Development of a Fly-By-Wire Flight Control System to Achieve Level 1 Handling Qualities on a BLACK HAWK
Helicopter
Igor Cherepinsky
Flight Controls
Sikorsky Aircraft
Stratford, CT
icherepinsky@sikorsky.com
CW4 Sean Christian Magonigal
UH-60M Fly-By-Wire CTT Pilot
Aviation Flight Test Directorate
Redstone Arsenal, AL
sean.magonigal@us.army.mil
Joseph Driscoll
Handling Qualities
Sikorsky Aircraft
Stratford, CT
jdriscoll@sikorsky.com
Stephen Silder
UH-60M Fly-By-Wire Project Pilot
Sikorsky Aircraft
Stratford, CT
ssilder@sikorsky.com
ABSTRACT
Sikorsky Aircraft and the US Army have completed development and test of a UH-60M helicopter equipped with a
modern Fly-By-Wire (FBW) Flight Control system. This paper describes the hardware architecture including the use of high
availability processing elements for full-authority flight critical vehicle control. The software architecture that uses a
partitioned operating system, and was utilized to facilitate safety and reduce software cycle time, will be described. Control
Law (CLAW) strategy using model following techniques with auto-mode logic was employed to further reduce pilot
workload and will be discussed in the paper.
A model based development process was used throughout the entire program for software and control law design
and test. To accomplish this, Sikorsky Aircraft developed a simulation environment and modeling tools that, together with
the U.S. Army Aeroflightdynamics Directorate (AFDD)-developed CONDUIT® [1] flight control analysis and optimization
tool, were integral parts of the process. The use of the flight control system integration lab (FCSIL) for pilot-in-the-loop and
automated testing was extensively used to reduce risk and speed up development. Early CLAW prototyping on Rotorcraft
Aircrew Systems Concepts Airborne Laboratory (RASCAL) [2] was used to ensure success on the first flight. The value of
the approach taken will be discussed in this paper.
More than four hundred hours have been flown by the two aircraft prototypes. The culmination of these activities led
to achievement of Level 1 Handling Qualities in Good and Degraded Visual Environments (GVE/DVE). Flight test results
and lessons learned are documented in this paper.
INTRODUCTION 
The BLACK HAWK helicopter is a versatile utility
helicopter with variants used by many branches of the US
military and international customers. Its design roots are in
the 1970‟s, when electronic augmentation was limited to
analog circuitry and digital processing was just gaining wide
acceptance. Since the initial design, the flight control system
on the BLACK HAWK has undergone several upgrades.
These upgrades were limited to addressing hardware
obsolescence issues and marginal improvements in handling
qualities via introduction of a more sophisticated autopilot
and a flight director system. In 2007 Sikorsky Aircraft and
the US Army set out to make the BLACK HAWK
helicopter, the safest and best handling helicopter in
production. By this time handling qualities requirements
were solidified by a multitude of research performed by the
Aeroflightdynamics Directorate of the US Army and NASA.
This research culminated in the ADS-33 specification [3].
The Fly-By-Wire system for the BLACK HAWK helicopter
is designed to meet Level 1 Handling Qualities (as defined
in ADS-33E) in Good and Degraded Visual Environment
(DVE/GVE). It is also designed to increase mission
availability, greatly improve maintainability and lower
operating costs.
SYSTEM DESCRIPTION
Presented at the American Helicopter Society 68th Annual
Forum, Fort Worth, Texas, May 1-3, 2012. Copyright © 2012 by
the American Helicopter Society International, Inc. All rights
reserved.
The UH-60M FBW Flight Controls (FCS) system was
designed to meet the following requirements:
 Level 1 Handling Qualities in GVE/DVE


No pilot selectable modes for the hands-on
flight (i.e. Flight Director mode selection is
permissible)
Probability of loss of control 1x10 -9 per flight
hour
Fly-By-Wire System Hardware Description
The UH-60M FBW system diagram is shown on Figure
1. The heart of the system is a triplex set of Flight Control
Figure 1 - UH-60M FBW System
Computers (FCCs). These FCCs form the computing
backbone of the system. Each FCC contains two identical
processing lanes that are setup to cross check each other
during various stages of processing. If differences between
lanes are detected, either the affected signal or the entire
FCC is declared invalid – depending on the magnitude and
computational stage where the discrepancy was detected.
The triplex set of FCCs is linked with a redundant, high
speed Cross Channel Data Link (CCDL). The IEEE-1394 is
used for that purpose. It is configured as a dual ring, which
provides a very high availability CCDL. Either four wire
breaks or multiple hardware faults must occur before an
FCC loses communications with the others.
The hydraulics and actuation sub-systems of the legacy
UH-60 were updated to meet the reliability requirements
stated above. The UH-60M FBW hydraulic system consists
of three full-time transmission driven hydraulic pumps
connected to switching and isolation valve assemblies.
These assemblies, under the control of the FCCs provide for
further reduction in pilot work load by fully automating the
task of leak detection and leak isolation. The servos used on
the FBW system are dual ram electro-hydraulic servo valve
(EHSV) controlled.
Unique trim, center mounted long pole controllers are
utilized in the pitch and roll axes of the aircraft. The yaw
axis consists of unique trim pedals. These controllers always
return to center (or “detent”) position. The use of such
controllers allows implementation of multiple modes,
utilizing the same control strategy: displace the controller
from center to command the aircraft and return the controller
to center to hold the newly achieved state. For the collective
axis, a conventional displacement control is utilized. During
early studies it was determined that fully proportional
collective allows implementation of all desired vertical
modes and allows easier transition from the conventional
aircraft.
To provide better situational awareness tactile cueing is
utilized on three out of four primary controls. The Active
Inceptor System (AIS) provides tactile cueing to pitch / roll
cyclic and collective controls for each pilot. The AIS also
provides electronic linking between the pilot and copilot
controls. Electronic linking can be disabled from either
control station by the use of a guarded switch on the cyclic
grip. The AIS itself is a triplex-redundant system;
encompassing all needed control electronics and software for
inceptor force and position loop closure. The FCS
commands desired tactile characteristics via a triplex
redundant IEEE-1394 bus. The AIS applies the commanded
characteristics and sends back to the FCS its status, forces
and control positions via the same IEEE-1394 bus.
The sensor compliment consists of a set of triplex
Inertial Measurements Units (IMU), dual Air Data
Computers (ADC), dual Embedded GPS/INS (EGI) and a
radar altimeter. The IMUs are the main source of angular
rates, attitudes and body accelerations. The system is
designed to operate in full augmentation mode with as little
as one IMU operational. The EGIs provide velocity and
position sensing. These signals are used throughout the flight
envelope to provide velocity and position stabilization, as
appropriate. The ADCs provide airspeed and pressure
altitude. The airspeed signal is used for gain scheduling and
airspeed hold mode. Pressure altitude is used for barometric
altitude hold functionality. In case of complete loss of
airspeed, the system reverts to a set of fixed gains that
provide slightly degraded handling qualities.
The FCS serves as a primary data path for the
propulsion system on the aircraft. The FCS is responsible for
digitization of the engine control quadrant inputs, their
redundancy management and engine mode selection. It is
also responsible for setting a reference speed. The engine
control system governs the power turbine and the rotor speed
to this reference. This arrangement allows a tight coupling
between the FCS and the engine controls, which in turn
allows rapid changes to rotor speed commanded by the FCS
in response to various external factors. It also enables
implementation of rotor speed governing modes during
various engine failure scenarios. The engine control system
collects all engine parameters and sends them to the FCS.
The FCS is responsible for fault detection and redundancy
management of these parameters, and their dissemination to
the rest of the avionics on the aircraft.
Fly-By-Wire System Software Description
To allow for a reduced software development cycle, a
time and space partitioned operating system is utilized in the
FCCs. Use of such an operating system allows mixing
multiple software criticalities (as defined by DO-178B) on
the same processor. It also allows for localized software
changes and required regression testing to be limited to a
particular partition. The partitioning scheme employed on
the UH-60M FBW is shown in Figure 2. Due to
documentation overhead, it was decided to support two
levels of criticality in the system. Any function that was
deemed to have criticality level „B‟ or higher was
implemented in a flight critical (level „A‟) partition. Any
function that was designated as level „C‟ or lower was
implemented in a non-flight critical (level „C‟) partition. The
system was broken into three parts – input, processing and
output. Each of the three parts was split into flight critical
and non-flight critical portions. Furthermore the processing
portion was split into primary control and augmentation
partitions. The resulting seven partitions are described
below.
The flight critical Input Signal Management (ISMa)
partition contains all software functions necessary for fault
detection and redundancy management of all signals that are
needed for safe flight of the aircraft. These signals include
angular rates, attitudes, linear accelerations, linear velocities,
vehicle position, engine signals and a variety of other flight
critical inputs.
The non-flight critical Input Signal Management (ISMc)
partition contains the remaining fault detection and
redundancy management functions. These functions include
processing of inputs from the Control Display Unit (CDU),
navigational radio inputs, inputs from the Flight Director
Control Panel (FDCP), and other miscellaneous non-flight
critical inputs.
The flight critical Primary Flight Control System
(PFCS) partition contains the algorithms for basic modelfollowing control laws. These control laws provide rate
command / attitude hold response type and ground handling
for the aircraft. This partition also contains the logic for the
hydraulic system as well as a full complement of initiated
built-in test algorithms. The UH-60 aircraft is equipped with
a moving stabilator. The control laws and logic for the
stabilator resides in this partition as well.
The flight critical Flight Augmentation and Cueing
System (FACSa) partition contains the remainder of the
model-following control law algorithms. These algorithms
provide several different response types, described
elsewhere in this paper. This partition also contains all the
task-tailoring and automoding logic, as well as algorithms
for tactile cueing.
The non-flight critical Flight Augmentation and Cueing
System (FACSc) partition contains the algorithms for the
Flight Director system as well as provisions for non-flight
critical tactile cues.
The flight critical Output Signal Management (OSMa)
partition contains the algorithms to monitor the health of the
FCC and based on that assessment enable or disable the
outputs of the FCC. It also contains the algorithms to
monitor the status of each primary servo and manage the
servo redundancy. The OSMa partition also contains various
other output functions such as communications with the
cockpit, communications with the engine control system and
others.
The non-flight critical Output Signal Management
(OSMc) partition contains the algorithms that provide output
to the CDU and the FDCP as well as manage and display
Figure 2 - UH-60M FBW Software Partitioning
Scheme
system maintenance fault information.
Control Law Modes
Early in the development cycle, it was decided that the
UH-60M FBW control system was to use the philosophy of
auto-moding. This was done to provide a consistent set of
response types and hold modes, without the need for pilot
selectable modes. Experiments performed in the
development simulator suggested that such an approach
reduced training time as well as reduced the possibility of
mode confusion. The approach presented one key challenge
– finding the right balance between high levels of
augmentation that are required in a DVE environment, while
allowing for agility needed in GVE.
Figure 3 shows the mode diagram for the UH-60M
FBW control system. The system introduces a concept of
“blended speed”. This speed is a combination of inertial
body velocities and forward air speed. Blended speed is the
primary mode trigger for the system. Other aircraft states
Figure 3 - UH-60M FBW Control Law Modes
such as main rotor speed, altitude, inertial sideslip and
inertial flight path are used as secondary mode triggers.
Position and/or rate of pilot controls are also used for
triggering certain mode transitions. A description of each
significant system mode is as follows:
Starting in hover, displacing cyclic control in any
direction produces an attitude command response type. If the
cyclic is released back to detent prior to the aircraft
exceeding 5 knots of blended speed, the aircraft decelerates
back to hover and position hold automatically engages. It
should be noted after several seconds of pure attitude
command response type, acceleration augmentation is
utilized to transform the vehicle response into acceleration
command. This combined response (Attitude / Acceleration
Command Hover Hold) provides augmentation similar to
classical Translational Rate Command (TRC) response type,
but avoids possible issues with the transition out of TRC at
higher velocities. The yaw axis in this portion of the flight
envelope provides heading rate command response type,
with heading hold automatically engaging once the pedals
are released to detent. The collective axis, through the use of
the active inceptor, provides a vertical velocity command
mode. Applying force to the collective controller results in
vertical velocity proportional to the amount and direction of
the applied force. Releasing force on the collective results in
the aircraft decelerating in the vertical axis and engagement
of hybrid altitude hold. The hybrid altitude hold mode
consists of an inner loop, holding inertial velocity and
inertially derived altitude and the outer loop provides radar
altitude influence. This mode assures good disturbance
rejection, while providing good ride quality and radar
altitude retention.
In the portion of the flight envelope above the 5 knot
blended speed window, but below high speed flight (defined
as 60 knots of blended speed) the pitch and roll axes of the
aircraft provide attitude / acceleration command (as
described above) with velocity hold automatically engaging,
when the cyclic is returned to detent. In the portion of
forward flight where inertial sideslip is below 10 degrees,
low speed turn coordination is provided. In low speed turn
coordination mode displacing roll cyclic results in attitude
command response type, with yaw axis automatically
commanding heading rate, such that the tail of the aircraft is
aligned with its flight path. The collective axis in this part of
the flight envelope provides flight path command response
type. In this mode, a prediction of the collective stick
position for level flight is used to provide a tactile detent.
Placing collective into the level flight detent automatically
engages hybrid altitude hold mode. This hybrid altitude hold
mode is similar to the mode used for near hover operations,
with the exception that barometric altitude retention can be
provided as well – depending on the actual value of blended
speed. Displacing the collective stick from the level flight
detent position results in an unaugmented aircraft response
in the vertical channel. Once the system detects that the
collective stick is no longer moving and it is outside of the
level flight position, a flight path hold mode engages. In this
mode, the aircraft holds the flight path angle acquired at the
time of mode engagement.
In the high speed portion of the envelope, the pitch axis
provides attitude / acceleration command response type,
with velocity hold engaging as the pitch cyclic is released to
detent. The roll axis provides attitude command / attitude
hold response type. This response type links position of the
roll cyclic control with the attitude of the aircraft. Displacing
the roll cyclic from detent results in an aircraft roll attitude
change proportional to the amount of cyclic stick
displacement; returning the roll cyclic back to detent results
in the aircraft returning to a wings level attitude. To allow
prolonged dwelling at attitudes other than level with
minimal pilot fatigue, a set of trim switches is used on the
cyclic grip to allow change in reference attitude without
displacing the control. The collective axis in forward flight
provides the same functionality as discussed above in the
low speed region. The functionality of pedals is altered in
this regime to provide a sideslip command response type.
Since the aircraft does not have a sideslip sensor, an estimate
is used. This estimate provides sideslip envelope limiting.
Such a limit is considered a “soft” limit. At full pedal
command, once the estimated sideslip reaches the sideslip
limits, the commanded yaw rate is reduced to only a few
degrees per second. This ensures that envelope exceedance
can not develop rapidly and allows ample time for the pilot
to recognize the condition and release the control.
Throughout the entire flight envelope the primary
controls are augmented by trim switches. The crew can at
any time adjust any of the aircraft hold states via a
corresponding trim switch (an example was discussed above
with respect to roll attitude). Ample feedback is provided to
the pilot via indications on the primary flight displays in
regards to the values of the aircraft references. Any time the
pilot is actively commanding a change to any of the
references, be it via the primary controls or trim switches, a
predicted target value is displayed on the appropriate portion
of the primary flight display. Great care has been taken to
ensure that all dynamic responses are simple and predictable
– first order, well damped response has been designed for all
modes.
Additional emergency modes are active throughout the
entire envelope as well. These modes monitor engine status
and main rotor speed. Upon detection of a single engine
failure and depending on aircraft state the system will
execute a single engine fly-away profile or slow down as
deemed appropriate. During this time, the system will
provide active rotor speed governing. At any time, the pilot
can override the system by simply displacing the cyclic from
detent or moving the collective controller outside of the rotor
governing detent. If a dual engine failure or a rapid decrease
in rotor speed is detected, the system enters autorotation
mode. In this mode, the aircraft will automatically govern
main rotor speed via collective by providing a rotor
governing detent and adjusting it as needed.
SYSTEM DEVELOPMENT
Control Law Architecture
To implement the features described in the previous
section, a model-following control law architecture was
selected. Figure 4 shows a simplified view of the modelfollowing architecture. Pilot inputs are transformed by the
command model D(S) into desired commands in terms of
aircraft state. These desired commands are sent through the
feed-forward and feedback paths. The feed-forward path
consists of the inverse aircraft dynamics P -1(S), which
transforms desired aircraft commands into control positions
that achieve these desired commands. The feedback path
consists of several controllers C(S). The controllers provide
long term regulation and disturbance rejection for
commanded aircraft state. The outputs of both of these paths
are summed and are sent through mixing, kinematics and
control limiting algorithms M(S). Vehicle specific
parameters in each of the control law blocks are scheduled
with airspeed, as airspeed has a profound effect on vehicle
dynamics.
GenHel Simulation
Sikorsky‟s General Helicopter Flight Dynamics
Simulation (GenHel) [4] was used extensively throughout
Figure 4 – Model-Following Architecture
the UH-60M FBW development process. GenHel was used
to obtain stability and control derivatives for the inverse
plant. It was also used to predict stability margins and
specification compliance for every major software version
update. In addition, GenHel was used to stimulate hardwarein-the-loop testing and piloted evaluations in the FCSIL.
To ensure the validity of the simulation model and
adequacy for piloted evaluations, extensive correlation
activities against available flight test data was undertaken.
Flight test data for the Primary Mission and External Lift
configurations were used to assess GenHel‟s capabilities
against trim characteristics, bandwidth and stability
characteristics, as well as control power and dynamic
response using stick pulses.
Hover IGE
Hover IGE
Hover IGE
Hover IGE
Hover IGE
Hover IGE
Figure 5 - Primary Mission Configuration Level
Flight Control Positions
Hover IGE
Figure 6 - Primary Mission Configuration Level Flight
Attitudes / Engine Torque
Figure 7 - Primary Mission Configuration Roll Stick
Sweep at 120 kts
Typical level flight trim comparisons for the Primary
Mission Configuration are shown in Figures 5 and 6. Figure
5 shows the comparison between the longitudinal, lateral
cyclic, main rotor collective, and tail rotor pitch. Figure 6
shows the pitch and roll attitude, sideslip, and engine torque
requirements as a function of forward speed.
In order to ensure accurate predictions of stability
margins and bandwidth, frequency data from the GenHel
simulation was correlated against frequency data from flight
test. Figure 7 shows an overlay of a frequency response plot
of roll attitude to a roll stick sweep at 120 knots for the
Primary Mission Configuration. The low frequency (1 to 20
rad/sec) match is very good which is the area of interest for
bandwidth data and control law development.
To assess GenHel‟s control power and dynamic
response, stick pulses were compared to flight test by
playing back into the simulation the pilot control inputs
recorded during test. A hover pitch pulse comparison is
shown in Figure 8 comparing the longitudinal stick pitch
pulse, pitch attitude and pitch rate. Shown in Figure 9 is a
lateral cyclic roll pulse at 120 knots forward speed. Again
the figure shows the comparison to the stick input, the roll
attitude and roll rate response.
Based on extensive flight test correlation and good
agreement of the simulation with flight test data, it was
determined that GenHel was a valid tool for FBW
development, margin prediction, specification compliance,
and pilot evaluations.
Figure 8 - Pitch Pulse Flight Test Correlation at
Hover (PMC)
Control Law Parameterization
The process of choosing parameters for the control laws
started with the generation of a family of linear models of
the aircraft. The GenHel simulation described above was
used to generate all linear models used during this process.
The family of linear models spanned the entire flight
envelope and covered primary mission and typical external
load configurations. All linear models were validated against
non-linear GenHel simulations as well as existing UH-60
Figure 9 - Roll Pulse Flight Test Correlation at 120
kts (PMC)
flight test data. As the program progressed, linear model
correlation and verification continued. Figure 10 shows an
example of the roll axis agreement between the models. In
this case a frequency response between GenHel, the linear
model and flight test data gathered on the prototype aircraft
was compared. Similar correlation was performed for every
axis, at every every distinct flight condition.
The initial step in control law parametrization consisted
of inverse plant parameter selection. A sensitivity study was
performed to ascertain the need for inverse plant parameter
scheduling. It was determined that the inverse plant
parameters needed to be scheduled with airspeed every 40
knots. The inverse plant parameters were picked from plant
and control matrices of linear models for each flight
condition. Correctness of the selected parameters was
verified using the non-linear simulation (GenHel), and at a
later stage, flight test data. Figure 11 shows a comparison
between the roll axis inverse plant at 120 knots and flight
test data. Throughout the entire envelope a good set of
inverse plant parameters was obtained. As demonstrated in
the later sections of the paper, a careful selection and
validation of the inverse plant parameters resulted in
excellent performance of the model-following control laws.
The next step was to setup a complete linear analysis
environment with the control laws and the plant model in the
loop. For parameter optimization, an AFDD developed
CONDUIT® tool set was chosen. This toolset allows rapid
analysis and optimization of gains and other parameters in a
graphical user interface (GUI) driven, interactive
environment. An iterative process followed, where a
candidate gain set was selected using CONDUIT®, verified
using the non-linear GenHel simulation, and implemented in
real-time, pilot in-the-loop simulation. Piloted evaluations at
this point consisted of classical handling qualities
evaluations as well as evaluations of the suitability of a
particular control scheme to the mission of the aircraft.
Resulting feedback from the pilots was taken back to the
CONDUIT® environment and rolled into the design
specifications. At several points in this process, new
CONDUIT® specifications were created to cover a
particular point of pilot feedback. This design process
worked very well, and after a few iterations a candidate set
of control laws were created. Figure 12 shows typical
results, in this case the aircraft is at 19,300 lb, clean
configuration at hover. It should be noted that yaw axis
bandwidth is set to Level 2 intentionally. During testing,
pilot feedback indicated that Level 1 yaw bandwdith was too
aggressive for this aircraft and it‟s mission.
At this point in the design, it was decided to prove out
the candidate control laws on AFDD‟s UH-60 RASCAL
test bed. This test bed allows evaluation of candidate control
laws on a full authority, research FBW system in a safe and
monitored environment. A full description of this effort is
beyond the scope of this paper, as it was published at the
completion of RASCAL testing [2].
Control Law Mode control
Due to absence of pilot selectable modes, the logic for
mode control presented several unique challenges. First and
Figure 10 - Roll Axis Correlation (hover) between
linear model, GenHel and flight test data
Figure 11 - Roll Axis Inverse Plant verification (120
kts)
foremost was to ensure that mode switching occurs
predictably and with no perceivable transients. It was also
important to keep the mode logic testable from the software
and qualification perspective. For these reasons, use of state
machines to implement all control law logic was chosen.
Multiple state machines are used to implement the logic
suite for the control laws. Each state machine performs a
unique part of mode control and is testable as a stand alone
entity. This approach allowed consistent implementation of
the entire auto-mode logic suite as well as providing
predictable results.
Figure 12 - CONDUIT results, UH-60 FBW 19K/Clean Hover
SYSTEM TESTING
System testing consisted of several phases. Initial
testing was conducted in a desktop simulation environment.
Further testing was conducted in a laboratory environment,
and finally once the prototype aircraft and the system were
ready, flight test began. The desktop portion of system
testing consisted of control law parameter optimization
(described above) and test vector generation (beyond the
scope of this paper). The other phases of testing are
described below.
exercising the entire flight envelope in normal flight mode,
extensive failure mode testing was also conducted. This
testing included engine failures, tail rotor failures, stabilator
manual mode reversion, active inceptor failures, flight
control computer failures, and CLAW degraded mode
Laboratory testing
A full hardware-in-the loop laboratory was constructed
to house all aircraft equipment and the simulation. A wraparound visual system was installed as well as a
representative UH-60 cockpit.
The Flight Control System Integration Lab (FCSIL) was
used extensively throughout the UH-60M FBW CLAW
development process. Piloted evaluations were conducted
for every software change and 50 to 100 hours of piloted
simulation were conducted for each software version that
was flight tested. Handling qualities evaluations were
developed that exercised the entire operating envelope as
well as all CLAW mode transitions. This ensured a smooth
transition from the lab to the flight test article. In addition to
Figure 13 – UH-60M FBW First Flight
testing. The CLAW degraded mode testing including
extensive evaluations during aggressive maneuvers while
transitioning from full up FACS mode to various levels of
degraded CLAWs.
Flight test approach and results
Shakedown and development flight testing was
conducted from August 2008 through December 2011 by a
combined Sikorsky/Army test team at Sikorsky‟s
Development Flight Center in West Palm Beach, Florida.
Initial shakedown flight testing (Figure 13) was conducted to
open safe structural and dynamic flight envelopes to permit
evaluation of flight loads, vibrations, rotor stability,
performance, engine operating characteristics, and handling
qualities. Airspeed, altitude, load factor and aircraft
configuration were gradually increased to expand the
operating envelopes. Critical measurements were monitored
using telemetry for all flights. Handling qualities testing was
conducted to demonstrate Level 1 ADS-33E-PRF
compliance on the UH-60M FBW aircraft. ADS-33E-PRF
testing included Mission Task Elements (MTEs) in the
primary mission configuration and a 5,200 pound external
load configuration in both GVE and DVE. Additional flight
regimes evaluated in the primary mission configuration
included quantitative ADS-33E-PRF requirements, trimmed
level flight, yardwork at sea level and altitude, stability
margin verification, controllability, sustained turns and
recoveries, simulated engine malfunctions, autorotational
flares, and selected maneuvers with a 5,200 and 9,000 pound
external slung loads. Maneuvers were performed at the
historically most severe combinations of airspeed, gross
weight, center of gravity, and altitude to demonstrate worst
case conditions and limit the amount of required testing.
For the low speed MTEs the courses were setup along
the runway at the Development Flight Center. For DVE
flights it was arranged to have all facility lights shut off
between the hours of midnight and 5:00 am to ensure
minimal ambient light. Filters were still used in conjunction
with the night vision goggles to provide a low 2 Usable Cue
Environment (UCE). Infrared LED lights were used to
illuminate the pilot cues on the courses. The hover board
cues illuminated through night vision goggles is shown in
Figure 14. A minimum of three pilots conducted each MTE
and provided their comments and handling qualities ratings.
The GVE MTEs for the primary mission configuration were
conducted in moderate winds of 20 knots gusting to 28 to 30
knots. All other low speed MTEs were conducted in calm
winds. All MTEs were rated Level 1 in GVE and DVE for
both the primary mission and external load configurations.
Another positive feature of the UH-60M FBW is its
ability to hold an inertial reference position. The position
hold requirement in ADS-33E-PRF states the aircraft must
perform a 360 degree hover turn within 30 seconds and
maintain position within a 10-foot diameter circle. Figure 15
shows the results of this maneuver. The aircraft maintained
position within +/- 1 foot while completing the turn in less
than 27 seconds reaching yaw rates of 15 to 16
degrees/second.
The ADS-33 MTE‟s were designed as a qualitative tool
with specified and precisely defined criteria. The stringent
desired performance criteria were specifically designed to
drive pilot aggressiveness to expose weakness in the
aircraft/control laws especially with the pilot in the loop. As
Figure 14 – Precision hover DVE course cues
Figure 15 – Position hold performance during 360
degree turn
an example, the hover task desired tolerance requires a
deceleration from translating flight to a spot over the ground
with moderate aggressiveness to maintain position within +/3 feet. During practice runs of this maneuver it was
discovered that some of the pilots became very active in the
lateral axis during the deceleration and stabilization to a
hover which resulted in decreased ride quality as the pilot
made small amplitude inputs about the stick detents to
capture and maintain the precise hover position. It was
determined that as the pilot backed out of the loop to allow
the hold modes to become active, the aircraft drift was less
than 1.5 ft which was well within the desired task criteria of
maintaining position within 3 ft. The ADS-33 MTE‟s can be
an excellent indicator of an aircrafts performance in the real
world mission environment.
The UH-60M FBW aircraft performance during DVE
ADS-33 MTEs was singular. Altitude maintenance was the
most notable reduction in pilot workload during the hover
and pirouette maneuvers. The radar altitude hold worked
well enough during these tasks that the pilot typically made
no collective inputs during the execution of the task. The
high level of confidence in altitude maintenance during
traditionally high workload tasks in low speed flight applies
directly to real world missions like sling load hook up in a
dusty pick-up zone at night. The lack of sufficient visual
cues, combined with harsh environmental conditions has
contributed to the loss of many aircraft. The UH-60M FBW
flight control system can make these real world mission
tasks safer and at the same time significantly reduce the pilot
workload providing more capacity for situational awareness
to other hazards. The slalom mission task element relates
well to the real world task of nap of the earth flight in a
hostile environment requiring frequent course and heading
changes. The low speed control laws of the UH-60M FBW
will make low speed maneuvering in a narrow ravine in poor
weather conditions much less stressful as the low speed turn
coordination keeps the tail behind the nose reducing the need
for most pedal inputs from the pilot and reducing the
probability of inadvertent tail contact with terrain or
obstacles. The UH-60M FBW could be flown more
aggressively in low light conditions with greater certitude
than the baseline UH-60M in optimal conditions.
Figure 16 – Level 1 Qualitative Requirements
(MTEs)
CONCLUSIONS
Figures 16 and 17 show ADS-33 requirements that were
demonstrated by the two prototypes. Full ADS-33 Level 1
GVE/DVE compliance was shown. The only exception to
Level 1 requirements was yaw bandwidth. As discussed
above, it was set to Level 2 performance to address multiple
pilot comments.
The UH-60M FBW does not fly like a typical partial
authority, augmented rate command, mechanical control
system helicopter. It was designed to be easier to fly than
most helicopters due to the automatic hold mode in every
axis anytime the wheels are off the ground with the active
inceptors in detent. The most noticeable change from the
baseline UH-60M is the attitude command system. In the
low speed flight regime, cyclic displacement from the detent
commands a given attitude and in order to maintain the
desired attitude the cyclic must remain displaced at the same
position. In the roll axis, from level flight at 80KIAS to
initiate a 30 deg angle of bank turn for 180 degrees of
heading change, the pilot must displace the lateral cyclic to
command the 30 degrees angle of bank and maintain that
same displacement throughout the turn returning the lateral
Figure 17 – Level 1 Quantitative Requirements
stick to the detent just prior to reaching the desired heading
resulting in the roll attitude returned to zero. Though this
was initially awkward to most pilots when they began flying
the UH-60M FBW, it became very intuitive after only a few
minutes of flight. All of the experimental test pilots on the
UH-60M FBW program were high time flight test pilots on
the UH-60L and baseline UH-60M aircraft concurrent with
the UH-60M FBW test program. None of the pilots had a
problem making the switch from one aircraft series to the
other even though there were stark differences in the levels
of automation from a UH-60L to the UH-60M FBW. For
many pilots, often the most difficult transition was
remembering to back out of the loop to let the automatic
features reduce the workload. In the low speed portion of
the flight envelope the UH-60MU was overall very easy to
fly to tight tolerances in heading, altitude, and speed. The
handling qualities in the degraded visual environment in
hover and low speed flight regimes were superior to the
current utility fleet.
Lessons Learned
A multitude of lessons were learned during the
development of this system, due to space and scope
limitations, key lessons learned are discussed below.
During CLAW development activities in the FCSIL,
pilots tend to prefer significantly higher bandwidths than is
warranted on the actual aircraft. The aircraft always feels
sportier than the simulation. Caution should be taken not to
rely on piloted simulation for setting rate schedules,
acceleration limits, bandwidth, or beep rates.
Flight control characteristics are extremely important to
the handling qualities of the aircraft. The method used for
tuning the active inceptors‟ force feel was at times
haphazard making several adjustments based solely on pilot
comments and then stacking several adjustments into a
single change. The method for flight control mechanical
characteristics must be deliberate and methodical. Having
the capability to tune the parameters during flight within a
specific range of variables was very helpful in the end.
Higher levels of automation do not necessarily equate to
“happy pilots”. The UH-60M FBW was designed so that the
upper level flight control modes were active anytime the
aircraft was flying with the sticks in the detent. The pilots
were often troubled when the “system” performance
degraded as the pilot entered the loop. When designing a
piloted vehicle, it should be able to have the pilot in the loop
or out of the loop with no significant change in performance.
More mission representative testing needs to be
conducted throughout a developmental flight test program to
help identify strengths, weaknesses, and flaws in the system
early. An excellent example of this lesson was discovered in
2009 during a photo opportunity when both of the UH-60M
FBW test articles were launched simultaneous to fly close
formation.
The pilot in the trail aircraft reported
significantly higher workload than when flying formation in
the UH-60L because of the always-armed hold modes that
became active every time the sticks were in the detent. The
trail aircraft pilot noted that he spent a great deal of energy
constantly moving the cyclic (“dithering”) about the detent
to avoid activating the pitch and roll attitude hold modes
(think of using your car‟s cruise control on a congested
highway).
A functional, fixed-based simulator in close proximity
to the developmental flight test location is highly desirable.
There were many times during flight test when the pilots
noted anomalies that were in fact a part of the system design.
False stops due to perceived issues could have been
significantly reduced if the pilots and engineers could have
efficiently rehearsed in a hardware-in-the loop fixed-based
simulator.
ACKNOWLEDGMENTS
The success of the UH-60M FBW project is due to the
outstanding efforts of many key contributors beyond the
authors of this paper. In particular, the authors would like to
acknowledge the significant contributions, guidance and
support received from Mr. David Arterburn whose vision
was to have a completely automated control system, Dr.
Chris Blanken, Dr. Mark Tischler, Mr. Frank Luria, Mr.
Jeffrey Harding, and the late Mr. Doug Kinkead whose
career at Sikorsky Aircraft was spent for the betterment of
the H-60 product line. In addition, the authors would like to
thank the pilots from Sikorsky Aircraft and the U.S. Army
ATTC who participated in the UH-60M FBW flight test
program.
REFERENCES
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[3] Anonymous,
“Aeronautical Design Standard,
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