Heading hold logic

The automatic heading retention mode of operation of an automatic flight control system for a rotary wing aircraft is discontinued and reestablished through the use of a logic circuitry responsive to the yaw rate, bank angle and airspeed of the aircraft and to the force applied to the cyclic pitch control member by the pilot. The logic circuitry provides for the isolation of information commensurate with left and right turns and summation of the yaw rate and bank angle in each direction to determine if automatic disengagement of the heading hold is warranted.

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 the operation thereof, 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 heading 
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 automatic control mode, for example a heading hold 
mode, any deviations from the desired heading 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 systems 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 automatic 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 by 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 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 discriminae between transient conditions, for 
which it is desired to compensate utilizing the heading hold control 
circuitry, and those conditions where the pilot wishes to disengage the 
heading hold control so as to execute a maneuver. 
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 techinque. In accordance with this invention, the heading 
hold mode of operation is automatically discontinued, and heading error 
signals thus isolated from the actuator portion of a flight control 
system, through the operation of a unique combination of logic circuit 
elements. The automatic disengagement of the heading hold mode will be 
accomplished when the sum of parameters corresponding to the actual bank 
angle and yaw rate exceed a predetermined minimum and the pilot is 
exercising some affirmative action such as, for example, applying a 
lateral force in excess of a predetermined minimum to the cyclic pitch 
stick. In accordance with a preferred embodiment, the heading hold mode 
will also be discontinued when a signal commensurate with a pilot 
initiated command is received from the cyclic pitch trim release switch or 
a roll trim switch; either of these signals operating in lieu of the 
signal commensurate with the force applied to the cyclic pitch stick to 
cause the desired discontinuance of the heading hold operative mode when 
the sum of the bank angle and yaw rate exceed the predetermined minimum. 
The input signals commensurate with a pilot commanded heading change, 
which are employed to disengage the heading hold mode of operation, are 
preferably generated separately for left and right turns. Also, the 
automatic disengagement of the heading hold operative mode will, in 
accordance with a preferred embodiment, be permitted to occur only when 
the forward component of airspeed of the aircraft is above a preselected 
minimum. 
The present invention also contemplates the utilization of the previously 
employed yaw pedal switches to disengage the heading hold mode when the 
aircraft is operating at a forward speed below the preselected minimum 
whereby the pilot may assume total manual control below this preselected 
airspeed by placing his feet on the yaw control pedals. 
In addition to the left and right turn logic isolation, and to the 
summation of the yaw rate and bank angle to disengage the heading hold 
mode in each turn direction, the present invention contemplates delaying 
the application of input signals commensurate with cyclic pitch stick 
force so as to insure that the system will be responsive to steady state 
input signals. Additionally, the cyclic pitch stick force, roll trim 
switch operation or cyclic pitch trim release signals are preferably 
delayed in the interest of imparting stability to the system. In order to 
accomplish trimming of the aircraft upon selection of a new heading, the 
removal of input signals derived from the roll trim switch may also be 
delayed.

Description of the Preferred Embodiment: 
With reference now to the drawing, means for accomplishing the automatic 
disengagement and reengagement of the automatic heading hold mode of 
operation of an automatic flight control system is depicted in functional 
block diagramm form in FIG. 1. As is well known in the art, the yaw 
channel of an automatic flight control systemm provides automatic heading 
retention and rate damping. For a somewhat more detailed description of a 
yaw channel of an automatic flight control system for rotary wing 
aircraft, reference may be had to copending application Ser. No. 649,331 
of Ronald E. Barnum filed Jan. 15, 1976 and assigned to the assignee of 
the present invention. 
A primary input to the flight control system yaw channel is derived from a 
directional gyro, not shown, which may be of the magnetically slaved type. 
The directional gyro or compass senses aircraft heading and, subsequent to 
selection of a desired heading, provides an output signal commensurate 
with heading error. In the heading hold mode of operation this error 
signal is delivered to a servo system wherein it is employed to maintain 
the aircraft on the desired heading; the control system operating to null 
the heading error signals. 
The flight control system yaw channel also receives, as an input, rate of 
turn signals from a yaw gyro. These yaw turn rate signals are employed as 
inputs to the heading hold logic of the present invention. In order to 
facilitate understanding of the invention, the yaw rate gyro has been 
depicted at two places in FIG. 1 and identified by reference numeral 10. 
The yaw rate gyro 10 will sense the rate of heading change and provide a 
signal commensurate with both the direction and rate of aircraft yaw. 
Additional inputs to the heading hold logic circuitry of FIG. 1 are derived 
from a roll vertical gyro 12, a cyclic pitch force sensor 14, and an 
airspeed sensor 16. As in the case of yaw rate gyro 10, roll vertical gyro 
12 has been depicted at two places on FIG. 1. Similarly, the cyclic stick 
force sensor 14 has been indicated as two separate inputs since the force 
sensor is capable of discriminating between lateral stick forces applied 
to the left and right. Additional inputs to the heading hold logic are 
provided from left and right "beeper" switches 18 and 20 resprectively, 
which are provided for use by the pilot in trimming roll attitude, and a 
cyclic stick trim release switch 22, through which the pilot overcomes 
cyclic stick forces in order to provide manual control of the pitch and 
roll axes of the aircraft. The final inputs to the heading hold logic 
circuit consists of signals provided by the closing of the yaw pedal 
switches 24 and signals generated by the airspeed sensor 16 in conjunction 
with a comparator 70. 
The yaw turn rate signals from yaw rate gyro 10 are applied to a pair of 
polarity sensitive circuits 30 and 32. As noted above, the magnitude and 
polarity of the output signals from yaw rate gyro 10 are indicative of the 
direction and rate of aircraft yaw; these signals being represented 
graphically within the circuits 30 and 32. The polarity sensitive circuit 
30 will provide an output signal commensurate with the rate of yaw in a 
left-hand direction while the polarity sensitive circuit 32 will provide a 
signal commensurate with the rate of yaw in the righthand direction. The 
output signals from polarity sensitive circuits 30 and 32 are respectively 
applied to summing circuits 34 and 36. 
The roll vertical gyro 12 will provide an output signal having a magnitude 
commensurate with the bank angle and a polarity corresponding to the 
direction of bank; i.e., left or right bank angle. The output signal from 
roll vertical gyro 12 is applied to a further pair of polarity sensitive 
circuits 38 and 40. The input signals to polarity sensitive circuits 38 
and 40 from roll gyro 12 are graphically depicted within each of the 
circuits. The polarity sensitive circuit 38 will provide an output signal 
commensurate with the magnitude of bank angles to the left while the 
polarity sensitive circuit 40 will provide an output signal commensurate 
with the magnitude of bank angles to the right. The output signals 
provided by the polarity sensitive circuits 38 and 40 are respectively 
applied as second inputs to summing circuits 34 and 36. 
The cyclic stick force sensor, which will be described in greater detail 
below in the discussion of FIG. 2, provides input signals to a pair of 
level comparator circuits 42 and 44. Circuits 42 and 44, which may 
comprise operational summing amplifiers configured as level comparators 
and time delay, provide for the delayed passage, to the remainder of the 
heading hold logic, of signals commensurate with the application of 
pressure by the pilot to the cyclic pitch stick. There will, however, be 
no delay upon the "drop out" of the stick pressure. The output of 
comparator circuit 42, which is a signal commensurate with pressure 
applied in the left-hand direction to the cyclic pitch stick, is delivered 
to an OR gate 46; circuit 42 providing an output signal commensurate with 
a 1 or 0. Similarly, comparator circuit 44 provides an output signal, 
commensurate with the presence or absence of pressure applied to the 
cyclic pitch stick in the right-hand direction, to an OR gate 48. The 
delay imparted to the stick force signals by level comparator circuits 42 
and 44 insures that the system will not respond to inputs of insufficient 
duration to cause an aircraft heading change. 
Second inputs to each of OR gates 46 and 48 are provided by the cyclic 
stick trim release switch 22. As noted above, the pilot will operate trim 
release switch 22 when manual control is desired thus releasing the 
attitude hold logic. 
Third inputs to each of OR gates 46 and 48 will respectively be derived 
from beeper switches 18 and 20. Switches 18 and 20 are roll trim switches 
which are operated by the pilot, in the manner known in the art, to make 
single axis attitude changes. Signals commensurate with the closing of 
switches 18 and 20 are respectively applied to the associated OR gate 
through either of delayed deactivation circuits 50 and 52. Circuits 50 and 
52 provide for the immediate application of signals commensurate with the 
closing of switches 18 and 20 on their respective OR gates but delay the 
drop out of the signals upon the opening of the beeper switches. The 
delayed drop out is incorporated in the system to allow sufficient time 
for the aircraft to establish a trimmed bank angle and turn rate; this 
being brought about by the autopilot roll attitude hold function after the 
roll attitude reference has been displaced by beeping action. 
The final input to both of OR gates 46 and 48 is a signal, generated in the 
manner to be described below, indicative of the existence of conditions 
which warrant the disengagement of the heading hold mode of operation. 
This fourth input signal to each of OR gates 46 and 48 thus, also in the 
manner to be described, functions to logically "lock" out the heading 
retention function of an automatic flight control system in the disengaged 
condition until such time as the yaw rate and bank angle signals have 
decayed to a level indicative of a return to a straight flight path. 
The signals passed by OR gates 46 and 48 are respectively delivered, via 
delay circuits 54 and 56, to respective switching circuits 58 and 60; 
switches 58 and 60 respectively being the left and right turn switches. 
The delay circuits 54 and 56 are incorporated in the system to reduce the 
susceptibility of the logic circuitry to noise. Considering the left turn 
logic only, the sum of the yaw rate and bank angle signals, as appears at 
the output of summing circuit 34, will be passed by switch 58 to a 
comparator 62 when switch 58 is in the closed condition. Switch 58 will, 
as will be apparent from the discussion above, be closed when pressure in 
excess of a predetermined minimum is laterally applied, in the left-hand 
direction, to the cyclic pitch stick for a period of time in excess of the 
delay established by level comparator circuit 42 or when either of the 
left beeper switch 18 or cyclic stick trim release switch 22 has been 
closed by the pilot. Thus, switch 58 will be closed in response to some 
positive action on the part of the pilot and, upon the closing of the 
switch, the summation of signals commensurate with the yaw rate and bank 
angle will be delivered to comparator 62. Comparator 62, which may be an 
operational amplifier level detector, will compare the input signal 
delivered thereto from summing circuit 34 with a preselected bias and will 
provide an output signal if the sum of the yaw rate and bank angle exceeds 
a predetermined minimum level. Comparator 62 thus provides, at its output 
terminal, either a 1 or a 0 respecitively indicative of the necessity of 
disengaging the heading hold operational mode or the lack of need for any 
action to disrupt the existing operational state. A comparator 64 in the 
right-hand logic performs the same function as comparator 62 with respect 
to left turns. As noted above, the outputs of comparators 62 and 64 are 
respectively delivered to OR gates 46 and 48 to cause the latching of 
switches 58 and 60 in the closed position. The outputs of comparators 62 
and 64 are also applied as the inputs to a further OR gate 66. 
Signals passed by OR gate 66, indicative of the existence of conditions 
which require the disengagement of the heading hold operational mode of an 
automatic flight control system, are delivered as a first input to AND 
gate 68. The second input to AND gate 68 is a signal, in the form of a 1 
or 0, indicative of whether the forward airspeed of the aircraft is above 
or below a preselected level. The binary signal indicative of airspeed is 
generated by a comparator 70 which may comprise an operational amplifier 
level detector, if the signal sensed is a voltage proportional to 
airspeed, or simply a set of electrical contacts connected to a bellows 
sensing the difference between dynamic and static pressure and closing at 
a predetermined pressure. Comparator 70 receives, as its input, an analog 
signal as provided by the airspeed sensor 16. The output stage of 
comparator 70 comprises a bistable circuit whereby a 1 will be delivered 
to AND gate 68 when the airspeed is above the preselected minimum value 
and a 1 will be delivered to a further AND gate 72 when the airspeed is 
below the preselected level. In a typical example the minimum airspeed 
will be 60 knots. When the aircraft is operating above 60 knots and either 
of comparators 62 and 64 is providing an output signal both inputs to AND 
gate 68 are in the positive or 1 state and gate 68 will provide a signal 
to OR gate 74. This signal will be passed by gate 74 and will operate, in 
the manner known in the art, to disengage the heading hold operational 
mode. The disengagement may be accomplished by providing the output of OR 
gate 74 to a yaw synchronizer to cause the synchronizer to switch to the 
"synchronized" condition. With the yaw synchronizer, which is a subsystem 
employed in prior art heading hold controls, in the synchronized state the 
output of the directional gyro will be employed to position the rotor of a 
control transformer within the synchronizer so as to hull the transformer 
output. During the time the synchronizer is being internally nulled, there 
will be no directional error signal delivered to the servo system coupled 
to the controlled autopilot mechanism. When the synchronizing signal from 
OR gate 74 is removed from the yaw synchronizer, the internal control 
transformer output will be in a nulled state and a new heading reference 
will have been established. The automatic control system will now use this 
new heading reference to guide the aircraft. 
In accordance with the disclosed embodiment of the present invention, at 
airspeeds below the preselected minimum value, the heading hold 
operational mode will be terminated only when the pilot places his feet on 
the yaw pedals. This action on the part of the pilot closes the yaw pedal 
switches 24 thereby delivering a second input to AND gate 72. This second 
input, with the airspeed below the preselected minimum value, will be 
delivered to OR gate 74 and thence to the yaw synchronizer. Thus, the 
disengagement of the heading hold mode of operation at airspeeds below the 
preselected minimum value will be accomplished when the pilot places his 
feet on the yaw pedals. This manner of operation is desirable because, 
during low speed or hovering flight in a helicopter, heading changes are 
not accomplished by establishing roll bank angles but by yawing the 
heading with the tail rotor control via the pedals. 
Referring now to FIG. 2, a cyclic pitch stick is depicted schematically at 
80. Stick 80 includes, mounted in the handle portion 82 thereof, a cyclic 
trim release switch 84; switchh 84 controlling among other functions the 
engagement and disengagement of a solenoid operated clutch 86. The clutch 
86 has been shown in two sections to facilitate understanding of the 
invention. Clutch 86, when in the normally engaged condition, serves to 
couple the movements of a pair of shafts 88 and 90 respectively to 
transducers 92 and 94. Transducers 92 and 94, in a preferred embodiment, 
comprise AC linear transducers. The current supply for transducers 92 and 
94 has been omitted from the drawing. Shaft 88 is directly coupled to the 
cyclic pitch stick 80 while shaft 90 is coupled to stick 80 via a spring 
96. Shaft 90 is directly coupled to the piston of a hydraulic stick trim 
actuator 98; pressurized fluid delivered to actuator 98 under the control 
of an electrohydraulic servo valve 102 positioning the piston as a 
function of the roll control stick trim requirements. The output of 
transducer 94, which is a trim position transducer, is applied as a first 
input to a summing circuit 100. The output of transducer 92, which is the 
control position transducer, is delivered as the second input to summing 
circuit 100. The outputs of transducers 92 and 94, which will be of 
opposite polarity, result from defluctions of the spring 96 caused by the 
application of force to the control stick 80. The output of summing circut 
100 is delivered to a comparator 102. Comparator 102 may comprise an 
operational amplifier with a preset trip level. In the manner to be 
described below, comparator 102 will provide a pair of output signals 
resprectively commensurate with whether the force applied to the cyclic 
pitch stick is in the left or right-hand direction. Thus, elements 84-100 
of FIG. 2 correspond to the cyclic stick force sensor 14 of FIG. 1 while 
comparator 102 is in fact equivalent to the level comparator circuits 42 
and 44 of FIG. 1. 
The operation of the cyclic stick force sensor of FIG. 2 is as follows: The 
force sensor produces an output by sensing the difference in output 
motions of transducers 92 and 94. There will be no output if the motions 
are equal which must be the case when the spring 96 is not compressed. Any 
commpression of spring 96 will be reflected as a difference in motion of 
transducers 92 and 94 and an output proportionate to force will thus be 
generated. The clutch 86 provides a means for initializing to various 
flight trim positions and giving transducers 92 and 94 a new nulled output 
position thus reducing system error stackup. 
As should now be obvious to those skilled in the art, the present invention 
comprises novel and improved heading hold logic which functions to control 
the disengagement and engagement of the heading hold mode of an automatic 
flight control system. In accordance with the disclosed embodiment of the 
invention, heading hold is disengaged by yaw pedal switches at low speeds, 
where the pedals are customarily employed to control heading, and by the 
lateral force applied to the cyclic pitch stick or the operation by the 
pilot of the cyclic trim release switch or bank angle beeper switches at 
high speeds, where lateral cyclic inputs control heading. The invention 
contemplates the employment of time delays to reduce the susceptiblility 
of the control to noise and to eliminate transients. The present invention 
is characterized by left and right turn logic isolation and by summation 
of the yaw rate and bank angle to determine if automatic disengagement of 
the heading hold is warranted. 
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.