U.S. patent application number 17/471774 was filed with the patent office on 2022-03-31 for aircraft performance optimization based on engine performance monitoring.
The applicant listed for this patent is GE Avio S.r.l.. Invention is credited to Sridhar Adibhatla, Mehdi M. Baladi, Kathleen K. Collins, Aniello Esposito.
Application Number | 20220097864 17/471774 |
Document ID | / |
Family ID | |
Filed Date | 2022-03-31 |
United States Patent
Application |
20220097864 |
Kind Code |
A1 |
Baladi; Mehdi M. ; et
al. |
March 31, 2022 |
AIRCRAFT PERFORMANCE OPTIMIZATION BASED ON ENGINE PERFORMANCE
MONITORING
Abstract
A system and method for optimizing performance of an aircraft or
boat through detection and trending of engine deterioration based
on the performance of the vehicle's gas turbine. The system and
method detects declines in engine power due to either of in-transit
events or over the extended lifetime of the engine, the declines
being due to routine engine part aging or an event. As engine power
gradually or suddenly deteriorates, the system and method lowers a
maximum operating line which defines the safe limits for peak
engine power consumption during flight. For in-transit events, the
system and method detects when actual power consumption is
approaching the current maximum operating line. The controller may
then automate changes to operations of entirely separate aircraft
systems, such as rebalancing electrical energy consumption by
various non-engine elements of the aircraft.
Inventors: |
Baladi; Mehdi M.;
(Cincinnati, OH) ; Esposito; Aniello; (Turin,
IT) ; Adibhatla; Sridhar; (Glendale, OH) ;
Collins; Kathleen K.; (Lynn, MA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
GE Avio S.r.l. |
Rivalta di Torino |
|
IT |
|
|
Appl. No.: |
17/471774 |
Filed: |
September 10, 2021 |
International
Class: |
B64D 31/06 20060101
B64D031/06; B64D 33/02 20060101 B64D033/02; B64D 45/00 20060101
B64D045/00; F02C 9/00 20060101 F02C009/00; G08G 5/00 20060101
G08G005/00 |
Foreign Application Data
Date |
Code |
Application Number |
Sep 30, 2020 |
IT |
102020000023104 |
Claims
1. A computer-readable, non-transitory storage medium storing
instructions that, when executed by a digital controller of an
aircraft comprising a turbine engine causes the aircraft to execute
a method comprising: receiving during a flight of the aircraft, at
the digital controller, a data from an engine sensor of the turbine
engine; determining at the digital controller, based on the data
received from the engine sensor, that a load on the engine exceeds
an allowed maximum operating power; and upon determining that the
load on the engine exceeds the allowed maximum operating power,
issuing a control command to an aircraft system other than the
engine--ASOTE to reduce a power load on the engine.
2. The computer-readable, non-transitory storage medium of claim 1,
wherein the method further comprises: issuing a control command to
the ASOTE to reduce, to modify, or to limit an operation of the
ASOTE, wherein the reduced, modified, or limited ASOTE operation
reduces the electrical load on an electrical power generator which
is driven by the engine.
3. The computer-readable, non-transitory storage medium of claim 1,
wherein the method further comprises: issuing a control command to
the ASOTE to rebalance an air pressure bleed of the aircraft,
wherein the rebalanced air pressure bleed from the aircraft
increases an air compression within a compressor of the turbine
engine.
4. The computer-readable, non-transitory storage medium of claim 1,
wherein the method further comprises at least one of: issuing a
control command to limit an operation of a foreign object damage
door of the aircraft; issuing a control command to limit an
operation of a cabin passenger support system to rebalance an
electrical power requirement of the aircraft; and issuing a control
command to limit an operation of an aviation electronics of the
aircraft.
5. The computer-readable, non-transitory storage medium of claim 1,
wherein the method further comprises: modifying the allowed maximum
operating power, the modification being responsive to a change in
the sensed performance of the engine over a significant engine
degradation time.
6. The computer-readable, non-transitory storage medium of claim 5,
wherein the method further comprises: storing in a memory of the
digital controller an operating log of sensed engine performance
over the significant engine degradation time; calculating based on
the operating line the allowed maximum operating power.
7. A computer-readable, non-transitory storage medium storing
instructions that, when executed by a digital controller of an
aircraft comprising a turbine engine causes the aircraft to execute
a method comprising: storing in a memory of the digital controller
an operating log of sensed engine performance over a significant
engine degradation time; calculating from the operating log a
change in the performance of the turbine engine over the
significant engine degradation time; determining an allowed maximum
operating power for the turbine engine, the determination being
responsive to a change in the sensed performance of the engine over
the significant engine degradation time.
8. The computer-readable, non-transitory storage medium of claim 7,
wherein the method further comprises: storing in the memory of the
digital controller at least one of a factory condition operating
power for the turbine engine and a new jet aircraft operating power
for the turbine engine, and determining the allowed maximum
operating power for the turbine engine based on: the at least one
of a factory condition operating power for the turbine engine and a
new jet aircraft operating power for the turbine engine; and the
change in the sensed performance of the turbine engine over the
significant engine degradation time.
9. The computer-readable, non-transitory storage medium of claim 7,
wherein the method further comprises: determining the changes in
the sensed performance of the turbine engine over the significant
engine degradation time based on sensed changes in performance of a
compressor of the turbine engine over the significant engine
degradation time.
10. The computer-readable, non-transitory storage medium of claim
9, wherein the method further comprises: determining the changes in
performance of the compressor based on changes in constant
efficiency islands of performance on a stall map of the engine
compressor performance over time.
11. The computer-readable, non-transitory storage medium of claim
9, wherein the method further comprises: determining at the digital
controller, based on the sensed data received from the engine
sensor during flight, that a load on the turbine engine exceeds an
allowed maximum operating power; and upon determining that the load
on the turbine engine exceeds the allowed maximum operating power,
issuing a control command to an aircraft system other than the
engine--ASOTE to rebalance a power load on the engine.
12. The computer-readable, non-transitory storage medium of claim
11, wherein the method further comprises: issuing a control command
to the ASOTE to reduce modify or limit an operation of the ASOTE,
wherein the modified or limited ASOTE operation rebalance the
electrical load on an electrical power generator which is driven by
the engine.
13. A controller for use in an aircraft comprising an aircraft
engine with a compressor, the controller comprising: a hardware
processor configured to: determine a level of degradation of the
aircraft engine over a flight time of the aircraft; and responsive
to the determined degradation of the aircraft engine over the
flight time, adaptively rebalance the power requirements of one or
more aircraft systems other than the engine--ASOTE to reduce a
power load on the engine.
14. The aircraft controller of claim 13, wherein the hardware
processor is further configured to modify the designated safety
threshold over a significant engine degradation time, the
modification being a response to a change in the sensed performance
of the turbine engine over the significant engine degradation
time.
15. The aircraft controller of claim 14, wherein the hardware
processor is further configured to: store in a memory of the
controller a log of sensed engine performance over the significant
engine degradation time; calculate based on the log of sensed
engine performance a maximum transient operating line, wherein the
designated safety threshold comprises the maximum transient
operating line.
16. The aircraft controller of claim 13, wherein the hardware
processor is further configured to: receive data from an engine
sensor of the engine during flight; determine based on the data
received from the engine sensor a load on the engine; determine
that the load on the engine exceeds a designated safety threshold;
and upon determining that the load on the engine exceeds the
designated threshold, issue a control command to the ASOTE to
rebalance, reduce, modify, or limit one or more operations of the
ASOTE, wherein the rebalanced, reduced, modified, or limited
operations of the one or more ASOTE reduces the load on the turbine
engine to below the designated safety threshold.
17. The aircraft controller of claim 16, wherein issuing a control
command to the ASOTE comprises issuing a command from the
controller to reduce an electrical power demand on an electrical
power generator which is driven by the turbine engine.
18. The aircraft controller of claim 16, wherein issuing the
control command to the ASOTE comprises the controller issuing a
command to at least one of: reduce a pressure requirement of an
aircraft bleed system; limit an operation of a foreign object
damage door of the aircraft; limit an operation of a cabin
passenger support system to rebalance an electrical power
requirement of the aircraft; and limit an operation of an aviation
electronics of the aircraft.
19. The aircraft controller of claim 16, wherein the controller is
further configured to: issue a plurality of control commands to a
plurality of ASOTE, wherein the plurality of control commands
collectively cause the plurality of ASOTE to reduce a power
required of the engine to below the maximum transient operating
line.
20. The aircraft controller of claim 16, wherein issuing the
control command to the ASOTE comprises the controller issuing a
command to modify an aircraft speed or an aircraft angle of ascent
or descent to reduce a thrust demand on the engine.
Description
TECHNICAL FIELD
[0001] The present invention relates to optimization of aircraft
performance. The present invention further relates to optimization
of multiple features of aircraft performance responsive to
monitoring of aircraft engine performance.
BACKGROUND
[0002] On transport vehicles, including but not limited to
airplanes, it is generally known to have sensors and associated
electronic processing systems which monitor various component units
of the vehicle, and which further modify operating performance of a
component unit in response to the monitoring. For example, on an
aircraft, electronic systems will monitor performance parameters of
a jet engine, and will further modify and sometimes limit engine
operations (such as limiting engine speed) for reasons of
operational efficiency and vehicular safety.
[0003] Aircraft are complex systems, and the performance and
operations of different units (such as engines, wings, interior
electronics) can have interactive and synergistic effects. However,
because the different component units of a large aircraft are often
made be entirely different companies, integration of component unit
monitoring to achieve aircraft-wide changes in operations has not
always been possible. For example, aircraft engines are often
designed, engineered, and manufactured by a first corporate entity;
and the engines so designed are then integrated into aircraft which
are designed and built by some entirely separate corporate
entity.
[0004] At the same time, it may be the case that changes in
real-time performance of operations in one aircraft components,
such as the engines, may be best addressed (in part or in whole) by
making suitable real-time changes in the operations of entirely
different units of the aircraft.
[0005] A further challenge for aircraft design, or more generally
vehicular design, is that component unit performance tends to
degrade over time. An aircraft engine or aircraft engine compressor
with 10,000 hour of operational time may no longer perform as well
as when the same engine was fresh from the factory. Or similarly,
an aircraft engine/compressor which is 10,000 hours from its last
maintenance may no longer perform as well as the same
engine/compressor immediately after maintenance. As such, the
overall design of the aircraft, and operational limitations for the
aircraft, may be based on anticipated worst-case/degraded engine
performance. This means the aircraft does not take advantage of the
more optimal engine performance when the engine is newly built, or
when the engine has just received maintenance. This results in
unnecessary limitations on the performance of the aircraft during
periods of peak or near-peak engine performance.
[0006] What is needed, then, is a system and method for increasing
aircraft performance and safety by determining and automating
modifications to aircraft operations, at a system-wide level, based
on sensed changes over time to operational performance of a
particular hardware subsystem such as the aircraft engines.
SUMMARY OF THE INVENTION
[0007] Controller(s): In at least one aspect, embodiments of the
present system and method entail a controller which detects
declines in overall engine power over the extended lifetime of the
engine, the declines being due to routine engine part aging. As
engine thrust and power gradually deteriorate over time periods of
weeks and months, the controller recalculates--and generally
reduces--a maximum operating line which defines the safe limits for
peak engine power consumption during flight.
[0008] In another aspect, embodiments of the present system and
method entail a controller which detects any deterioration, over
time, in the performance of an aircraft engine (or other transport
vehicle engines, such as boat engines). Based on the detected
variations, the controller may then automate certain changes to
operations of entirely separate aircraft systems, such as making
changes to electrical energy consumption by various elements in the
interior of the aircraft. For example, in response to reduced
engine performance, the controller may limit the electricity
consumed by non-essential electronics within the interior cabin of
the aircraft.
[0009] In another aspect, embodiments of the present system and
method may entail a controller which detects any deterioration,
over time, in the performance of one element of an aircraft engine
(such as the compressor). Based on the detected variations, the
controller may then automate certain changes to operations of other
engine elements (for example, foreign object debris (FOD) doors)
which are not conventionally associated with addressing compressor
performance.
[0010] In another aspect, embodiments of the present system and
method may entail a controller which detects any deterioration,
over time, in the performance of one element of an aircraft engine
(such as the compressor). Based on the detected variations, the
controller may then signal a pilot to take actions which may not
typically be part of conventionally-known or conventionally taught
pilot actions, but which may address operational stresses being
imposed on the aircraft engines.
[0011] In another aspect, embodiments of the present system and
method may entail a controller which detects any deterioration,
over time, in the performance of one element of an aircraft engine
(such as the compressor). Based on the detected variations, the
controller may then take system-wide or aircraft-wide actions or
issue pilot signals to address the degraded engine performance via
system-wide actions, based on the current state of engine
performance; in this way actions taken or signaled by a controller
will be tuned for and responsive to current engine performance, as
opposed to being tuned for or designed for anticipated, potential
worst-case engine performance.
[0012] In another aspect, embodiments of the present system and
method may entail a controller which detects can adjust overall
aircraft performance parameters based on whether the engines are at
peak or near-peak operational performance, versus times when engine
performance is naturally degraded due to extended periods of engine
usage.
[0013] Software: In at least one aspect, embodiments of the present
system and method entail software which operates one or more
aircraft controllers based on data from one or more aircraft
sensors. The software detects declines in overall engine power over
the extended lifetime of the engine, the declines being due to
routine engine part aging. As engine thrust and power gradually
deteriorate over time periods of weeks and months, the software
recalculates--and generally reduces or lowers--a maximum operating
line which defines the safe limits for peak engine power
consumption during flight.
[0014] In another aspect, embodiments of the present system and
method entail software which operates one or more aircraft
controllers based on data from one or more aircraft sensors. The
software detects any deterioration, over time, in the performance
of an aircraft engine (or other transport vehicle engines, such as
boat engines). Based on the detected variations, the software
issues instructions to the controller(s) which may then automate
certain changes to operations of entirely separate aircraft
systems, such as making changes to electrical energy consumption by
various elements in the interior of the aircraft. For example, in
response to reduced engine performance, the software may limit the
electricity consumed by non-essential electronics within the
interior cabin of the aircraft.
[0015] In another aspect, embodiments of the present system and
method may entail a controller which detects any deterioration,
over time, in the performance of one element of an aircraft engine
(such as the compressor). Based on the detected variations, the
controller may then automate certain changes to operations of other
engine elements (for example, FOD doors) which are not
conventionally associated with addressing compressor
performance.
[0016] In another aspect, embodiments of the present system and
method may entail software which operates one or more aircraft
controllers based on data from one or more aircraft sensors. The
software detects any deterioration, over time, in the performance
of one element of an aircraft engine (such as the compressor).
Based on the detected variations, the software may then signal a
pilot to take actions which may not typically be part of
conventionally-known or conventionally taught pilot actions, but
which may address operational stresses being imposed on the
aircraft engines.
[0017] In another aspect, embodiments of the present system and
method may again entail software which operates one or more
aircraft controllers based on data from one or more aircraft
sensors. The software detects any deterioration, over time, in the
performance of one element of an aircraft engine (such as the
compressor). Based on the detected variations, the software may
then issue instructions to the controller to take system-wide or
aircraft-wide actions, or the software, via suitable interfaces,
may issue pilot signals to address the degraded engine performance
via system-wide actions, based on the current state of engine
performance; in this way actions taken or signaled by a software
will be tuned for and responsive to current engine performance, as
opposed to being tuned for or designed for anticipated, potential
worst-case engine performance.
[0018] In another aspect, embodiments of the present system and
method may again entail software which operates one or more
aircraft controllers based on data from one or more aircraft
sensors. The software is configured to adjust overall aircraft
performance parameters based on whether the engines are at peak or
near-peak operational performance, versus times when engine
performance is naturally degraded due to extended periods of engine
usage.
[0019] For example, in response to reduced engine performance, the
software may limit the electricity consumed by non-essential
electronics within the interior cabin of the aircraft.
[0020] Methods: In another aspect, embodiments of the present
system and method may entail various methods of aircraft control,
implemented via controllers, software, firmware, and similar
technologies, the methods implementing the operations and
activities discussed immediately above with respect to controllers
and software embodiments.
BRIEF DESCRIPTION OF THE DRAWINGS
[0021] Advantageous designs of embodiment of the present invention
result from independent and dependent claims, the description, and
the drawing. In the following, preferred examples of embodiments of
the invention are explained in detail with the aid of the attached
drawings:
[0022] FIG. 1 is a schematic perspective diagram of an exemplary
aircraft including an exemplary turbine engine 200.
[0023] FIG. 2 is a schematic cross-sectional diagram of the
exemplary turbine engine 200 of FIG. 1 including an exemplary
turbine engine 200 compressor.
[0024] FIG. 3 is a schematic cross-sectional diagram of the
exemplary engine compressor of FIG. 2 including an exemplary
environmental control system (ECS) and controller.
[0025] FIG. 4A illustrates some exemplary aircraft systems other
than the engines (ASOTE).
[0026] FIG. 4B illustrates an exemplary method for managing
aircraft power consumption during flight.
[0027] FIG. 5A presents an exemplary compressor map (referred to
equivalently herein as a stall map 500) which may be used by an
engine controller in legacy systems.
[0028] FIG. 5B presents an exemplary compressor map (referred to
equivalently herein as a stall map 500) which may be used by an
engine controller according to the present system and method.
[0029] FIG. 5C presents an exemplary map which illustrates how a
maximum operating line for an engine may be updated over the
lifetime of an engine or aircraft, according to the present system
and method.
DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS
[0030] The following detailed description is merely exemplary in
nature and is not intended to limit the system and methods, nor the
elements or steps of the system and method, nor its applications,
and its uses disclosed herein. Further, there is no intention for
the scope to be bound or limited to or by any theory presented in
the preceding background or summary, nor in the following detailed
description.
I. TERMINOLOGY
[0031] It will be understand in this document that:
[0032] Description of various embodiments may use "comprising"
language, indicating that the system and method may include certain
elements or steps which are described; but that the system and
method may also include other elements or steps which are not
described, or which may be described in conjunction with other
embodiments, or which may be shown in the FIG. only, or those which
are well known in the art as necessary to the function of
processing systems. However, it will be understood by one of
skilled in the art that in some specific instances, an embodiment
can alternatively be described using the language "consisting
essentially of" or "consisting of."
[0033] As used herein, the term "forward" or "upstream" refers to
moving in a direction toward the engine inlet, or a component being
relatively closer to the engine inlet as compared to another
component. The term "aft" or "downstream" used in conjunction with
"forward" or "upstream" refers to a direction toward the rear or
outlet of the engine or being relatively closer to the engine
outlet as compared to another component.
II. EXEMPLARY AIRCRAFT
[0034] FIG. 1 illustrates an exemplary aircraft 100 according to
the present system and method. Aircraft 1 may include a fuselage 2,
a cockpit 4 positioned in the fuselage 2, and wing assemblies 6
extending outward from the fuselage 2. The aircraft 100 can also
include multiple engines 10, including turbine engine 200s 200
which by way of non-limiting example can be turbojet, turbofan,
turboprop, or turboshaft engines. While a commercial fixed wing
aircraft 100 has been illustrated, it is contemplated that aspects
of the disclosure described herein can be used in any type of fixed
wing, rotary wing, or convertible wing aircraft such as the V22
Osprey in a non-limiting example. Further, while two turbine engine
200s 200 have been illustrated on each of the wing assemblies 6, it
will be understood that any number of turbine engine 200s 200
including a single turbine engine 200 200 on the wing assemblies 6,
or even a single turbine engine 200 200 mounted in the fuselage 2
can be included. In an example where the aircraft 100 includes a
rotary wing aircraft, multiple fuselage mounted turbine engine 200s
can be included in the aircraft 100.
[0035] Controllers and Avionics:
[0036] The aircraft 100 may include one or more general aircraft
controllers 105, together referred to as the avionics suite 105,
which may be digital controllers with microprocessor(s) and memory.
Ad discussed further below, the aircraft engines 200 may include
one or more engine controllers 110 which may be Full Authority
Digital Controllers 110 (FADECS). Suitable internal communications
links 117 (electrical, fiber optic or wireless) enable digital and
possibly analog communications between the avionics suite 105 and
the FADECS 110.
[0037] The present system and method is particularly applicable to
relatively smaller commercial jet airplanes which are still in
widespread use, but may be applicable to larger commercial
aircraft, military aircraft, helicopters, and boat engines as well.
The illustration and exposition here of the present system and
method with turbine engine 200s for aircraft is exemplary only, and
should not be construed as limiting.
[0038] Turning to FIG. 2, the turbine engine 200 200 has a
generally longitudinally extending axis or centerline 12 extending
forward 14 to aft 16. The engine 200 includes, in downstream serial
flow relationship, a fan section 18 including a fan 20, a
compressor section 300 including a booster or low pressure (LP)
compressor 304 and a high pressure (HP) compressor 306, a
combustion section 28 including a combustor 30, a turbine section
32 including a HP turbine 34, and a LP turbine 36, and an exhaust
section 38.
[0039] The fan section 18 includes a fan casing 40 surrounding the
fan 20. The fan 20 includes a plurality of fan blades 42 disposed
radially about the centerline 12. The HP compressor 306, the
combustor 30, and the HP turbine 34 form a core 44 of the engine
10, which generates combustion gases. The core 44 is surrounded by
core casing 46, which can be coupled with the fan casing 40.
[0040] A HP shaft or spool 48 disposed coaxially about the
centerline 12 of the engine 200 drivingly connects the HP turbine
34 to the HP compressor 306. A LP shaft or spool 50, which is
disposed coaxially about the centerline 12 of the engine 200 within
the larger diameter annular HP spool 48, drivingly connects the LP
turbine 36 to the LP compressor 304 and fan 20. The spools 48, 50
are rotatable about the engine centerline and couple to a plurality
of rotatable elements, which can collectively define a rotor
51.
[0041] The LP compressor 304 and the HP compressor 306 respectively
include a plurality of compressor stages 52, 54, in which a set of
compressor blades 56, 58 rotate relative to a corresponding set of
static compressor vanes 60, 62 to compress or pressurize the stream
of fluid passing through the stage. In a single compressor stage
52, 54, multiple compressor blades 56, 58 can be provided in a ring
and can extend radially outwardly relative to the centerline 12,
from a blade platform to a blade tip, while the corresponding
static compressor vanes 60, 62 are positioned upstream of and
adjacent to the rotating blades 56, 58. It is noted that the number
of blades, vanes, and compressor stages shown in FIG. 1 were
selected for illustrative purposes only, and that other numbers are
possible.
[0042] The blades 56, 58 for a stage of the compressor can be
mounted to (or integral to) a disk 61, which is mounted to the
corresponding one of the HP and LP spools 48, 50. The vanes 60, 62
for a stage of the compressor can be mounted to the core casing 46
in a circumferential arrangement.
[0043] The HP turbine 34 and the LP turbine 36 respectively include
a plurality of turbine stages 64, 66, in which a set of turbine
blades 68, 70 are rotated relative to a corresponding set of static
turbine vanes 72, 74 (also called a nozzle) to extract energy from
the stream of fluid passing through the stage. In a single turbine
stage 64, 66, multiple turbine blades 68, 70 can be provided in a
ring and can extend radially outwardly relative to the centerline
12 while the corresponding static turbine vanes 72, 74 are
positioned upstream of and adjacent to the rotating blades 68, 70.
It is noted that the number of blades, vanes, and turbine stages
shown in FIG. 1 were selected for illustrative purposes only, and
that other numbers are possible.
[0044] The blades 68, 70 for a stage of the turbine can be mounted
to a disk 71, which is mounted to the corresponding one of the HP
and LP spools 48, 50. The vanes 72, 74 for a stage of the
compressor can be mounted to the core casing 46 in a
circumferential arrangement.
[0045] Complementary to the rotor portion, the stationary portions
of the engine 10, such as the static vanes 60, 62, 72, 74 among the
compressor and turbine section 300, 32 are also referred to
individually or collectively as a stator 63. As such, the stator 63
can refer to the combination of non-rotating elements throughout
the engine 10.
[0046] In operation, the airflow exiting the fan section 18 is
split such that a portion of the airflow is channeled into the LP
compressor 304, which then supplies pressurized air 76 to the HP
compressor 306, which further pressurizes the air. The pressurized
air 76 from the HP compressor 306 is mixed with fuel in the
combustor 30 and ignited, thereby generating combustion gases. Some
work is extracted from these gases by the HP turbine 34, which
drives the HP compressor 306. The combustion gases are discharged
into the LP turbine 36, which extracts additional work to drive the
LP compressor 304, and the exhaust gas is ultimately discharged
from the engine 200 via the exhaust section 38. The driving of the
LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the
LP compressor 304.
[0047] A portion of the pressurized airflow 76 can be drawn from
the compressor section 300 as bleed air 77. The bleed air 77 can be
drawn from the pressurized airflow 76 and provided to engine
components requiring cooling. The temperature (T) of pressurized
airflow 76 entering the combustor 30 is significantly increased. As
such, cooling provided by the bleed air 77 is necessary for
operating of such engine components in the heightened temperature
(T) environments.
[0048] A remaining portion of the airflow 78 bypasses the LP
compressor 304 and engine core 44 and exits the engine assembly 200
through a stationary vane row, and more particularly an outlet
guide vane assembly 80, comprising a plurality of airfoil guide
vanes 82, at the fan exhaust side 84. More specifically, a
circumferential row of radially extending airfoil guide vanes 82
are utilized adjacent the fan section 18 to exert some directional
control of the airflow 78.
[0049] Electrical Power and Cooling:
[0050] Some of the air supplied by the fan 20 can bypass the engine
core 44 and be used for cooling of portions, especially hot
portions, of the engine 200, and/or used to cool or power other
aspects of the aircraft 100. Electrical power for the aircraft 100
can be provided by one or more electrical generators which obtains
rotary power from the rotor 51 of engine 100. See FIG. 4 for
additional discussion.
[0051] In the context of a turbine engine 200, the hot portions of
the engine are normally downstream of the combustor 30, especially
the turbine section 32, with the HP turbine 34 being the hottest
portion as it is directly downstream of the combustion section 28.
Other sources of cooling fluid can be, but are not limited to,
fluid discharged from the LP compressor 304 or the HP compressor
306.
[0052] Turning to FIG. 3, the compressor section 300 can include an
LP bleed air line 92 fluidly connected to the stage 52 of the LP
compressor 304, and an HP bleed air line 94 fluidly connected to
the stage 54 of the HP compressor 306. An LP dump valve 96 can be
coupled to the LP bleed air line 92, and an HP dump valve 98 can be
coupled to the HP bleed air line 94 as shown, where the dump valves
96, 98 can be fluidly connected to atmosphere external to the
aircraft 1 (FIG. 1). It will be understood that the stages 52, 54
to which the bleed air lines 92, 94 are connected are exemplary and
can be any stages within the LP compressor 304 and HP compressor
306, respectively.
[0053] The aircraft 100 (FIG. 1) can further comprise an
environmental control system (ECS) 305 and ECS sensors 310, which
can be positioned anywhere within the engine 200 and are
illustrated in FIG. 3 as positioned at the stages 52, 54 of the
compressors 304, 306. It will be understood that the ECS 305 can
manage cabin air supply as well as other factors, such as air
pressurization or cabin temperature, in non-limiting examples. The
ECS sensors 310 can sense an engine operating parameter including,
but not limited to, engine horsepower, engine horsepower
extraction, environmental control system bleed, foreign object
debris door deployment, or a secondary engine parameter including,
but not limited to, engine age, transient effects, temperature
effects, minimum airspeed, altitude, attitude, engine inlet
distortion, or inclement weather conditions, and it should be
appreciated that the sensed parameters can also indicate a demand
for the use of bleed air from the compressors 304, 306. It will be
understood that at least one of the above-listed parameters may be
sensed by the sensors 310, or at least four, or all of the
parameters, or any desired combination, in non-limiting examples.
It will further be understood that other aircraft or engine
parameters which can impact operability of the aircraft 1 may also
be sensed by the sensors 310.
[0054] Engine Controller or FADEC:
[0055] An engine controller 110 (which can be a full authority
digital engine controller, or "FADEC") can also be included in the
turbine engine 200 200. An ECS signal 320 communicating the sensed
parameter can be sent from the sensors 310 to the ECS 305 or the
controller 110, and a controller signal 111 can be sent from the
controller 110 to the dump valves 96, 98 or to other components
(not shown) of the engine in operation.
[0056] The engine controller 110 can include a hardware
microprocessor 115 (or CPU 115) and long-term and/or short-term
memory 112.
[0057] Compressor Map and Stall Margin representation (see also
FIG. 5A and FIG. 5B below for further discussion): A compressor map
500, which is used by the controller 110 to operate the engine 200,
can be stored in the memory 112. The compressor map 500 can include
storage of data representations for a compressor stall line 515, an
operating line 530, and a stall margin 570. As used herein, the
compressor stall line 515 will refer to the pressure ratio 502
beyond which the airflow through the engine 200 can break down or
become unstable, causing an engine stall; and "operating line" 530
or "maximum transient operating line" 530 will refer to the
operating pressure ratio 502 which can safely be below that of the
stall line.
[0058] For purposes of this document and the appended claims, the
following terms are deemed equivalent: "operating line", "operating
power", "operating load", and "operating HPx". For example, the
terms "maximum transient operating line 530", "maximum transient
operating power 530", "maximum transient operating load 530" and
"maximum transient operating HPx" are deemed equivalent terms. For
another example, the terms "allowed operating line", "allowed
operating power", "allowed operating load" and "allowed operating
HPx" are deemed equivalent terms.
[0059] Pressure Ratio and Engine Power:
[0060] The "pressure ratio" can refer to any of the following in
non-limiting examples: the ratio 502 (=P2/P3 or =P3/P2) of air
pressures taken at the inlet and outlet of the compressor section
300, the ratio 502 (=P2/P3 or =P3/P2) of air pressures taken at the
inlet and outlet of the compressors 304, 306 individually, or the
ratio 502 (=P2/P3 or =P3/P2) of air pressures taken at any two
locations within the compressor section 300 as desired such as the
exemplary P2, P3 measurement locations shown in the figure. The
pressure ratio 502 P3/P2 may also be understood as one measure of
the engine power actually being generated during flight.
[0061] Stall Margin:
[0062] The stall margin 570 can be defined as the difference
between the compressor stall line 515 and maximum transient
operating line 530, and it should be understood that the percent
difference between the compressor stall line 515 and operating line
530 can also be used to define a minimum allowed stall margin
570.
[0063] In legacy engines 200, the stall line maximum transient
operating line 530 is set based on a fixed-value, maximum safety
factor stall margin 570, which assumes worst case operating
conditions for the aircraft and which leads to a corresponding
large stall margin. All other factors being equal, the greater the
stall margin 570, the less efficiently the engine 200 is operating.
The present system and method is operative, in some embodiments, to
safely reduce the stall margin and thereby increase engine
efficiency. See further discussion associated with FIG. 5 and other
discussion below.
III. EXEMPLARY AIRCRAFT SYSTEMS OTHER THAN THE ENGINE
[0064] Discussed above (FIGS. 2 and 3) are the engines 200 and
engine compressor 300 of an aircraft 100. As is well known in the
art (as well as to aircraft pilots and passengers), an aircraft 100
contains numerous operational components, sections, and systems 400
apart from the engines 200. There are referred to herein as
Aircraft Systems Other Than The Engine (ASOTE) 400. In various
embodiments, the present system and method adaptively maintains the
demands on aircraft engines 200 to stay within safe limits, taking
advantage of the presence of these ASOTE 400. In response to wear
and degradation of the engines 200 over time, the present system
and method adaptively rebalances the energy, the bleed, and the
power requirements of ASOTE 400 to optimize the power demands on
the engines 200. Rebalancing a power demand, whether for an
aircraft system or subsystem, or for the entire aircraft, may in
some embodiments entail reducing a power demand, such as reducing
an electrical by an aircraft system. In alternative embodiments,
rebalancing a power demand (such as an electrical power demand) may
entail reducing a power demand on one aircraft system or subsystem
while increasing a demand on another. In some instances,
rebalancing a power demand may entail increasing a power demand,
such as an electrical power demand, by at least one system or
subsystem of the aircraft 100.
[0065] FIG. 4 illustrates a handful of the numerous ASOTE 400 which
are typical of many if not most exemplary aircraft 100. It will be
appreciated by persons skilled in the arts that some ASOTE 400,
while not part of the engines proper, are most conveniently mounted
on or within the engines 200. This may be true, for example, of
generators, fuel pumps, and other aircraft elements which may
receive rotary power directly from the engines 200. The ASOTE 400
illustrated in FIG. 4 and discussed herein are exemplary only; in
various embodiments, the present system and method may be
implemented with controlled modifications to the operations of many
different ASOTE. Exemplary ASOTE may include, for example and
without limitation: [0066] One or more electrical generators 400.1:
A radial drive train 402 or gear shaft 402 may take rotary power
from the engine rotor 51, and transfer the rotary power to an
accessory gearbox (AGB) 405 (also referred to as an accessory drive
section 405). The accessory gearbox 45 in turn provides rotational
power for the one or more electrical generators 400.1, which can
then provide electrical power 411 to other powered components and
functions of the aircraft 100, including power for the wings 6
(control of flaps, etc.) and for the interior cabin 2 and cockpit
4. It will be noted that in FIG. 4 the gearbox 405 and generators
400.1 are not shown to scale with the engine 200 (which is included
in the figure to provide context only). It will be further noted
that while the gearbox 405 and generators 400.1 may be mounted
interior to the engine 200 (possibly on either of fan casing 40 or
core casing 46, the generator 400.1 is not considered to be part of
the engine 200; the generator 400.1 is an exemplary ASOTE.
(Aircraft 100 may also include a separate generator which provides
electrical power to the engine 200 itself; in various embodiments,
such an engine electrical generator (not shown in the figures) may
be considered part of the engine.)
[0067] It will be noted that if too much electrical power is
generated by the generators 400.1, engine airflow through the
compressors and engine thrust may be reduced. If the general
environment and operating conditions have placed the engines of
risk of stalling, a temporary reduction in the power demands on the
generators 400.1 may restore a sufficient margin of safety to avoid
engine stalls. [0068] Aircraft Air Bleed System 400.2: Aircraft
bleed systems 400.2 are used to "bleed" or remove excess air
pressure from the aircraft cabin 2, or to increase air pressure, as
necessary. [0069] Foreign Object Debris (F.O.D.) Door 400.3: --A
mechanical separator that separates rocks or clouds of the air.
Foreign objects are understood to be material objects which may be
in the air in front of the aircraft 100, and which pose a risk of
damaging the engines 200 if drawn into the fore of the engine 200.
The FOD door 400.3 provides a pathway and mechanism for ejection of
small foreign objects which may be drawn into the engines 200
during operation.
[0070] A disadvantage of using the FOD doors 400.3 is that the
doors 400.3 may divert some of the air pressure inside the engine
200, which in turn reduces the power that can be generated by the
engines and places increased wear on the engines (and in particular
on the compressors 300); under extreme circumstances, using the FOD
doors 400.3 may increase the risk of stalling. In many aircraft,
the FOD doors 400.3 may be opened or closed (under pilot or
aircraft control), and the amount of air diverted may be regulated
over some range. [0071] Interior cabin passenger support systems
400.4: The interior cabin passenger support systems include any
electronics which may contribute to passenger comfort and
convenience. These systems may include, for example and without
limitation: cabin lighting, air-to-ground communications systems,
temperature regulation, cooking and refrigeration systems, and
other systems which draw electricity.
[0072] Aircraft electricity is provided by electric generators
400.1. In some flight situations, then, it may be beneficial to
temporarily reduce overall electrical demands by cabin systems by
temporarily placing limits on those demands (for example, reducing
the intensity of cabin lighting, turning off cabin video displays,
suspending ground-to-air communications systems for passengers,
deactivating microwave ovens, reducing the demands of food
refrigeration, etc.). [0073] Aircraft Speed and Ascent/Descent
Angle 400.5: Making an aircraft accelerate, and increasing the
altitude or ascent angle of an aircraft 100, both demand
substantial engine power. Limits can be placed on the maximum
velocity, maximum acceleration, and maximum ascent angle for an
aircraft, in order to reduce the maximum demands on the engines
200. In some cases, limits on deceleration and/or angle of descent
may also contribute to constraining the maximum power loads on an
aircraft. [0074] Aviation Electronics 105 (400.6): Aircraft
electricity for control and aviation purposes is also provided by
electric generators 400.1 to avionics 105. Such avionics control
systems 105 may not be broadly amenable to significant power
reductions because they are required for routine flight operations
and safety. Nonetheless, aircraft engineers may identify some
elements of airplane avionics 105 where some limits may be safely
placed on their power demands. Such limitations can be
designed/selected with the criteria that while they might reduce
flight efficiency, they do not reduce flight safety. Any such
limitations, when invoked by the present system and method,
contribute to reducing the power demand on the engines 200.
IV. OVERVIEW OF THE PRESENT SYSTEM AND METHOD
[0075] The present system and method pertains to the use of one or
more electronic controller systems which may be employed to
maintain power consumption gas turbine engine 200s 200 within safe
limits. Typical of such gas turbine engine 200s 200 are the jet
engines used on modern aircraft such as turbo-propeller engines.
However, the present system and method also has applications for
airplanes employing turbofan or turbojet engines. The present
system and method also has applications for turbine engine 200s 200
and system controllers of helicopters 100, and also for shipboard
(that is, maritime) applications as well.
[0076] It will be noted in particular that the "one or more
controller systems" described here may include not only an
electronic controller of a turbine engine 200 200, but also other
electronic controllers for other element of an aircraft or other
transport vehicle 100.
[0077] In general, turbine engine 200s 200 regulate gas flows
(typically air flow 76) and chemical flow, along with employing
combustion, to generate thrust. In a time and spatial sequence
events, a turbine engine 200 200 increases air pressure via
compressors 300, combines the fuel with compressed air in a
combustion chamber 30, ignites/explodes the fuel/air mixture (also
in the combustion chamber 30), and then releases the heated exhaust
at reduced pressure via the turbine sections 32, 34, 36 and exhaust
chamber 38. This process generates thrust for the vehicle 100, and
also powers various rotary components (such as generators 400.1) to
provide additional power and services for the vehicle 100.
[0078] Compressor, Air Pressure, and Engine Stall:
[0079] The engine module that first accepts the incoming air is the
intake fan 40, 42 followed immediately by the compressor 300 (which
may be a multistage compressor 302, 304). A crucial aspect of
turbine engine 200 operations is that, during intake, and in
spatial and time progression of the air from fore to aft, the
ambient air pressure must increase. Somewhat in analogy to riding a
bicycle up a hill, it is never good to "go backwards", that is, it
is undesirable to have at any point a decrease in air pressure
going from fore to aft (until after the combustion process). Should
the air pressure actually fall at any point along with progress of
the air through the compressor 300, the engine 200 will stall, that
is, completely lose power, which can lead to fatal outcomes for an
aircraft 100.
[0080] To prevent engine stalls, legacy engines 200 employ sensing
of engine operational conditions to modify the operations of the
turbine engine 200s 200.
[0081] Engine Monitoring/Sensing and Real-Time Engine Performance
Changes:
[0082] All contemporary engines (not just turbine engine 200s, but
automotive engines and other combustion powered sources) contain
electronic sensors which monitor the performance and environment of
the engine. While aircraft engine sensors not shown in the figures,
they are present at multiple points throughout an aircraft engine
200, and may monitor a variety of parameters and phenomena,
including for example and without limitation: air speed at multiple
points throughout the engine, temperature (T) of the air at
multiple points through the engine, engine part rotation speeds,
engine part temperatures, and air pressures at multiple points
throughout the engines. Such measurements are typically made
frequently or continuously throughout the time of flight of an
aircraft 100.
[0083] Prognostic Health Monitoring System (PHMS):
[0084] A turbine engine 200 200 ages with time, so that some sensed
values may be particularly useful as long-term measures of engine
health. A prognostic health monitoring system 119 (PHMS), which may
be an element of an engine controller 110 or be a separate engine
part element in conjunction with the engine controller 110, and may
be implemented as a dedicated controller chip (for example, PROM),
or may be software or firmware which is stored in controller memory
112 and executed by the processor 115 of engine controller 110. The
PHMS 119 may detect, evaluate, and process various
health-indicative sensed conditions which may include, for example
and without limitation:
[0085] >P1 and P3:
[0086] These are sensed air pressures at the fore 14 and aft 16
stages of the engine compressor 300. P1 vs. P3 (or P3-P1) is
therefore a measure of the change in pressure across the turbine
from fore to aft.
[0087] >ITT:
[0088] Inner Turbine Temperature (T). The turbine heat causes
metallic deterioration over time, which in results in decreased
generation/extraction of high pressure from the compressor 200
(decreased values over time for P3-P1).
[0089] >NGR:
[0090] This is a corrected core rotations speed for the turbines,
as corrected with respect to air speed. As a turbine engine 200 200
degrades over time, the NGR will typically need to increase to
produce a given pressure.
[0091] Aircraft Optimization Based On Engine Performance:
[0092] As noted above, legacy systems employ PHMS 119 data to
modify engine operations as needed to prevent stall conditions. The
present system and method improves on legacy operations in multiple
ways, which include for example and without limitation:
[0093] (i) Reducing the power demands on the engines 200 by making
adjustments to a variety of aircraft systems other than the engines
(ASOTE) 400 (including for example the demands of various aircraft
systems for electrical power); and
[0094] (ii) Dynamically modifying, over time, the threshold
criteria (such as engine operating lines 530) which determine when
the aircraft engines 200 are at risk of stall, the threshold
modifications over time being indicative of declining engine
performance with use.
V. EXEMPLARY METHOD FOR MODIFYING REAL-TIME AIRCRAFT POWER
DEMANDS
[0095] In some embodiments, the present system and method entails
modifying real-time aircraft power demands in response to sensed
environmental circumstances which may increase flight demands on
the engines 200.
[0096] Gas turbine engine 200s 200 operate by regulating a chemical
flow (of jet fuel) and oxygen in the air to generate thrust and to
generate electrical power within the engine. A sequence of pressure
increases in the forward (aft) part of the engine 200, typically
through a multistage compressor 300, raises the pressure of
external air drawn into the engine. Further aft, the pressurized
air is mixed with the jet fuel in the combustion chamber 30 and a
spark (electrical energy) is applied, resulting in the explosion of
the fuel and air. This explosion provides thrust, as well as
providing rotational power for one or more electrical generators
400.1. It will be noted that the electrical generators 400.1
contribute to the power demands on the engine 200.
[0097] In some embodiments of the present system and method, an
analysis of the real-time performance of one or more aircraft
engines 200 during the flight time of an aircraft 100 may indicate
that the demands on an engine 200 are at or approaching a maximum
transient operating line 530. This puts the engine at risk of an
engine stall.
[0098] Aircraft Aging Responses:
[0099] The present system and method evaluates the whole system
loads (from all components of an aircraft 100) on the engines 200,
and rebalances the maximum allowed loads (represented by a maximum
transient operating line 530) over a significant engine degradation
time based, on the downward trend 595 of engine performance over
significant engine degradation. This aspect of the present system
and method is discussed further in conjunction with FIG. 5A-5C and
other figures below.
[0100] Real-Time (in-Flight) Aircraft System Power Adaptions:
[0101] The present system and method also provides for real-time
reductions in the overall power demands during the flight duration
of the aircraft 100, via either of: (i) providing instructions to
the aircraft pilot for manual load adjustments; or (ii) automating
the process of managing the loads, thereby taking some of the load
management work away from the pilot. As is well-known in the art, a
flight duration or flight time of an aircraft 100, for a single
flight, is typically on the order of hours (for example, half an
hour, one hour, or between one and twenty-four hours, though longer
flight durations may be envisioned).
[0102] Legacy aircraft systems adjust the engine performance in
response to engine aging. There are design and historical reasons
for this legacy limitation. In most aircraft (especially commercial
aircraft but also military aircraft), there are entirely separate
designer teams for the engines 200 versus the aircraft 100 as a
whole. (Often the engines and aircraft are designed and built by
entirely separate corporate entities, where the aircraft designer
purchases the engines from an engine builder.) The practical result
is a lack of completely integrated and automated control of the
engines in combination with the aircraft. A significant part of the
work determining power load adjustments, in response to demands on
the engines, is left to the pilots.
[0103] In an embodiment, the present system and method
automatically determines the broader power loads on the entire
aircraft system 100 via the controllers 105, 110. Those broader
power loads include electrical power drawn from the engines 200,
which reduces the available pressure in the engines and puts an
engine compressor close to the stall margin 580 and the maximum
transient operating line 530. The present system and method
provides control elements and means to limit the overall power
loads on the aircraft 100 so that those loads remain within safe
power limits.
[0104] In some embodiments, the present system and method further
functions to determine the broad deterioration trends 595 on the
engines over a significant engine degradation time. The system and
method may make the pilot(s) aware of the reductions in power over
time, so that pilots and mission planners can rebalance the engine
loads over significant engine degradation times (such as over
months or years). The present system and method further makes the
aircraft avionics 105 aware of the need to reduce power demands on
the engines 200 as the aircraft ages. Put another way, the present
system and method automates the process of managing aircraft-wide
power loads in response to aging engines.
[0105] Integrated Aircraft Control:
[0106] Full integration of engine control with whole aircraft
control requires digital controllers 110 for both the engines 200
and the aircraft 100 as a whole. In many previous legacy aircraft,
particularly turbo-prop jet engines (typically used in business and
aviation planes), the controlling mechanism for engine acceleration
and propeller pitch is a hydro-mechanism (liquid line or pneumatic
based control). Full integration of the control of the aircraft 100
and the engines 200 requires that the engines 200 have a Full
Authority Digital Engine Control (FADEC) 110. A FADEC 110 enables
digitally monitoring and trending of the engine 200 via the
controller 110. A FADEC 110 also enables detailed data and control
exchange with other aircraft avionics and controllers 105.
[0107] FIG. 4B is a flowchart of an exemplary method 420 for
managing aircraft power consumption to maintain safe limits and
prevent engine stalls from occurring. The method may be
implemented, for example and without limitation, via computer code
which is stored in a tangible medium on the aircraft 100 and
executed via one or more processors of the FADEC 110 and/or the
processors of other aircraft avionics and controllers 105.
[0108] The method beings with step 430, a monitoring step, which
may encompass two processes running substantially in parallel. In
step 430.1, the method monitors aircraft engine power consumption
(AEPC) 502 in real-time (that is, during flight), and/or values
which may serve as a proxy for AEPC 502. For example, one measure
of aircraft power consumption 502 may be the ratio of P3/P2. In
parallel, in step 430.2, the method monitors power consumption of
aircraft systems other than the engine (ASOTE) (400). While not
illustrated in the flowchart 420, additional, concurrent, ongoing
monitoring may entail measuring environmental and operational
conditions of and around the aircraft 100, including speed,
altitude, atmospheric conditions, and other data which pertain to
the loads the aircraft 100 is experiencing and is anticipated to
experience in upcoming minutes of the flight. Data for other
factors, such as anticipated or desired flight path and intended
flight speed, may be gathered as well.
[0109] The method 420 continues with step 435. In step 435, the
method evaluates whether the power demands on the engines 200 are
nearing, are at, or possibly are exceeding the maximum transient
operating line (MTOL) 530, which indicates the maximum power that
can be safely demanded from the engines 200 without risk of stall.
As discussed elsewhere in this document, the present system and
method may modify the maximum transient operating line 530 over
extended periods of time (weeks and/or months), generally in a
downward trend 595 as a result of routine engine wear and tear.
This means that the MTOL may be adjusted over weeks or months.
However, at a given time and for shorter stable time durations (for
example, for the duration of a single flight, and possibly over
multiple successive flights), the MTOL 530 is generally
pre-determined as a well-defined, stable threshold for the flight
duration(s).
[0110] If in step 435 the AEPC 502 is below the MTOL 530 (that is,
AEPC<MTOL), the method continues with the monitoring 430.
[0111] If in step 435 the MTOL 530 is above the AEPC 502 (that is,
AEPC>=MTOL), the method continues with steps 440, which
evaluates possible modifications to engine operations and/or
overall aircraft operations. The possible modifications are
intended to reduce power consumption by the engines 200, and/or to
modify engine operations to increase engine efficiency. The step
440 may evaluate, for example, whether current power demands can
best be addressed by changes to engine operations (as in step 445)
or changes to other aircraft power demands (steps 450/455). Step
440 may determine that some combination of changes to engine
operations (step 445) and changes to other aircraft power
consumption (steps 450/455) may be the optimal choice to reduce the
current aircraft engine power consumption 502.
[0112] In step 455, the present system and method implements
necessary changes to operations of the engines proper to reduce
power consumption of the engines 200 or to optimize the transient
performance of the engines 200. The details of such changes are
beyond the scope of this document.
[0113] In some embodiments, as shown in the flow chart 420, step
445 (modifying engine operations) may precede steps 450 and 455
(modifying other aircraft operations); however, that ordering is
exemplary only. In alternative embodiments, steps 450/455 may
precede steps 440, and in other embodiments step 445 and steps
450/455 may occur substantially in parallel and possibly in
coordination with each other. Further, step 440 may occur partly or
substantially in parallel with some aspects of steps 445, 450, and
455.
[0114] In step 450, the method 420 evaluates various possible
operations to the current activities of aircraft systems other than
the engines (ASOTE) 450. Since the ASOTE 450 generally require
electrical power, reductions in power consumption by the ASOTE 450
will result in reduced power demands on the aircrafts generators
400.1, and therefore reduced power demands on the engines 200. In
embodiments, the evaluations may entail consideration of both the
system-wide power reductions which may be obtained by altering or
reducing operations of various ASOTE 400; and also consideration of
various vehicle safety and passenger comfort priorities which may
be established (by system engineers or aircraft operators) for the
ASOTE 400. Possible changes to operations of ASOTE 400, which may
be evaluated, may include, for example and without limitation:
[0115] >>Modifying the air pressure operations of various air
bleed systems 400.2;
[0116] >>Limiting the operations of the foreign object door
(FOD) 400.3, since maximum operation of the FOD 400.3 reduces
available engine pressure and power;
[0117] >>Temporary reductions of power available to various
interior cabin passenger support systems 400.4;
[0118] >>Changes to current aircraft speed, ascent or descent
angle, or changes to anticipated course 400.5;
[0119] >>Temporary changes or limitations some aspects of
operations of aviation electronics, for example powering down
non-essential systems;
[0120] >>Drawing extra power or reserve power from aircraft
batteries (not shown in the figures).
[0121] In step 455, the method 420 implements or executes one or
more selected, temporary operational changes to the functions of
the ASOTE 455. It will be noted in this context that in some
practical applications, "temporary" operational changes may be
sustained at least through the duration of a flight (that is,
through the safe landing of the aircraft). In other cases,
temporary operational changes may only be in effect during part of
a flight.
[0122] Further Regarding Aircraft Systems Other than the Engines
(ASOTE) 400:
[0123] While the discussion above has distinguished between
engine/flight operational changes versus changes to the operations
of ASOTE 400, it will be noted that some alternations to flight
activity may potentially be viewed as falling at least partly into
both categories. For example, changes to flight velocity proper may
be obtained simply by reducing fuel supplied to the engines, which
may be viewed as part of modifying engine operations. However,
changes in course path and/or altitude (which may reduce the power
demands on the engines) may require changes both to airspeed and
also to wing flaps, rudder, and trim tabs, which are not part of
the engines. In general, and as understood in this document and in
the appended claims, changes in operations of the ASOTE 400 are to
be understood as changes to aircraft operations which reduce
demands on engine pressure, and/or optimize air compression in the
engines, without requiring direct control or alternation of the
operations of internal engine components, for example those shown
in exemplary engine 200 of FIG. 2 and exemplary turbine 300 of FIG.
3. As noted above, the FOD door 400.3 and air bleed systems 400.2
are considered to be parts of the aircraft systems other than the
engines 400.
[0124] Anticipatory Power Demand Processing:
[0125] In an alternative embodiment, step 435 above--the step of
determining that the demand for engine power 502 exceeds the
maximum transient operating line 530--may be an anticipatory step.
This means that the present system and method may evaluate a
current or recent trend in power consumption and engine performance
during the current flight; and may further evaluate anticipated
additional power demands in upcoming minutes of the flight (for
example, extra power required to climb over upcoming storm clouds).
Step 435 may then anticipate that while current engine power
consumption 502 is less than the MTOL 530, it is anticipated that
upcoming flight demands are likely to increase the engine power
consumption to be equal to our greater than the MTOL 530. The
present system and method can then implement actions, or recommend
actions to the pilot, to mitigate the expected change in power
demands. For example, the present system and method may recommend
to the pilots that they change their route to navigate around an
upcoming storm, rather than attempt to climb over the storm.
VI. ADJUSTMENTS TO ENGINE POWER CONSUMPTION THRESHOLDS OVER
TIME
[0126] In some embodiments, the present system and method makes
adjustments over time to threshold values for maximum jet engine
power consumption. These adjustments are made responsive to sensed
deterioration in aircraft engine performance over extended periods
of operation time (typically, variations over weeks, months, and
years). In some embodiments, the time frames for the threshold
value adjustments may be shorter, such as during the duration of a
single flight.
[0127] The threshold values are system-established limitations on
the amount of power which may be drawn from the engines without
incurring undue risk of engine stall. In some embodiments, the
present system and method adjusts the threshold values for a
maximum operating line 530 for the engines. Details are discussed
immediately below. For purposes of this document and the appended
claims, the following terms are deemed equivalent: "operating
line", "operating power", "operating load", and "operating HPx".
For example, the terms "maximum transient operating line 530",
"maximum transient operating power 530", "maximum transient
operating load 530" and "maximum transient operating HPx" are
deemed equivalent terms. For another example, the terms "allowed
operating line", "allowed operating power", "allowed operating
load" and "allowed operating HPx" are deemed equivalent terms.
VII. LEGACY STALL MAP (COMPRESSOR MAP)
[0128] FIG. 5A presents an exemplary legacy compressor map 500.1,
referred to equivalently herein as a stall map 500, which may be
used by the controller 110 to operate the engines 200, and which
can be stored in the memory 112. Persons skilled in the relevant
arts will appreciate that while the map 500 is illustrated in
graphical form, this is for convenience only; the compressor map
500 can be stored in the memory 112 in the form of data tables,
equations, and/or other digital representations known in the
art.
[0129] The compressor map 500.1 plots the compressor pressure ratio
502 (P3/P2) as a function of corrected compressor flow 76. Pressure
(typically measured in psi in the U.S.) determines the ability of
an air compressor 300 to perform a certain amount of work at any
given point in time. Flow 76 is indicative of the ability of an air
compressor to continuously perform a task (such as pressuring air
sufficiently to generate required thrust in a jet engine 200). In
the U.S., air flow 76 is typically measured in cubic feet per
minute (cfm) at a specific pressure, and can be viewed as "how
fast" the air flows past a given point in a compressor 300.
[0130] The compressor map 500 may include storage of data
representations for a surge line 515, a maximum operating line 530,
and a stall margin 570. As used herein, "surge line" 515 will refer
to the pressure ratio 502 wherein the airflow 76 through the engine
200 can break down or become unstable, causing an engine stall; and
"operating line" 530 will refer to the operating pressure ratio 502
which can safely be below that of the surge line. In addition,
"pressure ratio" 502 can refer to any of the following in
non-limiting examples: the ratio 502 (=P2/P3 or =P3/P2) of air
pressures taken at the inlet and outlet of the compressor section
300, the ratio 502 of air pressures taken at the inlet and outlet
of the compressors 304, 306 individually, or the ratio 502 of air
pressures taken at any two locations within the compressor section
300 as desired such as the exemplary P2, P3 measurement locations
shown in the figure.
[0131] >Stall Margin:
[0132] The compressor stall margin 570 can be defined as the
difference between the compressor stall line 515 and the operating
line 530; it will be understood that the percent difference between
the surge line 515 and operating line 530 can also be used to
define the stall margin 570. In legacy systems, the operating line
is pre-set as a fixed-value for a maximum safety factor stall
margin; the fixed-value is calculated before the engine 200 or
aircraft 100 are even put into service, and is calculated based on
the assumption of worst case operating conditions for the aircraft.
This leads to a corresponding large stall margin 570.1. All things
being equal, the greater the stall margin 570.1, the less
efficiently the engine 200 is operating.
[0133] >Uninstalled (Factory Condition) Operating Line 507:
[0134] The uninstalled operating line 507 reflects the P3/P2 ratio
for the engine 200, and therefore the power generated by the engine
200, if the engine is brand new, not installed as part of the
aircraft 100, and was run for example as part of ground-based
testing, that is, with no load on the engine 200. This can also be
understood as the horsepower generated by the engine 200 under the
same (ground-based, testing-only) conditions.
[0135] >Customer Bleed 508:
[0136] This line represents compressor flow exhausted (bled) from a
port on the engine compressor 102 during normal operations by the
aircraft when in routine flight, including but not limited to bleed
flow drawn by the aircraft systems other than the engine (ASOTE)
400.
[0137] >HPx New Engine (or Simply HPx) 510.1:
[0138] This line represents the power that is withdrawn from the
engines 200 during normal operations by the aircraft when in
routine flight, including but not limited to electrical power drawn
by the aircraft systems other than the engine (ASOTE) 400.
[0139] Acceleration and the Transient Operating Line 590.1:
[0140] In the course of normal operations, and even more so under
stressed operating conditions, an aircraft 100 will sometimes need
to accelerate. This places increased power demands on the aircraft
100. An exemplary transient operating line 590.1 represents a power
consumption by the engines 200 during acceleration. The word
"transient", as employed here in the phrase "transient operating
line", is indicative of the fact that the increased power demands
on the engine 200 and compressors 300 are typically expected to be
relatively transient in time or of brief duration, present only
during aircraft acceleration or other flight stresses which
increase power demands on the aircraft 100. At other times during a
flight, the power demands on the engines 200 (as reflected by a
real-time operating line) are expected to be at or near the power
demands reflected by the steady-state operating lines 505, as
indicated by routine operating line 506 for routine, level, and/or
unaccelerated periods of a flight.
[0141] It will be evident from FIG. 5A that during acceleration the
power consumption increases significantly compared to the power
consumption in routine, steady state flight 510.1. Further, persons
skilled in the art will appreciate that as an engine 200 ages and
degrades in the normal course of use over time, and for a given
level of operational demands on the aircraft 100, the transient
operating line 590.1 will have higher values.
[0142] Optimally, and while maintaining aircraft safety (by not
risking an engine stall) the power demands on the aircraft 100
could be made high enough to raise transient operating line 590.1
to just below the compressor stall line 515. In legacy aircraft
however, and as noted above, a maximum transient operating line
530.1 is established in anticipation of engine aging. The maximum
transient operating line 530.1 is established at a fix value at the
time the aircraft engine 200 is built or installed, resulting in a
large stall margin 530.1 from the outset of engine usage. As a
further result, and in the course of normal flight, there is a
large amount of underuse of available engine power during the early
and midlife periods of the engine. This is shown in the figure as
an engine power utilization gap 580.
VIII. ADJUSTABLE STALL MAP (OPTIMIZED COMPRESSOR MAP)
[0143] It is a goal of the present system and method to optimize
the stall margin 570 and the engine power utilization gap 580 over
the course of engine life. This results in more efficient
utilization of the engines 200, with concomitant resource and cost
savings, while still maintaining safe (stall-free) flight
operations.
[0144] FIG. 5B presents an exemplary compressor map 500.2, referred
to equivalently herein as a stall map 500.2, which may be used by
the controller 110 to operate the engines 200 according to the
present system and method. Many of the elements of stall map 500.2
are the same or similar to those of the stall map 500.1 of FIG. 5A
above, and the details will not be repeated here.
[0145] The compressor map 500.2 may be understood as implicitly
representing multiple maps 500, the multiple stall maps being
indicative of engine performance at different times (possibly weeks
apart, months apart, or even years apart) in the life cycle of the
engines 200.
[0146] Initial Maximum Operating Line:
[0147] The Initial Maximum Transient Operating Line 530.2
represents an allowed power demand on the engine compressors 300
when the engines 200 and/or aircraft 100 are first operational, and
operating at factory-fresh or near factory-fresh efficiency.
Because the engines 200 and/or aircraft 100 are new, this is viewed
as a safe operating line 530 for the engines 200, compressors 300
and aircraft 100 is the initial or early stages of use. As may be
seen in the figure, and according to embodiments of the present
system and method, the initial maximum transient operating line
530.2 is set to be very close to the compressor stall line (515).
To maintain a safety margin against stalls, the initial maximum
transient operating line 530.2 may be set at slightly below the
values of the compressor stall line 515. For example, in various
exemplary embodiments, the P3/P2 values of the initial maximum
transient operating line 530.2 may be set to 99%, 98%, or 95% of
the values on the stall line 515 (for corresponding W2R values
along the X-axis). In other embodiments, other percentages,
close-to-but-not-at 100%, may be envisioned as well.
[0148] Time-Adjusted Maximum Operating Line:
[0149] Over time, the present system and method detects, via
sensors 310, declines in engine performance due to routine engine
wear-and-tear. In response to sensed declines in engine
performance, such as reduced P3/P2 ratios or more generally reduced
power output over time (the time frames typically being weeks,
months, or years, though shorter times may be envisioned), the
controller 110 calculates an adjusted maximum transient operating
line 530.3.
[0150] With successive such adjustments and calculations reflecting
an extended use-life of the engines 200, the adjusted maximum
transient operating line 530.3 will typically show a downwards
trend 595 (typically approaching the fixed maximum transient
operating line 530.1 (see FIG. 5A)). However, because particular
engine usage and aircraft usage varies from aircraft to aircraft
and from engine to engine, some engines 200 will maintain over time
a better than expected power output. As such, even as a particular
engines 200 ages, it's adjusted maximum transient operating line
530.3 may still permit greater power demands on the engine 200 than
the fixed maximum transient operating line 530.1 of legacy
systems.
[0151] Engine or Compressor End-of-Life:
[0152] In some embodiments of the present system and method, a
fixed maximum transient operating line 530.1, based on historical
and/or calculated worst-case engine-use scenarios, may still be
stored in memory 112. (Historical lines 530.1 would be based on the
history of decline of other engines of the same design or similar
design, as measured in actual operation over time.)
[0153] Over an extended period of engine usage, as the adjusted
maximum transient operating line 530.3 approaches the fixed maximum
transient operating line 530.1 (as indicated by some threshold
difference level defined by system engineers or by aircraft
engineers), the controller 110 of the present system and method may
issue a warning or alert. The warning or alert may indicate a need
for urgent engine maintenance, or more generally may indicate that
the engine 200 and/or compressor 300 may be nearing the end of
useful operational life.
[0154] Efficiency Benefits:
[0155] As may be seen in FIG. 5B, during the early operational life
of the engine 200--when the maximum transient operating line is
very close to or substantially close to the compressor stall line
515--the stall margin 570.2 is very small compared to, for example,
the fixed and larger stall margin 570.1 of legacy systems. This
smaller stall margin can be established within safety limits, and
as a result a transient operating line 590.2 may safely place great
power demands on the engine 200, as compared for example to the
transient operating line 590.1 of legacy systems. This results, in
both early- and mid-life stages of the engine 200, in more
efficient use of the engines, more efficient use of fuel, and also
in the safe use of greater reserve power in emergency
situations.
[0156] Adaptable Operating Line and Significant Degradation
Time:
[0157] In some embodiments of the present system and method, and as
noted above, the controller 110 employs data from sensors 310 to
identify declines in engine power and performance due to routine
engine wear-and-tear, and particularly due to degradation over time
of the compressor 300. In response to sensed declines in engine
performance, such as reduced P3/P2 ratios or more generally reduced
power output over weeks, months, and/or years, the controller 110
calculates an adjusted maximum transient operating line 530.3.
Based on general historical experience with engines, experience
with engines of a same or similar design, computer simulations of
engine performance, and factory testing, design engineers may
determine for an engine 200 a significant engine degradation time
(SEDT), or equivalently "significant degradation time (SDT)" or
"engine degradation time (EDT)".
[0158] The significant degradation time (SDT) may be a time period,
such as a minimum time period or an average time period or similar,
when the engine 200 may be expected to display sufficient physical
or structural degradation so as to meaningfully impact (and
generally reduce) engine thrust and power. As indicated above, such
time periods may typically be on the order of weeks or months, but
in some cases longer or shorter periods. Generally, system design
engineers or aviation engineers will define the criteria for a
meaningful rebalance in engine power extraction for electrical
load, which may in some embodiments be a 0.5% reduction in engine
power extraction, a 1% reduction in engine power, or other
reductions such as 2%, 5%, or 10%. The change may also entail
maintaining the same level of engine power extraction for
electrical load while increasing compressor bleed extraction for
aircraft use to rebalance overall load on the engine compressor. In
embodiments, the present system and method will be configured to
adapt or adjust an engine's maximum transient operating line 530.3
at time intervals no longer than the SDT. In alternative
embodiments, the present system and method may recalculate or
adjust an engine's maximum transient operating line 530.3 when a
stored data log of engine performance indicates a specified
reduction over time in engine performance (for example, a 0.5%
reduction in engine power or compression ratio, a 1% reduction in
engine power or compression ratio, or other reductions such as 2%,
5%, or 10%). Typically, this will result in the maximum transient
operating line 530.3 being recalculated and updated at least at
time intervals of the significant degradation time (SDT), and
possibly more often.
[0159] FIG. 5C presents an exemplary compressor map 500.3 which
illustrates some aspects of an exemplary method to establish an
adjusted maximum transient operating line 530.3.
[0160] As above, the compressor map 500.3 plots the compressor
pressure ratio 502 (P3/P2) as a function of corrected compressor
flow 76. The compressor flow 76 is a corrected flow, for which a
suitable formula is identified in the course of development of the
engine 200 and compressor 300. In one embodiment, which is
exemplary only and non-limiting, the corrected compressor mass flow
ratio is calculated and plotted as:
CCF=(W2*SQRT(T2/518.67))/(P2/14.67)
[0161] where:
[0162] W2=air flow rate at the point where P2 is measured;
[0163] T2=the temperature at the point where P2 is measured;
and
[0164] P2=the air pressure.
[0165] An exemplary, non-limiting method for obtaining a maximum
transient operating line 530.3 is presented here. It will be noted
that this method may be repeated periodically, for example at
designated intervals of weeks or months, to adjust the maximum
transient operating line 530 for the engine. The method may also be
employed at other times, for example when pilots, other aircraft
specialists, or software have identified periods of time when the
aircraft 100 and its engines/compressors 200/300 have been subject
to greater than usual stresses. The steps of the method entail:
[0166] (1) Identifying lines of constant core speed, as corrected
for temperature and pressure.
[0167] (2) Identifying, in the stall map, a region or "island" of
constant efficiency 597. For a given period or life stage (LS)
during the entire operational life of the engine 200 and compressor
300, an island of constant efficiency represents a phase space for
compressor pressure ratio vs. compressor air flow where the
engine/compress 200/300 exhibits constant efficiency. This
determination is based on analysis of engine pressure, air flow,
and power delivered by the engine 200, and a detailed discussion is
beyond the scope of this document.
[0168] An island of constant efficiency 597 is expected to remain
reasonably stable in size over shorter periods of engine usage and
life (on the order of weeks or months), but will decline in area
over the extended length of the useful operational lifetime of the
engine 200 as represented in the figure by engine Life Stages 1, 2,
. . . , n (LS1, LS2, . . . , LSn).
[0169] (3) A transient operating line 530 may, during a given life
stage LS(k), be constructed from lines which are orthogonal to
selected lines of constant speed and tangent to an upper boundary
of the constant efficiency operating lines 597 as shown in the
figure.
[0170] During the lifetime of the engine 200, data from the sensors
310 is used as a guide to update the transient operating line 530
for an engine. In this way, when an engine 200 is new the aircraft
100 can take full and most efficient advantage of the high engine
efficiency; as the engine ages and its efficiency is reduced in the
normal course of operations, power demands on the engines can be
suitably rebalanced during progressive engine Life Stages LS(k) to
ensure engine operations and thrust remain within safe limits (so
as to avoid engine stalls).
IX. FURTHER EMBODIMENTS
[0171] The present system and method leverages prognostic health
management (PHMS) 119 tools to assess deterioration levels on
turbine engine 200s such as the GE Catalyst engine. In one
embodiment, once a predefined level of engine deterioration is
surpassed, the FADEC 110 will send a CAS message to alert the pilot
regarding potential limitations on HPx 510.1, ECS bleed, ability to
deploy FOD doors 400.3, or minimum aircraft speed at the altitude
ceiling to avoid risk of stall.
[0172] More specifically, and in an embodiment: The present system
and method utilizes PHMS 119 methods and tools such as (for example
and without limitation) comparing torque sensor output to ITT, NGR,
P1 vs P3, as well as other parameters with License Required
restrictions, to regularly monitor and analyze engine performance
data, and further to create a deterioration trend. Embedded models
or tracking filters maybe utilized to assess deterioration.
[0173] Analysis and trending could be on board the FADEC 110 or off
board with updates back to the FADEC 110 at regular intervals. At
some maximum, predefined threshold of deterioration, corresponding
to a certain level of surge margin consumption, the pilot will be
notified of potential actions to reduce surge risk based on
operational procedures.
[0174] Such action procedures could include, for example and
without limitation: (i) do not close ECS bleed from PL25 while
utilizing max HPx 510.1; (ii) turn off/reduce HPx 510.1 prior to
turning off ECS bleed; (iii) do not deploy FOD doors above a
predetermined altitude ceiling while flying slower than a preset
threshold; (iv) reduce cabin electrical consumption (for example,
by limiting comfort or support services for passengers); (v)
reducing other system-wide power consumption; or (vi) any
combination of the actions above.
[0175] In some embodiments of the present system and method, if
within a predefined window the pilot does not take necessary
action, the FADEC 110 may impose automated limits on HPx 510.1
based on altitude and speed thresholds through direct communication
with Starter Generator Power Control Unit (PCU) and notify the
pilot of the pre-imposed limitations. These actions should provide
more flexibility and optimization regarding the impact of
extractions on Operability margin. They will allow for higher
electrical load HPx 510.1 from the core while the engine is new (as
compared to reduced electrical loads HPx 510.1 as the engines
deteriorate with time).
[0176] In some embodiments of the present system and method, the
step of pilot notification and pilot response time for action may
be omitted. In such embodiments, the FADEC 110 may automatically
impose designated various automated limits on HPx 510.1, and may or
may not supply pilot notification as well.
[0177] Legacy technology sets an initial threshold (that is,
fresh-off-the-engine production line) for engine power consumption,
and therefore for maximum allowed stress on the engine compressors,
based on the anticipated future wear, tear, and degradation of
engine structure (in the ordinary course of use) over time. Such
estimates of future decline in engine function necessarily take
into account worst-case scenarios for the future use of the engines
in the ordinary course of flight. The legacy approach therefore
imposes significant limits on engine power consumption from the
first day of use, when in fact new engines and moderately aged
engines may have significant power to spare within the bounds of
safety. Legacy engine monitoring systems also address declines or
deficiencies in engine thrust over time by modifying internal
operations of the engines.
[0178] The present system and method for aircraft optimization
based on engine performance leverages PHMS 119 tools, as described
throughout this document, to assess engine conditions over time,
and further to optimize available power extractions to the unique,
particular condition of an aircraft, the aircraft engines, and the
aircraft engine compressors for each unique transport asset and its
unique missions. The present system and method therefor enables
more tailored engine capability, particularly in the early- and
midlife-stages of an aircraft's deployment, instead of limitations
imposed by the long-term, worst possible statistical case in the
fleet. The present system and method further provides for adjusting
a variety of aircraft systems and operations, including those which
are not at all engine operations or which may be viewed as systems
only peripheral to the engines, in order to reduce short-term
engine power demands while in flight.
CONCLUSION
[0179] Alternative embodiments, examples, and modifications which
would still be encompassed by the disclosure may be made by those
skilled in the art, particularly in light of the foregoing
teachings. Further, it should be understood that the terminology
used to describe the disclosure is intended to be in the nature of
words of description rather than of limitation.
[0180] Those skilled in the art will also appreciate that various
adaptations and modifications of the preferred and alternative
embodiments described above can be configured without departing
from the scope and spirit of the disclosure. Therefore, it is to be
understood that, within the scope of the appended claims, the
disclosure may be practiced other than as specifically described
herein.
* * * * *