U.S. patent application number 16/437456 was filed with the patent office on 2020-03-12 for hybrid electric aircraft propulsion system.
The applicant listed for this patent is PRATT & WHITNEY CANADA CORP.. Invention is credited to Richard Freer, Eric Latulipe.
Application Number | 20200079515 16/437456 |
Document ID | / |
Family ID | 67875361 |
Filed Date | 2020-03-12 |
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United States Patent
Application |
20200079515 |
Kind Code |
A1 |
Latulipe; Eric ; et
al. |
March 12, 2020 |
HYBRID ELECTRIC AIRCRAFT PROPULSION SYSTEM
Abstract
An aircraft comprising a hybrid electric aircraft propulsion
system. The system comprises a first sub-assembly having a first
electric propulsor assembly and a first thermal propulsor assembly,
the first thermal propulsor assembly having a first thermal engine,
a first generator and a first rotating propulsor, the first
electric propulsor assembly attached to the aircraft at a first
location and the first thermal propulsor assembly attached to the
aircraft at a second location. The system also comprises a second
sub-assembly having a second electric propulsor assembly and a
second thermal propulsor assembly, the second thermal propulsor
assembly having a second thermal engine, a second generator and a
second rotating propulsor, the second electric propulsor assembly
attached to the aircraft at a third location and the second thermal
propulsor assembly attached to the aircraft at a fourth
location.
Inventors: |
Latulipe; Eric; (Ste-Julie,
CA) ; Freer; Richard; (Saint-Basile-Le-Grand,
CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
PRATT & WHITNEY CANADA CORP. |
Longueuil |
|
CA |
|
|
Family ID: |
67875361 |
Appl. No.: |
16/437456 |
Filed: |
June 11, 2019 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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62727673 |
Sep 6, 2018 |
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62727678 |
Sep 6, 2018 |
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62727681 |
Sep 6, 2018 |
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62727683 |
Sep 6, 2018 |
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62729818 |
Sep 11, 2018 |
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62731384 |
Sep 14, 2018 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B64D 31/06 20130101;
B64D 27/10 20130101; B60L 53/24 20190201; H02P 2101/25 20150115;
H02P 5/74 20130101; B64D 31/00 20130101; F05D 2220/323 20130101;
H02K 19/34 20130101; H02K 47/04 20130101; B60L 50/51 20190201; F02K
3/06 20130101; F01D 15/10 20130101; B60L 50/13 20190201; Y02T 50/64
20130101; B64D 27/24 20130101; H02P 9/008 20130101; H02P 27/06
20130101; B64D 35/02 20130101; B60L 2220/20 20130101; H02K 2213/06
20130101; H02P 6/12 20130101; B64D 27/02 20130101; B60L 2210/42
20130101; F05D 2220/764 20130101; G08C 19/38 20130101; H02K 51/00
20130101; B64D 27/26 20130101; H02P 2101/30 20150115; B60L 50/60
20190201; B60L 2200/10 20130101; B64D 27/12 20130101; B64D 35/08
20130101; B64D 2027/026 20130101; H02K 41/03 20130101; H02P 17/00
20130101; H02P 6/005 20130101 |
International
Class: |
B64D 27/02 20060101
B64D027/02; B64D 27/24 20060101 B64D027/24; B64D 27/12 20060101
B64D027/12; B60L 50/13 20060101 B60L050/13; B60L 50/51 20060101
B60L050/51; B60L 50/60 20060101 B60L050/60; B60L 53/24 20060101
B60L053/24 |
Claims
1. An aircraft comprising a hybrid electric propulsion system, the
system comprising: a first sub-assembly having a first electric
propulsor assembly and a first thermal propulsor assembly, the
first thermal propulsor assembly having a first thermal engine, a
first generator and a first rotating propulsor, the first electric
propulsor assembly attached to the aircraft at a first location and
the first thermal propulsor assembly attached to the aircraft at a
second location; and a second sub-assembly having a second electric
propulsor assembly and a second thermal propulsor assembly, the
second thermal propulsor assembly having a second thermal engine, a
second generator and a second rotating propulsor, the second
electric propulsor assembly attached to the aircraft at a third
location and the second thermal propulsor assembly attached to the
aircraft at a fourth location.
2. The aircraft of claim 1, wherein the first location and the
third location are adjacent, and the second location and the fourth
location are adjacent.
3. The aircraft of claim 1, wherein the first location and the
fourth location are adjacent, and the second location and the third
location are adjacent.
4. The aircraft of claim 1, wherein the first sub-assembly is
mounted cross-wing to the aircraft, and the second sub-assembly is
mounted cross-wing to the aircraft.
5. The aircraft of claim 1, wherein: the first electric propulsor
assembly comprises a first electric motor, a third rotating
structure coupled to the first electric motor, and a first motor
inverter coupled between a direct current (DC) power source and the
first electric motor; and the second electric propulsor assembly
comprises a second electric motor, a fourth rotating structure
coupled to the second electric motor, and a second motor inverter
coupled between the DC power source and the second electric
motor.
6. The aircraft of claim 5, wherein: the first electric propulsor
assembly further comprises a third electric motor and a fifth
rotating structure coupled to the third electric motor; and the
second electric propulsor assembly further comprises a fourth
electric motor and a sixth rotating structure coupled to the fourth
electric motor.
7. The aircraft of claim 6, wherein: the first electric propulsor
assembly further comprises a third motor inverter coupled to the
third electric motor; and the second electric propulsor assembly
further comprises a fourth motor inverter coupled to the fourth
electric motor.
8. A method for controlling yaw of an aircraft, the method
comprising: operating an aircraft having a hybrid electric
propulsion system; detecting a fault to a first sub-assembly of the
hybrid electric propulsion system, the first sub-assembly having a
first electric propulsor assembly attached to the aircraft at a
first location and a first thermal propulsor assembly attached to
the aircraft at a second location; in response to detecting the
fault: shutting down the first electric propulsor assembly and the
first thermal propulsor assembly; operating the aircraft with a
second sub-assembly of the hybrid electric propulsion system, the
second sub-assembly having a second electric propulsor assembly
attached to the aircraft at a third location and a second thermal
propulsor assembly attached to the aircraft at a fourth location;
and controlling yaw of the aircraft with the second electric
propulsor assembly and the second thermal propulsor assembly.
9. The method of claim 8, wherein the first location and the third
location are adjacent, and the second location and the fourth
location are adjacent.
10. The method of claim 8, wherein the first location and the
fourth location are adjacent, and the second location and the third
location are adjacent.
11. The method of claim 8, wherein the first sub-assembly is
mounted cross-wing to the aircraft, and the second sub-assembly is
mounted cross-wing to the aircraft.
12. The method of claim 8, wherein the first thermal propulsor
assembly comprises a first thermal engine, a first generator and a
first rotating propulsor, and wherein the second thermal propulsor
assembly comprises a second thermal engine, a second generator and
a second rotating propulsor.
13. The method of claim 12, wherein: the first electric propulsor
assembly comprises a first electric motor, a third rotating
structure coupled to the first electric motor, and a first motor
inverter coupled between a direct current (DC) power source and the
first electric motor; and the second electric propulsor assembly
comprises a second electric motor, a fourth rotating structure
coupled to the second electric motor, and a second motor inverter
coupled between the DC power source and the second electric
motor.
14. A hybrid electric aircraft propulsion system comprising: a
first sub-assembly having a first electric propulsor assembly and a
first thermal propulsor assembly, the first thermal propulsor
assembly having a first thermal engine, a first generator and a
first rotating propulsor, the first electric propulsor assembly
attachable to an aircraft at a first location and the first thermal
propulsor assembly attachable to the aircraft at a second location;
and a second sub-assembly having a second electric propulsor
assembly and a second thermal propulsor assembly, the second
thermal propulsor assembly having a second thermal engine, a second
generator and a second rotating propulsor, the second electric
propulsor assembly attachable to the aircraft at a third location
and the second thermal propulsor assembly attachable to the
aircraft at a fourth location.
15. The system of claim 14, wherein: the first electric propulsor
assembly comprises a first electric motor, a third rotating
structure coupled to the first electric motor, and a first motor
inverter coupled between a direct current (DC) power source and the
first electric motor; and the second electric propulsor assembly
comprises a second electric motor, a fourth rotating structure
coupled to the second electric motor, and a second motor inverter
coupled between the DC power source and the second electric
motor.
16. The system of claim 15, wherein: the first electric propulsor
assembly further comprises a third electric motor and a fifth
rotating structure coupled to the third electric motor; and the
second electric propulsor assembly further comprises a fourth
electric motor and a sixth rotating structure coupled to the fourth
electric motor.
17. The system of claim 16, wherein: the first electric propulsor
assembly further comprises a third motor inverter coupled to the
third electric motor; and the second electric propulsor assembly
further comprises a fourth motor inverter coupled to the fourth
electric motor.
18. The system of claim 17, wherein: the first motor inverter is
selectively connected to the first electric motor and the third
electric motor, and the third motor inverter is selectively
connected to the first electric motor and the third electric motor;
and the second motor inverter is selectively connected to the
second electric motor and the fourth electric motor, and the fourth
motor inverter is selectively connected to the second electric
motor and the fourth electric motor.
19. The system of claim 17, wherein: the first generator comprises
a first generator stator connected to the first electric motor and
a second generator stator connected to the third electric motor;
and the second generator comprises a third generator stator
connected to the second electric motor and a fourth generator
stator connected to the fourth electric motor.
20. The system of claim 19, wherein: the first electric motor
comprises a first motor stator connected to the first generator
stator and a second motor stator connected to the first motor
inverter; the second electric motor comprises a third motor stator
connected to the third generator stator and a fourth motor stator
connected to the second motor inverter; the third electric motor
comprises a fifth motor stator connected to the second generator
stator and a sixth motor stator connected to the third motor
inverter; and the fourth electric motor comprises a seventh motor
stator connected to the fourth generator stator and an eighth motor
stator connected to the fourth motor inverter.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present application claims the benefit of U.S.
Provisional Patent Application No. 62/727,673 filed on Sep. 6,
2018, U.S. Provisional Patent Application No. 62/727,678 filed on
Sep. 6, 2018, U.S. Provisional Patent Application No. 62/727,681
filed on Sep. 6, 2018, U.S. Provisional Patent Application No.
62/727,683 filed on Sep. 6, 2018, U.S. Provisional Patent
Application No. 62/729,818 filed on Sep. 11, 2018 and U.S.
Provisional Patent Application No. 62/731,384 filed on Sep. 14,
2018, the contents of which are hereby incorporated by reference in
their entirety.
TECHNICAL FIELD
[0002] The present disclosure relates generally to aircraft
propulsion systems that use power from both an internal combustion
engine and an electric motor.
BACKGROUND OF THE ART
[0003] Hybrid electric aircraft propulsion systems combine internal
combustion and electric propulsion technologies. In an electric
propulsion system, electrical energy is converted to rotational
energy by an electric motor to drive a propulsion fan or a
propeller.
[0004] There are environmental and cost benefits to having at least
a portion of the power for an aircraft propulsion system come from
electric motors. Therefore, there is a need for improvement to
existing architectures.
SUMMARY
[0005] In accordance with a broad aspect, there is provided an
aircraft comprising a hybrid electric aircraft propulsion system.
The system comprises a first sub-assembly having a first electric
propulsor assembly and a first thermal propulsor assembly, the
first thermal propulsor assembly having a first thermal engine, a
first generator and a first rotating propulsor, the first electric
propulsor assembly attached to the aircraft at a first location and
the first thermal propulsor assembly attached to the aircraft at a
second location. The system comprises a second sub-assembly having
a second electric propulsor assembly and a second thermal propulsor
assembly, the second thermal propulsor assembly having a second
thermal engine, a second generator and a second rotating propulsor,
the second electric propulsor assembly attached to the aircraft at
a third location and the second thermal propulsor assembly attached
to the aircraft at a fourth location.
[0006] In accordance with another broad aspect, there is provided a
method for controlling yaw of an aircraft. The method comprises
operating an aircraft having a hybrid electric propulsion system
and detecting a fault to a first sub-assembly of the hybrid
electric propulsion system, the first sub-assembly having a first
electric propulsor assembly attached to the aircraft at a first
location and a first thermal propulsor assembly attached to the
aircraft at a second location. In response to detecting the fault,
the method comprises shutting down the first electric propulsor
assembly and the first thermal propulsor assembly, operating the
aircraft with a second sub-assembly of the hybrid electric
propulsion system, the second sub-assembly having a second electric
propulsor assembly attached to the aircraft at a third location and
a second thermal propulsor assembly attached to the aircraft at a
fourth location, and controlling yaw of the aircraft with the
second electric propulsor assembly and the second thermal propulsor
assembly.
[0007] In accordance with yet another broad aspect, there is
provided a hybrid electric propulsion system. The system comprises
a first sub-assembly having a first electric propulsor assembly and
a first thermal propulsor assembly, the first thermal propulsor
assembly having a first thermal engine, a first generator and a
first rotating propulsor, the first electric propulsor assembly
attachable to an aircraft at a first location and the first thermal
propulsor assembly attachable to the aircraft at a second location.
The system comprises a second sub-assembly having a second electric
propulsor assembly and a second thermal propulsor assembly, the
second thermal propulsor assembly having a second thermal engine, a
second generator and a second rotating propulsor, the second
electric propulsor assembly attachable to the aircraft at a third
location and the second thermal propulsor assembly attachable to
the aircraft at a fourth location.
[0008] Features of the systems, devices, and methods described
herein may be used in various combinations, in accordance with the
embodiments described herein.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] Reference is now made to the accompanying figures in
which:
[0010] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine, in accordance with an illustrative embodiment;
[0011] FIG. 2 is a block diagram of a hybrid electric aircraft
propulsion system, in accordance with an illustrative
embodiment;
[0012] FIG. 3 is a block diagram of a hybrid electric aircraft
propulsion system having two electric motors, in accordance with an
illustrative embodiment;
[0013] FIG. 4 is a block diagram of a block diagram of a hybrid
electric aircraft propulsion system having a dual stator generator,
in accordance with an illustrative embodiment;
[0014] FIG. 5 is a block diagram of a hybrid electric aircraft
propulsion system having dual stator electric motors, in accordance
with an illustrative embodiment;
[0015] FIG. 6 is a block diagram of a hybrid electric aircraft
propulsion system having selective connectivity, in accordance with
an illustrative embodiment;
[0016] FIG. 7 is a block diagram of a hybrid electric aircraft
propulsion system having additional propulsive power, in accordance
with an illustrative embodiment;
[0017] FIGS. 8a-8d are schematic diagrams of a top view of an
aircraft with examples of on-wing configurations for the hybrid
electric aircraft propulsion system;
[0018] FIG. 9 is a block diagram of a first example embodiment for
an electric propulsor assembly;
[0019] FIG. 10 is a block diagram of a second example embodiment
for an electric propulsor assembly;
[0020] FIG. 11 is a block diagram of a third example embodiment for
an electric propulsor assembly;
[0021] FIG. 12 is a block diagram of a fourth example embodiment
for an electric propulsor assembly;
[0022] FIG. 13 is a block diagram of a first example embodiment for
a thermal propulsor assembly;
[0023] FIG. 14 is a block diagram of a second example embodiment
for a thermal propulsor assembly;
[0024] FIG. 15 is a block diagram of an example embodiment of the
hybrid electric aircraft propulsion system of FIG. 8;
[0025] FIG. 16 is a schematic diagram of a motor inverter, in
accordance with an illustrative embodiment; and
[0026] FIG. 17 is a block diagram of an example controller, in
accordance with an illustrative embodiment.
[0027] It will be noted that throughout the appended drawings, like
features are identified by like reference numerals.
DETAILED DESCRIPTION
[0028] There is described herein a hybrid electric aircraft
propulsion system and method. The aircraft propulsion system uses
power generated by a thermal engine and power generated by an
electric generator.
[0029] There is also described an aircraft comprising a hybrid
electric aircraft propulsion system. The system comprising a first
sub-assembly and a second sub-assembly. In some embodiments, the
aircraft has a first wing and a second wing and the first and
second sub-assemblies are mounted to the first and second wings. In
some embodiments, each sub-assembly is mounted cross-wing on the
aircraft. In some embodiments, the aircraft has a single wing, for
example a flying wing aircraft, and the sub-assemblies are mounted
to an aircraft body at first and second locations.
[0030] Each sub-assembly has a first electric propulsor assembly
and a first thermal propulsor assembly, the first thermal propulsor
assembly comprising a first thermal engine, a first generator
coupled to the first thermal engine and to the first electric
propulsor assembly, and a first rotating structure coupled to the
first thermal engine. The first electric propulsor assembly is
attached to the aircraft at a first location and the first thermal
propulsor assembly is attached to the aircraft at a second
location. The second electric propulsor assembly is attached to the
aircraft at a third location and the second thermal propulsor
assembly is attached to the aircraft at a fourth location.
[0031] FIG. 1 illustrates an example thermal engine of a type
preferably provided for use in subsonic flight, namely a gas
turbine engine 10. The gas turbine engine 10 generally comprises in
serial flow communication, a fan 12 through which ambient air is
propelled, a compressor section 14 for pressurizing the air, a
combustor 16 in which the compressed air is mixed with fuel and
ignited for generating an annular stream of hot combustion gases,
and a turbine section 18 for extracting energy from the combustion
gases. High pressure rotor(s) 20 of the turbine section 18 are
drivingly engaged to high pressure rotor(s) 22 of the compressor
section 14 through a high pressure shaft 24. Low pressure rotor(s)
26 of the turbine section 18 are drivingly engaged to the fan rotor
12 and to other low pressure rotor(s) (not shown) of the compressor
section 14 through a low pressure shaft 28 extending within the
high pressure shaft 24 and rotating independently therefrom.
[0032] Although illustrated as a turbofan engine, the gas turbine
engine 10 may alternatively be another type of engine, for example
a turboshaft engine, also generally comprising in serial flow
communication a compressor section, a combustor, and a turbine
section. A turboprop engine may also apply. In some embodiments,
the thermal engine may be of a type other than a combustion engine,
such as a piston engine or a rotary engine. In addition, although
the engine 10 is described herein for flight applications, it
should be understood that other uses, such as industrial or the
like, may apply. Note that a constant volume combustion thermal
machine, other than a piston or a rotary engine, may also be
used.
[0033] Referring now to FIG. 2, there is illustrated an example
embodiment for a hybrid electric aircraft propulsion system 200.
The system 200 presents an AC-AC architecture, whereby alternating
current (AC) electric power is generated and applied directly to an
electric motor. In an AC-DC-AC architecture, the AC electric power
is converted to direct current (DC) and then reconverted to AC
electric power to drive the electric motor. The AC-AC architecture
is more efficient than the AC-DC-AC architecture as there are
losses incurred during the conversion stages from AC to DC and from
DC to AC.
[0034] Thermal engine 201 is operatively coupled to a generator
202. The generator 202 receives motive power (or mechanical energy)
from the thermal engine 201 and converts the motive power into
electrical power. The generator 202 outputs alternating current
(AC) electric power. The AC electric power is then provided
directly to an electric motor 204. The electric motor 204 converts
the AC electric power into mechanical energy in the form of a
rotary force. The rotary force is applied to a rotating structure
206 (i.e. a rotating propulsor), such as a propeller or a
propulsion fan of an aircraft.
[0035] The electric motor 204 also receives AC electric power from
a motor inverter 208. The motor inverter 208 is operatively coupled
to a battery 210 (or the aircraft electrical system). The battery
210 may be a dedicated battery provided for the hybrid electric
aircraft propulsion system 200. Alternatively, the electric power
supplied to the inverter 208 may come from: an auxiliary power
unit, a supplementary power unit, a backup power generator system,
or the aircraft electrical system that does not include batteries
or capacitors. The motor inverter 208 may thus be connected to the
battery 210 via a battery bus on the aircraft, or via dedicated
wiring and/or connectors. Any type of device containing one or more
cells that convert chemical energy directly into electrical energy
may be used as the battery 210. In some embodiments, the battery
210 is based on a non-chemical principal, such as using the
electricity of a supercapacitor in an energy storage mode.
[0036] The motor inverter 208 receives direct current (DC) electric
power from the battery 210 (or another direct current source) and
converts the DC voltage to AC voltage whose frequency and phase is
adjusted to enable the motor 204 to generate mechanical power. The
motor inverter 208 can also be used in the opposite sense to charge
the battery 210 when the electric motor 204 is being driven by
external machine forces, making it behave as a generator, in which
condition the inverter can convert AC voltage back into DC voltage.
In some embodiments, the motor inverter 208 may also be used in a
reverse mode, whereby AC electric power is converted to DC electric
power.
[0037] The electric motor 204 therefore has a first input
operatively coupled to the generator 202 to receive a first source
of AC electric power. The electric motor 204 also has a second
input operatively coupled to the motor inverter 208 to receive a
second source of AC electric power. Note that the electric motor
204 does not need to have two independent inputs (i.e. one from the
generator 202 and one from the inverter 208). The output of the
generator 202 and the output of the inverter 208 may be combined or
connected in parallel before a motor input interface. The electric
motor 204 will, in response to receiving either one of the first
source of AC electric power and the second source of AC electric
power (or both simultaneously), generate a rotating output for
driving the rotating structure 206.
[0038] A controller 212 is configured for selectively driving the
electric motor 204 using the first source of AC electric power from
the generator 202, the second source of AC electric power from the
motor inverter 208, or a combination thereof. When AC electric
power is received concurrently from both the generator 202 and the
motor inverter 208, a greater amount of power is available to the
electric motor 204. This in turn allows for a greater rotary force
to be applied to the rotating structure 206.
[0039] In some embodiments, the first source of AC electric power
provided by the generator 202 is the primary source of propulsion
power for the electric motor 204. For example, under low power
operating conditions, such as in cruise mode, all propulsive power
for the rotating structure 206 may be provided by the generator
202. Under high power operating conditions, such as in climb mode
or take-off mode, a boost of propulsion power may be provided by
the battery 210 through the motor inverter 208. A secondary or
additional source of electric power is thus available for the
electric motor 204 through the motor inverter 208 in order to
supplement the electric power provided by the generator 202.
[0040] In some embodiments, the motor inverter 208 is sized to
match the secondary power requirements of the electric motor 204,
i.e. the motor inverter 208 does not need to be a full-size motor
inverter in order to drive the electric motor 204 on its own. In
addition, there is no need for a generator converter to convert the
voltage from AC to DC, since the electric power generated by the
generator 202 is fed to the electric motor 204 without conversion.
This architecture thus avoids the need for two stages of conversion
during low power operating conditions. In some embodiments, the
battery 210 is recharged directly from the motor inverter 208.
[0041] In some embodiments, the motor inverter 208 is used to
recover from or prevent desynchronization of the generator 202 and
the electric motor 204. When used for recovery of
desynchronization, the motor inverter 208 may resynchronize the
frequency of the electric motor 204 to the frequency of the
generator 202, in response to a command or control signal received
from the controller 212, or in response to logic implemented in the
inverter 208 which monitors and seeks to control the frequency and
phasing of the generator 202 and motor 204. During
resynchronization, the controller 212 may temporarily disconnect
the generator 202 from the electric motor 204, for example by
opening a relay therebetween. The motor inverter 208 may then
adjust the power to the electric motor 204 so that the speed of the
electric motor 204 is modified in such a way to match the frequency
of the electric motor 204 to the frequency of the generator 202.
The electric motor 204 and the generator 202 can then be brought
back in phase with each other. Once the electric motor 204 is back
in phase with the generator 202, the connection between the
generator 202 and the electric motor 204 is restored.
[0042] When used for prevention of desynchronization, the motor
inverter 208 may actively monitor the phase of the electric motor
204. Upon detection of a mismatch in phase between the generator
202 and the electric motor 204, for example in the case where the
motor is slowed down by external forces, the motor inverter 208 may
send an electric signal that is in-phase with the electric motor
204, to provide additional power to bring the electric motor 204
back in phase with the generator 202. This feature may be put into
effect by the controller 212 or it can be incorporated in the
inverter 208.
[0043] In some embodiments, the motor inverter 208 is used to
recharge the battery 210 or provide additional electrical power to
the aircraft electrical systems that are connected to the inverter
208. For example, the motor inverter 208 can increase the power
demand on the generator 202, such that the power produced by the
generator exceeds the power required by the rotating structure 206
or load, and feed the excess power back to the battery 210 or
aircraft electrical system. Alternatively, the motor inverter 208
can increase the power going to the electric motor 204 by feeding
energy from the battery 210 to the electric motor 204. This in turn
increases the power available to the rotating structure 206. The
motor inverter 208 can either increase the available torque to the
rotating structure 206 (ex: for a variable pitch propeller,
additional torque may be required depending on the selected pitch
of the propeller), or it can act as a generator to extract energy
from a windmilling propeller by converting the AC voltage to a DC
voltage that is higher than DC bus voltage in order to recharge the
battery or feed electrical power back to the aircraft electrical
system, if so desired. The inverter 208 has the ability to control
the voltage on the DC bus, thereby controlling the power going to
the DC bus.
[0044] Referring to FIG. 3, there is illustrated another example
embodiment for a hybrid electric aircraft propulsion system 200. In
this example, two electric motors 204a, 204b are driven by the
generator 202. Each electric motor 204a, 204b, is associated with
its own rotating structure 206a, 206b, respectively. Two motor
inverters 208a, 208b are also provided, one to feed each one of the
electric motors 204a, 204b, respectively. It should be understood
that a single motor inverter 308 may also be used instead of two
separate motor inverters 208a, 208b. The motor inverter 308 would
then be sized for the power requirements of two electric motors
204a, 204b instead of just one of the electric motors 204a, 204b.
In addition, the motor inverter 308 would need to be disconnected
from electric motor 204a in order to perform a resynchronization of
electric motor 204b, and disconnected from electric motor 204b in
order to perform a resynchronization of electric motor 204a.
[0045] In some embodiments, the two motor inverters 208a, 208b are
interconnected, in case of a failure of one of the two motor
inverters 208a, 208b. For example, if one motor inverter 208a
fails, the other motor inverter 208b may be used to charge the
battery 210 or to drive the electric motor 204a associated with the
failed motor inverter 208a. This may also be achieved using a
series of connections between the electric motors 204a, 204b and
the motor inverters 208a, 208b to allow for selective connection
and disconnection of the electric motors 204a, 204b to the motor
inverters 208a, 208b.
[0046] In some embodiments, one or more relays 314a, 314b are
provided in the interconnection path between the generator 202 and
each one of the electric motors 204a, 204b, respectively. The
relays 314a, 314b are an example embodiment for allowing selective
connection and disconnection of the generator 202 to either one of
the electric motors 204a, 204b. Other means of
connection/disconnection may also be used.
[0047] The electric motors 204a, 204b may be provided on separate
wings of an aircraft. Alternatively, they may be provided on a same
wing of an aircraft. Also alternatively, one or both of the
electric motors 204a, 204b may be provided in the nacelle or hub of
a rotating propulsion device such as a propeller, fan, lift rotor
or thruster, or in the hub of a contra-rotating propeller, fan
rotor or thruster.
[0048] Three or more electric motors may be provided, whereby each
electric motor 204 is associated with a motor inverter 208.
Alternatively, two or more motor inverters 208 may be combined to
form a larger motor inverter 308 for all of the electric motors 204
or a subset thereof.
[0049] FIG. 4 illustrates yet another embodiment of the hybrid
electric aircraft propulsion system 200. In this example, the
generator 202 is a dual stator generator, having a first generator
stator 402a driving one electric motor 204a and a second generator
stator 402b driving the other electric motor 204b. Note that this
embodiment could also be applied to three or more sets of generator
stators, motors and inverters. Two or more of the inverters 208 may
be combined to form a larger motor inverter 308 for all the
electric motors 204 or a subset thereof. In some embodiments, motor
inverters 208a, 208b comprise a converter to perform the conversion
from AC to DC to recharge the battery 210. The electric motors
204a, 204b are driven independently from each other. Relays 314a,
314b may or may not be present in this embodiment, so as to allow a
selective connection/disconnection of the generator 202 and the
electric motors 204a, 204b.
[0050] FIG. 5 illustrates another example embodiment of the hybrid
electric aircraft propulsion system 200. In this example, one or
more of the electric motors 204a, 204b have dual stators. The
example illustrated shows electric motor 204a having a first motor
stator 502a and a second motor stator 504a, and electric motor 204b
having a first motor stator 502b, and a second motor stator 504b.
Motor stator 502a is operatively coupled to generator stator 402a,
and motor stator 502b is operatively coupled to generator stator
402b. Motor stator 504a is operatively coupled to motor inverter
208a, and motor stator 504b is operatively coupled to motor
inverter 208b. In this configuration, the generator 202 does not
share the same motor stator as the motor inverters 208a, 208b, thus
resulting in the generator power and the motor inverter power being
independent from one another. Should the motor inverter 208a be
used to recharge the battery 210, stator 504b would act as a motor
while stator 504a would act as a generator from which a conversion
from AC to DC is required to recharge the battery 210. Should the
motor inverter 208b be used to recharge the battery 210, stator
504a would act as a motor while stator 504b would act as a
generator from which a conversion from AC to DC is required to
recharge the battery 210.
[0051] Referring to FIG. 6, there is illustrated another embodiment
of the hybrid electric aircraft propulsion system 200. In this
example, additional connections are provided between the generator
202, the electric motors 204a, 204b, and the motor inverters 208a,
208b. Generator stator 402a may be selectively connected to both
stators 502a, 504a of the electric motor 204a by closing relays
314a and 314c and opening relay 314d. Generator stator 402b may be
selectively connected to both stators 502b, 504b of the electric
motor 204b by closing relays 314b and 314e and opening relay 314f.
Motor inverter 208a may be selectively connected to both stators
502a, 502b of the electric motor 204a by closing relays 314d and
314c and opening relay 314a. Motor inverter 208b may be selectively
connected to both stators 502b, 504b of the electric motor 204b by
closing relays 314f and 314e and opening relay 314b. Also, each
stator 502a and/or 502b can be connected to both the generator
stator 402a and the inverter 208A simultaneously, without opening
relay 314a if so desired. This is useful to be able to use the
inverters to either provide additional power or extract power to
recharge batteries, or provide power to the aircraft DC bus, or to
adjust and to maintain the motor and generator in phase with each
other. Likewise for the other motor and associated generator
stator, inverters, relays etc. This architecture provides the
ability for the generator 202 to drive both stators 502a, 504a or
502b, 504b of either electric motor 204a, 204b, and for the motor
inverters 208a, 208b, to drive both stators 502a, 504a or 502b,
504b of either electric motor 204a, 204b.
[0052] The architecture of FIG. 6 also provides the ability to
recharge the battery 210 directly from the generator 202 by closing
relays 314a, 314c, 314d and/or 314b, 314e, 314f, thereby providing
a direct path between the generator 202 and the motor inverters
208a, 208b, respectively. In such a case, one or both of motor
inverters 208a, 208b would need to be provided with an AC-DC
converter to transform the AC power received by the generator 202
into DC power for the battery 210. The architecture of FIG. 6 also
provides the ability for the battery to be used to provide
mechanical power to the thermal engine, by driving the stator 402a
or 402b as an electric motor instead of as a generator, if so
desired, for example if this is useful to start the thermal engine
or drive auxiliary systems such as pumps that are connected to the
generator shaft before the engine is started. The same could be
said of other embodiments.
[0053] Turning to FIG. 7, there is illustrated yet another
embodiment of the hybrid electric aircraft propulsion system 200.
In this example, a rotating structure 206c is operatively coupled
to the thermal engine 201 so as to provide additional propulsion to
the aircraft. Indeed, the motive power generated by the thermal
engine 201 may also be applied directly to a rotating structure
206c such as a propeller or propulsion fan of the aircraft in order
to supplement propulsion provided by rotating structures 206a,
206b.
[0054] It should be noted that while rotating structure 206c is
illustrated in combination with the embodiment of FIG. 4, it may
also be added to any of the other embodiments provided in FIGS. 2,
3, 5 and 6, as well as variants thereof. Variants thereof include,
for example, a three-stator generator connected to three electric
motors. It will be understood that the present embodiments apply to
any number of generator stators and motor stators. Another example
of a variant to the present embodiments is a three stator electric
motor, each stator connected to one of a generator stator, a first
motor inverter, and a second motor inverter. The first and second
motor inverters can be connected to a same battery or to separate
batteries. Other variants will be readily understood by those
skilled in the art.
[0055] The hybrid electric aircraft propulsion system 200 described
herein may be mounted to an aircraft using various configurations.
FIG. 8a is an example embodiment showing an aircraft 800 having a
fuselage 802, a first wing 804 and a second wing 806. The aircraft
800 is equipped with a hybrid electric aircraft propulsion system
comprising two sub-assemblies provided on-wing. A first
sub-assembly comprises a first electric propulsor assembly 808a
mounted to the first wing 804 and a first thermal propulsor
assembly 810a mounted to the first wing 804. The electric propulsor
assembly 808a and the thermal propulsor assembly 810a are connected
together. A second sub-assembly comprises a second electric
propulsor assembly 808b mounted to the second wing 808b and a
second thermal propulsor assembly 810b mounted to the second wing
806. The electric propulsor assembly 808b and the thermal propulsor
assembly 810b are connected together.
[0056] FIG. 8b is an example embodiment of the hybrid electric
aircraft propulsion system provided in a cross-wing configuration.
The first sub-assembly comprises the first electric propulsor
assembly 808a mounted to the first wing 804 at a first location and
the first thermal propulsor assembly 810a mounted to the second
wing 806 at a second location. The electric propulsor assembly 808a
and the thermal propulsor assembly 810a are connected together. The
second sub-assembly comprises the second electric propulsor
assembly 808b mounted to the second wing 808b at a third location
and the second thermal propulsor assembly 810b mounted to the first
wing 804 at a fourth location. The electric propulsor assembly 808b
and the thermal propulsor assembly 810b are connected together. In
this example, the first location and the third location are
adjacent, and the second location and the fourth location are
adjacent. In an alternative embodiment, the first location and
fourth location are adjacent, and the second location and third
location are adjacent, such that the first electric propulsor
assembly 808a is next to the second electric propulsor assembly
808b, and the second thermal propulsor assembly 810b is next to the
first thermal propulsor assembly 810a.
[0057] In some embodiments, instead of being mounted cross-wing,
the sub-assemblies are mounted cross-fuselage, such that each
sub-assembly has one of an electric propulsor assembly and a
thermal propulsor assembly on a given side of the fuselage. Any
cross-mounting of the sub-assemblies may be used in a manner that
balances the thrust produced across the aircraft.
[0058] The cross-mounting of the sub-assemblies, whether
cross-wing, cross-fuselage, or other, allows for a control of yaw
of the aircraft, particularly in the case of a fault to any of the
sub-assemblies 808a, 808b, 810a, 810b. In response to detecting a
fault, for example to the first sub-assembly, the first electric
propulsor assembly 808a and first thermal propulsor assembly 810b
may be shut down and the aircraft may be operated with the second
sub-assembly. The yaw of the aircraft can still be controlled using
the second electric propulsor assembly 810b and second thermal
propulsor assembly 808b.
[0059] FIG. 8c is an example embodiment of the hybrid electric
aircraft propulsion system provided in a cross-wing configuration
and having only one sub-assembly, composed of the first electric
propulsor assembly 808a mounted to the first wing 804 and connected
to the first thermal propulsor assembly 810a mounted to the second
wing 806.
[0060] FIG. 8d is an example embodiment of the hybrid electric
aircraft propulsion system provided as a combination of an on-wing
and off-wing configuration. The first thermal propulsor assembly
810a is mounted to the fuselage 802 and connected to the first
electric propulsor assembly 808a mounted to the first wing 804 and
the second electric propulsor assembly 808b mounted to the second
wing.
[0061] FIG. 9 illustrates a first embodiment for the first electric
propulsor assembly 808a, which may also be duplicated in the second
electric propulsor assembly 808b. The electric motor 204a is
coupled to the rotating structure 206a and to the motor inverter
208a. The motor inverter 208a receives DC electric power from a DC
power source, such as a battery, converts the DC electric power to
AC electric power, and provides the AC electric power to the
electric motor 204a. The electric motor 204a may also receive AC
electric power from a generator. The electric motor 204a converts
AC electric power into a rotary force that is applied to the
rotating structure 206a.
[0062] FIG. 10 illustrates a second embodiment for the first
electric propulsor assembly 808a, which may also be duplicated in
the second electric propulsor assembly 808b. Two separate electric
motors 204a, 204b, are provided and each one is coupled to a
respective rotating structure 206a, 206b. A single motor inverter
208a is coupled to both electric motors 204a, 204b to provide a
secondary source of AC electric power, as described in more detail
herein.
[0063] FIG. 11 illustrates a third embodiment for the first
electric propulsor assembly 808a, which may also be duplicated in
the second electric propulsor assembly 808b. Two separate motor
inverters 208a, 208b are provided, each one associated to a
respective electric motor 204a, 204b, to provide a secondary source
of AC electric power, as described in more detail herein. Each
motor inverter 208a, 208b receives DC electric power for conversion
to AC electric power. The motor inverters 208a, 208b may be
connected to a same DC power source or to separate DC power
sources. In some embodiments, the motor inverters 208a, 208b are
both connected to the aircraft battery through a battery bus.
[0064] FIG. 12 illustrates a fourth embodiment for the first
electric propulsor assembly 808a, which may also be duplicated in
the second electric propulsor assembly 808b. Motor inverters 208a,
208b may be selectively connected to both electric motors 204a,
204b via a plurality of relays 314a, 314b, 314c, 314d. It should be
understood that the relays 314a, 314b, 314c, 314d may be replaced
by other connection/disconnection means, such as but not limited to
a toggle relay that toggles between three positions.
[0065] FIG. 13 illustrates a first embodiment for the first thermal
propulsor assembly 810a, which may also be duplicated in the second
thermal propulsor assembly 810b. The thermal engine 201 is coupled
between the rotating structure 206c and the generator 202. The
motive power generated by the thermal engine 201 is used to rotate
the rotating structure 206c and is converted into AC electric power
by the generator 202.
[0066] FIG. 14 illustrates a second embodiment for the first
thermal propulsor assembly 810a, which may also be duplicated in
the second thermal propulsor assembly 810b. The generator 202 is a
dual stator generator. Each stator generator 402a, 402b
independently provides AC electric power.
[0067] FIG. 15 is a detailed example a hybrid electric aircraft
propulsion system 200, whereby a first sub-assembly 1502a comprises
a first thermal propulsor assembly 810a and a first electric
propulsor assembly 808a, and a second sub-assembly 1502b comprises
a second thermal propulsor assembly 810b and a second electric
propulsor assembly 808b. In this example, the sub-assemblies may be
mounted same-wing, as per the example illustrated in FIG. 8a, or
cross-wing, as per the example illustrated in FIG. 8b.
[0068] The first thermal propulsor assembly 810a comprises a
dual-stator generator 202a, a thermal engine 201a, and a rotating
structure 206c. The second thermal propulsor assembly 810b
comprises a dual stator generator 202b, a thermal engine 201b, and
a rotating structure 206f.
[0069] The first electric propulsor assembly 808a comprises a first
dual stator motor 204a, having a first motor stator 502a and a
second motor stator 504a. A rotating structure 206a is coupled to
the electric motor 204a. The first electric propulsor assembly 808a
also comprises a second dual stator motor 204b having a first motor
stator 502b and a second motor stator 504b. A rotating structure
206b is coupled to the second electric motor 204b. A first motor
inverter 208a is coupled to the motor stator 504a while a second
motor inverter 208b is coupled to the motor stator 504b. Although
not illustrated, the motor inverters 208a, 208b are coupled to a DC
power source such as battery 210. The motor stator 502a is coupled
to the first generator stator 402a and the motor stator 502b is
coupled to the second generator stator 402b.
[0070] The second electric propulsor assembly 808b comprises a
first dual stator motor 204c, having a first motor stator 502c and
a second motor stator 504c. A rotating structure 206d is coupled to
the first electric motor 204c. The second electric propulsor
assembly 808b also comprises a second dual stator motor 204d having
a first motor stator 502d and a second motor stator 504d. A
rotating structure 206e is coupled to the second electric motor
204d. A first motor inverter 208c is coupled to the motor stator
504c while a second motor inverter 208d is coupled to the motor
stator 504d. Although not illustrated, the motor inverters 208c,
208d are coupled to a DC power source such as battery 210. The
motor stator 502c is coupled to the first generator stator 402c and
the motor stator 502d is coupled to the second generator stator
402d.
[0071] It will be understood that all of the variants described
herein are also applicable to the sub-assemblies 1502a, 1502b.
[0072] In some embodiments, the various architectures for the
hybrid electric aircraft propulsion system 200 are used to address
possible failures within the system 200. Various examples of
failure modes are described below.
[0073] Based on the embodiments of FIGS. 2 to 7, propulsion power
for the rotating structure(s) 206 are available via the battery 210
in case of a failure of the thermal engine 201. The available power
is however limited by the power & energy capacity of the
battery 210 or the aircraft DC bus, and the power rating of the
motor inverter(s) 208, 208a, 208b, 308. Similarly, in the case of a
generator failure due to a short circuit, propulsion power for the
rotating structure(s) 206, 206a, 206b is available via the battery
210, but limited by the capacity of the battery 210 and the power
rating of the motor inverter(s) 208, 208a, 208b, 308. In the case
of a battery failure, propulsion power is available but limited by
the power of the generator 202 and the available fuel for the
thermal engine 201.
[0074] Based on the embodiments of FIGS. 3 to 7 where there are at
least two electric motors 204a, 204b, in the case of a failure to
one electric motor 204a, full power is available to the other
functional electric motor 204b. In the case of a failure to one
motor inverter 208a, full power is available to one electric motor
204b and partial power is available to the other electric motor
204a, limited by the power of the generator 202 and the available
fuel for the thermal engine 201. If the two motor inverters 208a,
208b are connected together, then full power may be available to
either electric motor 204a, 204b.
[0075] Based on the embodiments of FIGS. 4 to 7 where the generator
202 has dual stators 402, 404, in the case of a generator failure
due to an open-circuit on one of the two stators 402, 404, full
power is available to one of the two electric motors 204a, 204b.
Partial power is available to the other one of the two electric
motors 204a, 204b, limited by the capacity of the battery 210 and
the power rating of the motor inverter 208a, 208b associated
thereto. In addition, in case of battery depletion, energy from the
functioning generator stator can be transferred to the opposite
motor through the two motor inverters.
[0076] Based on the embodiment of FIG. 5 where each electric motor
204a, 204b has two stators, in the case of a failure to one stator
502a or 504a, some power is still available from the electric motor
204a. If the failure is to stator 502a, power is still available
from the battery 210 and is limited by the capacity of the battery
210 and the power rating of the motor inverter 208a associated
thereto and the power rating of the remaining motor stator 504a. If
the failure is to stator 504a, power is still available from the
generator 202 and is limited by the power of the generator 202 and
the available fuel for the thermal engine 201 and the power rating
of the remaining motor stator 502a.
[0077] Based on the embodiment of FIG. 6, if an inverter 208a or
208b fails, full power is still available for the associated
electric motor 204a or 204b from the generator stator 402a or 402b
with proper relay configuration, limited by generator rating.
Similarly, if a generator stator 402a or 402b fails, both motor
stators of the associated electric motor 204a or 204b may be
powered by the battery 210. The total power available would then
depend on the rating of the inverter 208a or 208b and of the
battery 210.
[0078] Based on the embodiment of FIG. 7, if both electric motors
204a, 204b fail and the aircraft is out of gas, power from the
battery can be transferred to the main shaft of the thermal engine
201 to produce propulsion power for the rotating structure 206c.
The motor inverters 208a, 208b may be connected to the generator
202 to act as a motor and add power to the rotating structure 206C.
If all electric systems fail, the thermal engine 201 can still
provide propulsion power for the rotating structure 206c.
[0079] Based on the embodiment of FIG. 8b, this configuration may
be used to minimize a yaw imbalance created by a failure to one of
the two cross-wing sub-assemblies (808a and 810a; 808b and 810b).
Yaw should be understood as the motion of the aircraft 802 along
the direction 814, which is a rotation perpendicular to central
axis 812 of the aircraft 802. Yaw motion is generally controlled by
the deflection of a rudder, or a hinged section at the rear of the
aircraft 802. When rotating structures on one side of an aircraft
are no longer in operation, a force imbalance is created from the
rotating structures still in operation on the other side of the
aircraft. The force imbalance is countered using the rudder. Using
the configuration illustrated in FIG. 8b, the force imbalance is
minimized in case of a failure to one sub-assembly as the remaining
rotating structures are still provided on each side of the
aircraft. In some embodiments, the size of the rudder may be
reduced due to the minimal impact to the yaw motion in case of a
power failure to one sub-assembly. In other embodiments, the size
of the rudder is maintained but additional performance is obtained
in case of a power failure to one sub-assembly. For example, the
take-off weight may be increased or the take-off distance may be
shortened due to the additional torque that may be generated by the
rudder, as compared to the performance obtained in case of a power
failure when the rudder must compensate for the force imbalance on
one side of the aircraft.
[0080] FIG. 16 is an example embodiment of a motor inverter 208.
There is illustrated one of many possible embodiments that can be
used with the hybrid electric aircraft propulsion system 200
described herein. Lines V+ and V- are connected to the positive and
negative terminals of the battery 210. Lines A, B, C are connected
to an electric motor 204, and represent the three phases of the AC
electric power provided thereto. In some embodiments, lines A, B, C
are connected to motor stator 504a or 504b.
[0081] With reference to FIG. 17, there is illustrated an example
embodiment of controller 212, comprising a processing unit 1702 and
a memory 1704 which has stored therein computer-executable
instructions 1706. FIG. 17 may also represent a control function
embedded in the inverter 208. The processing unit 1702 may comprise
any suitable devices configured to implement the system such that
instructions 1706, when executed by the controller 212 or other
programmable apparatus, may cause the functions/acts/steps as
described herein to be executed. The processing unit 1702 may
comprise, for example, any type of general-purpose microprocessor
or microcontroller, a digital signal processing (DSP) processor, a
central processing unit (CPU), an integrated circuit, a field
programmable gate array (FPGA), a reconfigurable processor, other
suitably programmed or programmable logic circuits, or any
combination thereof.
[0082] The memory 1704 may comprise any suitable known or other
machine-readable storage medium. The memory 1704 may comprise
non-transitory computer readable storage medium, for example, but
not limited to, an electronic, magnetic, optical, electromagnetic,
infrared, or semiconductor system, apparatus, or device, or any
suitable combination of the foregoing. The memory 1704 may include
a suitable combination of any type of computer memory that is
located either internally or externally to the controller 212, for
example random-access memory (RAM), read-only memory (ROM), compact
disc read-only memory (CDROM), electro-optical memory,
magneto-optical memory, erasable programmable read-only memory
(EPROM), and electrically-erasable programmable read-only memory
(EEPROM), Ferroelectric RAM (FRAM) or the like. Memory 1704 may
comprise any storage means (e.g., devices) suitable for retrievably
storing machine-readable instructions 1706 executable by processing
unit 1702. Note that the controller 212 can be implemented as part
of a full-authority digital engine controls (FADEC) or other
similar device, including electronic engine control (EEC), engine
control unit (EUC), and the like.
[0083] The methods and systems described herein may be implemented
in a high level procedural or object oriented programming or
scripting language, or a combination thereof, to communicate with
or assist in the operation of a computer system, for example the
controller 212. Alternatively, the methods and systems may be
implemented in assembly or machine language. The language may be a
compiled or interpreted language. Program code for implementing the
methods and systems may be stored on a storage media or a device,
for example a ROM, a magnetic disk, an optical disc, a flash drive,
or any other suitable storage media or device. The program code may
be readable by a general or special-purpose programmable computer
for configuring and operating the computer when the storage media
or device is read by the computer to perform the procedures
described herein. Embodiments of the methods and systems may also
be considered to be implemented by way of a non-transitory
computer-readable storage medium having a computer program stored
thereon. The computer program may comprise computer-readable
instructions which cause a computer, or in some embodiments the
processing unit 1702 of the controller 212, to operate in a
specific and predefined manner to perform the functions described
herein.
[0084] Computer-executable instructions may be in many forms,
including program modules, executed by one or more computers or
other devices. Generally, program modules include routines,
programs, objects, components, data structures, etc., that perform
particular tasks or implement particular abstract data types.
Typically the functionality of the program modules may be combined
or distributed as desired in various embodiments.
[0085] The above description is meant to be exemplary only, and one
skilled in the art will recognize that changes may be made to the
embodiments described without departing from the scope of the
invention disclosed. Still other modifications which fall within
the scope of the present invention will be apparent to those
skilled in the art, in light of a review of this disclosure.
[0086] Various aspects of the systems and methods described herein
may be used alone, in combination, or in a variety of arrangements
not specifically discussed in the embodiments described in the
foregoing and is therefore not limited in its application to the
details and arrangement of components set forth in the foregoing
description or illustrated in the drawings. For example, aspects
described in one embodiment may be combined in any manner with
aspects described in other embodiments. Although particular
embodiments have been shown and described, it will be apparent to
those skilled in the art that changes and modifications may be made
without departing from this invention in its broader aspects. The
scope of the following claims should not be limited by the
embodiments set forth in the examples, but should be given the
broadest reasonable interpretation consistent with the description
as a whole.
* * * * *