Automatic lock-positioning of foldable helicopter blades

The foldable rotor blades (12) of a helicopter are automatically adjusted to pitch angles where they can be locked as a prerequisite to folding by commands generated (98, 112) to cause trim actuators (39-41) to drive swash plate servos (17-19) to the correct positions, initially (98) in response to stored trim references (120) and eventually (112) in response to the difference (109) between stored swash plate servo positions (120) and current servo positions (20-22).

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 automatically 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 applications, 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 being folded. 
As is known, the pitch 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. 
It 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 has 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. Electric 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 blades while they are held in the proper 
position. This would have required the use of pilot indicators to show the 
pilot correct blade pitch positions for pin engagement, the pilot moving 
the controls very slowly to achieve indications and to provide minute 
adjustment 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, copending U.S. patent application of Ferris 
entitled BLADE FOLD RESTRAINT SYSTEM, Ser. No. 35,364, filed on May 2, 
1979 and now U.S. Pat. No. 4,284,387. 
DISCLOSURE OF INVENTION 
Objects of the invention include automatic positioning 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 the present 
invention, the pitch, roll and collective trim commands are provided in 
response to stored pitch, roll and collective trim reference 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 invention, 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 previous 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. 
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. Similarly, the 
foregoing and various other objects, features and advantages of the 
present invention will become more apparent in the light of the following 
detailed description of an exemplary embodiment thereof, as illustrated in 
the accompanying drawing.

BEST MODE FOR CARRYING OUT THE INVENTION 
Referring now to FIG. 1, a conventional helicopter with suitable apparatus 
to permit practice of the present 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 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 suitable 
electrically-controlled trim actuator 39-41 (either electric motor or 
hydraulic) which is responsive to a corresponding electric connection 
42-44 from an automatic flight control computer 45. Provision of suitable 
trim command signals on the connections 42-44 by the computer 45 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 flight. The automatic flight control computer may take 
the form of one of the computers described in a commonly owned, copending 
U.S. patent application of Murphy and Clelford, Ser. No. 938,583, filed on 
Aug. 31, 1978, now U.S. Pat. No. 4,270,168, and entitled 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 automatic flight control computer 
45, and the capability of receiving inputs from inertial devices such as 
accelerometers 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 automatic flight control computer 45 by a plurality of 
corresponding connections 55. The automatic flight control computer 45 may 
interconnect with other apparatus of the aircraft through multiplexed 
inputs and outputs which include analog conversion where necessary, all as 
is well known in the art and as described in the forementioned 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 control 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 hereinafter, if the tail rotor is tilted, as is shown in 
U.S. Pat. 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, 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 described using as an example 
one of the automatic flight control 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 (BG) 
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 performed when the aircraft is on the ground are reached 
in a third autopilot routine (AP 3) which is illustrated, as modified to 
practice blade folding, in FIG. 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 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 
provide 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 
connections 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 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. All of the foregoing is described 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 described 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 4, 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 mode, in 
which self health tests such as a check sum test routine and a scratch pad 
test are performed, as illustrated in FIG. 4 of the aforementioned Murphy 
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 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 operation. 
In FIG. 3, the blade 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 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 program 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 provides the 
initialization steps 67-68d 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 continue 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. 
In a subsequent pass through the blade fold executive 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 interrupts 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 background 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 may occur, or as a consequence of having 
satisfactorily completed the blade pitch positioning function of a blade 
folding operation (which is the only function that the automatic flight 
control computer 45, FIG. 1, performs 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 that the rotor is 
inadvertently moved from its folding index position, or if the pilot 
changes his mind and disengages 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, thereby 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 (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 background program 62 is test 86 
which determines if any of the tests which may be performed on the trim 
system 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 maintenance 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 positioning 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 successfully 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. 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 background 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 automatic 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 
automatically branches back to the blade fold background program 62 by 
means of interrupt release, in the normal fashion. As is apparent from a 
full understanding of all of the routines described hereinbefore and 
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 programming 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. 
Once the executive program has been set into the blade fold mode by step 79 
within the blade fold executive 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 calculation 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 reference, roll reference and 
collective reference values, to be utilized in generating trim commands on 
the connections 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 steps 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 30 (FIG. 1). 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 101 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 
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 characteristic 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. Therefore, 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 acutator, such as a pitch bias 
actuator, has reached a position equal to a stored center position. If 
not, a negative result of test 108 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 affirmative 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 individual 
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 constants K1-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 
adjustments to the forward, aft and lateral swash plate servo 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 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 interruptions. 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 
purposes. Eventually, the third autopilot routine of FIG. 1 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 (FIG. 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 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. Initially, 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 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 positions 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 
ultimately locked. 
Although the invention has been shown and described with respect to 
exemplary embodiments thereof, it should be understood by those killed 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.