Aircraft Concept Research Articles

Development of a medium/long-haul reference aircraft

For the evaluation of future aircraft architectures or novel technologies, a well-understood and sound reference to compare to is essential to prove the impact of intended changes. A reference aircraft serves as a basis for calibration of the aircraft and is used as a starting point for subsequent investigations, sensitivity studies or optimizations. The prime need for such a reference arose from the European Clean-Sky-2 project HLFC-Win (Hybrid Laminar Flow Control—Wing), where no industrial reference data for a conventional aircraft were available. The project investigates a HLFC system integrated into the wing aiming to reduce the wing drag and consequently reducing the fuel consumption. Therefore, this paper presents the D300-XRF1 designed by DLR, which is based on the AIRBUS research aircraft XRF1 (eXternal Research Forum). It is intended to be used as a reference aircraft database representing aircraft for medium- and long-range missions. A variable fidelity multidisciplinary design analysis and optimization (MDAO) aircraft design environment is set up to provide a consistent estimate of the geometry, mass breakdown, propulsion system, aerodynamics and aircraft performance. A conceptual aircraft design tool is utilized in the aircraft design environment to initialize the process, and higher-fidelity modules are used to enhance the results. The presented results describe the overall aircraft characteristics of the D300-XRF1 based on 2010 entry into service technology level and is offered to be used as a reference within the aviation research community, hopefully reducing similar design efforts in other research projects.The aircraft operates at a design cruise Mach number of 0.83, has a design range of 5500 NM and transports a payload of 31.5 t (300 PAX at 105 kg/PAX). The D300-XRF1 serves as a consistent reference aircraft database and provides a holistic overview of the aircraft’s performance that has been presented to and approved by AIRBUS. Therefore, this reference aircraft design can be used for future studies and to assess new technologies on a sophisticated level.

A Simulation Framework for Aircraft Take-Off Considering Ground Effect Aerodynamics in Conceptual Design

The development of novel aircraft concepts and propulsion technologies requires up-to-date physics-based methods and tools for conceptual aircraft design. In this context, a simulation model for the take-off manoeuvre is proposed in this article, to be employed in the conceptual design phase for aircraft whether of traditional or innovative configuration. The model is capable of evaluating the longitudinal dynamics, both translational and rotational, of the aircraft considered as a rigid body, and influenced by the aerodynamic effects introduced by the presence of the ground. The ground effect, indeed, induces variations in the aerodynamic forces depending on the distance and the attitude of the lifting surfaces from the ground, which may significantly influence the aeromechanical characteristics of the aircraft during the evolution of the take-off manoeuvre. The simulation model is based on the numerical solution of the equations of the dynamics of the rigid aircraft in the longitudinal plane and integrates a vortex lattice aerodynamic solver to evaluate the aerodynamic and aeromechanical characteristics of the aircraft considering the ground effect in each time-step. The proposed approach is configuration independent, as it can model the geometry, evaluate the aerodynamics, and simulate the dynamics of aircraft with any lifting architecture; furthermore, the simulation model is fast and flexible, making it effective for the conceptual phase of aircraft design. The paper proposes the description of the take-off manoeuvre of two aircraft with different airframes: one with a conventional tube-and-wing architecture and one with a box-wing lifting system. The results proposed highlight the potential of the simulation model to detect aeromechanic and dynamic differences during the development of the manoeuvre for different aircraft configurations, and to assess the significance of considering ground effect aerodynamics.

Aerodynamic Investigation of Blended Wing Body Configuration


 A blended wing body (BWB) configuration is an unconventional aircraft design in which the wing and fuselage are blended to form an aircraft. This design concept has inherent higher aerodynamic efficiency, environmental benefits and capacities. These advantages make the BWB configuration a feasible concept for commercial transport aircraft. In the present work, a 3-D BWB model is designed in SolidWorks and fabricated using a 3D printer. The numerical and experimental analyses are carried out with this BWB geometry. Aerodynamic characteristics and flow features obtained from the open-source CFD software OpenFOAM have been studied, analyzed, and compared with the wind tunnel results. Experimental and computational data compare well and the present BWB can operate at a high angle of attack. The coefficient of lift (CL) increases with AoA up to 45º. The CL starts decreasing beyond this AoA, and the present BWB geometry stalls at around AoA = 45º. The coefficient of drag (CD) increases with the increase in AoA due to the spreading of the separated region over the geometry. Lift/Drag (L/D) variation with AoA is also studied to find the optimum flight configuration of the present BWB geometry. Sectional pressure distribution at different spanwise locations, velocity contours, pathlines, surface limiting streamlines and tuft flow visualization are also presented to investigate the flow. The studies investigate the aerodynamics, flow field and optimal flight configuration for cruising a BWB geometry.

A tilt-wing VTOL UAV configuration: Flight dynamics modelling and transition control simulation

AbstractThis paper aims to present a vertical take-off and landing unmanned aerial vehicle (VTOL UAV) configuration and numerically simulate its flight transition from hover to cruise and from cruise to hover. It can tilt the canard and wing along with two attached propellers. Additionally, two fixed front propellers are pointing upwards. Multi-body equations of motion are derived for this concept of aircraft, which are used to compute the flight transition trajectory from hover to cruise configuration. Furthermore, a transition control algorithm based on gain scheduling is described, which stabilises the aircraft while it accelerates from hover to cruise, gradually tilting the wing along with its propellers, sequentially switching between equilibrium states, as the stability cost functions thresholds are reached. The transition control algorithm of the conceptual aircraft model is numerically simulated.

Review of target objects for the group-operated unmanned aerial vehicles application

Currently, the big number of companies, and specifically in Russia, is engaged in designing unmanned aerial vehicles (UAV). It is caused by the tendency to reduce human labor in various fields of activity, in particular, associated with a great risk to life and health. All these factors lead to the steady growth of different-type missions for the UAV application. Urgency of the research generates considerable interest towards the issues of the UAV application. As part of developing a method for determining UAV design parameters, the author analyzed the results in this domain, considering the operation in a group of aircraft. In order to propose a conceptual aircraft solution for designing a model, the analysis of target objects for the UAV group application was conducted. The article considers various target objects for the application of UAV groups. The result of the analysis for target objects can form the basis for the conclusion involved with the imposed requirements for the design of this UAV class. Compared to other specialists’ articles, this paper is distinguished by synthesizing the achievements in the sphere of group-operated UAV design, as well as by establishing the basic trends for their application while conducting warfare.

A Design Model for Electric Environmental Control System in Aircraft Conceptual and Preliminary Design

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

A Performance-Based Parametric Design Exploration Tool for Student Aircraft Design Projects

In this study, a performance-based calculation tool was developed to ease the process of exploring different conceptual aircraft conditions. The study aims to minimize human errors when calculating aircraft performance parameters manually. The parametric design tool, which will aid in designing unmanned aircraft, was built using Microsoft Excel. It was developed for the purpose of saving time and excessive effort that hand calculations require. Based on the criteria mentioned in the SAE aero design competition, five conceptual designs for the unmanned aircraft were drafted. One of the main criteria of this competition is creating an aircraft with a high payload fraction and a low empty weight. Using this Excel spreadsheet, we computed the values of thrust force, lift coefficient, drag coefficient, dynamic pressure, and drag force acting on the unmanned aircraft, allowing us to analyze and compare the performances of the different conceptual designs. To make sure that the Excel spreadsheet was reliable, it was tested and verified by comparing it with hand calculations.

An energy systems model of large commercial liquid hydrogen aircraft in a low-carbon future

Liquid hydrogen (LH2) aircraft have the potential to achieve carbon neutrality. However, if the hydrogen is produced using electricity grids that utilise fossil fuel, they have a non-zero carbon dioxide (CO2) emission associated with their well-to-wing pathway. To assess the potential of LH2 in aviation decarbonisation, an energy systems comparison of large commercial LH2, liquified natural gas (LNG), conventional Jet-A and LH2 dual-fuel aircraft is presented. The performance of each aircraft is compared towards 2050, over which three system changes occur: (1) LH2 aircraft technology develops; (2) both world average and region-specific grid electricity, which is used to produce the hydrogen, decarbonises; and (3) the International Air Transportation Association (IATA) emissions targets, which are used to restrict the passenger-range performance of each aircraft, tighten. In 2050, the emissions of all aircraft are thus constrained to 0.063 kg-CO2/p-km, relative to 0.110 kg-CO2/p-km for the unconstrained Jet A fuelled Boeing 787-8. It is estimated that, in this year, an LH2 aircraft powered by fuel cells and sourcing world average electricity can travel 6000 km, 20% further than the conventional Jet A aircraft that is also constrained to meet the IATA targets, but not as far as the LNG aircraft. At its maximum range, the LH2 aircraft carries 84% of the Jet A passenger demand. Analysis using region-specific hydrogen indicates that LH2 aircraft can travel further than LNG aircraft in North America only, accounting for 17% of the global demand. 1.59 times the current aviation energy consumption is required if all conventional aircraft are replaced with LH2 designs. Under stricter emissions constraints than those outlined by the IATA, LH2 outperforms LNG in Europe and the Americas, accounting for 41% of the global demand. Also in these regions, the range, energy consumption and passenger capacity of LH2 aircraft can be improved upon by combining the advantages of LH2 with LNG in dual-fuel aircraft concepts. The use of LH2 is therefore advantageous within several prominent niches of a future, decarbonising aviation system.

Design for bird strike crashworthiness using a building block approach applied to the Flying-V aircraft

The Flying-V aircraft concept promises better fuel-burn performance over conventional tube-and-wing configurations, integrating the passenger cabin and cargo volume into the lifting surface. However, the wing-fuselage and cockpit windows of the Flying-V are exposed to the flight direction, posing a new challenge to the design and certification of structures in terms of bird strikes. This study is a first step towards understanding the dynamic load path and contribution of each structural element on the bird strike resistance of the Flying-V leading-edge structures. The objective is to propose a building block approach to design the Flying-V's leading edge bird strike crashworthiness that complies with the EASA's certification CS25.631 using a 4lb bird impacted at a sea level cruising speed of 70 m/s. An additional requirement by the authors is to keep the structure within the elastic deformation during the impact of a 4lb bird to avoid the need for repairs in the Flying-V fuselage. Plasticity generated in the structure is regarded as damage and is used as a comparative parameter. At the highest building block level, a sensitivity analysis is performed to identify the effect of the thickness of each structural element on the plasticity and weight of the leading-edge structures. The trends are used to modify the baseline design and achieve a reduction of 80% of the plastic energy. The critical case of a 133 m/s impact of a 4lb bird at the cruise altitude of 37000 ft is also evaluated, and the results show penetration.

Structured Expert Judgment Elicitation in Conceptual Aircraft Design

Disruptive technologies and novel aircraft generations represent a potential approach to address the ambitious emission reduction goals in aviation. However, the introduction of innovative concepts is a time-consuming process, which might not necessarily yield an optimal design for a given flight mission and within the defined time frame. In order to address the need for a structured and more exhaustive search for novel concept generations, the Advanced Morphological Approach (AMA) and its further enhancement was introduced earlier. It implies the decomposition of design problems into functional attributes and appropriate technological alternatives. Subsequently, these are evaluated and combined into solutions, which are then projected onto a solution space. The current paper focuses on the technology evaluation step by deriving and integrating structured expert judgment elicitation (SEJE) techniques into conceptual aircraft design with the AMA. For this purpose, the first aim of the work is to justify the developed method by giving an overview and discussing the most prominent SEJE methods and their applications in aerospace. Then, the derived SEJE concept is described and applied in the form of an expert workshop on the use case of wing morphing architecture. As a result, a solution space of wing morphing architecture configurations is generated and analyzed. The workshop conduction and the expert feedback serve as valuable findings for both the further AMA enhancement and similar research.

Improvement of Take-Off Performance for an Electric Commuter Aircraft Due to Distributed Electric Propulsion

The need for environmentally responsible solutions in aircraft technology is now considered the priority for global challenges related to the limited supply of traditional fuel sources and the potential global hazards associated with emissions produced by traditional aircraft propulsion systems. Several projects, including research into highly advanced subsonic aircraft concepts to drastically reduce energy or fuel usage, community noise, and emissions associated with aviation, are currently ongoing. One of the proposed propulsion concepts that address European environmental goals is distributed electric propulsion. This paper deals with the detailed aerodynamic analyses of a full-electric commuter aircraft with fuel cells, which expects two primary electric motors at the wing tip and eight other electric motors distributed along the wingspan as secondary power sources. The main objective was the numerical estimation of propulsive effects in terms of lift capabilities at take-off conditions to quantify the possible reduction of take-off field length. However, the aircraft was designed from scratch, and therefore a great effort was spent to design both propellers (for the tip and distributed electric motors) and the wing flap. In this respect, several numerical tests were performed to obtain one of the best possible flap positions. This research work estimated a reduction of about 14% of the take-off field length due to only the propulsive effects. A greater reduction of up to 27%, if compared to a reference conventional commuter aircraft, could be achieved thanks to a combined effect of distributed propulsion and a refined design of the Fowler flap. On the contrary, a significant increment of pitching moment was found due to distributed propulsion that may have a non-negligible impact on the aircraft stability, control, and trim drag.

Assessment of Potential Commercial Success of Business Jets with Natural Laminar Flow

The present research investigates the potential application of the natural laminar flow wing technology to business jet market segments from light to long-range jets. A database of existing business jets was generated to determine the range extension as a potential driver of customer interest. A conceptual design of several configurations for each market segment was performed to investigate potential improvements in the aircraft flight range, operating costs, and price changes using the technology. An initial sizing module and a low-fidelity multidisciplinary design optimization were used to size all aircraft. Results demonstrated a 13–30% increase in the flight range depending on the aircraft concept. Long-range jets with natural laminar flow could not achieve significant range extension due to a combination of the design airspeed and maintenance costs. A rapid increase in acquisition prices for all aircraft suggested that super-midsize and large jets that combine relatively low empty weight and high range extension could be more favorable options compared to other segments, but a significant reduction of acquisition costs and an increase in operating flight hours are required to make the technology implementation successful.

Preliminary Design and Simulation of a Thermal Management System with Integrated Secondary Power Generation Capability for a Mach 8 Aircraft Concept Exploiting Liquid Hydrogen

This paper introduces the concept of a thermal management system (TMS) with integrated on-board power generation capabilities for a Mach 8 hypersonic aircraft powered by liquid hydrogen (LH2). This work, developed within the EU-funded STRATOFLY Project, aims to demonstrate an opportunity for facing the challenges of hypersonic flight for civil applications, mainly dealing with thermal and environmental control, as well as propellant distribution and on-board power generation, adopting a highly integrated plant characterized by a multi-functional architecture. The TMS concept described in this paper makes benefit of the connection between the propellant storage and distribution subsystems of the aircraft to exploit hydrogen vapors and liquid flow as the means to drive a thermodynamic cycle able, on one hand, to ensure engine feed and thermal control of the cabin environment, while providing, on the other hand, the necessary power for other on-board systems and utilities, especially during the operation of high-speed propulsion plants, which cannot host traditional generators. The system layout, inspired by concepts studied within precursor EU-funded projects, is detailed and modified in order to suggest an operable solution that can be installed on-board the reference aircraft, with focus on those interfaces impacting its performance requirements and integration features as part of the overall systems architecture of the plane. Analysis and modeling of the system is performed, and the main results in terms of performance along the reference mission profile are discussed.

Fluctuations by atmospheric turbulence in aircraft flyover auralisation

Long-term exposure to aircraft noise causes significant health issues among residents near airports. Therefore, noise impact assessments and noise control at the source are important aspects of the design of new aircraft. The design process of low-noise aircraft can be supported with auralisation of virtual flyovers. In order to render plausible aircraft auralisations, multiple propagation phenomena have to be considered, such as geometrical spreading, air absorption, Doppler effect, and reflections from the ground. Additionally, measurements of aircraft flyover show clear patterns of amplitude and phase fluctuations by atmospheric turbulence. Aircraft flyover auralisations should incorporate these clearly audible features of sound propagation in real-world conditions to be perceived as plausible. We present new approaches of time-variant filtering techniques to account for phase and amplitude fluctuations as a function of atmospherical conditions characterised by the von Kármán turbulence spectrum of wind and temperature fluctuations. Compared to earlier approaches, the proposed model is closer to the physical mechanisms. The application of the filters leads to a reduction of unnatural flanging and to a higher naturalness of level fluctuations. As the proposed method has shown to increase plausibility of aircraft flyover auralisation, its application in perception-based evaluation of future aircraft concepts is forseen.

Evaluation of flyover auralisations of today's and future long-range aircraft concepts

The European research project ARTEM (Aircraft noise Reduction Technologies and related Environmental iMpact) develops innovative aircraft noise reduction technologies such as advanced engine fan lining, metamaterials and low-noise high-lift systems applied to a vehicle with enhanced shielding of the engine noise, namely, a blended wing body. Using aircraft flyover auralisation in laboratory listening experiments, such future technologies can be evaluated with respect to human sound perception. To assess the reliability of such perception-based evaluations, the simulation chain should be validated with existing aircraft flyovers. This contribution presents a systematic and rigorous hierarchical validation of auralisations of current jet aircraft using field recordings. Uncertainty in the source modelling is considered by using two different prediction tools for partial sound sources. In addition to comparing computed noise indicators, a psychoacoustic validation is done in laboratory listening experiments with a 3D loudspeaker array. The validation comprises three levels: (i) direct comparison of auralisations with recordings to study the identifiability of auralisations, (ii) ranking of auralisations and recordings regarding plausibility, and (iii) subjective annoyance ratings to test whether auralisations and recordings differ with respect to noise effects. Further, first results on the comparison of a future concept with a current aircraft are presented.

Experimental Assessment of Aero-Propulsive Effects on a Large Turboprop Aircraft with Rear-Engine Installation

This paper deals with the estimation of propulsive effects for a three-lifting surface turboprop aircraft concept, with rear engine installation at the horizontal tail tips, conceived to carry up to 130 passengers. This work is focused on how the propulsive system affects the horizontal tailplane aerodynamics and, consequently, the aircraft’s static stability characteristics using wind tunnel tests. Both direct and indirect propulsive effects have been estimated. The former produces moments whose values depend on the distance from the aircraft’s centre of gravity to the thrust lines and propeller disks. The latter entails a change in the angle of attack and an increment of dynamic pressure on the tailplane. Several tests were also performed on the body-empennage configuration to investigate the propulsive effects on the aircraft’s static stability without the appearance of any aerodynamic interference phenomena, especially from the canard. The output of the experimental campaign reveals a beneficial effect of the propulsive effects on the aircraft’s longitudinal stability, with an increase in the stability margin of about 2.5% and a reduction in the directional stability derivative of about 4%, attributed to the different induced drag contributions of the two horizontal tail semi-planes. Moreover, the rolling moment coefficient experiences a greater variation due to the propulsion depending on the propeller rotation direction. The outcomes of this paper allow the enhancement of the technical readiness level for the considered aircraft, giving clear indications about the feasibility of the aircraft configuration.

Определение критической частоты вращения ротора высокого давления для двухконтурного турбореактивного двигателя с интегрированным стартером-генератором

Today, the development of aircraft within the electric aircraft concept framework is relevant. In this case, the key technology is the integrated starter-generator, which makes it possible to refuse the mechanical, pneumatic, and hydraulic power take-off from the aircraft engine. In the well-known scientific articles, integrated starter-generators are designed without considering the features due to their location inside the aircraft engines. This study starts a set of projects devoted to the design of integrated starter-generator for a bypass turbojet engine, considering the features of this technical decision. The purpose of this study is to develop approaches to determine critical rotation frequency of high-pressure rotor considering starter-generator integration. Models to determine the critical rotation frequency of a high-pressure rotor with an integrated starter-generator rotor are developed based on non-traditional models of rotor dynamics by mathematical modeling. An area for the integrated starter-generator location has been determined and a constructive method for its implementation is proposed for a bypass turbojet engine. In this study, mathematical expressions are obtained to determine high-pressure rotor critical speed in case of a detailed study of dynamics and in case of an integrated starter-generator preliminary calculations. Based on the value of the critical speed of the high-pressure rotor calculated from the obtained expressions, firstly, it is possible to check the correctness of the choice of the air gap of the integrated starter-generator and strength calculations of the integrated starter-generator in the most difficult operating mode. Secondly, it is possible to determine safe operation area of the bypass turbojet engine with an integrated starter-generator. The results of the study enhance the theoretical foundations of the design of the integrated starter-generators and make it possible to avoid errors in their design and further operation.

SystemXF: a novel approach for holistic system modeling in aircraft conceptual design

In aircraft conceptual design, systems are typically represented by historical regressions without considering system architecture. For novel technologies, this is no longer possible. Alternative approaches are either hard coded solutions for special system architectures or lack the analysis capabilities necessary for conceptual aircraft design. The novel SystemXF approach intends to close this gap by providing a framework for system modeling that is not technology specific and allows for the integration of different analysis disciplines within a single model. A central XML file is used for system specification which can be edited using a SysML interface. A mass estimation and a safety analysis tool are available. An example system which is inspired by a real pitch control system is used to illustrate the method and to show the applicability. This modeling example shows that the proposed model is capable to represent and analyze realistic systems in conceptual aircraft design. The analysis tools can be used for system studies and to provide inputs for a conceptual aircraft design process.

Geometrically Nonlinear Coupled Adjoint Aerostructural Optimization of Natural-Laminar-Flow Strut-Braced Wing

Novel aircraft concepts employing ultrahigh-aspect-ratio wings, such as the strut-braced wing (SBW) configuration, are promising ways to achieve the next-generation sustainable and fuel-efficient aviation goals. However, as the wing aspect ratio increases, the wing increasingly exhibits more flexibility, higher deformation, and geometrically nonlinear behavior that cannot be accurately simulated by conventional sizing methods and typical linear structural analysis models. This paper establishes a framework for SBW aircraft conceptual design, conceptual optimization, and aerostructural optimization. The presented aerostructural optimization method has medium-fidelity and physics-based features. A geometrically nonlinear structural analysis solver and a quasi-three-dimensional aerodynamic solver are coupled for the aerostructural optimization of composite natural-laminar-flow SBW aircraft. A medium-range (MR)-SBW aircraft is initially designed and optimized in the conceptual design stage. A gradient-based aerostructural optimization is performed using the proposed tool for minimizing the fuel mass of the initially sized and optimized MR-SBW aircraft. The optimization results in a more than 10% reduction in fuel mass, a more than 8% reduction in aircraft maximum takeoff mass, and a more than 30% reduction in wing and strut structural weight by optimizing the wing box structure, the wing planform, and the airfoil shape while satisfying the constraints on structural failure, wing loading, and aileron effectiveness.

Component-Level Modeling of More Electric Auxiliary Power Units for Cooperative Control

Today, the more electric aircraft (MEA) concept is gaining tremendous popularity. As a key component of the MEA, a more electric auxiliary power unit (MEAPU) integrated model with high accuracy and real-time performance is essential when conducting cooperative control and hardware-in-the-loop (HIL) test research. This paper proposes a novel MEAPU integrated model consisting of a MEAPU component-level-model (CLM) and a starter-generator (SG) model. Firstly, a MEAPU CLM was built and a continuous scaling method for the component characteristic map in the CLM is proposed to improve the model’s accuracy. Then, a double winding induction starter-generator (DWISG) model based on the electromagnetic theory, which is quite time consuming, was simplified using the pulse width modulation (PWM) rectifier linearization method. Finally, considering the coupling relationship between the MEAPU CLM and DWISG, an accurate real-time MEAPU integrated model was built and its simulation results were analyzed. Compared with the test results, the error of the proposed model was less than 0.5%; meanwhile its single-step simulation time was less than 20 ms, which can meet the demands of cooperative control and HIL test research. Furthermore, the continuous scaling method and PWM rectifier linearization method were found to be effective for modeling other MEAPUs and more electric engines (MEE).