U.S. patent application number 10/260470 was filed with the patent office on 2003-05-15 for flight lock actuator with dual energy sources.
Invention is credited to Barba, Valentin G..
Application Number | 20030089826 10/260470 |
Document ID | / |
Family ID | 26948016 |
Filed Date | 2003-05-15 |
United States Patent
Application |
20030089826 |
Kind Code |
A1 |
Barba, Valentin G. |
May 15, 2003 |
Flight lock actuator with dual energy sources
Abstract
A flight lock actuator that can be powered by two sources of
stored energy when aircraft power has been switched off. A
mechanical energy storage means and an electrical energy storage
means provide a fully redundant energy storage system that stores
sufficient energy to complete the actuator's extension stroke when
aircraft power is removed. The actuator has a motor control system
that limits the stroke velocity for both the extension and
retraction strokes, including a damper feature capable of
effectively braking the actuator during the back-driven extension
stroke.
Inventors: |
Barba, Valentin G.;
(Scotrun, PA) |
Correspondence
Address: |
FISH & NEAVE
1251 AVENUE OF THE AMERICAS
50TH FLOOR
NEW YORK
NY
10020-1105
US
|
Family ID: |
26948016 |
Appl. No.: |
10/260470 |
Filed: |
September 26, 2002 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60348881 |
Nov 13, 2001 |
|
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Current U.S.
Class: |
244/129.1 |
Current CPC
Class: |
E05B 81/25 20130101;
B64C 1/1407 20130101; H02K 7/06 20130101; E05B 2047/0023
20130101 |
Class at
Publication: |
244/129.1 |
International
Class: |
B64C 001/00 |
Claims
What is claimed is:
1. A method for providing improved reliability in an aircraft door
flight lock actuator comprising: storing energy in a mechanical
energy storage means and an electrical energy storage means;
powering the actuator using the energy stored in the mechanical
energy storage means and the electrical energy storage means to
complete an unlocking stroke in the absence of aircraft power; and
controlling a linear velocity of the actuator.
2. The method defined in claim 1, wherein storing energy in the
mechanical energy storage means comprises deforming a compression
coil spring during a powered locking stroke of the actuator.
3. The method defined in claim 1, wherein storing energy in the
electrical energy storage means comprises charging at least one
capacitor during a powered locking stroke of the actuator, and
during a subsequent powered stall of the actuator.
4. The method defined in claim 1, wherein storing energy in the
electrical energy storage means comprises charging a rechargeable
battery during a powered locking stroke of the actuator, and during
a subsequent powered stall of the actuator.
5. The method defined in claim 1, wherein the mechanical energy
storage means and the electrical energy storage means are fully
redundant.
6. The method defined in claim 1, wherein controlling the linear
velocity of the actuator comprises: sensing a rotational speed of
an actuator motor; sensing a first current supplied to the motor;
reducing the first current if the rotational speed is higher than a
maximum speed, or if the first current is higher than a maximum
current.
7. The method defined in claim 6, wherein controlling the linear
velocity of the actuator further comprises: shunting a second
current generated by the motor into a damper circuit to place an
electrical load on the motor if the first current is substantially
zero and the rotational speed is higher than the maximum speed.
8. The method defined in claim 6, wherein sensing the rotational
speed of the motor comprises measuring a frequency of a Hall effect
sensor signal.
9. The method defined in claim 6, wherein sensing the rotational
speed of the motor comprises measuring a back electromotive force
generated by the motor.
10. The method defined in claim 6, wherein reducing the first
current comprises reducing a voltage supplied to the motor.
11. The method defined in claim 6, wherein reducing the first
current comprises pulse-width-modulating a power signal supplied to
the motor.
12. A system for providing improved reliability in an aircraft door
flight lock actuator comprising apparatus for: storing energy in a
mechanical energy storage means and an electrical energy storage
means; powering the actuator using the energy stored in the
mechanical energy storage means and the electrical energy storage
means to complete an unlocking stroke in the absence of aircraft
power; and controlling a linear velocity of the actuator.
13. The system defined in claim 12, wherein the apparatus for
storing energy in the mechanical energy storage means comprises
apparatus for deforming a compression coil spring during a powered
locking stroke of the actuator.
14. The system defined in claim 12, wherein the apparatus for
storing energy in the electrical energy storage means comprises
apparatus for charging at least one capacitor during a powered
locking stroke of the actuator, and during a subsequent powered
stall of the actuator.
15. The system defined in claim 12, wherein the apparatus for
storing energy in the electrical energy storage means comprises
apparatus for charging a rechargeable battery during a powered
locking stroke of the actuator, and during a subsequent powered
stall of the actuator.
16. The system defined in claim 12, wherein the apparatus for
storing energy in the mechanical energy storage means and the
electrical energy storage means are fully redundant.
17. The system defined in claim 12, wherein the apparatus for
controlling the linear velocity of the actuator comprises apparatus
for: sensing a rotational speed of an actuator motor; sensing a
first current supplied to the motor; reducing the first current if
the rotational speed is higher than a maximum speed, or if the
first current is higher than a maximum current.
18. The system defined in claim 17, wherein the apparatus for
controlling the linear velocity of the actuator further comprises
apparatus for: shunting a second current generated by the motor
into a damper circuit to place an electrical load on the motor if
the first current is substantially zero and the rotational speed is
higher than the maximum speed.
19. The system defined in claim 17, wherein the apparatus for
sensing the rotational speed of the motor comprises apparatus for
measuring a frequency of a Hall effect sensor signal.
20. The system defined in claim 17, wherein the apparatus for
sensing the rotational speed of the motor comprises apparatus for
measuring a back electro-motive force generated by the motor.
21. The system defined in claim 17, wherein the apparatus for
reducing the first current comprises apparatus for reducing a
voltage supplied to the motor.
22. The system defined in claim 17, wherein the apparatus for
reducing the first current comprises apparatus for
pulse-width-modulating a power signal supplied to the motor.
23. A linear actuator comprising: a brushless electric motor for
rotating a shaft in either rotational direction; a ball screw
assembly for converting rotation of the shaft to linear motion of a
follower member and vice versa, the follower member moving in
either linear direction with the respective rotational direction of
the shaft; control circuitry for selectively powering the motor to
rotate in either rotational direction; and a mechanical energy
storage assembly for resiliently urging the follower member to move
in a predetermined one of its linear directions, the motor being
powerful enough to overcome the resilient urging of the storage
assembly when the control circuitry powers the motor to produce
motions of the follower member opposite the predetermined one
linear direction.
24. The linear actuator defined in claim 23 further comprising:
electrical energy storage circuitry for storing electrical energy
during powering of the motor to produce motion of the follower
member opposite the predetermined one linear direction, the control
circuitry being adapted to selectively use electrical energy from
the storage circuitry to power the motor to produce motion of the
follower member in the predetermined one linear direction.
25. The linear actuator defined in claim 23 further comprising: a
first shock absorbing stop for stopping motion of the follower
member in the predetermined one linear direction adjacent a first
location.
26. The linear actuator defined in claim 25 further comprising: a
second shock absorbing stop for stopping motion of the follower
member opposite the predetermined on linear direction adjacent a
second location.
27. The linear actuator defined in claim 23 wherein the control
circuitry comprises: circuit components for limiting the speed of
the motor.
28. The linear actuator defined in claim 27 wherein the circuit
components comprise: circuitry for selectively applying electrical
current generated by the motor to an electrical load to thereby
retard the motor.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application claims the benefit of U.S. Provisional
application No. 60/348,881, filed Nov. 13, 2001.
BACKGROUND OF THE INVENTION
[0002] This invention relates to actuators used in aircraft door
lock mechanisms. More specifically, this invention relates to
systems and methods for improving the reliability of aircraft door
flight lock actuators. Flight lock actuators are used in aircraft
door lock mechanisms to secure a lock mechanism in the locked
position during flight, or whenever aircraft power is supplied to
the actuator. During flight, for safety reasons, it is conventional
practice to maintain the flight lock actuator in a powered stall
against its locked position stop. When aircraft power is removed at
the end of the flight, the flight lock actuator is conventionally
returned to its unlocked position by a spring system. For safety
reasons, limit switches and brakes are not permitted in flight lock
actuators.
[0003] Actuators built according to the present art suffer
occasional failure due to damage from abruptly impacting mechanical
stops at the end of an actuator's stroke, especially while being
back-driven by a spring system during the unpowered extension
stroke to the unlocked position. These repeated mechanical shocks
to the internal mechanism of a flight lock actuator can cause
jamming and mechanical failure of the actuator. In addition,
actuator failure may also result from damage to a brush-type
actuator motor due to prolonged periods of powered stall in the
locked position. The aircraft flight lock actuator is a key safety
element in an aircraft. Any failure in the door lock mechanism,
including the flight lock actuator, should be avoided.
[0004] Therefore, it would be desirable to provide a redundant
stored energy system to power the flight lock actuator's extension
stroke to the unlocked position stop. It would be further desirable
to extend and retract the flight lock actuator to its mechanical
stops in a controlled manner, so as to eliminate failure due to
damage from abrupt impacts. It would be further desirable to power
the flight lock actuator using a motor less prone to suffer damage
from prolonged periods of powered stall.
[0005] In view of the foregoing, it is an object of this invention
to provide a flight lock actuator using systems and methods that
significantly improve its reliability.
SUMMARY OF THE INVENTION
[0006] These and other objects are accomplished in accordance with
the principles of the present invention by providing an aircraft
flight lock actuator having a redundant energy storage system, a
motor control system and a brushless motor drive.
[0007] The redundant energy storage system utilizes a mechanical
energy storage system and an electrical energy storage system to
store sufficient energy for extending the actuator to its unlocked
position after the removal of aircraft power. The electrical energy
storage system stores electrical energy during the actuator's
powered retraction to the locked position and during the period of
powered stall in the locked position. The mechanical energy storage
system also stores energy during the powered retraction stroke. The
dual energy storage systems are fully redundant, which provides
that energy stored in either system alone would be sufficient to
drive the actuator to its unlocked position if the other energy
storage system fails.
[0008] A motor control system is provided that senses the
rotational speed of the flight lock actuator motor, and which
limits both the retract and extend strokes to a desired maximum
velocity to reduce mechanical shock. The motor control system also
limits, to a desired maximum, the current supplied to the actuator
motor during a period of powered stall to prevent motor damage from
overheating. The motor control system also includes a damper
feature that is capable of effectively braking the actuator during
a back-driven extension stroke to ensure a controlled arrival into
the extended position mechanical stop.
[0009] Further features of the present invention, its nature, and
various advantages will be more apparent from the accompanying
drawings and the following detailed description of the preferred
embodiments.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] FIG. 1 is a simplified elevational view, partly in section,
of an illustrative aircraft flight lock actuator mechanical system
in accordance with the present invention.
[0011] FIG. 1A is an enlargement of a portion of FIG. 1.
[0012] FIG. 2 is a simplified schematic block diagram of an
illustrative aircraft flight lock actuator electrical system in
accordance with the present invention.
DETAILED DESCRIPTION
[0013] As shown in FIGS. 1 and 1A, an illustrative embodiment of
the flight lock actuator mechanical system includes rear housing 1,
center housing 2, guide tube 3, guide bushing 4, and front cover 5.
Disposed within rear housing 1 and center housing 2 is a brushless
DC electric motor 20, which includes stator 25 and rotor 26. Rotor
26 includes rotor shaft 28, which rotates with rotor 26. Ball screw
shaft 7 is press fit into rotor shaft 28, so that ball screw shaft
7 also rotates with rotor 26. The right-hand portion of ball screw
shaft 7 is threaded to provide the central element of a ball screw
assembly. Ball screw shaft 7 is axially and rotationally fixed to
sleeve 14. Pin 15 is captured within the inner race of bearing 19,
and extends diametrically through shaft 7 and sleeve 14 to ensure
the fixed relationship between elements 7 and 14.
[0014] The rotational assembly including rotor 26, shaft 7, and
sleeve 14 is rotatably supported by bearings 19. Sleeves 11 and 14
are slidable inside the inner races of bearings 19 parallel to the
longitudinal axis of shaft 7. The amount by which sleeves 11 and 14
can slide in this manner is limited by the compressibility of two
sets, 12 and 13, of Belleville washers. Each set of Belleville
washers is disposed between a set of two flat washers. The first
set of Belleville washers 13 is captured between a radially
outwardly extending flange on sleeve 11 and the inner race of the
adjacent bearing 19, and the second set 12 is captured between a
radially outwardly extending flange on sleeve 14 and the inner race
of the adjacent bearing 19.
[0015] Belleville washer sets 12 and 13 nominally axially center
sleeves 11 and 14 between bearings 19. However, by resiliently
deforming, the sets of Belleville washers allow shaft 7 to
temporarily shift to the left or right when the axially translating
sleeve of ball screw 7 hits its outbound or inbound stop,
respectively. The Belleville washers thereby act as resilient shock
absorbers for the rotational assembly.
[0016] The output assembly includes ball nut 27, ball nut coupling
9, and output ram 6. Output ram 6 is attached to ball nut coupling
9, which is in turn attached to ball nut 27. The entire output
assembly is able to translate axially inside guide tube 3. As motor
20 drives ball screw 7 to rotate, ball nut 27 is prevented from
rotating by a key on collar 8 (attached to ball nut elements 9/27),
which key is slidably engaged in an axial slot in guide tube 3. As
ball screw 7 rotates, ball nut 27 is driven (via balls (not shown)
between elements 7 and 27) to translate axially inside guide tube
3, causing the output assembly (especially output ram 6) to drive
an external load through bearing 16. Conversely, linear translation
of the output assembly including ball nut 27 causes ball screw 7
and motor 20 to rotate.
[0017] Helical compression spring 24 may be provided as means for
mechanical energy storage. Disposed within guide tube 3,
compression spring 24 is trapped between washer 10, set against
center housing 2, and collar 8 that moves axially with ball nut 27.
In FIGS. 1 and 1A, ball nut 27 is shown in its fully retracted
position. This is the condition of the actuator mechanical system
in which spring 24 is in its most compressed state. Spring 24 urges
output ram 6 to extend from the fully retracted position
illustrated in FIGS. 1 and 1A to a fully extended position in which
collar 8 contacts stationary bushing 30 in the right-hand end (as
viewed in FIGS. 1 and 1A) of guide tube 3. Spring 24 applies a
spring force to output ram 6 in the direction of extension
regardless of the output ram's position in guide tube 3. This
includes a residual spring force applied by spring 24 in the
extension direction when output ram 6 is in its fully extended
position.
[0018] When motor 20 is driven in the appropriate direction by
electrical power from the aircraft that includes the flight lock
actuator, motor 20 is able to overcome the force of spring 24 and
retract ball nut 27 all the way to contact another stationary stop
at the left-hand end (as viewed in FIGS. 1 and 1A) of guide tube 3.
Moreover, as long as power is thus applied to motor 20, it is able
to hold ball nut 27 in the fully retracted position with spring 24
substantially compressed. When power from the aircraft is removed
from the flight lock actuator, spring 24 is able to drive, even
without reverse driving of motor 20 as described below, assembly
6/9/27 back to its other stop at the other (right-hand) end of
guide tube 3, thereby fully extending output ram 6. Alternate means
may also be used for mechanical energy storage within the scope of
the present invention.
[0019] Output ram 6/9/27 is sealed with respect to guide tube 3
using seal 18 disposed in guide bushing 30. Wiper seal 17 may also
be disposed in front cover 5. Guide bushing 4/30 remains stationary
and functions as the internal extension stroke mechanical stop.
Washer 10 functions as the internal retraction stroke mechanical
stop. When ball nut 27 hits washer 10 at the end of the retraction
stroke, motor 20 is able to continue to rotate briefly as
Belleville washers 12 compress. Thereby reducing the impact of ball
nut 27 on washer 10, and allowing motor 20 to stop somewhat
gradually, rather than instantaneously. Belleville washers 12
therefore cushion the end of the retraction stroke, thereby greatly
reducing the risk of damage to any part of the apparatus at the end
of retraction strokes. Belleville washers 13 function similarly to
cushion the end of extension strokes.
[0020] A block schematic diagram of an illustrative embodiment of
the flight lock actuator electrical system is shown in FIG. 2.
Aircraft DC power (e.g., 28 volts DC) is received via terminal 112
when switch 110 is closed. Switch 110 is typically closed
automatically when an aircraft enters a predetermined condition
(e.g., a forward ground speed of a certain number of miles per
hour, etc.). Switch 110 automatically re-opens when the aircraft is
no longer in a condition that causes the switch to close.
Electro-magnetic interference filter 111 is connected to terminal
112 to protect the flight lock actuator's electrical system from
conducted and radiated interference from the aircraft's electrical
system, and vice versa. Other terminals 114 and 116 (e.g., 28V
RETURN and CASE ground) may also exist in the system.
[0021] From terminal 112, aircraft power is conducted to motor
controller 130 through diode D1. The presence or absence of
aircraft power at terminal 112 is sensed by motor controller 130
via the RET/EXT (RETRACT/EXTEND) input to the motor controller. If
the signal on lead RET/EXT is "high" (e.g., 18 to 29 VDC), motor
controller 130 drives motor 20 to retract the actuator to its
retracted position (in which the actuator locks a door-opening
mechanism of the aircraft). Motor 20 will stall the actuator in the
retracted position as long as aircraft power continues to be
supplied.
[0022] As means for electrical energy storage, capacitor C may be
provided in electrical energy accumulator 140. When aircraft DC
power is present at terminal 112, capacitor C charges through
resistor Rc and diode D2. Zener diode Z sets the capacitor charge
voltage upper limit at a value appropriate to proper circuit
operation. When capacitor C has charged to the Zener diode
breakdown value, all further charging current, limited by charging
resistor Rc, bypasses capacitor C and flows to RETURN through Zener
diode Z. Other suitable means for electrical energy storage may
also be used within the scope of the present invention. For
example, a rechargeable battery may be alternately provided in
electrical energy accumulator 140 and charged using aircraft power
from terminal 112.
[0023] Sensing line RET/EXT signals the removal of aircraft power
by going "low" (e.g., .about.0V), which signals motor controller
130 to cause any subsequently applied electrical power to rotate
motor 20 in the direction required to extend the actuator to its
unlocked position. After removal of aircraft DC power, power for
motor 20 and for motor controller 130 during the extension stroke
is provided by capacitor C through diode D3. The capacitor voltage
will decay as current is drawn from electrical energy accumulator
140. However, the flight lock actuator electrical components are
preferably sized so that the capacitor does not discharge below a
motor controller 130 operational voltage value before the extension
stroke of the actuator has been completed. Voltage regulator 150
supplies a constant control voltage for powering the control
circuitry of motor controller 130 (as distinct from powering motor
20).
[0024] During either the retraction or extension stroke, motor
controller 130 preferably limits stroke velocity by limiting
current through motor 20. For example, such current controller
limiting may be achieved by pulse-width-modulation of the motor
power signal, or by reducing the voltage available to motor 20.
Current through motor 20 is measured by a voltage drop across
resistor R1. In order to limit impact velocity at the mechanical
stops, the rotational speed of motor 20 is preferably limited,
while still meeting a maximum allowable stroke time with
appropriate margin. Motor 20 is typically equipped with Hall effect
sensors 31 that signal the rotor's angular position to motor
controller 130. Rotational speed of motor 20 may be ascertained
using the frequency of the Hall effect sensors' signal.
Alternately, because a permanent magnet motor generates a back-EMF
proportional to its rotational speed, the generated back-EMF may be
used by motor controller 130 to ascertain the rotational speed of
motor 20.
[0025] If motor controller 130 senses a motor rotational speed in
excess of a desired maximum speed, it may reduce the voltage
available to motor 20, or may pulse-width-modulate the motor power
signal, so as to drop the motor rotational speed to the desired
range. If current limit circuitry 160 senses that the current
passing through motor 20 is higher than a predetermined value that
indicates motor 20 is in a state of powered stall, current limit
circuitry 160 may signal motor controller 130 to reduce the voltage
available to motor 20, or appropriately pulse-width-modulate the
motor power signal, so that the current supplied during a period of
powered stall does not overheat motor 20.
[0026] The flight lock actuator's extension stroke is typically
subject to a substantial aiding force from spring 24 internal to
the actuator and possibly also from springs external to the
actuator in the aircraft door lock mechanism. In order to limit
excessive extension stroke velocity caused by this aiding force,
the flight lock actuator electrical system, in addition to limiting
motor rotational speed by limiting the current to motor 20, may
seek to effectively brake motor 20 using damper circuit 170. During
the actuator's extension stroke, if current to motor 20 has been
reduced to substantially zero, while motor rotational speed remains
above a desired maximum speed, the flight actuator electrical
system may shunt back-EMF (electro-motive force) generated by the
motor into damper circuit 170 to place an electrical load on motor
20. By temporarily transforming motor 20 into such a loaded
electrical generator, a braking effect is achieved on motor 20.
Logic circuitry 180, having inputs from current limit circuit 160
and motor controller 130, monitors the predetermined condition for
shunting current generated by motor 20 to damper circuit 170.
[0027] To elaborate the last points, a permanent magnet motor
generates a back-EMF proportional to its rotational speed. This
generated back-EMF, or the signal frequency of Hall effect sensors
31, may also be used by motor controller 130 to ascertain the motor
rotational speed. Logic circuitry 180 may monitor motor current
information from current limit circuit 160 and motor rotational
speed information from motor controller 130. During the extension
stroke, a motor current that is substantially zero, in combination
with motor rotational speed in excess of a desired maximum speed,
are signals indicating that the aiding force has driven motor 20 to
an excessive speed despite motor controller 130 reducing the motor
voltage or pulse width to near zero.
[0028] When logic circuitry 180 senses this condition, it reacts by
throwing the motor controller 130 bridge into a full wave rectifier
mode (i.e., all MOSFETs "off") to shunt all current generated by
motor 20 through the MOSFETs' internal bypass diodes to a load
resistor in damper circuit 170. By operating motor 20 as a
generator in conjunction with the load resistor in damper circuit
170, the motor speed is reduced to a desired speed that ensures a
controlled arrival into the mechanical stop.
[0029] The aircraft door flight lock actuator of the present
invention utilizes a redundant energy storage system having
mechanical and electrical energy storage means to store energy
during the actuator's powered retraction stroke to the locked
position, and subsequent period of powered stall in the locked
position. Once aircraft power is removed from the flight lock
actuator, the stored energy is used to power an extension stroke to
the unlocked position. A brushless motor is used to power the
actuator, and the actuator electrical system ensures that the motor
does not overheat by limiting the current supplied to the motor to
a desired maximum. The actuator electrical system ensures the
controlled arrival of the actuator into its mechanical stops by
limiting motor rotational speed using a current limiting method and
a damper feature that effectively brakes the actuator motor.
Utilizing these systems and methods, a flight lock actuator with
substantially improved reliability is provided.
[0030] It will be understood that the foregoing is only
illustrative of the principles of the invention, and that various
modifications can be made by those skilled in the art without
departing from the scope and spirit of the invention. For example,
although the foregoing describes an illustrative aircraft door
flight lock actuator that retracts to lock and extends to release,
it should be obvious to those skilled in the art that the present
invention is equally adaptable to an actuator that extends to lock
and retracts to release.
* * * * *