Preliminary Aircraft Design Research Articles

A simplified automated approach for a preliminary safety assessment of distributed electric and hybrid propulsion systems

AbstractTo reduce the environmental impact of future aircraft, the usage of disruptive hybrid electric propulsion systems, in conjunction with distributed propulsion systems, is envisioned. The complexity of those novel powertrain concepts increases significantly. Besides nominal system characteristics, this particularly applies for off-nominal behaviour and consequences. As the safety impact on aircraft level is not always obvious, the suitability of a given component failure rate has to be analysed already in a preliminary design stage. Within this work, a method shall be introduced that allows a simplified automated safety assessment of any powertrain architecture. All relevant combinations of single and multiple failures are identified and a combined failure rate is derived. Thereafter, the impact of single and multiple component failures is evaluated on aircraft level. A selection of metrics for top-level-aircraft functions is used to implement an impact-driven analysis for each relevant failure case. The metrics used assess the adverse effects on climb performance, lateral controllability and range degradation. For all failure combinations, these effects can be validated against requirements imposed on the configuration. The possibility to formulate failure rate-related requirements follows the idea to target an equivalent level of safety compared to existing regulations, independently of the nature of the failure scenario. Results of this study show the necessity for a reevaluation of the configurations assessed, either regarding their architecture or the assumed component failure rate. The method allows an early detection of configuration-specific shortcomings of complex powertrain architectures already in a preliminary aircraft design stage.

Improved modelling of low-pressure rotor speed in commercial turbofan engines: A comprehensive analysis of machine learning approaches

Accurate engine thrust modelling offers notable opportunities for reducing fuel consumption, mitigating emission effects, managing air traffic, and optimizing the preliminary design of aircraft. This study focuses on the development of machine learning models to predict the low-pressure rotor speed (N1) parameter, which is closely related to the thrust of turbofan engines, for commercial aircraft during the cruise phase. The analysis utilizes a Flight Data Recorder (FDR) dataset from 1086 actual flights involving twin-engine, narrow-body, short-to-medium haul commercial aircraft equipped with CFM56-7B turbofan engines. To achieve accurate N1 parameter predictions, the study employs machine learning models, including Deep Learning (DL), Random Forest (RF), Gradient Boosted Machines (GBM), and Artificial Neural Networks (ANN). These models are evaluated using statistically significant indicators, demonstrating that machine learning models yield highly accurate results. Among the models tested, the DL model provides the most precise estimates of the N1 parameter. This study not only advances the understanding of engine thrust modelling through machine learning but also provides practical insights for the aerospace industry. The findings underline the potential of machine learning techniques in delivering superior prediction accuracy, which can be integrated into real-time flight management systems, ultimately contributing to more sustainable and efficient aviation operations.

Knowledge-based model generation for aircraft cabin noise prediction from pre-design data

The climate targets set out in the “European Green Deal” call for the consideration and implementation of climate-friendly propulsion concepts and sustainable fuels in future aircraft configurations. This puts the use of very efficient propeller engines into the focus of aircraft design. However, these pose major challenges, especially to cabin acoustics, due to high, tonal sound pressure levels on the fuselage. The design of noise control measures requires the competence to predict the resulting interior noise as early as possible in the design process of the vehicle. This requires a high level of geometric and structural detail regarding the fuselage structure and cabin components. With increasing frequency range, the necessary structural details also increase, which have to be resolved in the simulation models due to the associated decreasing structural wavelength. Especially in the context of aircraft pre-design, there is usually not enough information available for a detailed vibro-acoustic modeling of the fuselage and cabin components, which allows a meaningful prediction of the vibrations and thus the cabin noise. Therefore, the knowledge-based tool Fuselage Geometry Assembler (FUGA) is developed for the targeted enrichment of preliminary design data with knowledge for detailed numerical analyses. This paper describes the knowledge-based geometry and model generation in FUGA, which can consider the necessary (increasing) level of detail for the vibro-acoustic prediction already in the preliminary design. For this purpose, aircraft data sets in the preliminary design data format Common Parametric Aircraft Configuration Schema (CPACS) form the modeling basis. Originating from the aircraft preliminary design, these initially describe the outer shell of the vehicle and are extended by detailed structural information that defines the geometric boundary conditions for component placement in cabin design. For the cabin components, the open-source geometry kernel Open Cascade Technology (OCCT) is used to provide geometries at the level of detail required for subsequent analyses. The geometry models are then discretized in open source (e.g. Gmsh) or commercial meshers and further used for numerical analysis. Finally, the prediction of cabin noise is demonstrated as a Proof of Concept using the example of a short-haul propeller-driven aircraft and the feasibility of the proposed method is indicated by investigating the sensitivity of resulting simulation models to the fuselage skin thickness.

An efficient PanAir integrated framework for automated analysis

The work proposed here is an automated pre and post-processor integrated to PanAir that is is a high-order aerodynamic panel method-based software for flow analysis developed in 70s but still in active use especially for preliminary aircraft design. With the integrated environment proposed in this work, manipulation of input and output data to and from PanAir is bypassed successfully that is otherwise requires manual manipulations and use of third party software. The integrated environment is validated over a Cessna 210 aircraft with a modified NLF (1)-0414 airfoil. The flow around the aircraft is analyzed using PanAir together with the integrated environment and results show that pre and post processing times reduced and ease in PanAir use is increased significantly.

Integrated Model for Predicting Demand of Supersonic Transports Under Low-Boom Constraint

Commercial supersonic aircraft could return between 2035 and 2040 due to the desire for fast travel and advancements in technology that could reduce the cost of supersonic transport. Nevertheless, several challenges remain to make supersonic travel a reality. This paper presents an integrated model that ties the aircraft preliminary design process, the economics of aircraft development and operations, and passenger preferences to estimate worldwide demand for subsonic overland, Mach cutoff, and low-boom aircraft design concepts. This comprehensive and iterative methodology takes into account the uncertainty of market demand for supersonic travel and the complexities of supersonic network operations. Based on the study of four conceptual supersonic designs, we found that low-boom aircraft could attract 28% more demand worldwide compared to an optimized Mach cutoff design because of the additional time-saving benefits on overland routes. The results of this work offer valuable guidance to aircraft designers by explicitly tying the technical parameters of a design and the potential market demand. In addition, the integrated model developed may streamline the initial design development cycle by quantifying fleet size as part of an integrated modeling process.

Analytic Solutions for Volume, Mass, Center of Gravity, and Inertia of Wing Segments and Rotors of Constant Density

In the preliminary design of aircraft lifting surfaces, accurate mass and inertia properties can be difficult to obtain. Typically, such methods as computer-aided design or statistical processes are used to determine these properties. These methods require significant time and effort to implement. The present paper presents an exact analytic method for calculating the volume, mass, center of gravity, and inertia properties of wing segments and rotors of constant density. The influence of taper, spanwise thickness distribution, airfoil geometry, and sweep are included. The utility of the method is presented, and the accuracy is evaluated with various test cases via percent difference with a corresponding computer-aided design model. These case studies demonstrate the present method to be accurate to within about 1% for typical wing geometries and within about 1.3% for typical propeller geometries.

Physical aspects of vortex-shock dynamics in delta wing configurations

Delta wing configurations with double- and triple-leading edges introduced within the North Atlantic Treaty Organization Applied Vehicle Technology -316 task group are examined to investigate the dynamics of vortices and shocks, with potential implications for the preliminary aircraft design. The numerical simulations are conducted for the configurations at Ma∞=0.85 and Re∞=12.53×106 using the Reynolds-averaged Navier–Stokes k−ω shear stress transport (SST) model across a range of incidence angles. The detailed analysis focuses on the case with α=20° using the scale-adaptive simulation based on the k−ω SST model. This study considers shock-vortex interaction and breakdown with buffeting to study the transient flow physics over the wing. Additionally, insights into vorticity strength and destruction are gained through the enstrophy transport equation. The findings reveal that the inboard vortex (IBV) development is impeded by counter-rotating secondary vortices from IBV and the midboard vortex. A key distinction is observed for the first time between the double-delta and triple-delta wings, in that the double-delta wing experiences shock-induced vortex breakdown, with the transient nature of this breakdown leading to an adjustment in the shock position, causing a shock buffet. In contrast, the breakdown in the triple-delta wing is linked to a stationary shock induced by the kink in the planform. This study highlights the crucial role of the orientation of the shock relative to the vortex axis in characterizing the aerodynamic performance of the planforms.

Neural-Network-Based Model for Trailing-Edge Flap Loads in Preliminary Aircraft Design

Accurately predicting the forces and moments acting on trailing-edge devices under different flight conditions is a critical aspect in the design of the kinematics and actuation for high-lift or variable-camber applications. However, accurate modeling without elaborated computational fluid dynamics (CFD) analyses in the subsonic and transonic regimes needs a sophisticated model. Thus, the objective of this paper is to create such a model that accurately predicts the forces and moments acting on flaps during different flight conditions while remaining applicable to the preliminary aircraft design. The target values in this model are the three-dimensional (3D) forces and moments on the flap, which were obtained through 3D CFD simulations. The chosen input values required for the model include two-dimensional airfoil data, and wing geometry data for three different aircraft types: short-, medium-, and long-range, including a high-aspect-ratio configuration. Among several potential approaches, a neural network was deemed to be the most promising for predicting the target values. The neural network was used as a regression tool to accelerate the model development process in the preliminary aircraft design. Consequently, multiple studies were conducted on how the setup of the neural network, including the number of neurons, activation functions, and initialization, influences the results. The results reveal that the developed neural network accurately predicts the flap forces and moments with a mean deviation of under 2% for the vertical force Fz and the lateral force Fx and under 4% for the moment My.

Stream Life Cycle Assessment Model for Aircraft Preliminary Design

The growing environmental public awareness and the consequential pressure on every industrial field has made environmental impact assessment increasingly important in the last few years. In this scope, the most established tool used in the specialized literature is the life cycle assessment. Applying this method to the life cycle of an aircraft requires it to be broken down into at least four phases: production, operation, maintenance and disposal. In the assessment, the evaluation of the environmental impact of fuel consumption can be performed linearly and has already been studied over many years, while calculating the impact of other life phases is more complicated, and it is still under study. This paper describes a simple and effective method developed to assess the environmental impact of an aircraft at a preliminary design stage and the implemented model that resulted from it. A detailed consideration of all life cycle phases is essential to serve as a reference for the ecological assessment of novel aircraft concepts. Thereby, the developed method is based on some parametric equations that take into account preliminary information, such as the mass breakdown, the technology used and some program considerations. The results obtained have been compared with those of the literature for verification and validation and have proved to be quite reliable. In fact, the comparison with known analyses, conducted on individual aircraft in a very precise manner, has showed that the proposed model is capable of giving results that fell within ±10% of the reference values. This is due to the broad generality of the model, which does not require a large number of specific data as a starting point to obtain reasonably reliable results for use during project development. In the near future, the use of this model can assist the design of aircraft architectures that comply with the European Green Deal of reducing net greenhouse gas emissions by at least 55% by 2030 and of having no net emissions of greenhouse gases by 2050.

Control moment coefficient methodology validation for eVTOL sizing

This paper presents a novel approach for the preliminary design of electric Vertical Takeoff and Landing (eVTOL) aircraft that utilizes the new Control Moment Coefficient (CMC) to size electric motors and to determine the rotor location and incidence angle. The CMC is determined for both thrust and arm length in eVTOL aircraft design, and is used to measure the moment produced by the rotors in the roll, pitch, and yaw axes. Analyzing its dimensionless value thus allows insights into an eVTOL aircraft's controllability. To test our methodology, two eVTOL aircraft were used in flight tests, one of which had up to 126% higher CMC values than the other. The results of the flight tests showed that a higher CMC value yielded many benefits, including an increased margin of safety between the rotors and the saturation level, reduced tracking error, and reduced control effort (or energy consumption).Furthermore, the 126% increase in the dimensionless CMC related to the pitch resulted in a 30% increase in the Pulse-Width Modulation - PWM margin of safety of the rotors at the saturated level while still maintaining a reasonable tracking error and a 97% decrease in the pitch control effort. Our research suggests that incorporating higher CMCs into the preliminary design of an eVTOL aircraft can significantly improve its safety and controllability. We hope that our findings will encourage further exploration of this promising approach in future.

Meshless Large-Eddy Simulation of Propeller–Wing Interactions with Reformulated Vortex Particle Method

The vortex particle method (VPM) has gained popularity in recent years due to a growing need to predict complex aerodynamic interactions during the preliminary design of electric multirotor aircraft. However, VPM is known to be numerically unstable when vortical structures break down close to the turbulent regime. In recent work, the VPM has been reformulated as a large-eddy simulation (LES) in a scheme that is both meshless and numerically stable without increasing its computational cost. In this study, we build upon this meshless LES scheme to create a solver for interactional aerodynamics. Propeller blades are introduced through an actuator line model following well-established practices for LES. A novel, vorticity-based actuator surface model (ASM) is developed for wings, which is suitable for propeller–wing interactions when a wake impinges on the surface of a wing. This ASM imposes the no-flow-through condition at the airfoil centerline by calculating the circulation that meets this condition and by immersing the associated vorticity in the LES following a pressure-like distribution. Extensive validation of propeller–wing interactions is presented by simulating a tailplane with tip-mounted propellers and a blown wing with propellers mounted midspan.

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.

On the use of filament-based free wake panel methods for preliminary design of propeller-wing configurations

With distributed propulsion and electric vertical take-off and landing aircraft on the rise, fast and accurate methods to simulate propeller slipstreams and their interaction with aircraft components are needed. In this work, we compare results obtained with a filament-based free wake panel method to experimental and previously validated numerical data. In particular, we study a propeller-wing configuration at zero angle of attack and the aerodynamics of the blade-resolved slipstream interaction with the wing. We use a prescribed wake on the wing and a free wake on the propeller, which greatly accelerate the computations. Results indicate that, while forces are overpredicted due to the inviscid nature of the panel method, the free wake is able to capture the slipstream deformation and shearing with remarkable success. We find that a filament-based free wake panel method can be a useful tool for propeller-wing interaction in preliminary aircraft design.

Methodology for Preliminary Flight Control Actuator Design

In this paper, a methodology for the preliminary design of decentralized civil-use flight control actuators is presented that is able to perform the calculations with a very small amount of input data. This enables the user to achieve good actuator key characteristics regarding the weight, power, and volume in the preliminary aircraft design stage without the need for an empirical database. Optionally adjustable setting parameters of the individual components allow the component design to be customized in detail. Both electromechanical and electrohydrostatic actuators are considered. The methodology presented here aims at sizing each individual component of the actuators in terms of volume, mass, and power consumption. The components are broken down into simple geometric shapes where possible. The methodology is evaluated on several actuators of different performance and weight classes from 3.2 kg up to 66.9 kg and provides more accurate results than other common approaches, which is suitable in the preliminary aircraft design stage.

Preliminary hybrid-electric aircraft design with advancements on the open-source tool SUAVE

The large variety of possible propulsion architectures for environmentally friendly and sustainable aircraft concepts create opportunities for the overall aircraft design. In particular, hybrid-electric aircraft have an increased number of additional components that must be implemented in the aircraft design tool. Primary and secondary energy sources can be used to achieve the necessary reduction in emissions. The preliminary aircraft design tool SUAVE offers the basic setup necessary to integrate the electrification of the propulsion system into the aircraft design. For this purpose, SUAVE has been extended to enable an iterative sizing loop for hybrid-electric aircraft. Hence, many propulsion architectures are easily adaptable. Additional components can be included and their interactions are considered. Furthermore, the fidelity levels of the models are modifiable. Moreover, the top-level aircraft requirements are transformed into a sizing chart to calculate the necessary power and wing loading. Thereby, a reasonable hybridization factor of installed power can be identified. This definition of hybridization implies a specific energy management strategy. As a result, hybridization and the energy management strategy are directly coupled and have to be considered together. Finally, in the postprocessing, the raw data is processed and can be evaluated with figures of merit. This procedure converts the results into a descriptive value and facilitates the comparison to other configurations. The figures of merit include evaluations for emissions, operational aspects and development costs and effort.

Parametric model of aircraft external geometry as a component of FAST-OAD aircraft preliminary design system.

Rapid evolutions in fields such as aerospace propulsion, materials engineering, geometry modelling tools or optimization tools provide the opportunity to search for novel configurations of future aircraft in an expanding domain. In order to effectively explore the space of variables of constantly expanding size, one needs tools that give the ability to automate this process. Such a tool is the Future Aircraft Sizing Tool for Overall Aircraft Design or FAST-OAD. FAST-OAD allows for rapid estimation of aircraft size, based on set, so-called Top Level Aircraft Requirements (TLARs) by using multi-disciplinary and multi-fidelity analysis and optimization.An essential component of such a system is a module capable of modelling the external geometry of the calculated case. This article describes an attempt to use a powerful commercial Computer Aided Design software, Siemens NX, to model the aircraft external geometry based on the results of the analysis performed by FAST-OAD. This approach, compared to previous works, brings some limitations but, on the other hand, gives the possibility to base the geometry on a mathematically consistent model and helps to better understand the set of parameters necessary to describe the geometry correctly.The main objective of this research is to provide a tool that allows the automatic generation of aircraft exterior geometry based on the output parameters received from the FAST-OAD package.An additional goal of this activity is to determine the parameter set necessary to properly define the external geometry, but at the same time not containing redundant, dependent geometric parameters. The minimum number of independent parameters necessary to completely describe the external geometry is sought.

Influence of Electric Wing Tip Propulsion on the Sizing of the Vertical Stabilizer and Rudder in Preliminary Aircraft Design

During preliminary aircraft design, the vertical tail sizing is conventionally conducted by the use of volume coefficients. These represent a statistical approach using existing configurations’ correlating parameters, such as wing span and lever arm, to size the empennage. For a more detailed analysis with regard to control performance, the vertical tail size strongly depends on the critical loss of thrust assessment. This consideration increases in complexity for the design of the aircraft using wing tip propulsion systems. Within this study, a volume coefficient-based vertical tail plane sizing is compared to handbook methods and the possibility to reduce the necessary vertical stabilizer size is assessed with regard to the position of the engine integration and their interconnection. Two configurations, with different engine positions, of a hybrid-electric 19-seater aircraft, derived from the specifications of a Beechcraft 1900D, are compared. For both configurations two wiring options are assessed with regard to their impact on aircraft level for a partial loss of thrust. The preliminary aircraft design tool MICADO is used to size the four aircraft and propulsion system configurations using fin volume coefficients. These results are subsequently amended by handbook methods to resize the vertical stabilizer and update the configurations. The results in terms of, e.g., operating empty mass and mission fuel consumption, are compared to the original configurations without the optimized vertical stabilizer. The findings support the initial idea that the connection of the electric engines on the wing tips to their respective power source has a significant effect on the resulting torque around the yaw axis and the behaviour of the aircraft in case of a power train failure, as well as on the empty mass and trip fuel. For only one out of the four different aircraft designs and wiring configurations investigated it was possible to decrease the fin size, resulting in a 53.7% smaller vertical tail and a reduction in trip fuel of 4.9%, compared to the MICADO design results for the original fin volume coefficient.

A discrete adjoint full potential formulation for fast aerostructural optimization in preliminary aircraft design

Preliminary aircraft design is often carried out with the help of multidisciplinary optimization processes. Because there is a strong coupling between the flow and the structure of the aircraft, and because new composite materials have a unique capability of being designed for anisotropic directional stiffness and strength, these processes must model the aeroelastic behavior of the aircraft in order to be effective. Since many design variables are involved during the preliminary design stage, the optimization problems, formulated using high or medium fidelity models, can be solved using the adjoint method. Moreover, the level of fidelity of the fluid model and the associated simulation technique must also be selected with care, as they tend to be the main contributors to the computational cost. The novel contribution of the present work is twofold. Firstly, the discretized gradients of the full potential flow equation, which is a medium-fidelity model, are derived analytically so that they can be used in adjoint optimization problems. Moreover, the full potential flow solution and the computation of the gradients are implemented in an open-source and readily available finite element code. Secondly, aerodynamic shape and aerostructural optimization calculations are carried out on example wings to demonstrate the effectiveness and the computational efficiency of the proposed method. Overall, the results show that the newly implemented discrete adjoint nonlinear potential flow formulation is able to quickly optimize both the shape and the structural parameters of a typical wing. More specifically, the twist distribution along the wingspan is adapted to reduce the induced drag, the airfoils become more supercritical so that the shock strength and the associated wave drag are reduced, and the thickness of the structural elements is tailored to the loads to reduce the internal stresses. The present methodology is therefore able to deliver results at a low computational cost which are sufficiently accurate for the early design stages. Furthermore, the results obtained using the proposed methodology could be used as a starting point for the optimization calculations performed in later design stages.

Prediction of Installed Jet Noise from Wave-Packet Models Tuned with Freejet Data

In this work, a kinematic wave-packet model is used to predict installed-jet noise. Large-eddy simulation results of freejets, for Mach numbers 0.4 and 0.9, are used to obtain parameters of wave packets representing large-scale turbulent structures, which were used to provide a model source for the Lighthill analogy used to predict far-field noise spectra. The source amplitude in the model is calibrated using noise measurements for a freejet, and such a wave-packet source is used to predict noise of the same jet in an installed configuration using a tailored Green’s function. Results from the prediction model are compared to installed-jet experimental data for four different observer positions and a large range of frequencies. Overall, the model predicts both directivities and amplitudes similar to the experimental data, with a hump in the generated noise for lower Strouhal numbers and a clear peak near a Strouhal number of 0.2. This low-order model is fast and flexible, and it is expected to be helpful in preliminary aircraft design.

Methodology for inerting system analysis in the long-range aircraft preliminary design phase

Methodology for inerting system analysis in the long-range aircraft preliminary design phase