IEEE Electrification Magazine - December 2017 - 40

40

Figure 1. The architecture of common hydraulic flight-control actuation. TRU: transformer rectifier unit; PSI: pounds per square inch.

Actuator
Position
Feedback
Flight
Sensor
Inputs

Command
Pilot
Inputs

Flight-Control
Computer

28 Vdc

I E E E E l e c t r i f i cati o n M a gaz ine / DECEMBER 2017

Other Electric Power
Consumption

Electric Power
(Generator)

Aircraft Engine
(Power Source)

Hydraulic Power
(Reservoir/Pump)

Other Hydraulic
Power Consumption

115 Vac

3,000 PSI

Power Conversion
(TRU)

28 Vdc

EHSV Position
Feedback

Actuation Power Systems

Actuator Controller

EHSV

EHSV Control
Current

High/Low
Pressure

Hydraulic Actuator

Low/High
Pressure

28 Vdc

Flig

ht S

urfa

ce

required to hold these loads. If the
flight-control surface could be held in
one position, a brake on the back of the
motor could be used to lock the actuator and hold the static load. But, because
the actuator position is constantly dithering, a brake cannot be set, and the
actuator must rely on the motor to continuously produce rated torque. This
constant load not only presents a thermal challenge for the motor; it also
presents a thermal challenge to the
controller. These are some of the reasons that hydraulic actuators have
been used to control flight surfaces
since the 1930s (SAE 2015).

In addition to thrust, the main engines
provide power for actuation. Conventional aircrafts use the engine to power
the system from both hydraulic pumps
and electric generators. On the A380, Airbus was able to increase electrical power
generation and distribution and eliminate one of the three hydraulic systems
by managing redundancy with electrically powered loads. Most conventional
commercial aircraft have three hydraulic
systems while being equipped with two
main electrical systems. This provides a
total of five power sources for flight-critical actuation. The A380 aircraft maintained the two main electrical systems
while reducing the hydraulic systems
from three to two sources, thereby
reducing the total to four power sources
(van den Bossche 2006). By eliminating
one hydraulic system, aircraft weight
was reduced while maintaining the necessary actuation redundancy for safety.

Actuation System Applications
Primary Flight Controls
Figure 1 shows the architecture of a common flight-control actuation system. The
system provides relatively low power
(watts) to the electronics that control the
flight computer as well as the actuator
controller. The source for high power
(kilowatts) comes from the hydraulic
supply generated by the engine-driven
pumps. In simple terms, the electrical
system controls the brain, and the hydraulic system controls the brawn.



Table of Contents for the Digital Edition of IEEE Electrification Magazine - December 2017

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