U.S. patent application number 11/505235 was filed with the patent office on 2008-10-02 for compound aircraft control system and method.
This patent application is currently assigned to Piasecki Aircraft Corporation. Invention is credited to Andrew S. Greenjack, Joseph F. Horn, Frank N. Piasecki.
Application Number | 20080237392 11/505235 |
Document ID | / |
Family ID | 39792552 |
Filed Date | 2008-10-02 |
United States Patent
Application |
20080237392 |
Kind Code |
A1 |
Piasecki; Frank N. ; et
al. |
October 2, 2008 |
COMPOUND AIRCRAFT CONTROL SYSTEM AND METHOD
Abstract
The Invention is a control system for a compound aircraft. A
compound aircraft has features of both a helicopter and a fixed
wing aircraft and provides redundant control options. The control
system allows an authorized person to select any of plurality of
control biases each which is designed to achieve an overall
operational objective. The control system applies the selected
control bias in allocating the control function among the redundant
control options.
Inventors: |
Piasecki; Frank N.;
(Haverford, PA) ; Greenjack; Andrew S.; (Thornton,
PA) ; Horn; Joseph F.; (State College, PA) |
Correspondence
Address: |
LIPTON, WEINBERGER & HUSICK
P.O. Box 203
Exton
PA
19341
US
|
Assignee: |
Piasecki Aircraft
Corporation
Essington
PA
|
Family ID: |
39792552 |
Appl. No.: |
11/505235 |
Filed: |
August 16, 2006 |
Current U.S.
Class: |
244/6 |
Current CPC
Class: |
G05D 1/0858 20130101;
B64C 2027/8236 20130101; B64C 27/82 20130101; B64C 2027/829
20130101; B64C 27/26 20130101 |
Class at
Publication: |
244/6 |
International
Class: |
B64C 27/22 20060101
B64C027/22 |
Claims
1.-4. (canceled)
5. A compound aircraft, said compound aircraft comprising: a. a
fuselage; b. a main rotor rotatably connected to said fuselage,
said main rotor being configured to exert a main rotor lifting
force to said fuselage; c. a wing connected to said fuselage, said
wing being configured to exert a wing lifting force to said
fuselage, said main rotor and said wing in combination being
configured to exert a total lifting force to said fuselage; d.
means to select an operational objective by a user of the compound
aircraft from among a plurality of operational objectives; e. an
inceptor, said inceptor being configured to receive a command for a
trimmed flight; f. a microprocessor, said microprocessor being
operably connected to said main rotor and to said wing, said
microprocessor being operably connected to said means to select
said operational objective and to said inceptor, said
microprocessor being configured to receive said selected
operational objective from said means to select said operational
objective, said microprocessor being configured to receive said
command for said trimmed flight from said inceptor, g. a computer
memory operably connected to said microprocessor, said computer
memory storing a database accessible to said microprocessor, said
database including a plurality of combinations of trim settings,
each of said plurality of combinations of trim settings being
configured to achieve said command for said trimmed flight when
applied by said microprocessor, a one of said plurality of
combinations of trim settings corresponding to said selected
operational objective, said microprocessor being configured to
select said one of said plurality of combinations of trim settings
corresponding to said selected operational objective, said
microprocessor being programmed to allocate said total lifting
force between said rotor lifting force and said wing lifting force
for said trimmed flight, said microprocessor being programmed to
base said allocation of said total lifting force for said trimmed
flight upon said selected one of said plurality of combinations of
trim settings.
6. The compound aircraft of claim 5, the compound aircraft further
comprising: a sensor, said sensor being operably connected to said
microprocessor, said sensor being configured to detect an aircraft
condition, said sensor being configured to communicate said
aircraft condition to said microprocessor, said microprocessor
being programmed to base said allocation of said total lifting
force between said main rotor lifting force and said wing lifting
force based upon said aircraft condition.
7.-8. (canceled)
9. The compound aircraft of claim 5 wherein a. said main rotor is
configured to selectably apply a rotor rolling moment to said
fuselage, said rotor rolling moment being variable; b. said wing is
configured to selectably apply a wing rolling moment to said
fuselage, said wing rolling moment being variable, said main rotor
and said wing in combination exerting a total rolling moment on
said fuselage; c. said microprocessor is programmed to allocate
said total rotting moment between said main rotor rolling moment
and said wing rolling moment for said trimmed flight, said
microprocessor being programmed to base said allocation of said
total rolling moment for said trimmed flight upon said selected one
of said plurality of combinations of trim settings.
10. The compound aircraft of claim 9 wherein said main rotor is
configured to provide a main rotor forward thrust, said main rotor
forward thrust being variable, the compound aircraft further
comprising: a. means for providing a non-rotor forward thrust, said
non-rotor thrust being variable, said non-rotor forward thrust and
said main rotor forward thrust in combination defining a total
forward thrust; b. said microprocessor being operably connected to
said means to provide said non-rotor forward thrust, said
microprocessor being programmed to allocate said total forward
thrust between said main rotor forward thrust and said non-rotor
forward thrust for said trimmed flight, said microprocessor being
programmed to base said allocation of said total thrust for said
trimmed flight upon said selected one of said plurality of
combinations of trim settings.
11. The compound aircraft of claim 10 wherein said main rotor has a
collective setting and a cyclic setting, said selected combination
of said trim control settings including said collective setting and
said cyclic setting and wherein said configuration of said wing to
provide said wing lifting force comprises: a flap operably
connected to said wing, said flap having a variable angle with
respect to said wing, said microprocessor being programmed to
select said angle of said flap, said angle of said flap being
defined by said selected combination of trim control settings.
12. The compound aircraft of claim 9 wherein said adaptation of
said wing to provide said wing rolling moment comprising: a pair of
ailerons operably connected to said wing, said pair of ailerons
being adapted to have differential positions, said microprocessor
being configured to select said differential positions of said pair
of ailerons, said differential positions of said pair of ailerons
being defined by said selected combination of trim control
settings.
13. The compound aircraft of claim 12 wherein said pair of ailerons
is a pair of flaperons.
14. The compound aircraft of claim 10 wherein said means to provide
non-rotor forward thrust comprises: a propeller having a variable
propeller pitch, said propeller being rotatably attached to said
fuselage, said microprocessor being operably connected to said
propeller, said microprocessor being programmed to select said
propeller pitch, said propeller pitch being defined by said
selected combination of trim control settings.
15. The compound aircraft of claim 14, the compound aircraft
further comprising: a. a rudder connected to said fuselage, said
rudder being configured to impart a variable rudder yaw moment to
said fuselage, said configuration of said rudder to impart said
variable yaw moment comprising said rudder having a variable rudder
angle with respect to said fuselage, said propeller being adapted
for rotation and adapted to generate a stream of air when said
propeller is rotating, said rudder being located within said stream
of air generated by said propeller when said propeller is rotating;
b. a sector operably connected to said fuselage, said sector being
configured to impart a variable sector yaw moment to said fuselage,
said configuration of said sector to impart said sector yaw moment
comprising said sector being configured to variably receive said
stream of said air when said propeller is rotating, said sector
being adapted to variably redirect said stream of said air, said
variable rudder yaw moment and said variable sector yaw moment in
combination defining a total yaw moment; c. said sector and said
rudder being operably connected to said microprocessor, said
microprocessor being programmed to allocate said total yaw moment
between said rudder yaw moment and said sector yaw moment for said
trimmed flight, said microprocessor being programmed to base said
allocation of said total yaw moment for said trimmed flight upon
said selected one of said plurality of combinations of trim
settings.
16.-31. (canceled)
32. The compound aircraft of claim 6 wherein said command for said
trimmed flight is a one of a plurality of possible commands for
said trimmed flight and wherein said plurality of combinations of
trim settings is a one of a plurality of pluralities of
combinations of trim settings, each of said plurality of
pluralities of combinations of trim settings corresponding to a one
of said plurality of operational objectives.
33. A compound aircraft, said compound aircraft comprising: a. a
fuselage; b. a main rotor rotatably connected to said fuselage,
said main rotor being configured to exert a main rotor lifting
force to said fuselage; c. a wing connected to said fuselage, said
wing being configured to exert a wing lifting force to said
fuselage, said main rotor and said wing in combination being
configured to exert a total lifting force to said fuselage; d.
means to select an operational objective by a user of the compound
aircraft from among a plurality of operational objectives; e. an
inceptor, said inceptor being configured to receive a command for a
maneuvering flight; f. a microprocessor, said microprocessor being
operably connected to said main rotor and to said wing, said
microprocessor being operably connected to said means to select
said operational objective and to said inceptor, said
microprocessor being configured to receive said selected
operational objective from said means to select said operational
objective, said microprocessor being configured to receive said
command for said maneuvering flight from said inceptor; g. a
computer memory operably connected to said microprocessor, said
computer memory storing a database accessible to said
microprocessor, said database including a plurality of combinations
of weighting factors, each of said plurality of combinations of
weighting factors being configured to achieve said command for said
maneuvering flight when applied by said microprocessor, a one of
said plurality of combinations of weighting factors corresponding
to said selected operational objective, said microprocessor being
configured to select said one of said plurality of combinations of
weighting factors corresponding to said selected operational
objective, said microprocessor being programmed to allocate said
total lifting force between said rotor lifting force and said wing
lifting force for said maneuvering flight, said microprocessor
being programmed to base said allocation of said total Lifting
force for said maneuvering flight upon said selected one of said
plurality of combinations of weighting factors.
34. The compound aircraft of claim 33 wherein said command for
maneuvering flight is a one of a plurality of possible commands for
maneuvering flight and wherein said plurality of combinations of
weighting factors is a one of a plurality of pluralities of
combinations of weighting factors.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] The Invention is a control system for a compound aircraft.
For trimmed flight, the control system selects a combination of
trim control settings for the redundant controls of the compound
aircraft to achieve a pilot's command consistent with a
user-selectable objective, such as speed maximization, fuel
consumption minimization, vibration reduction or lifecycle cost
reduction. For maneuvering flight, the control system distributes
control among the redundant control options of the compound
aircraft and may perform that distribution consistent with the
user-selectable objective. The Invention is also a method for
controlling a compound aircraft.
[0003] 2. Description of the Related Art
[0004] A `compound` aircraft is an aircraft that includes features
of both fixed wing aircraft and rotary wing aircraft. The compound
aircraft includes the elements of a helicopter, including at least
one main rotor and a mechanism to overcome the torque response of
the rotating main rotor. The compound aircraft also includes
elements of a fixed-wing aircraft, such as a wing. The wing may be
equipped with ailerons, flaps or a combination of flaps and
ailerons known as `flaperons.` The compound aircraft may be
equipped with a separate thrust mechanism to drive the aircraft
forward, such as a propeller in a ducted fan. Through the use of
appropriate vanes or sectors that change the configuration of the
duct, the ducted fan may serve as the mechanism to overcome the
torque response of the rotating rotor blades and to provide yaw
control.
[0005] A compound aircraft offers several advantages over a
conventional helicopter. Those advantages include achieving higher
flight speeds and delayed onset of retreating blade stall and
leading blade compression effects. Although the advantages of a
compound helicopter are well known, no compound helicopters have
been placed in regular operation in commercial or military fleets.
One reason is the control complexity of the compound aircraft.
[0006] The pilot of a conventional helicopter has only limited
controls. The controls available for a conventional helicopter
having a single main rotor and a tail rotor are:
[0007] Throttle--The pilot can control the amount of power supplied
to the rotor blades and to the tail rotor.
[0008] Collective pitch--The pilot contemporaneously can change the
pitch of all main rotor blades by an equal amount using the
collective pitch control, also known as the
`collective.`Contemporaneously changing the pitch angle of all main
rotor blades increases or decreases the lift supporting the
helicopter. Increasing the collective and the power will cause the
helicopter to rise. Decreasing the collective and the power will
call the helicopter to sink.
[0009] Cyclic pitch--The pilot may use the cyclic pitch control,
also known as the `cyclic,` to cause the pitch angle of the main
rotor blades to change differentially as the main rotor rotates
through 360 degrees. The cyclic pitch control is used to control
the pitch and roll of the helicopter. For example, increasing the
pitch angle of a rotor blade when the rotor blade is retreating
toward the rear of the helicopter and decreasing the pitch angle
when the rotor blade is advancing toward the front of the
helicopter will cause the main rotor plane of rotation to tilt
forward and hence will cause the helicopter to move forward.
[0010] Tail rotor pitch control--For a conventional helicopter
having a tail rotor mounted on a boom, a pedal-operated yaw control
changes the pitch of the tail rotor blades so that the tail rotor
presents more or less force countering the torque response of the
rotating main rotor. The pitch of the tail rotor blades therefore
controls the yaw of the conventional helicopter.
[0011] For a conventional helicopter and for a particular throttle
setting, there is only one combination of trim control settings for
the collective, cyclic and tail rotor pitch controls to achieve any
particular desired trimmed condition of the helicopter. The pilot
of the conventional helicopter therefore has few control
options.
[0012] A compound aircraft will have the aforementioned controls
and in addition will have other controls. For example, the compound
aircraft may feature the following controls:
[0013] Flaperon controls--The flaperons (a combination of flaps and
ailerons) are located on the wings. When deflected differentially
like ailerons, the flaperons may cause the aircraft to roll. When
deflected in unison like flaps, the flaperons may increase or
decrease lift generated by the wing. In hovering flight, the
flaperons may be deployed to reduce the effective wing area and
hence reduce the downward force on the wings from the downwash of
the main rotor.
[0014] Forward thrust control--The compound aircraft may be
equipped with a ducted fan or other mechanism to provide forward
thrust. Thrust provided by the ducted fan or by another mechanism
that is not the main rotor is referred to in this application as
"non-rotor forward thrust."
[0015] Rudder/stabililator--The compound aircraft may be equipped
with a rudder and with an elevator or stabilator. The rudder
controls the yaw of the aircraft, in cooperation with the tail
rotor, ducted fan, or other mechanism countering the torque
reaction of the rotating main rotor. An elevator or stabilator
controls the pitch of the compound aircraft, in cooperation with or
instead of the cyclic pitch control.
[0016] The pilot of the compound aircraft is presented with a
variety of control combinations to achieve a desired flight
condition. For example, if the pilot desires to increase the
forward speed of the compound aircraft, the pilot can increase the
non-rotor forward thrust using the forward thrust control, can use
the cyclic pitch, stabilator and throttle controls to pitch the
aircraft forward, or can use any combination of forward thrust
control, stabilator, cyclic pitch control and throttle. Each of the
possible combinations of trim control settings offers advantages
and disadvantages. A combination of trim control settings that is
optimal for one objective (for example, minimizing fuel
consumption) may not be optimal for another objective (for example,
minimizing vibration).
[0017] Only one combination of trim control settings for the
compound aircraft will be optimal for achieving a particular
trimmed condition or for implementing maneuvering flight commands
consonant with also achieving a particular operational objective.
The prior art does not disclose a control system for a compound
aircraft that allows selection among a plurality of objectives and
that then automatically optimizes control settings to achieve pilot
control commands consistent with the selected objective.
BRIEF DESCRIPTION OF THE INVENTION
[0018] The Invention is a control system for a compound aircraft. A
user selects an overall objective for the control system, such as
reducing vibration, increasing performance and speed, reducing
lifecycle costs, reducing loading of one or more components,
reducing fuel consumption, or any combination of these objectives
or of any other desired objectives. The overall objective is
pre-selected from among a plurality of overall objectives by the
pilot or by another authorized person; for example, by the owner of
the compound aircraft. The control system receives a command from a
pilot for trimmed flight. The control system also receives
information from sensors relating to current aircraft condition
(such as attitude, altitude, vertical speed, airspeed, main rotor
speed, control surface positions, acceleration and angular
rates).
[0019] The control system compares the pilot command to the current
aircraft condition as detected by the sensors and consults a
look-up database of combinations of trim control settings. The
control system applies the user-selectable overall objective in
consulting the look-up database. The control system selects one of
the combinations of trim control settings for trimmed flight from
the look-up database. The selected combination of trim control
settings provides a control setting for each of the various control
effectors of the compound aircraft to achieve the pilot's intended
trimmed flight condition consistent with the pre-selected overall
objective. As used in herein, the term `control effector` means
collectively all of the various flight control surfaces and engines
of the compound aircraft. The control system applies the selected
combination of trim control settings to the control effectors of
the compound aircraft, including the redundant control
effectors.
[0020] The sensors monitor the current condition of the aircraft
and provide constant feedback to the control system. The control
system continuously selects and applies different combinations of
trim control settings from the look-up database as needed to
achieve the selected overall objective for trimmed flight. If the
pilot control inceptors are in `detent,` which is a neutral
position that does not indicate a commanded change in aircraft
condition, a feedback controller regulates the aircraft control
effectors so that the aircraft stays in trim and the selected
overall objective is achieved.
[0021] In the event the pilot maneuvers the aircraft, the control
system will receive a pilot command from a control inceptor
operated by the pilot and will filter the pilot command using a
`command filter` to determine the commanded change in aircraft
condition. The command filter determines the dynamic response and
thus the handling qualities of the compound aircraft. The control
system compares the filtered pilot command to the condition of the
aircraft as detected by the sensors and selects a combination of
control effector settings to achieve the maneuver.
[0022] The control system applies `weighting factors` to control
the distribution of control among the redundant control effectors
in maneuvering flight. The control designer can select a
combination of weighting factors to achieve an overall objective
for maneuvering flight, such as minimization of certain structural
loads. An authorized person, such as the pilot or owner of the
aircraft, may select an overall objective for maneuvering flight
from among a plurality of objectives. The selected overall
objective for maneuvering flight may be the same as or different
from the selected overall objective for trimmed flight. The control
system may consult a look-up database and select a combination of
weighting factors associated with the selected overall objective
for maneuvering flight. The control system applies the selected
combination of weighting factors in allocating control among the
redundant control surfaces. The control system will supply the
selected control settings to the appropriate actuators to achieve
the maneuver. When the control inceptors are returned to detent,
the aircraft will once again reach trim, with the appropriate
control settings to achieve the selected overall objective for
trimmed flight.
[0023] The pilot may change the overall objective for trimmed or
maneuvering flight and hence the applied control trim settings or
weighting factors during flight. For example, the pilot may change
the overall objective for trimmed flight from `reduce vibration` to
`maximize speed.` The control system then will select a different
combination of trim control settings to accomplish the new overall
objective for trimmed flight. Alternatively, the pilot may not be
authorized to change the overall objective for trimmed or
maneuvering flight and the function of selecting the overall
objective may be reserved to another person, such as the owner of
the aircraft.
[0024] In an important application of the Invention, a pilot will
fly a compound aircraft using only the familiar helicopter flight
controls of collective pitch, cyclic pitch and tail rotor pitch
(pedal yaw control), just as if the pilot were flying a
conventional helicopter. The control system receives the
collective, cyclic and tail rotor inputs from the pilot and infers
the intent of the pilot. The control system then selects an
appropriate combination of trim control settings for the
collective, cyclic, flaperon, forward thrust, elevator, sector,
rudder and any other available control effector to best achieve the
pilot's intent, consistent with the pre-selected overall objective
for trimmed or maneuvering flight. A pilot skilled in flying a
conventional helicopter may therefore pilot a compound aircraft
using the control system of the Invention and achieve the selected
overall objective without simultaneously applying the skills of a
fixed-wing pilot.
[0025] The control system of the Invention may be a component of a
fully authorized "fly-by-wire" system in which the control system
operates all aircraft flight controls. Alternatively, the control
system of the Invention may be configured to operate only a portion
of the controls of the aircraft. For example, the pilot may
directly control the collective, cyclic pitch and throttle
controls, while the control system of the Invention automatically
controls the flaperons and forward thruster.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] FIG. 1 is a perspective view of a compound aircraft.
[0027] FIG. 2 is a side view of a compound aircraft.
[0028] FIG. 3 is a rear view of a compound aircraft.
[0029] FIG. 4 is a schematic representation of the control system
of the invention.
[0030] FIG. 5 is a schematic representation of information flow
through the control system of the Invention.
[0031] FIG. 6 illustrates the overall control system
architecture.
[0032] FIG. 7 illustrates the longitudinal/vertical control
subsystem.
[0033] FIG. 8 illustrates the speed command/speed hold
subsystem.
[0034] FIG. 9 illustrates the engine torque limiting module
[0035] FIG. 10 illustrates the lateral/directional control
subsystem.
[0036] FIG. 11 illustrates the turn coordination/yaw rate command
module.
DESCRIPTION OF AN EMBODIMENT
A. Compound Aircraft Features
[0037] The apparatus of the Invention is a control system for a
compound aircraft 2. As shown by FIGS. 1 and 2, the compound
aircraft 2 includes features of both a helicopter and a fixed wing
aircraft. Those features include a fuselage 4, a main rotor (or
rotating wing) 6, a hub 8 about which the main rotor 6 rotates and
wings 10. Rotation of main rotor 6 about hub 8 induces main rotor
lift 12. Movement of air across wings 10 in response to the forward
motion 14 of the compound aircraft 2 generates wing lift 16. Rotor
lift 12 and wing lift 16 provide lift to the compound aircraft
2.
[0038] Wings 10 feature a wing control surface known as a
`flaperon` 18. Flaperon 18 may be moved differentially, in which
event flaperons 18 act as ailerons. When used as ailerons, the
flaperons 18 in conjunction with wings 10 impart a rolling moment
to fuselage 4. The flaperons 18 also may be moved in unison, in
which event the flaperons 18 act as flaps. When used as flaps,
flaperons 18 change the aerodynamic characteristics of the wing 10
and change wing lift 16.
[0039] FIG. 3 is a rear view of the compound aircraft 2. As shown
by FIGS. 1, 2 and 3, the tail of the compound aircraft 2 features a
forward thruster 20. Forward thruster 20 is preferably a ducted fan
22. Ducted fan 22 features a shroud 24. Shroud 24 improves safety
and reduces the likelihood of damage to the propeller 26 resulting
from contact between the propeller 26 and the ground.
[0040] Propeller 26 rotates about a ducted fan axis of rotation 28,
which is generally parallel to the forward direction 14 of compound
aircraft 2. Propeller 26 is directly connected to the drive system
for main rotor 6, and so the speed of rotation of propeller 26 is
directly proportional to the speed of rotation of main rotor 6 and
is not independently controllable. The pitch of propeller 26 is
variable, allowing adjustment of the amount of thrust provided by
ducted fan 22.
[0041] Sectors 30, shown by FIG. 3, form an adjustable segmented
duct to selectably change the direction of thrust of ducted fan 22.
Sectors 30, in conjunction with rudder 32, serve to selectably
direct the thrust of ducted fan 22 to apply a torque to fuselage 4
contrary to the torque applied by main rotor 6. FIG. 3 shows the
sectors 30 in a deployed position and ready to direct ducted fan 22
thrust to counter the torque of the main rotor 6.
[0042] Rudder 32 is adapted to cooperate with sectors 30 to control
the direction of thrust of ducted fan 22. Rudder 32 is in the air
stream of fan rotor 26 and therefore is capable of affecting yaw of
the compound aircraft 2 at any speed.
[0043] Elevator 34 corresponds to the elevator of a fixed wing
aircraft. Elevator 34 is in the air stream of fan rotor 26 and is
capable of affecting the pitch of the compound aircraft 2 at any
speed.
B. Control System Overview
[0044] FIG. 4 describes the operational relationship between the
physical components of the control system of the Invention. A pilot
operates control inceptors 36. The control inceptors 36 correspond
to the flight controls of a conventional helicopter, with a
collective pitch control, a cyclic pitch control, pedal yaw control
and throttle. A person skilled at flying a helicopter therefore may
operate the compound aircraft 2 without simultaneously applying the
skills of a fixed-wing pilot. The control inceptors 36 are
connected to a microprocessor 38.
[0045] Sensors 40 monitor the condition of the compound aircraft 2
and are connected to the microprocessor 38. The sensors 40 may
monitor compound aircraft 2 variables such as airspeed, weight and
balance parameters, ambient atmospheric conditions, engine torque,
propeller 26 torque, vertical speed, pitch rate and attitude, roll
rate and attitude, and yaw rate.
[0046] Microprocessor 38 is operably connected to actuators for
each of the control effectors of the compound aircraft 2. Those
control actuators include the cyclic pitch actuator 42, the
collective pitch actuator 44, the throttle actuator 46, the first
and second flaperon actuators 48, 50, the elevator actuator 52, the
rudder actuator 54, the sector actuator 56 and the propeller pitch
actuator 58. Each of the actuators is adapted by conventional means
to operate its associated control effector under the command of the
microprocessor 38.
[0047] Computer memory 60 is connected to microprocessor 38.
Computer memory 60 includes a plurality of selectable overall
operational objectives for the compound aircraft 2. Memory 60 also
contains a trim schedule comprising a look-up database of
combinations of trim control settings selected to achieve each of
the selectable overall operational objectives for any given
condition of the aircraft and any given pilot command. The
microprocessor 38 consults the database and selects the combination
of trim control settings applicable to the condition of the
aircraft, the pilot command, and the selected overall operational
objective. The microprocessor constantly updates the selection of
the combination of trim control settings based on feedback from the
sensors detecting the changing aircraft conditions.
[0048] The microprocessor applies the selected combination of trim
control settings for trimmed flight. "Trimmed flight" includes
coordinated, level flight and also may include a steady climb or
descent or a coordinated turn.
[0049] The control system 62 of the Invention uses a "unique trim"
concept; that is, when the pilot places the control inceptors in a
neutral, or detent, position, the control system 62 of the compound
aircraft 2 automatically goes to a trimmed flight condition. The
way in which the compound aircraft 2 is trimmed is selected by the
control system 62 of the Invention based upon the selected overall
operational objective and based upon the condition of the aircraft
2 as detected by the sensors.
[0050] The selected combination of trim control settings defines
the control effector positions and trimmed attitude of the compound
aircraft 2 required to achieve the optimal trim to accomplish an
overall operational objective. Sensors 40 determine the deviation
by the compound aircraft 2 from the selected combination of trim
control settings. Feedback paths in the control system 62 can add
or subtract to the effector positions or attitude commands. Since
the compound aircraft 2 is closed-loop stable, over time if the
pilot keeps the control inceptors in detent, the aircraft 2 will
eventually settle into a trimmed condition very close to the
optimal trim.
C. Control System Information Flow
[0051] FIG. 5 provides an overview of the flow of information
through the control system 62 and the principal functions of the
control system 62. The control system 62 receives a pilot command
64. The pilot command 64 is generated by the control inceptors 36
of FIG. 4 and represents an instruction by the pilot of the
compound aircraft 2. The control system 62 also receives a variety
of aircraft condition information 66. The aircraft condition
information 66 is generated by sensors 40 as shown by FIG. 4 and
provides the microprocessor 38 with the status of the compound
aircraft 2.
[0052] The control system 62 examines the sensor information 66 to
determine the condition of the compound aircraft 2 and to evaluate
the pilot command 64 to determine how the condition of the aircraft
will be affected by the pilot command 64. The control system 62 is
fully authorized to operate all of the control effectors of the
compound aircraft 2. The control system 62 includes subsystems for
selecting optimal trim 68, speed command/speed holding 70,
longitudinal/vertical control 72, lateral/directional control 74
and turn coordination 76. The control system 62 determines the
appropriate subsystem to apply based on the pilot command 64 and
aircraft condition information 66 received. The control system 62
then applies the protocols of the appropriate subsystem to
determine the appropriate aircraft control settings 78 to
accomplish the pilot command 64 and transmits those control
settings 78 to the appropriate actuators illustrated by FIG. 4.
[0053] The control system 62 relies on feedback to implement the
commands of the pilot and preferably will incorporate explicit
model-following control architecture, as is known in the art. In
such a system, a pilot commands that the compound aircraft 2 assume
a selected flight condition. The control system 62 determines the
condition of the aircraft 2 utilizing sensors 40. The control
system 62 determines the changes to the condition of the aircraft
required to reach the commanded condition. The control system 62
applies an inverse model to determine the specific control settings
78 required to achieve the commanded condition and applies those
control settings 78 to the control effectors of the aircraft 2. To
compensate for disturbances and modeling or inversion error, the
control system 62 measures the changing state of the aircraft 2
using the sensors 40 and feeds back the information to update the
control settings 78 and achieve the commanded condition of the
aircraft 2.
[0054] For a conventional helicopter without the redundant controls
of a compound aircraft, the prior art model following/model
inversion process is straightforward. The model inversion
determines the single combination of trim control settings 78 that
will achieve the desired change in the state of the aircraft and
dynamically updates that combination of trim control settings 78 to
accommodate changing conditions and modeling errors.
[0055] For a compound aircraft 2 with redundant controls, the model
inversion process is more complex. Because of the redundancy, many
different combinations of trim control settings 78 can achieve a
particular change in aircraft state. For any particular change in
aircraft state, the forces to achieve that change in state can be
allocated among the applicable control effectors and the control
settings adjusted accordingly.
[0056] The control subsystems 68-76 illustrated by FIG. 5 require
different aircraft condition information 66 and require adjustment
of different combinations of control actuators 42-58, from FIG.
4.
D. Control System Architecture
[0057] Each of the control subsystems 68-76 is discussed below and
is illustrated in more detail by FIGS. 6-11. The following terms
have the following meanings in FIGS. 6-11 and in the discussion
below.
.beta..sub.p is the propeller pitch in degrees. .delta..sub.coll is
the collective control to the mixer in inches. .delta..sub.e is the
elevator deflection in degrees. .delta..sub.FO is the symmetric
flaperon deflection in degrees. .delta..sub.Flat is the
differential elevator deflection in degrees. .delta..sub.lat is the
lateral control to the mixer. .delta..sub.long is the longitudinal
control to the mixer. .delta..sub.yaw is the yaw control to the
VTDP mixer. .phi. is the roll attitude in radians. .theta. is the
pitch attitude in radians. .OMEGA. is the rotor speed in
radians/second. .tau..sub.y is the yaw response time constant.
a.sub.y is lateral acceleration in ft/sec.sup.2. `c` subscript
means `post-command filter.` `cmd` subscript means `command.` FADEC
means "Fully Automatic Digital Electronic Control. The FADEC
controls fuel to the engine to regulate rotor speed. GW means gross
weight. HP.sub.e is the engine power in standard horsepower.
HP.sub.VT is the power utilized by the VTDP in standard horsepower.
K.sub..theta. is the ratio of forward thrust from the rotor to one
plus forward thrust from the propellers. K.sub.D is the derivative
gain. K.sub.I is the integral gain. K.sub.P is the proportional
gain. P.sub.amb is the ambient pressure in pounds per square inch.
P is the roll rate in radians/second. q is the pitch rate in
radians/second. r is the yaw rate in radians/second. s is the
Laplace operator. T.sub.amb is the ambient temperature. .tau..sub.h
is the vertical response time constant. `trim` subscript means
`optimal trim value.` U is a pilot command that is filtered to
avoid exceedence of operating parameters. V is the forward speed in
knots or feet/second. VTDP means `vectored thrust ducted propeller`
and is the ducted fan. V.sub.Z is the vertical speed feet/second.
Downward is positive. .omega..sub.n.theta. is the pitch response
natural frequency. W is the vertical body velocity in feet/second.
.chi..sub.Bp is the relationship between propeller pitch and the
amount of thrust generated by the propellers, and varies with
airspeed. .chi..sup.CG is the longitudinal center of gravity (CG)
position.
[0058] FIG. 6 is a more detailed diagram of the overall control
system 62 architecture. Optimum trim schedule 68 is a lookup
database of combinations of trim control settings 78 and provides
control settings and aircraft attitude for trimmed flight. Optimum
trim schedule 68 schedules the combinations of trim control
settings 78 based on control inceptor input, aircraft condition
information 66 and the selected overall operational objective. The
aircraft condition information 66 for the optimum trim schedule 68
may include airspeed, weight and balance parameters and ambient
conditions. The output of the optimum trim schedule 68 comprises
optimum trim control settings 78 for all control subsystems 70-76.
As noted, the overall operational objective is selectable and may
include minimizing fuel consumption, minimizing fatigue damage or
any other objective or combination of objectives. As shown by FIG.
6, the optimal rotor speed (.OMEGA..sub.TRIM) selected by the
optimal trim schedule subsystem 68 is supplied directly to the
FADEC and determines engine power output.
[0059] As shown by FIG. 6, the pilot manipulates the control
inceptors 36 of cyclic pitch, collective pitch, pedal yaw control
and throttle. The pilot command is processed and the relevant
components, as indicate by FIG. 6, are directed to the subsystems
of speed command/speed hold control 70, longitudinal/vertical
control 72, lateral/directional control 74 and turn
coordination/yaw rate command 76. Each of the subsystems 70-76 also
receives the trim control settings selected by the optimal trim
schedule subsystem 68 and relevant aircraft condition information
66 from the sensors 40. Each of the subsystems 70-76 synthesizes
the information received and determines control settings to
accomplish the control task assigned to the subsystem, described in
more detail below.
E. Control Mixing
[0060] The resulting signals from the subsystems 68-76 must be
mixed as shown by FIG. 6. Control mixing is a conventional
technique used on rotary-wing and fixed-wing aircraft to minimize
cross-coupling effects. Cross-coupling is a change in aircraft 2
attitude or velocity that results from the displacement of a pilot
control inceptor which is not consistent with the primary objective
of that inceptor. For example, on a helicopter the primary function
of the collective lever is to increase or decreases thrust on the
main rotor 6 and thereby change the vertical velocity of the
vehicle. The control system achieves this effect by changing
collective pitch of the main rotor 6 proportional to the collective
lever. However, changes in collective pitch also change the torque
of the rotor 6 and the resultant torque reaction causes the
aircraft to yaw. Control mixing can be used to allow tail rotor
collective pitch to also vary with the pilot's collective lever in
order to reduce the collective-to-yaw cross-coupling. Control
mixing can be achieved mechanically or in the software of a
fly-by-wire control system.
[0061] As other examples of control mixing, both the
longitudinal/vertical control 72 and the lateral/directional
control 74 subsystems control the flaperons 18 to accomplish the
different tasks of those subsystems 72, 74. The flaperon 18 control
signals generated by both subsystems 72, 74 must be mixed in the
wing mixing module 82 to generate a consistent control signal to
the flaperon actuators 48, 50. Similarly, the collective and cyclic
control signals from the speed command/speed hold subsystem 70, the
longitudinal/vertical control subsystem 72 and the
lateral/directional control subsystem 74 must be mixed in the rotor
mixing module 84 prior to sending control settings 78 to the
collective and cyclic actuators 42, 44.
[0062] For a compound aircraft equipped with a Vectored Thrust
Ducted Propeller (VTDP) 22, the VTDP 22 provides selectable forward
or reverse thrust and yaw control using a combination of rudders
32, sectors 30, and the propeller 26. The rudder 32 and sectors 30
are in the slipstream of the propeller 26, so the forces generated
by these control surfaces are coupled to the propeller pitch. The
VTDP control mixer 86 is used to determine the combination of
rudder 32, sector 30, and propeller 26 pitch to achieve the desired
thrust and yaw moment for the given flight condition and pilot
control inputs. In hover and low speed flight, the VTDP 22 is
typically configured so that the sectors 30 and rudder 32 are fully
deflected and yaw control is achieved using variations in propeller
26 pitch (this is the "low speed mode"). In forward flight the
sectors 30 are retracted, and yaw control is achieved primarily by
deflecting the rudder 32 ("high speed mode"). The VTDP control
mixer 86 must change the control surface to achieve the low speed
mode, high speed mode, and transitions between the two modes. The
inputs to the VTDP control mixer 86 consist of the yaw control
input, a propeller 26 pitch setting (used for forward thrust
control), and calibrated airspeed of the aircraft 2. The yaw
control input and propeller 26 setting may come from the pilot
control inceptors, the electronic flight control system, or some
combination of the two. The output of the VTDP control mixer 86
includes the propeller 26 pitch, and the rudder 32 and sector 30
deflections. The VTDP control mixer 86 also may include constraints
to ensure proper controllability and to ensure load limits are not
exceeded. For example, decreasing propeller 26 pitch setting in
high speed flight may be used to slow down the aircraft 2, but the
reverse thrust is limited since it reduces flow over the rudder 32
and thereby decreases yaw controllability. There are also airspeed
limits on when the low speed mode can be used since large
deployment of sectors 30 at high speeds can result in large
loads.
F. Longitudinal/Vertical Control
[0063] FIG. 7 illustrates the architecture of the
longitudinal/vertical control subsystem 72. The coupled
longitudinal and vertical control subsystem 72 controls the
longitudinal pitch attitude (.theta.) and vertical motion (V.sub.Z)
of the compound aircraft 2. The longitudinal/vertical control
system 72 decreases the pilot workload by selecting control
settings 78 to achieve and hold a longitudinal pitch attitude in
response to a pilot command 64. The longitudinal/vertical control
system 72 also will select control settings 78 to achieve a pilot
command 64 for the aircraft to move vertically and a pilot command
64 to hold a commanded altitude.
[0064] The inputs to this subsystem are the commanded pitch
attitude, the commanded vertical speed, and the trim pitch
attitude. The aircraft condition information 66 used by the
longitudinal/vertical control 72 may include the aircraft pitch
rate and attitude, vertical speed and airspeed. The aircraft
control settings 78 generated by the control system 62 include
longitudinal cyclic, collective, elevator deflection and symmetric
flaperon deflection.
[0065] In operation and as shown by FIG. 7, the
longitudinal/vertical control 72 receives a command for a change in
aircraft pitch (.DELTA..theta..sub.CMD) and a change in vertical
speed (.DELTA.V.sub.ZCMD). The longitudinal command filter 88 will
consider the existing pitch rate and pitch attitude and determines
the desired dynamic response of the aircraft in the pitch and
vertical axes. The longitudinal command filter 88 allows the
aircraft to meet the ADS-33 Handling Qualities Specifications, if
applicable. The pitch response is represented by a second order
filter. The command filter calculates the desired pitch attitude,
attitude rate, and attitude acceleration. The desired attitude and
attitude rate are compared to measured values and multiplied by
gains and summed with the desired pitch acceleration. The summed
value is the pitch pseudo-control, which represents the commanded
pitch attitude acceleration. The modified command is indicated as
U.sub..theta. on FIG. 7.
[0066] As shown by FIG. 7, the pilot command 64 for a change in
vertical speed is treated similarly. The vertical response is
represented by a first-order filter. The time constant can be
selected to meet ADS-33 heave response specifications. The outputs
of the filter are the desired vertical speed and vertical
acceleration. The desired vertical speed is compared to the
measured value and passed through a proportional plus integral
compensator. This is summed with the desired vertical acceleration
to calculate the vertical axis pseudo-control, which represents the
commanded vertical acceleration of the aircraft.
[0067] From FIG. 7, the pitch and vertical axis pseudo-controls are
passed though a model inversion. The inversion module 92 determines
what changes to the control settings (.delta.) must be made to
achieve the filtered command (U). The inversion module 92 considers
the trim settings from the optimum trim schedule 68 and the
condition of the aircraft from the sensors 40. The inversion module
92 is `airspeed scheduled` because the results of the model vary
with air speed. The inversion module 92 allocates the pitch
commands between the redundant controls that affect the pitch of
the compound aircraft 2; namely, the longitudinal cyclic control
and the elevator 34 deflection control. For example, the inversion
module 92 may allocate more of the pitch control duties to the
elevator 34 and less to the longitudinal cyclic to reduce the loads
on the rotor hub 8.
[0068] To implement the model inversion of the
longitudinal/vertical control 72, the pitch pseudo-control (which
represents the second derivative of the pitch Euler angle) and the
vertical pseudo-control (which represents the vertical acceleration
in the inertial frame) are converted to pitch acceleration and
vertical acceleration in the body axes. Also, the desired pitch
attitude rate and vertical speed are converted to body axes.
q . c = U .theta. cos .phi. w . c = U v z + u cos .theta. .theta. .
c cos .theta. cos .phi. q c = .theta. . c cos .phi. w c = V z c + u
sin .theta. cos .theta. cos .phi. Eqn . 1 ##EQU00001##
At any given airspeed the linearized short-period longitudinal
dynamics can be represented by:
[ .DELTA. w . q . ] = [ Z w Z q + u 0 M w M q ] [ .DELTA. w q ] + [
Z .delta. long Z .delta. coll Z .delta. e Z .delta. F 0 M .delta.
long M .delta. coll M .delta. e M .delta. F 0 ] [ .DELTA. .delta.
long .DELTA. .delta. coll .DELTA. .delta. e .DELTA. .delta. F 0 ]
Eqn . 2 ##EQU00002##
This represents a linear state space model of the form:
{dot over (x)}=Ax+Bu Eqn. 3
The B matrix in this case is wide due to the redundant controls
effectors. Normally, model inversion is achieved by taking the
inverse of B:
u=B.sup.-1({dot over (x)}.sub.des-Ax.sub.des) Eqn. 4
However, in this case the matrix is square and cannot be inverted.
In fact, since there are redundant controls, there are many
different combinations of controls that will achieve the desired
pitch and vertical body accelerations. One possible solutions is to
use a left inverse of B,
B.sup.L=B.sup.T(BB.sup.T).sup.-1 Eqn. 5
which results in the control vector u with minimum norm. However,
one might want to put different weighting on the magnitudes of each
of the different control effectors. A weighted left inverse of the
B matrix can be represented as:
B.sup.+=W(BW).sup.T[(BW)(BW).sup.T].sup.-1 Eqn. 6
the control law can then be represented by:
u=B.sup.+({dot over (x)}.sub.des-Ax.sub.des) Eqn. 7
which gives the a control vector that achieves the desired
accelerations while minimizing the norm of the vector Wu.
[0069] In general, the longitudinal dynamics of the aircraft will
vary significantly with airspeed. An airspeed scheduled model of
the longitudinal dynamics can be represented by:
[ .DELTA. w . q . ] = A lon ( V ) [ .DELTA. w q ] + B lon ( V ) [
.DELTA. .delta. long .DELTA..delta. coll .DELTA..delta. e
.DELTA..delta. F 0 ] Eqn . 8 ##EQU00003##
The control law is given by:
[ .DELTA. .delta. long .DELTA. .delta. coll .DELTA. .delta. e
.DELTA. .delta. F 0 ] = B lon + ( V ) ( [ w . c q . c ] - A lon ( V
) [ w c - w trim q c ] ) where w trim = V tan .theta. trim Eqn . 9
##EQU00004##
The weighted left inverse is defined by:
B + = W lon ( B lon W lon ) T ( B lon W lon ) T ( B lon W lon W lon
T B lon T ) - 1 w lon = [ w long w coll w e w .delta. F ] Eqn . 10
##EQU00005##
[0070] The parameters w.sub.long, w.sub.coll, w.sub.e, and
w.sub..delta..sub.F are selected to get the desired distribution of
control to the longitudinal cyclic, collective, elevator, and
flaperons respectively. For example, the w.sub.e term can be
increased and the w.sub.long term increased to select more elevator
relative to longitudinal cyclic when maneuvering the compound
aircraft 2 in pitch. The distribution of the control among the
redundant effectors may be scheduled, stored in computer memory 60
and selected by the microprocessor 38 to achieve user-selected
overall operational objectives, such as limiting stress on a
particular part or subsystem, minimizing vibration, reducing
lifecycle costs, maximizing fuel economy, maximizing speed, or any
other operational objective.
[0071] The required changes to the control settings 78 from the
airspeed-scheduled inversion module 92 are directed to the three
mixers 82-86, as described above relating to FIG. 6.
G. Speed Command/Speed Hold
[0072] FIG. 8 illustrates the operation of the speed command/speed
hold subsystem 70. The speed command/speed hold system 70 allows
the control system 62 to respond to longitudinal acceleration
commands ({dot over (V)}.sub.CMD) and holds the existing forward
speed (V.sub.Z) of the aircraft when no change in command is
received. The overall purpose of the speed command/speed hold
system 70 is to simplify the pilot task of controlling forward
speed. In a helicopter, the forward acceleration and deceleration,
and hence forward speed, is controlled by the pitch attitude
(.theta.) of the main rotor disk. In the compound aircraft 2, the
forward speed of the aircraft 2 is controlled not only by pitch
attitude of the main rotor disk but also by the propeller blade
pitch (.beta..sub.P). The speed command/speed hold 70 control
subsystem integrates the propeller blade pitch control with the
pitch attitude of the aircraft 2 so that the pilot can control
forward speed with a single inceptor.
[0073] The engine torque limiting module 94, illustrated by FIG. 8
and shown in detail in FIG. 9, constrains the pilot command 64 for
forward acceleration or deceleration by imposing engine torque
(power) limits. Engine torque limits prevent the control system 62
from calling for control settings 78 that require more than 100% or
less than 0% of the available output of the engine. The aircraft
condition information 66 required by the torque limiting module
include the pilot command 64 for forward acceleration ({dot over
(V)}.sub.CMD), airspeed, engine torque, propeller blade torque and
ambient conditions. The output of the torque limiting module is a
constrained acceleration ({dot over (V)}.sub.CMD).sub.LIM.
[0074] The constrained acceleration command is integrated to create
a command forward speed. This is compared to the measured forward
speed to calculate an error signal which is passed through a PI
compensator and the resulting signal is added to the commanded
acceleration to create a pseudo-control for forward speed. The PI
compensator provides the Speed Hold function of the controller.
[0075] The forward speed pseudo-control represents the desired
forward acceleration. A simplified linear model of the speed
dynamics can be represented as:
{dot over
(V)}=X.sub..beta..sub.p.DELTA..beta..sub.p-g.DELTA..theta. Eqn.
11
Acceleration is proportional both to main rotor disk pitch attitude
and the change in propeller pitch. The equation can be inverted to
calculate the change in pitch attitude and propeller pitch needed
to achieve the desired forward acceleration:
.DELTA..theta. cmd = - K .theta. g U v . .DELTA..beta. p = ( 1 - K
.theta. ) X .beta. p U v . Eqn . 12 ##EQU00006##
[0076] The X.sub..beta..sub.p term represents the sensitivity of
auxiliary thrust due to propeller pitch changes. It is a function
of flight condition, and in this case is scheduled with airspeed.
It could be scheduled with other parameters if necessary, but an
exact value of the propeller sensitivity is not needed.
[0077] The change in propeller pitch is added to the trim propeller
pitch specified by the optimal trim schedule. Likewise the
.DELTA..theta..sub.cmd term is added to the optimal trim pitch
attitude in the Longitudinal AFCS. The propeller pitch setting can
be limited to observe torque limits on the propeller gearbox using
a similar scheme used for engine torque limiting.
[0078] The gain K.sub..theta. represents the amount of pitch
attitude used for acceleration relative to auxiliary thrust. So if
K.sub..theta.=1 the aircraft accelerates like a helicopter using
changes in pitch attitude and the propeller pitch only changes to
follow the optimal trim schedule. If K.sub..theta.=0, the aircraft
accelerates using auxiliary thrust while the pitch attitude follows
the optimal trim schedule. The designer can select K.sub..theta. to
achieve optimal distribution of control in maneuvering flight. The
gain may be scheduled with current aircraft condition and selected
overall operational objective.
[0079] The propeller sensitivity schedule 96 of FIG. 8 represents
the relationship between propeller pitch and the amount of thrust
generated by the propeller 26. This relationship is referred to by
the term X.sub..beta..sub.p and is a function of airspeed.
[0080] The speed command/speed hold subsystem 70 determines the
changes to the propeller pitch (.beta..sub.P) and to the pitch
attitude (.theta.) of the aircraft necessary to achieve the
allocated commanded forward acceleration or deceleration. If a
compound aircraft 2 is accelerating from the minimum power speed or
is descending, it can accelerate very quickly because excess power
is available. If the compound aircraft 2 is climbing, operating
near the maximum speed or performing an aggressive turn, then
available excess power and hence acceleration is limited. The
maximum acceleration may even be negative because the pilot needs
to bleed off airspeed to perform a maneuver. This subsystem
calculates constraints on the acceleration based on the current
power and available power of the aircraft.
[0081] The engine power (or torque) is measured and filtered. The
maximum power available from the engines is typically a function of
ambient conditions. Helicopter turbine engines have less power
available at high altitude or high ambient temperatures. In some
cases the power may be limited by transmission limits. A schedule
is used to determine maximum power. Maximum power is compared to
the measured power to determine the power margin. For the
deceleration limit, the power is compared to the minimum power,
which would be set to 0 or some small value to prevent over speed
of the rotor and drive system.
[0082] A simple relationship is used to estimate acceleration
limits from the power margin. The kinetic energy of the aircraft is
represented by:
K.E.=1/2mV.sup.2 Eqn. 13
[0083] The power increment required to accelerate can be estimated
by taking the derivative of the kinetic energy equation and also
including an efficiency factor, .eta..
.DELTA. P = mV V . .eta. Eqn . 14 ##EQU00007##
[0084] Substituting the power margin for .DELTA.P and allowing for
conversion from horsepower:
V . lim = 550 .eta..DELTA. HP m V Eqn . 15 ##EQU00008##
[0085] When implementing equation 15, it is necessary to put a
lower limit on the speed term to avoid division by zero.
[0086] The relationship expressed in equation 15 provides an
approximate expression for the acceleration limits. If the aircraft
is accelerating at a limit defined by Equation 15, the power margin
may not approach zero in steady-state. This may allow the aircraft
to exceed the power limit in steady-state or operate too
conservatively below the power limit. If the aircraft is operating
at an acceleration limit, a correction mechanism is engaged. The
power margin is passed though an integrator compensator and added
to the acceleration limit. This step forces the power margin to
approach zero if the aircraft is operating at an acceleration
limit.
H. Lateral/Directional Control
[0087] FIG. 10 is a schematic diagram of the coupled
lateral/directional control subsystem 74. The operation of the
lateral/directional control subsystem 74 is directly analogous to
the operation of the longitudinal/vertical control subsystem 72
illustrates by FIG. 7 and discussed above. The purpose of the
lateral/directional control subsystem 74 is to integrate the
redundant controls of the compound aircraft 2 for roll and yaw.
Those redundant controls include the lateral cyclic, differential
flaperons, sector, rudder and propeller pitch.
[0088] The inputs to this subsystem are the commanded roll
attitude, the commanded yaw rate, and the trim roll attitude. The
outputs of the system are the lateral cyclic, differential
flaperons, and yaw control. The lateral/directional control system
74 uses a model following/model inversion architecture. The model
and the control subsystem 74 are configured to achieve attitude
command/attitude hold in roll so that the aircraft will respond to
a command to achieve a desired roll and roll rate and will hold a
selected roll angle. The model and the control subsystem 74 also
are configured to achieve rate command heading/heading hold in yaw,
so that the aircraft will achieve a commanded yaw rate and will
hold a specified yaw angle in flight. The model and control system
can be configured to exhibit any desired dynamic response, such as
the requirements of the ADS-33 specification set. The desired roll
response is second order and can be designed to meet the ADS-33
roll bandwidth requirements. The desired yaw response is first
order can be designed to meet the ADS-33 yaw bandwidth
requirements.
[0089] The pilot commands 6 to the lateral/directional subsystem 74
include the pilot-commanded roll rate (.DELTA..PHI..sub.CMD) and
the pilot-commanded yaw rate (.tau..sub.CMD). The subsystem 74 also
receives the trim roll attitude from the optimum trim schedule 68.
The aircraft condition information 66 inputs include the roll
attitude and roll rate, yaw attitude, airspeed, and pitch attitude.
The outputs of the system are supplied to the rotor mixing module
84, the wing mixing module 82 and the VTDP mixing module 86, as
illustrated by FIG. 6. As in Equations 8 to 10, a weighted
pseudo-inverse is used to distribute control among the three
lateral-directional controls.
I. Turn Coordination/Yaw Rate Command
[0090] FIG. 11 illustrates in detail the turn coordination/yaw rate
command subsystem 76 shown on FIG. 6. The turn coordination/yaw
rate command subsystem 76 is desirable because the turning
characteristics and requirements of the compound aircraft 2 flying
at low speed are different from the turning characteristics and
requirements when the aircraft 2 is flying at high speed. The Yaw
Rate Command mode is desirable for speeds below 50 knots and the
Turn Coordination mode is desirable for speeds above 60 knots. A
simple blending scheme is used for transitions between 50 and 60
knots. At low speed (below 50 knots) the turn coordination/yaw rate
control system 76 will allow the aircraft 2 to turn at a yaw rate
(r) of up to 60 degrees/second. At high speed (greater than 60
knots) the control system allows a maximum lateral acceleration of
20 ft/sec.sup.2. For speeds between 50 and 60 knots, the yaw rate
and lateral acceleration commands are blended. The output of the
turn coordination/yaw rate command subsystem 76 (r.sub.CMD) is fed
directly to the lateral/directional control subsystem 74, as shown
by FIG. 6.
[0091] The turn coordination mode calculates an effective yaw rate
for turn coordination. This yaw rate can be fed directly to the
Lateral/Directional AFCS, so no changes to this AFCS need to be
made as the aircraft transitions from low speed to high speed mode.
The commanded lateral acceleration is compared to measured value.
The turn coordination controller can be derived from the equation
of motion for the lateral body velocity:
{dot over (v)}=a.sub.y-ur+pw+g sin .phi. cos .theta. Eqn. 16
With some simplifying assumptions one can derive the control
law:
r cmd = G TC ( a y cmd - a y ) + g sin .phi. cos .theta. V Eqn . 17
##EQU00009##
The yaw rate command is calculated to correct the lateral
acceleration error and then fed to the Lateral-Directional AFCS. J.
Return to Trimmed Flight after a Maneuver
[0092] At the end of a maneuver, the compound aircraft 2 returns
automatically to trimmed flight optimized to achieve the selected
overall operational objective. The microprocessor 38 constantly
updates the combination of trim settings selected from the trim
schedule based on feedback in the form of the measured condition of
the compound aircraft 2 as detected by sensors 40. The selected
combination of trim settings changes relatively slowly because the
parameters on which the selection is based (parameters such as
airspeed, altitude and vehicle gross weight) change relatively
slowly.
[0093] During a maneuver of the compound aircraft 2, the dynamic
inversion controller is constantly receiving feedback and is adding
or subtracting control effector deflections to achieve the
commanded maneuver. The corrections of the dynamic inversion
controller are based mainly on the feedback of aircraft condition
variables that change rapidly with time, for example angular rate
and aircraft attitude.
[0094] When the compound aircraft 2 reaches equilibrium after a
maneuver, the aircraft conditions as detected by the sensors 40
(and especially the fast parameters such as aircraft attitude) will
have reached near steady-state and the feedback signals will
approach zero or a small value. In addition, once the aircraft 2 is
in trimmed flight, the pilot will have moved the control inceptors
back to the detent position, indicating that the pilot is
commanding trimmed flight. The control system 62 will adjust the
control effectors consistent with the constantly-updated
combination of trim settings to achieve the selected overall
operational objective.
K. Simulation Study
[0095] The benefits of a compound aircraft 2, including a compound
aircraft 2 having the control system 62 of the Invention, have been
demonstrated. A simulation study was performed that compared the
performance of an AH-60 Blackhawk helicopter to a similar Blackhawk
helicopter equipped as a compound aircraft 2. The simulation study
evaluated conditions of minimum power usage and minimum vibration,
among others, in trimmed flight for comparable aircraft with
comparable loads and for comparable altitudes and air
temperatures.
[0096] The simulation study showed that the compound aircraft 2 was
capable of higher speeds than the comparable helicopter and that
the compound aircraft 2 was capable of operating with less power
than the helicopter for any particular airspeed. When optimized for
minimum vibration, the simulated compound aircraft 2 offered
significant vibration reduction compared to the helicopter. The
simulation study showed that reduced vibration did not necessarily
coincide with reduced power.
[0097] The simulation study demonstrates that no one set of control
settings is optimal for all possible missions of the compound
aircraft 2 and that control settings 78 may be optimized to
accomplish an overall objective, such as minimizing vibration or
minimizing power. The study also demonstrated that the optimum
control settings 78 to accomplish an objective vary with the flight
condition of the aircraft 2, such as airspeed.
L. Control of Maneuvering Flight
[0098] The control system 62 of the Invention may be used to
achieve a user-selected overall operational objective for
maneuvering flight as well as trimmed flight. For maneuvering
flight, the objective of achieving a selectable overall operational
objective may be accomplished by selecting appropriate weighting
factors for distribution of control among the various control
effectors.
[0099] An investigation conducted concerning the Invention
demonstrated that selection of weighting factors can limit
structural loads during critical maneuvers. For example, the
weighting factor "K.sub..theta." is applied in the Speed
Command/Speed Holding subsystem to allocate forward thrust between
the main rotor and the propeller. Different values of the
K.sub..theta. parameter may be scheduled and stored in computer
memory 60. The appropriate value for K.sub..theta. may be selected
by the microprocessor 38 to achieve a selectable overall
operational objective based on the current condition of the
aircraft and on the control inceptor input.
[0100] Other weighting factors applicable to the other control
effectors appear in the model inversions of FIGS. 6-11 and are
indicated collectively in the figures and the equations by the
symbol "w." In the discussion above for longitudinal control and
vertical control, the weighting factors are indicated as
"w.sub.long, w.sub.coll, w.sub.e, and w.sub..delta..sub.F" for
weighting factors for longitudinal cyclic, collective, elevator and
flaperons, respectively. Similar weighting factors exist for the
Lateral/Directional system; namely, w.sub.lat, w.sub.dir, w.sub.df,
relating to lateral cyclic, directional control and differential
flaperons, respectively. Other weighting factors also may be
applied.
[0101] The weighting factors may be selected by the microprocessor
38 from a weighting factor schedule stored in computer memory 60,
just as a combination of control settings for trim is selected from
the trim schedule. The weighting factor schedule may be configured
to select particular weights based on an overall objective to be
achieved (i.e., vibration reduction, life cycle cost reduction,
prevention of over stressing a component) and upon current flight
conditions such as airspeed, altitude, vehicle gross weight and
center of gravity position. Other parameters may be used to select
the appropriate weighing factor from the schedule.
M. Other Applications
[0102] The control system 62 of the Invention may be applied in any
situation where an objective can be accomplished through any one of
a plurality of combinations of control variables and where the
different combinations of control variables achieve different
overall objectives. In the control system 62 of the Invention, the
user selects among the overall objectives. The control system 62
then selects the appropriate combination of control variables to
accomplish the objective consonant with the selected overall
objective.
[0103] A user may select an overall operational objective by any
conventional means, such as by selecting an icon on a computer
display, by throwing a switch, or by affixing a jumper to a circuit
board.
[0104] In describing the above embodiments of the invention,
specific terminology was selected for the sake of clarity. However,
the invention is not intended to be limited to the specific terms
so selected, and it is to be understood that each specific term
includes all technical equivalents that operate in a similar manner
to accomplish a similar purpose.
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