Turboelectric Propulsion System Research Articles

Optimization of Lift-Plus-Cruise Vertical Take-Off and Landing Aircraft with Electrified Propulsion

Conventional aircraft sizing methods face challenges in analyzing all-electric or hybrid-electric novel aircraft configurations, such as those for urban air mobility applications. The vast design space containing both continuous and discrete design variables and competing design objectives necessitates searching for not necessarily a unique optimal design but rather an array of Pareto-optimal designs. This paper uses the Parametric Energy-Based Aircraft Configuration Evaluator, an aircraft sizing framework for novel aircraft and propulsion system architectures, to pursue multi-objective optimization of a lift-plus-cruise urban air mobility aircraft with all-electric, hybrid-electric, and turbo-electric propulsion system architectures using Nondominated Sorting Genetic Algorithm II. The optimization cases considered include multiple objective functions, such as maximum takeoff mass, mission time, energy and propulsion mass fraction, and energy used per unit distance per unit payload. Optimization was performed for trip distances of 80, 120, and 150 km and battery specific energy levels of 350 and 400 Wh/kg. Except when mission time was an objective function, Pareto-optimal designs occurred near the lower bound of the velocity range. Raising that bound had an impact on the architectural composition of the final generation. All-electric architectures appeared exclusively for lower trip distances, hybrid-electric designs appeared for longer trip distances, and turbo-electric designs only appeared for combinations of longer trip distances, a higher minimum cruise speed, and a lower battery specific energy level. The sizing results were sensitive to assumptions regarding the overload capacity of lift propulsor motors in postfailure conditions.

Design Space Exploration of Lithium-Ion Battery Packs for Hybrid-Electric Regional Aircraft Applications

Distributed hybrid and electric propulsion systems are one of the most promising technologies to reduce aircraft emissions, resulting in research efforts to investigate new architectures and the design of optimal energy management strategies. This work defines the optimal requirements in terms of battery pack sizing and cell technology for a hybrid-electric regional wing-mounted distributed propulsion aircraft through the application of a design space exploration method. The propulsion system considered in this study is a series-parallel hybrid turboelectric power train with distributed electric fans. A set of six lithium-ion battery cell technologies was identified and experimentally characterized, including both commercially available and prototype cells at different combinations of specific energy and power. A model of the aircraft was developed and used to define the optimal energy management strategy for the hybrid turboelectric propulsion system, which was solved using dynamic programming. The design space exploration was conducted by varying the cell technology and battery storage system size; and the effects on fuel consumption, energy management strategy, and thermal management were compared.

Electric Propulsors for Zero-Emission Aircraft: Partially superconducting machines

Be it through combustion in a turbine or via direct conversion with fuel cells, hydrogen electric is considered the most attractive path toward a future with zero-emission aircraft. An internal combustion engine, with hydrogen as the fuel, results in very low amounts of nitrogen oxide emissions, providing an attractive option for a turboelectric propulsion system. Additionally, new lean combustion strategies show promising results to further minimize these emissions. On the other hand, a fuel cell-based electrical architecture, with no direct emissions, can be used to power a fully electric propulsion system. Whichever method is used, electric machines play a key role in hydrogen-powered aircraft as it is the unit that efficiently converts electrical power to mechanical power. With promising trends pointing toward hydrogen as the key to electrification of large transport aircraft, efficient and “flightweight” megawatt (MW)-class motors and, in some cases, tens of MW are needed. NASA’s Advanced Air Transport Technology (AATT) program anticipated this and launched several NASA Research Announcements, leading to the continued development and availability of enabling technologies. Other agencies are also stepping up to the challenge with integral R&D, such as thermal management-related work supported by Advanced Research Projects Agency–Energy. In this article, we discuss technology options, developments, challenges, and trends toward an electric propulsor that will power the future of aviation.

Preliminary Design of Integrated Partial Turboelectric Aircraft Propulsion Systems

This paper presents a novel tool for the modeling of partial turboelectric propulsion systems together with a corresponding case study for a commercial single-aisle aircraft. In order to reduce the environmental impact of air traffic, radically new aircraft and propulsion concepts with a high market penetration are needed. Partial turboelectric propulsion systems seem to offer a promising option to achieve this. For the development of these propulsion systems, a preliminary design tool with a homogeneous and sufficiently high fidelity, both for turbomachinery and electric components, is needed. To address this, the authors of this publication have developed a tool based on the GasTurb software. The models developed, in particular for the electric components which together form the electric powertrain, are described here. In the case study, which demonstrates the coupling of the developed tool with an aircraft design environment, a conventional turboprop baseline aircraft is compared to a derived aircraft which features a partial turboelectric propulsion system with wingtip propellers. The latter are intended to reduce the induced drag, enabling a reduction of the aircraft's total shaft power demand compared to the conventional baseline aircraft. The comparison between the partial turboelectric aircraft and the baseline aircraft indicates that fuel reduction increases with power split. However, primarily increasing electric powertrain masses and a stagnating drag reduction result in lower additional fuel reductions for higher power splits. Despite these conclusions, the predicted induced drag reductions need further refinement as they were found to be optimistic. In summary, this publication presents a methodology and a set of physics-based component models for the preliminary design of partial turboelectric propulsion systems, so that the electric components can be investigated and optimized at the same high level of detail as the gas turbine.

Evaluation of a Regional Aircraft with Boundary Layer Ingestion and Electric-Fan Propulsor

This paper presents a conceptual study of a regional aircraft with a turboelectric propulsion system and boundary-layer ingestion. The aircraft has an additional electric propulsor installed at the aircraft tail cone, which is driven by generators installed on both underwing engines, aiming to ingest the fuselage boundary layer and improve the aircraft overall performance. The proposed configuration is an aircraft reengining of the reference aircraft platform, the Embraer 175-E1, targeting minimizing airframe modifications. Parametric variations of the thrust split ratio as well as the electric fan pressure ratio were performed in order to create a design space of possible solutions for the proposed concept and also to provide insights of the aircraft key variables trends. From that, a feasible and optimized configuration in terms of efficiency was selected and compared to the reference aircraft. As the main conclusion, it was determined that the studied propulsion system has the potential to provide specific air range benefits in the order of 4 to 7%, which may not be enough to justify a new development.

Development of a Turbo Electric Distribution System for Remotely Piloted Aircraft Systems

Turboelectric distributed propulsion systems are paving the way for more electric aircraft systems (TeDP). This type of system provides a solution for some of the drawbacks of current low-energy-density batteries, which limit the ability of long-endurance electric aircraft. However, turboelectric propulsion requires the use of advanced turboelectric motors, superconductive materials and cryogenic cooling technologies, which are still under development and may be in production in the near future. This paper investigates a turboelectric propulsion system that can be considered an initial step in the production of TeDP in a remotely piloted aircraft system with the use of existing technology. This is achieved by replacing the gear and the starter motor of a turboprop with a high-speed permanent magnet electric machine to generate electrical power and propelling the aircraft through a distributed electric propulsion system. In this theoretical study, an augmentation to Breguet’s range and endurance equation is developed. This study confirmed that the new system is 31% lighter than the turboprop engine. Then the effect of the weight savings is used in the distributed electric propulsion (DEP) aerodynamic studies and found that there is a drastic increase in the range for a TeDP developed with the high-speed machine.

Performance analysis of turbo-electric propulsion system with fuselage boundary layer ingestion

A propulsion system analysis of an aircraft concept featuring a fuselage tail cone integrated turbo-electrically powered fan is presented. The aircraft has two underwing podded geared turbofan engines and an aft-fuselage mounted boundary layer ingesting fan. The fuselage fan is driven by an electric motor that is powered by power offtake from the main power plants under the wing. A long-range tube-and-wing aircraft with a year 2050 entry into service is used as a reference. A coupled multidisciplinary method for system level assessment of the turbo-electric boundary-layer ingesting propulsion system is presented. A correlation-based method is used to predict fuselage drag and a series of optimizations are carried out for a range of fuselage fan diameters. The optimal level of ingested drag is 30%-57% of the total fuselage drag, resulting in a net reduction in mission fuel burn of 0.6%-3.6% depending on technology assumptions. Further analysis reveals that installation effects, mainly increased mass, offset some of the gains of boundary layer ingestion for the smallest fuselage fan sizes. For larger fan sizes, it is instead the losses in the electric machinery together with the lower efficiency of the fuselage fan, compared to the main engine fan, that compensate for the gains from boundary layer ingestion. However, the by far strongest effect for determining the optimal level of ingested drag is that it is very difficult to obtain a net benefit from ingesting the half of the total fuselage drag contained in the outer part of the boundary layer. The benefit is outweighed by losses in the electric transmission system, installation effects and efficiency deficits of having an aft-mounted fan instead of larger size under-wing main engines. This is true also under the assumption of radical technology such as superconducting electric machinery. It is concluded that the studied aircraft architecture, despite having a high theoretical potential, faces a large difficulty in beneficially ingesting the significant amount of the fuselage drag contained in the outer part of the boundary layer. This severely limits its potential to substantially reduce the fuel burn compared to a conventional twin-engine tube-and-wing aircraft in the year 2050 timeframe.

Baseline architecture design for a turboelectric distributed propulsion system using single turboshaft engine operational scenario

Abstract Reducing the carbon footprints of aerial transportation became a major target for both industries and academia. Various solutions have been proposed to develop cleaner alternative methods for green transportation. Full electric, hybrid electric, and turboelectric propulsion system architectures intend to reduce greenhouse gas emissions and fuel consumption of today’s aero gas turbine engines. In this study, a turboelectric propulsion system, which is considered as the most promising technology for future aviation is selected for modelling and simulation. As the main power supply, a high fidelity mathematical model of GE T700 turboshaft engine is constructed in MATLAB/Simulink to emulate the technology of today. Selected aero gas turbine’s mathematical model is combined with NASA’s Baseline electrical power distribution architecture which is firstly designed for N-3X turboelectric aircraft. MATLAB/Simulink model is utilized to analyses a single-engine operational scenario of twin-engine aircraft which is a major design consideration due to single-engine failure. Power requirements, distribution percentages, preliminary power assessment for power electronic systems and nominal power capacities of each electrical unit of a turboelectric propulsion system are obtained using GE T700 as the main power supply.

Baseline architecture design for a turboelectric distributed propulsion system using single turboshaft engine operational scenario

Abstract Reducing the carbon footprints of aerial transportation became a major target for both industries and academia. Various solutions have been proposed to develop cleaner alternative methods for green transportation. Full electric, hybrid electric, and turboelectric propulsion system architectures intend to reduce greenhouse gas emissions and fuel consumption of today’s aero gas turbine engines. In this study, a turboelectric propulsion system, which is considered as the most promising technology for future aviation is selected for modelling and simulation. As the main power supply, a high fidelity mathematical model of GE T700 turboshaft engine is constructed in MATLAB/Simulink to emulate the technology of today. Selected aero gas turbine’s mathematical model is combined with NASA’s Baseline electrical power distribution architecture which is firstly designed for N-3X turboelectric aircraft. MATLAB/Simulink model is utilized to analyses a single-engine operational scenario of twin-engine aircraft which is a major design consideration due to single-engine failure. Power requirements, distribution percentages, preliminary power assessment for power electronic systems and nominal power capacities of each electrical unit of a turboelectric propulsion system are obtained using GE T700 as the main power supply.

Comparative Analysis and Optimization of Technical and Weight Parameters of Turbo-Electric Propulsion Systems

According to Flightpath 2050, the aviation industry is aiming to substantially reduce emissions over the coming decades. One possible solution to meet these ambitious goals is by moving to hybrid-electric drivetrain architectures which require the electric components to be extremely lightweight and efficient at the same time. It has been claimed in several publications that cryogenic and in particular superconducting components can help to fulfill such requirements that potentially cannot be achieved with non-cryogenic components. The purpose of this work was to make a fair comparison between a cryogenic turbo-electric propulsion system (CEPS) and a non-cryogenic turbo-electric propulsion system (TEPS) on a quantitative level. The results on the CEPS were presented in detail in a previous publication. The focus of this publication is to present the study on the TEPS, which in conclusion allows a direct comparison. For both systems the same top-level aircraft requirements were used that were derived within the project TELOS based on an exemplary mission profile and the physical measures of a 220-passenger aircraft. Our study concludes that a CEPS could be 10% to 40% lighter than a TEPS. Furthermore, a CEPS could have a total efficiency gain of up to 18% compared to a similar TEPS.

An Approach to Maximize Torque Density in a Brushless Doubly Fed Reluctance Machine

Brushless doubly fed reluctance machines (BDFRM) are an attractive option for megawatt-scale turbo-electric propulsion systems due to use of a partially rated power converter, reduced maintenance, and absence of permanent magnets. Conventional design approach of a BDFRM uses equal electrical loadings on both stators and an initial phase offset $\delta$ of $\pi /2$ between the two stator currents. However, this approach leads to poor torque density because the machine saturates even before the stator reaches the current density limit. This article proposes an approach to maximize the torque density by finding appropriate electrical excitations on both stator windings for chosen machine dimensions as well as for specific flux density, current density, and grid voltage limits. A single-objective optimization problem is formulated. The optimized solution establishes that an initial phase offset $\delta$ of $2\pi /3$ produces 75% more torque as compared to the conventional $\delta = \pi /2$ design. The new design allows a higher stator currents injection without saturating the machine. The results are verified through extensive simulations using a finite element model. Operating with the optimized design also leads to higher machine efficiency and better power factor on the secondary stator, thus reducing the converter rating.

Conceptual Design and Electromagnetic Analysis of 2 MW Fully Superconducting Synchronous Motors With Superconducting Magnetic Shields for Turbo-Electric Propulsion System

With air traffic predicted to increase in the future, CO2 emission must be decreased to prevent global warming. Our research group previously proposed electric propulsion systems based on superconducting technology, which is promising for reducing emissions. In this study, propulsive 2 MW fully superconducting motors were conceptually designed. At operating temperatures of 20 and 64 K, we observed the influence of the superconducting magnetic shield on the electromagnetic properties, especially the output power density, AC loss, and efficiency. The superconducting magnetic shield can reduce leakage magnetic field to the outside of the motor by inducing a current, and the shield is reproduced with a conductor of extremely high conductivity for simplicity in electromagnetic analysis. As a result, at an operating temperature of 20 K, the superconducting motors with the superconducting magnetic shields showed a high power density greater than 40 kW/kg and high efficiency of 99.7%. Even at an operating temperature of 64 K, the motors with the superconducting shield showed a power density and efficiency of 20 kW/kg and 99.4%, respectively.

Conceptual Design and Numerical Analysis of 10 MW Fully Superconducting Synchronous Generators Installed With a Novel Casing Structure

Currently, feasibility studies on fully turboelectric propulsion systems for electric aircrafts are being conducted worldwide. A fully superconducting rotating machine can realize fully turboelectric propulsion systems with light weight and high power density. In the simple casing model considered in a previous study, the field winding forms a large percentage of the generator weight. Therefore, to reveal the relationship between the operation temperature of the field winding and output power density, we compared an inner cryostat model having a field winding operation temperature of 20 K with the simple casing model having a field winding operation temperature of 64 K in terms of the output power density, efficiency, and increasing temperature. The stator and rotor rooms of the inner cryostat model were insulated by introducing the vacuum layer. The inner cryostat casing structure can have a lower refrigerant gas temperature, e.g., 20 K. Consequently, the temperature of the simple model increased from 5.4 to 7.1 K and that of the inner cryostat model increased from 3.0 to 6.1 K. The simple and inner cryostat models’ power densities were 18.8 and 21.7 kW/kg, respectively. Both models achieved 99% efficiency. No major difference was observed between both models in terms of efficiency and output power density.

Assessment of a Turbo-Electric Aircraft Configuration with Aft-Propulsion Using Boundary Layer Ingestion

In this paper, a turbo-electric propulsion system was analyzed, and its performance was assessed. The aircraft considered here was a single-aisle, medium-range configuration targeting a capacity of 150 Pax. The propulsion concept comprised two boosted geared turbofan engines mounted under-wing. Those main engines were supported by an electrically driven aft-propulsor contributing to the thrust generation and by taking advantage of ingesting the boundary layer of the fuselage for potentially higher levels of propulsive efficiency and allowing for the improved operation of the main engines. The performance assessment as carried out in the context of this paper involved different levels: Firstly, based on the reference aircraft and the detailed description of its major components, the engine performance model for both main engines, as well as for the electrically driven aft-propulsor was set up. The methodology, as introduced, has already been applied in the context of hybrid-electric propulsion and allowed for the aforementioned aircraft sizing, as well as the subsequent gas turbine multi-point synthesis (simulation). A geared turbofan architecture with 2035 technology assumptions was considered for the main engine configuration. The present trade study focused on the design and performance analysis of the aft-propulsor and how it affected the performance of the main engines, due to the electric power generation. In order to allow for a more accurate description of the performance of this particular module, the enhanced streamline curvature method with an underlying and pre-optimized profile database was used to design a propulsor tailored to meet the requirements of the aft propulsor as derived from the cycle synthesis and overall aircraft specification; existing design expertise for novel and highly integrated propulsors could be taken advantage of herein. The resulting performance characteristics from the streamline curvature method were then fed back to the engine performance model in a closely coupled approach in order to have a more accurate description of the module behavior. This direct coupling allowed for enhanced sensitivity studies, monitoring different top-level parameters, such as the thrust/power split between the main engines and the aft propulsor. As a result, different propulsor specifications and fan designs with optimal performance characteristics were achieved, which in return affected the performance of all subsystems considered.

Design Study of 10 MW REBCO Fully Superconducting Synchronous Generator for Electric Aircraft

Future aircraft are expected to have low noise pollution, low emission, and low fuel consumption. A fully turbo-electric propulsion system was proposed by NASA with these aims. The advantages of fully superconducting rotating machines are their weight and efficiency, owing to a high current density resulting in no iron core and no joule heat. In this paper, 10-MW fully superconducting generators with REBa 2 Cu 3 Oy (RE: Rare Earth, Y, Gd, Eu, and others, REBCO) superconducting tapes were studied in relation to their output power density and efficiency. The target output power density is over 20 kW/kg. Fully superconducting generators of 10 MW were designed, then a numerical simulation of the rated operation was conducted. The ac loss of the superconducting windings, iron loss of the yoke, total efficiency, dry weight, and output power density were evaluated, and their dependence on the operating temperature and the synchronous rotating speed were investigated. Here laser-scribing of the wires into a multifilamentary structure and transposed parallel conductors were introduced for ac loss reduction and current capacity enhancement. Consequently, the efficiency reached 99.2% in the case where the REBCO superconducting tapes were scribed into a 10-filament structure. The best model achieved a power density of 26.7 kW/kg. The iron yoke accounts for approximately half the total weight of all the models. The result suggests that the development of superconducting shields brings significant improvement to the output power density.

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.

Coupled aeropropulsive design optimisation of a boundary-layer ingestion propulsor

AbstractAirframe–propulsion integration concepts that use boundary-layer ingestion (BLI) have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA’s STARC-ABL aircraft configuration, which offers the potential for fuel burn reduction by using a turboelectric propulsion system with an aft-mounted electrically driven BLI propulsor. So far, attempts to quantify this potential fuel burn reduction have not considered the full coupling between the aerodynamic and propulsive performance. To address the need for a more careful quantification of the aeropropulsive benefit of the STARC-ABL concept, we run a series of design optimisations based on a fully coupled aeropropulsive model. A 1D thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation to model the aft propulsor at a cruise condition and the effects variation in propulsor design on overall performance. A series of design optimisation studies are performed to minimise the required cruise power, assuming different relative sizes of the BLI propulsor. The design variables consist of the fan pressure ratio, static pressure at the fan face, and 311 variables that control the shape of both the nacelle and the fuselage. The power required by the BLI propulsor is compared with a podded configuration. The results show that the BLI configuration offers 6–9% reduction in required power at cruise, depending on assumptions made about the efficiency of power transmission system between the under-wing engines and the aft propulsor. Additionally, the results indicate that the power transmission efficiency directly affects the relative size of the under-wing engines and the aft propulsor. This design optimisation, based on computational fluid dynamics, is shown to be essential to evaluate current BLI concepts and provides a powerful tool for the design of future concepts.

Modeling Boundary Layer Ingestion Using a Coupled Aeropropulsive Analysis

The single-aisle turboelectric aircraft with an aft boundary layer propulsor aircraft concept is estimated to reduce fuel burn by 12% through a turboelectric propulsion system with an electrically driven boundary layer ingestion propulsor mounted on the fuselage tail cone. This motivates a more detailed study of the boundary layer ingestion propulsor design, which requires considering the fully coupled airframe propulsion integration problem. However, boundary layer ingestion studies thus far have accounted only for the aerodynamic effects on the propulsion system, or vice versa, but have not considered the two-way coupling. This work presents a new approach for building fully coupled aeropropulsive models of boundary layer ingestion propulsion systems. A one-dimensional thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation. This aeropropulsive model is used to assess how the propulsor design affects the overall performance for a simplified model of the single-aisle turboelectric aircraft with an aft boundary-layer propulsor. The results indicate that both propulsion and aerodynamic effects contribute equally toward the overall performance and that the fully coupled model yields substantially different results compared to the uncoupled model. Although boundary layer ingestion offers substantial fuel burn savings, it introduces propulsion-dependent aerodynamic effects that need to be accounted for.

Implications of distribution voltage configuration on TeDP electrical systems

Purpose – 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. Design/methodology/approach – 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. Findings – 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 pow...

Development of a 3D Sizing Model for All-Superconducting Machines for Turbo-Electric Aircraft Propulsion

Distributed propulsion in aircraft brings many advantages in terms of efficiency and noise reduction. While the distribution can be done mechanically through the use of gears and transmissions, electrical propulsion allows for lower maintenance needs, higher efficiency, and lower emissions through the complete decoupling of the gas turbines and the propulsion fans. Such systems have been investigated in the past and NASA is executing on a development plan to bring turbo-electric propulsion systems for transportation aircraft by 2035. The very high specific power required for the airborne generators and motors can only be achieved by using superconductors. Analytical 2-D sizing models have been created and showed very promising results. NASA is now funding the development of higher fidelity models for superconducting machines in which an actual 3-D representation of the geometry is considered. The magnetic flux distribution is calculated using Biot-Savart's law coupled with the magnetic moment method for the backiron. The code also includes thermal and mechanical models allowing for a full and accurate design. The paper describes the model architecture and the methods used to perform high-temperature superconducting machine sizing and optimization.