Aircraft Synthesis Research Articles

Benefits and Challenges of Electric Propulsion for Commuter Aircraft

This paper presents the Library for Unified Conceptual Aircraft Synthesis (LUCAS) sizing and optimization framework, which is largely based on Stanford’s Stanford University Aerospace Vehicle Environment (SUAVE). LUCAS is applied to a commuter mission carrying 19 passengers. Results show that energy usage benefits depend on mission range and electrical component technology and that reserve requirements play a significant role in the feasibility of electrified aircraft. It is found that storing more energy in batteries rather than fuel reduces the mission energy requirement at the expense of aircraft weight growth. All-electric aircraft could provide onboard energy savings of upward of 30% for a 180 km mission, but this would only become feasible if battery specific energy reached the high value of 900 W⋅h/kg. Range could be increased to 550 km if part of the reserve mission requirements is relaxed. Under more realistic technology assumptions, a hybrid design storing energy in fuel and batteries could fly commuter missions with energy savings of 6%. A beneficial use case of hybrid configurations is a mode of operation where the aircraft uses only batteries during the main mission and fuel to store all the reserve energy. If reserves are not required for a particular flight, energy savings are still maximized, while the availability of fuel ensures certification requirements are met with fewer penalties.

Effect of Progressive Integration of On-Board Systems Design Discipline in an MDA Framework for Aircraft Design with Different Level of Systems Electrification

The on-board design discipline is sometimes ignored during the first aircraft design iterations. It might be understandable when a single on-board system architecture is considered, especially when a conventional architecture is selected. However, seeing the trend towards systems electrification, multiple architectures can be defined and each one should be evaluated during the first tradeoff studies. In this way, the systems design discipline should be integrated from the first design iterations. This paper deals with a progressive integration of the discipline to examine the partial or total effect of the systems design inside an MDA workflow. The study is carried out from a systems design perspective, analyzing the effect of electrification on aircraft design, with different MDA workflow arrangements. Starting from a non-iterative systems design, other disciplines such as aircraft performance, engine design, and aircraft synthesis are gradually added, increasing the sensibility of the aircraft design to the different systems architectures. The results show an error of 40% in on-board systems assessment when the discipline is not fully integrated. Finally, using the workflow which allows for greater integration, interesting differences can be noted when comparing systems with different levels of electrification. A possible mass saving of 2.6% of aircraft MTOM can be reached by properly selecting the systems technologies used.

Benchmarking New CEASIOM with CPACS adoption for aerodynamic analysis and flight simulation

PurposeThe purpose of this paper is to present the status of the on-going development of the new computerized environment for aircraft synthesis and integrated optimization methods (CEASIOM) and to compare results of different aerodynamic tools. The concurrent design of aircraft is an extremely interdisciplinary activity incorporating simultaneous consideration of complex, tightly coupled systems, functions and requirements. The design task is to achieve an optimal integration of all components into an efficient, robust and reliable aircraft with high performance that can be manufactured with low technical and financial risks, and has an affordable life-cycle cost.Design/methodology/approachCEASIOM (www.ceasiom.com) is a framework that integrates discipline-specific tools like computer-aided design, mesh generation, computational fluid dynamics (CFD), stability and control analysis and structural analysis, all for the purpose of aircraft conceptual design.FindingsA new CEASIOM version is under development within EU Project AGILE (www.agile-project.eu), by adopting the CPACS XML data-format for representation of all design data pertaining to the aircraft under development.Research limitations/implicationsResults obtained from different methods have been compared and analyzed. Some differences have been observed; however, they are mainly due to the different physical modelizations that are used by each of these methods.Originality/valueThis paper summarizes the current status of the development of the new CEASIOM software, in particular for the following modules: CPACS file visualizer and editor CPACSupdater (Matlab) Automatic unstructured (Euler) & hybrid (RANS) mesh generation by sumo Multi-fidelity CFD solvers: Digital Datcom (Empirical), Tornado (VLM), Edge-Euler & SU2-Euler, Edge-RANS & SU2-RANS Data fusion tool: aerodynamic coefficients fusion from variable fidelity CFD tools above to compile complete aero-table for flight analysis and simulation.

Conceptual design, performance and stability analysis of a 200 passengers Blended Wing Body aircraft

The Blended Wing Body (BWB) is a type of innovative aircraft, based on the flying wing concept. For this aircraft, the literature has reported performance improvements compared to conventional aircraft: economy of fuel, reduction of the weight of the structure, increased payload capacity and less impact on the environment. However, most BWB studies have focused on large aircraft and it is not sure whether the gains are the same for smaller aircraft. The main objective of this study is to perform the conceptual design of a 200 passengers BWB and compare its performance against an equivalent conventional A320 aircraft in terms of payload and range. Moreover, an emphasis will be placed on obtaining a stable aircraft, with the analysis of static and dynamic stability over its flight envelope. This kind of aircraft has a lack of stability due to the absence of vertical tail. Most studies of stability were already realized on reduced size models of BWB, but there is no study on a 200 passengers BWB. The design of the BWB was realized with the platform called Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods (CEASIOM). The airplane, the engines and the control surfaces were obtained in the geometrical module AcBuilder. This design platform, suitable for conventional aircraft design, has been modified and additional tools have been integrated in order to achieve the aerodynamic, performance and stability analysis of the BWB aircraft. The aerodynamic coefficients are calculated from Tornado program. The BWB flight envelope was created based on aeronautical data of A320 aircraft. From this flight envelope, we have got back several thousand possible points of flight. The static and dynamic stability was studied using the longitudinal and lateral matrices of stability and the Flying Qualities Requirements for every flight point.

Aerodynamic wing shape optimization based on the computational design framework CEASIOM

PurposeA collaborative design environment is needed for multidisciplinary design optimization (MDO) process, based on all the modules those for different design/analysis disciplines, and a systematic coupling should be made to carry out aerodynamic shape optimization (ASO), which is an important part of MDO.Design/methodology/approachComputerized environment for aircraft synthesis and integrated optimization methods (CEASIOM)-ASO is developed based on loosely coupling all the existing modules of CEASIOM by MATLAB scripts. The optimization problem is broken down into small sub-problems, which is called “sequential design approach”, allowing the engineer in the loop.FindingsCEASIOM-ASO shows excellent design abilities on the test case of designing a blended wing body flying in transonic speed, with around 45 per cent drag reduction and all the constraints fulfilled.Practical implicationsAuthors built a complete and systematic technique for aerodynamic wing shape optimization based on the existing computational design framework CEASIOM, from geometry parametrization, meshing to optimization.Originality/valueCEASIOM-ASO provides an optimization technique with loosely coupled modules in CEASIOM design framework, allowing engineer in the loop to follow the “sequential approach” of the design, which is less “myopic” than sticking to gradient-based optimization for the whole process. Meanwhile, it is easily to be parallelized.

신속한 항공기 개발을 위한 통합 개념설계 프로세스에 대한 연구

Integrated aircraft synthesis process for rapid analysis and design is described in this paper. Data flow between different analysis fields is described in details. All the data are divided into several groups according to importance and source of the data. Analysis of design requirements and certification regulations is carried out to determine baseline configuration of an aircraft. Overall design process can be divided into initial sizing, conceptual and preliminary design phases. Basic data for conceptual design are obtained from initial sizing, CAD and geometry analysis. Basic data are required input for weight, aerodynamics and propulsion analyses. Results of this analysis are used for stability and control, performance, mission, and load analysis. Feasibility of design is verified based on analysis results of each discipline. Design optimization that involves integrated process for aircraft analysis is performed to determine optimum configuration of an aircraft on a conceptual design stage. The process presented in this paper was verified to be used for light aircraft design.

Reduced-Order <i>H<sub>∞</sub> </i>Controller Design Using the <i>μ</i>-Synthesis Applied to Lateral Control of Highly Flexible Aircraft

A lateral flight reduced-order controller for a flexible aircraft is presented. It is based on the reduced-order model of the flexible aircraft and theμ-analysis and synthesis toolbox. The first step deals with the order reduction of the full-model by using the balanced-truncation model reduction method. Then the reduced-orderHcontroller is achieved by using theμ-analysis and synthesis theory and applied to the reduced-order model and full-order model of the flexible aircraft. Finally, numerical results are presented and discussed. The simulation results show the efficiency of the reduced-order controller based on the reduced-order model since the close-loop system of the full-order model meets a set of realistic specifications in the frequency and time domains.

Integrated flight control system development using CEASIOM

Flight control system design typically involves hardware and software development, requiring the definition of both continuous and discrete design parameters. The control surface topology is an example of a discrete optimization problem, whereas the tuning of control system gains is an example of a continuous one. Contemporary design practice classically involves decoupling these two problems, however, they are intrinsically linked; for example, the maximum performance attainable by the control system software is determined by the control system topology employed. The cascaded effect is that no definitive sequential roadmap of development can be established, resulting in the requirement for an iterative framework using dynamic programming or optimization methods. This paper presents a framework for aircraft design together with the Flight Control System Designer Toolkit, and the Computerised Environment for Aircraft Synthesis and Integrated Optimisation Methods. The software platform was developed for the European Framework 6 Program Simulating Stability and Control and integrates both the reliability-driven hardware analysis and the performance-driven software design. A case study is presented, based on the NASA Boeing 747–100 model, in order to demonstrate the potential impact of introducing control design early into the design process and to show the intrinsic nature of the hardware–software FCS problem.

Multifidelity Design Optimization of Low-Boom Supersonic Jets

The practical use of high-fidelity multidisciplinary optimization techniques in low-boom supersonic business-jet designs has been limited because of the high computational cost associated with computational fluid dynamics-based evaluations of both the performance and the loudness of the ground boom of the aircraft. This is particularly true of designs that involve the sonic boom loudness as either a cost function or a constraint because gradient-free optimization techniques may become necessary, leading to even larger numbers of function evaluations. If, in addition, the objective of the design method is to account for the performance of the aircraft throughout its full-flight mission while including important multidisciplinary tradeoffs between the relevant disciplines the situation only complicates. To overcome these limitations, we propose a hierarchical multifidelity design approach where high-fidelity models are only used where and when they are needed to correct the shortcomings of the low-fidelity models. Our design approach consists of two basic components: a multidisciplinary aircraft synthesis tool (PASS) that uses highly tuned low-fidelity models of all of the relevant disciplines and computes the complete mission profile of the aircraft, and a hierarchical, multifidelity environment for the creation of response surfaces for aerodynamic performance and sonic boom loudness (BOOM-UA) that attempts to achieve the accuracy of an Euler-based design strategy. This procedure is used to create three design alternatives for a Mach 1.6, 6-8 passenger supersonic business-jet configuration with a range of 4500 n mile and with a takeoff field length that is shorter than 6000 ft. Optimized results are obtained with much lower computational cost than the direct, high-fidelity design alternative. The validation of these design results using the high-fidelity model shows very good agreement for the aircraft performance and highlights the need for improved response surface fitting techniques for the boom loudness approximations.

수직이착륙기의 착륙접근시 단일엔진고장 및 비행전이 영역 해석

본 연구에서는 착륙 접근시 단일엔진 고장 분석 및 비행전이 해석을 통하여 수직이착륙기에서 요구하는 최적의 엔진추력과 날개하중비를 구하고자 한다. 항공기 해석 모듈들은 기존의 항공기 통합 사이징 프로그램을 기반으로 하고 있으며, 항공기 설계 및 해석을 하나로 묶는 컴퓨팅 체계 구축은 사용자 용도에 맞게 주문 제작이 용이한 반 완성 프로그램인 EMDIOS를 설계프레임웍으로 사용하였다. 시뮬레이션 결과, 고도 40 ft 근처가 단일엔진고장시 가장 위험한 영역임을 확인할 수 있었으며, 비행전이 영역에 대한 파라메타 연구를 통하여 엔진 추력은 클수록, 반면에 엔진틸트각 회전속도와 날개 하중비는 작을수록 비행전이시의 고도 상실이 작아짐을 알 수 있었다. The objective of this study is to find the optimal thrust condition and wing loading of a vertical take-off and landing (VTOL) fixed-wing aircraft through a single engine failure analysis during landing approach and an analysis of transition flight. The aircraft analysis modules used in the study are based on the aircraft synthesis program. To achieve the computing infrastructure for aircraft design and analysis, the EMDIOS was employed as a design framework, which is a semi-completed application program and ready to customize. Simulation results reveal the most critical height at the event of single engine failure is approximately 40 ft. And, in order to avoid a significant loss in altitude during the transition, the thrust to weight ratio must be kept high, while both the engine tilt speed and the wing loading must be kept low, as confirmed by the analysis results.

Structural optimization for joined-wing synthesis

The joined wing is an innovative aircraft configuration with a rear wing, or horizontal tail, that is attached near the top of the vertical tail and sweeps forward to join the trailing edge of the main wing. This study evaluates two structural design methods for application in an aircraft synthesis code and quantifies the differences between these methods for joined wings in terms of weight, stress, direct operating cost, and computational time. A minimum weight optimization method and a fully stressed design method are used to design joined-wing structures. Both methods determine the sizes of 204 structural members, satisfying 1020 stress constraints and five buckling constraints. Monotonic splines are shown to be a very effective way of linking spanwise distributions of material to a few design variables. Five beam buckling constraints for the horizontal tail are included in both design methods. Without this constraint on buckling, the fully stressed design is shown to be very similar to the minimum weight structure. Adding a beam buckling constraint for the horizontal tail increased the structural weight by 13% and produced a fully stressed design that is 0.9% heavier than the minimum weight structure. Using the minimum weight optimization method to design the structure and to save 0.9% in weight required 20 times the computational time. Furthermore, the minimum weight structure produced only a 0.02% savings in direct operating cost. This study suggests that a fully stressed design method based on nonlinear analysis is adequate for a joined-wing synthesis study. The same joined wing considered in this study was shown, in an earlier study, to be slightly more expensive to operate than a conventional configuration designed for the same medium range transport mission. Since the same fully stressed design method was used in this earlier study, this work supports the comparisons of joined-wing and conventional aircraft performance presented in the earlier study. Of course, a different set of mission specifications and design assumptions may produce joined wings that perform significantly better.

Application of global sensitivity equations in multidisciplinary aircraft synthesis

The present paper investigates the applicability of the Global Sensitivity Equation (GSE) method in the multidisciplinary synthesis of aeronautical vehicles. The GSE method provides an efficient approach for representing a large coupled system by smaller subsystems and accounts for the subsystem interactions by means of first-order behavior sensitivities. This approach was applied in an aircraft synthesis problem with performance constraints stemming from the disciplines of structures, aerodynamics, and flight mechanics. Approximation methods were considered in an attempt to reduce problem dimensionality and to improve the efficiency of the optimization process. The influence of efficient constraint representations, the choice of design variables, and design variable scaling on the conditioning of the system matrix was also investigated.

Low cost computer graphics in engineering education

The Apple II is a personal computer which provides 6 color, raster graphics via a 280 h × 192 v frame buffer. A user-programmable MOS Technology 6502 CPU addresses 48 K bytes of RAM, 12 K of ROM. and 4 K of memory-mapped I/O. This system has been evaluated in a variety of applications for which one normally assumes higher resolution, higher cost graphics devices are needed. Examples were drawn from computer aided design, color microfilm previewing, cartography, and aircraft synthesis. The results were encouraging: low resolution raster scan graphics is surprisingly effective for a wide range of applications. As a result, a computer graphics laboratory, with a cluster of 10 Apples as its nucleus, was formed and has been functioning successfully for 4 terms.

Computerized aircraft synthesis.

A concept and philosophy of computerized aircraft synthesis is discussed, and Computer Program SYNAC (SYNthesis of AirCraft) is examined in some detail by considering examples of both its construction and its application. An aircraft synthesis computer program involves three principal functions: determination of force characteristics, determination of aircraft performance, and configuration control. There are three groups of variables: independent configuration size and shape variables; intermediate weight-, propulsion-, and aerodynamic-force characteristics variables; and dependent performance variables. Program SYNAC is organized in seven major modules: input, configuration control, geometry, weight, propulsion, aerodynamics, and performance. The program yields accuracies comparable with conventional aircraft synthesis techniques. Computer time per problem is currently on the order of 1 min, and ultimate times of y - of a minute are expected. A typical application of Program SYNAC involves the maximum-range optimization (in terms of wing area, wing thickness ratio, wing aspect ratio, wing taper ratio, and engine size) of a variable wing-sweep configuration for an air-to-ground mission.