proceedings of the institution of mechanical engineers, part g_ journal of aerospace...

Original Article

A methodology for vehicle and missionlevel comparison of More Electric Aircraftsubsystem solutions: Application to theflight control actuation system

Imon Chakraborty1, Dimitri N Mavris1, Mathias Emeneth2 andAlexander Schneegans2

Abstract

As part of the More Electric Initiative, there is a significant interest in designing energy-optimized More Electric Aircraft,

where electric power meets all non-propulsive power requirements. To achieve this goal, the aircraft subsystems must be

analyzed much earlier than in the traditional design process. This means that the designer must be able to compare

competing subsystem solutions with only limited knowledge regarding aircraft geometry and other design characteristics.

The methodology presented in this work allows such tradeoffs to be performed and is driven by subsystem requirements

definition, component modeling and simulation, identification of critical or constraining flight conditions, and evaluation

of competing architectures at the vehicle and mission level. The methodology is applied to the flight control actuation

system, where electric control surface actuators are likely to replace conventional centralized hydraulics in future More

Electric Aircraft. While the potential benefits of electric actuation are generally accepted, there is considerable debate

regarding the most suitable electric actuator electrohydrostatic or electromechanical. These two actuator types form

the basis of the competing solutions analyzed in this work, which focuses on a small narrowbody aircraft such as the

Boeing 737-800. The competing architectures are compared at both the vehicle and mission levels, using as metrics

subsystem weight and mission fuel burn, respectively. As shown in this work, the use of this methodology aids the

decision-making process by allowing the designer to rapidly evaluate the significance of any performance advantage

between the competing solutions.

Keywords

More Electric Aircraft, More Electric Initiative, electric actuator, electrohydrostatic actuator, electromechanical actuator

Date received: 28 February 2014; accepted: 13 June 2014

Introduction: Consideration of MoreElectric subsystems in conceptual design

The conceptual design phase of the traditional aircraftdesign process typically gives only limited consider-ation to the aircraft subsystems (In this work, bysubsystems the authors mean Aircraft EquipmentSystems such as actuation systems for flight controlsurfaces, landing gear, thrust reversers, nosewheelsteering, and wheel braking, the environmental con-trol system (ECS) and ice protection system (IPS),engine starting, etc.), which are addressed mainly inthe late preliminary and detailed design phases.1 Thisapproach has proven adequate when applied to air-craft with conventional subsystem solutions due to thepresence of a continuously updated historical data-base and a relative lack of coupling/interactionsbetween different subsystems.

This will not be the case, however, for conceptualdesign of More Electric Aircraft (MEA) or, in thelimit, All Electric Aircraft (AEA), for which there isno equivalent historical data, and where subsysteminteractions may be significant. As indicated byFaleiro, MEA/AEA design will necessitate a newdesign paradigm that considers the aircraft and itssubsystems as a system of systems.2 More thoroughconsideration must therefore be given to the

Proc IMechE Part G:

J Aerospace Engineering

2015, Vol. 229(6) 10881102

! IMechE 2014

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DOI: 10.1177/0954410014544303

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1Aerospace Systems Design Laboratory, School of Aerospace

Engineering, Georgia Institute of Technology, Atlanta, GA, USA2PACE America Inc., Seattle, WA, USA

Corresponding author:

Imon Chakraborty, Aerospace Systems Design Laboratory, School of

Aerospace Engineering, Georgia Institute of Technology, Atlanta,

GA 30332, USA.

Email: [email protected]

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requirements, interactions, and effects of the aircraftsubsystems even in the conceptual design phase.

This paper presents a requirements-driven method-ology aimed at MEA/AEA design that integrates theanalysis of subsystem requirements and effects with theconceptual phase aircraft sizing process (Figure 1(a)).For each subsystem, the functional requirements areset by identifying constraining operating scenarios.The subsystem is sized to those requirements usingconceptual phase aircraft parameters from the aircraftsizing analysis (dimensions, estimated takeoff grossweights, etc.). Fallouts such as subsystem weight(Wsub), non-propulsive power requirement (Pnp), anddrag coefficient change (CD, if any) can be fed back tothe aircraft and propulsion system sizing analyses forthe subsequent iteration (if Cycle? Yes). The con-vergence of this iterative process yields a design whereconceptual design parameters such as aircraft weight,wing area, propulsion requirements, etc. were com-puted by representing subsystem requirements andeffects explicitly.

This paper focuses on the application of this meth-odology to electric flight control actuation systems. Asseen in Figure 1(b), the steps involved mirror the gen-eral steps of Figure 1(a) and form the subject of subse-quent sections of the paper. An overview of flightcontrol actuation systems is described in the next sec-tion. The Control surface actuation requirementssection derives the actuation requirements of the con-trol surfaces, which feed into the actuator models that

are described in the EHA and EMAmodels section.The Electric flight control actuation system architec-tures section identifies two configurations of interestand develops their corresponding architectures. TheVehicle level analysis section shows a comparisonof the two configurations at the vehicle level, basedon subsystem weight. The Mission level analysis sec-tion compares the two configurations at the missionlevel, based on fuel burn. Finally, conclusions drawnfrom this work and avenues identified by the authorsfor future work are presented.

Overview of flight control actuationsystems: Hydraulic versus electric

Most commercial and military aircraft in servicetoday use centralized hydraulic systems for actuatingflight control surfaces, landing gear, brakes, thrustreversers, etc. These systems, utilizing electric orengine-driven pumps to pressurize hydraulic fluid to2035MPa (30005000 psi), have been maturedthrough decades of aeronautical experience, withredundancy achieved through the incorporation ofmultiple physically segregated independent systems.

However, the inevitable technology saturationreached by hydraulics coupled with significantimprovements in the power densities of power elec-tronics and electric drives3,4 have led to a renewedinterest in the research and development of electric

Figure 1. Integrating subsystem sizing into the traditional aircraft sizing activity in the conceptual design phase. (a) Methodology

linking subsystem sizing with traditional conceptual design phase aircraft sizing, with feed-back information flow to enable design

cycling in response to subsystem changes and (b) Application to flight control actuation system with electrically actuated flight

control surfaces.

EHA: electrohydrostatic actuator; EMA: electromechanical actuator.

Chakraborty et al. 1089


actuators for flight control actuation as part of theMore Electric Initiative. Compared to centralizedhydraulics, electric actuators have the advantage ofPower on Demand, consuming power only whenthe control surface moves or overcomes a load.Moreover, they require only electrical energy input(Power By Wire) and allow the removal of bulkyhydraulic lines, offering potential aircraft levelweight savings. An additional advantage stems fromthe fact that electric actuators are Line ReplaceableUnits, and are much easier to remove/install thanhydraulic actuators.

The majority of research and development activ-ities related to electric actuation has been on the mili-tary side, starting as far back as the 1950s with theBritish V-bombers.4 More recent efforts have beenundertaken in the United States,5 involving theC-141 Starlifter,6 NASAs F-18 Systems ResearchAircraft,7,8 the AFTI F-16,9 and most recently theF-35 Joint Strike Fighter.10 On the civil aviationside, a notable MEA is the Airbus A380, where elec-tric actuators are used for certain control surfaces.11

This 2H/2E configuration has been claimed to offera 450 kg (1000 lb) weight savings compared to a con-ventional 3H configuration.12 With the exception ofthe Joint Strike Fighter, other flight test programs andanalyses have focused on only single control surfacetypes at a time, e.g., flaperon,7,8 aileron,6 spoiler,13,14

or high lift devices.15 In this work, since the entireflight control actuation system must be represented,all the aircraft flight control surfaces were consideredsimultaneously.

Electric actuator concepts fall mainly into twocategories, within which there may be architecturalvariationsthe electrohydrostatic actuator (EHA)and the electromechanical actuator (EMA). Eachhas advantages and disadvantages relative to theother, and there is significant debate (and no consen-sus) within the community regarding the type moreapplicable to flight control surface actuation. TheEHA and EMA form the competing technologies ana-lyzed in this work.

The competing subsystem solutions or architectureswere formed by associating the competing technolo-gies (EHA or EMA) to the aircraft control surfaces, amultitude of which are present even on small narrow-body aircraft. These associations were developedtaking into consideration flight-criticality of the con-trol surfaces, control surface redundancy, etc.

The current analysis utilized MATLAB/Simulinkand Pacelab SysArc, an interactive system architec-ture investigation platform developed by PACEAerospace Engineering and Information TechnologyGmbH (Organization website: www.pace.de; PacelabSysArc: www.pace.de/products/preliminary-design/pacelab-sysarc). It comprises a Knowledge Designerwhere system components (control surfaces, actu-ators, etc.) were mathematically modeled, and anEngineering Workbench, using which the

configuration architectures were developed. For add-itional details regarding the use of Pacelab SysArc foraircraft architectural analysis, the interested reader isreferred to Chakraborty et al.16 The proposed meth-odology can of course be applied using other analysistools as well.

Control surface actuation requirements

Even for aircraft of similar size, the precise control sur-face configuration is often influenced by organizationaldesign philosophy, especially for high-lift devices.17 Todemonstrate thismethodology, the control surface con-figuration and dimensions of the Boeing 737-800 air-craft were utilized. Each wing contains an outboardaileron and four flight spoilers for roll control. Aft ele-vators and a trimmable horizontal stabilizer (THS)provide pitch control, and a single rudder controlsyaw. The high lift system comprises two inboardKrueger flaps, four outboard leading edge slats, andtwo double-slotted trailing edge flaps (TEFs) on eachwing. The speedbrake function is performed by thesymmetric deployment of all flight spoilers in the air,and is supplemented on the ground by the deploymentof two ground spoiler panels on each wing.

For hinged control surfaces, the required actuationload can be characterized by the maximum hingemoment Mh,max. The hinge moments developed forthe ailerons, elevator, and rudder may be expressedin standard form by

Mh q Sf cf Ch,0 Ch,M eff Ch,M,

1

where q is the dynamic pressure, M is the Machnumber, eff is the effective angle of attack seen bythe main surface, the control surface deflection, andSf and cf are the control surface planform area andchord. Ch,0, Ch,, and Ch, are hinge moment coeffi-cients representing respectively the effect of main lift-ing surface camber, eff, and on the hinge moment.

The hinge moment coefficients were determinedbased on the geometries of the control surfaces andassociated main lifting surfaces using the empiricalmethodology of Roskam.18 Hinge moments predictedusing these coefficients were validated against pub-lished hinge moments from flight tests of the NASAF-18 Systems Research Aircraft.7 This comparison,shown in Figure 2, yielded acceptable agreementbetween predicted and observed hinge moments. Forthe flight and ground spoilers, the method proposedby Scholz19 was used to determine hinge momentcoefficients.

For each primary control surface, the flight condi-tion (represented by q, eff, and ) that yields the max-imum hinge moment can be obtained from therelevant Federation Aviation Regulations (FARs),and have been identified by Scholz.19 For the aileron,it is given by FAR 25.349, and corresponds to the

1090 Proc IMechE Part G: J Aerospace Engineering 229(6)


maximum hinge moment from among three flightconditions abrupt and full deflection at design man-euver speed (VA), a deflection yielding the same rollrate at design cruise speed (VC), and a deflection yield-ing a third of that roll rate at the design dive speed(VD). For the elevators, FAR 25.255 specifies therequired positive and negative recovery load factorsthat must be available following a runaway failureof the THS. For the rudder, the applicable FARs areFAR 25.149 (one engine inoperative, OEI) and FAR25.351 (permissible rudder deflections). FAR 25.351and FAR 25.255 were evaluated over the aircraft alti-tude-speed envelope to determine the most criticalflight condition yielding the maximum hingemoment (Figure 3). For the flight spoilers, the max-imum hinge moment occurs when they are extendedduring an emergency descent performed at the designdive speed (VD). Since weight-on-wheels is requiredfor ground spoiler deployment, the maximum ratedtire speed was used to compute their actuation loadrequirement.

High lift device actuating loads strongly depend onflap mechanism kinematics, which vary significantlybetween both aircraft and manufacturers.17 In lieu of adetailed analysis of any one mechanism, the maximumactuation loads for the Krueger flaps and slats werederived using force and moment coefficients from pub-lished wind tunnel analyses.20,21 For the TEFs, therequired actuation load was set to the ratings of theexisting leadscrew typeactuators (Boeing 737 linear ball-screw flap actuator, Triumph Gear Systems, TriumphGroup, Inc.,High-LiftActuationSystemsproduct data-sheet. Available online: www.triumphgroup.com/companies/triumph-gear-systems-park-city/Documents, accessed 25 February 2014.).

The maximum angular rates of ailerons, elevators,rudder, and flight spoilers were set based on theobserved trend of higher rate requirements formodern automatic flight control systems.22

Representative rates for the high lift devices wereobtained by timing videos of these devices deploying.Table 1 summarizes the derived actuationrequirements.

EHA and EMA models

Linear type actuators (as opposed to rotary actuators)were considered in this work. The relationship

(a) (b)

Figure 2. Comparison of hinge moments predicted using Roskams method with flight test data from NASAs F-18 SRA. (a) Abrupt

roll doublet - full stick (left): M 0.72, h 7,620 m (25,000 ft), q-bar 13.60 kPa (284 lbf/ft2) and (b) Abrupt aileron reversal - full stick(left): M 0.84, h 7,620 m (25,000 ft), q-bar 18.29 kPa (382 lbf/ft2).

Figure 3. Aircraft flight envelope with maximum hinge

moment flight condition for ailerons, elevators, rudder, and

flight spoilers.



between control surface angular movement and actu-ator ram linear movement for hinged control surfacesis governed by the kinematics of the control surface toactuator linkage (Figure 4(a)). This allows the actu-ators output force FL, ram speed v, and ram positionx to be computed from the control surfaces hingemoment, angular rate, and angular position.

The actuator is sized based on the required load-speed envelope, shown notionally in Figure 4(b). Itsstall load F0, no-load (maximum) speed vmax, andstroke xmax can be computed from the maximumcontrol surface hinge moment Mh,max, maximumangular rate _max, control surface angular rangemax, and the linkage kinematics (Figure 4(a)).

Electrohydrostatic actuator

In the EHA, a bidirectional motor is directly mated toa bidirectional, fixed-displacement pump, which actu-ates a hydraulic cylinder whose piston is connected tothe actuator output ram (Figure 5(a)). The cylinderpressure difference p required to achieve an outputforce FL, and a piston velocity _x with acceleration x isgiven by

p 1Apmr x c _x FL 2

where Ap is the piston cross-sectional area, mr themass of reciprocating components (piston outputrodbalance rod), and c the viscous friction

(a) (b)

Figure 4. Control surface linkage kinematics and actuator load-speed envelope. (a) Actuation of hinged control surface using linear

actuator and (b) actuator load versus speed envelope.

(a)(b)

Figure 5. Notional schematic of electrohydrostatic and electromechanical actuators. (a) Electrohydrostatic actuator (EHA) and

(b) electromechanical actuator (EMA).

Table 1. Summary of flight control surface actuation

requirements for example small narrowbody aircraft.

Control surface Actuating load Rate

No.

per aircraft

Ailerons 4,200 Nm 60/s 2

Elevators 7,600 Nm 60/s 2

Rudder 8,200 Nm 60/s 1

Flight spoilers 4,200 Nm 60/s 8

Ground spoilers 3,800 Nm 40/s 4

Trailing edge flaps 51,000 N 102 mm/s 4

Leading edge slats 6,300 N 60 mm/s 8

Krueger flaps 5,600 Nm 16/s 4



coefficient. The required piston area was computed byevaluating equation (2) at the stall load condition andwith maximum pressure difference, i.e., withFL F0, _x x 0, p pmax. Cylinder compo-nents were sized considering the loads applicable foreach, e.g., piston rod diameter based on tension/com-pression and buckling, cylinder wall thickness basedon tensile and circumferential (Hoop) stresses, etc.23

The cylinder mass mc was computed by summing thecomputed masses of the constituent components andthe enclosed hydraulic fluid, and considering thecylinder configuration (single or tandem).

The required total pump flowQtot is given by the sumof the ideal flow Qideal and the leakage flow Qlkg, i.e.

Qtot Qideal Qlkg Ap _x clkgp 3

where clkg is the leakage coefficient. The pump shaftspeed !p and required shaft torque p are given by

!p QtotDpump

, p p Dpump

h4

where Dpump is related to the pump displacement(Pump displacement is defined as the fluid volumepumped per unit revolution (mm3/rev or in3/rev)) asDpump pump displacement=2 and h is thepump efficiency. Since the leakage flow is typicallymuch smaller than the ideal flow, the maximum flowrate corresponds to the maximum ram velocity, i.e.,Qmaxtot Apvmax. For the EHA, the motor and pumpshaft speeds and torques are physically identical, i.e.,!m !p, m p. The required maximum speed andshaft torque of the motor are then given by

!maxm ApvmaxDpump

, maxm pmax Dpump

h5

A parametric pump model was used to computethe pump mass mp based on the pump displacement.This included the mass of the hydraulic fluid presentwithin the pump, while the mass of the hydraulic fluidin the lines connecting the pump to the cylinder wasaccounted for using the mass calibration factor dis-cussed in a subsequent section.

Electromechanical actuator

In the EMA, a bidirectional motor drives a reductiongearbox (GB) which in turn drives a ballscrew (BS)that converts the rotational motion of the GB outputshaft to the linear motion of the actuator output ram(Figure 5(b)). The gearing ratios of the BS and the GBwere defined as

Gbs !bsv

, Ggb !m!bs

6

where !bs is the angular speed of the rotating memberof the BS. The gearing ratios can be used to relate

actuator no-load speed and stall load to maximummotor shaft speed and shaft torque:

!maxm vmax Gbs Ggb, maxm F0

bs Gbsgb Ggb7

where bs and gb are the BS and GB efficiencies. Forthis work, a GB efficiency of gb 0:9 was used. TheBS efficiency was computed as bs tanbs=tanbs bs, where bs tan1 bs is the frictionangle and bs tan11=Gbsrbs is the screw helixangle. The radius of the screw rbs was selected as thelarger of the radius computed from buckling consid-eration and that from consideration of combined axialand torsional loading.23

For the GB, the number of stages required toattain the overall GB ratio Gbs was computed asn ceil logGgb=log Gmaxstg

, where Gmaxstg 2Lcc=

dmin 1 is the maximum stage reduction ratio for agiven center-to-center distance Lcc and minimum per-missible pinion diameter dmin. The GB mass mgb wascomputed after determining the gear face width fromstrength considerations,23 and accounting for thenumber of driver-driven pairs.

Motor and power electronics

A brushless DC (BLDC) motor was considered forboth the EHA and EMA cases, as this type ofmotor has been successfully flight tested in electricactuators.8 Using parametric BLDC torque-speedcharacteristics,24 the required motor output torque(maxout ) was determined based on the following criteria:

1. The peak output torque given by equations (5)and (7) must be met with a torque margin appliedto satisfy acceleration requirements against actu-ator inertia.

2. The actuator force-speed (F v) envelope, whenreduced to a torque-speed ( !) envelope by theactuator gearing, must be contained within themotor torque-speed envelope.

3. Themotorsmaximumcontinuous torquemustmeetthe actuators steady-holding load requirement.

Evaluating the input electrical power over theboundary of the motor ! envelope (accountingfor losses whose magnitudes depend on the operatingpoint in the envelope), the maximum electrical inputpower Pmaxin was determined. Torabzadeh-Tari

25 pro-vides an empirical estimate of the mass of a BLDCmotor based on maxout and P

maxin as

mem 1034:7

maxout

0:841222 26:7 f1 3750:28

0:98 Pmaxin 4910:11kg

8



where f1 is the source fundamental frequency. Thesame author also provides an estimate for theweight of the power electronics as

mpe 103400

VACin

0:022290 0:3 fsw 60000:8

0:3 Pin 20001:1kg

9

where fsw is the switching frequency, VACin is the input

AC voltage (for this work, the actuators are poweredfrom 115 volts alternating current (VAC) buses), andPinis obtained by scaling the motors maximum inputpower Pmaxin by the assumed efficiency of the power elec-tronics. For actuator designs comprising two motors,each was assumed to have its own power electronic unit.

Actuator weight computation

Since the component mass estimation described aboveaccounts for only the major actuator components, butneglects other components such as fittings, clamps,etc., simply adding these masses would likely under-predict total actuator mass. To compensate, an addi-tive (likely under-predicted) estimate of actuator masswas first obtained by adding the computed masses of

the main components while factoring in the number ofcomponent units N: present in one actuator, i.e.

mEHA mc Np:mp Nem:mem Npe:mpemEMA mbs Ngb:mgb Nem:mem Npe:mpe

10

To determine the magnitude of the under-predic-tion, the masses predicted using equation (10) werecompared against published masses of electric actu-ators available from previous analyses or flight tests.For each such case, the actuator was sized to the samestall load, no-load rate, and stroke as the publishedactuators. Additional information regarding the actu-ator layout (such as electrohydraulic (EH) or electro-mechanical redundancy) was incorporated wheneveravailable (Abbreviations: 3 three phase, PMM permanent magnet machine, FDP fixed displace-ment pump, EH electrohydraulic, TC tandemcylinder, BLDCM brushless DC motor, BS ball-screw, SRM switched reluctance machine, M/P motor/pump, M/GB motor/gearbox.). A calibrationfactor of 1.15 (i.e., 15%) applied to the predictedcylinder, pump, BS, and GB masses produced reason-able agreement, as shown in Tables 2 and 3 for EHAand EMA, respectively.

Table 2. Comparison of predicted versus published EHA weight following calibration.

Description F-18 SRA flaperon Typical tandem EHA

Source Navarro7 Sadeghi and Lyons26

Stall load 59.16 kN (13,300 lbf) 142.34 kN (32,000 lbf)

No load rate 195.6 mm/s (7.7 in/s) 203.2 mm/s (8.0 in/s)

Stroke 114.3 mm (4.5 in)

Layout 1 3 PMM! 1 FDP Dual EH channel, 2 TC/surfacePublished weight 18.8 kg (41.5 lb) 72.3 kg (159.5 lb)

Predicted weight 19.1 kg (42 lb) 74.4 kg (164 lb)

EH: electrohydraulic; PMM: permanent magnet machine; FDP: fixed displacement pump; TC: tandem cylinder;

3: three-phase.

Table 3. Comparison of predicted versus published EMA weight following calibration.

Description F-18 SRA flaperon Transport spoiler C-141 aileron

Source Jensen et al.8 Fronista and Bradbury14 Thompson6

Stall load 58.72 kN (13,200 lbf) 222.4 kN (50,000 lbf) 84.74 kN (19,050 lbf)

No load rate 170.2 mm/s (6.7 in/s) 177.8 mm/s (7.0 in/s) 118.1 mm/s (4.65 in/s)

Stroke 104.8 mm (4.125 in) 152.4 mm (6.0 in) 137.9 mm (5.43 in)

Layout 2 3BLDCM! 1 BS 1 5 SRM! 1 BS 2 M/GB! 1 BSPublished weight 11.8 kg (26 lb) 17.7 kg (39 lb) 15.9 kg (35 lb)

Predicted weight 11.8 kg (26 lb) 18.1 kg (40 lb) 16.8 kg (37 lb)

BLDCM: brushless DC motor; SRM: switched reluctance machine; M/GB: motor/gearbox; BS: ballscrew; 3: three-phase; 5: five-phase.



Electric flight control actuation systemarchitectures

The EHA and the EMA each have certain advantagesand disadvantages relative to the other, which is thecause of the EHA versus EMA debate within theflight control actuation community.

If designed to the same stall load, no-load rate, andstroke, the EMA will generally be lighter than theEHA. This is seen in Figure 6, which comparesweight-optimal EHA and EMA designs for a rangeof hinge moments. The weight advantage of the EMAis seen to grow with the magnitude of the sizing hingemoment. The EMA is also potentially more energy-efficient as it converts electrical power directly tomechanical power without an intermediate step ofhydraulic/fluidic power.

However, with regard to flight control surface actu-ation, a critical hazard for the EMA is the possibility ofcertain single-point jam failures.27 Since a mechanicaljam is involved, it is not possible to use multiple EMAs

per control surface to provide redundancy either. Thishas caused the EMAs suitability for actuating at leastthe flight-critical control surfaces to be questioned bysome, since a catastrophic outcome may result fromthese control surfaces jamming at an adverse position(e.g., fully deflected or hardover position).

Considering this same criterion, the EHA enjoys amajor advantage in that it can be designed to be fail-safe. If a failure occurs, the hydraulic cylinder can bedesigned to enter a damped/bypass mode, preventingthe actuators output ram from jamming but simul-taneously providing damping against flutter. Thismeans that EHAs can be used in parallel to meetthe failure probability criteria applicable to flight-cri-tical control surfaces. They can also be integrated withcentralized hydraulics (e.g., Electrical BackupHydraulic Actuator on Airbus A38011).

The EMA may, however, be suitable for the actu-ation of flight and ground spoilers. The ground spoi-lers are disabled in flight and are not flight-critical(i.e., their operability is not necessary for the contin-ued safe flight of the aircraft). Further, there may becases where the layout and sizes of wing control sur-faces are such that Failure Mode and Effect Analysis(FMEA) can be used to establish the fact that theflight spoilers are also not flight-critical. In suchcases, less stringent failure probability requirementswould be applicable to these control surfaces, andthey may presumably be EMA-driven. In fact, EMAdesign for spoiler actuation was analyzed by Atallahet al.13 and Fronista and Bradbury.14

A similar argument holds for the high lift devices,and an EMA-driven high lift system was analyzedexperimentally and from a safety/reliability stand-point by Bennett et al.15,28 In line with this reasoning,the two control surface actuation configurationsshown in Figure 7 were selected for further analysis.

. Configuration 1: Other than the TEFs, all controlsurfaces are driven by EHAs. Two parallel EHAs

Figure 7. Configuration 1: TE flaps actuated by EMA-LS. All other surfaces actuated by EHA. Configuration 2: Ailerons, elevators,

rudder actuated by EHA. LE slats, Krueger flaps, and spoilers actuated by EMA. TE flaps actuated by EMA-LS.

EHA: electrohydrostatic actuator; EMA: electromechanical actuator; TEF: trailing edge flap.

Figure 6. Weight-optimal EHA and EMA designs versus sizing

hinge moment.

EHA: electrohydrostatic actuator; EMA: electromechanical

actuator.



are used to drive each aileron and elevator, whilethree are used for the rudder. Slats, Krueger flaps,flight spoilers, and ground spoilers are driven by asingle EHA. Each TEF is driven by two parallelElectromechanically Actuated Leadscrews (EMA-LS: Electromechanically Actuated Leadscrew. Avariant of the EMA in which the nut is the trans-lating member instead of the rotating member.This is similar to the leadscrews used in the con-ventional Boeing 737-800 TEF system.).

. Configuration 2: The slats, Krueger flaps, flightspoilers, and ground spoilers are each actuatedusing a single EMA. Actuation for the remainingcontrol surfaces is identical to Configuration 1.

A detailed functional hazard analysis was beyondthe scope of the current work, but design practicesfrom Boeing29 and Airbus30 were adopted to deter-mine the nature of redundancy incorporated into thetwo architectures.

1. The stringent reliability requirements of the flight-critical primary control surfaces was met using acombination of actuator redundancy and internalredundancy.31 Actuator redundancy was providedby using two parallel actuators in activeactiveconfiguration for the elevators and ailerons, andthree parallel actuators in activeactive-standbyconfiguration for the rudder. Only EHAs wereconsidered for these surfaces due to their fail-safe characteristics, and parallel EHAs werepowered from different electric buses. Each EHAdriving a flight-critical control surface wasassumed to have internal redundancy, in whichtwo motor-pump pairs drove a tandem cylinder(2M/P ! 1 TC).

2. For non-flight-critical control surfaces such asspoilers, slats, and Krueger flaps, surface redun-dancy already existed since there were multiplepanels per wing. In this case, each panel was pow-ered by a single actuator (either EHA or EMA).The actuators were connected to the electric busesin a way that ensured symmetric availabilityof control surfaces following the failure of anyone bus.

3. TEF panels were provided with actuator redun-dancy, and were driven by two EMA-LS units atthe two extremities (in practice, this also preventspanel warping due to aerodynamic loads), pow-ered from different buses. Each EMA-LS hadinternal redundancy,6 in which two motor-gearboxpairs drove a BS through a summing element. Thisarrangement was selected as being representativeof the Distributed Flap Drive Actuation phil-osophy that is currently being investigated fortrailing edge high-lift systems in the interest ofweight savings and functional flexibility.32 Insuch an arrangement, flap synchronization

between left and right wings will be ensuredthrough electronic sensing and signalling, in lieuof the mechanically ensured (or sometimeshydraulically ensured) synchronization found oncommercial aircraft currently in service.

Vehicle level analysis

For the two configurations shown in Figure 7, a gra-dient-based constrained optimization was performedto obtain the weight-optimal EHA and EMA designsfor the relevant control surfaces. For the EHA, max-imum pump pressure and pump displacement wereused as design variables, while for the EMA, the BSand GB gearing ratios were used. For the EHA, thereis an iterative loop between available pressure differ-ence, cylinder sizing (and mass), and required pressuredifference. A constraint was therefore enforced toensure that the required pressure difference did notexceed the pumps maximum pressure difference.For the EMA, the BS RPM was limited to a max-imum permissible value.14 For both EHA andEMA, maximum limits were enforced for motorRPM and permissible current under stall load.

Table 4 shows the weight breakdown of the twoarchitectures, which includes the weight of allweight-optimal actuators and also the associatedwiring from the main electrical buses to the actuators.The wiring weight was computed by physically pos-itioning the actuators in Pacelab SysArc, and theninvoking the tools routing algorithm. This algorithmconverts logical connections between components(e.g., between a bus and an actuator) to physicalequivalents, using connecting elements (electricwiring). Wire gauge is determined from voltage dropand current considerations, and length is determinedbased on user-defined allowable pathways, throughwhich electrical wiring is constrained to run.

Configuration 2 enjoys a weight advantage of5.87% over Configuration 1, but it should be notedthat this weight advantage would be only 0.11% ofthe maximum takeoff weight of the Boeing 737-800(79,000 kg).

Mission level analysis

Both configurations were evaluated over the course ofthe representative mission shown in Figure 8. Primarycontrol surface movements depend on the phase offlight, and during the mission simulation, these werecharacterized using the normalized magnitudes ofdeflection in one or both directions and the cyclesper second parameter which represents the frequencyof movement.33 The full control surface deflectionsshown in Table 5 for the Ground segment representthe control surface check that is performed after start-ing the engines. The deflections and rates shown forthe Cruise segment are representative of an



assumed encounter with moderate turbulence thatlasted for a fraction of the cruise segment, andrequired more aggressive control surface movements.Primary control surface movements for the remainderof the cruise segment were assumed to be negligible.The high lift devices were assumed to deploy or retractin a manner consistent with what is observed in atypical commercial aircraft flight.

To compute the power consumed at the bus by theelectric actuator, the control surface hinge momentand angular rate were converted to the requiredforce (Freq) and speed (vreq) at the control/actuatorlinkage using the control surface gearing Gcs. Thesewere then propagated through the mathematicalmodel of the actuator to get the required torque(req) and shaft speed (!m,req) of the electric motor,effectively using the overall actuator gearing (Gact).The motor torque and shaft speed requirementswere then evaluated in conjunction with the assumedtorque-speed characteristics of the BLDC motor24 tofind the motor input power requirement. Finally, thiswas scaled using the assumed power electronics

efficiency pe to give the power draw at the bus (Analternate method is to estimate the power requirementat the control surface, i.e., the product of hingemoment and angular rate, and then scale successivelyby the efficiencies of all actuator components encoun-tered to obtain the power draw at the bus. However,this method will predict zero power draw from the busif the control surface is held fixed against a non-zeroload, while in reality electric actuators are known toconsume power in this condition. The methoddescribed above and used in this analysis does notsuffer from this discrepancy.). This is shown in equa-tion (11) for a single-channel actuator

Mh,! !Gcs

Freq, vreq !Gact

req,

!m,req !em,!

Pemin !pe

Ppein Pbus

11

Certain control surface motions (e.g., flap deploy-ment/retraction) occur only during a fraction of theduration of a mission segment. For these, the averagepower consumption over the duration of the missionsegment Pseg was computed using the mean powerconsumed during the motion P and the activity ratio as follows

tactivetseg

2 0, 1, ) Pseg P 0:1 P

12

As only the flight control actuation subsystem wasbeing considered, the non-propulsive power require-ments of the other subsystems (Table 6) were assumedto be identical for both configurations. The ECS wasassumed to have a bleedless architecture, and enginebleed air was used only for the cowl IPS, similar tothe Boeing 787 arrangement. The power and bleedair demands of these systems were reported by

Table 4. Weight comparison between Configuration 1 and Configuration 2.

Control surface Actuators Configuration 1 weight (kg) Configuration 2 weight (kg)

Name No. Act./surf. Design Unit Group Design Unit Group

Aileron 2 2 EHA 15.8 63.2 EHA 15.8 63.2

Elevator 2 2 EHA 25.6 102.4 EHA 25.6 102.4

Rudder 1 3 EHA 27.0 81.0 EHA 27.0 81.0

Flight spoiler 8 1 EHA 10.2 81.6 EMA 9.0 72.0

Ground spoiler 4 1 EHA 8.4 33.6 EMA 8.9 35.6

Krueger flap 4 1 EHA 19.9 79.6 EMA 13.4 53.6

LE slat 8 1 EHA 16.3 130.4 EMA 9.7 77.6

TE flap 4 2 EMA-LS 44.5 356.0 EMA-LS 44.5 356.0

Total weight of actuators! Total! 927.8 Total! 841.4Electrical wiring Length (m) AWG Weight (kg) Length (m) AWG Weight (kg)

571 4/0 544 571 4/0 544

Total architecture weight! Total! 1471.8 Total! 1385.4

EHA: electrohydrostatic actuator; EMA-LS: electromechanically actuated leadscrew; EMA: electromechanical actuator.

Figure 8. Representative flight profile for mission level ana-

lysis. In addition to the 3000 NM (380 min) flight shown,

shorter flights of 350 NM (63 min), 850 NM (132 min), and

1350 NM (202 min) were also evaluated with representative

cruise altitudes.



Tagge et al.34 for a Boeing 767-type airplane, andwere scaled using guidelines reported by Cleveland35

to Boeing 737-800 size. Power demands for instru-mentation, communication, navigation (tabulated asinstruments), equipment and furnishings, lights,pumping fuel, and miscellaneous loads were also setin a similar manner. The power demand for theassumed electrical wing de-icing system was estimatedusing a rapid evaluation method developed by Meierand Scholz36 which was shown to have good agree-ment with the Boeing 787 wing IPS power demand.Centralized hydraulics were assumed to be retained toactuate the landing gear, landing gear bay doors,brakes, nosewheel steering, and cargo doors. Thepower consumption for these users for differentflight phases was reported by de Tenorio37 for thecase of the similar-sized Airbus A320, and has beenincorporated as a shaft-power offtake (to run engine-driven hydraulic pumps) in this work.

The propulsion system characteristics were gener-ated using the Environmental Design Space tool38

developed for the U.S. Federal AviationAdministrations Office of Environment and Energy(FAA/AEE) and NASA. At the core of this tool is the

Numerical Propulsion System Simulation (NPSS)39

tool, which was used to generate an engine deck thatis representative of a CFM56-7B27 type engine, butdoes not utilize or contain any proprietary information.

The difference between the two configurations interms of mission fuel burn may be caused by a differ-ence in either the non-propulsive power requirementor the subsystem weight. While both effects were con-sidered in this work, the magnitude of the first wasquite small, as seen from Table 7. Further, the peakpower demand for flight control actuation (a factorinfluencing generator sizing) does not occur for anyof the flight conditions of Table 7. Tagge et al.34

note that this critical peak power flight condition isgenerally the OEI crosswind landing condition,where the combination of high load and rate require-ments justifies a conservative sizing approach based oncorner power (Corner power for a control surface isthe product of maximum hinge moment and max-imum angular rate.). Referring to Table 1 and assum-ing operability of both ailerons, both elevators, therudder, and four out of eight flight spoilers in theOEI condition, the combined corner power evaluatesto around 51 kW for both configurations.

Table 6. Non-propulsive power requirements (excluding flight control actuation).

System Type Unit Ground Takeoff Climb Cruise Descent Landing OEI

ECS Bleedless kVA 129 129 210 210 210 129 204

Instruments Electric kVA 8 8 8 8 8 8 8

Equipment Electric kVA 60 60 60 65 0 0 0

Fuel EMDP kVA 7.5 7.5 7.5 7.5 7.5 7.5 7.5

Fuel EDP kW 7.5 7.5 7.5 7.5 7.5 7.5 7.5

Wing IPS Electric kVA 17 17 25 25 25 17 17

Cowl IPS Bleed air kg/s 2.2 2.2 2.2 0 2.2 2.2 1.1

Lights Electric kVA 10 10 10 10 10 10 10

Hyd. users EDP kW 1.43 0 0.67 0 0.44 9.83 9.83

Misc. Electric kVA 5 5 5 5 5 5 5

Demand Electric kVA 124 124 171 173 139 92 260

per Mechanical kW 4.9 4.1 4.5 4.1 4.4 9.5 15.2

Engine Bleed kg/s 1.1 1.1 0 0 1.1 1.1 1.1

ECS: environmental control system; OEI: one engine inoperative; IPS: ice protection system; EMDP: electric motor driven pump; EDP: engine driven

pump.

Table 5. Characterization of aileron, elevator, rudder, and flight spoiler movements.

Control! Aileron Elevator Rudder Flt. spoiler

Motion! / cps / cps / cps cpsGround 1/1 0.4 1/1 0.4 1/1 0.4 1 0.4Takeoff 0.12/0.2 0.4 0.2/0.2 0.4 0.2/0.2 0.4 0.12 0.4Climb 0.12/0.2 0.2 0.2/0.2 0.2 0.2/0.2 0.2 0.12 0.2Cruise 0.5/0.5 0.4 0.3/0.3 0.4 0.3/0.3 0.4 0.5 0.4Descent 0.12/0.2 0.3 0.2/0.2 0.3 0.2/0.2 0.3 0.12 0.3Landing 0.4/0.53 0.3 0.53/0.53 0.3 0.53/0.53 0.3 0.4 0.3



The difference in mission fuel burn is primarilydriven by the difference in weight between the twoconfigurations (Table 4). In addition to the3000NM mission shown in Figure 8, several inter-mediate trip distances that are more commonlyencountered in small narrowbody aircraft operationswere also considered, for which cruise altitudes wereselected based on a review of typical cruise altitudesused by airlines for such distances (Online live flighttracking resource, www.flightaware.com). The com-parison of mission fuel burn for all flight profiles con-sidered is presented in Table 8.

Conclusions and future work

Motivated by the need to consider MEA or AEA sub-systems in greater detail in the conceptual designphase, a methodology was presented that integratesthe analysis of subsystem requirements and effectswith the conceptual stage aircraft sizing process.

This requirements-driven methodology analyzesand sizes the subsystems based on functional require-ments in conjunction with aircraft design parametersthat are either available or easily estimable in the con-ceptual design phase. Since information regardingsubsystem characteristics can be fed back into the air-craft and propulsion system sizing process, the con-verged design point obtained from this iterativescheme is one where the effect of aircraft characteris-tics on the subsystems and the feedback effect of sub-system characteristics on aircraft and propulsor sizinghave been accounted for explicitly. Such a method-ology will give the conceptual phase designer ofMEA/AEA the ability to rapidly evaluate the aircraftand mission level effects of subsystem architectures,

allowing the subsystem concept space to be exploredmore thoroughly. The designer will therefore be ableto make more informed decisions regarding subsys-tem architectures before progressing to the subse-quent design phases, where design freedom isprogressively reduced as configuration characteristicsare frozen, and after which major design changesbecome expensive in terms of both time and cost.

The methodology was applied to the conceptualdesign phase analysis of the flight control actuationsystem of a small narrowbody aircraft (e.g., Boeing737 or Airbus A320). Rather than comparing theeffect of transitioning from hydraulic to electricflight control actuation systems, this paper comparedtwo candidate electric actuation solutions. It wasfound that a configuration that employed EHAs forflight-critical control surfaces and EMAs for non-flight-critical ones enjoyed a marginal aircraft-levelweight advantage over another predominantly-EHAconfiguration. A concomitant fuel burn advantage,driven primarily by the weight advantage, was alsodetected. The conceptual phase designer exercisingthis methodology would be in a position to weighthe observed advantages against additional consider-ations such as safety/reliability. In this specific case,considering the possibility of a single-point jammingfailure of an EMA and the potentially hazardousflight conditions that may result from a jammed spoi-ler or high-lift device, the designer may be inclined toaccept the marginal weight penalty of the predomi-nantly-EHA configuration in exchange for its fail-safe nature.

Additional factors that were not directly con-sidered in this work would also significantly influencesuch a design decision. Given that electric motors ingeneral and power electronics in particular are quitesensitive to operating temperature from the point ofview of performance, reliability, and life, thermal per-formance requirements would be significant designdrivers for electric actuators. The thermal challengeis compounded by the absence of centralized hydrau-lic fluid to act as a heat sink, the likely infeasibility ofactive cooling systems for actuators from a weight

Table 7. Actuation power requirements for PFCS and SFCS for Configurations 1 and 2 (averaged

over mission segment duration).

SegmentConfiguration 1 (kW) Configuration 2 (kW)

PFCS SFCS Total PFCS SFCS Total

Ground 0.02 1.19 1.21 0.02 1.20 1.22

Takeoff 0.24 0.22 0.46 0.23 0.25 0.48

Climb 0.67 0.55 1.22 0.67 0.59 1.26

Cruise 0.99 0.99 0.98 0.98

Turbulence 4.96 4.96 4.92 4.92

Descent 0.97 0.54 1.51 0.96 0.57 1.53

Landing 3.26 0.22 3.48 3.24 0.26 3.50

PFCS: primary flight control surfaces; SFCS: secondary flight control surfaces

Table 8. Mission fuel-burn comparison (Configuration 2

relative to Configuration 1).

Trip Distance! 350 NM 850 NM 1350 NM 3000 NM

Block fuel (%) 0.06 0.08 0.09 0.11



perspective, and the degradation in the convectiveheat transfer coefficient with altitude. In order to befeasible therefore, an electric actuator must bedesigned to tolerate the most constraining thermalloads (which can be initially misidentified8) with pas-sive cooling alone. Clearly, a thermal performanceadvantage enjoyed by an actuator would significantlyinfluence a design decision.

Based on the experience and results obtained fromthis work, the authors have identified three main ave-nues for further research, which will be the subjects offuture publications. The first avenue of research will beaimed at determining how the relative performance ofthe two actuation configurations varies as aircraftgross weight increases. In other words, how significantwould the weight difference be if, instead of a Boeing737 or Airbus A320-sized aircraft, a Boeing 747-sizedor Airbus A380-sized aircraft is considered? As seenfrom Figure 6, the magnitude of the EMAs weightadvantage over the EHA increases as the actuationload increases. However, not only do larger-sized air-craft have larger overall control surface planformareas, they also generally have multiple control surfacesections (e.g., a split rudder, or multiple aileron sec-tions per wing). The actuators driving these split sur-faces would be sized to smaller loads, but they wouldbe larger in number. Second, some aircraft may havecontrol surfaces such as inboard high-speed ailerons(e.g., Boeing 747) that others do not (e.g., Boeing 737).Thus instead of attempting to scale the results (whichwould disregard or mask these configurational differ-ences), future work will attempt to bound the prob-lem by assessing the two configurations applied to asmall number of aircraft which taken together encom-pass a broad range of aircraft gross weights.

The second avenue of research will consider otheraircraft subsystems such as the ECS, IPS, etc. in add-ition to the flight control actuation system, using thesame requirements-driven subsystem sizing methodsdiscussed in this work. This will allow the conceptualdesigner to set mission requirements (payload, range,etc.) and compare a spectrum of aircraft subsystemarchitectures. For example, at the two extremesmight be an AEA and another candidate with conven-tional subsystem architecture, with several MEAoptions in between. More Electric subsystem solutionsapplied to multiple subsystems may in this case resultin large differences in the non-propulsive powerrequirements and weights of subsystems relative to abaseline with conventional architecture. In this case,the feedback loop of the proposed methodologywill allow the snowballing effects of the subsystemchanges on the overall aircraft characteristics to becaptured. Depending on the case at hand, the differ-ence in aircraft characteristics between uncycled(feedback loop not active) and cycled (feedbackloop active) designs may be quite substantial, andsomething of considerable interest to the conceptualphase designer.

The third avenue of research will consider the effectof uncertainty in the decision-making process, inother words, it will deal with decision making in thepresence of uncertainty. This is predicated on the factthat in the conceptual design phase there are signifi-cant levels of uncertainty associated with the param-eters and metrics that the current analysis is based on.For example, the actuator sizing process is driven bythe estimation of control surface aerodynamic loads.These loads are significantly affected by design par-ameters that are seldom considered at the conceptualdesign phase (e.g., flap mechanism kinematics, etc.).Therefore, rather than having a single sizing load, onewould in reality have a probabilistic distribution ofloads that would feed into the actuator sizing process.This uncertainty, coupled with uncertainties in com-ponent material properties, etc., would ultimatelyyield distributions for the metrics of interest (weight,fuel-burn, etc.) on which the design decisions will bebased. The explicit consideration of these uncertain-ties in the decision-making process will allow for adecision that is more robust, i.e., less sensitive to theeffect of the uncertainties.

Funding

This research received no specific grant from any fundingagency in the public, commercial, or not-for-profit sectors.

Conflict of interest

None declared.

Acknowledgements

The authors would like to thank Mr. Christopher Perullo,Research Engineer II at the Aerospace Systems Design

Laboratory, for supporting their investigation on the pro-pulsion side with his expertise and experience with the EDSand NPSS tools.

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