26TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES

A SIMULATION FRAMEWORK FOR POWER SYSTEMS ARCHITECTING

Susan Liscouët-Hanke*# *Research-Engineer, Airbus #PhD candidate, Université de Toulouse, INSA - UPS, Laboratoire du Génie Mécanique de Toulouse, France

Keywords: Aircraft Power Systems, Architecting, Preliminary Design, Modeling & Simulation, More Electric Aircraft

Abstract GEN Generator IFE In-flight Entertainment Today, the preliminary design process of power ISA International Standard Atmosphere system architectures of civil aircraft is usually RAT characterized by the separation into ATA (Air WIPS Wing Transport Association) chapters. However, the complexity of an aircraft energy network, the 1 Introduction large number of influencing design-parameters and system interfaces require a common and The increasing complexity of modern civil transparent process if a meaningful evaluation aircraft in a constantly evolving, ultra of different system architectures with regards to concurrently environment of technological, the overall aircraft efficiency is to be achieved. regulatory, economic and ecological challenges The development of a dedicated leads the aircraft manufacturers to seek for more methodology, a simulation framework, and and more optimized solutions for aircraft adapted modeling techniques is the objective of architectures. A global ‘aircraft level’ approach the presented research. This article focuses on is seen as the most promising, overcoming the the dedicated modeling approaches that are today’s system per system optimization. developed in order to analyze systems at Energy saving is one of the most aircraft-level. Three different modeling important aspects in a highly competitive techniques illustrate on the one hand the effort market with increasing fuel prices and the required to develop adapted models to fit in the growing importance of environmental friendly proposed analysis environment. On the other transport. Thus, the optimization of the aircraft hand, the added value of such an integrated power system architecture with regards to modeling approach is demonstrated with the energy consumption is an important issue [1] examples of the sizing beside the improvement of the analysis and the link of power systems and aerodynamic performance. simulation to a global aircraft thermal model. The aircraft power system architecture is a complex network of interacting systems Abbreviations fulfilling all different functions. In order to structure in this difficult task, the so-called APU ATA-chapters classification, established in 1936 ATA Air Transport Association by the ATA Air Transport Association is used CCS Commercial Cabin Systems to identify the sub-system responsibilities and ECS Environmental Control System define interfaces between design teams. The EOP Extent Of Protection ATA breakdown is based on a conventional EPS Electric Power System system architecture, which has not changed ETOPS Extended-range Twin-Engine significantly until the need for optimization Operational Performance Standards arrived at the system architecture level. Today, ENG Engine while the technology is available to change the 1 S. Liscouët-Hanke

system architecture in a revolutionary way, the systems are coupled via their energy exchanges cemented ATA breakdown reveals its weakness: (electric, hydraulic, pneumatic, mechanical, the efficient integration of new technologies is thermal, fuel-flow). Additionally, the energy only possible at multi-ATA level; the flow depends on the time axis (different flight conventional interfaces change or disappear. phases), the operation mode (normal, degraded This impacts the system design in the same way or failure mode), the chosen technology (power as the organizational structure: emerging topics type and power consumption characteristics) as like power and heat management are only well as on the safety requirements (installation solvable at multi-ATA and multi-domain of redundant systems etc.). The aircraft power (Systems, Structure, Power-plant) level. A system architecture is linked to the overall functional approach is clearly the most aircraft performance via its contribution to promising solution [1], [2]. Especially with • mass, regards to the comparability of different • drag and solutions, a functional approach will enable to • fuel consumption. open the design space and to start to shift from Additionally, the design and thus the power an evolutionary design to more revolutionary requirements and interaction between systems solutions. depends on aircraft level design parameters, e.g. The variety of possible solutions and number of passenger seats, aircraft geometry, new technologies (e.g. more electrical aircraft mission profile, and on system / technological systems, bleed-less power system architectures) parameters, e.g. the power supply type (electric, and the large number of influencing design hydraulic or pneumatic), pressure level in the parameters require an efficient tool enabling the hydraulic circuits, voltage level etc. system designer or aircraft architect to analyze Regarding the increasing complexity and the impact of system architecture changes at the changing interfaces for e.g. more electric aircraft level (mass, drag, fuel consumption) and architectures, the developed methodology [4] is the impact of aircraft level changes (e.g. based on a formalization of the design process certification, safety or ETOPS requirements) on (Inverse Engineering Principle) of the overall the system design. A model-based, parametric architecture (thus, of the integrated design pre-design process that allows the quantification processes of each system within the architecture and minimization of design margins and the of systems) following a functional approach. better understanding of interdependencies This approach allows technology choices at the between the systems with regards to the power end of the decision chain and enables the architecture would be of great benefit for the highlighting of key parameters that are aircraft manufacturers in its role as architect and interesting for analysis on aircraft level. The system integrator. functional approach also enables the To cope with these challenges is the implementation of a generic process, which then objective of an Airbus internal R&T project allows the comparison of different architectures (part of the Common Virtual Bird initiative [3]), or of aircraft level parameter sets in a consistent started in 2005. A methodology has been way. developed [4] and implemented into a It is proposed to classify the systems of simulation framework prototype [5]. The aircraft power system architecture, as depicted present paper aims to illustrate selected in Fig. 1, according to their major function modeling techniques and to demonstrate the within the power architecture: added value of this approach via two application examples. • Power Generation System, like 2 Methodology & Prototype Implementation engines, auxiliary power unit (APU) for ground operations and a ram air turbine The aircraft power system architecture is a (RAT) for emergency cases. Fuel cells system of interacting sub-systems with a high are candidates for future solutions. level of functional couplings (Fig. 1). The 2 A SIMULATION FRAMEWORK FOR AIRCRAFT POWER SYSTEMS ARCHITECTING

• Power Transformation and can be derived from the simulation of the three- Distribution Systems, which transform dimensional power profiles (Fig. 2). Then, the and distribute the secondary power required versus the available power from the provided by the power generating engine and other power generating systems can systems to the consumer system: be derived. conventionally, the electrical power In a second step, the performance of the system (EPS) and the hydraulic power so defined architecture can be assessed for system transform the mechanical power different off-design mission profiles. of the engine into electric and hydraulic power. Pneumatic power is distributed to the dedicated systems after temperature and pressure control. • Power Consuming Systems, which fulfill the different functions of the aircraft: e.g. the flight control system, wing ice protection system (WIPS),

environmental control system (ECS), fuel system, fuel inerting system, Fig. 2: Three-dimensional definition of the power- exchange interfaces of the power system modules commercial cabin systems (CCS), systems, primary flight Summarizing, the two following steps are control systems, high-lift system, proposed to allow the elaboration of aircraft- equipment cooling systems. level energy balanced system architectures: • the power to be installed, corresponding to the weight, (outcome of the sizing process taking into account various security margins and other requirements) has to be balanced against • the power actually consumed during a flight mission, corresponding to the fuel consumption, (to be assessable in the performance process). Schematically, this calculation principle is depicted in Fig. 4-a.

2.1 Power System Modules In principle, each system of the power systems architecture can be represented in the form of a general power system module (Fig.3). The main interfaces between the system modules are the power interfaces. Each System depends on different types of parameters: • Aircraft Parameters (global parameters)

Fig. 1: Energy Coupling of a Conventional Aircraft • System Design Parameters (local Power Systems Architecture parameters, that can influence The architecture sizing process starts from the neighboring systems) functional requirement of the Power Consuming • Operational Parameters (describe the Systems. After their definition and choice of off-design characteristics of a system). technology, the sizing requirements of the Power Transformation & Distribution Systems

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Aircraft case those “parasitic” couplings of systems, Parameters which lead to snowballed effects on aircraft Energy Energy level, are often neglected as out of responsibility Exchange Power Exchange of one system domain. One example is the System Module Mass increased cooling demand that leads to higher Drag ram air demand, which increases the aircraft System drag and thus the fuel consumption. Or, when Parameters cooling requirements are fulfilled via a liquid Fig. 3: General Power Systems Module cooling system, the weight of this system and the snowballed weight of power supply and thus Some of these parameters can be varied and fuel consumption via increase power off-takes build thus the variable-space for optimization on the engines have to be considered. and trade-offs. The bidirectional energy interfaces 2.2 Implementation schematize the two different calculation modes A prototype of a simulation framework that required for each module: corresponds to the above-described • Sizing Mode: sizing characteristics like requirements has been developed (see Fig. 4) in number of required components, duct a Matlab/Simulink/Stateflow® environment. diameters, generator sizes are outcome Additional code is coupled, depending on the of analysis of computation results for systems module. E.g. C-code is used for the one or more sizing scenarios (ambient Engine-Model, a java application for the aircraft conditions, failures case etc.). performance tool, Dymola/Modelica® import is • Performance Mode: the module achieved for the ECS module. calculates the dedicated output variables The simulation framework contains the (mainly the energy flow) for different parametric power system modules as listed off-design conditions with fixed above embedded in the integrated automated characteristics from the sizing mode for sizing and performance calculation process (Fig. the whole mission profile. 4-a). The process is guided via a graphical user Each system module is developed to ensure the interface (Fig. 4-b). For the coupling to the energy balance principle. This allows aircraft level, the power systems computation is automatic assessments of e.g. the required linked to an Airbus in-house mission and cooling demand or heat rejection to the performance tool (Fig. 4-c). Common environment. This aspect becomes crucial for parameters (Fig. 4-d,e), e.g. environmental more electric or bleed-less architectures. In this (c)

(d) (b) (e)

(a)

Fig. 4: Simulation Framework Overview 4 A SIMULATION FRAMEWORK FOR AIRCRAFT POWER SYSTEMS ARCHITECTING

conditions, mission profile data and aircraft In a last step, the actual technology data, are made available for all systems in order solution is regarded: represented by an to ensure consistency. efficiency ηWIPS,i. The power demand from the dedicated Power Distribution System (Eq.1), 3 Modeling Approaches of Selected Systems here electric or pneumatic, can be computed for In the following section the modeling principles the design point. As well, the systems mass can of selected systems are presented. These be defined. systems are chosen in order to illustrate the 1 different modeling principles used for the WIPS , lPP EOPiperEOP ⋅⋅= (1) η implementation: a functional/deterministic ,iWIPS approach for the WIPS, a statistical approach With: for the CCS, a logic approach for the EPS. The PperEOP [W/m] required power at slat surface per analysis capabilities offered by these modeling m-spanwise extension approaches within the simulation framework are lEOP [m] spanwise length of EOP [-] technology efficiency factor illustrated in the use-case example developed in ηWIPS,I i [-] technology index Section 4. According to the system’s technology, 3.1 Wing Ice Protection System additional design parameters can be chosen: e.g. The function of the WIPS is to protect dedicated for the case of a conventional pneumatic surfaces against ice accretion. In this way, the thermal anti-icing system, the trade can be made maneuverability of the aircraft will be ensured between the temperature, pressure and airflow even during (descent or climb provided to the slat surface. through clouds). To build up the performance mode, the Different candidate technologies exists power demand of the WIPS is computed to fulfill the ice protection function: e.g. depending of the operational parameters: pneumatic thermal anti-icing, electro-thermal • maximum altitude for icing conditions, anti-icing, electro-thermal de-icing, electro- which depends on the ambient mechanic de-icing. Hybrid solutions (e.g. conditions – on hot days, icing can occur combined anti- and de-ice [6]) are possible as up to higher altitude as for International well. Standard Atmosphere (ISA) normal or The first step in the design process is the cold day conditions; definition of the so-called Extent of Protection • WIPS operation definition (e.g. only for (EOP). The EOP depends on one side from the climb, descent and approach but not for wing geometry and on the other side on the take-off or on ground), which depends chosen Ice Protection technology principle on the aircraft-level sizing philosophy. (anti-ice or de-ice, thermal or impulse). As an example, Fig. 6 shows the power demand In a second step, the power required at of two different technologies for different the slat-surface to proceed anti- or de-icing has mission conditions and operation modes. The to be determined, which could be done with e.g. conventional WIPS (pneumatic) enables no load more detailed simulation or test data. The shedding in failure cases, whereas an electro- design point of the WIPS is driven by thermal concept with de-ice function for failure certification requirements: 45 min holding flight case or high altitudes increases the operational under icing conditions [7] which defines the flexibility. However, a combined anti-ice de-ice required power. It depends on the system system has higher weight due to the increased technology if the power demand in off-design EOP. The impact of these choices on the electric conditions can be modulated or not. generator sizing will be discussed in Section 4. Conventional pneumatic ant-ice systems do not modulate the power demand (on/off valve).

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required power WIPS electro-thermal minimum (ISA) operating airline (low cost, luxury liner) as well [norm.] WIPS electro-thermal - degraded (HOT ) as on the final passenger behavior (leisure WIPS electro-thermal - normal (ISA) travel, business travel, etc.). 1,6 WIPS pneumatic (ISA) Taking into account the above-described WIPS pneumatic (HOT) 1,4 characteristics of theses systems, the model is 1,2 based on statistical and technology dependant 1 component data. 0,8 The number of components e.g. reading 0,6 light, IFE units depends on the number of 0,4 passengers, thus on the aircraft size and cabin 0,2 layout. For the cabin lights, the cabin area to be 0 illuminated is the main design driver. Based on groundtaxi a statistical analysis [8] of around 30 aircraft of climb different airlines, the statistical distribution of cruise installed components for the following cabin descent configurations can be defined: mission time [norm.] approach landing • minimum, roll-out • most likely or Fig. 6: WIPS Operational Power requirements for different system configurations. • maximum. This approach is illustrated in Fig. 7 for the example of beverage makers (part of the galley 3.2 Commercial Cabin Systems equipment).

The passenger comfort requirements drive the 1 sizing of the CCS, such as lighting (cabin light, minimum maximum individual reading lights), galleys (e.g. ovens, 0.75 refrigerators, beverage makers) and In-flight most likely Entertainment (IFE). 0.5 As air transport becomes more and more widespread, the requested operational reliability 0.25 of these systems increased and most of the

Freq. of occurrence[norm.] 0 airlines have changed their requirements 0 0.25 0.5 0.75 1 significantly. In terms of power architecture Amount of installed beverage-makers [norm.] design, this means, that CCS, not essential for Fig. 7: Example for statistical distribution of installed the flight mission, impact differently the failure cabin system equipment [8]. cases analysis (refer to Section 4). The number of components corresponding to With regards to the power architecture minimum can be found in equal or less than 10% synthesis the CCS impact the design of the of the examined aircrafts. More equipment than following systems significantly: on the maximum line was found in equal or less • Environmental Control System (ECS): of 10% or the airlines. The most likely-value all heat dissipated in the cabin builds a represents the average number of components heat-load that impacts the air- counted in the aircraft test examples. conditioning and especially the The user can chose between these three temperature control design. values, which are then representing a dedicated • Electric Power System (EPS), which level of airline comfort standard: minimum delivers the required energy to the cabin corresponds to charter airline with e.g. low systems. catering standards, maximum represents long The amount of installed equipment and thus the range or high catering/comfort standards. The amount of power required and the amount of nominal power demand and the weight of the heat dissipated depends significantly on the type installed systems are defined by those choices. of aircraft (long range, short range) and the 6 A SIMULATION FRAMEWORK FOR AIRCRAFT POWER SYSTEMS ARCHITECTING

For the operational power demand, so- systems). The latter is of increasing importance called usage-factors are defined per system and for high-voltage and high-power EPS per flight-phase. For these usage-factors, architectures. defined between 0 and 1, as well, most probable The key driver for the EPS sizing is, statistical data are available. beside the EPS architecture itself, the power Summarizing, the following Eq.(2) required by the electric power consuming defines the power-space of the CCS. systems. For conventional architectures, the

n most important are the CCS, the different technical loads (fuel system pumps, ECS partly, ,CCSelec )( = ∑ ,iusage )( ,inom ⋅⋅ cPtctP , jOM (2) i=1 , more electric actuation systems). For With: bleed-less architectures the WIPS and the ECS n [-] number of different cabin sub-systems are the major consumers. cusage,i [-] usage factor, depending on the flight The key sizing parameters of the EPS phase are the number of generators (GEN) per engine Pnom,i [W] nominal power per sub-system i (ENG), the chosen voltage level for AC and DC cOM,j [-] operation mode factor to be defined for each operation mode j (1=normal, and the location of the power centers. The latter 0=minimal). mainly influences the mass of the feeders, which is linked to the trade-off between voltage The application of this statistical modelling drop, feeder-temperature and diameter. approach within the PWR-Platform is illustrated In order to illustrate the modeling in Section 4.1. principle that allows automated sizing of the generators, a simplified approach of the main 3.3 Electric Power System channel of the EPS is presented (see Fig. 8) in The EPS is a key system in the aircraft power this article. system architecture and becomes more and more important as the amount of required AC consumer DC consumer Distribution electric power increases and especially for more consumer-power centre ηC-P electric architecture concepts. From a functional (C-P) point of view, the electric power system Power centre DC channel (simplified) ηP • generates electric power in different AC channel forms (AC, DC, different voltage levels) Distribution n P t)( from transformation of mechanic power power centre – generation ∑ elec consumer,, i ηP-G (P-G) i=1 (engines, auxiliary power unit (APU), Generation Ram Air Turbine (RAT)), GEN1 … GENn GEN1 … GENn ηGEN • modulates electric power in different forms and voltage levels (AC to DC, DC ENG1 ENG2 to AC), Fig. 8: Schematic of example architecture of the main • stores electric energy in batteries (e.g. channel of the electric power system. for emergency use or APU starting) and • distributes electric power to the The Fig. 8 shows the example of an aircraft with dedicated consumer systems. two engines and two AC-generators per engine. As well, electric power can be provided directly, However, the principle is valid for any number via an external supply (for ground operations) or of generators or engines, when a symmetric via e.g. a system, which generates architecture is chosen. The following electric power. assumption is made: all consumer power is The power interfaces of the EPS-module equally distributed between the available are the electric power demand of the consumer sources of electric power. systems, the electric power sources, mechanic The generator size is then defined in Eg.(3). power demand from the engines, APU or RAT Safety and reliability issues mainly drive the and thermal exchange (heat-load for cooling sizing of the EPS and thus of the generators. 7 S. Liscouët-Hanke

In this way, it is possible to evaluate the level of ⎧ ⎫ m ⎪ 1 ⎛ n ⎞ ⎪ confidence for each system separately. The P = max⎨ ⎜ ∑ P )t( ⎟ ⎬ sized,GEN N ⎜ i,cons,elec ⎟ complete architecture is assessed by comparing = 1j ⎪ j,GEN ⎝ = 1i ⎠ ⎪ ⎩ j⎭ ‘manual’ trade-studies with the simulation (3) results. Concluding, manual calculations at With: system level tend to overestimate and this is Pelec,cons,i(t) [W] electric power of consumer systems then propagated to the aircraft level. Often, at the generator input, function of adding margins compensates lack of consistency time in the design assumptions. The minimisation of NGEN [-] number of generators n [-] number of electric power consuming design margins and the understanding of their systems impact is one of the major benefits of a model- m [-] number of tested operation scenarios based early design trade-off phase. Therefore, Eq.(3) has to be analyzed for the major sizing scenarios: e.g. normal conditions, 4 Application Examples one GEN failure, one ENG failure. For each The developed simulation framework allows a failure scenario j, the possible frequency of broad field of possible application: e.g. the occurrence induces dedicated power comparison of the fuel consumption between requirements of the consumer systems. These two different system architectures taking into dedicated load profiles of the consumer systems account all snowball-effects and propagated (defined in the 3-dimensional power space) are changes (mass, drag, aircraft mission-re- combined with the number of available evaluation), the sizing analysis of aircraft family generators in order to identify the sizing case. concepts, and many more. In the following sections, two examples of closer interest for 3.4 Validation more electric aircraft are focused: the generator The validation of the developed models is a key sizing example and the global thermal analysis. step when aiming acceptance of simulation for decision-making. Here, the validation has two 4.1 Electric Generator Sizing Analysis aims: As an example, the impact on the generator • Confidence-building in the modelling sizing is described in this chapter for two approaches, different power system architectures, • Quantification of the model uncertainty. summarized in Tab. 1. However, the validation of models for future Tab. 1: Power Systems Architectures used for Test Case system solutions cannot be done with traditional methods. As well, the comparison of the System Architecture 1 (A1) Architecture 2 (A2) simulation with existing architectures is not “Conventional” “Bleed-less” obvious, as design paradigms as well as ECS Pneumatic (electric Electric recirculation fans) component technology have changed. WIPS Pneumatic anti-ice Electro-thermal anti- Therefore, a hybrid approach is taken here: and de-ice • System models representing a Actuation Hydraulic Electro/hydraulic “conventional” configuration are EPS 2 generators 4 generators compared with existing systems. The occurring differences are analysed The simulation results are depicted in Fig. 9, taking into account the changing showing the profile of total required power per requirements and hypothesis. GEN for different operation scenarios (normal, one GEN failed, one ENG failed) for the • System models representing new system conventional architecture (a) and the bleed-less configurations are calibrated to current architecture (Fig. 9-b). The load shedding state of the art technology and applied to the CCS and ECS is identical for both knowledge. Evolution potential is architectures, in the case one GEN failed, only outlined.

8 A SIMULATION FRAMEWORK FOR AIRCRAFT POWER SYSTEMS ARCHITECTING

CCS loads are shed. Due to the different system power consumption, here only showed numbers of generators, the GEN failure is the for few selected subsystems of one energy type, sizing condition for the A1. For A2, the ENG and contribution to sizing brings benefit when failure is sizing as only two GEN instead of setting the sizing scenarios and enable the three are left to carry the load. implementation of dedicated power A detailed look on the sizing case for A1 management strategies. is given in Fig. 9-c. As ECS and technical loads are not to be shed in this failure case, the CCS 4.2 Aircraft Thermal Analysis shedding philosophy drives directly the Beside the application illustrated in the previous generator sizing. As the design point is in cruise section, the developed simulation framework is conditions, load shedding in this phase of the also conceived to fit more multidisciplinary mission would be difficult to accept by airlines. analysis. One of these of specific interest for An EPS architecture with four GEN would be more electric aircraft in combination with an option, adding more operational flexibility composite structures is the aircraft global but adding also complexity and probably thermal analysis. In the frame of another Airbus weight. internal research project, a global thermal model For A2, the ENG failure is the sizing of the aircraft has been developed. This model case, as only two of four GEN remain. (see Fig. 10-a) enables the coupling of the Significant is the contribution of the WIPS to systems with the aircraft environment: the the GEN size. It is obvious, that load shedding structure and ambient air. The power system strategies, e.g. the here chosen switch to a de-ice simulation framework presented here is used for mode, downsize the generator. As well, the a test case simulation to compute the systems further load shedding of the CCS only during heat rejection for different flight missions. Via WIPS operation phases would bring benefit the mass and an approximate surface of with only low impact on passenger comfort. components, combined with the power loss Summarizing, this detailed view on the simulation of the systems, the heat rejection of (a) Conventional Architecture, 1 GEN per ENG (b) Bleed-less Architecture, 2 GEN per ENG - Design point search - - Design point search -

(c) (d) Conventional Architecture, 1 GEN per ENG Bleed-less Architecture, 2 GEN per ENG - Generator Failure: Details - - Engine Failure: Details -

Fig. 9: Engine Generator Sizing Analysis 9 S. Liscouët-Hanke

the systems can be computed. This enables the principle. The coupling of probabilistic methods calculation of structure, fuel and air for uncertainty management and sensitivity temperatures (see Fig. 10-b and Fig. 10-c) for analysis is in progress. Preliminary test studies two different mission points, which will be in showed already the further benefit brought to return an input to the systems sizing calculation. the model development and the end user of this The coupling of the structural and the systems tool. module will allow the identification of thermal critical points already earlier in the design 6 Acknowledgements process, the development of adapted thermal The author thanks Prof. J-C. Maré and S. Pufe management strategies and the establishing of a for their support and advice, and the various topological view on in their specialists in Airbus and EADS Innovation environment. Works. The author is very grateful to Zonta (a) International that rewarded this research work Systems within structural with the Amelia Earhart Fellowship Award in model 2007 and 2008.

References [1] Faleiro L. Beyond the More Electric Aircraft. America, American Institute of Aeronautics and Astronautics, September 2005. [2] Pahl G and Beitz W. Engineering Design – A systematic approach. Springer, 1996. (b) (c) Ground Case Take-off [3] Bachelier P. Virtual Product Development: a strategic axis at Airbus. International Conference on Trends in Product Life Cycle, Modeling, Simulation and Synthesis PLMSS, Bangalore, India, 2006 [4] Hanke S, Pufe S and Maré J-C. A methodology for aircraft power system architectures analysis International Workshop on Aircraft System Fig. 10: Application example - analysis of systems heat Technologies, p59-68, Hamburg, Germany, March rejection on the structure temperature. 2007. [5] Hanke S, Pufe S and Maré J-C. A Simulation 5 Conclusions and Outlook Framework for Aircraft Power Management European Conference for Aerospace Sciences, The presented simulation framework Bruxelles, Belgium, July 2007. implements a methodology, which is a great [6] Sinnett M. 787 No-Bleed Systems: Saving Fuel and asset for architecting (definition, pre-sizing and Enhancing Operational Efficiencies. Aero Magazine. The Company, Issue 28 Q4, 2007. optimizing) aircraft power systems. It allows [7] Federal Aviation Administration (FAA). FAR25, parametric studies during the preliminary design Appendix C, 2008. phases of modern aircraft. Different modeling [8] Christoffers J. Common Virtual Bird–Power methodologies (deterministic, statistics, logic) Management: Development of the Cabin Systems are combined in a common platform, respecting Module. Technical Report. EADS IW, 2007. the specific characteristics of each system while helping the architect to integrate them into an Copyright Statement energy-balanced design. The common The authors confirm that they, and/or their company or simulation framework enables the challenging institution, hold copyright on all of the original material of design margins, to analysis of more different included in their paper. They also confirm they have architecture configurations in less time and the obtained permission, from the copyright holder of any third party material included in their paper, to publish it as elaboration of more integrated architectures. part of their paper. The authors grant full permission for Efforts are still necessary to model all the publication and distribution of their paper as part of systems of interest following the presented the ICAS2008 proceedings or as individual off-prints from the proceedings. 10