FLIGHT TEST EXPERIENCE WITH AN ELECTROMECHANICAL
ACTUATOR ON THE F-18 SYSTEMS RESEARCH AIRCRAFT
Stephen C. Jensen, NASA Dryden Flight Research Center, Edwards, CA
Gavin D. Jenney, PhD, PE," Bruce Raymond, PE," Dynamic Controls, Inc, Dayton, OH
David Dawson, USAF Wright Laboratory, Wright-Patterson AFB, OH
Abstract
Development of reliable power-by-wire
actuation systems for both aeronautical and space
applications has been sought recently to
eliminate hydraulic systems from aircraft and
spacecraft and thus improve safety, efficiency,
reliability, and maintainability. The Electrically
Powered Actuation Design (EPAD) program was
a joint effort between the Air Force, Navy, and
NASA to develop and fly a series of actuators
validating power-by-wire actuation technology on
a primary flight control surface of a tactical
aircraft. To achieve this goal, each of the EPAD
actuators was installed in place of the standard
hydraulic actuator on the left aileron of the
NASA F/A-18B Systems Research Aircraft (SRA)
and flown throughout the SRA flight envelope.
Numerous parameters were recorded, and overall
actuator performance was compared with the
performance of the standard hydraulic actuator
on the opposite wing. This paper discusses the
integration and testing of the EPAD
electromechanical actuator (EMA) on the SRA.
The architecture of the EMA system is discussed,
as well as its integration with the F/A-18 Flight
Control System. The flight test program is
described, and actuator performance is shown to
be very close to that of the standard hydraulic
actuator it replaced. Lessons learned during this
program are presented and discussed, as well as
suggestions for future research.
Introduction
Power-by-wire (PBW) actuation is the next
major breakthrough in aircraft control. Just as the
fly-by-wire flight control system eliminated the
need for mechanical interfaces, power-by-wire
actuators eliminate the need for central hydraulic
systems. Control power comes directly from the
aircraft electrical system. This has several
advantages. Central hydraulic systems are
complicated and difficult to maintain. Removing
these systems would greatly reduce the amount of
support equipment and personnel required to
maintain and operate current air and space
vehicles. In addition, PBW actuators have the
potential to be more efficient than their
hydraulic counterparts. A central hydraulic
system must generate and sustain significant
hydraulic pressure (3,000 to 6,000 pounds per
square inch) at all times, regardless of demand.
PBW actuators only use electrical power when
needed. Finally, PBW actuation systems can be
made far more fault tolerant than those
depending on a central hydraulic supply. Once a
hydraulic line is compromised, it usually leads to
the loss of that entire hydraulic circuit. As a
result, multiple hydraulic circuits are required to
maintain some level of redundancy. With a PBW
system, a failed actuator can simply be switched
off, isolating the problem to a single surface.
Types of PB W Actuators
There are several different types of PBW
actuators, including electrohydrostatic actuators
(EHA) and electromechanical actuators (EMA).
EHAs use a reversible, electrically driven pump-
motor to directly pump self-contained hydraulic
fluid to a piston. This drives the ram in the same
fashion as a standard hydraulic actuator
(Figure l(a)). An EMA has no internal hydraulic
fluid, instead using electric motors to directly
drive the ram through a mechanical gearbox
(Figure 1(b)). Compared to an EHA, the EMA
has certain advantages. It is lighter, smaller, and
less complex than an equivalent EHA because of
the absence of an internal hydraulic system. Since
there is no hydraulic fluid in the load path, the
EMA tends to be stiffer than an equivalent EHA.
The EMA tends to be more efficient because
there are no windage losses or pump
inefficiencies.Finally,sincethereis noleak
potentialwithanEMA,it is bettersuitedtolong-
termstorageor spaceapplications.
(a) Electrohydrostatic Actuator (EHA).
(b) Electromechanical Actuator (EMA).
Figure 1. Examples of power-by-wire
actuators.
Electrically Powered Actuation Design
Program Goals
The Electrically Powered Actuation Design
validation program (EPAD) was managed by the
Air Force Research Laboratory, Control Systems
Development & Applications Branch, Wright-
Patterson Air Force Base, Ohio, and was a joint
partnership between the Air Force, Navy and
NASA. The objective of the EPAD program was
to establish the credibility of electric actuation as
the method of control for a primary control
surface on a tactical aircraft. The EPAD program
consisted of the design, development, and flight
test of three aileron actuators on the NASA
F/A-18B Systems Research Aircraft (SRA) [1].
The first actuator was the Smart Actuator, a
hydraulic actuator with loop closure and failure
detection performed locally at the actuator
instead of in the F-18 flight control computers
(FCC) as would normally be the case [2].
Communications to the Smart Actuator were by
fiber optics. The second actuator was an EHA,
with an external controller [3]. The third and
final actuator was an EMA, the subject of this
paper. The flight test objectives of the EPAD
program were to measure actuator performance
under actual flight conditions, and subject the
actuator to combined surface loads (inertial,
aerodynamic, and aeroelastic) and environments
(noise, temperature, vibration, and
electromagnetic interference (EMI)). The
actuator was to be subjected to a series of realistic
maneuvers including rapid flight changes, trim
changes, and real flight dynamics.
Note that use of trade names or names of
manufacturers in this document does not
constitute an official endorsement of such
products or manufacturers, either expressed or
implied, by the National Aeronautics and Space
Administration.
Acronyms and Symbols
Bdc
BIT
dB
EHA
EMA
EMI
EPAD
FBW
FCC
g
IBIT
IBOX
MCT
PBIT
PBW
PCME
PCU
q
brushless dc
built-in-test
decibels
electrohydrostatic actuator
electromechanical actuator
electromagnetic interference
Electrically Powered Actuator
Design validation program
fly-by-wire
flight control computer
acceleration of gravity
initiated BIT
interface box
MOS-controlled thyristor (MOS =
metal oxide semiconductor)
periodic BIT
power-by-wire
power, control, and monitor
electronics
power conversion unit
dynamic pressure, lb/ft 2
SRA
V ac
V dc
Systems Research Aircraft
Volts alternating current
Volts direct current
EMA System Description
Standard F/A-18 Aileron Actuator
The standard F/A-18 aileron actuators are
dual-redundant hydromechanical servo-
mechanisms. The F/A-18 flight control system is
divided into four identical channels, with each
aileron being driven by two separate channels.
The system can withstand one electrical failure
and one hydraulic failure and still function. If
either two hydraulic or two electrical failures are
detected, the system will revert to a "trail
damped" mode, fairing into the airstream with
enough dynamic stiffness to prevent flutter.
The ailerons on the F/A-18 are really
flaperons, being used for both roll control and as
flaps (Figure 2). If an actuator failure occurs with
flaps down, the aircraft flight control logic will
slowly bring the opposite flaperon up to maintain
aircraft symmetry while the failed surface is
blown to a faired position.
EC98-44672-3
Figure 2. F/A-18B Systems Research
Aircraft, showing ailerons in flap
configuration.
The EPAD EMA System
Architecture Overview
The EPAD EMA system was designed to be
a simplex replacement for the standard F/A-18
actuator that could be implemented without
modification to the standard aircraft flight
control system (Figure 3). All loop closure and
failure detection occurs between the actuator and
the power control and monitor electronics
(PCME) unit located in the left wing. Two
interface boxes (IBOXs) were required to both
satisfy the loop closure and failure detection
requirements of the aircraft FCCs and, at the
same time, convert the rate commands generated
by the FCCs into a position command usable by
the actuator. A power conversion unit (PCU) was
installed to rectify the 3-phase, 115 V ac aircraft
supply into the ±135 V dc (270 V dc differential)
power required by the actuator. The existing
aircraft instrumentation system acquired data
from the IBOXs, the PCU, and additional aircraft
sensors and telemetered it to the ground for real
time monitoring and recording. Location of the
various components on the aircraft is shown in
Figure 4.
F:
Figure 3. EPAD EMA system layout.
Figure4.EPADEMA hardwarelocations.
ElectromechanicalActuator (EMA)
TheEMA was designed and built by MPC
Products (Skokie, Illinois). This actuator was
designed to meet the same performance
specifications as the standard F/A-18 hydraulic
aileron actuator. The EMA consists of two
3-phase brushless dc (Bdc) motors driving a single
ball screw through a velocity-summing
differential. Mechanical stroke was 4.125 in. and
maximum load was 13,200 lb. The actuator
weighed approximately 26 lb, and was rated at
approximately 5 horsepower maximum output.
The production actuator has the same maximum
load capability, and weighs approximately 17 lb.
Maximum current draw for the EMA was
30 amperes (A) at 270 V dc, with a potential
70 A transient peak. An antirotation device was
incorporated inside the actuator to prevent the
ball screw from turning. The actuator is pictured
in Figure 5.
Photo courtesy of MPC Products, Inc.
Figure 5. EPAD Electromechanical actuator.
Power, Control, and Monitor
Electronics (PCME)
The power, control, and monitor electronics
unit was designed and built by Lockheed Martin
Control Systems (Johnson City, New York). This
unit combined both the low-power actuator
control and monitoring functions and the high-
power, high-speed motor commutation functions
inside the same unit. The unit provided closed-
loop control of the actuator using ram position,
motor velocity, and motor current. The PCME
conducted fault monitoring and continuously
monitored system performance. If a fault were
detected, it would transition the system into a
trail-damped mode, matching the behavior of the
standard hydraulic actuator. It also performed
both periodic built-in-test (PBIT) and initiated
built-in-test (IBIT) functions.
Commutation was provided by a series of
MOS-controlled thyristors (MCTs), which
provided a trapezoidal torque function. Actuator
power was controlled using pulse-width
modulation (PWM), performed by an additional
MCT. The PCME is pictured in Figure 6.
EC93 -41023 -3
Figure 6. Power, control, and monitor
electronics (PCME).
Interface Box (IBOX)
Dynamic Controls Incorporated (Dayton,
Ohio) designed and built the IBOXs specifically
for the EPAD program. The IBOXs served as the
interface between the FCC and the PCME. They
allowed the use of a research actuator on the F-18
without requiring modification of the aircraft
flight control system. The IBOXs collected data
from the PCME and transmitted it to the aircraft
instrumentation system by means of a MIL-STD-
1553B [4] databus. An IBOX is pictured in
Figure 7.
4
program.It wasdesignedtosupplypowertothe
actuatorat4-135 V dc, up to 100 A. This power
was produced by rectifying the 3-phase, 115 V ac
supply produced by the aircraft generators. The
PCU also served to block any regenerated power
coming back from the EMA system. A PCU is
pictured in Figure 8.
EC93-41023-9
Figure 7. EPAD interface box (IBOX).
Experiment Integration
Iron Bird Testing and Simulation
Before the EMA system was installed on the
SRA, it was first installed on a hardware-in-the-
loop test bench which replicated the attach
points and kinematics of the left F/A-18 aileron
(Figure 9). The avionics were installed on the
F-18 Iron Bird (which is a retired F-18 airframe).
This setup was used to perform system
integration, verification, validation, and failure
modes and effects testing without tying up the
aircraft. In addition, several mission profiles and
failure scenarios were "flown" both by engineers
and pilots by connecting the Iron Bird with the
Dryden F- 18 simulator. This simulation proved
invaluable in assessing the hazards of system
failures at various points in the flight envelope,
as well as generating emergency procedures.
EC95-43334-1
Figure 8. EPAD power conversion unit
(PCU).
Power Conversion Unit (PCU)
The PCU was also developed by Dynamic
Controls Incorporated, specifically for the EPAD
EC94-42668-5
Figure 9. EPAD aileron test bench with EMA
actuator installed.