Note: Descriptions are shown in the official language in which they were submitted.
1~8219
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Description
Multi-Axis Force Stick, Self-Trimmed
Aircraft Flight Control System
.
Technical Field
This invention relates to aircr~ft control systems,
and more particularly to an aircraft control system
providing an entirely new mode of controlO
Background Art
In both fixed wing and rotary wing (helicopter)
aircraft, it is common for the pilot to use a variety
of positionable controls, such as s-ticks, levers,
wheels and pedals, to position the control or aero-
dynamic surfaces of the aircraft, thereby to control
the aircraft attitude, altitude, speed and the like.
In the simplest of systems, the controls are connected
by cables to the control surfaces (such as pedals
connected by cables to the rudder of a fixed wing
light plane). In more complicated systems, the con-
trols may have mechanical connections which are
boosted by hydraulic servos, or the like.
~ As aircraft systems become more and~more compli-
; cated, the useful space in the cockpit which is
accessible to the pilots becomes more nearly filled with
instruments, switches and the like. Thus, the controls
themselves compete for cockpit space with other
apparatus.
In a typical aircraft, there ls a control wheel
on a stick that controls the roll and pitch of the
aircraft, pedals that control a rudder, and a throttle
console for controlling the engine thrust. In a heli-
copter, there is typically a cyclic pitch stick for
controlling pitch and roll of the aircraft, pedals
S-3371
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for controlling yaw, and a collective pitch stick for
controlling vertical lift. These controls and their
mechanic~l connections to control surfaces or servo
mechanisms responsive thereto, together encumber the
cockpit space to a large degree. For instance, the
presence of the control wheel or stick in front of the
pllot seat renders it impractical to have electronic
displays or the like immediately in front of the pilot
due to the need for moving the wheel or the stick into
various positions in that space, and due to the fact
that the presence of such apparatus blocks the vision
of the pilot in certain angles. The presence of foot
pedals renders it difficult to provide forward and
downward visibility to the pilots, as would be useful
in helicopters employed in logging operations, con-
struction and the like. Additionally, whenever pas-
sengers sit in one of the pilot seats, inadvertent
control inputs can be provided by unwanted passenger
contact with the controls. Access into and out of the
pilot seats is also encumbered to varying degrees by'
these controls.
In systems employing pilot and copilot controls,
it is essential that the controls be synchronized
positionally to each other, so that one pilot can take
over from the other without providing abrupt inputs
into the control system. For this reason, each of
the pilot controls is normally mechanically connected
with the corresponding copilot control. For the
most part, such interconnections are mechanical, be-
cause hydraulic or electric sensors and actuatorsnecessary to avoid mechanical connections are too slow
and cumbersome for such use.
In order to avoid some of the deficiencies noted
above, attempts have been made in the past to provide
"side-arm" controllers which may be operated by a
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pilot while his hand is resting on the arm of a seat.
Also, in aircraft or spaee craft in which the pilots
have to withstand high gravitational forces, the cushio-
ning of the pilot in a seat has led to the use of some
side-arm controllers. Typical side-arm controllers
whieh have found some measure of success are limited to
two axes, usually ineluding piteh and roll. However,
- this leaves throttles or eollective pitch sticks and
pedals to be dealt with in the traditional fashion,
thereby recluiring the pilot to reaeh outside of his
seat for hand eontrols, and forcing his position to be
established relative to foot pedals. This also fails
to clear the clutter out of the cockpit to the fullest
extent.
There have also been attempts to make side-arm
controllers operative in more than two axes. These may
inelude pitch, roll and yaw, or pitch, roll and colleetive
(or throttle, in the ease of a fixed wing plane).
However, there has been a universal failuré in side-arm
eontrollers designed to eontrol three or more axes due
to the eross eoupling between the axes. Thus, if one is
eontrolling piteh and roll with fore~aft and right-left
motion, one eannot also eontrol colleetive piteh in a
helieopter with an up-down motion of the same stiek,
sinee any tendeney to move the stiek fore and aft also
results in the stiek moving up and down to a certain
degree (and vice versa)~ It is believed that this is an
inherent problem of the manner in which the human hand is
conneeted to the human forearm, with essentially a pivot
at the wrist. This eonfliets with the pivotal action of
a side-arm eontroller having three or more axes sinee
the natural human wrist motion eauses eoupling between
stiek motions in the different axes. The same is true
with respeet to twisting motions when they are eombined
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with fore-aft and right-left motions.
In order to reduce aircraft weight, to provide
redundancy in systems for additional reliability
and survivability, and to take advantage of modern
technology (such as computers), there has been some
investigation of "fly-by-wire" systems, which are
characterized by sensors and actuators connected either
electrically or optically (or both) to avoid mechanical
interconnections in an aircraft. In such a case, the
typical mechanical linkage operating a booster servo
to position thecontrol surfaces of an aircraft might
be supplanted by an electrical position sensor which
in turn controls an electro/hydraulic actuator. How-
ever, there has heretofore been difficulty in pro- -
viding fly-by-wire systems which can cause synchroni-
zation between the pilot and copilot controls without
undue added complexity and cockpit-mounted apparatus.
Thus in fly-by-wire systems adapted for use in air-
craft having controls common a~ this time, mechanical
interconnections between the pilot and copilot controls
and electrical transducers connected to the single
mechanical connection are typically proposed. This is
necessitated by the fact that the position (such as
the cyclic stick in a helicopter or the wheel of a
fixed wing aircraft) must be in the same position at
both the pilot and copilot stations if transfer of
control is to be shifted between the pilot and the
copilot. But the motion or position of such controls
cannot be readily synchronized other than mechanically
due to the inherent difficulty of suitably fast
follow-up systems which do not take up too much space.
Disclosure of Invention
Objects of the invention include providing air-
craft controls which reduce cockpit clutter, permit
improved visibility, reduce pilot fatigue and support
fly-by-wire and/or fly-by-light control systems.
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In accordance with a particular embodiment of
the invention there is provided a control system for pro-
viding principal manual control to an aircraft having
four control axes, the four control axes including pitch,
roll, yaw and lift~speed. The contro~ system includes
a plurality of positionable aerodynamic surfaces, the
positions of which control the aircraft in the four con-
trol axes. Control means are operable by a pilot to
provide aerodynamic surface positioning command signals
and positioning means are connected between the control
means and the aerodynamic surfaces and operative in res-
ponse to the positioning command signals applied thereto
to control the positioning of the aerodynamic surfaces.
In accordance with the invention, the control means
further includes a multi-axis force stick adapted to be
held by the hand of the pilot for providing output signals
indicative of forces applied to the stick in at least
three distinct stick axes, each of the stick axes corres-
ponding to a related one of the aircraft control axes.
A plurality of signal processing channels are each connec-
ted for response to the output signal related to a cor~es-
ponding one of the stick axes each providing a related
positioning command signal to the positioning means which
is both a proportional function and an integral function
of the related force stick output signal for providing
manual control of the aircraft by flying to a continuously
updated trim point established by force inputs to the
force stick, the integral portion of each control signal
establishing the trim point for each axis and the propor-
tional portion of each signal causing deviation from the
trim point in such axis.
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This invention is predicated on our discovery
that coupling between the axes of side-arm controllers
having three or more axes of control is eliminated by
the use of a force-responsive control stick, with no
perceptible motion required in order to provide the
necessary force inputs. This invention is also
predicated on our discovery that a multi-axis force
stick provides an improved input to an aircraft control
system when utilized as a trim adjustment, in a system
employing forward integral and proportional paths with
a suitable response characteristic, which is on the
order of pilot reaction time to aircraft response to
the inputs made thereto through the force stick.
According to the invention, a control stick oper-
able in more than two axes is responsive to force
applied in the pluraLity of axes by the pilot to pro-
vide proportional and relatively fast integral inputs
to rapid, full authority control surface position
actuators.
According to the invention, a force stick, operable
in threè or four axes, responsive to force within a
suitable control range of forces, and without any
motion which is perceptible to the pilot while con-
trolling the aircraft in flight, is used as an input
to a control system. In still further accord with the
present invention~ electrical signals from a control
stick are utilized to provide proportional and integral
commands to actuators which adjust the position of the
control surfaces of the aircraft, whereby the electrical
inputs provided by the pil~t adjust a continuously
updated trim point in each of the controlled axes.
The present invention (the use of a multi-axis
force stick together with a proportional and integral
control system) provides the capability for a pilot to
control an aircraft in response to his perceptions of
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changes in attitud~, altitude, speed, heading and the
like, with control inputs provided by the pilot only
in the event that a change in the aircraft response is
desired. This comprises a wholly new concept of
aircraft flight control (flying to trim).
The present invention provides, for the first
time, the capability of employing a single control
stick (such as on a side-arm controller) to control
three or four axes without any coupling between the
axes. The invention significantly reduces pilot
fatigue since uncomfortable positioning and excessive
motion of the pilot's body is not required, as is true
of common, position-related control systems. Because
it permits flying to a constantly updatable trim point
in each of the controlled axes, the invention eliminates
the need for synchronization between the pilot and co-
pilot sticks. The invention allows elimination of the
large conventional sticks, pedals and thé like which
hinder visibility of instruments and of the outside
world and which take up excessive space. The invention
reduces pilot workload without any compromise of air~
craft maneuverability. The invention makes it possible,
for the first time, to fly an aircraft without use of
the feet, and with one free hand. The invention further
permits the provision of highIy sophisticated aircraft
control systems at a cost which is inherently capable
of being less than the cost of conventional stick and
pedal systems. The invention may be readily implemented
in the light of the specific teachings thereof which
follow hereinafter by employment of apparatus and
technology which is well within the skill of the art.
The foregoing and other objects, features and
advantages of the present invention will become more
apparent in the light of the following detailed
description of exemplary embodiments thereof, as
illustrated in the accompanying drawings.
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Brief ~escription of Drawings
Fig. 1 is a perspective illustration of a side-arm
controller in accordance with the invention;
Fig. 2 is a simplified schematic diagram of an
aircraft control system, for a helicopter, in accordance
with the present invention;
Fig. 3 is an illustration of response character-
istics which may be employed in the control system of
Fig. 2;
Fig. 4 is a simplified schematic illustration of
one manner of implementing a characteristic of Fig. 3
in the system of Fig. 2;
Fig. 5 is an illustration of another response
characteristic; and
Fig. 6 is a partial schematic diagram of a
modification to the system of Fig. 2 for providing
power remaining indications.
Best Mode For Carrying Out The Invention
Referring now to Fig. 1, a side-arm controller 10
according to the invention may comprise a stick 12
mounted on a suitable sensing ~ransducer assembly 13
which is disposed on an arm 14 of a pilot seat 16.
The arm 14 may be pivoted as at 18 so as to be rotat-
able upwardly and out of the way, thereby to provide
access to the seat or to remove the side-arm controller
10 from the vicinity of a passenger's hand, if desired.
- As illustrated in Fig. 1, the side-arm controller 10
has four axes including fore-aft, right-left, up-do~n
and twist. The fore-aft axis may relate to the pitch
of the aircraft, and thereby control the longitudinal
cyclic pitch channel of a helicopter or the elevator
of a fixed wing aircraft. The right-left axis of the
controller 12 may be used to control roll, and there-
fore control the lateral cyclic pitch channel of ahelicopter or the ailerons of a fixed wing aircraft.
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The twist axis of the controller 10 may be used to
control yaw, and therefore control the tail rotor pitch
channel of a helicopter., or the rudder of a fixed wing
aircraft. The up-down axis of the controller 10 may
control lift/speed, and therefore control the collective
pitch channel of a helicopter or the throttle and/or
engine/propeller blade pitch of a fixed wing aircraft.
In accordance with one aspect of the invention,
the controller 10 is a force controller capable of
responding to measurably distinct forces.applied there-
to by the pilot, in any one or all four of the axes,
(or three axes if desired), while requiring no motion
of the stick, other than a minimal amount necessary to
detect the force and which is imperceptible to the pilot
in contr~st to the forces applied by him. The response
of the stick to forces, and the capability of the stick
to sense the applied forces while itself permitting no
motion of any consequence in the.direction of any applied
force, avoids any conflict between the natural motion
and position reflexes of a human hand and.forearm, and
therefore supplies the capability to provide inputs to
all four of the axes without coupling between any of
the axes (that is, without an upward motion also tending
to be a rearward motionl and the like). A force stick
of this type, hav.ing imperceptible motion~ is readily
~vailable in the market, one of which being the Model
404-G517, produced by Measurement Systems, Inc.,
Norwalk, Connecticut, U.S.A. Other sticks could
readily be utilized; the only requirement for the prac-
tice of the present invention being that the stick besufficiently stiff .in all axes and have sufficiently
sensitive force measurement characteristics so that a
suitable range of force sensitivity (such as on the
order of between 0 and 40 lbs. in either direction of
each axes) can be achieved while the motion required
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to sense such forces (such as by strain gages measuring
the minute deflection resulting from the applied force)
is imperceptible to the pilot while maneuvering in
flight. By `'imperceptible`', it is meant that the motion
S which results from adequate force inputs to control the
aircraft is so slight that there is essentially no
sensation of motion, and thus there is no coupling between
axes as a consequence of hand motion.
As described hereinbefore, one aspect of the ~ -
invention is the discovery that a multi-axis stick,
capable of use for controlling three or four axes of
aircraft response without coupling between axes, is
achieved by using a stick which is responsive only to
force, with imperceptible motion. However 7 the application
of nearly constant force is inherently fatiguing. Main-
taining constant forces in three or four axes at the same
time can obviously be an additional source of fatigue.
Furthermore, it has been found that rapid maneu-
vering in a plurality of axes, such as turning a heli
copter 180 during hover in gusty wind conditions, is a
difficult task to perform when all four axes of the he-
licopter are being controlled in a single hand. Although
all of the human factor relationships, including func-
tioning of the hand itself and pilot reaction to aircraft
response, are not fully understood, it is believed that
this difficulty occurs as a consequence of the need for
coordinated commands in two or more axes during such
complex maneuvers. The force stick of the invention
differs from conventional control systems in which the
aircraft responds to positions of the controls, which po-
sitions can easily be adjusted in a minute fashion, with
the aid of the eye and with the aid of relative human
member position reaction (e.g. where the hand is with
respect to the knee). And it differs from conventional
control which
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allocate different tasks to different body members that
are accustomed to handling those tasks, only the stick or
wheel requiring single hand coordination for the pitch
and roll axes of control.
S The ~oregoing problems with a multi-axis force
stick are overcome by a second aspect of the invention:
the provision of a control system which has a close trim
follow-up. That is, any input provided by the pilot is
used to establish a new trim or reference point of con-
trol for the related axis. Thus, with the invention the
pilot responds to his observations of aircraft attitude,
speed, altitude, and changes therein, provided to him by
visual observation or by instruments, and essentially
adjusts the current trim position of the aircraft control
surfaces to provide corrections thereto.
Referring now to Fig. 2, a control system incor-
porating the present invention employs a four-axis force
stick of the type described with respect to Fig. 1. The
force stick 10 has a plurality of outputs 20~23 that
provide signals of which the voltage is a known function
of force applied in the vertical, longitudinal, lateral
or twist axis of the stick 10. In the stick 10 de-
scribed with respect to Fig. 1, each axis is bilateral,
providing voltages in respectively opposite polarities
for vertical motion in the up and down directions, for
longitudinal motion in the fore and aft directions, for
lateral motion in the right and left directions, and
for twist motion in the clockwise and anticlockwise
directions. Also, for the force stick described herein-
before, the voltages are nearly linear functions of
force. However, this need not necessarily be so, since
a plurality of signal conditioning circuits 24-27, one
for each of the outputs 20-23, may be employed to provide
a specific voltage to force
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relationship on signal lines 28-31, which comprises
the actual signal input to the control system.
An example of signal conditioning provided by the
circuit 26 is illustrated in Fig. 3~ Therein, the
abscissa is the lateral force, either to the left or to
the right, and the ordinate is the voltage at the output
of the circuit 26 on the line 30. The signal condition-
ing is of course a voltage to voltage conditioning,
depending upon the force to voltage relationship of the
signal on the line 22. ~owever, in terms of desired
functional result in the example herein, Fig. 3 il-
lustrates that a dead band of about one-half pound
(0.23 Kgm~ in either the right or the left direction
may be provided, so as to reduce inadvertent pilot in-
puts and any potential for drift about the lateral-zero
center of sti~k force. This is essential to avoid
long-term integration of minute inadvertent signals, as
described hereinafter. Then, a rather sensitive region
in either direction may be provided for forces between
a half pound and four pounds. This may increase linearly
from zero volts to 8/10 of a volt ~in the correct
polarity). Above forces of about 4 lbs. in either
direction, the output of the circuit 26 (Fig. 2) as
illustrated in Fig. 3 may increase in an increasing
fashion with force, so as to provide very sensitive
operation at low forces yet provide for fast, full au-
thority response in the control system when needed. In
Fig. 3, the voltage to force relationship is shown as
~eing nonlinear, with increasing slope. However, the
particular shape may be tailored to suit any imple-
mentation of the present invention, depending upon the
other factors of the control system, such as the char~
asteristics of the hydraulic servos, as well as the
aircraft flight characterlstics and desired aircraft
response, all as is within the skill of the art.
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An example of how the signal conditioning of the
type illustrated in Fig. 3 may be readily achieved, is
illustrated in Fig. 4, which depicts how suitably biased
and limited ampli~iers might be arranged to provide a
composite conditioning of the signal as illustrated in
Fig. 3. In Fig. 4, the signal conditioning circuit 26
may comprise six amplifiers 26a-26f. The dead-band
amplifiers 26a and 26b each have zero gain until a
voltage representative of a half pound of force is
10! reached, after which these amplifiers provide linear
gains of one. This simply provides a dead band of
~.5 lbs. The vernier gain amplifiers 26c and 26d pro-
vide the low force sensitivity region, by providing the
zero gain for forces of the opposite direction, and
for any voltage passed by the dead-band amplifiers 26a?
26b a linear gain of 2/10 of a volt per lb. to a max-
imum of 8/10 of a volt1 where the output is then clamped
or limited. The high gain amplifiers 26e and 26f provide
the high gain for high forces, which is depicted as
nonlinear with increasing slope in Fig. 3. These
therefore have zero gain until the output of the vernier
gain amplifiers 26e, 26f attains 4 volts, after which
the gain increases to the lirnit of the input signal.
The output of the vernier and high gain amplifiers 26c-
26f are summed in a summing junction 26g, which may
comprise a special summing amplifier or may comprise
the input network to proportional and/or integral gain
device described with respeck to Fig. 2 hereinafter.
~he pitch and yaw channels may have signal con-
ditioning to provide characteristics similar to those
described with respect to Fig. 3. In fact, the invention
has been implemented with a ~itch channel characteristic
identical to that of the roll channel characteristic
illustrated in Fig. 3, and has been implemented with a
yaw channel signal conditioning
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characteristic which differs from the roll characteristic
of Fig. 3 only in that the gain is .225 volts per inch
lb. of torque~ and the dead band is ~.27 inch lbs.
The collective channel, on the other hand, may
have a different shape curve, one characterized with ne-
gative change in slope with respect to force. As illus-
trated in Fig. 5, the vertical channel may require ~0 lbs.
of force for maximum stick inputs (rather than 20 lbs. as
in the right-left and fore-aft axes). A dead band of
~1 lb. may be employed, and linear gain in the up
direction may be on the order of .19 volts per lb. while
the gain in the down direction may be on the order of .8
volts per lb. but existing over an expanse of 8 lbs. in
the negative direction. Additionally, Fig. 5 illustrates
that, to accommodate the droop ln the relationship
between collective pitch and airspeed, the slopes in
Fig. 5 may best be decreasing (rather than increasing as
in the case of the pitch, roll and yaw channels). Exa-
mination of Fig~ ~, in any event, illustrates the ease
with which the positive or negative dead band may be in-
dependently adjusted, and several gains and limits may
be combined for either direction in any of the axes,
to provide a desired voltage characteristic with respect
to the force applied to the stick in either direction of
any axis. Similarly, by means of table look up, or by
means of calculations utilizing constants which are
looked up in a table, based upon the magnitudes of vol-
tage on the lines 20-23, the characteristics of the
type illustrated in Figs. 3 and 5 can be provided
digitally in a suitable digital computer, such as that
disclosed and claimed in United States Patent 4,270,168.
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Referring to Fig. 2, the conditioned signals on
the lines 28-31 are fed to a plurality of amplifiers
32-39, the amplifiers 32 35 being proportional ampli-
fiers, and the amplifiers 36-39 being integrating
amplifiers. The amplifiers 32-39 therefore provide
proportional plus integral gain of the pilot input to
the control surfaces of the aircraft. Each of the
amplifiers provides an output on a corresponding line
40-47 which are summed in related summing junctions
50-53 along with corresponding negative feedback sig-
nals on related lines 54-57. The output of each sum-
ming junction is a positional error signal on a related
line 60-63 that drives a suitable amplifier 64-67 which
in turn controls the electromagnetic ~alve 70-73 of a
hydraulic servo 74-77. Three of the servos 74-76 drive
mechanical inputs to a mixer 84 which in turn controls
the mechanical inputs 86-88 to a swash plate 90 that
in turn controls the pitch of the blades of a main
rotor 92. The yaw servo 77 drives the mechanical con-
nection 94 to a pitch beam 96 that controls the pitch
of the blades of the tail rotor 98.
~ach of the servos 74-77 is provided with a cor-
responding position sensor 100--103 that provides an
electric signal on a related line 104-107 indicative
of the position of the mechanical output 80-82, 94 of
the corresponding servo. These signals are applied
through related amplifiers 108-111, for proper scaling
and isolation, to the feedback lines 54-57. At any
given moment of time, each of the servos is at a
particular position, and if a position is being com-
manded that differs therefrom, this will manifest it-
self as a signal on one of the lines 60-63 which,
through the amplifiers 64-67, will provide mzgnetic
force in the electromechanical valve 70-73 to divert
the valve and create an imbalance in the servo so that
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hydraulic fluid under pressure applied by a conduit 112
from a source 113 will move the hydraulic pist~n and
thereore the mechanical outputs 80-82 and 94 for the
desired action~ All of the servo and helicopter ap-
paratus 64-113 is conventional in nature. However, the
servos 74-77 must comprise high speed, full authority
electrically controlled servos, rather than the elec-
trically trimmed, mechanical booster servos of the type
heretofore used in the art to control the aircraft sur-
faces. Servos suitable for use in incorporating thepresent invention are readily available.
Considering operation of one of the axes of the
control system illustrated in Fig. 2, the new mode of
flight control becomes apparent. For instance, should
the pilot desire greater collective pitch, he will
press upwardly on the stick so as to provide an elec-
trical signal on the vertical axis output 20 as a
function of the amount of force that he exerts verti-
cally on the stick. This signal will be levèl-converted
in accordance with the signal conditioning circuitry
24 (i.e. that shown by way of example in Fig. 5) to
provide a pilot command signal on the line 28. Instan-
taneously, the proportional amplifier 32 will amplify
the signal on the line 28 and apply it on the line 40
as an input to the summing junction 50. This will
automatically cause an imbalance in the output of the
summing junction 50 since the servo 74 cannot move the
mechanical linkage 80 instantaneously, and therefore
the position sensor lO0 will be providing a signal over
the line 54 to the summing junction indicative of the
instantaneous original position of the mechanical link-
age 80. Thus, the summing junction 50 will provide a
signal on the line 60 which then is amplified by the
amplifier 64 and causes an imbalance in the electro-
magnetic valve 70 to cause the servo 74 to drive thelinkage 80 in the desired direction. The servos 74-77
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are selected to be capable of moving the control sur-
faces through 100% o~ their authority in a very short
time, on the order of one second. Depending upon the
gains of the signal conditioning circuitry 24 and the
amplifiers 32, 64, some pressure exerted by the pilot
can result in a signal of sufficient magnitude at the
electromagnetic valve 70 so that the servo 74 will apply
maximum hydraulic pressure to its piston and thereby
maximum accelerating force to the mechanical linkage 80.
On the other hand, if the pilot utilizes a small signal,
the initial proportional component of that signal which
is passed through the proportional amplifier 32 through
the summing junction 50 and the amplifier 64, may be
only slight and therefore only cause a nudging of the
piston within the servo 74.
A system only employing proportional gain, as
has just been described, would work perfectly fine except
for the fact that the pilot would have to conti~uously
maintain a force that would equal the desired position
of the mechanism for balancing with the feedback signal
on the line 54 (for example), even during long-term
flight with no changes in the control surfaces. This
could obviously result in fatigue over many tens of
minutes. And, the fatigue is worse in that the forces
must be applied in several axes (four, if the invention
is employed in a four-axes mode), all at the same time.
Under initial consideration, the foregoing fati-
gue problem would appear to be readily resolved by a
trimmed system of the type used in conventional aircraft
controls. In such systems, the controls are positioned
until the aircraft flight parameters are as desired,
and then the various controls are trimmed to their cur-
rent positions. This sort of trim position holds the
control stic~, wheel or pedal in a physical relationship
with respect to the aircraft that represents the
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desired corresponding position of the control surfaces
to which they are attached. When the pilot desires to
alter the position of the control surfaces with respect
to any of the controls, he reengages the particular
control at the position where it has been held in trim.
He can then move the control against the spring detent
or the like to a different position and reengage the
trim, or as is the usual case, he can release the trim
with respect to the particular control which he desires
to adjust, move it to a new position and then reengage
the trim~ However, it is literally impossible to trim
one axis at a time in a three or four axes, single
handed force stick. ~his is due to several factors:
first, in a force actuated, proportional system, if
the trim engage is effected by means of buttons on the
single-handed controller itself, the mere movement of
a thumb or a finger to engage the button will alter the
forces in one or more of the axes, so that trim is
effected at an undesirable point; second, it is almost
impossible to reestablish in the force stick the force
command to which trim may have been engaged, when it
is desired to release thetrim and utilize the force
stick to establish a new trim point - e~en if force
meters were provided, release of trim in three or four
axes at one time would require balancing the actual
force in the stick to the trim force by visual com-
parison, which would be nearly impossible; and, third,
actual trimming of the force stick itself would require
a highly exotic micro-sensitive position or farce
holding servo mechanism in each of the four axes, thus
mitigating all the advantages which a side-arm con-
troller can provide by introducing new complexities to
the system. Finally, timed base trim by ramping
electronic signals, to maintain the same actuator
position as the force stick, is not practical because
~: , : ~ .. , , :
2 ~ 9
- 18 -
the pilot must remove his force gradually to match the
trim ramp. For all practical purposes, this is an im-
- possible task. Any mismatch between the pilot's
removal of force from the controller and the ramp
causes unacceptable aircraft transients. Further, the
difficulty of adjusting all four axes at one time during
high pilot workload maneuvers, such as ground-related
maneuvering of a helicopter (e.g., loading ships),
take off or landing of any aircraft in high cross winds,
and the like, is aggravated when forces in three or
four axes all must be sirnultaneously adjusted by a
single hand.
Another aspect of the invention is providing
follow-up to the commands given by the multi-axis force
stick. In one example of the invention, the follow-up
is provided by feed forward integral gains provided by
the amplifiers 36-39 in parallel with the proportional
amplifiers 32-35. Thus, in the example of operation
given hereinbefore, once the pilot provides a force
input which indicates a desired change in the position-
ing of the linkage 80, the instantaneous effect is
created by a signal to the proportional amplifier 32,
as described hereinbefore. But before the servo 74
can reach a position to cause the feedback on the line
54 to equal the proportional command on the line ~0,
the integrating amplifier 36 will commence to have an
increasing output on the line 44 of the same polarity
as the signal on the line 40. The integrating amplifiers
36-39 are provided with time constants so as to
be able to assume the entire pilot input in a time frame
which is commensurate with the pilot's reaction to
aircraft response, which is on the order of a second
or so. Thus, in a typical case, if the pilot dasires
to trim the control surface by some amount, a very
slight input thereto, which is immediately backed-off,
' :
2 ~ ~
.. - 19 -
may achieve the desired result since the servo 74 will
initially respond to the proportional signal on the line
. 40 and the steady state condition will be quickly reached
by a signal on the line 44 balancing the feedback sign~1
on the line 54. In the case of a desire for a large but
slow change in the position of a control surface, the
pilot may provide a very small force so that the signal
from the stick on the line 20 is very small, and the
signal to be integrated by the integrating amplifier 36
may be commensurately small. However, if the pilot
continues to apply a small force over a period of time,
. the continuing presence of the signal on the line 20
. will cause the integrator 36 to continually build up
its output (up to a limited maximum, as described
hereinafter) so that the signal on the line 44 can
easily exceed, by several orders of magnitude, the
signal on the line 40. This would cause the servo 74
to continue to slew the position of the linkage 80 until
: the feedback signal on the line 44 matches that provided
by the proportional gain on the line 40 and the integral
gain on the line 44.
In practice, it has been found that the combination
of a force transducer (with imperceptible motion) and
the proportional plus integral control over the-servo
in response to the applied force permits the pilot to
apply a force until he senses a desired response and to
then reduce the force back to zero as the integral gain
.~ portion of the system balances up with the feedback sig-
: nal. Thus, each of the four axes depicted in Fig. 2
have a floating trim point wherein each servo mechanism
74-77 has caused the positioning of the corresponding
: mechanical linkage 80-82 and 94 to a position where
the related feedback signal on a line 54-57 balances
with the integral gain signal on the line 44-47. Con-
trol over the aircraft is, at all times, in a mode in
which the pilot adjusts this floating trim position in
1~8219
- 20 -
any axis by providing a commensurate force in the
desired direction for a sufficient period of time and
with a sufficient magnitude to achieve the.desired change
in the floating trim point for that axis, at the desired
speed of change. The overall effect from the pilot's
viewpoint is that there is a unique trim point, namely,
zero force on the controller (actually, force levels
within the dead-band region). Further, the specific
full scale or saturation level of the force.-controller
tends to be de-emphasized, since the integral control
produces a control surface velocity (typically, air-
craft acceleration command) for any constant force
application.- Hence, the pilot does not have to apply
full controller force to attain maximum aircraft
maneuvering. This replaces the maximum maneuver for
full control application found in conventional dis
placement controllers. Thus, with the invention, the
pilot can fly with a loose feel on the stick, or hands
off in steady-state flight. Because of the.possibility
of producing large commands by integrating.very small
signals provided by the force stick, it is essential
that the signal conditioning means provide a dead band
for each polarity of each axis of the stick.
i In Fig. 2, the line 31 in the yaw channel is con-
nected to an additional integrating amplifier.117
which provides an integral of the twist force on a
line 118 to wheel steering mechanism 119. This is not
: essential to the invention, but is illustrative of the
. fact that, if steering pedals are eliminated in an
; 30 aircraft (such as to provide earth visibility around
the pilot's feet and to reduce control system weight),
the stick could be used for ground steering as well as
in-flight maneuvering.
At the top of Fig. 2, a line 114 provides a signal
indicative of the fact that the aircraft has touched
1 1~82~.~
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down, that is, that a wheel or skid is in contact with
the ground. Such a signal ma~ be provided by a "squat
switch", or derived in some other fashion from the
wheel or skid support mechanism on the aircraft. Such
a signal is commonly provided in many aircraft for a
variety of purposes, such as preventing the operation
of automatic flight control stability equipment while
the aircraft is on the ground. The signal on the line
114 is applied to each of the integrating amplifiers
1~ 36-39 and is connected to operate as an integrator hold
signal: depending upon the implementation of the in-
vention, this ~ignal may simply deactivate an elec-
tronic switch connected in the integrator feedback path
so as to isolate the integrating capacitor from the
input to the amplifier. Thus when the aircraft touches
the ground, the floating trim point is held constant at
the moment, and the pilot then completes maneuvering
solely through the proportional path. When the air-
craft is shut down, the floating trim points are all
electrically reduced to zero, either by specific
initiali~ation resets, or otherwise, as is within the
skill of the art. Then, when operation of the aircraft
is resu~edl the signal on the line 114 holds all of the
integrators at ~heir initialization value, which is
zero. Therefore, any stray controller inputs provided
during taxiing or while par~ed will not cause any
command integrations to occur. Thus, it is assured
that the trim point of all of the control surfaces is
at the neutral position during takeoff, so that no
unwanted control inputs can exist at the start of
takeoff. Takeo~f is therefore effected bv the pilot
through the proportional loop alone. The signal on
the line 114 is also provided to an inverter 116 that
causes complementary operation to the integrating
amplifier 117 used for steering the aircraft while on
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- 22 -
the ground (if such apparatus is required).
Referring now to Fig. 6, an indication of power,
or authority remaining may be required in systems em-
ploying the present invention. In conventional systems,
S the mechanical linkage actually moved by the pilot as he
maneuvers a stick, lever, wheel or pedal includes
position responsive means to activate warnings of the
fact that the limit of authority in a given axis has
been reached. As a substitute for such position-responsive
means~ electronic means may be provided as illustrated
in Fig. 6. For instance, the summation of the propor-
tional and integral outputs may be provided by a summing
circuit 50a which does not have the position feedback
of the line 54 added to it, thereby providing a desired
position command signal on a line 60a. This can be
compared in a summing junction 120 with a suitable
reference voltage indicative of 100% authority for the
given channel, such as from a source 122 9 to provide a
signal on a line 124 indicative of the remaining authority.
This signal may operate a meter 126 to provide a constant,
quantitative indication of remaining authority to the
pilot, and may also be applied to a level detector 128,
the output of which on a line 130 will be indicative of
the fact that 90% (or some other fraction) of total
authority is currently being commanded in that axis.
This may be combined, as in an OR circuit 132, with
discrete indications of reaching the threshold limit of
authority in other axes, such as provided on lines 134,
to gene~ate a caution signal on a line 136 that can oper-
ate a caution light 138 as well as a stick shaker 140 or
other conventional alarm. The stick shaker in combina-
tion with the warning lights and power (or authority)
remaining indicator replaces the control stop feel a~d
control meters used to warn pilots of reaching the ~ull
limit
.
~8~
- 23 -
of authority (the control stops). In fact, having the
force stick shake as the particular axis approaches
a limit is more desirable as a warning cue than waiting
to reach the stop as in positional systems.
The invention, employing the particular force stick
described hereinbefore with respect to Fig. 1, has been
successfully used in controlling a liyht helicopter. In
that embodim~nt, the signal conditioning circuits 24-27
had characteristics as described hereinbefore with
respect to Figs. 3-5. The gains of the amplifiers
32-39 were adjusted so as to provide response times
which are in the one-half to two second range. For
instance, the constant Kc for the integrating amplifier
36 was selected as 1.25, and, with a maximum force input
applied vertically on the force axis stick 10, so that
a maximum voltage appeared on the line 20, the minimum
time for full travel of the servo 74 in eithsr direction
was about 1.5 seconds. The constant Kp in the amplifier
37 was selected as 0.5, and provided a minimum time for
full travel of the servo 75 in either direction of about
2 seconds. The constant KR in the amplifier 38 was
selected as 1.0, and provided a minim-m time, for full
travel of the servo 76 in either direction, of about 1
second. And, the constant Ky in the amplifier 39 was
selected as 1.25 and provided a minimum time, or full
travel in either direction of the servo 77, of about
0.8 seconds. The gains are relative to the gain of
the corresponding proportional channel; however each
of these gains are adjusted in dependence upon the gain
relationship provided by the signal conditioning
circuits 24-27 and the characteristics (such as the
servo mechanism gain) provided in the remainder of the
system, all as is well known in the art.
The description thus far has been essentially in
terms of analog controls, employing amplifiers having
suitable gains, limits and integrating characteristics,
2 ;1 g
- 24 -
and the summing of analog voltages to drive the servo
valves. However, the invention rnay as well be practiced,
and in many cases will be preferably practiced, in a
system in which the signal conditioning, the integrating,
the summing and the like are all provided by one or more
digital computers. One example is the dual computer
system disclosed in the aforementiond U.S. patent. To
practice the invention in an aircraft utilizing such com-
puters, the voltage outputs of the force stick 10 would
be accessed through various multiplexed inputs to the
analog to digital converter in Fig. 1 of said application
and the electromagnetic valves 70-73 would be driven as
is illustrated in Figs. 1 and 2 of said patent. Obviously,
if a twin computer system were used, both computers would
be connected in each of the axes. On the other hand, it
is of course possible to use only a single computer, if
desired.
The signal processing, as alluded to briefly
hereinbefore, could either be so:Lely by table look up,
or a combination of table look up for constants followed
by calculations employing the constants. All of the
digital techniques required for implementing the func-
tions described hereinbefore with respect to Fig. 2
are well known in the art, being currently utilized in
various systems for similar aircraft control but not
control in the new mode provided by the present invention.
The invention is readily used with automatic
flight control systems, such as autopilots which control
altitude, speed and heading, and such as stability
augmentation systems that compensate for external inputs
to aircraft attitude, such as by wind gusts and the
like. The coupling of automatic flight control systems
to an aircraft control system employing the invention
is rendered quite simple, since the flying to a trim
115~2~
- 25 -
point is already achieved with the invention, the trim
point being correctable by the automatic pilot as a
function of gyro outputs, and being stabilized by the
stability augmentation systems as a function of rate
gyro outputs. For instan~e, the autopilot functions
could be summed to the input of the appropriate inte-
grating amplifiers 36-39, and the stability inputs
could be summed into the proportional amplifiers 32-35
or into the summing junctions 50-53. This would cause
the autopilot trim point to coincide with the actual
trim point memorized in the integral path of the system.
On the other hand, both autopilot functions could be
simply summed in the summing junctions 50-53, if desired,
in such case 9 the deviations corrected for by the auto-
pilot would be electrical input signals indicative of
the offset from the trim point established in each axis
by the integral path thereof. In either case, the elec-
trical signals from the automatic flight control equip-
ment must be suitably conditioned to take into account
the differences between a positional, mechanical system ~ -
of the conventional type and the system described herein.
For instanc~ the magnitude of stability signals should
be kept low, on the order of 5% or 10% of authority, and
the autopilot signals should have a limited rate of change,
although operating with full authority. If the automatic
stability signals are summed in after the integral path,
and added to the floating trim point of the invention,
the trim point for stability signals should be conti-
nuously updated by the autopilot signals, so as to permit
the mean point of limited stability authority to follow
the variations in the autopilot trim point. And, all
the techniques known and used for such automatic flight
control systems are directly applicable to the present
invention, requiring no other special considerations
when being employed in an aircraft control system accor-
ding to the invention.
- 2~ -
The invention has been described pxincipally in
terms of a rotary wing aircraft (helicopter). However,
the principles of the invention are equally applicable
to control systems utilized for fixed wing aircraft.
In the case of fixed wing aircraft, the longitudinal
axis would control the elevator, the lateral axis would
control the ailerons~ and the twist axis would control
the rudder. The vertical axis could be used to control
speed and/or lift (e.g., engine thrust or propeller
pitch), to suit any particular implementation of the
invention, as desired, Naturally, the time constants
and signal conditioning derived for such system would
be based on the skill of the art in providing servo
control to the aerodynamic surfaces of a fixed wing
aircraft. However, there is nothing special that need
be taken into account when providing a fixed wing air-
craft control system in accordance with the invention,
other than is described hereinbefore.
If desired, the function of the mechanical mixer
8~ could be performed by electric signal combinations,
in a fly by-wire system incorporating the invention.
Then, the signals would directly drive the main servos
in the swash plate 90 (not sho~n), which would be
electromagnetic (or the like) rather than mechanical.
And, the four axes of the force stick would not then
have a one-to-one correspondence with any particular
servo. The significant fact is that the invention pro-
vides integral and proportional control over an aero-
dynamic axis of the aircraft in response to force inputs
to a corresponding axis of a stick having at least three
axes.
Similarly, although the invention has been shown
and described with respect to exemplary embodiments
thereof, it should be understood by those skilled in
the art that the foregoing and various other changes,
omissions and additions may be made therein and thereto,
without departing from the spirit and the scope of the
invention.
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