methodology for design and evaluation of more …methodology for design and evaluation of more...

Methodology for Design and Evaluation of More Electric AircraftSystems Architectures within the AVACON Project

M. Jünemann, F. ThieleckeHamburg University of Technology, Institute of Aircraft Systems Engineering

Neßpriel 5, 21129 Hamburg, Germany

F. Peter, M. HornungBauhaus Luftfahrt e.V.

Willy-Messerschmitt-Str. 1, 82024 Taufkirchen, Germany

F. Schültke, E. StumpfRWTH Aachen University, Institute of Aerospace Systems

Wüllnerstraße 7, 52062 Aachen, Germany

Abstract

Within the research project AdVanced Aircraft CONfiguration (AVACON) a collaborative conceptual design of amid-range aircraft with ultra high bypass ratio (UHBR) engines and an envisioned entry-into-service in the year2028 is conducted. In this paper, the approach for overall architecting, sizing, and evaluation of aircraft on-boardsystems in AVACON is presented. To this end, important contributions to conceptual system design methodsproposed in the literature are reviewed to identify directions for methodological improvements. The role allocationof contributing partners and their methodologies for system design activities is described. Having differentcontributors guarantees, that tasks from overall aircraft to detailed subsystem design with a successive increasein the fidelity of system models is covered. In addition, a state-of-the-art baseline architecture is defined, whichwill serve as a starting point to conduct trade-off studies for investigation of the potential of system architectureconcepts and innovative technologies. Substantial system design requirements and novel boundary conditions,implied by the advanced aircraft configuration, are derived to give an outlook for planned technology studies.

1 INTRODUCTION

One of the major challenges during the complex taskof aircraft design is to coordinate the data flow andharmonize calculation results obtained by the relateddesign disciplines. This is especially true for distributedand collaborative design teams. Hence, it is necessaryto further advance processes for digital aircraft deve-lopment by synchronizing and improving methods anddata interfaces among the involved discipline experts.Within the project AdVanced Aircraft COnfigurations(AVACON), a consortium of industry partners, researchfacilities, and academia improve their joint capabilitiesfor the overall aircraft design (OAD) process during theconceptual phase. An optimized baseline concept ofa more-electric mid-range transport aircraft with ultrahigh bypass ratio (UHBR) engines and an envisionedentry-into-service in the year 2028 is conducted col-laboratively within the ongoing research project. Theobtained virtual aircraft model lays the foundation fortechnology studies regarding concepts for aircraft in-tended to serve the new mid-range aircraft market.

The idea of improving the design process in order todecrease development costs and the risk of poor de-sign decisions in the early design stage has alreadybeen subject to past and recent research [1–4]. In thecourse of the AGILE project, for instance, a collaborati-ve multidisciplinary design optimization environment isdeveloped and tested [4]. Hereby, the major emphasisis on developing and testing frameworks, which sup-port coordination and management of a collaborativemultidisciplinary design optimization process. This isachieved by means of consistent and harmonized toolinterfaces and the usage of knowledge-based techno-logies. Although an improvement of cross-site commu-nication of dedicated sizing and simulation tools is alsoregarded within the AVACON project, the main focuslies on improving a more expert-centric rather than atool-centric design optimization process. This expert-centric process captures cross-disciplinary effects andconsiders an increasing level of detail throughout theproject.

When studying the impact of innovative system archi-tectures on vehicle-level metrics like fuel efficiency and

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Deutscher Luft- und Raumfahrtkongress 2018DocumentID: 480197

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maintenance costs for a not yet investigated aircraftconfiguration, it is of considerable interest for the sys-tem design engineer to be supplied with sufficientlyaccurate boundary conditions specified by the indivi-dual design disciplines. For instance, a good estimateof the maximum hinge moment coefficient of the flapobtained by aerodynamic calculations is required forthe design of the high-lift actuation system. In addition,the flap geometry and its position on the wing sur-face estimated by stability and control considerationsis a determining factor. So is the available installationspace for the actuation system in the wing dictated bystructural wing design analyzes. Thus, novel systemboundary conditions induced by the considered aircraftconfiguration and technologies have to be identified. Aconstellation of partners for system design activitiesis presented in this paper. Contributions of differentpartners guarantee to consider these new constraintsand identify coupling effects between overall aircraftdesign and detailed subsystem design.

Based on an assessment of a suitable baseline sys-tems architecture, more detailed technology studiescan be conducted. Hereby, the baseline system is si-zed by means of available handbook methods adjustedwith state-of-the-art technology factors. The trade-offswill be focused on more electric aircraft (MEA) systemsarchitectures including concepts for power consumersand their supply systems. Assessment on subsystem-level and across system-system boundaries with vary-ing level of fidelity will contribute to deriving holisticallyoptimized MEA system architectures. System parame-ter sensitivities are then provided for overall aircraftdesign synthesis.

Since it is a major objective of the AVACON projectto improve design methods and processes throughoutthe project, the paper begins with a review of alreadyexisting methods for the conceptual design of aircraftsystems in Section 2. In this regard, potential researchtopics for methodological improvements are identified.In Section 3, the overall system design (OSD) me-thodology and the distribution of roles for partnerscontributing to system design activities is outlined. Tothis end, selected methods from these partners areoutlined. The reader is provided with the descriptionof considered technologies for system design trade-off studies and the definition of the baseline systemarchitecture in Section 4. The paper concludes withan outlook for the upcoming system design activitieswithin the scope of AVACON in Section 5.

2 REVIEW OF EXISTING APPROACHES TO CON-CEPTUAL SYSTEM DESIGN

A primary objective of conceptual system design is toprovide the overall aircraft design synthesis with suffi-

ciently good estimates of relevant system parametersand support early architectural decision-making onsystem-level. For aircraft systems, conceptual designactivities may thereby be subdivided into sizing andsimulation, architecting, and assessment.

Sizing and simulation

In order to estimate the size of a system or its com-ponents in terms of mass, maximum design load, orrequired installation space, several methods with va-rying level of fidelity can be found in literature. Foradvanced sizing methods, their results yield parame-ters for system component models, with which thesystem behavior can then be simulated to analyze andassess system performance over a given flight missionprofile.

In this paper, the reviewed approaches for sizing and si-mulation models are assigned a ranking order from 0 to2 to indicate their level of detail as it is listed in Table 1.Level-0 methods provide empirical system mass andpower consumption estimates. Level-1 methods rela-te to simple and advanced physics-based sizing mo-dels, which make use of regression-based componentscaling laws like power-to-mass ratios. The system iscommonly sized based on propagation of its designloads through the system’s power path componentsin reverse to the actual power flow. Hereby, individualcomponent efficiencies are considered. Thus, for theanalysis of a predefined overall systems architecture,the components are sized starting from power con-sumers continuing to power distribution componentsand ending up with the sizing of power generation sys-tems. The fidelity of Level-1 simulation approachesthereby ranges from simple power consumption esti-mates for mission segments like takeoff, climb, cruise,descent, and landing to mission-dynamic analyses ofquasi-static system models.

Current sizing and simulation methods in literaturecombine Level-0 and Level-1 methods. They levera-ge knowledge-based model libraries in order to reuseand scale parametrized component models. Systems,which are not sensitive to overall aircraft parametersor which have nearly constant power requirements li-ke the avionics system, are typically sized by meansof Level-0 methods. Systems with a high impact onoverall systems weight and a variable power consump-tion are modeled in more detail using Level-1 methods.However, if detailed conceptual system analyses arenecessary during early design evaluations, Level-2methods are used. They comprise detailed geometry-based system sizing optimization approaches to yieldthe physical dimensions of the components. Materialproperties like mass densities are then used to estima-te the mass of components. Analysis tools like FiniteElement Method (FEM) can thereby support the design

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engineer to calculate component masses. For Level-2system simulations, physical (acausal) modeling tech-niques are applied to solve ordinary differential equa-tions (ODE) or differential-algebraic equations (DAE)which represents the transient component and systembehavior. For physical component modeling and corre-sponding control design, extensive system knowledgeis required. Although these methods are not necessaryto perform conceptual architecture studies, they ena-ble the consideration of dynamic system requirementsearly in the system design process. Doing so can leadto more accurate component sizing results [5].

Level Characterization

0 - At system-level- Regression-based equations- No simulation of behavior- No volumetric sizing

1 - At system or component-level- Simple or advanced physics-based models- Static simulation models

(quasi-stationary system behaviour)- Scaling-law-based component sizing

(power densities and efficiencies)

2 - At component-level- Physical (acausal) modeling- Dynamic simulations

(transient system behaviour)- Geometry-based component sizing

(volumetric mass densities and efficiencies)- Optimization techniques for system sizing

TABLE 1: Definition of system model fidelity levels

Level-0 system mass estimation methods for aircraftdesign are well known and published for example byRoskam [6], Torenbeek [7], Raymer [8], and Luftfahrt-technisches Handbuch [9]. These methods provideempirical equations for aircraft systems, which contri-bute a relevant amount to the aircraft’s empty weight.These equations are of the form

(1) msys = f (θ AC,θ sys) .

That is, the system mass prediction msys is a functi-on of aircraft-related parameters θ AC (e. g. maximumtakeoff weight, wing span, etc.) and system-relatedparameters θ sys (e. g. flap chord length, landing gearstrut length, etc.).

Koeppen [10] compares a collection of these methodsand identifies the need for physics-based system mo-deling, if novel systems architecture concepts are as-sessed on aircraft level. As one of the first, he pro-poses a Level-1 approach for conceptual sizing and

simulation of the overall aircraft systems architecturein order to perform studies which capture couplingeffects between system-system boundaries and theirimpact on overall aircraft design. In his approach, air-craft on-board systems are sized through an analyticalbottom-up approach which avails knowledge-basedscaling laws for component weight prediction. An as-sessment of the influence of system parameters onoverall aircraft figures like takeoff weight is then con-ducted by coupling the systems design tool with theconceptual aircraft design environment PrADO [11].Koeppen’s focus, however, lies on analytical mass pre-diction methods rather than on simulating system’spower consumption. Therefore, Liscouet-Hanke [12]further advances Koeppen’s approach by introducinga model-based power-flow-oriented sizing and simu-lation method. The simulation results are aggregatedon overall aircraft level to perform MEA systems archi-tecture studies. Volumetric sizing and thermal aspectsare also addressed in her dissertation. Budinger [13]proposes a scaling-law based approach for derivationof estimation models, which are used for model-baseddesign. He considers parameters for installation spaceestimation, power sizing, and component dynamics.The approach is exemplified with sizing results of anelectromechanical actuator. Lammering [14] descri-bes a method to integrate suitable power flow orientedLevel-0 and Level-1 system sizing and simulation me-thods into the multidisciplinary integrated conceptualaircraft design optimization (MICADO) environment.The integrated approach enables the assessment ofsystem concepts on overall aircraft level consideringaircraft resizing effects. Chakraborty [15] further ad-vances these methods and complements them withheuristics for 3D power distribution network design. Inaddition, he investigates the use of probabilistic un-certainty quantification techniques for aircraft systemdesign. Chiesa et al. [16] also propose an analyticalsizing approach for generic sizing of several systemsarchitecture variants. These Level-0 and Level-1 me-thods are implemented within the tool ASTRID. Pra-kasha et al. [17, 18] and Fioriti et al. [19] apply thisframework, specifically to identify coupling effects bet-ween system parameters and overall aircraft designparameters.

Using Level-2 methods during conceptual design, Balset al. [20] present the concept of a virtual iron birdto model and simulate power demands of aircraft sys-tems architectures by means of the modeling languageModelica and an associated model library. Thereby, in-verse Level-2 models are employed, which solve theDAE for the inputs given the values for the outputs. Mo-reover, modeling and analysis of model uncertainty isconsidered in order to estimate simulation accuracy. AtLinköping University, researchers have been focusedon implementing Level-2 methods into the concep-tual aircraft design process. Krus et al. [21] present

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the application of the multi-domain simulation softwa-re Hopsan NG, which supports multi-core simulation.Due to this multi-core functionality it is suitable for acoupled dynamic simulation of aircraft systems, flightdynamics, and propulsion system during conceptualdesign. Likewise, parametric definition of aircraft sys-tems like the aircraft fuel system (c. f. Lopez et al. [22])is used to integrate conceptual system design into theknowledge-based aircraft geometry design tool RAPID.Within the RAPID framework, a joint CAD model of theaircraft and its systems is used for conceptual aircraftdesign and optimization [23]. In addition to the geome-try model, a model generated with Modelica is updatedby the geometry model and is executed to simulate thedynamics of the control surfaces [24]. The underlyingparametric actuation system model sizing approachis presented by Munjulury et al. [25] and Munjuluryand Puebla [26]. Safavi et al. [27] further advance thissimulation-centric framework to a collaborative MDOenvironment for aircraft conceptual system design. Anentire systems architecture sizing and simulation fra-mework with Level-2 methods, though, has not beendeveloped so far.

Most of the approaches stated before use simple Level-1 methods to conduct systems sizing. Yet, there is stilla lack of an integrated process for advanced Level-1and Level-2 methods during overall conceptual systemdesign. Detailed system design methods for system-specific domains, for instance, proposed by Dober-stein [28] (landing gear), Bauer [29] (primary flightcontrol), Pfennig [30] and Benischke [31] (secondaryflight control), Scholz [32] (flight control and hydrau-lics), Dunker [33] (hydraulics), Lüdders [34] (fuel cell),Annighöfer [35] (avionics), and Sielemann [36] (envi-ronmental control system) facilitate detailed systemdesign and optimization within an individual systemdomain. Seizing their results on the overall systemsand aircraft level as well as managing varying fidelityof analysis results has not been considered in the lite-rature so far. However, van Driel et al. [37] present aframework to combine component simulations of diffe-rent types of power flows to an overall aircraft powersystem optimization. He concludes that the implemen-tation of a local-global optimization is necessary tocombine system-specific optimization with the overallpower systems architecture-level optimization as it isalso described by de Tenorio et al. [38].

Architecting

In principal, aircraft systems architecting is concernedwith the efficient selection of an aircraft systems archi-tecture configuration from a large space of componentalternatives, that fulfill the required system functions inan optimal manner. Model-based systems engineeringtechniques have been introduced in conceptual sys-tems design, e. g., by Mohr [39]. Since their introduc-

tion, the idea of automatically generating alternativesof systems architecture configurations from a designspace of available system technologies is subject torecent literature. Armstrong [40] presents a methodfor automated generation of systems architecture al-ternatives from a functional viewpoint. His approachstarts with the definition of top-level functions, whichare then mapped to several combinations of physicalcomponents capable of fulfilling these functions. Theselected components themselves require, for instance,power supply. Thus, they induce new functions to befulfilled. This approach is called functional inductionapproach. An algorithm thereby performs a full enume-ration of all component combinations that may fulfill therequired product functions and considers predefinedconstraints that exclude specific component combina-tions. In his dissertation, de Tenorio [1] uses the sameidea and examines the additional use of the modelinglanguage SysML for architecture meta-modeling. Li-kewise, Jackson [41] adapts Armstrong’s approach todevelop a method for robust conceptual architectingstudies under parameter uncertainty. Chakraborty andMavris [42] integrate these design space explorationcapabilities into an architecting algorithm, which auto-matically connects architecture components based onidentified heuristics that consider redundancy require-ments. Fioriti et al. [43] use the design inputs statedby Chakraborty to perform a full enumeration of possi-ble systems architecture configurations. However, hediscards infeasible solutions manually in order to sizeand assess all feasible architecture alternatives. Anapproach for architecting fault-tolerant systems, whichconsiders redundancy aspects by means of reliabili-ty block diagrams (RBD) is proposed by Raksch [44].Based on this work, Bornholdt et al. [45] present theGeneSys methodology for function-driven design offault-tolerant aircraft systems architectures, which isoutlined in Section 3.3. As part of the methodology,the design space of possible architecture variants isgenerated and then down-sized by a rapid preliminarysystem safety assessment (PSSA) using RBDs. Theremaining feasible architectures, which comply withpredefined safety requirements, are then sized andevaluated to identify the dominant architecture concept.Similar algorithms for architecture design space explo-rations for aircraft system modeling frameworks arepresented by Garriga et al. [46] and Judt et al. [47–49].A generic algorithm to optimize the architecture ofcyber-physical systems (Level-2 models) is proposedby Finn et al. [50]. They demonstrate their approachwith a candidate search of an optimal aircraft environ-mental control system configuration.

The algorithms reviewed so far explore the designspace by generating possible architecture variants anddownsizing the design space by means of evaluatingconstraints (e.g. safety or redundancy requirements),using heuristics or relying on expert judgment. Auto-

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mated design space explorations can be supportive forsystem-specific trade-offs as can be seen from studiesby Bornholdt et al. [45]. A full enumeration of the entireaircraft systems architecture design space, however,implies an enormous effort in setting up the problem.In addition, automated design space explorations onOSD-level seem to over-examine the architecture defi-nition problem. They often yield dominant architectureconfigurations, that have already been envisioned bysystem designers. Hence, it is concluded that automa-ted design space explorations can be helpful. Yet, amanual design space definition by the tacit knowled-ge of an experienced engineer may lead to the sameconclusions more efficiently. As a result, the alreadyproposed methods are regarded as sufficient for futurearchitecture design studies.

Assessment

Available alternatives for architectures can be evalua-ted using either system-level or aircraft-level metrics.System-level metrics like the cumulative system massof the architecture, power consumption, and systemacquisition costs are typically used for comparison ofcompeting system concepts. In addition, Bornholdt etal. [45] use the technical dispatch reliability (TDR) asan operational assessment criterion. Hereby, RBDsare employed to calculate the probability of the sys-tem being able to comply with the master minimumequipment list (MMEL). As the MMEL is typically notavailable during the early design phase of novel aircraftconcepts, this criterion is only meaningful for retrofitstudies.

The system-metrics mentioned before are typically op-posing. Hence, the system designer has to select thebest alternative based on the generated Pareto-frontof non-dominated alternatives. However, the system-metrics can also be combined to a single aircraft-levelmetric. For this purpose, the impact of mass and powerconsumption on fuel efficiency can be evaluated bycoupling static system simulation models with detailed2D-mission simulation codes [12, 14, 15, 17, 51]. Forexample, Dollmayer [51] presents the mission simula-tion tool SysFuel, that interacts with the gas turbinesimulation program GasTurb. This interaction enablesthe estimation of the impact of systems on the actualengine thermodynamic state for each mission pointduring the flight simulation. Lüdders [34] advances thisapproach by implementing simple functions to predictaircraft resizing effects of system mass changes onstructural elements like wing, empennage or landinggear. Chakraborty [15] describes a similar procedurefor early assessment of aircraft systems by means of amission performance analysis and aircraft resizing ru-les. Lammering [14] proposes an integrated approachfor assessment of system’s impact on aircraft resizingeffects. To this end, a systems design tool is integrated

into MICADO. A detailed mission simulation is perfor-med taking into account average power consumptionestimates for every mission segment. In addition, thedetailed aircraft cost model proposed by Lammeringet al. [52] is able to combine system acquisition costs,fuel efficiency, and maintenance aspects into a sin-gle economic metric, namely direct operating costs.Bineid [53] proposes a top-down method for concep-tual aircraft system design considering reliability andmaintainability as contradicting aspects. Hereby, RBDsare utilized to predict aircraft dispatch reliability, allo-cate failure rates and component dispatch reliabilityfrom system to component level, and comparing actu-al to allocated component dispatch reliability in orderto identify the dominant architecture with respect toavailability (reliability and maintainability).

Despite the abundance of methods for system assess-ment on overall aircraft level, a technique to considerassembly costs as a target metric hasn’t been propo-sed in literature.

Future research directions

Conceptual systems design has advanced considera-ble in recent years. The approaches implement model-based systems engineering techniques and modelswith increasing fidelity level. However, a full OSD frame-work with methods ranging from Level-0 to advancedLevel-1 or Level-2 methods is still not available. In thiscontext, knowledge-based engineering (KBE) techni-ques become even more meaningful to ensure thereusability of component models, design pattern and3D-models. Moreover, dynamic system requirementson system components should be studied in more de-tail during conceptual design, since their considerationcontribute to more accurate predictions.Thermal models have already been addressed in the li-terature. The ongoing electrification of aircraft systems,however, requires that heat rejection from aircraft com-ponents is modeled in more detail in order to studyheat management strategies.The concept of propagating data from bottom-up sizing(detailed system design) to top-level assessment (over-all aircraft design) raises the question of how to mana-ge the data flow and the interpretation of aggregatedresults on the OAD level. For instance, this is the case,if the majority of the aircraft systems are modeled withsimple Level-1 methods, whereas a specific systemunder investigation is modeled with detailed Level-1 orLevel-2 methods. The overall systems design evaluati-ons then require to collate analysis data with varyinglevel of detail. In literature, however, means to estima-te the accuracy of an aggregated overall assessmentresult has not been scrutinized. The employment ofuncertainty management and analysis techniques withconsideration of parameter and model uncertainty cansupport the evaluation of study result accuracy within

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multi-fidelity design environments.Although system design tools have already been tes-ted in collaborative and distributed MDO frameworks,an integrated system and engine sizing procedure hasnot been proposed. Especially in the context of UHBRengines, an integrated design of aircraft systems andpower plant is required to study the sizing effects ofconventional and all-electric power extraction from theengine as it is discussed in Section 4.In addition, there is still a lack of a standardized da-ta exchange format for the definition and characte-rization of systems architectures comparable to theCommon Parametric Aircraft Configuration Schema(CPACS) [54].An approach to conceptual design of aircraft systemswith a special focus on design for optimal assemblycosts is missing. A consistent methodology conside-ring and trading opposing assessment criteria for fuelefficiency, assembly, and availability should be develo-ped.

In summary, following directions can be identified fromliterature for improvement of conceptual system designmethods in future research:

• Develop an overarching multi-fidelity methodologyfor conceptual system design, which uses methodsfrom Level-0 to Level-2

• Consider effects of dynamic system requirementson component sizing

• Advance thermal models applied during conceptualsystem design

• Establish an uncertainty quantification and manage-ment strategy for multi-fidelity aircraft system design

• Investigate an integrated system and engine sizingprocedure on the overall aircraft level

• Define a standardized data exchange format forsubstantial systems architecture information

• Examine advanced assessment criteria for more in-tegrated architecture studies considering the effectsof fuel efficiency, assembly, and maintenance

3 OVERALL SYSTEMS DESIGN METHODOLOGYWITHIN THE AVACON PROJECT

A holistic approach to conceptual system design co-vers tasks ranging from overall aircraft to detailed sub-system level. For this purpose, different methods fromcontributing partners are combined for the systemsdesign activities within the AVACON project. The rela-ted allocation of responsibilities for this constellation isdepicted in Section 3.1. Methods for all-electric Level-0 system sizing, overall systems architecting, sizingand assessment, and integrated system design of a

hybrid laminar flow control system are outlined in Sec-tions 3.2, 3.3, and 3.4, respectively.

3.1 Constellation of system design partners

System design tasks are subdivided into overall air-craft design (OAD), overall systems design (OSD), anddetailed subsystem design (DSD). As depicted in Figu-re 1, Bauhaus Luftfahrt e.V. (BHL) is mainly concernedwith design activities on OAD and OSD level. This in-cludes, for instance, technology assessments on OADlevel and systems design with Level-0/1 methods. Themain focus of the Institute of Aircraft Systems Engi-neering (FST) at Hamburg University of Technology(TUHH) is to govern the physics-based sizing, simu-lation and assessment approach on the OSD- andDSD-level. Moreover, FST conducts detailed designstudies (Level-2) for selected system domains, whichare discussed in Section 4.

AIRBUS

Detailed Subsystem Design(DSD)

Overall Systems Design(OSD)

Overall Aircraft Design(OAD)

CollaborativeRoles

RWTH

HLFC designSupervision

TUHH

BHL

FIGURE 1: Collaborative roles within AVACON system de-sign workpackage

Both partners thereby contribute towards applicationand improvement of sizing and simulation methods.BHL uses a dedicated Level-0/1 sizing method forall-electric systems architectures, which is outlinedin Section 3.2. Instead, FST improves Level-1 sizingmethods and successively integrates Level-2 sizingmethods into the overall systems design and assess-ment methodology GeneSys presented in Section 3.3.This partnership constellation enables a design pro-cedure that starts with top-down analyses serving asa starting point for detailed bottom-up system designactivities. The results from detailed studies are thenfed back to OAD synthesis by means of system para-meter sensitivities. With the sensitivities, a holisticallyoptimized systems architecture for the target aircraftcan be derived.For the sizing of the hybrid laminar flow control (HLFC)system, an integrated approach is necessary. The sys-tem laminarizes the flow by means of boundary layersuction and thereby improves aerodynamic performan-

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ce. However, the suction distribution and the amountof suction flow determines the size of the system com-ponents like the air compressors. Thus, a sizing andassessment on overall aircraft level has to be con-ducted to find the optimal system size. To this end, adedicated method developed by the Institute of Aero-space Systems (ILR) at RWTH Aachen is applied. It isdescribed in Section 3.4.The industry partner AIRBUS revises the definitionof requirements and assumptions for state-of-the-artcomponent properties to assure the quality of results.

For exchange of systems architecture information andquantitative data of requirements, an interface bet-ween involved system design partners is needed. It istherefore an objective to contribute to a standardizeddata exchange format suitable for future implemen-tation in the CPACS standard, that holds substantialinformation about systems architecture properties.

3.2 Methodology for Level-0 sizing of all-electricaircraft systems architectures (eFlow)

The performance assessment of aircraft concepts withan envisioned entry-into-service date for 20 to 50 yearsin the future is a frequent task at Bauhaus Luftfahrt e.V.(BHL). Hereby, the incorporation of a realistic systemsarchitecture and the assessment of its performance isa vital component of a sound overall aircraft evaluation.For the investigation of future system architecture con-cepts, alternative energy sources (e.g. fuel cells) aswell as major changes in the consumers (e.g. hybridand all-electric propulsion architectures) have to beconsidered. Often there are not many details availablefor the concepts in this early design stage. This ma-kes it necessary to provide a methodology for sizingand assessment with a minimum amount of input da-ta. For this reason, the Energy Flow Analysis (eFlow)methodology has been developed at BHL. The eFlowmethodology is a set of methods to estimate weightand power consumption for all electric systems archi-tectures including advanced components. These me-thods are mainly based on Level-0 methods. Level-1methods are applied for components that are of majorimportance for new architectures as, for instance, cir-cuit brakers, converters, and cables. The methods areimplemented in MATLAB. They enable the calculationof power loads for a mission comprised of the seg-ments taxiing, takeoff, climb, cruise, descent, diversion,and landing. Furthermore, the consumers are groupedinto non-essential, essential, and vital loads in order tosize the architecture according to critical failure cases.An interface for an automated import of data from thepreliminary aircraft design platform APD Pacelab isavailable. As mentioned, the ability to execute a preli-minary estimation of loads and masses of the systemsarchitecture with minimum input data is important inorder to consider its impact in early conceptual design

stages. Therefore, all implemented methods can beexecuted using data from reference cases, which arescaled by a few aircraft characteristics of the conceptunder consideration (e.g. maximum take-off weight,wing and empennage area, passenger number, de-sign range, and mission segment duration). Electricalconsumers of conventional systems architectures likeavionics and instrumentation are modeled with me-thods taken from the literature [55]. For substantialelectrical consumers of a MEA systems architecture,methods from literature have been gathered to esti-mate their power demand and mass. For example,the flight control actuators are considered as electro-mechanical actuators [56] within eFlow. The landinggear is modeled assuming a local hydraulic system foreach gear [57] and the wing de-icing system is imple-mented as an electro-thermal de-icing system [58].

Within the scope of AVACON, the existing set of me-thods will be further improved and extended. Methodsfor the non-electrical components and for conventionalsystem elements will be implemented (e.g. environ-mental control, flight controls and de-icing) in orderto enable a sound analysis of a baseline systems ar-chitecture. One of the major objectives of AVACON isto improve the exchange and cooperation of aviationresearch entities in Germany. For this reason an in-terface with CPACS has been implemented to reduceeffort and necessary time to analyze the design itera-tions of the aircraft concepts in AVACON. In order toexploit all information included in the AVACON aircraftconcepts, an extended mode for the methods will be re-garded, which enables the incorporation of all availableaircraft data (e.g. specific flight control surface geome-tries). Furthermore, a transfer of tool implementationto Python will be considered. This could enable theincorporation of CPACS support functions used by theaircraft conceptual design framework VAMPzero deve-loped by the German Aerospace Center (DLR) as wellas increased flexibility by enhanced object-orienteddata structure.

3.3 Methodology for overall systems design andassessment (GeneSys)

The modular methodology toolbox for overall aircraftsystems design and assessment GeneSys is proposedby Bornholdt et al. [45]. Main steps of the GeneSysprocedure are illustrated in Figure 2. The first twosteps are concerned with exploring the possible de-sign space. Based on a top-down functional break-down starting from top-level aircraft requirements, thecombinatorial design space of architecture variantsfrom a set of physical components, that may fulfillthe required system functions, are identified (Step 1).A rapid PSSA by means of aircraft system RBDs isthen performed to withdraw architecture variants, thatdo not comply with predefined system safety require-

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ments (Step 2). The remaining systems architecturesare then sized using either Level-0/1 sizing methodsor surrogate models of system-specific Level-2 sizingtools (Step 3). An assessment of feasible architectureswith respect to mission performance and predicted air-craft resizing effects (Step 4) can be performed usingthe dedicated mission simulation tool SysFuel+. If amaster minimum equipment list is available, an ope-rational assessment based on the technical dispatchinterruption rate (TDIR) can be conducted (Step 5).

AutomatedSafety Assessment

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Generation ofarchitecture variants

1

System sizingand simulation

3

Assessment onoverall aircraft level

4

Operationalanalysis

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zusätzlicheDaten

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Prelim. system safetyassessment

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- Definition of safety requ.- Construction of RBDs

Mapping to physicalcomponents

-

- Analysis of requirements- Functional breakdown

- 2D-mission simulation- Aircraft resizing effects

A) SysFuel+ interfaceB) Response Surf.

- MMEL entries- Tech. Dispatch Interruption Rate

FIGURE 2: Aircraft system design and assessment proce-dure of the GeneSys methodology

GeneSys has already been applied, inter alia, to opti-mize electro-hydraulic power distribution systems [59],for conceptual design of an all-electric actuated flightcontrol system and its corresponding electrical powerdistribution system [45], and for a power allocationoptimization of a Sakurai flight control system con-cept [60].

Systematic improvements for GeneSys

In the course of AVACON, the GeneSys methodologywill be advanced to a framework with following additio-nal capabilities:

• Overall systems architecting and system sizing withimproved methods using models ranging from lowto high fidelity

• Management of multi-fidelity analyses consideringparameter and model uncertainty

• Advanced interfaces for integration in an overar-ching collaborative and distributed aircraft designenvironment

• Assessment based on refined operational criteria

• Visualization of architectures based on a standardi-zed systems architecture data exchange format

A major objective of AVACON is to increase the in-tegration of aircraft design disciplines. To this end,the GeneSys methodology will be improved for app-lications within collaborative and distributed aircraftdesign setups by means of dedicated interfaces, forinstance, to CPACS as depicted in Figure 3. Integratedstudies can then be conducted in order to identifycoupling effects between systems and overall aircraftdesign. Hereby, an integrated systems and enginesizing approach will be investigated with respect toconventional and all-electric engine secondary powerextraction, as it is discussed in Section 4.

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FIGURE 3: Planned interface of GeneSys for integrated as-sessment on overall aircraft level

GeneSys is fed with information from a collection ofsizing methods with a large variety of fidelity. A first esti-mate for overall systems architecture mass and powerconsumption can be obtained by means of Level-0sizing methods based on empirical relations. For rapidphysics-based architecture studies, an analytical Level-1 method proposed by Koeppen [10] is available. Theunderlying concepts will be improved and updated inthe course of the project. The foundation of GeneSys,however, is its accessibility to Level-2 design tools forsystem-specific domains, which have been developedat FST. Although Level-2 design tools have alreadybeen integrated in GeneSys by means of simple surro-gate models during prior studies, a strategy to manageresults with varying level of detail from several systemanalyses has not been developed. Thus, it is pursued

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within the course of AVACON. In addition, the impactof uncertainty on the interpretation of study results willbe quantified by means of uncertainty analysis andmanagement capabilities. To this end, more advancedsurrogate modeling techniques, suitable for propaga-ting uncertainties from subsystem to overall aircraftlevel, can be applied.

3.4 Method for integrated assessment of HLFCsystem architectures on OAD level

The main functionality of a hybrid laminar flow control(HLFC) system is the reduction of viscous aerodyna-mic drag. This is achieved by a laminarization of theflow along the aircraft surfaces using optimized ae-rodynamic surface shapes and applying active flowsuction on the surface skin. The aerodynamic benefit,however, comes with an increase in mass and powerconsumption of the suction system. Compressors, pi-pes and wiring are required to provide a suction massflow that energizes the transient area of the boundarylayer. Hence, it is necessary to perform an integra-ted sizing and assessment of the HLFC system toconsider cross-disciplinary coupling effects betweenaerodynamics and system design on OAD level.

Such an integrated HLFC design method has be-en developed by the Institute of Aerospace Systems(ILR) at RWTH Aachen University. It is used withinthe AVACON project. The method follows a sectionalapproach (quasi 3D approach) which divides the 3Dobject of the wing into several 2D sections. The geome-tric properties of each section is used to calculate therelated flow characteristics. This yields the pressuredistribution as well as lift and drag coefficients for eachsection along the wing span. The method supportscalculations considering fully turbulent flow as well aspartly laminar flow. A detailed description of the HLFCdesign process is given by Risse [61].In order to set a specified suction distribution along thechord of a wing section, the wing has to be perforatedand compressors have to generate a certain suction-pressure. These compressors require electrical powerprovided by the electric generators of the aircraft. Thesuction flow is thereby led through pipes to the outflowvalve at the wing-fuselage intersection. The sizing ofthese components is based on the simplified suctionconcept developed during the ALTTA project [62,63].For this concept, Pe and Thielecke [64] propose me-thods and equations for the HLFC system components.They are implemented in the HLFC system method ofILR. The input design variables of the HLFC systemdesign module are the pressure and suction distribu-tion for the design Mach number and design altitudeof the HLFC system. The most important outputs ofthe HLFC system sizing are the total HLFC systemmass mHLFC,tot and the total electrical power requiredto drive the suction system PHLFC,tot . The total system

mass is composed of the mass of the compressors,motors, variable frequency drives, the pipes of theducting system, and the electric wiring. The requiredsystem power depends only on the required power ofthe compressor and the efficiency of the motor andvariable frequency drive.

In the following, the influence of the mass flow onboth the aerodynamic performance and on HLFC sys-tem mass and power consumption is discussed. Tothis end, a preliminary sizing study of a wing-installedHLFC system for the AVACON research baseline ispresented.

Discussion of preliminary system sizing results

For the AVACON baseline aircraft, the HLFC systemcan be installed in different ways. Either a central archi-tecture with a single large compressor or a decentrali-zed architecture with several smaller compressors canbe used. If a decentralized architecture is chosen, thenumber of compressors can be adapted. Further, it canbe selected if the compressors are either connected toa separate or a collective duct. For the AVACON rese-arch baseline, a decentralized architecture with threecompressors and a collective ducting system is chosen.Based on this reference configuration, several studieswere carried out to test the influence of the suctionsystem on the aerodynamic performance (measuredin relative change of lift-to-drag ratio L/D), the sys-tem mass mHLFC,tot, and the system power consumpti-on PHLFC,tot. In this study, the cruise Mach number ofMa = 0.83 and the initial cruise altitude ICA = 35000ftof the AVACON research baseline are taken as missi-on design point for the HLFC system. Thereby, thesystem is sized for different distributions of the suctionvelocity along the chord of the wing section.

The results are presented in Figure 4, Figure 5, and Fi-gure 6. First, the interrelationship between a possibleaerodynamic performance increase and the requiredmass flow obtained by an aerodynamic analysis isshown in Figure 4. It can be observed that differentsuction distributions can result in the same aerodyna-mic improvement. Thus, an optimization of the suctionalong the wing chord is necessary. The pareto-optimalpoints of optimal suction-distributions is highlightedwith a solid line. Based on this pareto-front, it can beobserved that an increase in the suction mass flowimproves the aerodynamic performance.

In Figure 5 and Figure 6, the influence of the suctionmass flow on the HLFC design in terms of systemmass and power demand is shown. As already seenin Figure 4 for the relative L/D change, not every in-put value (mass flow) of the system sizing results in adiscrete output value (system mass). This means thatthe system parameters mass and power consumptiondepend on the mass flow on the one hand, but also on

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the suction distribution itself and the pressure distribu-tion (i.e. lift coefficient) on the other hand. However, anincrease in suction mass flow yields higher values forsystem mass and power consumption, since a morepowerful suction system requires, for instance, morepowerful compressors.

As can be seen from the study results, the HLFC sys-tem design objectives, i.e. high aerodynamic systemperformance and low system mass and power con-sumption, are contradictory. Thus, only an integratedassessment of these aerodynamic and system designsensitivities on OAD level can lead to a sound HLFCsystem design optimization.

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FIGURE 6: Boundaries of total HLFC system power as afunction of required mass flow

4 CONSIDERED SYSTEM TECHNOLOGIES FORARCHITECTURE TRADE STUDIES AND BASE-LINE SYSTEMS ARCHITECTURE

Employing over-wing UHBR engines in conjunctionwith a hybrid laminar flow control (HLFC) system forthe main wing implies novel requirements for sys-tem design. In the following, concepts for relevantsystem domains are described, that satisfy these re-quirements. In addition, system technologies for theAVACON baseline systems architecture are selectedwith respect to a close resemblance to the A350 air-craft systems architecture.

4.1 Power supply systems

Aircraft systems require to be supplied with either me-chanical, hydraulic, electrical or pneumatic power. Acommon source is the aircraft engine. It provides se-condary power for the aircraft systems beside the pri-mary propulsive power. The engine is regarded as themain source for secondary power within the scopeof AVACON as well. The concept of a MEA systemsarchitecture is thereby considered as the target con-figuration for all architecture trade-off studies. In thiscontext, the two extremes of engine secondary powerextraction architectures – i.e. a conventional and anall-electric concept – are described in the following.They serve as starting points for more detailed powersupply architecture studies.

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Conventional engine secondary power extraction

A conventional concept to extract secondary powerfrom the engine is shown in Figure 7a. Hereby, engineshaft power is converted to electrical and hydraulicpower by an integrated drive generator (IDG) and anengine driven pump (EDP), respectively. These powerconversion components are mounted on an accessorygearbox (AGB) driven by the engine shaft. For conven-tional electrical power supply of large transport aircraft,the combination of a 115/200V AC system with eithera nominal 400Hz constant frequency or a variable fre-quency (VF) ranging from 360 to 800Hz, and a 28V DC

power supply is still an industry standard for existingaircraft. However, concepts for high voltage alternatecurrent (HVAC) at 230V AC and high voltage direct cur-rent (HVDC) at +/- 270V DC are considered to enablelighter and more efficient power distribution and ma-nagement systems [65]. The EDP generates hydraulicpower from mechanical shaft power at a constant no-minal pressure level, which is defined as 3000 psi and5000 psi for traditional and modern large transportaircraft, respectively. Electrical and hydraulic powersupply systems are designed with redundant supplycircuits to comply with safety requirements for fault-critical system functions like flight control actuation.

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FIGURE 7: Examples of power path architectures from power generation to power consumption

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Therefore, an optimal architecture of the electrical andhydraulic system is tightly coupled with the architectureof the primary and secondary flight control actuationsystem and has to be determined with dedicated ar-chitecture trade studies.

Moreover, bleed-air is extracted from the intermediateand high pressure compressor of the engine in or-der to provide pneumatic power to systems like theenvironmental control system (ECS) or the ice pro-tection system (IPS). The air extracted from the twocompressor stages is mixed and regulated to requiredtemperature and pressure levels.

Additional energy conversion components ensure safeoperation in case of failure of a single circuit or du-ring ground operation. For the hydraulic system, forinstance, an electric motor pump (EMP) converts elec-trical to hydraulic power, whereas a constant speedmotor/generator (CSM/G) converts hydraulic to electri-cal power. An air driven pump (ADP) can be employedto transform pneumatic to hydraulic power. An auxi-liary power unit (APU) provides additional bleed airfor ground supply, engine start capability, and pneu-matic power supply redundancy. The correspondingAPU starter/generator supplies electric power duringground and abnormal operations.

All-electric engine secondary power extraction

Although the described conventional engine seconda-ry power extraction principle has been the dominantconcept for large transport aircraft, there are at leasttwo drivers in the light of the AVACON target confi-guration for a more electric engine power extractionconcept:

1. Further electrification of aircraft systems due tobenefits originating from power management stra-tegies, advanced maintenance and monitoring ca-pabilities, and improved system reliability

2. Reduced pneumatic power extraction efficiency dueto unfavorable bleed-air conditions when extractedfrom an advanced UHBR geared turbofan engine

The all-electric engine power extraction concept depic-ted in Figure 7b overcomes these shortcomings. Theextraction of compressed bleed air from the engineis thereby replaced by a motorized turbo-compressor(MTC) supplied by the EPS. In this example, electricalpower is provided by a variable speed constant fre-quency starter/generator (VSCF S/G), which makesuse of dedicated power electronics for transformationfrom variable to constant frequency. It also ensures en-gine start functionality. More advanced concepts consi-der an engine-mounted starter/generator to eliminatethe AGB and thereby provide even higher efficiencyand reliability. However, this advanced all electric engi-ne (AEE) concept is not considered within the scope

of AVACON. Between the two extremes, more electricengine (MEE) concepts are subject to recent rese-arch. They consider to partly replace the mechanically-driven hydraulic and pneumatic power generation withelectrical machines. Since this replacement reducescomplexity of the power extraction concept, it improvesdispatch reliability and maintainability [66]. Hence, theMEE is a potential candidate for systems architecturestudies in AVACON.

Electro-hydraulic power system architecture

With regard to more electric engine power extraction,the hydraulic system can be considered as an elec-tric consumer. Instead of EDPs, the electro-hydraulicsystem power is then supplied by compact hydraulicpower packages (HPP) which integrate an EMP withessential components of conventional hydraulic sys-tems in a single module. While Trochelmann et al. [5]conduct a trade-off with respect to central and dis-tributed electro-hydraulic power generation systemswith zonal HPPs for a short-range aircraft at a nominalpressure level of 3000 psi, a related study has notbeen performed for a mid-range aircraft. Hence, theseconcepts will be subject of AVACON trade-off studies.

Indeed, a conventional engine power extraction con-cept and a power supply architecture designed to ob-tain Extended Twin Operations (ETOPS) type approvalis selected for the AVACON baseline systems architec-ture. It is characterized by following technologies:

• 2 hydraulic circuits with 5000 psi constant pressurecontrol supplied by 1 EDP per engine

• 3 electrical distribution networks

– 230V AC VF supplied by 2 engine generatorsper engine

– 115V AC converted from 230V AC VF by autotransformer units (ATU)

– 28V DC network supplied by the 230V AC net-work with transformer rectifier units (TRU)

• A 2H/2E power distribution architecture for the 230VAC electrical circuit and 5000 psi hydraulic circuitsatisfies the safety requirements for flight controland landing gear functions.

• Bleed air extraction from intermediate and highpressure compressor stage and pre-conditioningvia a precooler and dedicated valves

• 1 APU with a starter/generator 230V AC at 400Hzconstant frequency for electrical backup

• 2 EMPs for ground power supply and electro-hydraulic backup

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4.2 Environmental control system and ice pro-tection system

For large aircraft, the selection of technology conceptsfor ECS and IPS depends on the the underlying powersupply system architecture. The primary functionalityof the ECS is to supply fresh air for on-board occu-pants as well as to provide pressurization, temperaturecontrol, and ventilation for selected compartments. Tosatisfy these requirements, conventional ECS are sup-plied with bleed air, which is conditioned by bootstrapair cycle machines. Via a mixing unit and a ducting net-work with integrated recirculation and ventilation fans,the conditioned air is distributed to the correspondingcompartments and regulated according to requiredconditions. However, if bleed air is eliminated due to aMEE power extraction concept, the air cycle machinehas to be provided with compressed outside air deli-vered by a dedicated ram air inlet. To this end, an MTCcan be utilized, which is powered by the electrical sys-tem. For this concept, an additional use of vapor cyclesystems have to be considered in order to guaranteeconditioning during ground operations.The main function of the IPS is to enable safe opera-tions in icing conditions by protecting, inter alia, wingleading edge and cockpit windshields. For conventio-nal systems architectures of large transport aircraft,the wing ice protection system (WIPS) uses hot bleed-air from the engines to heat part of the leading edgearea. Impinging supercooled water droplets are eva-

porated. In addition, electrical heater wires embeddedinto the windshields heat the cockpit windows. In theabsence of bleed-air extraction, the functions of theWIPS can be adopted by electro-thermal solutions.Electrical heater mats, which are electrical resistorsconnected to the high voltage electrical circuits, areconsidered in this case.An electrical IPS constitutes the main peak power con-sumer of a MEA architecture. Thus, if a more electricengine secondary power extraction is chosen, hybriduse of bleed air and MTC will be investigated to en-hance engine power extraction efficiency.

For the baseline systems architecture, the followingsystem variants are selected:

• Bleed-air supplied ECS with bootstrap air cycle ma-chines

• Thermal bleed-air WIPS and electrical heating forcockpit windows

4.3 Flight control system

The control surface layout resembles a conventionalsplit of primary and secondary flight controls for theAVACON baseline and target aircraft. It is defined bythe responsible OAD work package. However, the se-lection of actuation concepts and their allocation to thepower supply are part of system design activities.

Primary flight control system (PFCS). The main functi-

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on of the PFCS is to provide roll, pitch, and yaw controlduring flight. In addition, load-alleviation systems havebeen developed to relieve structural loads for maneu-vering and gust turbulence, which enables certificationfor higher maximum takeoff weights. Since the entireFCS has to be designed for fault-tolerant operation,flight control actuators have to be allocated to indepen-dent power sources (e.g. to multiple circuits of a powertype) to comply with safety requirements.To satisfy these requirements, several actuation tech-nologies are available, which are described in detailby Maré and Fu [67]. Next to traditional hydraulic ser-vo actuators (SA), electro-hydrostatic actuators (EHA)are considered in the course of AVACON. Hereby, theactuator’s position is controlled by an electro-hydraulicmotor-pump unit. A hybrid solution of EHA and SA isthe electrical backup hydraulic actuator (EBHA). It isconnected to both electric and hydraulic circuits. In ad-dition, the use of electro-mechanical actuators (EMA)for front-line operation, which are connected to theelectrical power supply system, will be examined.

Secondary flight control system (SFCS). The mainfunctionality of the high lift system is to augment liftduring takeoff and landing. More advanced systemsalso provide differential flap setting and variable cam-ber for performance optimization, especially duringcruise, as it has been developed for the Airbus A350XWB-900 [68]. Hence, the high lift actuation systemhas to be designed to assure safe control surface mo-tion respecting maximum hinge moment, maximumdeflection rate, available installation space, and con-trol surface kinematics. To this end, the SFCS caneither be realized as a centralized architecture, wherea central drive unit (power control unit, PCU) drives thecontrol surfaces via a mechanically-driven shaft trans-mission system, or a distributed actuation concept,for which the actuation can be fulfilled by distributedmechanical, electro-hydraulic or electro-mechanicalpower trains [69]. If differential flap setting functionalityis a system requirement for centralized mechanically-driven flaps, an active or passive differential gear boxhas to be considered to enable a controlled asynchro-nous flap movement of adjacent surfaces.

An UHBR engine and HLFC technology installationimplies additional design constraints. Due to its largeengine diameter, a close mounting of the engine un-der the wing is inevitable. This may cause prematureseparation of the air flow and calls for employmentof flow control devices [70]. In addition, the slat me-chanism has to be selected such that its functionalityis not affected by the nacelle. A gapless droop nosedevice for inboard slats will be taken into account ins-tead of rack and pinion or Shielded Krüger kinematicsfor inboard slats. This also minimizes noise producti-on [68]. The jet wake may also impact the load on theflap, which may favor new moveable concepts with me-

chanically decoupled flaps [70]. If an over-wing engineconfiguration in conjunction with a HLFC system is con-sidered, their laminarization effects may facilitate thedesign for thin aircraft wing profiles. Thus, installationspace for flight control actuation can be limited. Mo-reover, the HLFC technology requires Shielded Krügerflap kinematics for the leading-edge high lift system.Thus, Shielded Krüger flaps will be assumed as an ad-ditional kinematics concept for outboard slats in futurestudies.

As stated before, a 2H2E-configuration is selected forthe baseline systems architecture considering two hy-draulic circuits (G,Y) and two electric circuits (E1,E2).The baseline flight control system architecture with cor-responding allocation to the supply system is depictedin Figure 8. Hereby, flaps with dropped-hinge kine-matics are assumed. For the inboard slats, gaplessdroop nose devices driven by geared rotary actuators.Rack and pinion kinematics with sealed slats for theoutboard leading-edge high lift devices are assumed.

4.4 Landing gear system

An over-wing configuration of UHBR engines is consi-dered for the AVACON target aircraft. This configurati-on enables the integration of unconventional landinggear concepts, which will be examined in a dedica-ted AVACON work package. The related methodology,concepts and requirements are outlined in detail byKling et al. [71]. However, for the AVACON baselinesystems architecture, a wing-mounted landing gearwith redundant 2H-supply for braking functionality anda 1H-supply for wheel steering and for extension/re-traction functionality is considered.

5 CONCLUSION AND OUTLOOK

In this paper, a methodology for system design andassessment of aircraft systems architectures for theAVACON baseline and target aircraft has been descri-bed. It combines expertise of partners contributing tosystem design activities with design and assessmentmethods ranging from low to high fidelity. To this end,an empirical method for all-electric system architec-ture sizing, an overall system architecting, sizing andassessment tool chain, and a dedicated HLFC sys-tem sizing and assessment approach of contributingsystem design partners have been presented. Basedon a comprehensive literature review of available ap-proaches to aircraft conceptual system design, a list ofdirections for methodological improvements in futureresearch has been proposed.A majority of these points will be addressed within thescope of the AVACON project. In addition, technologytrade studies will contribute to derive a holistically op-

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timized aircraft systems architecture for the AVACONtarget aircraft. As a next step, the AVACON baselinesystem architecture will be sized with improved low tomedium-fidelity methods using state-of-the-art tech-nology factors. Starting from this baseline, selectedsystem-specific trades will be conducted to iterativelyconverge at a dominant system architecture. This com-prises studies to optimize more-electric power supplysystems. In this regard, the identification of coupling ef-fects between system and engine sizing in the light ofconventional and all-electric engine power extractionwill be considered.

CONTACT INFORMATION

Marc JünemannHamburg University of TechnologyInstitute of Aircraft Systems EngineeringNeßpriel 5, 21129, Hamburg, [email protected]

ACKNOWLEDGEMENT

The results of the presented paper are part of the workin the research project AdVanced Aircraft CONfigu-ration (AVACON), which is supported by the FederalMinistry of Economic Affairs and Energy in the nationalLuFo V program. Any opinions, findings and conclu-sions expressed in this document are those of theauthors and do not necessarily reflect the views of theother project partners.

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