ELECTRIC ACTUATION FOR FLIGHT & ENGINE …€¦ · Electric Actuation for Flight & Engine Control:...

SAE-ACGSC Mtg 99Slide 1

Electric Actuation for Flight & Engine Control:Evolution & Challenges

Feb.28-March 2, 2007 Boulder Meeting

ELECTRIC ACTUATION FOR FLIGHT & ENGINE CONTROL:EVOLUTION & CHALLENGES

Amit KulshreshthaMoog Inc.

&Jean-Jacques CHARRIER

Hispano-Suiza




Rationale for Electric Powered Actuation in Aircraft• Reduced Weight

– Electric Powered Actuation allows the aircraft certification with 3 or, 2 Electric & 2 Hydraulic compared to conventional aircraft requiring 3 Electric & 3 Hydraulic

– A weight saving of ~ 1000 Lbs.for A380 and 400 Lbs for F-35 has been attributed to Electric Actuation due to savings from additional Hydraulic Systems. Weight savings by elimination of a hydraulic system are dependent upon the aircraft size.

• Improved Performance & Optimization– Hydraulic Pump/System is a continuous load on the engine whether Hydraulic Power is

used for actuation or, not while Electric load is on demand/when needed. – 25% reduction in peak non-propulsive power usage with 5% reduced fuel consumption:

2000 Lb wt. Reduction for A340 results in fuel savings of 55 Lbs/hr ~ Total 550 Lbs for 10 hr flight

• Improved Maintainability & Survivability/Robustness– Elimination of a hydraulic system improves reliability significantly due to low MTBF of

the hydraulic system~engine driven pumps, pressure seals and leakage etc.– Efficient Segregation & Independence of the Actuation Power provides Robustness

Developments in enabling technologies-electric motors, electronics etc. offer improved performance & optimization with lower maintenance cost




Conventional A330/340 Flight Control- Actuation System

Source: Airbus/SAE Presentation

3 Hydraulic Systems (Y, G & B ) distributed amongst all Actuators




More Electric-A380 Flight & Actuation Control (2H/2E) System

A380 More Electric Actuation with 2H/2E allows ~1000Lbs Wt Saving and use of higher pressure (5000 psi) allowed further ~2000 Lbs resulting in total 3000 Lbs Wt Saving

Source: Airbus/SAE Presentation




Electric vs Electrohydraulic Actuation

Variable-DeliveryConstant-PressureHydraulic Source

270 VDC ElectricalPower Source

270 VDC ElectricalPower Source

Variable-SpeedBrushless Motor

Flow ControlBy Throttling(Servovalve)

Variable-SpeedBrushless Motor

HydraulicCylinder

MechanicalTransmission

Fixed-Displacement

PumpHydraulicCylinder

VCA

EMA

EHA

Mechanical/HydraulicEnergy Conversion

Mechanical/ElectricalEnergy Conversion

Mechanical/ElectricalEnergy Conversion

RPM Flow

RPM

Velocity

Velocity

Velocity

Flow

Mech/HydConversion

Hyd/MechConversion

Elect/MechConversion

Elect/MechConversion

Hyd/MechConversion

Oil Cooling

Air Cooling

Air Cooling




Electric Actuation in Aircraft & Engine Control: What Does it Mean?• Actuation system derives power from aircraft & Engine Control and provides required

force/moment with a defined displacement within specified time to move a load/controlled element. Typical actuation system in aircraft is run by central hydraulic systems powered by engine driven hydraulic pumps and/or, central electric motor driven pumps. Engine control is performed by engine fuel based actuators similar to hydraulic servo actuators.

• Electric actuation system uses local electric power drive driving a mechanical transmission system for Electro-Mechanical Actuators (EMA) or, hydraulic transmission system such as Electro Hydro-static Actuators (EHA) or, Integrated Actuator Package (IAP).

RVDT#1

ROLLER SCREW/BALL SCREW

ELECTRIC MOTORCOMMUTATIONSENSOR-RESOLVER

MECHANICAL TRANSMISSION

SENSORINTERFACE

AUX POWER28 VDC

ELECTRONICS MOTOR CONTROL UNIT (EMCU)

CONTROLELECTRONICS

115/230VAC,Var Freq

BIT & MONITORELECTRONICS

POSITION /VELOCITYCOMMAND

EHM

Mode

device

Ram

Manifold

Electronics

Motor

PumpAccumulator

Mode

device

Ram

Manifold

Electronics

Motor

PumpAccumulator

ELECTRIC DRIVE

Electro-Mechanical Actuator (EMA) Electro-Hydro-Static Actuator (EHA)




Electrical Powered Actuation-EMA

• EMA based design can be made fault tolerant to motor & electronics failures~fail op/fail safe design with combination of control moment or, torque summing at control surface or, speed summing.

– EMA technology has been successfully deployed in thrust vector controls in launch vehicle/missiles and aircraft hi-lift/ non-flight critical controls.

– EMA design require prevention of back driving by adding brakes or, ‘no-backs’. – Technology capable of 400 deg f temp motor & electronics is enabling its use in severe

environment & extended storage life 20+ years.– Mechanical linkage between power source (electric motor) & load (control surfaces) &

mech. Transmission (ball screw or roller screw are gears with no fluids/leakage but subject to jamming; newer hi torque/low speed may eliminate gear.

– Additional work in progress for jam tolerant, failsafe with auto-centering after failures to address electric motor failures, loss of lubrication at –55~-65 deg C and mechanical jamming.




Electrical Powered Actuation-EHA

• Combines advantages of electric power & hydraulic actuation; integrated with conventional hydraulics – EHV for EBHA.

• Transmission linkage between electric motor & loads is hydraulic fluid making system immune to jamming.

• Proven hydraulic actuation with historical data on known failure modes of the system with ability to either disengage failed motor or, lock hydraulic actuator/fluid to prevent flutter depending upon failure.

• Requires digital control of high power electronics to control electric motor torque & speed as needed to meet load requirements.

• Storage life of 10+ years




Electrical Powered Actuation-IAP

• Integrated Actuator Package (IAP)– Uses electric power with swashplate controlled variable

displacement pump connected to constant speed motor drive.– Eliminates high power motor control electronics with low-power

servo control of the swashplate for flow control to hydraulic RAM.– Hydraulic transmission linkage between electric motor & load

eliminates jamming problems. – Constant load on pump results in heating of hydraulic fluid requiring

cooling of IAP unlike EMA/EHA where electric power is consumed only when needed.




Electro-hydrostatic Actuation (EHA) Evolution for Electrical Back-up Hydraulic Actuator: EBHA

Source~Airbus/SAE




Issues & Perception for EMAs in Flight Controls?

• EMAs are prone to Jamming and may result in the (i) Loss of the Control & (ii) Control Surface stuck up in failed/extended position

• EMAs including EMCU are not very Reliable and have more failure modes requiring monitoring for Fault/Jamming Detection, Isolation & Reconfiguration

• EMAs including their Power Converters-Electronics Motor Control Units (EMCU) generate significant heat and consume too much Electric Power-Energy Inefficient

• EMAs don’t provide adequate primary flight control bandwidth in FBW of Air Vehicles and are Heavierwith added redundant motors requiring structural strengthening

EMA should be designed to provide level of ‘availability’ & ‘integrity’ as provided by current primary flight control actuation systems by using Jam & Fault Tolerant Design , Parametric Design Optimization & Optimal Control for Efficient Actuation

Some of the possible solutions for high integrity/high availability EMA through newer technologies & fault tolerant designs are presented here




Primary Actuation Control: Remote Actuator Control• Accomplished using Actuator Mounted Remote

Electronics Unit or, Electronic Motor Control Unit

• Digital Data Bus Interface simplifies Aircraft Wiring between FCC and Actuator Control Electronics-REU or, EMCU based upon proven Safety Critical Data Bus: Could replace ~300 wires between FCC and Primary Actuators to 30-60 wires

• Analog Control Loop and Monitoring for Simple Servo Control & DSP based Control for Electric Motor Control with Independent Health Monitoring Interface for Fail/Safe Design

– Design/Verification of Software/Complex Hardware to DO-178B/DO-254 Level A is necessary condition for certification but not sufficient condition if it does not address ‘generic error free design’

– Civil Certification of processor/firmware (FPGA) based Motor Control Electronics –where a ‘generic’ failure may result in failure of multiple surfaces/actuators may result in architectural mitigation and/or, generic failures tolerant designs i.e., Dissimilar Control & Monitoring Option etc.

B787 Hydraulic Actuator with Digital Data Bus Interface

A400M Spoiler EBHA with Actuator Mounted EMCU and Hybrid (Analog & Digital Data Bus) Interface




Aircraft Flight Control Requirements: Load vs SpeedS

peed

(in/

s)

Typical Actuator Force vs. Speed: Single System

0 deg F, min capability70 deg F, min capability40 deg F, min capabilityRequirementStall Load

2 2.5 31 1.50 0.50

1

2

3

4

5

6

7

8

9

10

x 104Force (lbf)

Load

, % o

f Max

imum

80.0 100.040.0 60.00.0 20.0

0

20

40

60

80

100

Velocity, % of Maximum

HP1

HP2

Corner HP100% Load, 100% Velocity

Maximum Power Point~37% of Corner HP66.67% Load, 57.7% Velocity

Typ. EMA / EHACapability

Typical Load/Speed Curve for Flight Control ActuatorActuator Sizing Power Point : Current Generation Electric vs. Hydraulic Actuation

Bandwidth not a significant issue for Hydraulic or, IAP, it could be an issue for EMA and EHA motor load inertia and the acceleration requirements

Need for design optimization with Electric Actuation due to limited power availability compared to Hydraulic Actuator




Electric Motor Technology- Evolution

• Current Electric Actuators are mostly based upon Surface (Mounted) Permanent Magnets (SPM): PM-BLDC or, PMAC-Synchronous Motors;both designs are magnetically non-salient with relatively high torque ripple- with poor efficiency, a problem at low speed operation

• Many of the new generation designs are using Interior Permanent Magnet (IPM) motor where magnet is embedded in the rotor providing salient pole design with wider constant operating range

– IPM provides 12-15% additional torque compared to traditional SPM used. Baseline design with hybrid cars




Electric Motors Technology- Switched Reluctance Motors

• Switched Reluctance Motors (SRM) are rotating electric machines where both stator and rotors have salient poles. The motor is excited by current pulses applied across each of the phase, forcing the motor to rotate.

• SRM don’t use Permanent Magnet (PM) and so are prime candidates for High Temp environment where PM start loosing their magnetization; some hybrid i.e., PM-SRM motor designs provide fault tolerant motor for moderate applications

• More Electric Engine (MEE) are based upon SRM design based Starter Generator installed in the core of the engine; can be designed for very high temp with ceramic based winding




Electric Motors Technology-High Torque/Low Speed Motors for Direct Drive EMA

Roller Screw

Ref:KTH Report 2005Hi Torque Density Motor Design can provide Gearless Design




Mechanical Transmission System-Power Screw Comparison

Ball Screw Roller Screw ComparableLowest System Weight xJam Resistance xBacklash Control xWear xExternal Preloading xInternal Preloading xEfficiency xLowest Cost xHighest Stiffness xContamination Sensitivity xSpeed Sensitivity xLubrication Retention xSide Load Capability xBallistic Tolerance x

AdvantageParameter




Mechanical Transmission System- Typical Efficiency Comparison

Ball Screw Example




“Towards More Electric Propulsion System”

• Aircraft manufacturer’s moving toward “more electric” as means to reduce costs and weight– Believe “pneumatics”---big cost and weight driver--- has matured

and no significant gains possible– Power electronics & variable speed motors--- “enabling”

technologies for “more electrics”– First affected aircraft --- A380 and B787

• Propulsion System impact– A380- Electric Thrust Reverser (E-TRAS), but basic engine

subsystems (GP7200 and Trent 900) not affected– B787- significant changes being implemented in areas of ECS

(Environmental Control System), flight controls, main engine starting, and hydraulic power

• significant challenges for engine due to very large power extractions (starter generators),




Interest for a More Electrical Engine




• STUDY VEHICLE : CFM56 ENGINE




Engine Control - Component Environment• Thermal Environment

– dependent upon location (fan or core compartment)– Fuel carrying components typically located forward of combustor– Max air temperatures typically 300-400 F max but radiant heating

may be significant for some components

• Vibration Environment– Dominant input from rotor one per rev externally mounted

components though blade passing effects can also be significant

• Fuel Temperatures– Range is –65°F to ~300° F– Dominant thermal effect for most fuel carrying component

• Fire Certification getting tougher- FARs say “flame direction toward most critical components …”




Engine Control - EMAs:Unique Design Goals

• Jam tolerant-prefer jam immune designs

• Back drivable when power is removed

• Back driveable after any jam when power is removed

• No lost time of delayed response when operating through a jam

• Be able to report the first jam

• Light, stiff and low cost




Interest for a More Electrical Engine• Major Sub-systems for Turbo Fan Control System

Variable Stators ValvesAim is to adjust compressor input flow

Weight Fuel MeteringAim is to adjust a parameter effecting thrust

Variable Bleed ValveAim is to adjust flow into booster stage

• Large Commercial

– Fuel, Variable Stators, Variable Bleed Doors, Turbine Active Clearance Control, Booster Anti-ice Stability Bleed

– Fuel powered actuation/EHSV controlled via FADEC

• Small Commercial

– Fuel, Variable Stators and Stability Bleed

– Fuel Powered/EHSV Controlled

• Military

– Fuel, Variable Stators, Fan IGVs, and A8

– Fuel powered/EHSV controlled





• Specifics of electrical engine control systems– Redundancy of the electrical circuits so as to comply with the engine certification rule:

CS-E 50 (c):

• The Engine Control System must be designed and constructed so that:1) The rate for Loss of Thrust (or Power) Control (LOTC/LOPC) events, consistent with the safety

objective associated with the intended aircraft application, can be achieved:

2) In the Full-up Configuration, the system is essentially single Fault tolerant for electrical and electronic failures with respect to LOTC/LOPC events.

– High reliability: in order to maintain same level: ~ 5.10-7/EFH per function reinforce the need of electrical circuit redundancy

– Criticality of power electronic box heat dissipation– High density

– Unfavorable thermal ambience ~ 194°F

– Impact on the needed cooling system

– Criticality of actuator thermal ambient temp : 320°F/356 °F

– The electrical system can not be heavier than the existing system




Variable Geometrie

Power Electronic Channel A

VGPE #B

VSV RDE

VSV RD1

VSV RD2

VBV LEFT

VBV RIGHT

TBV

LPTACC

HPTACC

IGNITER LEFT

IGNITER RIGHT

DECU

Power Generation and Distribution 1

Power Generation and Distribution 2

PMG#A

PMG#B

AIRCRAFT ELECTRICAL NETWORK

Interest for a More Electrical Engine• GENERAL ARCHITECTURE

– 2 PERMANENT MAGNET GENERATORS

– A single power distribution box per lane, one of those is interconnected to the aircraft electrical network

– A single electronics controlling/commanding all GVs per lane

– GV ACTUATORS




• VSV ACTUATORS - ISO CFM56





BetterControl performance

IdenticalWeight

CompatibleHydraulic system : piping routing concern

Electric system : electronic boxes installation concern

Installation

BetterHeat exchanger 50% size reduction

Volume

BetterLeakage reduction by 2.10-5 D&C / FC

Availability (D&C)

IdenticalConsider as Input data

Safety

TBDCost

Electrical system Vs hydraulic system PerformancesCriteria's




Power Generation for the Engine Control SystemPower Generation Architecture: REDONDED PMG

3 ph A/C Network

AGB

CONVERSION AC-3ph / DC

I/F BUS DC


I/F BUS DC

CORRECTION THD

CONTROL EQUIPMENTS

FADECGV

IGNITER

BUS DC 200/300 V


I/F BUS DCBUS DC 200/300 V

BUS TRANSFERT

SWITCH




Classical Fault Tolerant Dual Electric Drives: Force Vs. Velocity Summing

Force Summing

• Baseline – power screws mechanically tied together at output

• Failed motor may require mechanical disconnect

• 50% loss of output force

• Output velocity preserved

Velocity Summing

• Motors connected to power screw via differential geartrain

• Failed motor required engagement of a brake

• 50% loss of output velocity

• Output force preserved




Current Generation Fault Tolerant EMA Architecture With Dual Wound Stator & Common Rotor

Dual wound fault-tolerant PMAC machine

10-poles, 12 slots dual wound PMAC machine, with physically, magnetically and thermally

isolated phases.

Phasor diagram

RVDT#1

RVDT#2

ROLLER SCREW/BALL SCREW

M2M1

R1 R2

DUAL WOUNDELECTRIC MOTOR

COMMUTATIONSENSOR-RESOLVER

MECHANICAL TRANSMISSION

SENSORINTERFACE

SENSORINTERFACE

AUX POWERAUX POWER28 VDC28 VDC

ELECTRONICS MOTOR CONTROL UNIT (EMCU) #1

CONTROLELECTRONICS

CONTROLELECTRONICS

115/230VAC,Var Freq

115/230VAC,Var Freq

ELECTRONICS MOTOR CONTROL UNIT (EMCU) #1



The use of one common rotor driven by two redundant stator windings with their 3 phase motor inverter drives does not result in added inertia

load of the second rotor and helps to improve performance




Traditional Fault Tolerant – Dual Redundant Electric Drive• Two independent electro-mechanical paths until there is a mechanical

summing ( velocity or, torque) to drive the mechanical ransmission/actuation or, dual tandem EHA

POSITIONCONTROLLER

SPEEDCONTROLLER

CURRENTCONTROLLER

PWMGENERATOR

ELECTRICCOMMUTATION

POWERDRIVER

CURRENT DETECTION

SPEED GENERATOR

POSITION SENSOR

HULL SENSOR

REDUNDANCY I

DRIVEGEAR RUDDER

θg Igγg

POSITIONCONTROLLER

SPEEDCONTROLLER

CURRENTCONTROLLER

PWMGENERATOR

ELECTRICCOMMUTATION

POWERDRIVER

CURRENT DETECTION

SPEED GENERATOR

POSITION SENSOR

HULL SENSOR

REDUNDANCY IIθg Igγg

Motor

Motor

P4P3P2P1

Pump

Pump

A1 > A2 A3 = A4

ControlElectronics

InverterElectronics

FCCCommand

Electrical Feedback

1a°

1b°

1c°

ControlElectronics

InverterElectronics

Electrical Feedback

1c°

1b°

1a°FCCCommand

Dual Redundant EMA

Dual Tandem EHA

-

-

BRAKE& MO TO RCONTROLNVERTER

BRIDGE

RESOLVERSIGNALC OND

DEMOD;CMD/MON

POW ERCONDITIONING UNIT;AC~DC CONVERTERACTIVE FRONT END

DC LINKVOLTAGE270 VDC

115 VAC, 380 Hz ~ 900 Hz

BUS 'B 'BUS 'A '

LVPS

D ATA BUS;CMD/MON

BRAKECO NTRO L& M OTOR

DRIVE-INVERTER

BRIDGE

RESOLVERSIGN ALCOND

D EMOD;CMD/MON

LVPS

DATA BUS;CMD/MON

MOTORCOIL 'A '

ELECTRICBRAKE 'A '

MOTORRESOLVER

AC TUATORPOSITION

RESOLVER#1

AC TUATORPOSITION

RESOLVER#2

-

-

MOTORCOIL 'B '

ELECTRICBRAKE 'B'

MOTORRESOLVER

ACTUATORPOSITION

RESOLVER#1

ACTUATORPOSITION

RESOLVER#2

DIGITALDATA

BUS 'B '

D IG ITALDATA

BUS 'A '

28 VDCBUS 'B'

28 VDCBUS 'A '

POW ERCONDITIONING UNIT;AC~DC CONVERTERACTIVE FRO NT END

DC LINKVOLTAGE270 VDC

ELECTRICDRIVE UNIT(EDU); Eachactuator w ith

dual wound coilmotors is

controlled bydual redundant

ElectronicControl Units

(ECUs).

ECU 'A '

ECU 'B '




Next Generation Fault Tolerant EMA Architecture: Multi-Phase Electric Drives

• Electric Drives with Multi-Phase Motors Provide Fault Tolerance against Stator Coil or, Inverter IGBT or, Gate Drive Failures

• Multi-Phase based EMAs do not suffer from added rotor/load inertia associated with redundant motor based fault tolerant approach still providing same availability. Use of PMAC, SRM and PM-BLDC Motors has been demonstrated and test flown in few cases

ab

cd

evdc+-

Five-phase motor drive

Motor load

Five-phaseSynchronous

Motor

ωr

Five-PhaseSync MotorController

tes

tds

tcs

tbs

tas

Gate Drivers

tbs

tcs

tds

tes

tas

E+D+C+B+A+

E-D-C-B-A-

DC-LinkOver-Voltage

Protection

3-PhasePower Supply

L Li

C

iAiB

iCiD

iE

Poly-phase faulttolerant electric motor;phase motor coils withreliable rotor

Roller/Ballscrew

Monitor ElectronicsType B

Motor ControlProcessor; Type B

Motor ControlProcessor; Type A

Monitor ElectronicsType A

Eb Estimation

Eb Estimation




Electronics Motor Control Unit (EMCU)-Functional Elements

• Power Converter-Electronics Motor Control Unit (EMCU) typically includes:– (i) Power Conditioning – dependent upon AC vs. DC input,– (ii) Control Electronics and, – (iii) Power Drive/Motor-Inverter Drive.

DC

AC

AC~DC Power Conversion:Passive vs. Active Control

Motor-Inverter Drive: Control& Power Drive Electronics

DC-link

AC

DC

EnergyStorage

DC

AC

AC

DC

AC

AC

Direct Power Conversion (single stage)Indirect Power Conversion

Control Electronics for Signal Processing & Control/Optimization

Electronic Control of Electric Drives allows Optimizing Actuation Control System




ELECTRONIC MOTOR CONTROL UNIT; EXPLODED VIEW

BUS CAPACITORS & REGEN POWER DISSIPATION CONTRIBUTE TO A LARGE EMCU SIZE




Power Converter:Active AC~DC & AC~AC Regenerative Conversion with options for putting Regen Power on bus

Vienna Rectifier AC~DC

Conversion

One Cycle Control (OCC) Analog AC~DC

Conversion

Selection of Active AC~DC conversion is dependent upon

the electric power quality specs. & wt./vol. constraints while

AC~AC provides ‘motor inverter’ function for actuator control

Extract from ETH & UCI Presentation

C-

C+ uC+*

* uC-

iM

TSi=RN

UN,i

iN,i

UU,i

DF+

DM+ DN+

DF-

DM- DN-

OU21

+

OU21

−

VoltageDivider

RloadE

Dcp

DcnDbn

Scn

Dan

SbnSan

Dbp

Scp

Dap

SbpSapia, b, c

Region Selection Circuit

NVa Vb Vc

Logic

S Q

R Q

⏐⏐⏐

⏐⏐⏐

S Q

R Q

⏐⏐⏐

⏐⏐⏐

+-

+-

2

2To Driver ofThree-phaseHalf-Bridge

Ref+

Av(s)+- -

++

++

++

Voltage loopCompensator

Qn

Vm

Reset Signal

Extendedone-cyclecontrol core

AnalogSwitches

Current SignalSelector ⎟

⎠⎞

⎜⎝⎛

τ−•

tiVm

LaLbLc

AB

C

BU

A

B

C

a

b

c

p

n

(c)

Spa Dap

u

A

B

C

p

n

(b)

Sapa SpA DaP

Sana

a

b

c

a

b

c

(a)

A

C

i

Matrix Converter-Conventional 18 IGBT

Reduced Switch/IGBT Very Sparse Matrix

Converter(9~12 switches)

Back-Back Converter with DC Link and Boost Inductors

RE-GENERATIVE CONVERTER TOPOLOGY: CAN PUT REGEN. POWER ON AC BUS




Actuator Mounted Motor Electronics Control Unit (EMCU)

Electronic Motor Control Unit (EMCU)

Cmd. Data BusInterface

Motor Drive Bridgeand Brake Drive: SiC

type switchingelements with Drivers

Resolver &Position-Sensor

Feedback

Motor ControlElectronics & SensorFeedback: Resolver/

RVDT

Aircraft HighIntegrityElectricPower

Monitor Electronics forMotor Control -Data

Bus I/FFCC:FLIGHT

CONTROLCOMPUTER

Motor-Dualwound- with

BrakesLow Voltage PowerSupply (LVPS)28VDC

Electric Drive Unit: Electronics integrated with Motor

DIGITALDATA BUS

BRAKE CONTROL POWER

MOTOR CONTROL POWER270VDC or,115/230VAC

Wiring PWM/Square Wave based Power Drive/IGBTs and Motor Winding results in distortion of the waveform at motor winding with consequences in high voltages and premature failures of the electric drives

Induced Voltage in Motor Windings with PWM Inverter Drives vs. Sine wave

Electric Motor with Power Electronics Source: ARL & University of Nottingham

Power Drive Electronics (PDE) with Electric Drive helps to Optimize the Weight & Power/Performance




Makes the electrical conversion needed via software programming

Senses what they are plugged into...

Senses what is plugged into them...

Controls

I/O I/O

Thermal

PEBBFunctions In SoftwareInverterBreakersFrequency ConverterMotor ControllerPower SupplyActuator Controller

Integrated Modular Architecture for Power Converter

* PEBB defined by IEEE (Power Engineering Society)

‘Modular’ & ‘Partitioned’ design with defined interfaces allows selective replacement and upgrade without recertification

High Integrity Data-Bus supporting ‘temporal’ and ‘space’ partitioning simplifies cost effective designs




Power Converter-Modular Design: Functional Partitioning

.1- 1us10 us

Power In

100 us1 ms

10 ms

Power ModulePower Module

Power Filter

Power Filter

Power Filter

Power Filter

Load Controller

Load Controller

Flight Control

Computer

Flight Control

ComputerInner LoopInner Loop

SmartGate

Drives

SmartGate

Drives

SmartSensorsSmart

Sensors

SmartSensorsSmart

Sensors

SmartSensorsSmart

Sensors

ModulatorModulator

DIGITAL

Control Electronics

MIXED SIGNAL

SerialBus

SerialBus

Power Drive Electronics

SerialBus

SerialBus

Courtsey: VirginiaTech/ IEEE Power Electronics

Building Block (PEBB)

Standardization

High Speed & High Integrity Serial Bus is one of key element for it




USAF/AFRL Electric Actuation Program Objectives

Source: AFRL/USAF SAE AE-7 Presentation




Aircraft Electric Power System & Electric Actuators

• Primary flight control actuation system requires highly dependable (including availability with associated integrity) electric power as Electric Actuation moves forward to be primary actuation for aileron, elevator and rudder.

• High reliability/availability of electric power requirements translate to minimum number of elements between electric power generation and its utilization i.e., primary flight control actuation system and electric power available in presence of common cause/generic error in electric power/generation control (if any?) system.

• Flight Control/Fly by Wire System requires high availability i.e. ‘non interruptible’ electric power free of ‘generic errors’ in either by design or architectural mitigation in electric power generation and its distribution systems as it relates to FBW system

• Electric actuation has been designed to work with either 270VDC or, 115/230VAC, Variable Freq electric power; it has been indicated that on one to one comparison, 270VDC electric power/switch-gear (solid state) equipment is more expensive, larger and heavier compared to 115/230VAC equipment.

• Next generation aircraft electric power system should be designed to ‘absorb’ the ‘regen’ power to minimize heat dissipation in electric motor control electronics. Regen energy could be sufficiently high for primary actuators during braking/electric power interrupts when aero loads tend to back-drive the surface/actuators.




Power Drive Electronics-Wide Band Gap Device ~SiC Availability for Low Loss & Hi Temp

• SiC key property includes its low losses even at high temperature operation compared to Si

• SiC based Diodes are commercially available and provides 30% reduction in power dissipation for motor inverter drive in recently built prototypes

• SiC based devices are being sampled , configured for parallel operation and prototype built for 500V, 50kW Toyota Prius motor drive

PROPERTY SILICON SILICON CARBIDE BANDGAP (eV @300K) 1.1 2.9 MAX. OPERATING TEMP. 425 deg. K >900 DEG K

BREAKDOWN VOLTAGE (1E6 V/CM) 0.3 4.0 THERMAL CONDUCTIVITY (W/CM-DEG C) 1.5 5.0 PROCESS MATURITY HIGH LOW INTRINSICALLY RAD HARD NO YES




Future Electric Drives-Advanced Motor Control Technologies

• Energy Optimization Control for static and dynamic energy minimization to reduce thermal power dissipation and improve the reliability of the drive.

• Provide capability for motor/drive to estimate its position, speed and torque without adding additional sensor thus improving reliability of the system.

• Provide PHM capability to reduce the current dependence on redundancy in sensors/actuators by developing realistic models of the performance degradation.

The future of Aircraft Electric Drives is dependent upon adding more value while improving reliability and maintaining high level of performance with safety/integrity at no additional cost

CurrentRegulatedPWM-VSI

Position& VelocityObserver

MotionController

&

FieldOrientedTorque

Controller

Signal separationfilters

βφ−βφ

AC Machinewith Rotor Saliency

ωΩθΩ

Tem or Tem^ *

θr^

ωr^

Td^

iqds-fδ*

vqds-βδ*

iqds-fδ

iqds-αδ

Self-sensing via observer-based tracking of intrinsic machinesaliencies with inverter-based carrier injection

PWMVoltageSource

Inverter

Band pass filter

AC Machinewith Rotor Saliency

ωnθr

iqds-enδ

iqds-f + iqds-eδ

Self-sensing carrier excitation in a motor drive

SaliencyImage

TrackingObserver

ωnθrTd^ δ

iqdsδ

vqdsδ

vqdsδ∗

CurrentRegulator

vqds-dδ∗

++

vqds-cδ∗

Band stop filter

iqds-fδ*

iqds-fδ




Jam TolerantMechanism &

SwitchableLoad Path

Fault TolerantMotor

Windings vs.Redundancy

Direct Drive/High Torque/Low Speed

Motor

Fault TolerantElectronic

Motor ControlArch. & I/F

MagneticBearings &

Bearing-lessMotor

JAM & FAULTTOLERANTACTUATION

SYSTEM

Prognostic &Health

Monitoring

MagneticGearings

EnergyStorage for‘Return to

Neutral’ whenPower Off

Enabling Technologies & Key Components for Safe & Reliable Jam/Fault Tolerant EMA

• Different Customers may require some or, all issues to be included in their applications depending upon the criticality of the application and overall architecture and their perception on Jamming

• An inter-disciplinary approach is needed to address development and integration of various technologies to provide a modular approach to allow one or, more features may be integrated for an application/customer




Electric Actuation Technology: USAF/AFRL Roadmap

Summarizing, Electric Flight Control Actuation offers an opportunity for simplified and lower integration with appropriate data bus and electronics control allows

optimal control to improve thermal power with duty cycle

Source: AFRL/Joe Weimer




SUMMARY• Electric Actuators-EHA & EMA have been developed and being deployed on first

generation More Electric Aircraft

• Current generation EMA suffer from concerns on possible Jamming and many jam tolerant and jam immune mechanizations have been developed and demonstrated; still to undergo demonstrated reliability and safety requirements

• High Torque Density Electric Motors employing IPM design and concentrated windings have been demonstrated for aircraft use and axial design motors are under evaluation to improve it further

• The cost and failure rates of electric actuators is higher than hydraulics/fuel-draulic actuators but they provide cost, weight and reliability improvements at overall air-vehicle design level.

• Fault tolerant design and deterministic health monitoring are key enabling technologies for ensuring availability of control function with electric actuation at affordable price

• Distributed control of Electric Actuators with integration of Control function in Flight/Engine Control Computer and Power Electronics is a cost effective approach

ELECTRIC ACTUATION FOR FLIGHT & ENGINE …€¦ · Electric Actuation for Flight & Engine Control:...

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