Sizing Of Aircraft Research Articles

Initial sizing of a hybrid hydrogen fuel cell commercial aircraft with electric aft-fuselage propulsion

Abstract Hybrid hydrogen fuel cell (HH2FC) aircraft is one of the viable technological pathways to reduce carbon emissions in the aviation industry. For a configuration with aft-fuselage propulsion, a novel method is established for the initial sizing of the HH2FC aircraft in a situation where empirical data is lacking. The carbon emission reduction capabilities and potential for reducing fuel and carbon tax costs can be preliminarily assessed by the proposed initial sizing method. A case study on a narrow-body aircraft indicates that, at the current technological level, the carbon emissions and cost of the HH2FC aircraft will increase to achieve the same payload and range as the conventional aircraft. However, with the improvement of hydrogen technologies, such as fuel cell power density and hydrogen tank gravimetric density, HH2FC aircraft will reduce 6.6% carbon emissions compared to conventional aircraft. As a result, both economic and environmental benefits can be achieved in the future.

Global–local multidisciplinary optimisation for the evaluation of local constraints on finer meshes in preliminary aircraft design

Multidisciplinary design optimisation (MDO) is a methodology increasingly being used in the preliminary design of aircraft. To limit the computational cost of the procedure, it is generally based on coarse models, which do not accurately capture the internal deformation of details with a complex geometry. Therefore, it is not possible to apply constraints in these areas and designers are limited to a conservative pre-sizing of these parts, which are then kept fixed during the optimisation. In this paper we expose the limitations of this approach and present a novel methodology for the preliminary sizing of aircraft, based on global–local MDO. The commonly used coarse model is used together with finer local models, for the parts where additional accuracy is needed. The global–local analysis solves the internal deformation field with sufficient accuracy for the evaluation of local constraints. Furthermore, thanks to the formulation we introduce to compute the coupled sensitivities, the optimiser successfully finds a locally feasible design.

A Bifurcation Theoretic Integrated Methodology for Aircraft Conceptual Design

A bifurcation theoretic integrated methodology for aircraft conceptual design is presented in this paper. The methodology uses the extended bifurcation analysis technique to formulate a unified design framework that can be applied for configuration sizing of aircraft, assessment of aircraft performance and stability, and evaluation of aircraft maneuverability. Unlike conventional approaches for conceptual design, the proposed methodology incorporates a full six degrees-of-freedom flight dynamics model for stability and control analysis. Maneuverability evaluation is done with the help a suitable controller. Configuration sizing of a six-seater general aviation aircraft, followed by performance, stability and maneuverability assessment of the aircraft is successfully carried out in this paper to demonstrate the effectiveness of the proposed methodology.

Overall Aircraft Design: From the open-source FAST-OAD to its training branch

Reducing the environmental footprint of commercial aviation to a manageable level, is a formidable challenge for the whole aeronautical community. In particular, new aircraft configurations could drastically reduce the environmental footprint. Training of engineers in Overall Aircraft Design (OAD) is, therefore, an essential step to be considered in order to take environmental issues into account from the preliminary design phase. ONERA and ISAE-SUPAERO have developed FAST-OAD (Future Aircraft Sizing Tool – Overall Aircraft Design), an open source platform for the analysis, sizing and optimization of aircraft, emphasizing on user-friendliness and modularity. The open source version provides a flexible and practical framework that allows researchers and industrial actors all around the world, to extend the entire aircraft design process to new challenging aircraft configurations, in-house models and higher fidelity approaches. In this contribution, the authors introduce, with a qualitative perspective, a training branch of FAST-OAD, to help trainers to build classes and tutorials.

Aerodynamic Performance Benefits of Over-the-Wing Distributed Propulsion for Hybrid-Electric Transport Aircraft

The goal of this study is to analyze how the aeropropulsive benefits of an over-the-wing distributed-propulsion (OTWDP) system at the component level translate into an aeropropulsive benefit at the aircraft level, as well as to determine whether this enhancement is sufficient to lead to a reduction in overall energy consumption. For this, the preliminary sizing of a partial-turboelectric regional passenger aircraft is performed, and its performance metrics are compared to a conventional twin-turboprop reference for the 2035 timeframe. The changes in lift, drag, and propulsive efficiency due to the OTWDP system are estimated for a simplified unducted geometry using a lower-order numerical method, which is validated with experimental data. For a typical cruise condition and the baseline geometry evaluated in the experiment, the numerical method estimates a 45% increase in the local sectional lift-to-drag ratio of the wing, at the expense of a 12% reduction in propeller efficiency. For an aircraft with 53% of the wingspan covered by the OTWDP system, this aerodynamic coupling is found to increase the average aeropropulsive efficiency of the aircraft by 9% for a 1500 n mile mission. Approximately 4% of this benefit is required to offset the losses in the electrical drivetrain. The reduction in fuel weight compensates for the increase in powertrain weight, leading to a takeoff mass comparable to the reference aircraft. Overall, a 5% reduction in energy consumption is found, albeit with a uncertainty due to uncertainty in the aerodynamic modeling alone.

Sizing of Tilt-Wing Aircraft with All-Electric and Hybrid-Electric Propulsion

This paper demonstrates integrated vehicle and propulsion system sizing and performance analysis using Parametric Energy-Based Aircraft Configuration Evaluator, an aircraft sizing methodology and framework integrating discipline analyses for aerodynamics, propulsion, and weight estimation with parametric geometry definition, resizing, and energy-based mission performance analyses. The framework is used to demonstrate vehicle and propulsion system sizing and analysis for a vertical takeoff and landing tilt-wing urban air mobility aircraft family with two variants. The first variant features an all-electric propulsion system architecture, while the second variant features a hybrid-electric architecture in which turbogenerators are used to supply cruise power requirements, offset battery power draw in vertical flight, and recharge the batteries. Both variants are sized for a representative urban air mobility mission profile while considering varying battery technology levels and trip distances.

Hybrid-electric and hydrogen powertrain modelling for airplane performance analysis and sizing

This paper describes a framework for parametric modelling of hybrid-electric powertrain components for innovative airplane configurations. These models are used in scalability studies and performance analysis of novel propulsion architecture. The methodology involves culmination of these models in a set of tools specifically developed to study the initial and conceptual sizing of hybrid-electric aircraft. This allows quick parametric evaluation of various configurations based on components at different technology readiness levels, such as batteries and fuel cells. Characteristics and performance of the power-train components are evaluated using computational analysis as well as laboratory tests. This information is used to develop numerical models described in the paper and to validate the sizing of fundamental propulsion components. Applications to two variants of a commuter aircraft are given, one using a serial hybrid-electric architecture based on a thermal engine, and the other using a fuel-cell system fed by a gaseous hydrogen tank.

Structural Analysis and Optimization of Wings Subjected to Dynamic Loads

Dynamic loads arising from the Title 14 Code of Federal Regulations (14 CFR; “Aeronautics and Space,” Federal Aviation Administration, U.S. Dept. of Transportation, Subchapter C Aircraft, Part 25 Airworthiness Standards: Transport Category Airplanes, https://ecfr.federalregister.gov/current/title-14/chapter-I/subchapter-C/part-25 [retrieved 23 August 2021]) are critical drivers of aircraft design. The aircraft structural design process is an iterative one involving computationally expensive simulations. Existing methods simplify the dynamic loads to equivalent static ones and design the structure using shell-based analysis. Other literature approaches use beam models for structural sizing but make assumptions on the geometry, material distribution, or both. In this work, the use of the Variational Asymptotic Method is explored to 1) systematically reduce the three-dimensional (3-D) wingbox to a one-dimensional (1-D) model, 2) solve the 1-D model using a nonlinear beam theory, and 3) recover the stress field on the wingbox. An adjoint-driven gradient-based structural optimization framework is developed to size the structure for dynamic loads subjected to strength-based failure considerations. The proposed structural analysis method is applied to the wingbox structural analysis and sizing of a novel hybrid-electric propulsion aircraft: NASA’s Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) concept. Results in the paper show that the sizing of aircraft wings for 14 CFR specified maneuvers 2.5 and static maneuvers using the proposed approach produces a 6% difference in weight compared to a shell-based method but with a speedup of 7.8 times, and it is efficient at sizing the wingbox structure for dynamic loads generated by gust maneuvers.

Vertical tail sizing of propeller-driven aircraft considering the asymmetric blade effect

An engineering approach is presented to analyse the asymmetric blade thrust effect with the help of analytical and semi-empirical methods. It is shown that the contribution of the asymmetric blade thrust effect in the lateral-directional stability of multi-engine propeller-driven aircraft is significant particularly in critical flight conditions with one engine out of service. Also, in some cases where the engines are rotating in one direction, the asymmetric blade effect has substantial effects on the handling qualities of the aircraft even in normal flight conditions. Overall, due to the significant contribution of this phenomenon in the lateral-directional stability of propeller-driven airplanes, it is important to consider it in the design of the vertical stabilizer and rudder. The resulting analytical method has been used to determine the vertical tail incident angle and desired rudder deflection in accordance with the most critical flight condition for two different cases and validated to ensure the accuracy of the result. In this work, the aerodynamic coefficients as well as the stability and control derivatives have been predicted using analytical and semi-empirical methods validated for light aircraft.

Conceptual Design and Energy Storage Positioning Aspects for a Hybrid-Electric Light Aircraft

Abstract This work is a feasibility study of a 19-passenger hybrid-electric aircraft, to serve the short-haul segment within the 200–600 nautical miles. Its ambition is to answer some dominating research questions, during the evaluation and design of aircraft based on alternative propulsion architectures. The potential entry into service (EIS) is foreseen beyond 2030. A literature review is performed to identify similar concepts under research and development. After the requirements' definition, the first level of conceptual design is employed. The objective of design selections is driven by the need to reduce CO2 emissions and accommodate aircraft electrification with boundary layer ingestion engines. Based on a set of assumptions, a methodology for the sizing of the hybrid-electric aircraft is described to explore the basis of the design space, incorporating a parametric analysis for the consideration of boundary layer ingestion effects. Additionally, a methodology for the energy storage positioning is provided to highlight the multidisciplinary aspects between the sizing of an aircraft, the selected architecture (series/parallel partial hybrid), and the storage characteristics. The results show that it is not possible to fulfill the initial design requirements (600 nmi) with a fully-electric aircraft configuration, due to the far-fetched battery necessities. It is also highlighted that compliance with airworthiness standards is favored by switching to hybrid-electric aircraft configurations and relaxing the design requirements (targeted range, payload, battery technology). Finally, the lower degree of hybridization (40%) is observed to have a higher energy efficiency (−12% energy consumption) compared to the higher degree of hybridization (50%) and greater CO2 reduction, with respect to the conventional configuration.

Mass, primary energy, and cost: the impact of optimization objectives on the initial sizing of hybrid-electric general aviation aircraft

For short take-off and landing (STOL) aircraft, a parallel hybrid-electric propulsion system potentially offers superior performance compared to a conventional propulsion system, because the short-take-off power requirement is much higher than the cruise power requirement. This power-matching problem can be solved with a balanced hybrid propulsion system. However, there is a trade-off between wing loading, power loading, the level of hybridization, as well as range and take-off distance. An optimization method can vary design variables in such a way that a minimum of a particular objective is attained. In this paper, a comparison between the optimization results for minimum mass, minimum consumed primary energy, and minimum cost is conducted. A new initial sizing algorithm for general aviation aircraft with hybrid-electric propulsion systems is applied. This initial sizing methodology covers point performance, mission performance analysis, the weight estimation process, and cost estimation. The methodology is applied to the design of a STOL general aviation aircraft, intended for on-demand air mobility operations. The aircraft is sized to carry eight passengers over a distance of 500 km, while able to take off and land from short airstrips. Results indicate that parallel hybrid-electric propulsion systems must be considered for future STOL aircraft.

Impact of Battery Performance on the Initial Sizing of Hybrid-Electric General Aviation Aircraft

Studies suggest that hybrid-electric aircraft have the potential to generate fewer emissions and be inherently quieter when compared to conventional aircraft. By operating combustion engines together with an electric propulsion system, synergistic benefits can be obtained. However, the performance of hybrid-electric aircraft is still constrained by a battery’s energy density and discharge rate. In this paper, the influence of battery performance on the gross mass for a four-seat general aviation aircraft with a hybrid-electric propulsion system is analyzed. For this design study, a high-level approach is chosen, using an innovative initial sizing methodology to determine the minimum required aircraft mass for a specific set of requirements and constraints. Only the peak-load shaving operational strategy is analyzed. Both parallel- and serial-hybrid propulsion configurations are considered for two different missions. The specific energy of the battery pack is varied from 200 to 1,000 W·h/kg, while the discharge time, and thus the normalized discharge rating (C-rating), is varied between 30 min (2C discharge rate) and 2 min (30C discharge rate). With the peak-load shaving operating strategy, it is desirable for hybrid-electric aircraft to use a light, low capacity battery system to boost performance. For this case, the battery’s specific power rating proved to be of much higher importance than for full electric designs, which have high capacity batteries. Discharge ratings of 20C allow a significant take-off mass reduction aircraft. The design point moves to higher wing loadings and higher levels of hybridization if batteries with advanced technology are used.

Approach to the Weight Estimation in the Conceptual Design of Hybrid-Electric-Powered Unconventional Regional Aircraft

The present work deals with the development of an innovative approach to the weight estimation in the conceptual design of a Hybrid-Electric-Powered (HEP) Blended Wing Body (BWB) commercial aircraft. In the last few decades, the improvement of the environmental impact of civil aviation has been the major concern of the aeronautical engineering community, in order to guarantee the sustainable development of the system in presence of a constantly growing market demand. The sustained effort in the improvement of the overall efficiency of conventional aircraft has produced a new generation of vehicles with an extremely low level of emissions and noise, capable of covering the community requirements in the short term. Unfortunately, the remarkable improvements achieved represent the asymptotic limit reachable through the incremental enhancement of existing concepts. Any further improvement to conform to the strict future environmental target will be possible only through the introduction of breakthrough concepts. The aeronautical engineering community is thus concentrating the research on unconventional airframes, innovative low-noise technologies, and alternative propulsion systems. The BWB is one of the most promising layouts in terms of noise emissions and chemical pollution. The further reduction of fuel consumption that can be achieved with gas/electric hybridisation of the power-plant is herein addressed in the context of multidisciplinary analyses. In particular, the payload and range limits are assessed in relation to the technological development of the electric components of the propulsion system. The present work explores the potentialities of an energy-based approach for the initial sizing of a HEP unconventional aircraft in the early conceptual phase of the design. A detailed parametric analysis has been carried out to emphasise how payload, range, and degree of hybridisation are strictly connected in terms of feasible mission requirements and related to the reasonable expectations of development of electric components suitable for aeronautical applications.

Integrated Sizing and Optimization of Aircraft and Subsystem Architectures in Early Design

The aerospace industry’s current trend towards novel or More Electric architectures results in some unique challenges for designers due to both the scarcity or absence of historical data and a potentially large combinatorial space of possible architectures. These add to the already existing challenges of attempting to optimize an aircraft design in the presence of multiple possible objective functions while avoiding an overly compartmentalized approach. This paper uses the Integrated Subsystem Sizing and Architecture Assessment Capability to pursue a multi-objective optimization for a large twin-aisle aircraft and a small single-aisle aircraft using the Non-Dominated Sorting Genetic Algorithm II (NSGA-II) algorithm with parallel function evaluations. One novelty of the optimization setup is that it explicitly considers the impacts of subsystem architectures in addition to those of traditional aircraft-level design variables. The optimization yields generations of nondominated designs in which substantially electrified subsystem architectures are found to predominate. As a first assessment of the impact of epistemic uncertainty on the results obtained, the optimization is rerun with altered sensitivities for the thrust-specific fuel consumption penalties due to shaft-power and bleed air extraction. This analysis demonstrated that the composition of architectures on the Pareto frontier is sensitive to the secondary power extraction penalties, but more so for the small single-aisle aircraft than the large twin-aisle aircraft.

Conceptual sizing of next supersonic passenger aircraft from regression of the limited existing designs

Despite previous versions, there are no current supersonic passenger transport aircraft. Much aircraft research is focused on hypersonic flight and the new technologies therein and is therefore unlikely to add to commercial versions anytime soon. This study re-examines conceptual sizing of a supersonic transport aircraft based on extant supersonic designs in order to ignite research into whether a commercially-viable design might exist. Key metrics are developed using distances between likely airport network nodes, an assumed number of passengers, and a reduction in transport time to one-third of current journeys. The study uses multiple response regression of known designs to develop key performance formulae, which are then optimized to set performance values so as to estimate an initial aircraft size, including an expected value analysis to guide the next conceptual design iteration. Twenty years ago a NASA Langley Research optimization system was used to examine non-linear regression of supersonic aircraft designs and to optimize such a design around similar performance criteria. In contrast, this work is the first supersonic transport aircraft sizing to use commercially-available Excel add-on software and standard design-for-sixsigma analysis techniques; notably for the sensitivity analyses to guide the next design iteration.

Preliminary weight sizing of light pure-electric and hybrid-electric aircraft

Abstract The lack of consolidated preliminary design techniques coping with the characteristics of most recent electric and hybrid-electric power plants is often an obstacle for aircraft manufacturers and for owners and operators as well, making the design process less straightforward and hampering comparisons with respect to more traditional designs. In this paper, a technique for the preliminary weight sizing of electric aircraft in the General Aviation category is explained. This is based on existing procedures typical to conventionally-powered aircraft, integrated in a common framework to suitably tackle the issues raised by the peculiar features of electrically-powered aircraft. Then, an expansion of the design method to the case of a series hybrid propulsion system is investigated. Results in virtual environment on a realistic design are also presented.

PANDORA - A python based framework for modelling and structural sizing of transport aircraft

Over the last years a multidisciplinary aircraft predesign process chain was established at the DLR, including different numerical tools for the modelling and structural sizing of fuselage structures. To improve the flexibility and performance of this structural analysis part in the MDO process a new tool development has been started in 2016 called “Parametric Numerical Design and Optimization Routines for Aircraft” (PANDORA). The PANDORA framework is using the interpreted high-level programming language Python and is focused on using dedicated open-source packages. Within PANDORA a lot of new packages have been implemented, like a new interface to access CPACS data, a python based FE pre- and postprocessor, a FE data converter to build an interface between PANDORA and different FE solver and a visualization interface using “The Visualization Toolkit” (VTK). Some further packages to generate a FE model based on a CPACS file using the geometry core “Open Cascade” (OCC) and a new FE sizing algorithm is also under development. To simplify the usage of PANDORA and to keep it comprehensible - a graphical user interface (GUI) has been added using the PYQT toolkit. In this paper the current state of the PANDORA development is presented and initial applications are shown.

Optimal performance and sizing of a battery-powered aircraft

This paper presents an analytical and experimental framework for investigating the performance of fixed-wing electrically-driven propeller aircraft. After the experimental derivation of a novel constant power discharge model for lithium-polymer (Li-Po) batteries, closed-form expressions for flight endurance and range are derived by equating power required in steady level flight with power supplied by the battery pack. Best endurance and best range airspeed are obtained. A methodology is also described to optimally size the battery capacity for a given set of battery and aircraft characteristics. The proposed approach is validated by application to a test case relative to a small-size unmanned aerial vehicle.

An integrated approach to the preliminary weight sizing of small electric aircraft

Electric propulsion has received attention in aviation as witnessed by studies in hybrid designs and by the production of aircraft with support electric motors to be used in limited parts of the mission with ancillary roles. Until the recent past, the main limit to a wider adoption of electric propulsion, which besides having a lower environmental impact with respect to internal combustion engines (ICE) in terms of noise and emissions, can also improve reliability and on-board comfort, was the need for mass and volume-inefficient battery packs as devices for energy storage. However, thanks to the level of technology now reached by batteries, it is becoming possible to design and build electrically propelled aircraft at least in the category of light or general aviation. Due to the relative novelty of this technology, only few examples of similar aircraft exist today, mainly modifications of more traditional concepts, and thinking of a completely new electric aircraft is made difficult by the lack of a consolidated design framework, differently from the case of traditional ICE-powered models. This paper tries to cope with some basic aspects typical to electrically propelled aircraft, to the aim of setting up a stable and reliable preliminary sizing procedure allowing designers and aircraft companies to quickly size up and compare all-electric designs. To this aim, a statistical analysis of the basic characteristics of existing aircraft is presented first, showing a good correlation level between some of them. Next a method for the preliminary sizing of weights is shown, obtained starting from a more usual step-by-step procedure typically adopted for ICE-propelled aircraft. Due to the peculiar characteristics of electrically powered aircraft, the new procedure involves an integrated use of the case-specific mission profile and sizing matrix. The validity of the proposed procedure is testified by example analyses on two realistic designs of lightweight aircraft.

Optimization of Aircraft Lifting Surfaces Using a Cellular Division Method

The development of a biologically inspired methodology for topology, shape, sizing, and control surface optimization of aircraft lifting surfaces is presented. The methodology is based on the map L-systems modeling of cellular division to generate the substructure topology. This is combined with variables for aerodynamic shape, structural sizing, and control surfaces (number, size, location, and aeroelastic trimmed settings) and constraints on stiffness, strength, local skin panel buckling, static aeroelastic response (roll performance, pitch rate, trimmed angle of attack), and flutter requirements. A dual-objective function is formulated with weight and and is solved with a bilevel optimization algorithm. The map -system rules that develop the topology are evolved using a genetic algorithm in the outer optimization loop on and weight, which obtains the optimal parameter settings for topology, shape, and control surfaces, whereas the inner loop performs weight minimization with the structural sizing variables. The methodology is demonstrated on the design of a generic fighter aircraft wing.