Moving
Towards
a
More
Electric
Aircraft
J.A.
Rosero,
J.A.
Ortega,
E.
Aldabas,
&
L.
Romeral
ABSTRACT
The
latest
advances
in
electric
and
electronic
aircraft
technologies
from
the
point
of
view
of
an
"all-elect~ric"
aircraft
are
presented
herein.
Specifically,
we
describe
the
concept
of
a
"More
Electric
Aircraft"
(MEA),
which
involves
removing
the
need
for
on-engine
hydraulic
power
generation
and
bleed
air
off-takes,
and
the
increasing
use
of
power
electronics
in
the
starter/generation
system
of
the
main
engine.
Removal
of
the
engine
hydraulic
pumps
requires
fully-operative
electrical
power
actuators
and
mastery
of
the
flight
control
architecture.
The
paper
presents
a
general
overview
of the
electrical
power
generation
system
and
electric
drives
for
the
MIEA,
with
special
regard
to
the
flight
controls.
Some
discussion
regarding
the
interconnection
of
nodes
and
safety
of
buses
and
protocols
in
distributed
systems
is
also
presented.
INTRODUCTION
Conventional
aircraft
architectures
used
for
civil
aircraft
embody
a
combination
of
systems
dependent
on
mechanical,
hydraulic,
pneumatic,
and
electrical
sources.
The
resulting
conventional
equipment
is
the
product
of
decades
of
development
by
system
suppliers.
In
a
conventional
architecture
(Figure
1
is
a
basic
schematic)
fuel
is
converted
into
power
by
the
engines.
Most
of
this
power
is
used
as
propulsive
power
to
move
the
aircraft.
The
remainder
is
converted
into
four
main
forms
of
non-propulsive
power
[1]:
Pneumatic
power,
obtained
from
the
engines'
high-pressure
compressors.
This
kind
of
energy
is
conventionally
used
to
power
the
Refereeing
of
this
work
was
handled
by
L.M.
Kaplan.
Manuscript
received
November
30,
2005;
revised
May
23.
2006.
Author's
Currnt
Address;
LA.
Rosero,
J.A. Ortega,
E.
Aldabas,
&
L.
Romeral
0885/8985/07/
USA
$25.00
0
2007
IEEE
IEEE
A&E
SYSTEMS
MAGAZINE,
MARCH
2007
Environmental
Control
System
(ECS)
and
supply
hot
air
for
Wing
Anti-Icing
(WAI)
systems.
Its
drawbacks
are
low
efficiency
and
a
difficulty
in
detecting
leaks.
"
Mechanical
power,
which
is
transferred
(by
means
of
the
mechanical
gearboxes)
from
the
engines
to
central
hydraulic
pumps,
to
local
pumps
for engine
equipment
and
other
mechanically
driven
subsystems,
and
to
the
main
electrical
generator.
"
Hydraulic
power,
which
is
transferred
from
the
central
hydraulic
pump
to
the
actuation
systems
for
primary
and
secondary
flight
control;
to
landing
gear for
deployment,
retraction,
and
braking;
to
engine
actuation;
and
to
numerous
ancillary
systems.
Hydraulic
systems
have
a
high
power
density
and
are
very
robust.
Their
drawbacks
are
a
heavy
and
inflexible
infrastructure
(piping)
and
the potential
leakage
of
dangerous
and
corrosive
fluids.
"
Electrical
power,
which
is
obtained
from
the
main
generator
in
order
to
power
the
avionics,
cabin
and
aircraft
lighting,
galleys,
and
other
commercial
loads
(such
as
entertainment
systems).
Electrical
power
does
not
require
a
heavy
infrastructure
and
is
very
flexible.
Its
main
drawbacks
are
that
conventionally
it
has
a
lower
power
density
than
hydraulic
power,
and
results
in
a
higher
risk
of
fire
(in
the
case
of
a
short
circuit).
Each
system
has
become
more
and
more
complex,
and
interactions
between
different
pieces
of
equipment
reduce
the
efficiency
of
the
whole
system.
A
simple
leak
in
the
pneumatic
or
hydraulic
system
may
lead
to
the
outage
of
every
user
of
that
network,
resulting
in
a
grounded
aircraft
and
flight
delays.
The
leak
is
generally
difficult
to
locate
and
once located
it
cannot
be
accessed
easily.
The
trend
is
to
move towards
"all-electric"
aircraft,
which
means
that
all
power
off-takes
from
the
aircraft
are
electrical
in
nature,
thus
removing
the need
for
on-engine
hydraulic
3
Fligh~t
Hydraulic
Power
Mechanical
kaloN
Power
r
Elcrc
Electrical
Geeao,
Power
Ice
Protection
Pneumatic
Comfpressor
Power
ECS
approaches
to
on-board
energy
power
management
and
drive
systems
(Figure
3).
These
are
now
being
carefully
considered,
and
it
is
believed
that
electrical
systems
have
far
more
potential
for
future
improvement
than
conventional
ones
regarding
energy
efficiency.
fIot __, Ulvt
o
Fig.
1.
Schematic
of
conventional
power
distribution
power
generation
and
bleed
air
off-takes.
The
removal
of
bleed
air
off-takes
requires
new
high-voltage
electrical
networks
and
new
solutions,
such
as
air-conditioning,
wing
ice
protection,
or
electric
engine
start-up.
Removal
of
the
engine
hydraulic
pumps
requires
fully-operative
electrical
power
actuators
and
a
mastery
of
flight
control
architecture.
The
"all-electric"
aircraft
is
not
a
new
concept:
the
concept
of
an
electric
aircraft
has
been
considered
by
military
aircraft
designers
since
World
War
11
[2],
although
until
recently
the
lack
of electrical
power
generation
capability,
together
with
the
volume
of
the
power conditioning
equipment
and the
advanced
control
required,
rendered
the
approach
unfeasible
-
especially
for
commercial
and
civil
transport
applications.
Since
the early
1990s,
research
into
aircraft
power
system
technologies
has
advanced
with
the
aim
of
reducing
or
eliminating
centralized
hydraulics
aboard
aircraft
and
replacing
them
with
electrical
power.
Several
programs
have
been
started
with
the
aim
of
driving
the
research
on
this
field
[3],
such
as
Totally
Integrated
More
Electric
Systems
(TIMES),
devoted
to
use
previously
developed
systems
into
electrical
aircraft,
US
Air
Force
MEA
Program
that
investigates
for
providing
more
electrical
capability
for
fighter
aircrafts,
and
Power
Optimized
Aircraft
(POA),
which
tries
to
optimize
the
management
of
electrical
power
on
aircraft
in
order
to
reduce
non-propulsive
power
and
reduce
fuel
consumption,
while
increasing
the
reliability
and
safety
of
onboard
systems
and
reducing
maintenance
costs.
Nowadays,
novel
ways
of
generating.
distributing,
and
using
power
onboard
are
examined
at
the
aircraft
level.
Hybrid
or
bleed-less
air
conditioning
systems,
"More
Electric
Engines"
(MEEs),
fuel
cells,
variable
frequency
generators,
complex
embedded
digital
systems
and
distributed
system
architectures
are
just
a
few
of
the
technologies
vying
for
space
on
forthcoming
aircraft;
the
concept
is
known
as
"More
Electric
Aircraft"
(MEA)
as
presented
in
Figure
2.
Recently,
worldwide
research
into
the
future
development
of
commercial
aircraft
has
given
rise
to
more
advanced
Fig.
2.
Current
trends toward
the
MEA
Power
ft
Au
Compressor
P a
Fig.
3.
Schematic
of
MEA
Power
Distribution
Steps
toward
a
MEA
are
being
taken
in
two
different
ways:
"
Removing
current
air
and
hydraulic
engines
and
further
increasing
electrical
power
generation
capability.
This
requires
significant
changes
in
electrical
generation
and
network
techniques,
and
in
fault protection.
*
Substituting
hydraulic
actuators
for
electromechanical
actuators.
This
reduces
weight
and
decreases
maintenance
and
production
costs.
The
MEA
initiative
emphasizes
the
utilization
of
electrical
power
in
place
of
hydraulic,
pneumatic,
and mechanical
power
to
optimize
the
performance
and
life cycle
cost of
the
aircraft.
The
MEA
requires
a
highly
reliable,
fault
tolerant,
autonomously
controlled
electrical
power
system
to
deliver
higher
quality
power
and
electrical
levels
to
the
aircraft's
4
IEEE
A&E
SYSTEMS
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MARCH
2007
Loa
_J s
111111111111010.
Engine
syýpenis
OMMOMMM+
Ice
Protection
- I
EC
8
4
Central
Hydraulic
Pump
Flight
controls
loads. Also, reliable
high
integration
and
safety
of
the
electrical power system leads
to
the use
of
distributed
generation
and control
architecture.
The advantages of
More
Electric
systems
are
not
confined
to aircraft.
Other
transport
systems,
such
as
marine
propulsion,
are
also
moving
in
this
direction
[4].
The
next
sections
briefly
discuss
a
general overview
of
the
electrical
power
generation system
and electric
drives
on the
MEA,
especially
with
regard
to
the
flight
controls.
A
brief
introduction
to
the
safety
aspects
of
the
flight
controls
has
also
been
included.
ELECTRONIC
POWER
SYSTEMS
The first factor
to
take
into
account
is
the
large
amount
of
power
electronics
for
power
conversions
and
power users
that
MBA
will
involve:
at
least
1.6
MW
for
a
next-generation
300
pax
aircraft. The
development
of
efficient
and
secure
power electronics
technologies
is
a
great challenge.
However,
not
only are
power
electronics
necessary,
but
also
efficient
control
of
the
electronics
must
be
developed.
One
major evolutionary
technological advance
that
has
contributed
greatly
to
the
feasibility
of
an
electric
aircraft
non-propulsive
power
system
has
been
the
development
of
reliable, solid-state,
high
power-density,
power-related
electronics. Generator
power
control
units,
inverters,
converters,
and
motor
controllers
consist
of
state-of-the-art
silicon-based
power
semiconductor
switching devices
that
include
integrated
gate
bipolar
transistors
(IGBTs). It
is
expected
that advanced
composition,
high-performance
multi-layer
ceramic capacitors
will
dramatically improve
the
power
density
of
future
inverters,
converters,
and
motor
controllers. Improved, high-efficiency
electric
circuit
topologies
are
also
the
subject
of
on-going
research.
Some
of
the
higher
power
level
equipment
is
actively
cooled
through
the
use
of
oil
circulation
or
forced
air
convection.
The
extent
of
the
use
of
active,
fluid-based
cooling
systems
is
extremely
application-specific
and
is
yet
to
be
determined.
Lightweight,
simplified, passive
(non-pumped
fluid-based)
thermal
management
techniques
are
also
a
focus
of
research
and
will
be used,
wherever
feasible,
to
maintain
high
reliability.
Power
Distribution
and
Management
Systems
(PDMS)
provide
fully
automatic monitoring,
control, protection,
and
switching
of
aircraft
electrical loads
under
normal
and
emergency
conditions
with
load
management,
including
automatic
load
shedding
and
restoration,
to
make
best
use
of
available
power.
These
systems
comprise
the
Primary Power
Buses,
located close
to the
generators,
with
high
power
contactors
and
circuit
breakers,
and the
Secondary
Power
Distribution
Buses,
located
in
the avionics bay,
which
provide
the
monitoring
and
control
of
the
system,
and
contain
some
same
circuit
breakers
and
remote
power
switches.
The
use
of
programmable
solid-state
devices
and
switching
power
devices
in
place
of
traditional
electromechanical circuit
breaker
technology provides
benefits
to
the
aircraft
in
terms
of
load
management,
fault
isolation,
diagnostic
health
monitoring,
and
improved
flexibility
to
accommodate
modifications
and
system
upgrades.
With these advancing technologies,
it
will
be
feasible
to
use
high
power-density
electrical
power components
to
drive
the
majority
of
aircraft
subsystems. These
will
become
easier
to
maintain (supported
by
less
equipment
and
manpower),
more durable,
lower
in
cost,
and
higher
in
performance.
The
engine primarily provides
thrust,
but
it
also
produces
all
other
power
(Figure
1).
In
a
MEA,
current
engine
accessories
that
derive power form
gearbox
mounted
pumps
will be
replaced
with
electronically-driven
electrical
machines. Vibration
resistance,
electromagnetic
compatibility,
and
size
constraints
are
key
design challengers
of
embedding
electrical machines
into
the
engine.
The
integration
into
a
harsh
environment
of engine
off-takes
for
aircraft
system
needs
without
significantly affecting
engine
performance
is
also
a
difficult
task.
By
deleting
air off-takes,
virtually
the
only
requirement
the
engines have
to
satisfy
is
to
provide electrical
power.
Whilst
the
hot-air
bleed
ducts
and the
pre-cooler
are
removed,
several
other
integration
issues arise, such
as
generator
thermal management, mechanical
integration,
new
electric starting
requirements,
and
electrical
power
conversion,
(whether
the
chosen
solution
is
a
conventional
gearbox-mounted
generator
or
an
embedded
power-optimized
generator).
Conceptually,
electrical
power
for
an MEA
would
be
produced
by
a
starter/generator directly
driven
by
the
gas
generator
spool
of
the
main
engine.
Power
is
transferred
out
of
the
engine through
wires
that
feed
into
a
fault-tolerant
electrical
network
to
drive
the
aircraft
subsystems.
Electronic
power
converters
would
transform
the
electrical
power
and
no
accessory gearboxes
would
be
necessary.
Elimination
of
gearing
and
associated
gear
separation forces
enhances
the
use
of
advanced
magnetic bearing
systems
[5]1,
which
could
be integrated
into
the
internal
starter/generator
for both
the
main engine
and
auxiliary
power
units.
For
many
years, electrical
power
for
aerospace
applications
has
been
generated
using
a
variable
ratio
gearbox-mounted
wound-field
synchronous
machine
to
obtain
a
three-phase
115
V
AC
system
at
a
constant
frequency
of
400
Hz.
This
machine
is
known
as
a
Constant
Frequency Integrated Drive
Generator
(IDG),
and today
it
is
still
the
most
commonly
used.
However,
operating
experiences
under
the
new
requirements
of
lower
cost,
increased
reliability, easier maintenance, and
higher
operating
speed and
temperatures have
shown
that
a
replacement
for
the
gearbox
using
power
electronics has
obvious advantages.
A
high
quality
three-phase
AC-DC
conversion
plus
subsequent
DC-AC
conversion
is
one
of
the
steps involved
in
achieving
these
objectives.
The
resulting
system
is
known
as
variable speed
constant
frequency
(VSCF)
system, and
it
results
in
promising technology
that
meets these
requirements.
Figure
4
shows
a
typical
block
diagram
of
a
VSCF
system.
In
the
motoring
mode,
the
constant
frequency system
IEEE
A&E
SYSTEMS
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20075
5
supplies
the
machine
through
the
power
converter,
and
the
system
acts
as
a
starter for
the
aircraft
engine.
In
the
generating
mode, the
main
engine
moves the
machine,
providing
electrical
power
at
a
variable
frequency
which
is
transformed
into
a
constant
frequency
by
the
power
converter.
The
bidirectional
power converter
can
be
built
using
a
DC
link
in
a
back-to-back
topology
- a
mature
technology
in
use
in
civil
aircraft (Boeing,
MacDouglas,
etc.)
or
by
using
a
direct
AC-to-AC
converter.
This
is
a
new
technology
that
is
increasingly
used
in
military
fighter
aircraft.
The
matrix
converter
[6]
is
a
clear alternative
to
any
other
AC-to-AC converter
for
aerospace
applications.
The
converter consists
of
nine
bi-directional
switches arranged
as
three sets
of
three
so
that
any
of
the
three
input phases
can
be
connected
to
any
of
the
three
output
lines.
The
switches
are
then
controlled
in
such
a
way
that
the
average
output
voltages
are
a
three-phase
set
of
sinusoids;
of
the
required
frequency
and
magnitude.
Some
of
the
advantages
of
the
converter
that
make
it
a
promising
technology
for
the
near
future
are
as
follows:
"
A
higher
power
ratio
with
a
lower
size
and
weight.
"
Unity
power
factor
control.
"
It
is
free from
bulky reactive
components
(especially
large
electrolytic
capacitors).
Electromagnetic
interferences
due
to
large
currents
and
voltages
high
frequency
switching
are
the
main
disadvantage
of power
electronics
supplying
actuation systems.
These
interferences
can
be
alleviated
by
reducing
the length
of
electrical
cables
supplying
power
and even
more
by
integrating
the
matrix
converter
into de
motor-actuator
system.
Moreover,
the
ability
of
matrix
converter
to
supply
almost
sinusoidal currents
helps
to
reduce
these
interferences
as
well.
Application
of
higher
voltages
is
also
investigated,
which
allows
reducing
the
weight
for
the
power
used.
230/400
VAC
400
Hz
could
be
relevant
for
some
electrical
subsystems
because
of
its
lighter
weight
generator
system.
270
VDC
is
commonly
used
as
DC
link
bus
voltage,
whereas
the
motor
controllers
can
use
even
higher
level,
540
VDC.
Another
solution
to
generate
electrical
power
for
the
aircraft
consists
in
variable-frequency
(VF) power
generation,
which
allows
designers
to discard
the
complex
and
difficult-to-maintain equipment
necessary
to
convert
variable-speed
mechanical
power
produced
by
the
engines
to
constant-frequency
electrical
power traditionally
used
by
aircraft systems.
By
this
way,
variable-frequency power
generation
increases reliability
of
the whole
system.
Of
course,
aircraft's
systems
such
as
fuel
and
hydraulic
pumps
and
EHA/EMA
actuators
have
to
be
designed
to
be
compatible
with
VF
generation
and
distribution.
Yelrtr
V-rable
O-tart
Frenuert
Frequer
Speed
Fqcontrol
Or
Syncroflrizabon
Speed
Fig.
4
VSCF
Starter/Generator
System
Variable-frequency
power
generation
is
now
coming
for
large
aircrafts,
and
it
is
expected
that
power generation
reliability
will
be
increase
by
about
50
percent,
although
the
challenges
related
to
advanced
electromagnetic
technology,
high-speed
electronic
voltage
regulation
and
system
protection
to
maintain high-level
power quality
over
the wide
output
range have
to
be
solved.
The
switched
reluctance
machine
is
very
promising
as an
integral
starter/generator
system
in
future
aircraft
integral
engines.
The
simple
rotor
construction
and high
power
density
of
the
machine permit
high
speed
and
high
temperature
environment
operation. The
possibility
of
direct-driving
and,
hence,
the
elimination
of
gear
boxes
and
hydraulic
accessories
in
the
aircraft
may
give
it
in
an
advantage
over
the
classical
synchronous
and
induction
machine
technologies.
Reduction
of
an
aircraft's
multiple
secondary
power
subsystems
to
a
single
electric
subsystem
is
another
challenge
under
development.
There
are
numerous
generator
and
distribution
choices
to be
made
for
this
architecture,
such
as
ECS
and
Electro-Thermal
WAI, but
careful
application of
the necessary
system
integration
must
be done,
and
analysis
tools
to
design
and
verify
the
integrity
of
the
new
hardware-
and
software-based
systems
are
necessary.
Apart
from generators
and
loads,
other
elements
are
needed
for
the
control and
management
of
high-power
electrical
energy.
Power
electronics
and
control
are
seen
as
the
major
and
most
crucial
technologies
for
an
MEA,
which
faces
the
challenges
of
reduced
package
size,
higher power
capability,
reduced
acquisition
cost,
and
high
efficiency.
ELECTROMECHANICAL
ACTUATORS
Subsystems
of
the
MEA
include
power
electronics,
power
controllers,
converters,
inverters,
and
associated
components,
which
have
a
direct
impact
on
the
viability
of
the
MEA,
especially
in
the case
of
control
actuators.
The
basic
building
blocks
for
control
actuators
are
solid-state
power
electronics
and variable
speed
motor
drives.
Fully fault-tolerant
Control
Management
and
communications
for
decentralized
systems
are
also
required
to
link
and
control
the
wide
range
of
variables
used.
In
the
area
of
Actuation Systems,
alternative
architectures
incorporating
electro-hydrostatic,
hybrid
and
6
IEEE
A&E
SYSTEMS
MAGAZINE,
MARCH
2007
6
electromechanical
actuation
for
primary
and
secondary
flight
control
(as
well
as
new
landing
gear,
braking,
nacelle
actuation,
and
horizontal
stabilizer
architectures)
are being
examiined.
A
large
number
of
actuators
have
been
studied,
most
of
them
electromechanical
except
flight
control
actuators
due
to
the
showstopper
jamm-ing
case.
In
the
last
decade,
a
lot of
research
has
allowed
Electro-Hydrostatic
Actuator
(EHA)
technology
to
be
mastered.
One
result
of
this
on
new
aircraft
such
as
the
Airbus
A380
or
Boeing
B7E7
is
the
replacement
of
the
hydraulic
circuits
by
EHA
networks.
These
are
used
as
a
back-up
for
other
hydraulic
systems,
although
there
is
increasing
interest
in
the
use
of
electric
drives
to
substitute
hydraulics
and
electro-hydraulic
systems
in
aircraft.
In
such
systems,
an
electric
motor
directly
drives
a
pump,
a
fan,
or
an
actuator.
I---------------------------------------------
'Superv/ision
Subsystem
------------------------
I
II
I
Poe--4
ioq,-
See]
-op
7
Iovetr
ICirfe
orme
000
Motion
Subsystem
Fig.
5.
Direct
Drive
architecture
for
EMA
In
fact, the
next
step
from
the
present
hydraulic
or
electro-hydraulic
actuation
(EHA)
in
a
centralized
system,
to
the
use
of
Electromechanical
Actuators
(EMAs)
in
de-centralized
systems
(while
maintaining
the
same
level
of
safety)
is
today
of
major
importance
for
aeronautics.
The
objective
is
to
reduce
production
and
maintenance
costs.
Furthermore,
these
highly
safe
and
reliable
EMA
technologies,
which
are
jamming
free,
will
help
to
satisfy
the
social
demand
for sustainable
transport.
EMA
technologies
are
already
being
used
in
aeronautics,
but
for
safety
reasons
they
are
limited
to
Secondary
Flight
Controls
or
military
aircraft
[7].
Their
application
to
Primary
Flight
Controls
will
allow
reductions
in
the
weight
of
drives,
gas
consumption,
and
polluting
emissions.
The
major
step
in
moving
from
EHAs
to
jam-free
EMAs
is
the
prevention
of
potential
jamming
cases
by
appropriate
technology
and
monitoring,
thus
giving
the
system
aircraft
availability
for
dispatch
and
failure
sizing
cases.
Electromechanical
Actuator
drives
for
flight
controls
are
based
on
a
Direct
Drive
architecture
built
up
by
an
electric
motor,
(usually
a
Permanent
Magnet
Synchronous
Motor,
PMSM)
directly
connected
to
the
roller-screw
that
moves
the
actuator
(Figure
5).
The
power
stage
can
be
built
up
either
by
standard
inverters
or
by
new
matrix
converter
architectures.
The
complete
control
block
diagram
for
an
EMA
drive
includes
the
position,
speed,
torque
and
flux
controls,
and
also
the
supervisory
and
communication
systems.
From
the
previous
statements,
it
is
clear
that
not
only
power
electronics,
but
also
electric
machines,
are
becoming
more
and
more
important
in
the
general
electric
aircraft
power
system,
both
for
generation
and
load
control.
Specifically,
the
PMSM
is
increasingly
being
used
for
actuator
drives,
due
to
its
high
efficiency
throughout
the
full
speed
range,
high
power
ratio,
and
ease
of
refrigeration,
compared
to
classical
wound
machines
[8].
The
drive operating
the
flight
control
must
ensure
continuity
in
operation
even
in
the
case
of
a
fault. Dual
redundant
power
drive
electronics
providing
motor
drive,
speed
closed-loop,
and
control
management
can
help
to
overcome
this
issue.
With more
electronics
in
the
actuators,
it
is
also
possible
to
predict
how
long
an
actuator
will
last,
introducing
the
predictive
maintenance
instead
of
preventive
maintenance
today
used
by
airlines.
The
drive
should
also
be
able
to
diagnose
and
report
the
nature
of
the
fault.
The
system
must
also have
the
following
general
characteristics
[9]:
"
Testability,
to
make
verification
and
real-time
check-out
easier.
*
Reduced
complexity
and
low
maintenance
costs,
by
the
decomposition
of
the
main
CPU
into
smaller
distributed
controls
for
every
EMA,
many
of
them
consisting
of
identical
hardware.
"
Intelligent
software
running
in
every
control
node,
which
must
be
able
to
exchange
information
by
means
of
standard
interfaces.
To
achieve
the
above
specifications,
control
and
diagnosis
of
the
EMA
needs
to
rely
on
modem
electronics.
As
in
other
fields,
digital
computer
control
systems
have
been
incorporated
into
aircraft
avionics
system
design.
Digital
systems
are
more
reliable,
lighter
and
more
adaptable
to
change
or
modifications,
as
well
as
providing
self-test
capability.
For
these
reasons,
embedded
digital
control
systems
are
going
to
be
extensively
used
in
the
aeronautical
industry.
The
growth of
electronics
has
also
led
to
drive-by-wire
control
systems
in
which
there
are
no
physical
connections
(mechanical,
pneumatic,
or
hydraulic)
between
sensors,
controls,
and
actuators.
Similarly,
a
fly-by-wire
(FBW)
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7