABC domain control structure enables seamless transition from healthy mode to fault mode with only algebraic operations at the controller output unlike DQ/VSD domains. Furthermore, single-phase operating mode in ABC domain provides an additional fault mode for emergency operation which is unavailable in DQ/VSD control domains. The Fault mode operation under open circuit fault of one phase in each winding set is demonstrated. Simulation and experimental results validate the normal and fault condition operation of the proposed control method.