Note: Descriptions are shown in the official language in which they were submitted.
CA 02591645 2007-06-28
VARIABLE GRADIENT CONTROL STICK
FORCE FEEL ADJUSTMENT SYSTEM
This is a division of co-pending Canadian Patent
Application No. 2,380,346 filed August 4, 2000.
TECHNICAL FIELD
The present invention is directed to the field of
control stick force adjustment systems as used in
aircraft and, more particularly, to an improved variable
gradient control stick force feel adjustment system.
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BACRGROIIND OF TEE INVENTION
Control stick force adjustment systems are used in the
aircraft aviation field to provide pilots with a better 'feel' and
control over their aircraft by adjusting the tension of the manual
s control system (e.g, control stick, cyclic stick, steering,
peddles, etc.,) at varying air speeds. For example, In a
conventional force trim mechanism for a helicopter cyclic system,
only a fixed force gradient is provided. In simple terms, for
every increment of cyclic displacement, the pilot -feels a
proportional force. It is of course desirable that a certain
amount of force is encountered in any direction a pilot moves the
controller (be it left or right, forward or backward). Force on a
cyclic stick provides the pilot, and ultimately the aircraft, with
stability during airborne operations. The force, typically, is
is generated by a four bar linkage that compresses or extends a spring
cartridge. Two linkage assemblies are utilized, one for lateral
motion and another for longitudinal motion. By. moving the spring
cartridge grounding points, the position where the pilot using the
cyclic stick feels zero force can be moved. Actuators called Force
Trim Actuators are also used to move the spring cartridge grounding
points. Because the linkages of the conventional lateral or
longitudinal force trim system move in a fixed plane, these
linkages are considered two-dimensional.
It should be appreciated that the force encountered in the
typical helicopter operation is a substantially linear
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relationship. When operating an airplane, however, a pilot
normally encounters a much stiffer control stick because a much
higher spring force is required as the aircraft travels at higher
airspeeds. Instead of moving the stick forward and backward, or
s the steering assembly left or right, with a normal force of one
pound per inch, a pilot should encounter approximately 3 pounds per
inch. Without the additional force, an aircraft flying at high
speeds could undergo very erratic and dangerous aircraft movement.
Many prior controller force adjustment systems utilize
electric motors to place a higher torque on the control stick,
resulting in a higher tensioned feel. Although the force trim
systems for some aircraft incorporate a spring tension against any
force exerted by the pilot against the pilot-controlled directional
gear, automated control is the predominant technology in later
is model aircraft. For example, in current tilt rotor aircraft
applications, a variable force field actuator takes a given
parameter (e.g., tilt rotor position or airspeed) and uses an
electric motor to in-turn cause an increasing or decreasing force
against the pilot-controlled directional system, based on inputs to
the electric motor by a controller. Such a system is not only
heavy but also very expensive because of the electronics in
controlling the motor and the redundancy that may be required with
automated systems in order to safeguard against potential system
failures.
is Many problems in achieving variable tension on manual
controllers are unique to a tilt rotor aircraft because it
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functions as both an airplane and a helicopter. Because a tilt
rotor aircraft operates as both, it is desirable to have the feel
of the tilt rotor aircraft change as it is converted from an
airplane to a helicopter, and vice a versa, during flight. The way
s that the 'feel' and resulting handling capabilities are accomplished
currently in tilt rotor aircraft systems (such as the Bell XV1S and
the V22 tilt rotor aircraft), is to use the heavier, more expensive
variable force field actuator systems, as described above. It would
be more desirable in tilt rotor aircraft applications, and for the
aircraft industry as a whole, to have access to a less complicated,
lighter and more reliable variable gradient cyclic force feel
system, such as disclosed in the present invention.
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SUMMARY OF THE INVENTION
In accordance with one aspect of the present
invention there is provided a gradient actuator
operationally coupled to lateral and longitudinal springs
of a latitudinal and longitudinal control mechanism of a
tilt rotor aircraft, the gradient actuator further
coupled to a Nacelle position sensor via a controller for
tracking the Nacelle position of the tilt rotor aircraft
during operation so as to provide a force feel to a pilot
of the aircraft according to the position of the Nacelle,
wherein the gradient actuator is controlled by the
controller to change the moment arm length of a lateral
spring cartridge assembly with respect to a lateral
actuator and a longitudinal spring cartridge assembly
with respect to a longitudinal actuator, the change in
moment arm length being based on input to the gradient
actuator from the controller, the controller receiving a
position signal from the Nacelle position sensor.
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BRIEF DESCRIPTION OF THE DRAWINGS
The foregoing and other objects, aspects and advantages are
better understood from the following detailed description of a
preferred embodiment of the invention with reference to the
drawings, in which:
Figure 1 is a perspective view of a cyclic control stick
mechanism for a helicopter incorporating the variable gradient
control system of the present invention;
Figure 2 is a plan view of the position of the gradient
to actuator, linkages and moments arms when the aircraft is in
helicopter mode;
Figure 3 is a plan view of the position of the gradient
actuator, linkages and moment arms when a tilt rotor aircraft is in
airplane mode;
is Figure 4 is a superimposed plan view of the gradient actuator,
linkages and moments arms in airplane mode over helicopter mode;
Figure 5 is a graphical illustration of gradient force on the
y axis, plotted against a Nacelle angle on the x axis; and
Figure 6 is a schematic illustration of interacting components
20 for one system configuration given the teachings of the present
invention.
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DETAILED DESCRIPTION OF THE INVENTION
One preferred embodiment of a system in accordance with the
present invention is practiced with the lateral and
longitudinalitudinal pilot-actuated control system for a tilt rotor
s aircraft. The operation of a standard force trim system in a tilt
rotor aircraft will - first be generally discussed in order to
provide a frame of reference to compare the benefits of the
invention. It should be understood from the teachings of the
invention that it can be applied generally to all manual control
to stick system across the entire aircraft industry. For purposes of
this description, the terms 'control stick' or 'stick' are meant to
generically apply to manual control systems (e.g, control sticks,
cyclic, steering mechanisms, etc.) commonly found in the aircraft
(helicopter, airplane and tilt rotor aircraft) industry.
is If a pilot took the control stick in a conventional force trim
helicopter system and moved.it one inch, he would encounter about
one pound of force. If the pilot moved the same control stick 2
inches, he would encounter about 2 pounds of force. The same force
would be encountered in either left or right movement of the stick.
20 If the pilot is flying a standard airplane and moved the control
stick therein either forward or backward one inch, he would
encounter about 2 pounds per inch of force. If the control stick
were moved 2 inches, he would encounter 3-4 pounds per square inch
of force. In a conventional airplane, the pilot is encountering
25 about 2 pounds of force for every inch of controller movement. In
a helicopter, however, the force per square inch relationship is
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much more linear.
When a pilot flies in a airplane, he desires a much stiffer
feel over the control-stick controller than in a helicopter. The
same high tension force is not be desirable in a helicopter where
s faster movement and mechanical response is desired. A tilt rotor
aircraft requires the. ability to do both in order to have variable
force. Rather than solely using an electric motor to artificially
place a higher torque on the controller, resulting in a higher-
tensioned feel, the present invention uses a variable gradient
actuator in combination with a three-dimensional phasing linkage to
cause a moment arm on which the typical spring cartridge mechanism
is attached to change its distance with respect to the
Latitudinal/Longitudinal adjustment mechanism for the aircraft and
stick. Based on the simple engineering principle 'moment is equal
is to force times length a moment of one foot pound is equal to one
foot moment arm times one pound of force', instead of changing the
force on the stick electronically through motors, a mechanical
variance in the moment arm relationship to the directional hardware
and/or pilot control mechanism is changed.
Referring to Figure 1, a perspective view of the pilot
operated section of the latitudinal and longitudinal control system
10 .in an aircraft incorporating the force feel adjustment
improvement of the present invention is illustrated. The system 10
includes two spring cartridges, a lateral spring cartridge 5 and a
longitudinal spring cartridge 6. The lateral spring cartridge 5 is
tied to the control sticks 2, and controls the lateral motion of
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the sticks 2 (side to side motion). Through an additional linkage,
the sticks 2 are tied to the longitudinal spring cartridge 6, which
provides longitudinal motion (forward and back motion) to the
sticks 2. From the perspective view of the Figure, it can be seen
that lateral spring cartridge 5 is attached to a bell crank 7 that
is linked at 9 and 10 to the sticks 2 for lateral motion. The
spring cartridge's 5 opposite end, or what is referred to as ground
end, is attached to a lateral trim actuator 15. The lateral trim
actuator 15 may allow a pilot to reset to zero, or neutral force,
to the sticks 2 position by using a beep switch 3 located on the
sticks 2. Reset causes the bell crank 7 on the actuator 15 to move,
changing the systems zero point. As the bell crank 7 moves, the
sticks 2 move along with it such that a new zero point would be
achieved. The same is true for longitudinal functions. From the
is Figure, it can be seen that the longitudinal spring cartridge 6 is
also associated with a longitudinal trim actuator 16. The
longitudinal spring cartridge 6 and longitudinal trim actuator are
linked with the sticks 2 and associated with the beep switch 3, as
indicated above.
20 The actuators 15, 16 have a clutch (not shown), which
controls, or clutches, the movement of each associated bell crank
in and out. No spring force is required at all in some flight
modes such as helicopter applications, which require very
responsive stick action for precise aircraft motion. In such an
25 application, the bell cranks are actually declutched, allowing the
sticks 2, 3 to move together against minimal force. Free motion,
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however, is completely unacceptable for a tilt rotor aircraft
operating in airplane mode because a small cyclic reflection could
cause very dramatic aircraft motion. The clutch for each actuator
must therefore be disabled in airplane mode.
s For a control stick system to meet the requirements of the
tilt rotor, it must.., have a control stick force that increases with
increased speed when converting from helicopter to airplane
(otherwise known as a variable gradient force). The invention,
through the linkage arrangement described herein, changes the
moment arm length of the lateral and longitudinal springs 5, 6
through a gradient actuator 17 and linkage arrangement 18, 19 to
the lateral and longitudinal springs 5, 6, respectively.
Referring to Figure 2, what is illustrated is a plan view of
the gradient actuator 17 and its associated linkages 18, 19 to the
is lateral 5 and longitudinal 6 springs. The moment arm 21 for the
roll, or lateral, axis (or lateral setting) is approximately 2.2
inches in Helicopter mode. The moment arm 22 for the pitch axis
(or longitudinal setting) is approximately 1.8 inches.
There is a need for higher force on the sticks 2 when a tilt
rotor aircraft must move into airplane mode. Referring to Figure
3, to accommodate this needed change in force, the gradient
actuator 17 increases the length of the moment arms 21, 22 so that
the springs 15, 16 can also accommodate the change. Through
movement by the gradient actuator 17, the roll axis can be
increased to about 3.4 inches and the pitch axis to about 3.44.
Referring to Figure 4, the invention in Airplane mode is shown
CA 02591645 2009-09-10
superimposed over Helicopter mode. The respective positions of the
moment arms, 21 and 22 for each mode are what accomplishes the feel
and control advantages of the invention.
Referring to Figure 5, a graphical illustration shows what a
s pilot may feel as the aircraft is transitioning from airplane mode
to helicopter mode.. Two curves illustrate a force that decreases
for longitudinal stick position (or pitch) from seven pounds per
inch down to only about 2-3/4 pounds per inch as the angle of the
tilt rotors, with respect to the horizon, is increased (or as the
tilt rotor aircraft is otherwise moved from airplane mode into
helicopter mode). The pilot will encounter a lateral feel that
undergoes a similar change, from about 3-1/2 pounds in force in
airplane mode down to about 1.8 pounds of force in helicopter mode.
The size of the change may be made dependent on what is desired by
is the pilot. It is conceivable that the load on the cyclic stick
could actually go down to 0 by having the moment arms move to a 0
moment arm length, if such a change were desired by the pilot. Such
diverse operation would be coordinated by the controller, through
the gradient actuator 17.
Referring to Figure 6, sensors 61, 62, located in each Nacelle
53,54 of the tilt rotor aircraft (not shown) send information to
the controller 67, which then causes the gradient actuator 17 to
make the necessary adjustments to the moment arms 21, 22. In
addition to Nacelle placement, the controller 67 can also receive
input based on air speed 63, which would be used to determine
moment arm placement. Tailored pilot settings may also be input to
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the controller 67, manually 64 aad/or from memory 65. Presently,
however, because of the complexity of measuring air speed and the
possibility for controller failure or pilot miscalculation, it is
most simple and reliable to tie moment arm adjustments directly to
s the Nacelle angle. Such an arrangement could conceivably be made
with minimal electronic control by slaving the gradient actuator 17
to sensors /transducers 61 or 62 located at either Nacelle.
Furthermore, the Nacelle angle can be sensed very reliably and
redundantly with the placement of different independent sensors
throughout the system.
Because both lateral and longitudinal system force feel values
are set by the same parameter, a gradient actuator can be used to
vary the force gradient of both systems. Great benefit is derived
through the use of the lateral and longitudinal three-dimensional
is phasing linkages driven by a single gradient actuator as described
herein. This configuration results in a simple, light-weight
system. It should be appreciated that if independent varying of
lateral and longitudinal force feel values is desired, two separate
actuators in response to signals from a control mechanism as
described herein can also be used.
The variable gradient system of the present invention takes
the conventional system and varies the length of the moment arm.
This is accomplished by using a single actuator (which is not
independently back-drivable) to pivot the moment arms for the
is lateral and longitude motions of the stick so that desired tensions
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are accomplished. The conventional fixed gradient force trim
system described in the background, for example, can be converted
into a variable gradient force adjustment system with the
application of a three-dimensional phasing linkage arrangement. By
s adding hinge points to the bell cranks of a conventional force trim
system, the plane of the two dimensional linkage can be rotated via
a gradient actuator linked to the hinge points. As the operating
plane of the two dimensional linkage is rotated out of the original
plane of operation, the effective moment arm of the bell cranks
to with respect to the gear is reduced. The reduction in effective
moment arm reduces the amount the spring cartridge is compressed or
extended by the cyclic. The reduction continues until the linkage
is rotated to about 90 from its original position. At about 90
the effective moment arm is zero and cyclic movement has minimal
is effect on the spring cartridge. In effect, the three dimensional
linkage system allows the cyclic force felt by the pilot to be
continuously phased to zero force.
A pilot of a tilt rotor aircraft flying in helicopter mode can
now move into airplane mode and realize a gradual increase of
20 controller force and stability. The present system is simple,
relative to current systems, in that the redundancy required by
most force trim systems be overcome, or otherwise eliminated. An
aircraft using the present system would only lose the variable
gradient force should the gradient actuator fail, leaving the
25 controller in one force position. Such a condition wouldn't change
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with speed of the aircraft, but would be a much more benign failure
than having a system that goes completely limp. Furthermore, from
the standpoint of cost, weight, reliability and simplicity, the
present invention is a major improvement over current systems.
While the invention has been described in detail above, it
should be understood that it has been presented by way of example
only, and not limitation. Thus, the breadth and scope of a
preferred embodiment should not be limited by any embodiments
described above, but should be defined only in accordance with the
io following claims and their equivalents.
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