Note: Descriptions are shown in the official language in which they were submitted.
Description
.
Automatic Lock-Positioning
of Foldable Helicopter Blades
Technical Field
This invention relates to positioning, or adjusting
the pitch angle of foldable helicopter blades prior to
locking the blades in their pitch axis to enable folding
the blades near the.helicopter fuselage during nonuse
of the helicopter, and more particularly to automati-
cally positioning the pitch of the.main rotor blades to
enable them to be locked in position prior to.folding.
.
Background Art
In certain helicopters used for specific applica-
tions, such as helicopters based on seagoing ships, it
has long been known to fold the main rotor blades to
positions adjacent the fuselage of the helicopter during
nonuse of the helicopter. This facilitates storage of
each helicopter in a relatively small space as well as
rendering helicopters which are stored in the open less
vulnerable to wind gusts and the like during storage.
In order to fold the blades, it is necessary that
each blade assume a predetermined position with respect
to the blade fold hinge, and with respect to the
fuselage of the aircraft. Therefore, before folding
blades, the main rotor is indexed to a.predetermined
position which puts all of the blades.on the rotor in
a position where each may be folded to a position
alongside the fuselage. Thereafter, the pitch angle
of each of the blades is adjusted to a desired position
and the pitch angle is locked by. means of pins, so that
the pitch angle of the blades will not thereafter vary
as the blades are beins folded.
S-~13
~-3~56
13
As is kn~wn, the p~tch of the main rotor blades of
a helicopter is adjusted by push rods which are urged
against, and therefore raised and lowered by a swash
plate which can tilt varying amounts in any azimuthal
direction. I~ is the tilting of the swash plate which
causes the blades to achieve the nominal collective
pitch with the desired varying cyclic pitch superposed
thereon. As the blades rotate about the main rotor in
flight, the push rods connected to the blades and
rolling on the swash plate assume various positions in
- dependence on the tilt of the swash plate and the
azimuthal position of the rotor. Thus it is blade
motion as the blades rotate which actually achieves the
variation pitch in dependence upon the then-current
position of the swash plate. Therefore, adjusting the
pitch of the rotor blades prior to folding requires
positioning of the swash plate, in a fashion similar
to that achieved by the pilot controls and/or automatic
flight control system during flight.
In the earliest systems, the pitch lock pins were
generally displaced by hydraulic pressure, and the pitch
positions of the blades were slewed back and forth by
operation of manual controls (such as the cyclic pitch
stick and the collective pitch stick) until each blade
had passed by the pitch lock. The pin was able to snap
into place and thereby prevent further pitch change of
the blade. However, on-the-fly snap-in of locking pins
results in excessive wear. Furthermore, hydraulically
actuated lock pins are cumbersome and impede the ability
to properly design a rotor head for a helicopter. Elec-
tric motor actuated pins, on the other hand, are well
suited to rotor head design, but require that the pins
be given a sufficient time to engage the bladPs while
they are held in the proper position. This would have
required the use of pilot indicators to show the pilot
13
~ 3 --
correct blade pitch positions for pin engagement, the
pilot moving the controls very slowly to achieve indica-
tions and to provide minute adjustments in pitch position
once the indicators are lit until pin engagement was
achieved. A motorized blade fold lock of a modern type is
disclosed in a commonly owned, U.S. patent 4,284,387 of
Ferris entitled BLADE FOLD RESTRAINT SYSTEM.
Objects of the invention include automatic posi-
tioning of helicopter blade pitch angle to facilitate
folding of the blades.
According to the present invention, an automatic
flight control system of a helicopter provides commands
to the pitch, roll and collective pitch trim actuators to
thereby position the pitch angle of the main rotor blades
for locking, prior to folding the blades. In further
accord with one embodiment of the present invention, the
pitch, roll and collective trim commands are provided in
response to stored pitch, roll and collective trim refe-
rence values which were determined to be correct in a
prior blade folding operation, and further in response
to reference values of swash plate servo positions which
were determined to be correct in a prior blade folding
operation. In still further accord with the present in-
vention, the swash plate servo positions stored from a
previous blade folding operation are utilized as the
desired goal for positioning in response to the pitch,
roll and collective trim commands stored from the pre-
vious blade folding operation, and the pitch, roll and
collective trim commands are updated, by using an inverse
mixer matrix, if necessary in order to achieve positioning
of the swash plate as indicated by the swash plate servo
positions stored in a previous blade folding operation.
In accordance with one embodiment of the in-
vention there is provided a helicopter having foldable
main rotor blades, pitch angles of which are positionable
by push rods in dependence upon the vertical position and
tilt of a swash plate cooperating with the push rods.
; --
- 3a -
The blades are lockable at specific pitch angles as a pre-
requisite to folding. The swash plate is positioned by a
plurality of servos~ each of the servos being seperately
operable in response to corresponding outputs of a mixer.
The mixer, in turn, receives inputs from pitch, roll and
collective channels, each of the channels including an
electrically operated trim actuator for providing a corres-
ponding input to the mixer in response to a related trim
command signal from signal processing means. In accordance
with the invention there is provided a position detector
for each of the servos, each position detector providing
a servo position signal indicative of the position of a
corresponding one of the servos. The signal processing
means comprises means for storing a plurality of pitch,
roll and collective reference signals related to predeter-
mined positions of the servos at which rotor blade pitch
angle is correct for locking, and for providing trim command
signals to the pitch, roll and collective trim actuators
in respective response to the pitch, roll and collective
reference signals. Thus, the mixer is caused to drive the
servos substantially to the predetermined positions as in-
dicated by corresponding ones of the servo position signals.
According to a further embodiment of the present
invention, an automatic flight control system provides
commands to the pitch, roll and collective trim channels
of the main rotor pitch controls in response to the posi-
tions of the swash plate servos at a point in time just
prior to unfolding the helicopter blades, the information
being stored in the form of the deviation of such posi-
tions from nominal positions that are stored in permanent
read-only memory and need not be stored in nonvolatile me-
mory. In further accord with the present invention, mo-
vement of the desired swash plate position is achieved by
calculation of the pitch, roll and collective commands
necessary to be provided to the mixer (which converts
these three commands into separate and distinct commands
relating to the forward, aft and lateral swash plate servos),
~14-~3
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thereby eliminating the need to store flight commands uti-
lized in one blade positioning operation for use as a
starting point in a subsequent blade folding operation.
ln further accord with the present invention, an integra-
tor is used in each of the swash plate servo channels to
determine a desired swash plate position in order to
achieve desired blade pitch for locking, to eliminate long
term, steady state errors therein, the pitch, roll and
collective trim commands being generated in response to
the respective swash plate servo position integrators.
In accordance with another aspect of the invention, po-
sitioning of the pitch of the tail rotor blades, to pro-
vide a known coupling between the tail rotor blades and
the longitudinal cyclic pitch (or pitch) axis of the main
rotor, is achieved in a system in which the pitch trim
position detector does not reflect actual pitch, but only
pitch relative to the last synchonization of yaw trim
position.
1~14:i3
The present invention avoids the necessity for the
pilot to continuously move the collective stick and the
cyclic stick while awaiting pins to randomly fall into
locks, which causes excessive wear. The invention
overcomes the need for the pilot to hold the collective
and cyclic sticks in desired positions as indicated by
control panel indicator lights while slower pin locking
mechanisms may engage the pins. The invention provides
a competent manner of prepositioning the pitch angle of
the main rotor blades so as to allow use of relatively
slow pin engagement means, without wear.
The present invention may be implemented in a
variety of ways utilizing apparatus and techniques
which are well within the skill of the art, in the
light of the teachings which follow hereinafter. Sim-
ilarly, the foregoing and various other objects, fea-
tures and advantages of the present invention will
become more apparent in the light of the following
detailed description of an examplary embodiment thereof,
as illustrated in the accompanying drawing.
il~;l4 ~3
Brief Description of Drawings
Fig. 1 is a simplified schematic block diagram of
a helicopter blade pitch control system, employing an
automatic flight control system computer, in which the
present invention may be practiced;
Flg. 2 is a simplified logic flowchart of a portion
of an automatic flight control system program in which
routines related to folding the blades may be reached
when the aircraft is on the ground;
Fig. 3 is a simplified logic flowchart of an
executive routine for controlling the operation of an
automatic flight control system during blade fold
operations;
Fig. 4 is a simplified logic flowchart of a blade
fold background program, from which computer interrupts
may reach utility programs in practicing the present
invention;
Fig. 5 is a simplified logic flowchart of a blade
fold position calculation routine;
Fig. 6 is a variation of the logic flowchart of
Fig. 2, which incorporates improvements of the present
invention;
Fig. 7 is a simplified logic flowchart of a routine
for retaining positions of locked blades for use in
subsequent blade locking operations;
Fig. 8 is an alternative to the simplified logic
flowchart of Fig. 3, incorporating features utilized
in the present invention;
Fig. 9 is an alternative to the logic flowchart
of Fig. 5, representing a routine for calculating
blade folding positions and commands in accordance with
the improvement of the present invention; and
Fig. 10 is a simplified logic flowchart of a rou-
tine for centering yaw pedals (tail rotor pitch angle)
in a system in which the tail rotor pitch position
indication is a function of the point of trim engagement.
l~i41~
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Best Mode for Carrying Out the Invention
Referring now to Fig. 1, a conventional helicopter
with suitable apparatus to permit practice of the pres-
ent invention includes main rotor blades 12, each of
which is pivotable about a longitudinal axis in response
to the position of a related push rod (not shown) which
contacts a swash plate 13. When the main rotor is
stationary, the degree of tilt and azimuthal position
of tilt of the swash plate 13 in combination with
vertical positioning of the swash plate determines a
position for all of the push rods, thereby establishing
a particular pitch angle for each of the blades. When
the rotor blades 12 are rotating (the swash plate is
always nonrotating), the push rods cause the pitch
angle of each of the rotor blades to vary in a cyclic
fashion as a consequence of rotation thereof relative
to the swash plate 13.
The swash plate 13 is connected by mechanical
linkage 14-16 to a plurality of servos 17-19 which are
spaced about the swash plate and which can therefore
control the positioning of the tilt axis, the degree
of tilt and vertical position of the swash plate 13.
Although spacing of the servos can vary from one case
to the next, in the example herein it is assumed that
the servos have the common configuration of a servo
17 being located forward of the rotor axis, a servo
18 being located aft of the rotor axis, and a servo
19 being located to the side of the rotor axis, they
being referred to herein as forward, aft and lateral
servos. In order to provide closed-loop feedback
control, and otherwise to determine the pitch angle of
all of the rotor blades (by determining the tilting of
the swash plate 13), a plurality of position detectors
20-22 are provided for the respective servos. These
position detectors may comprise potentiometers working
4i3
from a regulated power supply, or linear differential
voltage transformers, or any other suitable position
detector as is known in the art.
Each of the servos 17-19 has a mechanical input
member 23-25 from a mixer 26 which receives mechanical
inputs 27-29 from each of three flight control axes:
pitch, roll and collective (or lift). Depending upon
the particular helicopter in which the present invention
is practiced, these axes may have series actuators or
they may not. For instance, the pitch axis input 27
to the mixer 26 is provided by a pitch bias actuator
30 which is in series with a servo 31, whereas the roll
input 28 and collective input 29 are provided directly
by related servos 32, 33~ Or, the servos 31-33 may
have direct series inputs for automatic flight control
inputs, which can be accommodated in the same fashion
as the series actuator 30 (described hereinafter). The
servos 31-33 are typically boost servos which have
mechanical inputs 34-36 from the two axes of a cyclic
pitch stick 37 and from a collective pitch stick 38,
respectively. Thus motion of one of the sticks moves
the input to the servo, which hydraulically boosts the
motion so that the desired activity is achieved with
relatively low force applied to the sticks 37, 38.
The input to each of the servos 31-33 has a suit-
able electrically-controlled trim actuator 39-41 (either
electric motor or hydraulic) which is responsive to a
corresponding electric connection 42-44 from an auto-
matic flight control computer 45. Provision of
suitable trim command signals on the connections 42-44
by the computer 44 can cause the actuators 39-41 to
control the pitch angle of the main rotor blades,
thereby to adjust the angle when on the ground in order
to enable locking prior to folding of the blades, or to
control the flight profile of the aircraft when in
1~1413
g
flight. The automatic flight control com~uter may take
the form of one of the computers described in commonly
owned U.S. Patent 4,270,168, Murphy and Clelford, and en-
titled SELECTIVE DISABLEMENT IN FAIL-OPERATIONAL, FAIL-
SAFE MULTI-COMPUTER CONTROL SYSTEM. A computer of that
type has output and input connections, 46, 47 to and from
a control panel 48 which may include indicators 49 and
switches 50 to allow a pilot to interchange with the au-
tomatic flight control computer 45, and the capability of
receiving inputs from inertial devices such as accelero-
meters and gyros, and various position indicators. As
seen in Fig. 1, each of the position detectors 20-22 and
similar position detectors 51-54 (relating to corresponding
actuators and servos 30-33) may be connected into the au-
tomatic flight control computer 45 by a plurality of cor-
responding connections 55. The automatic flight control
computer 45 may interconnect with other apparatus of the
aircraft through multiplexed inputs and outputs which in-
clude analog conversion where necessary, all as is well
known in the art and as described in the aforementioned
Murphy and Clelford patent.
Although not shown further in detail herein,
the aircraft may also include a servo operated yaw channel
56 which is suitably connected to the automatic flight con-
trol computer by connections 57. This channel includes a
pitch beam for controlling the pitch angle of the tail
rotor blades, the pitch beam being positioned by a servo
in a well known fashion. As is described briefly here-
inafter, if the tail rotor is tilted, as is shown in
U.S. Patent No. 4,103,848, there may be coupling between
the yaw axis and the pitch axis of the helicopter which
requires some consideration in the blade folding operation
However,
1~i141~
-- 10 --
the coupling is not itself part of the present invention,
is conventional and known and is therefore not described
further herein.
The manner of carrying out the invention is de-
scribed using as an example one of the automatic flightcontrol computers disclosed in the aforementioned Murphy
and Clelford patent. In that patent, two identical
computers are disclosed as working together in a
particular fashion, but the utilization of one of them
to perform the intended functions (without the
inter-computer functions) is readily achieved in the
light of the teachings which follow hereinafter.
In the computer of the aforementioned patent, all
of the flight control functions are performed during
specific interrupts. To reach these programs, a general
background routine, referred to as a background (sG)
program is interrupted in a real time fashion, and each
interruption causes a particular sequence of utility
programs to be performed. The programs relate to
generating automatic pilot commands, stability commands,
bias commands, stick force commands and the like. The
programs also provide many functions to determine the
operational health of each computer and the health
status communicated to it by the other computer, to
determine the manner in which the two computers may
handle the work load. In one of the routines reached
in the aforementioned computer, functions to be per-
formed wlen the aircraft is on the ground are reached
in a third autopilot routine (AP 3) which is illus-
trated, as modified to practice blade folding, inFig. 2 herein.
In Fig. 2 the four-digit reference numerals are
the corresponding reference numerals found in Fig. 14
of the aforementioned Murphy and Clelford patent, the
two-digit reference numerals are peculiar to the
11~1413
present disclosure. In Fig. 2, the third autopilot
routine is reached through an entry point 1401 and
a test 1402 determines if the particular computer is
operating in a simplex mode or not. If it is not,
then both computers are operating together and, in
accordance with the twin computer reliability scheme
of the aforementioned Murphy and Clelford patent, a
pitch outer loop calculation 1403 and a collective
outer loop calculation 1404 may be performed. But in
the aforementioned patent, if only one computer is
operating, it is not permitted to perform potentially
disastrous functions such as operating the autopilot,
so that the calculations 1403 and 1404 are bypassed by
an affirmative result of test 1402. However, in an
embodiment of the invention employing only a single
computer, the test 1402 may be eliminated so that the
pitch and collective outer loop calculation routines
1403 and 1404 will always be performed. Of course, use
of a single computer requires other steps to determine
the reliability of computer operation. In fact, the
pitch outer loop calculation is a calculation which,
with pitch output routines (of said patent), will pro-
vide the pitch trim command signal on the connection
42 (Fig. 1 herein). Similarly, the collective roll
and yaw routines (including the collective outer loop
calculation routine 1404 and other routines of said
patent not shown herein) will provide the trim command
signals on the connectlons 43, 44 and 57 (Fig. 1), in
manners which are described hereinafter.
In Fig. 2, a test 1405 determines if the aircraft
is on the ground. This tests a status indicator bit
or word indicative of pressure on the helicopter
wheels, the rotor being locked, and other factors. If
test 1405 is negative, then the computer may perform
air null routines 1406 which reestablish the nulls of
13
- 12 -
various inertial sensors, and may perform other routines
not related to the present invention. Then a step 1408
may transfer failure and fault codes to a maintenance
display and the program will advance to other functions
through a real time return point 1409, which is the
manner of releasing the real time interrupt through
which the third autopilot routine of Fig. 2 is reached,
to return to a background program. A11 of the foregoing
is de-cribed in far more detail in the context of an
entire automatic flight control computer system in the
aforementioned Murphy and Clelford patent.
In Fig. 2, if the aircraft is on the ground, as
indicated by an affirmative result of test 1405, then
a test 60 will determine if the executive control mode
of the computer is set in a blade fold mode or not. In
the present example, the blade fold executive mode is
deemed to include the service executive mode (in contrast
with a nonservice mode and a maintenance mode, as de-
scribed in the aforementioned Murphy and Clelford
patent). The nature and purpose of this, and the manner
of establishing it are described with respect to Figs.
3 and ~, hereinafter. In a first instance, test 60
will normally be negative so that the blade fold
executive routine 61 (described with respect to Fig. 3
hereinafter) will be reached. Depending on how the
routine 61 proceeds, the computer may have its executive
mode switched into the blade fold mode, in which case
the routine 61 will lead to a blade fold background
program 62 by releasing the real time interrupt within
which the third autopilot routine of Fig. 2 has been
reached. Thereafter, the basic computer background
program is the routine 62, and all of the normal
computer functions are reached by interrupting the
routine 62 (in contrast with interrupting a general
background program, when the computer is in the service
~1413
- 13 -
mode, in which self health tests such as a check sum
test routine and a scxatch pad test are performed, as
illustrated in Fig. ~ of the aforementioned ~urphy and
Clelford patent). In Fig. 2, if the blade fold
executive routine 61 does not determine that the blade
folding operation is ready to proceed, it will lead to
other routines, such as ground null routines 1407, the
codes to maintenance display routine 1408 and then end
the real time interrupt through the real time return
10 point 1409. Eventually, the routine 61 may establish
the executive in the blade fold mode in which case the
test 60 will be affirmative, leading to a blade fold
position calculation routine 63 which is described with
respect to Fig. 5 hereinafter. This is the routine
that actually provides the pitch, roll and collective
trim commands necessary to position the pitch angle of
the main rotor blades to enable them to be locked in
anticipation of a folding opexation.
In Fig. 3, the b]ade fold executive routine 61 is
reached through an entry point 64 and a first test 65
determines whether a pitch lock enable flag has been
set yet or not. In a first pass through the routine of
Fig. 3, the pitch lock enable flag normally will not
have been set so that a negative result of test 65 will
reach a test 66. This determines if the rotor has been
indexed to the desired position for blade locking, as
indicated by a suitable flag bit, and whether the pilot
has activated a blade fold switch. Normally, the test
66 will be negative during a first pass through the
routine of Fig. 3 so that a plurality of blade fold
initiation steps will be reached, over and over, until
the rotor is indexed to the correct position and the
pilot has engaged the blade fold switch. The initiation
steps include a step 67 to reset a new values stored
flag, a step 68a to reset a swash positions initiated
1413
- 14 -
flag, a step 68b to reset a fold start flag (a local flag
utilized only in the routine of Fig. 3 as described
hereinafter), and steps 68c, 68d to reset a thirty
second timer and a ten second timer. Then, the pro-
gram will return to the third autopilot routine of Fig.3 through a return point 70. The short pass through
the blade fold executive routine 61 which merely pro-
vides the initialization steps 67-6~d will be performed
whenever the aircraft is on the ground unless and until
the pilot decides to initiate a folding operation by
first indexing the rotor to the correct azimuthal
position for blade folding, and thereafter performing
the very first step, which is to engage the blade fold
switch so that test 66 can be affirmative.
In Fig. 3, when test 66 is affirmative, a test 71
determines whether or not a local fold start flag has
been set. This flag simply ensures that certain
functions are performed once and only once in each
blade fold operation. Since the fold start flag is
reset in test 67 during pre-folding initialization,
the initial pass through step 71 is always negative.
This causes a test 72 to determine whether or not the
pilot has engaged a pitch position switch, which is
the second stage of blade folding controlled by the
pilot. If the switch has not been engaged, a negative
result of test 72 will reach a step 73 which provides
a signal that causes the indicators 49 (Fig. 1) to
indicate pitch position enable so that the pilot knows
that the sequence has reached the stage where he should
engage the pitch position switch if he wishes to con-
tinue with the blade folding operation. In such case,
the step 73 is the only step performed in the blade
fold executive routine 61 during the current cycle, and
the third autopilot routine is returned to through the
return point 70.
- 15 -
In a subsequent pass through the blade fold execu-
tive routine 61, the pilot will have eventually engaged
the pitch position switch so that step 72 is affirmative.
In such case, a step 74 causes the trim system to be
engaged (that is, so that the computer can cooperate
with the trim valves 39-41, Fig. 1, so as to adjust the
pitch position of the blades for the folding operation).
Then a step 75 resets the indicate pitch position enable
flag which was set in step 73 so the pilot knows that
his engagement of the pitch position switch has been
recognized. And a step 76 will set the fold start flag
so that, in subsequent passes through the blade fold
executive routine 61, test 71 will be affirmative,
therefore bypassing the steps 73-76.
In Fig. 3, once the fold start flag has been set,
a test 77 determines if the real time interrupt (the
interruption of the program which causes reaching the
third autopilot routine of Fig. 2) has been released;
if not, a step 78 causes releasing of the real time
interrupt. This requires simply enabling all inter-
rupts of the same or lower priority as that of the real
time interrupts; and, the program will simply advance
in a fashion that does not lead to an interrupt return
(so that the return to the normal background program
which is in effect when the executive mode of the
computer is in the service mode will not be reached).
Then the executive of the computer is set into the
blade fold mode by a step 79, and the blade fold back-
ground program is reached through a transfer point 80.
In Fig. 4, the blade fold background program 62,
reached through the transfer point 80, is basically a
closed-loop that can only be exited by taking the
executive out of the blade fold mode, essentially
ending the blade fold operation. This may be done as
a consequence of changes in operation or failures which
- 16 -
may occur, or as a consequence of having satisfactorily
completed the blade pitch positioning ~unction of a
blade folding operation (which is the only function
that the automatic flight control computer 45, Fig. 1,
perf~rms during blade folding). Specifically, the
blade fold background program 62 begins with a test 81
to determine if the blade folding operation is still
in progress, as indicated by the blade fold switch
still being engaged and the rotor still being indexed.
This is the same as test 66 in Fig. 3~ In the event ~lat
the rotor is inadvertently moved from its folding index
position, or if the pilot changes his mind and dis-
engages the blade fold switch, then a negative result
from test 81 will cause the blade fold background
program to advance to a step 82 to resynchronize the
trim system to desired trim positions, ihereby to
eliminate any blade positioning which may have occurred
as a result of the blade positioning routines being
performed for several cycles before the pilot changes
his mind. A step 83 causes the blade fold indication
to be reset, a step 84 sets the executive mode into the
nonservice mode and a step 85 causes the program to
branch to an initialization portion of the program at
or close to that which occurs for a power on reset
2S (which may be somewhere in the regime of steps 400-405
in Fig. 4 of the aforementioned Murphy and Clelford
patent), in any fashion which is suitable depending
upon the particular computer and implementation of the
present invention. The steps 82-85 effectively shut
down the blade fold positioning operation and cause the
automatic flight control system to reinitiate for
normal flight modes.
In Fig. 4, a second test in the blade fold back-
ground program 62 is test 86 which determines if any
of the tests which may be performed on the trim system
1 3
- 17 ~
have failed, causing a trim system failure flag to be
set. If there has been failure of the trim system, the
test 86 will reach a step 87 to set appropriate main-
tenance codes, which may depend upon the particular
nature of the failure, and cause the blade fold mode
to be ended by steps 82-85 as described hereinbefore.
Another test 88 will determine if blade fold posi-
tioning has been in process for less than thirty
seconds. If not, the positioning process has taken
too long and therefore cannot be completed because of
some condition of the helicopter external of the
computer or some inability of the computer to provide
correct positions. In such case, a negative result
of test 88 will reach a step 89 that sets an appropriate
failure code, and the blade fold operation is terminated
by steps 82-85 as described hereinbefore.
In Fig. 4, the normal way of exiting the blade fold
background program 62 is by means of a test 91 which
tests a flag (generated as described hereinafter) which
determines that the swash plate has been positioned
within tolerance and therefore the blade pitch angle
positioning for a blade fold operation has been suc-
cessfully completed. An affirmative result from step
91 leads to a step 93 which sets a pitch lock enable
flag; this is a flag that indicates successful
completion of blade pitch positioning, enabling pitch
lock motors to drive pitch lock pins so as to retain
the established blade pitch angle during the folding
operation. Then, the blade fold positioning operation
of the computer is terminated by steps 82-85 as
described hereinbefore.
Once the blade fold background program 62 of Fig.
4 has been entered by the transfer point 80, it will
generally continuously cycle through the tests 81, 86,
88 and 91, returning to the test 81, and so forth.
13
-- 18 --
This is a locked program loop which can be exited only
as a consequence of test results as described immediately
hereinbefore, or by means of program interrupts. The
program interrupts are real time interrupts that causes
the computer routine to jump out of the blade fold back-
ground program and to perform all of the normal utility
programs, including the third autopilot routine of Fig.
2 and the routines reached therein. Thus while the
automa~ic flight control system computer 45 is in fact
being utilized to provide commands on the connections
42-44 (Fig. 1) so as to drive the swash plate for
correct blade pitch angles to permit folding of the
blades, the utility programs, and particularly those
related to blade folding, are all reachable by means of
the normal real time interrupts. When these programs
have been completed in each cycle, the computer auto-
matically branches back to the blade fold background
program 62 by means of interrupt release, in the
normal fashion. As is apparent from a full under-
standing of all of the routines described hereinbeforeand hereinafter, the only function of the blade fold
background program is to monitor the desirability of
retaining the computer executive in the blade fold
mode. And, this is utilized in the third autopilot
routine of Fig. 2 simply to either cause passage
through the blade fold executive program 61 or passage
through the blade fold position calculation routine
63, as described hereinafter. Of course, other pro-
gramming arrangements could be selected to provide
similar functions, with or without utilizing a
background program. This in turn may depend somewhat
upon the particular automatic flight control system
computer utilized to implement the present invention,
in accordance with the skill of the art.
13
-- 19 --
Once the executive program has been set into the
blade fold mode by step 79 within the blade fold execu-
tive program 61 of Fig. 3, each pass through the third
autopilot routine of Fig. 2 will cause the test 60 to
be affirmative so that the blade fold position calcula-
tion routine 63 will be reached, through an entry point
95 in Fig. 5. A first test 96 determines if position
values have been initiated, by interrogation of the
flag reset in step 68a, Fig. 3. If not, a step 97
causes the stored values of pitch, roll and collective
references and the stored values of desired swash plate
servo positions to be read into the working portion of
the computer from nonvolatile read/write memory. These
are provided in the nonvolatile read/write memory at
the conclusion of a preceding blade folding operation
as is described with respect to Fig. 3, hereinafter.
Then, a series of steps 98 cause the pitch refer-
ence, roll reference and collective reference values,
to be utilized in generating trim commands on the con-
nections 42-44 (Fig. 1), to be respectively set equal
to the stored pitch, stored roll and stored collective
values which were determined in the previous blade
operation. Then a step 99 sets the position initiated
flag which was tested in step 96 so that in subsequent
passes through the blade fold position calculation
routine 63, the stepe 97-99 will be bypassed.
If any implementation of the invention includes a
series actuator, such as the pitch bias actuator 30,
Fig. 1, it may be desirable to force such actuator to
a known position so as to achieve repeatability of
swash plate positioning, without any adverse effect by
the series actuator. Therefore, a step 100 may, in
each pass through the blade fold position calculation
routine 63, cause the pitch bias reference value to be
equal to the center position of the pitch bias actuator
- 20 -
30 (Fig. l). This provides a simple method of utilizing
the regular pitch bias command generation (such as that
shown in the aforementioned Murphy and Clelford patent)
to center the series actuator in an open loop fashion.
Then a step lOl will increment the thirty second timer
which was previously reset in step 68c of Fig. 3. A
test 102 determines if the thirty second timer has been
incremented to less than its maximum value; if not, this
indicates that the blade fold position calculation has
been in process over many cycles spanning thirty seconds
in time, which is an indication that something is wrong.
Therefore a negative result of test 102 will lead to a
step 103 which sets a failure code, a step 104 which
provides a fault indication to the pilot, and a step
105 which forces the trim system to be engaged (so
that it can be resynchronized in the fashion described
with respect to Fig. 4 hereinafter as the necessary
consequence of an excessive time terminating the blade
fold position operation of the computer). In such case,
the program will revert to the third autopilot routine
of Fig. 2 through the transfer point 70.
In a normal case, the thirty second timer will not
have timed out, so the test 102 will be affirmative
leading to a step 106 which increments a ten second
timer. Then a test 107 determines if the ten second
timer has been incremented in a sufficient number of
cycles so as to have reached its maximum count or not.
If the setting of the ten second timer is less than
its maximum, an affirmative result of test 107 causes
the program to advance directly back to the third
autopilot routine by means of the return point 70.
This provides ten seconds within which successive
computer program cycles can utilize the reference
positions established in steps 98 to cause trim
commands to be generated and utilized, in a fashion
13
- 21 -
described with respect to Fig. 1 hereinbefore, so as to
position the swash plate 13 to desired reference values.
There is no point in determining whether or not these
reference positions have been reached until there has
been sufficient time for them to be reached. Since the
blade pitch positioning system has a response charac-
teristic of 10% of authority per second, a time frame
of ten seconds will ensure that the rotor blades could
be rotated from any blade pitch angle to a desired
blade pitch angle within the ten second time frame,
since 100% of authority would be encompassed. There-
fore, when the ten second timer has been incremented
sufficiently to reach its maximum count, test 107 will
be negative causing a test 108 to determine if the
series bias actuator, such as a pitch bias actuator,
has reached a position equal to a stored center position.
If not, a negative result of test 10~ will cause the
third autopilot program to be resumed through the return
point 70. But assuming that the series actuator can be
suitably positioned within the ten second time frame or
shortly thereafter, test 108 will eventually be affirm-
ative leading to steps 109, which determine the error
between the desired swash plate servo positions and the
stored swash plate servo positions which are believed
to be correct to enable blade locking to occur. Then
a series of tests 110 determine if all of the errors
in swash plate servo position are less than 0.2% of
the maximum range of positions. If any of them is not
within 0.2~, a negative result of one of the tests 110
will cause correction of the position by means of steps
111 and 112. In steps 111, pitch, roll and collective
corrections relating to the current error in the in-
dividual swash plate servos (forward, aft and lateral)
are generated by a matrix which is inverse to the
function of the mixer 26 (Fig. 1). In other words, the
13
- 22 -
constants Kl-K9 used in steps 111 to cause pitch, roll
and collective correction factors to be generated, are
those that indicate adjustments to the pitch, roll and
collective commands which will result in suitable ad-
justments to the forward, aft and lateral swash plateservo positions taking into account the effect that the
mixer 26 (Fig. 1) has in transposing the aircraft axis
commands to the swash plate servo commands. This is
referred to herein as an inverse mixer matrix. Then,
steps 112 cause the pitch, roll and collective reference
values, which were established initially in steps 98,
and which control the commands provided to the pitch,
roll and collective trim valves 39-41, to be updated
by addition thereto of the correction factors generated
in steps 111. Then the program can advance back to
the third autopilot program of Fig. 2 by means of the
return point 70.
Eventually, after several passes through the blade
fold position calculation routine 63 of Fig. 5, if the
factors being used and all of the equipment operations
are proper, the forward, aft and lateral swash plate
servos will be positioned within 0.2~ of the previously
stored values, indicating that blade locking can take
place. Therefore, the tests 110 will be affirmative
causing a step 113 to set a flag indicative of the
fact that the swash plate has been positioned within
tolerance; and, a step 114 will cause an indication
that blade lock is enabled, so that the pilot can
operate the blade lock pin insertion motors or other
means. And then, the third autopilot routine is returned
to by means of the return point 70.
In Fig. 2, completion of each pass through the
blade fold position calculation routine 63 will lead
to the codes to maintenance display routine 1408 and to
the real time interrupt return 1409. This causes release
- 23 -
of the real time interrupt so that the computer reverts
to the blade fold background program 62 as illustrated
in Fig. 4. Because the basic loop of the blade fold
background program routine consists only of four tests,
it is well assured that at computer speeds normally
encountered, all of these four tests will be made many
times before the next real time interrupt causes the
program to revert to the routines caused by inter-
ruptions. Thus it is well assured that test 91 will
be made, and since the swash in tolerance flag was set
in step 113, as described hereinbefore with respect to
Fig. 5, this test will be affirmative. Therefore, step
93 will set the pitch lock enable flag (which is a
significant advancement in the routine as described
hereinafter) and steps 82-85 will cause the computer
to pass out of the blade fold background program and
refer to the nonservice mode for reinitialization pur-
poses. Eventually, the third autopilot routine of
Fig. 2 will again be reached during one of the real time
interruptions, and test 60 will not be negative so that
the blade fold executive routine 61 (E`ig. 3) will again
be reached. In this case, test 65 in Fig. 3 will be
affirmative since the pitch lock enable flag has been
set in step 93 of Fig. 4. An affirmative result of
25 test 65 will reach a test 116 which determines if a
one-time local flag has been set, indicative of whether
or not the new values have been stored as yet. Ini-
tially, test 116 will be negative so that a plurality
of tests 117-119 may be reached to determine if all
of the blades have been locked in pitch angle. If any
of the tests 117-119 is negative, the computer will
revert to the third autopilot routine through the
return point 70. In subsequent passes through the
third autopilot routine (Fig. 2) which lead to the
blade fold executive routine 61 (Fig. 3), eventually
L41~
- 24 -
all of the blades should be locked so that all of the
tests 117-119 will be affirmative. This will lead to
a series of steps 120 which cause the present values of
pitch, roll and collective trim reference to be stored
for use in subsequent blade folding operations and
cause the final positioning of the forward, aft and
lateral swash plate servos to be stored for use in a
subsequent blade fold operation. These are the values
which are accessed by step 97 in the blade fold position
calculation routine 63 (Fig. 5). Once the steps 120
are completed, the new values stored flag is set in a
step 121. Therefore, in any subsequent passes through
the blade fold executive routine 61 in Fig. 3, when
the pitch lock enable flag is still set, step 116 will
be affirmative, so that no functions are performed by
the blade fold executive program.
Thus there has been described a system which posi-
tions the swash plate servos by means of autopilot trim
commands so that the blade pitch angle of helicopter
blades can be locked before folding. When the blades
are locked, the autopilot trim references used to
command the swash plate servos are stored for use in a
subsequent blade folding operation, along with the
final swash plate servo positions at which the blades
were ulti~ately locked.
In accordance with the invention, an improved
system for automatically positioning the pitch angle
of rotor blades to enable locking prior to folding of
the blades, stores the positions of the swash plate
servos which occur at the time of spreading (unlocking~
the blades, for use in a subsequent blade folding
operation. In the embodiment of the invention to be
described hereinafter, no autopilot trim references
are stored for use in subsequent blade folding
operations; instead, only the swash plate servo
ll~ 3
positions are stored, and the autopilot trim commands
required to cause the swash plate servos to reach such
positions are calculated from the stored servo positions.
This avoids the necessity of storing trim command
references in the nonvolatile read/write memory, space
within which is at a premium. In addition, in the
embodiment of the invention to be described, the
precise swash plate servo reference position information
is stored as the deviation from nominal position values.
Only the deviations need to be stored in the nonvolatile
read/write memory, the nominal values being, in a sense,
wired into the computer by being present in a read only
memory, space within which is not at a premium.
The blade fold positioning described with respect
to Figs. 2-5 hereinbefore all takes place during the
blade folding operation. In the embodiment of the
invention to be described, the storage of servo posi-
tion deviations occurs at a totally different time in
the history of the helicopters: that is, at the time
that the blades are to be unfolded or spread. In order
to achieve this, the third autopilot program is modified
as illustrated in Fig. 6 so that there are three dif-
ferent routes in which the third autopilot program can
proceed when the aircraft is on the ground. Specific-
ally, in Fig. 6, when test 1405 indicates that the
aircraft is on the ground, the test 60 can determine
whether the executive is in the blade fold mode or not.
In the normal case, it is not. This reaches a test
125 which examines a flag bit indicative of whether
the blades are now being unlocked or not. This flag
can respond to a spread command switch energized by
the pilot, or to some other function at an appropriate
point in a procedure of unlocking the blades after
respreading the blades for use. During a spread
operation, there will be one pass through the third
1 3L13
- 26 -
autopilot program where test 125 will be affirmative.
This will cause the program to advance to a locked
blade position retention routine 126, which is described
with respect to Fig. 7 hereinafter. But when the third
autopilot routine is reached with the aircraft on the
ground in other than a spread operation, test 125 will
always be negative causing the blade fold executive
routine to be reached in the manner described hereinbe-
fore. However, in the case of the embodiment of the
invention to be described, a blade fold executive
routine 61a, which differs in some respect from the
blade fold executive routine 61 illustrated in Fig. 3,
is required,and this is~described with respect to Fig.
8, hereinafter.
In Fig. 7, the locked blade position retention
routine 126 is reached through an entry point 130. A
series of steps 131 generates deviations of the forward,
aft and lateral servo positions from nominal values of
forward, aft and lateral positions found in a read only
memory. This is achieved by reading the values indi-
cated by the servo position detectors 20-22 (Fig. 1)
and subtracting therefrom the corresponding nominal
values read out of a read only memory. Then a series
of tests 132-134 check each of these deviations to
ensure that each is less than some maximum permissible
deviation. This may be done by comparing the absolute
value of each with an unsigned maximum deviation, or
it may be done by testing each to be sure that it is
not more positive than a positive value nor more
negative than a negative value, all as is well known
in the art. If any of the deviations is excessive, a
negative result from one of the tests 132-134 will
cause the locked blade position retention routine to
reach a step 135 which will set failure codes and a
step 136 which will provide an indication that there
- 27 -
is fault in the system. sut if all of the deviations
are less than the maximum permissible value, each of
the tests 132-134 will be affirmative so that a step
137 will be reached in which a check sum value for
the deviations is calculated by summing all three
deviations together. The deviations calculated in
steps 131 and the check sum calculated in step 137 are
all stored in nonvolatile read/write memory in step
138. These deviations will therefore be available at
a later time when the blades are to be folded, for use
in calculating the necessary automatic flight control
` system trim commands so as to reposition the pitch
angle of the blades to the positions they are at when
the deviations were stored. Then a step 139 provides
an indication to the pilot that the pitch angle posi-
tions for blade folding have been updated so that he
may engage a switch to cause the next sequence of blade
spreading to occur (which forms no part of the present
invention, and is not described further herein). Then,
the program reverts to the third autopilot routine of
Fig. 6 through a return point 140.
As the computer passes repetitively through a
third autopilot routine during normal operations, when
the aircraft is on the ground, when blade folding has
not reached the stage where the executive of the
computer ls set in a blade fold mode, and the blades
are not being unlocked, test 1405 is affirmative, test
60 is negative and test 125 is negative so that the
blade fold executive program 61a is reached as il-
lustrated in Fig. 8. The only differences betweenthe blade fold executive program 61a and the blade fold
executive program 61 illustrated in Fig. 3, are that
the storage of values for use in the next operation by
means of tests and steps 116-121 is eliminated, since
these functions are instead performed during blade
~141~
- 28 -
spread as described with respect to Fig. 7; and, instead
of resetting the new values stored flag in step 67 of
Fig. 3, the initialization includes setting a pedal com-
mand equal to zero in an initialization step 143. The
remainder of Fig. 8 is the same and performs the same
function as described herei.nbefore with respect to
Fig. 3.
In Fig. 8, the blade fold executive program 61a will
perform no function whatsoever once pitch positioning
has been complete and the blade fold background program
62 (Fig. 4) has reached step 93 and set the pitch lock
enable flag. This will cause test 65 (Fig. 8) to always
be affirmative and thereby bypass the remainder of the
blade fold executive routine 61a.
During the blade fold sequencing, the blade fold
executive program 61a will eventually reach step 79 (as
is described with respect fo Fig. 3, hereinbefore) so
that the computer has its executive set into the blade
fold mode. Then, in subsequent passes through the third
autopilot routine of Fig. 6, the blade fold position
calculation routine 63a is reached, as is described with
respect to Fig. 9 hereinafter. This routine includes,
in the exemplary embodiment described herein, a center
pedals subroutine 144, as is described with respect to
Fig. 10 hereinafter.
In Fig. 9, the blade fold position calculation
routine 63a is reached through an entry point 145 and
a first test 146 determines if the swash plate servo
positions have been initiated or not by interrogating
a once-only flag, set as described hereinafter. In
the first pass through the routine, the result of test
146 is negative so that a series of steps 147 is reached
to cause reading from nonvolatile memory the three servo
position deviations and their corresponding check sum.
Then in a series of StQpS 148, values of desired forward,
11~141~
- 29 -
aft and lateral swash plate positions are set equal to
the corresponding nominal positions which are provided
from a read only memory. In a test 149, the swash
plate position deviations read in step 147 are summed
together and the result is compared against the check
sum read in step 147, to see if there has been any
apparent error in the storage and retrieval of the
swash plate position deviations in nonvolatile memory.
If the data is still correct, then the deviations are
added to the desired positions in a series of steps 150.
But if the check sum test 149 indicates error, then the
steps 150 are bypassed and only the nominal values can
be used as desired values. This use of the nominal
swash plate positions causes the servos to position
the pitch angle of the blades close to the lock position,
to aid maintenance personnel in manually adjusting them,
upon initial computer operation, or after maintenance.
In either event, a step 151 sets the swash positions
initiated flag utilized in test 146, and the swash in
tolerance flag is reset in a step 151a. Therefore in
subse~uent passes through the routine 63a, the tests
and steps 147-151a will be bypassed and the program
will advance from test 146 directly to a step 152 which
increments the thirty second timer. A test 153 deter-
mines if the thirty second timer has been incremented
to less than its nominal value; if not, this means an
excessive amount of time has elapsed since the first
pass through the blade fold position calculation
routine 63a, so that a negative result of test 153 will
lead to steps 154-156 which set failure codes, indicate
a fault on the control panel,and cause the trim to be
engaged to enable re-synching of the trim after the
abortive conclusion of the attempt to position the
pitch angle of the blades for folding.
- 30 -
If the thirty second timer has not timed out, an
affirmative result of test 153 will reach a step 157
in which the ten second timer is incremented. A test
158 determines if the ten second timer has been
incremented to less than its maximum value. IL it
has, an affirmative result of test 158 reaches a
series of steps 158a to initiate integrator registers
to the desired swash plate servo positions, and thence
to the center pedals subroutine 144, which is described
with respect to Fig. 7, hereinafter. When the sub-
routine 144 has been passed through, a series of steps
159 generate the inverse matrix of the mixer utilizing
the current integrator values for the forward, aft
and lateral swash plate servos to generate pitch, roll
and collective reference values for use by the computer
in generating pitch, roll and collective trim commands
for application on the connections 42-44 (Fig. 1) to
the respective servo trim actuators 39-41 (Fig. 1).
And then the program will revert to the third autopilot
routine through a return point 160. In subsequent
passes through the blade fold position calculation
routine 63a, eventually test 158 will indicate that
more than ten seconds has elapsed since the first trim
reference values were generated in step 159. This
25 means that the servos 31-33 (Fig. 1) have had ten
seconds at their 10% per second rate to achieve up to
100~ of their permissible motions, whereby any trim
reference value calculated in steps 159 should have
been achieved at this point. When the ten second
timer reaches maximum, a negative result from test
158 will lead to steps 162 which compare the desired
swash plate servo positions with the actual swash plate
servo positions indicated by the position detectors
20-22 (Fig. 1). If the swash plate servos have been
positioned very close to the desired amount, a series
11~141
-- 31 --
of tests 163-165 may all be affirmative. But if any of
the servo position errors is in excess of 0.2% of the
total servo positioning range, then a negati~e result
of any one of the tests 163-165 will lead to steps 166
in which the integrators for the forward, aft and
lateral swash plate servos are incremented by an
integration constant times the corresponding error.
This integration is one of the features of the in-
vention and avoids any long term errors, including
overshoots, which occur as a result of mechanical
errors in trim actuation and linkages that determine
the swash plate servo positions, and variations due
to non-zero series actuator positioning and yaw sensor
errors. In the normal course of events, successive
passes through the blade fold position calculation
routine 63a (reached through the third autopilot
program in successive computer real time interrupts,
as described hereinbefore) will cause the swash plate
position servo integrators to be integrated to values
which, after inverse matrix calculation (159), provided
correct pitch, roll and collective references which
drive the system to the desired swash plate servo
positions. During the first ten seconds when the steps
and tests 162-166 are bypassed, and subsequently when
the steps and tests 162-166 are included, the center
pedals routine 144 (described hereinafter with respect
to Fig. 7) is also performed. Thus, eventually, the
pedals should be centered as described hereinafter and
eventually the errors should all be less than 0.2%.
Thus an affirmative result from all three tests 163-165
will lead to a test 168 which determines if the pedals
are in fact centered. If not, the program simply exits
through the return point 160. But if the pedals are
centered, an affirmative result of test 168 will lead
to a step 169 which sets the swash in tolerance flag
- 32 -
which is utilized in the blade fold background program
of Fig. 4 to recognize the end of the blade fold mode
of operation ln the computer. And a step 170 commands
the blade lock to engage and indicates to the pilot
that the blade lock is enabled.
In the embodiment of the invention described with
respect to Figs. 1-5, hereinbefore, a simple method is
shown for centering a series actuator in an open loop
fashion, to eliminate discrepancies in swash plate
servo position as a consequence of variations in series
actuator position. There are other design criteria
of helicopters which may provide problems that need
to be overcome in order to provide the improved blade
pitch angle positioning according to the invention.
For instance, in the case of an aircraft having a
canted tail rotor (such as is disclosed in U.S. Patent
No. 4,103,848) any variation in tail rotor pitch and/or
speed will affect the pitch axis of the helicopter.
Therefore, there may be coupling between tail rotor
blade pitch angle commands and longitudinal cyclic
pitch col~mands in order to compensate for the effect
of the tail rotor on the pitch axis. In a typical case,
the coupling is such that there is essentially no
effect on the pitch axis when the tail rotor pedals
are centered, but increments of positive and negative
variations can be applied when the pedals are pushed
to the right or to the left respectively. Therefore,
the centered pedal position can be taken as a neutral
position insofar as eliminating tail rotor coupling
into the pitch axis is concerned to enable utilizing
the trim system to position the pitch angle of the
main rotor blades in accordance with the invention.
In any case in which there is a position detector
directly coupled to the yaw trim servo, the yaw trim
servo could be positioned to a center position in the
1~;141~
- 33 -
simple fashlon described with respect to the series
actuator in Fig. 5 hereinbefore. sut in a case where
the position detector utilized to close the servo
loop is engagable at different positions to represent
a selective synchronized trim position, the position
detector only indicates relative position with respect
to whatever position of the trim servo it was engaged
at. Therefore, the relative position detector does
not provide an indication of where the yaw trim servo
is located. In addition, all helicopters are provided
with coupling between the collective pitch and the
tail rotor pitch. Yet another problem which may occur
is the inability of a trim system to provide a command
e~ual to 100~ of authority (from maximum negative to
maximum positive, equal to driving from full left
pedal to full right pedal), Therefore, utilization of
the present invention requires a certain degree of
accommodation to one or more of the foregoing problems.
In accordance with another aspect of the invention,
a center pedals subroutine 144 which accommodates the
foregoing problems is reached in Fig. 10 through an
entry point 173. In this subroutine, the yaw trim
piston is first driven all the way to the left stop
(if it can be) and then it is driven back through 50%
of authority as indicated by the relative position
detector. This overcomes the problem of not knowing
the actual trim piston position represented by the
relative position detector. But, if the pedals are
initially set at or near a full right position,
limiting of the trim command will preclude providing
a large enough command signal to drive the pedals
fully to the left. The subroutine 144 therefore
senses a case where the relative position detector
indicates a relative position in excess o~ 90% of
full authority and will reposition the piston by 40%
4:~
~ 34 -
of the full authority as indicated by the relative
position detector. This overcomes the problem of
having less than full authority in the yaw trim
channel. Additionally, the subroutine of Fig. 10
takes into account that for collective stick positions
equal to 50~ or greater, there is coupling between the
collective stick and the yaw pedal stops, which can
affect the stop-referenced, relative pedal positioning
to be achieved in the center pedals routine. When it
can be determined that the yaw trim piston has been
driven as far to the left as it can, then either a 40
or a 50% of authority correction is utilized as a
command synchronizing position to drive the pedals to
or near center position.
A first test in the subroutine 144 of Fig. 10 is
a test 174 to determine if any pedal command has been
generated. During the initial phases of the subroutine,
there is no pedal command since it is reset to zero in
the initialization step 143 (Fig. 8). Therefore, an
affirmative result of test 174 leads to the portion of
the program which attempts to drive the yaw trim piston
fully to theleft so that it can be backed off therefrom
to the center, utilizing a relative position detector.
A step 175 resets the pedals centered flag that is
eventually interrogated in test 168, Fig. 9; when this
subroutine has completed its task, the pedals centered
flag will be set so that the blade lock enable signal
can be generated. Then a test 176 determines if pedal
force is greater than 2 lbs. (about 1 kilogram); when
the pedal force does reach about 2 lbs., this is an
indication that the trim detent spring is being ex-
tended as a consequence of the yaw trim piston being
pushed to theextreme left. Thus it is therefore known
that the pedals have reached a stop. Until this occurs,
a negative result of test 176 will reach a test 177 to
1~14 L~
- 35 -
determine if the relative yaw position exceeds 90~ of
total authority. If it does, this is an indication
that its zero position must have been set at or near
~ull right pedal, and that the trim system will not
have a sufficient authority to drive the yaw trim
piston into the left pedal stop and therefore no 2 lb.
indication will be available. If test 177 is negative,
a yaw trim synch command is generated equal to an
original yaw trim synch command plus an increment in
a step 17~. The increment is such as to command the
pedals to move to the left. This increment can be
chosen so as to cause the pedals to advance slowly to
the left stop in a desired fashion. And then, the
subroutine 144 will revert to the blade fold position
calculation routine 63a of Fig. 9 by means of a return
point 179. If, on the other hand, before 2 lbs. is
sensed in the test 176, the relative yaw position
exceeds 90% of full authority in the left hand direc-
tion, test 177 will be affirmative so that a step 180
will generate a pedal command equal to the relative
yaw position minus 40% of authority, which will bring
the yaw trim piston back near the center position. To
the extent that the yaw trim piston is not centered,
the pitch, roll and collective reference signals
generated in the step 159 (Fig. 9) will be incorrect,
so that the swash plate servo errors generated in step
162 will be significant. However, these errors will
be integrated out in the steps 166 so that pitch, roll
and collective reference signals will eventually be
generated to drive the swash plate servos to the
desired positions. Thus, some error in pedal centering
is tolerable in accordance with the present invention
due to the integration of swash plate servo error in
generating the trim commands.
- 36 -
If on the other hand, the pedal force reaches 2
lbs. the test 176 will be affirmative so that a test
181 will be reached to determine if the collective
stick position has been driven to a point below 45%
of authority. If it has not, no functions are per-
formed in the subroutine 144 but the program reverts
through the return point 179 to the blade fold position
calculation routine of Fig. 9. By choosing the blade
pitch angle for blade locking of the blades to cor-
respond to low collective pitch command, the trimsystem, by means of the collective reference signal
generated in step 159, will eventually drive the col-
lective stick position below 45% authority. When the
collective pitch has been driven to a point below 45%
of authority, an affirmative result of the test 181
will reach a step 182 which causes a pedal command to
be generated equal to the relative yaw position minus
50~ of yaw authority. Thus, the first portion of the
subroutine 144 in Fig. lO simply tries to place the
yaw piston in a known position and then uses the
relative yaw position to generate a command to drive
the yaw trim piston back to center (or near it).
When either of the steps 180, 182 has established
a pedal command, its non-zero condition will cause a
negative result of test 174 which will reach a step
183 in which the yaw trim synch is now set equal to
the recently generated pedal command. This pedal
command is, in either instance, a relative pedal com-
mand since it is based upon the relative yaw position
detector. The yaw trim piston is therefore driven
closed loop, the loop being closed by the relative
position detector. This occurs in every pass through
the subroutine by means of a test 184 which determines
if the relative yaw position now being indicated by
the relative yaw position detector is within some
- 37 -
tolerance of the yaw trim synch being repetitively
commanded in each cycle. When the relative yaw
position equals the command given, a step 185 sets
the pedals centered flag which is utilized in the
blade fold position calculation routine of Fig. 9.
Depending upon the starting conditions and response
times of the various portions of the system, it is
possible to have the swash plate almost correctly
positioned prior to centering of the pedals; in that
case,thefinal positioning achieved by the blade fold
position calculation routine of Fig. 9 may only be
taking out the errors which result from the pedals
not being centered until the pedals become centered.
In other cases, the pedals may become centered long
before the swash plate positions are near their cor-
rect setting In either event, it takes both the
positioning of the swash plate servos and the centering
of the pedals in order to effectively recognize that
the blades have been correctly positioned by setting
the swash in tolerance flag, in step 169 of Fig. 9.
Although the invention has been shown and de-
scribed 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.
