Automatic heading synchronization control system

The automatic heading retention mode of operation of a flight control system for a rotary wing aircraft is discontinued and reestablished through the use of logic circuitry responsive to the bank angle of the aircraft and to the force applied to a control stick by the pilot. When the bank angle of the aircraft exceeds a preselected minimum and the lateral force applied to the cyclic pitch stick also exceeds a predetermined minimum the logic circuit operates to isolate the heading hold control from heading error signals and the heading retention mode of operation is automatically resumed when both of the bank angle and stick force fall below the preselected minimum levels. The invention may also encompass the application to the control mechanism of signals which will cause the aircraft to perform coordinated turns above a preselected minimum airspeed.

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
The present invention relates to facilitating the exercise of control over 
aircraft and particularly rotary wing aircraft. More specifically, this 
invention is directed to means for automatically controlling the 
engagement and disengagement of the heading hold mode of an automatic 
flight control system. Accordingly, the general objects of the present 
invention are to provide novel and improved methods and apparatus of such 
character. 
2. Description of the Prior Art 
While not limited thereto in its utility, the present invention is 
particularly well suited for incorporation in an automatic flight control 
system for helicopters. When flying a helicopter a pilot must manipulate 
three separate control elements; i.e., the collective pitch stick, the 
cyclic pitch stick and the yaw pedals. Since a helicopter is an inherently 
unstable aircraft, manual flight control requires constant manipulation of 
these controls to maintain a predetermined attitude. Accordingly, to 
facilitate their use, automatic flight control systems for rotary wing 
aircraft have been devised. An early version of such an automatic flight 
control system is shown and described in U.S. Pat. No. 2,845,623. 
Automatic flight control systems for helicopters include, in a yaw 
"channel", means for implementing a heading hold mode of operation. When 
in the heading hold mode the aircraft will fly "hands off" along a linear 
course selected by the pilot. A heading hold mode of operation, and the 
apparatus which permits such operation, are well known in the art and will 
be briefly discussed below in the description of the preferred embodiment 
of the present invention. 
When the aircraft is in an automatic control mode, for example a heading 
hold mode, any deviations from the desired attitude will be sensed and 
appropriate corrective action automatically taken. In order to insure 
rapid and accurate aircraft response to a pilot input when the automatic 
flight control system is engaged, for example when an evasive maneuver is 
suddenly required, it is necessary that provision be made for 
disengagement of the automatic control or selected functional modes 
thereof. Such disengagement must be accomplished both automatically and 
quickly in order to insure that, when manual control is resumed, the 
automatic flight control system will not be resisting the input commands 
generated by the pilot. To this end, considering the heading hold mode of 
a helicopter automatic flight control system, sensor switches have 
customarily been provided on the yaw pedals. In order to assume manual 
control and disengage the heading hold mode, it was necessary for the 
pilot to place his feet on the yaw pedals thus actuating the pedal 
switches and thereby generating disengagement control signals. The 
employment of yaw pedal switches, however, required that the pilot have 
his feet on the pedals to make a turn and precluded the positioning of the 
pilot's feet on the yaw pedals during the heading hold mode. The necessity 
for the pilot to reposition his feet in order to shift between the manual 
and automatic control modes resulted, among other disadvantages, in a time 
delay incident to the resumption of manual control. 
As a further requirement of an automatic flight control system for rotary 
wing aircraft, it is necessary that the circuitry which supervises the 
automatic engagement and disengagement of the heading hold mode be capable 
of distinguishing between a pilot input and a sudden comparatively large 
magnitude attitude change such as might be incident to a wind gust. Thus, 
the aforementioned disadvantages incident to the employment of yaw pedal 
mounted sensor switches can not be overcome merely by sensing the 
magnitude of an attitude change, such as a bank, since such attitude 
magnitude change sensing would be unable to discriminate between transient 
conditions, for which it is desired to compensate utilizing the heading 
hold logic, and those conditions where the pilot wishes to disengage the 
heading hold logic so as to execute a maneuver. 
As a further complicating factor, aircraft, and particularly rotary wing 
aircraft, have a threshold speed of operation above which maneuvers such 
as turns should preferably be coordinated. As employed herein the term 
"coordinated turn" is synonymous with a bank turn, as opposed to a flat 
turn, and is a turn wherein the aircraft does not exhibit any lateral 
acceleration as characterizes a slip or skid. Thus, the yaw channel of the 
heading hold logic of an automatic flight control system should, above a 
preselected air speed, be capable of commanding and controlling automatic 
coordinated turns when in the manual mode; i.e., with the heading hold 
disengaged. In such a coordinated turn the yaw control will be 
automatically repositioned to produce the proper turn for the bank angle 
established by the pilot through manipulation of the cyclic pitch stick. 
Thus, should the pilot institute a maneuver, the heading hold mode should 
be automatically disengaged but sufficient control retained so as to 
produce a coordinated turn if the air speed is above a preselected level. 
Below this preselected air speed the disengagement of the heading hold 
mode should result in the pilot having total manual control. 
SUMMARY OF THE INVENTION 
The present invention overcomes the above-described disadvantages of the 
prior art by providing a novel and improved technique for engaging and 
disengaging the automatic heading hold mode of operation of an aircraft 
flight control system and by providing apparatus for use in the practice 
of that novel technique. In accordance with this invention the heading 
hold mode of operation is automatically discontinued, and heading error 
signals isolated from the actuator portion of a flight control system, 
when a preselected bank angle is exceeded and a lateral force in excess of 
a predetermined minimum is applied to a pilot operated control lever 
which, in the case of a helicopter, will be the cyclic pitch stick. Upon 
discontinuance of the heading hold mode the aircraft will, except during 
pilot commanded turns above a preselected airspeed, be under manual 
control. The manual control mode will be continued until both the bank 
angle and stick force fall below their preset minimum values. 
Also in accordance with the invention, turns executed by the pilot at 
speeds in excess of a preselected minimum will be coordinated by 
automatically generating and delivering signals to the flight control 
system actuator which oppose any tendency of the aircraft to slip or skid. 
These signals are generated by a lateral accelerometer and a gyroscope 
sensitive to aircraft roll. A switching circuit permits delivery of these 
sensor generated signals to the actuator only when the control is in the 
manual mode and the airspeed exceeds the selected minimum value.

DESCRIPTION OF THE PREFERRED EMBODIMENT 
With reference now to the drawing, a yaw channel for an automatic flight 
control system is depicted in functional block diagram form. The yaw 
channel provides, as will be described in detail below, automatic heading 
retention and rate damping. The yaw channel shown in the drawing also 
causes the aircraft to automatically perform coordinated turns in response 
to heading change commands initiated above a preselected air speed. 
The inputs to the flight control system yaw channel of the present 
invention include a directional gyro 10 which may be of the magnetically 
slaved type. The directional gyro or compass 10 senses aircraft heading 
and, subsequent to selection of a desired heading, provides an output 
signal commensurate with heading error. 
A yaw rate gyro 12 generates a second input signal for application to the 
yaw channel of the present invention. The yaw rate gyro 12 senses the rate 
of heading change and develops a signal commensurate with the direction 
and rate of aircraft yaw. In the manner to be described below, and as is 
conventional in the art, the rate signal provided by yaw gyro 12 is phased 
such that it opposes any heading change thereby damping the response of 
the aircraft to heading error influences. 
Additional inputs to the yaw channel depicted in the drawing are derived 
from a roll vertical gyro 14, a lateral stick force sensor 16, a lateral 
accelerometer 18, a roll rate gyro 20 and an air speed sensor 22. The 
outputs of the lateral accelerometer, roll rate gyro and air speed sensor 
are utilized during coordinated turns. The roll vertical gyro and lateral 
stick force sensor generate input signals which are applied to a heading 
synchronization logic circuit, indicated generally at 24, to automatically 
control engagement and disengagement of the heading hold mode of 
operation. 
During heading retention; i.e., in the automatic heading hold mode; heading 
error signals provided by the directional gyro 10 are coupled, via a 
control transformer in a yaw synchronizer 26, to a summing circuit 28. As 
will become obvious from the discussion below, the yaw synchronizer 
functions to null heading error signals when a new heading is desired. The 
yaw synchronizer, which includes a control transformer and a motor for 
positioning the transformer rotor, is part of previously available 
automatic flight control systems and does not comprise part of the present 
invention. The output of the control transformer in the yaw synchronizer 
26 will be zero at the time of engagement of the heading hold operational 
mode. In the heading hold mode, deviations from the selected flight path, 
as indicated by a change in the output of directional gyro 10, will result 
in a heading error signal appearing at the output of the yaw synchronizer 
26. The output of synchronizer 26 will, in the heading hold mode, be 
proportional to the difference between the actual heading and the desired 
heading; the desired heading being the heading at which synchronizing 
ceased. Heading error signals from synchronizer 26 are coupled, via 
summing circuit 28, to an amplifier 30. The output of amplifier 30 is 
delivered as the input to a servo drive motor 32. The servo driver 32 
commands the rotation of the yaw channel servo valve 34 thereby resulting 
in a displacement of a power piston and the flight controls connected 
thereto. The aircraft will, accordingly, be turned in the proper direction 
until the output of directional gyro 10 restores the yaw synchronizer 
control transformer output to null. 
As noted above, the yaw rate gyro 12 provides rate damping about the yaw 
axis. To accomplish such damping, the output of yaw rate gyro 12 is also 
applied to summing circuit 28 in such a manner as to oppose any heading 
change command signals appearing at the output of the yaw synchronizer 26. 
It will be understood by those skilled in the art that the yaw channel, as 
described above, is in accordance with the prior art. It will also be 
understood that the heading hold mode is energized by the pilot in 
switching the automatic flight control system to the "on" condition. In 
accordance with the present invention, additional inputs to the yaw 
channel are provided by the roll vertical gyro 14 and the stick force 
sensor 16. The output of the roll vertical gyro is applied to a comparator 
36 in the heading synchronization logic circuit 24 wherein it is compared 
with a bias signal commensurate with a preselected roll attitude as 
established by the setting of a potentiometer R1. This preselected roll 
attitude may, for example, be a bank angle of approximately 2.degree.. 
Comparator 36 has a pair of outputs which will, respectively, be at a high 
level depending on whether the signal corresponding to the actual bank 
angle is greater or less than the bias signal. Comparator 36 may, for 
example, comprise an operational amplifier, which receives the bias and 
sensor generated input signals, and a monostable multivibrator. Typically 
the output of the amplifier will be maintained at zero potential until the 
input from gyro 14 exceeds the bias from potentiometer R1 at which time 
the amplifier output will go positive and the multivibrator will be set. 
The first output of comparator 36, which will be in a "high" state when 
the output of the roll vertical gyro 14 indicates that the preselected 
bank angle has been exceeded, is applied as a first input to AND gate 38. 
The other output of comparator 36, which will be in the "high" state when 
the bank angle commensurate with the setting of potentiometer R1 is 
greater than the actual bank angle, will be applied as a first input to an 
AND gate 40. 
The output of lateral stick force sensor 16 is applied as a first input to 
a second comparator 42. Comparator 42 may be identical to comparator 36. 
The sensor 16 will comprise a standard force measuring device coupled to 
the cyclic pitch stick. The output of stick force sensor 16 will be 
compared, in comparator 14, with a bias voltage commensurate with a 
preselected minimum force; the minimum force bias voltage being developed 
at the wiper arm of a second potentiometer R2. When the output of force 
sensor 16 exceeds the preset bias, indicative of the pilot having moved 
the cyclic pitch stick incident to commanding a turn, a first output of 
comparator 42 will go "high". This first output of comparator 42 is 
applied as a second input to AND gate 38. The second output of comparator 
42 will go "high" when the output of force sensor 16 is less than the 
preselected bias; this condition indicating that any force applied to the 
cyclic pitch stick has resulted from outside influences for which the 
automatic flight control system must compensate. This second output of 
comparator 42 is applied as the second input to AND gate 40. 
The outputs of AND gates 38 and 40 are respectively delivered as the "set" 
and "reset" inputs to a bistable multivibrator circuit 44. The 
multivibrator or flip-flop 44 will thus be set when the actual bank angle 
exceeds the preselected minimum and the force applied to the cyclic pitch 
stick also exceeds a preselected minimum. Conversely, flip-flop 44 will be 
reset when the bank angle returns to a magnitude less than the preselected 
minimum and the force applied to cyclic pitch stick is less than the 
preselected minimum. The flip-flop 44 provides, at its output terminal 
when in the set condition, a control signal which is delivered to yaw 
synchronization circuit 26 and to an air speed switch 46. 
The control signal from flip-flop 44, when applied to the yaw 
synchronization circuit 26, will turn the heading hold of the automatic 
flight control system to the "synchronized" condition. With circuit 26 in 
the "synchronized" state, the output of directional gyro 10, as sensed at 
the secondary of the control transformer in synchronization circuit 26, 
will be amplified and employed to drive the motor in circuit 26. This 
motor will reposition the rotor of the control transformer so as to null 
the control transformer output. Accordingly, at the time the synchronizing 
signal from flip-flop 44 is removed from yaw synchronization circuit 26, 
the control transformer output will be nulled and a new heading will have 
been selected. Similarly, during the synchronized condition, because of 
this nulling action, no error signal will be delivered to summing circuit 
28 from directional gyro 10. During the manual control mode; i.e., with 
circuit 26 in the "synchronized" state; the output of yaw rate gyro 12 
will continue to be applied to the servo amplifier 30 via summing circuit 
28 and, as the aircraft turns, the yaw rate signal will provide rate 
damping. 
The yaw channel of an automatic flight control system in accordance with 
the present invention also causes the aircraft to automatically perform, 
when in the manual mode and above a preselected speed, coordinated turns; 
i.e., turns wherein the aircraft does not exhibit any lateral 
acceleration. An input signal from airspeed sensor 22, having a magnitude 
proportional to the forward speed of the aircraft, is applied to a further 
comparator 48 wherein it is compared with a bias signal derived from 
potentiometer R3. The bias signal will, in a typical example, be selected 
to correspond to 60 knots. Comparator 48 will provide an output control 
signal, which is delivered as a second input to air speed switch 46, 
whenever the speed of the aircraft is above 60 knots. When both input 
signals are present; i.e., when the air speed is above the preselected 
level and the output of heading synchronization logic circuit 24 is 
indicative of the manual operational mode, circuit 46 will be switched 
from its normally open state to the closed state. This permits the signals 
from the lateral accelerometer 18 and roll gyro 20 to be applied, via 
summing circuit 28, to the input of amplifier 30. 
Continuing with a discussion of the automatic coordinated turn mode, the 
roll rate gyro produces an output signal proportional to the aircraft roll 
rate; i.e., the gyro 20 provides a signal which is employed to anticipate 
the turn and aid in establishing the proper tail rotor pitch when 
commencing the turn. Thus, the output signal from gyro 20 will be phased, 
in the manner known in the art, such that it will aid in repositioning the 
yaw servo valve 34 to initiate the proper turn for the bank angle 
established by the pilot through movement of the cyclic pitch stick. When 
the aircraft stops rolling the roll rate gyro output decays to zero and 
the output of the lateral accelerometer 18 will coordinate the turn rate 
to the bank angle. The lateral accelerometer 18 senses lateral 
acceleration during a turn, indicative of the aircraft slipping into the 
turn or skidding out of the turn, and provides an output signal having a 
magnitude proportional to the "error" and a phase commensurate with the 
direction of the "error". This acceleration "error" signal, via the yaw 
servo system, drives the tail rotor pitch to correct the "error". 
Accordingly, should the aircraft slip or skid during a turn executed at an 
air speed in excess of 60 knots the resulting output developed by the 
lateral accelerometer 18 will cause repositioning of the yaw servo 34 to 
reestablish the coordinated turn. The outputs of lateral accelerometer 18 
and roll gyro 20 are summed in a summing circuit 50 prior to being passed 
to summing circuit 28 via switching circuit 46. 
To briefly summarize, at a speed above a preselected level, for example 60 
knots, should the pilot desire to turn the aircraft to a new heading he 
will manipulate his controls whereby the output signals from cyclic 
lateral stick force sensor 16 and roll vertical gyro 14 will exceed the 
preselected bias levels thus causing the heading hold to be disengaged and 
switch 46 to be closed. Upon closing of switch 46, turning of the aircraft 
will be coordinated; i.e., the turn rate will be coordinated to the pilot 
selected bank angle by application of the outputs from either or both of 
lateral accelerometer 18 and roll gyro 20 to the input of servo amplifier 
30. At this time, due to disengagement of the heading hold, heading error 
signals are nulled and thus there will be no heading error input to 
summing circuit 28. 
While a preferred embodiment has been shown and described, various 
modifications and substitutions may be made thereto without departing from 
the spirit and scope of the invention. Accordingly, it is to be understood 
that the present invention has been described by way of illustration and 
not limitation.