Technical Field
The present invention relates generally to brake systems
for vehicles.
Background of the Invention
Various types of braking systems are known. For example,
hydraulic, pneumatic and electromechanical braking systems have been developed for
different applications. In the past, however, it has not been shown to employ reliably
particularly with respect to an electromechanical braking system in a vehicle such
as an aircraft.
An aircraft presents a unique set of operational and safety
issues. For example, uncommanded braking due to failure can be catastrophic to an
aircraft during takeoff. On the other hand, it is similarly necessary to have virtually
fail-proof braking available when needed (e.g., during landing).
If one or more engines fail on an aircraft, it is quite
possible that there will be a complete or partial loss of electrical power. In the
case of a braking system, issues arise as to how the brakes will be actuated in
an emergency landing.
In view of such shortcomings associated with conventional
braking systems, there is a strong need in the art for an braking system which may
be employed reliably even on a vehicle such as an aircraft.
DE-A1-4227157 describes an automatic brake control system
for an aircraft. The system includes redundant control computers, redundant position
sensors, and redundant power supplies.
Summary of the Invention
According to the present invention, a braking system utilizes
redundancy features to provide safe and reliable braking.
In a preferred embodiment of the invention, the braking
system is configured to operate on power provided by multiple power sources. Different
modes of braking are available based on whether a failure has occurred in one or
more power sources. Additionally, system redundancy allows for failure in one or
more primary components without total loss of braking capacity. Proportional braking
is provided even in an emergency braking mode.
In a preferred embodiment, the present invention provides
a manner for arranging the components and power connection points within a braking
system architecture in order to better maintain isolation of the power busses, and
thereby improve overall integrity of the system, while still meeting system redundancy,
performance and safety requirements as in the past. In addition, the invention provides
a manner for connecting and efficiently using available power in emergency braking
and parking modes.
In accordance with one particular aspect of the invention,
there is provided a braking system comprising: a brake including a plurality of
brake actuators for effecting a braking torque on a wheel of a vehicle; and a plurality
of brake controllers for providing drive control signals to the brake actuators
in response to an input brake command signal to effect the braking torque, the plurality
of brake controllers being configured to function redundantly so as to provide the
drive control signals to effect the braking torque even in the event one of the
plurality of brake controllers becomes inoperative, characterized in that the limit
for the maximum braking force applied by a remaining actuator of the brake is increased
to compensate for an actuator of the brake that has been disabled.
To the accomplishment of the foregoing and related ends,
the invention, then, comprises the features defined in the claims. The following
description and the annexed drawings set forth in detail certain illustrative embodiments
of the invention. These embodiments are indicative, however, of but a few of the
various ways in which the principles of the invention may be employed. Other objects,
advantages and novel features of the invention will become apparent from the following
detailed description of the invention when considered in conjunction with the drawings.
Brief Description of the Drawings
- Fig. 1 is a environmental view of an electromechanical braking system in an
aircraft in accordance with the present invention;
- Fig. 2 is a general block diagram of the electromechanical braking system in
accordance with the present invention;
- Fig. 3 is a detailed block diagram of the electromechanical braking system in
accordance with the present invention;
- Fig. 4A is a timing diagram illustrating operation of the electromechanical
braking system in a first alternate braking mode in which a primary AC power source
has failed;
- Fig. 4B is a timing diagram illustrating operation of the electromechanical
braking system in a second alternate braking mode in which an essential primary
AC power source has failed;
- Fig. 4C is a timing diagram illustrating operation of the electromechanical
braking system in an emergency braking mode in which all primary power sources have
failed;
- Fig. 4D is a timing diagram illustrating operation of the electromechanical
braking system in a park (ultimate) braking mode in which all primary power sources
are unavailable;
- Fig. 5A is a timing diagram illustrating operation of the electromechanical
braking system during failure of a brake system control unit;
- Fig. 5B is a timing diagram illustrating operation of the electromechanical
braking system during failure of an electromechanical actuator controller;
- Fig. 6 is a detailed block diagram of a particular embodiment of an electromechanical
braking system in accordance with the present invention;
- Fig. 7 is a detailed block diagram of a particular embodiment of a brake system
control unit in accordance with the present invention;
- Fig. 8 is a detailed block diagram of a particular embodiment of an electromechanical
actuator controller in accordance with the present invention;
- Fig. 9 is a detailed block diagram of an electromechanical braking system in
accordance with another embodiment of the present invention;
- Fig. 10 is a detailed block diagram of an electromechanical braking system in
accordance with a third embodiment of the present invention; and
- Fig. 11 is a detailed block diagram of an electromechanical braking system in
accordance with a fourth embodiment of the present invention.
Description of the Preferred Embodiments
The present invention will now be described with reference
to the drawings, wherein like reference labels are used to refer to like elements
throughout.
Referring initially to Fig. 1, an electromechanical braking
system 30 in accordance with the present invention is shown within a jet aircraft
32 (illustrated in phantom). As will be explained in more detail below, the system
30 is designed as a brake-by-wire system compatible with the performance, safety,
electrical and mechanical interfaces, redundancy, and other requirements of an aircraft
such as a commercial transport. The system 30 operates based on power provided from
a plurality of power sources. Power is segregated within the system 30 such that
the system 30 is capable of providing satisfactory braking even upon failure of
one or more power sources. Moreover, the system 30 has built in redundancy which
allows the system 30 to continue to operate satisfactorily even in the case of failure
of one or more system components.
In the exemplary embodiment, the system primary components
include four electromechanical brakes 34. The aircraft 32 in the present embodiment
includes a pair of wheels 36 mounted to a landing gear under the left wing of the
aircraft and a pair of wheels 36 mounted to a landing gear under the right wing
of the aircraft. Each wheel 36 includes a respective brake 34 for providing braking
action thereto.
The system 30 further includes two redundant digital brake
system control units (BSCUs) 40. As will be described in more detail below, the
BSCUs 40 carry out the brake control and antiskid processing functions. The BSCUs
40 are located in the electronics bay 42 of the aircraft 32, and preferably are
packaged into one enclosure with a firewall therebetween.
The system 30 also includes four redundant electromechanical
actuator controllers (EMACs) 44 which convert brake clamp force commands from the
BSCUs 40 to servo motor control signals which ultimately provide actuator braking
forces. The EMACs 44 preferably are packaged similar to the BSCUs 40, with two EMACs
44 per enclosure located near the top of the gear strut of each respective landing
gear.
A pilot of the aircraft 32 provides brake commands to the
braking system 30 via a pair of left and right brake pedal transducers 46 included
in the cockpit. The transducers 46 provide brake command signals to the BSCUs 40
which are proportional to the desired amount of braking. The output of each transducer
46 is coupled to the BSCUs 40 via a cable 48. Communications between the BSCUs 40
and the EMACs 44 occur over a communication bus 50 connected therebetween.
Each of the EMACs 44 is designed to provide electrical
power to the electromechanical actuators within the corresponding brakes 34 via
a respective power cable 52. In addition, each brake 34 has an associated torque
sensor and wheel speed sensor as described below. The outputs of the sensors are
provided to the respective EMACs 44 via cables 54. The EMACs 44 condition the signals
and provide them to the BSCUs 40 as feedback signals to carry out the brake control
and antiskid processing functions.
Fig. 2 is a simplified block diagram of the braking system
30 as employed within the aircraft 32. The BSCUs 40 and the EMACs 44 are shown collectively
as an electromechanical braking controller 60. The controller 60 receives as its
primary inputs the brake command signals from the transducers 46, and the outputs
of the torque and wheel speed sensors 62 included as part of the brake 34 on each
wheel 36.
The braking system 30 receives power from three primary
power busses and a secondary power buss included within the aircraft 32. As is known,
an aircraft 32 oftentimes will include multiple power busses. In the exemplary embodiment,
the aircraft 32 includes primary power busses PWR1, PWR2 and PWRess. Each power
buss preferably is independent of one or more of the other power busses to provide
a level of redundancy. For example, the power buss PWR1 consists of an alternating-current
(AC) power source AC1 and a commonly generated direct-current (DC) power source
DC1. Similarly, the power buss PWR2 consists of an AC power source AC2 and a commonly
generated DC power source DC2; and the power buss PWRess consists of an AC power
source ACess and commonly generated DC power source DCess.
The power buss PWR1 (i.e., AC1 and DC1 ) may be derived
from power generated by the left wing engine in the aircraft 32, for example. Similarly,
the power buss PWR2 (i.e., AC2 and DC2) may be derived from power generated by the
right wing engine 34. In this manner, if the left wing engine or the right wing
engine fails, power is still available to the system 30 via the power buss corresponding
to the other engine.
The power buss PWRess (i.e., ACess and DCess) may be derived
from power generated by the parallel combination of the left wing engine and the
right wing engine. In such manner, power from the power buss PWRess will still be
available even if one of the engines fail.
The aircraft 32 further includes an emergency DC power
buss represented by a DChot power source. The DChot power source is a battery supply
on board the aircraft 32. The battery may be charged via power from one of the other
power sources, or may be charged separately on the ground.
As will be appreciated, various circumstances can arise
where power from one or more of the power busses will become unavailable. For example,
the left wing engine or the right wing engine could fail causing the PWR1 (AC1/DC1)
and PWR2 (AC2/DC2) power sources to go down, respectively. Alternatively, power
generating equipment such as a generator, inverter, or other form of power converter
could fail on one of the respective power busses resulting in the AC1/DC1, AC2/DC2
and/or ACess/DCess power sources becoming unavailable. In addition, a failure can
occur in the cabling providing the power from the respective power sources to the
system 30, thus effectively causing the respective power source to no longer be
available. For this reason, the routing of the power cables for the different busses
preferably occurs along different routes throughout the plane to avoid catastrophic
failure on all the power buss cables at the same time.
Turning now to Fig. 3, the braking system 30 is illustrated
in more detail. As will be explained in more detail below, the system 30 utilizes
power buss partitioning in accordance with the present invention in order to reduce
and/or eliminate the risk of impairing or failing a power buss or supply as the
consequence of a system or component failure. In addition, the invention present
provides a method for connecting and efficiently using the available power in the
system 30 in parking and emergency modes.
The braking system 30 as shown in Fig. 3 has an exemplary
architecture for satisfying typical redundancy, performance and safety requirements
within an aircraft. Such architecture is presented by way of example to illustrate
the context in which the principles of the present invention may be employed. It
will be appreciated, however, that the present invention has utility with other
architectures and is not limited to the particular architecture shown. The manner
in which the present invention provides for power buss partitioning and efficient
braking in the parking and emergency modes can be applied to other architectures
as well.
As noted above, the system 30 includes two BSCUs 40 designated
BSCU1 and BSCU2, respectively. BSCU1 and BSCU2 are redundant and are both configured
to provide an input/output interface to the aircraft 32 electronics within the cockpit,
for example, via a bus 70. In addition, BSCU1 and BSCU2 each contain circuitry for
performing top level brake control and antiskid algorithm processing functions.
BSCU1 and BSCU2 each receive proportional brake command signals from the transducers
46 via cable 48.
BSCU 1 and BSCU2 are each designed to receive the proportional
brake command signals from the transducers 46 and process the signals based on the
aforementioned brake control and antiskid algorithms to produce a brake command
signal which is provided to the EMACs 44. The particular brake control and antiskid
algorithms employed by the BSCUs 40 can be conventional, and hence further detail
based thereon is largely omitted in the present description for sake of brevity.
BSCU1 and BSCU2 each provide brake commands and otherwise
communicate with the EMACs 44 via the aforementioned communication bus 50. As noted
above, the system 30 includes four redundant EMACs 44 respectively labeled EMAC
Left1, EMAC Left2, EMAC Right1 and EMAC Right2. As shown in Fig. 3, each EMAC 44
is coupled to the communication bus 50 so as to be able to receive brake commands
from each of the BSCUs 40 and otherwise communicate with the other devices coupled
to the bus 50. The EMACs 44 receive the left and right brake commands from the BSCUs
40 and provide control signals to actuator modules within the brakes 34 as discussed
below to drive the actuator modules to their commanded position or clamp force.
In this manner, controlled braking may be effected.
Each brake 34 included in the system 30 includes four separate
actuator modules (designated by numerals 1-4), although a different number may be
employed without departing from the scope of the invention. Each actuator module
1-4 includes an electric motor and actuator (not shown) which is driven in response
to electrical control signals provided by a respective EMAC 44 to exert mechanical
braking torque on a respective wheel 36. Each EMAC 44 controls half of the actuator
modules 1-4 for the wheels 36 on either the left wing landing gear or the right
wing landing gear. Thus, EMAC Left1 provides control to actuator modules 1 and 3
of each of the wheels 36 in the left side landing gear (representing the left brakes)
via cable 52. Similarly, EMAC Left2 has its output coupled to the remaining actuator
modules 2 and 4 of the wheels 36 in the left side landing gear via cable 52. EMAC
Right1 similarly provides power to the actuator modules 1 and 3 for the wheels 36
in the right side landing gear (representing the right brakes), and EMAC Right2
provides power to the remaining actuator modules 2 and 4 in the right side landing
gear via another cable 52.
Thus, when the system 30 is fully operational (i.e., during
normal operation) each of the EMACs 44 receives brake commands from BSCU1 and BSCU2
which will be generally redundant. Nevertheless, the EMACs 44 may be configured
to give commands provided by BSCU1 priority or vice versa. In the event commands
are not received from one of the BSCUs 40, the EMACs 44 are configured to default
to the other BSCU 40. During normal operation, all four actuator modules 1-4 will
receive brake control signals from their respective EMAC 44 to provide full braking.
Although not shown in Fig. 3, the outputs of the wheel
speed and torque sensors 62 (if used) for each brake 34 are coupled to the respective
EMACs 44 via the cables 54 (Fig. 2). The EMACs 44 are configured to condition the
signals and provide the measured wheel speed and torque to the BSCUs 40 via the
communication bus 50. The BSCUs 40 in turn use such information in a conventional
manner for carrying out brake control and antiskid processing.
As is shown in Fig. 3, EMAC Left2 and EMAC Right2 differ
from the remaining EMACs in that they also receive left and right proportional brake
commands directly from the transducers 46 via a separate cable 72 (not shown in
Fig. 1). As is discussed in more detail below, such direct input of the brake commands
from the transducers 46 is used during emergency braking operations. Also, EMAC
Left2 and EMAC Right2 receive a parking brake control signal from a switch located
in the cockpit via the cable 72 for carrying out a parking brake operation as described
below.
Continuing to refer to Fig. 3, both BSCU1 and BSCU2 are
designed to operate on DC power. However, BSCU1 is coupled to the DC 1 power source
and BSCU2 is coupled to a different power source, namely the DC2 power source. Thus,
different power busses (e.g., PWR1 and PWR2) are used to supply operating power
to the respective BSCUs 40. Similarly, EMAC Left1 and EMAC Right1 are designed to
operate on power from the different power busses PWR1 and PWR2, respectively. Specifically,
EMAC Left1 receives AC operating power from the AC1 source and DC operating power
from the DC1 source. EMAC Right1 receives AC operating power from the AC2 source
and DC operating power from the DC2 source.
EMAC Left2 and EMAC Right2 are configured to operate on
power from the PWRess power buss. Specifically, both EMAC Left2 and EMAC Right2
receive AC operating power from the ACess source and DC operating power from the
DCess source. In addition, EMAC Left2 and EMAC Right2 are designed to operate in
an emergency mode based on power provided by the DChot bus as discussed below.
The system 30 is designed to carry out built-in testing
among the EMACs 44 to detect the loss of power from any of the primary power busses
PWR1, PWR2 and PWRess. Such built-in testing can be carried out by configuring the
EMACs 44 to poll each other via the communication bus 50, for example. If an EMAC
44 fails to respond to polling by another, for example, it can be assumed that power
from the particular power buss servicing the EMAC 44 is unavailable or that the
EMAC 44 itself has failed. The polling EMACs 44 then communicate such information
to the BSCUs 40 via the bus 50. The BSCUs 40 in turn command the functioning EMACs
44 to revert to an alternate mode of braking. Other techniques for detecting the
loss of power on one of the power busses or the failure of one of the components
can be used without departing from the scope of the invention as will be appreciated.
For example, the BSCUs 40 may instead be configured to
poll each EMAC 44 via the communication bus 50. If an EMAC 44 fails to respond,
the BSCU(s) 40 recognize the problem EMAC 44 and in turn command the functioning
EMACs 44 to revert to an alternate mode of braking.
Braking Modes:
The braking system 30 includes five primary operating modes
for purposes of the present invention, including a normal mode, alternate mode 1,
alternate mode 2, emergency mode and park (ultimate) mode. In each mode braking
is available despite failure of a power buss, etc., as will now be explained with
reference to Figs. 4A-4D and 5A-5B.
Figs. 4A-4D and 5A-5B illustrate the state of respective
power busses and components within the system 30 with respect to time during different
failure modes. A line level "A" in the figures indicates that the power buss or
component is available and operational. A line level "IN" indicates that the power
buss or component is inactive or unavailable. With respect to a line level between
"A" and "IN", this indicates that the brakes or components are partially available
or operational as will be further described below.
Normal Mode:
Normal mode operation is defined as operation during which
power from all the primary power busses PWR1, PWR2 and PWRess is available, and
the BSCUs 40 and EMACs 44 are functional. Referring initially to Fig. 4A, normal
mode operation is shown at a time prior to a failure time tf. As is shown, all of
the power busses are available, the BSCUs 40 and EMACs 44 are receiving power and
are operational. Moreover, each of the actuator modules 1-4 in the left brakes and
right brakes are powered and operational.
Alternate Mode 1:
Alternate mode 1 is defined as operation during which the
power buss PWR1 or PWR2 is unavailable due to failure, for example, but the power
buss PWRess remains available.
Fig. 4A illustrates a particular example where, at a failure
time tf, the power buss PWR1 (AC1/DC1) fails. As noted above, such failure may occur
due to engine failure, power converter failure, broken power cable, etc. Since BSCU1
is powered by the power buss PWR1, BSCU1 will stop functioning at time tf as represented
in Fig. 4A. However, since BSCU1 and BSCU2 are redundant and BSCU2 still receives
operating power from the power buss PWR2 (AC2/DC2), brake control operation and
antiskid processing may still be carried out.
Since BSCU2 receives operating power from power buss PWR2
and therefore does not require power from power buss PWR1, BSCU2 is isolated from
power buss PWR1 as well as BSCU1. Thus, a failure of power buss PWR1 and/or BSCU1
will not produce a consequential failure of power buss PWR2. For example, a short
circuit or breakdown of the power buss PWR1 and/or BSCU1 will not result in a catastrophic
failure of power buss PWR2. Of course, the same is true with respect to the reverse
situation if BSCU2 and/or PWR2 were to experience a failure. Power bus PWR1 would
remain available to BSCU1 as it is isolated within the braking system 30 from the
failed BSCU2 and/or power buss PWR.
Since EMAC Left1 receives power from the power buss PWR1,
it also becomes unavailable at time tf. Because EMAC Left1 becomes unavailable,
the actuator modules 1 and 3 controlled by the EMAC in the left brakes are disabled.
Nevertheless, each of the remaining EMACs 44 remain operational. Accordingly, two
of the four actuator modules (i.e., 2 and 4) remain available for braking as controlled
by the EMAC Left2. Ordinarily this would result in a loss of 50% of the total available
braking force on the left wheels 36. However, the EMACs 44 are designed to increase
the upper force limit exerted by the respective actuator modules 1-4 in the alternate
mode.
For example, the limit for the maximum braking force applied
by each of the remaining two actuators 2 and 4 is increased by the EMAC Left2 by
60%. Hence, the total braking force for the left brakes can achieve 80% of the normal
braking capability. In another example, the maximum braking force limit can be adjusted
by some other amount.
The aforementioned built-in testing detects the loss of
the power buss PWR1. This results in the BSCU2 informing the EMAC Left2 to increase
the braking force limit. Even absent such compensation, 50% braking is still available.
Thus, as is shown in Fig. 4A, partial braking for the left brakes is available even
after time tf.
The risk that the power buss PWR2 may become disabled as
a consequence of the failure of power buss PWR1 (or the failure of EMAC Left1 itself)
is avoided in accordance with the present invention. The remaining EMACs 44 and
the power provided thereto are isolated within the system 30 from the power buss
PWR1.
Similar operation to that shown in Fig. 4A would occur
if the power buss PWR2 (AC2/DC2) failed rather than the power buss PWR1. In such
case, however, BSCU1 would remain operational and BSCU2 would fail. Similarly, EMAC
Right 1 would fail and the remaining EMACs 44 would continue to operate. The actuator
modules 1 and 3 in the right brakes would be disabled, but the EMAC Right2 would
increase the maximum force limit of the actuator modules 2 and 4, similar to that
previously described.
Alternate Mode 2:
Alternate mode 2 is defined as operation during which the
power buss PWRess is unavailable due to failure, for example, but the power busses
PWR1 and PWR2 remain available.
For example, Fig. 4B illustrates how the power buss PWRess
fails at time tf while power busses PWR1 and PWR2 remain active. In such case, EMAC
Left2 and EMAC Right2 are considered unavailable by the system 30 as shown. Although
EMAC Left2 and EMAC Right2 receive power from the DChot bus, such power is utilized
only in the emergency mode discussed below.
Since EMAC Left2 and EMAC Right2 are not operational, the
actuator modules 2 and 4 for each of the brakes 34 for the left and right wheels
36 are disabled. In this case, only 50% of the actuator modules 1-4 are active for
each of the brakes 34. Nevertheless, failure of the PWRess is detected and the BSCUs
40 instruct the remaining EMAC Left1 and EMAC Right1 to increase the force limits
of the active actuator modules 1 and 3 so as to provide a higher percentage of the
normal braking force. Again, this reduced braking function in the left and right
brakes is reflected in Fig. 4B.
It will again be appreciated that according to the present
invention, failure of the power buss PWRess and/or EMAC Left2 or EMAC Right 2 will
not result in a consequential failure of the power buss PWR1 or PWR2 or the remaining
EMACs since the power from power buss PWRess is provided separately to the EMAC
Left 2 and EMAC Right 2. The power to the EMACs Left 1 and Right 1 is provided separately
by the other power busses, and hence avoids consequential failure. Again, the reverse
is also true.
Emergency Mode:
The emergency mode is defined as failure of all the primary
power sources PWR1, PWR2 and PWRess. Only the DChot power source remains available.
Fig. 4C illustrates the emergency mode where all the primary
power sources PWR1, PWR2 and PWRess fail at or before time tf. In such case, both
BSCUs 40 become disabled as does EMAC Left1 and EMAC Right1. Only EMAC Left2 and
EMAC Right2 remain active on a limited basis by virtue of the DChot power source.
EMAC Left2 and EMAC Right2 are configured to recognize such condition and are designed
to operate under condition on the brake commands provided directed thereto from
the transducers 46 via cable 72.
Under such condition, only actuator modules 2 and 4 remain
active in each brake 34. EMAC Left2 and EMAC Right2 are designed to use the pedal
input commands received directly from the transducers 46 to achieve proportional
brake force application using the actuator modules 2 and 4 in each brake 34. Such
pedal input commands may derive power from the DChot source via the connecting cables
72 and 48, and the system 30 preferably is designed to provide the most direct electrical
path between the transducers 46 and the brakes 34 to minimize the number of intermediate
components, and hence decrease the possibility of component failure in that path.
Since only actuator modules 2 and 4 remain active in each
brake, it is preferable that EMAC Left2 and EMAC Right2 be configured to control
the upper force limit of each actuator module under such condition in order to optimize
braking while avoiding wheel lock-up since antiskid protection is not available.
In addition to controlling the upper force limit, or in the alternative, the EMACs
44 may be configured to operate the actuator modules in a pulse mode to avoid wheel
lock-up. It is noted that in the emergency mode, both BSCUs 40 are disabled, and
hence antiskid protection is not available.
Park (Ultimate) Mode:
In the park (ultimate) mode, only power from the DChot
source is available as represented in Fig. 4D. This may be because the aircraft
32 is on the ground with the remaining power systems shut down. Alternatively, all
the primary power busses PWR1, PWR2 and PWRess may have failed similar to the emergency
mode discussed above.
For the same reasons discussed above in relation to Fig.
4C and the emergency mode, only EMAC Left2 and EMAC Right2 remain active in the
park (ultimate) mode. Moreover, these particular EMACs are only partially active
in the sense that they are operating based on power from the DChot source. Operation
differs from the emergency mode in the following respects.
As mentioned above, the cockpit includes a parking brake
switch selectively activated by the pilot. The parking brake switch is coupled to
EMAC Left2 and EMAC Right2 via the cables 48 and 72, for example. EMAC Left2 and
EMAC Right2 are both configured to provide a predetermined fixed braking force via
the enabled actuator modules 2 and 4 in each of the brakes 34 upon closing of the
parking brake switch. Power from the DChot source is used only to actuate the actuator
modules 2 and 4 into position. Thereafter, a mechanical holding device within the
actuator module holds the actuator mechanism in place so as to no longer require
power from the DChot source. In this manner, the park mode uses power only during
activation or when the park brake is released in order to conserve power in the
aircraft battery.
Release of the parking brake is implemented by removing
the brake clamping force as a result of the EMAC Left2 and EMAC Right2 disabling
the mechanical holding device and driving each actuator module 2 and 4 to a running
clearance position. Specifically, the parking brake switch in the cockpit being
moved to a release position causes the EMAC Left2 and EMAC Right2 to release the
parking brake.
In the event the power buss PWRess is available, the system
can be designed to operate on power from DCess in order not to discharge the aircraft
battery serving as the DChot Source.
The park (ultimate) mode is considered to be a final means
of applying brakes in an aircraft emergency situation in order to stop the aircraft.
The EMACs are configured preferably such that the park mode overrides any normal
braking commands unless the normal braking command torque level is higher than the
park torque level. If the remainder of the system 30 fails due to the BSCUs 40 or
the main power busses PWR1, PWR2 and PWRess failing, for example, it is noted that
operation of the park (ultimate) mode is neither prevented nor delayed.
Referring now to Fig. 5A, a case where one of the BSCUs
40 fails is illustrated. For example, Fig. 5A shows how BSCU1 may fail at time tf
due to component failure. Since BSCU1 and BSCU2 are redundant, the EMACs 44 will
continue to receive brake commands from BSCU2. Hence, the system 30 will continue
to operate in a normal mode. Although not shown, if BSCU2 were also to fail for
some reason (e.g., component failure), the EMACs 44 are configured to revert to
emergency mode operation. More specifically, in the absence of commands from the
BSCUs 40, EMAC Left2 and EMAC Right2 are configured to operate proportionally in
the emergency mode based on the direct inputs from the brake pedal transducers 46
as described above.
The failure of the BSCU1 may create a short circuit or
other adverse condition which could cause the power buss PWR1 to fail due to its
connection to BSCU1. In accordance with the present invention, however, BSCU2 and
the power buss PWR2 are isolated within the braking system 30. Thus, failure of
BSCU1 and/or power buss PWR1 will not result in a consequential failure of power
buss PWR2. The same principles apply if BSCU2 was to fail instead.
Although not shown, if BSCU2 was also to fail for some
reason (e.g., component failure), the EMACs 44 are configured to revert to emergency
mode operation. More specifically, in the absence of commands from the BSCUs 40,
EMAC Left2 and EMAC Right2 are configured to operate proportionally in the emergency
mode based on the direct inputs from the brake pedal transducers 46 as described
above.
Fig. 5B illustrates how if EMAC Right1 fails at time tf1
due to component failure, for example, the remaining EMACs 44 continue to operate
such that the right brakes continue to provide at least partial braking. If EMAC
Left1 were to then fail at time tf2, for example, partial braking would again still
be available in the left brakes. Thus, the present invention provides protection
against component failure much in the same way as protection against failure of
the power systems.
As in the case of a failed BSCU, the failure of one of
the EMACs could potentially produce a short circuit or other adverse condition which
could cause its respective power buss connected thereto to fail. In accordance with
the present invention, however, the remaining EMACs in addition to providing for
redundancy, receive power from a power buss which is isolated from the failed power
buss within the braking system 30. Thus, a consequential failure of the remaining
power buss(es) is avoided.
Fig. 6 illustrates in detail the particular configuration
of the braking system 30 in accordance with one example of the present invention.
Fig. 7 represents an exemplary architecture for the BSCUs 40. However, it will be
appreciated that each BSCU 40 can have a variety of configurations yet still satisfy
the objects of the invention. Fig. 8 represents an exemplary design of an EMAC 44
and actuator 34 for carrying out the above described functions. Again, however,
the particular design illustrated in Fig. 8 is not intended to limit the scope of
the invention. For example, the actuator 34 may utilize force sensors in place of
position sensors.
Turning now to Figs. 9-11, alternative embodiments of the
present invention will now be discussed. Referring initially to Fig. 9, an electromechanical
braking system which incorporates redundant centralized controllers with power drive
circuits is designated 80. In the exemplary embodiment, the system 80 includes two
identical centralized controllers 82a and 82b. Each controller 82a and 82b includes
a BSCU controller as discussed above, combined with power drive circuits (EMACs)
for each brake actuator to be driven by the BSCU controller. Thus, in the embodiment
of Fig. 9 the BSCU 40 and EMACs 44 are combined into a centralized controller 82.
As shown in Fig. 9, the controllers 82a and 82b are redundant
in that each receives brake commands from the transducers 46 via cable 48. The output
of each controller 82a and 82b is coupled to the brake actuator modules 1 and 2
for each wheel 36 in both the left wheel brakes and the right wheel brakes. The
outputs from the torque and wheel speed sensors 62 for each of the wheels 36 is
coupled to both controllers 82a and 82b.
Each controller 82a and 82b processes the brake commands
received via cable 48 and outputs brake actuator drive signals to the actuator modules
1 and 2 for each wheel, thus providing a fundamental form of redundancy. If the
BSCU in one of the controllers (e.g., 82a) was to fail, the BSCU in the other controller
(e.g., 82b) would still function to provide full braking capabilities. If a given
EMAC within one of the controllers 82 was to fail, the corresponding EMAC in the
other controller would still be available to provide the necessary drive signals
to the respective brake actuator module.
The controllers 82a and 82b preferably are contained in
their own respective enclosures within the aircraft. Power is provided to the respective
controllers 82a and 82b via different power busses as in the previous embodiment,
or via the same power buss. The advantage of providing power via different power
busses is that if one power buss was to fail, the controller 82 driven by the other
power buss would remain active.
Fig. 10 shows an electromechanical braking system 84 which
utilizes redundant BSCUs 40 as in the embodiment of Fig. 3. In addition, the left
brakes and the right brakes each include redundant EMACs 44. In this embodiment,
however, the EMACs 44 are located within the landing gear adjacent the actuators
34. Moreover, power is provided from a centralized power converter located withing
the root of the wing of the aircraft.
More particularly, redundant BSCUs 1 and 2 receive brake
command signals from the transducers 46 via cable 48 as in the previous embodiments.
The BSCUs 1 and 2 provide brake control signals to each of a plurality of redundant
EMACs 44 included for each of the left wheel brakes and the right wheel brakes.
In the exemplary embodiment, the left wheel brakes are controlled by two EMACs,
namely EMAC1 and EMAC2. The right wheel brakes are controlled by two EMACs, namely
EMAC3 and EMAC4. EMAC1 and EMAC2 each receive brake control signals from both BSCUs
1 and 2, and provide redundant drive signals to each of actuators 1 and 2 for both
left wheels 36. Similarly, EMAC3 and EMAC4 each receive brake control signals from
both BSCUs, and provide redundant drive signals to each of actuators 1 and 2 in
both right wheels 36.
If one of the BSCUs (e.g., BSCU1) was to fail in the embodiment
of Fig. 10, full brake control would still be available by virtue of the other BSCU
(e.g., BSCU2). If one of the EMACs (e.g., EMAC3) was to fail, the other EMAC (e.g.,
EMAC4) would still be available to provide the appropriate drive signals to the
actuators.
Power is provided to the BSCUs via different power busses
as in the embodiment of Fig. 3, or the same power buss as discussed above. In the
exemplary embodiment, power is provided to the EMACs via a power converter 88 located
in the wing root of the aircraft. The converter 88 receives AC and DC power from
one or more power busses and converts the power into a operating line voltage Vemac
which is delivered to EMACs 1 thru 4. Preferably, the converter 88 is designed to
receive power from two or more different power busses in order to provide redundancy
in the event one of the power busses was to fail.
Fig. 11 illustrates another embodiment of an electromechanical
braking system which is designated 90. Similar to the embodiment of Fig. 10, the
system 90 includes redundant BSCUs 1 and 2 for processing brake commands received
from the pedal transducers via cable 48. The EMACs 44 are again located in the landing
gear adjacent the brake actuator modules which, in this example, consist of three
actuator modules 1-3 per wheel 36. EMAC1 receives brake control signals from both
BSCU1 and BSCU2, and in turn drives actuators 1 thru 3 for the left wheels. EMAC2
also receives brake control signals from both BSCU1 and BSCU2, and instead drives
actuators 1 thru 3 in connection with the right wheels. In this example, the EMACs
are located at the bottom of the landing gear, closer to the respective actuator
modules 1-3. This allows the length of the power cables between the EMACs and the
actuator modules to be minimized.
The various embodiments described herein provide for different
levels of redundancy in the event of equipment failure, power failure, or both.
In many instances a particular number of redundant BSCUs, EMACs, etc. are described.
However, it will be appreciated that different numbers of redundancy in BSCUs, EMACs,
etc., are possible depending upon the number of wheels, brakes, actuators, etc.
The present invention is intended to include any and all such possible numbers.
Although the invention has been shown and described with
respect to certain preferred embodiments, it is obvious that equivalents and modifications
will occur to others skilled in the art upon the reading and understanding of the
specification. For example, although the present invention has clear utility in
connection with an aircraft, the braking system described herein can also be used
on other type vehicles without departing from the scope of the invention. Also,
although the invention is described primarily in the context of an electromechanical
braking system it will be appreciated that the invention has application to other
systems such as hydraulic, pneumatic, etc. The present invention includes all such
equivalents and modifications.