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More electric aircraft (MEA) is defined as the extensive usage of electric power in aircraft. The demand for electric power in new generation aircraft rises due to environmental and economic considerations. Hence, efficient and reliable starter/generators (SGs) are trending nowadays. The conventional main engine starting system and power generation system can be replaced with an individual SG. The constraints of the SG should be investigated to handle the aviation requirements. Even though the SG is basically an electric machine, it requires a multidisciplinary study consisting of electromagnetic, thermal and mechanical works to cope with aviation demands. This study aims to review conventional and new-generation aircraft SGs from the perspective of electric drive applications.

First of all, the importance of the MEA concept has been briefly explained. Also, the historical development and the need for higher electrical power in aircraft have been indicated quantitatively. Considering aviation requirements, the candidate electrical machines for aircraft SG have been determined by the method of scoring. Those machines are compared over 14 criteria, and the most predominant of them are specified as efficiency, power density, rotor thermal tolerance, high-speed capability and machine complexity. The features of the most suitable electrical machine are pointed out with data gathered from empirical studies. Finally, the trending technologies related to efficient SG design have been explained with numeric datasets.

The induction motor, switched reluctance motor and permanent magnet synchronous motor (PMSM) are selected as the candidate machines for SGs. It has been seen that the PMSM is the most preferable machine type due to its efficient operation in a wide range of constant power and speed. It is computationally proven that the using amorphous magnetic alloys in SG cores increases the machine efficiency more. Also, the benefits of high voltage direct current (HVDC) use in aircraft have been explained by a comparison of different aircraft power generation standards. It is concluded that the HVDC use in aircraft decreases total cable weight and increases aircraft operation efficiency. The thermal and mechanical tolerance of the SG is also vital. It has been stated that the liquid cooling techniques are suitable for SGs.

The demand for electrical power in new generation aircraft is increasing. The SG can be used effectively and efficiently instead of conventional systems. To define requirements, constraints and suggestions, this study investigates the SGs from the perspective of electric drive applications.

With the objective to assess potentially performant hybrid-electric architectures, this paper aims to present an aircraft performance level evaluation, in terms of range and payload, of the synergies between a hybrid-electric energy system configuration and a cryogenic fuel system.

An unmanned aerial vehicle (UAV) is modeled using an aircraft performance tool, modified to take into account the hybrid nature of the system. The fuel and thermal management systems are modeled looking to maximize the synergistic effects. The electrical system is defined in series with the thermal engine and the performance, in terms of weight and efficiency, are tracked as a function of the cooling temperature.

The results show up to a 46 per cent increase in range and up to 7 per cent gain on a payload with a reference hybrid-electric aircraft that uses conventional drop-in JP-8 fuel. The configuration that privileges a reduction in mass of the electric motors by taking advantage of the cryogenic coolant temperature shows the highest benefits. A sensitivity study is also presented showing the dependency on the modeling capabilities.

The synergistic combination of a cryogenic fuel and the additional heat sources of a hybrid-electric system with a tendency to higher electric component efficiency or reduced weight results in a considerable performance increase in terms of both range and payload.

The potential synergies between a cryogenic fuel and the electrical system of a hybrid-electric aircraft seem clear; however, at the present, no detailed performance evaluation at aircraft level that includes the fuel, thermal management and electric systems, has been published.

The purpose of this paper is to implement a closed loop regulated bidirectional DC to DC converter for an application in the electric power system of more electric aircraft. To provide a consistent power supply to all of the electronic loads in an aircraft at the desired voltage level, good efficiency and desired transient and steady-state response, a smart and affordable DC to DC converter architecture in closed loop mode is being designed and implemented.

The aircraft electric power system (EPS) uses a bidirectional half-bridge DC to DC converter to facilitate the electric power flow from the primary power source – an AC generator installed on the aircraft engine’s shaft – to the load as well as from the secondary power source – a lithium ion battery – to the load. Rechargeable lithium ion batteries are used because they allow the primary power source to continue recharging them whenever the aircraft engine is running smoothly and because, in the event that the aircraft engine becomes overloaded during takeoff or turbulence, the charged secondary power source can step in and supply the load.

A novel nonsingular terminal sliding mode voltage controller based on exponential reaching law is used to keep the load voltage constant under any of the aforementioned circumstances, and its performance is contrasted with a tuned PI controller on the basis of their respective transient and steady-state responses. The former gives a faster and better transient and steady-state response as compared to the latter.

This research gives a novel control scheme for incorporating an auxiliary power source, i.e. rechargeable battery, in more electric aircraft EPS. The battery is so implemented that it can get regeneratively charged when primary power supply is capable of handling an additional load, i.e. the battery. The charging and discharging of the battery is carried out in closed loop mode to ensure constant battery terminal voltage, constant battery current and constant load voltage as per the requirement. A novel sliding mode controller is used to improve transient and steady-state response of the system.

Static inverters are very important for the emergency energy distribution system of aircraft and similar machines. At the same time, the electrical energy produced at high frequency for electrical devices is used to reduce the weight of the cables in the aircraft and spacecraft because of the skin effect. In the high-frequency system, a thinner cable cross-section is used, and a great weight reduction occurs in the aircraft. So, fuel economy, less and late wear of the materials (landing gear, etc.) can be obtained with decreasing weight. This paper aims to present the development of a functional multilevel inverter (FMLI) with fractional sinus pulse width modulation (FSPWM) and a reduced number of switches to provide high-frequency and quality electrical energy conversion.

After the production of FSPWM for FMLI with a reduced component, which, to the best of the authors’ knowledge, is presented for the first time in this study, is explained step by step, and eight operating states are given according to different FSPWMs operating the circuit. The designed inverter and modulation technique are compared by testing the conventional modular multilevel inverter on different loads.

According to application results, it is seen that there is a 50% reduction in cross-section from 100 Hz to 400 Hz with the skin effect. At 1000 Hz, there is a 90% cross-section reduction. The decrease can be in cable weights that may occur in aircraft from 10 kg to 100 kg according to different frequencies. It causes less harmonic distortion than conventional converters. This supports the safer operation of the system. Compared to the traditional system, the proposed system provides more amplitude in converting the source to alternating voltage and increases the efficiency.

FSPWM is developed for multilevel inverters with reduced components at the high frequency and cascaded switching studies in the power electronics of aircraft.

Although the proposed system has less current and power loss as mentioned in the previous sections, it contains fewer power elements than conventional inverters that are equivalent for different hardware levels. This not only reduces the cost of the system but also provides ease of maintenance. To reduce the cable load in aircraft and create more efficient working conditions, 400 Hz alternative voltage is used. The proposed system causes less losses and lower harmonic distortions than traditional systems. This will reduce possible malfunctions and contribute to aircraft reliability for passengers and cargo. As technology develops, it is revealed that the proposed inverter system will be more efficient than traditional inverters when devices operating at frequencies higher than 400 Hz are used. With the proposed inverter, safer operation will be ensured, while there will be less energy loss, less fuel consumption and less carbon emissions to the environment.

The proposed inverter structure shows that it can provide energy transmission for electrical devices in space and aircraft by using the skin effect. It also contains less power elements than the traditional inverters, which are equivalent for different levels of hardware.

This paper aims to automatically generate load shedding sequences due to insufficient power supply, to ensure flight safety and complete flight task.

In this paper, a power allocation and load management model, including logical and physical submodels of the distribution system, is first established according to different requirements of the loads in different flight phase and the current total power supply. Then, an optimal load management scheme based on an improved ant colony algorithm is proposed to automatically generate load shedding sequences for both safety-critical and nonsafety critical loads, to achieve a reliable and safe power supply.

To verify the efficiency and feasibility of the algorithm, the proposed method is verified in a virtual simulation platform. Simulation result illustrates that the proposed algorithm is efficient and feasible.

The proposed method can provide guidance on load power supply when the civil aircraft is under abnormal power supply situation.

An optimal load management scheme is proposed by considering different requirements of the loads in different flight phase and the current total power supply.

This paper aims to analyze the viability of a solar power system as a supplemental power source for commercial and business aircraft.

First, a model is established to estimate the potential available power from suitable aircraft surfaces for various meteorological conditions, ground and flight mission characteristics. A proposed aircraft system architecture and an associated parametric conceptual sizing model are presented. This supplemental solar power system sizing model is integrated into an aircraft multidisciplinary design optimization environment to evaluate the aircraft-level impact on mission fuel burn. A parametric study for a business jet aircraft is performed to analyze various solar cell types and power densities for converters. Trade-off studies are performed between efficiency and weight.

Considering today’s efficiency and power-to-weight ratio of the system components, overall fuel burn reduction can be achieved. Therefore, the technology development work can start now to target short to mid-term applications. In addition, promising system integration scenarios are identified, such as the use of solar power for autonomous operation of the air conditioning system on ground, which yield potential further benefit. In conclusion, a supplemental solar power system seems a promising candidate for more efficient aircraft operation.

The presented novel supplemental solar power system architecture concept and its foreseen aircraft integration show potential benefits for near term applications. The results show that the break even for this technology is already reached and therefore build the foundation to further investigate the technology integration challenges. Clear directions for future research and development are outlined enabling the advancement of the technology readiness level.

This article is aims to inform aircraft propulsion system designers of the implications which fundamental power distribution design assumptions have on the effectiveness and viability of turboelectric distributed propulsion (TeDP) systems. Improvements and challenges associated with selecting alternating or direct current for normal- and superconducting distribution systems are presented. Additionally, for superconducting systems, the benefits of bi-polar DC distribution are discussed, as well as the implications of operating voltage on the mass and efficiency of TeDP grid components.

The approach to this paper selects several high-level fundamental configuration decisions, which must be made, and it qualitatively discusses potential implications of these decisions.

Near term TeDP architectures which employ conventionally conducting systems may benefit from alternating current (AC) distribution concepts to eliminate the mass and losses associated with power conversion. Farther term TeDP concepts which employ superconducting technologies may benefit from direct current (DC) distribution to reduce the cryocooling requirements stemming from AC conduction losses. Selecting the operating voltage for superconducting concepts requires a divergence from the present day criteria employed with terrestrial superconducting transmission systems.

The criteria presented in the paper will assist in the early conceptual architecting of TeDP systems.

The governing principles behind the configuration of multi-MW airborne electrical microgrid systems are presently immature. This paper represents a unique look and the motivating principles behind fundamental electrical configuration decisions in the context of TeDP.

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