U.S. patent application number 16/018754 was filed with the patent office on 2018-10-18 for low noise compressor and turbine for geared turbofan engine.
The applicant listed for this patent is MTU Aero Engines AG, United Technologies Corporation. Invention is credited to Detlef Korte, Bruce L. Morin, David A. Topol.
Application Number | 20180299477 16/018754 |
Document ID | / |
Family ID | 55961262 |
Filed Date | 2018-10-18 |
United States Patent
Application |
20180299477 |
Kind Code |
A1 |
Topol; David A. ; et
al. |
October 18, 2018 |
LOW NOISE COMPRESSOR AND TURBINE FOR GEARED TURBOFAN ENGINE
Abstract
A gas turbine engine has a fan section including a fan, a
compressor section including a low pressure compressor and a high
pressure compressor, a turbine section including a low pressure
turbine and a high pressure turbine, and a gear reduction. The low
pressure compressor and the low pressure turbine have a number of
blades in each of at least one of a plurality of blade rows. The
blades are rotatable at least some of the time at a rotational
speed in operation. The number of blades in at least one row and
the rotational speed are such that the following formula holds true
for at least one row of the compressor rotor turbine:
5500.ltoreq.(number of blades.times.rotational
speed)/60.ltoreq.10000, the rotational speed being in revolutions
per minute.
Inventors: |
Topol; David A.; (West
Hartford, CT) ; Morin; Bruce L.; (Longmeadow, CT)
; Korte; Detlef; (Karlsfeld, DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation
MTU Aero Engines AG |
Farmington
Muenchen |
CT |
US
DE |
|
|
Family ID: |
55961262 |
Appl. No.: |
16/018754 |
Filed: |
June 26, 2018 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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15662528 |
Jul 28, 2017 |
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16018754 |
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15270027 |
Sep 20, 2016 |
9733266 |
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15662528 |
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15014363 |
Feb 3, 2016 |
9650965 |
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15270027 |
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14967478 |
Dec 14, 2015 |
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15014363 |
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14591975 |
Jan 8, 2015 |
9624834 |
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14967478 |
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14144710 |
Dec 31, 2013 |
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14591975 |
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14016436 |
Sep 3, 2013 |
8714913 |
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14144710 |
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13630276 |
Sep 28, 2012 |
8632301 |
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14016436 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y10T 29/49236 20150115;
F05D 2200/14 20130101; F02C 7/32 20130101; F05D 2220/36 20130101;
F02K 3/06 20130101; F02C 3/10 20130101; F05D 2260/40311 20130101;
F02C 7/24 20130101; Y10T 29/4932 20150115; Y02T 50/60 20130101;
Y02T 50/673 20130101; F01D 5/12 20130101; F05D 2260/96 20130101;
F02C 3/04 20130101; F01D 5/02 20130101; G01P 3/48 20130101; F02C
7/36 20130101; F05D 2200/13 20130101; F05D 2220/32 20130101 |
International
Class: |
G01P 3/48 20060101
G01P003/48; F02C 7/32 20060101 F02C007/32; F02K 3/06 20060101
F02K003/06; F02C 7/24 20060101 F02C007/24; F01D 5/12 20060101
F01D005/12; F02C 3/04 20060101 F02C003/04; F02C 7/36 20060101
F02C007/36; F01D 5/02 20060101 F01D005/02; F02C 3/10 20060101
F02C003/10 |
Claims
1. A gas turbine engine comprising: a fan section including a fan,
and a low fan pressure ratio of less than 1.45, wherein the low fan
pressure ratio is measured across a fan blade alone; a compressor
section including a low pressure compressor and a high pressure
compressor; wherein the fan delivers air into a bypass duct defined
within a nacelle, and a portion of air into the compressor section,
with a bypass ratio defined as the volume of air delivered into the
bypass duct compared to the volume of air delivered into the
compressor section, and the bypass ratio being greater than 10; a
turbine section including a low pressure turbine and a high
pressure turbine; wherein the low pressure turbine includes an
inlet, an outlet and a pressure ratio of greater than 5, the
pressure ratio being pressure measured prior to the inlet as
related to pressure at the outlet prior to an exhaust nozzle; a
gear reduction including an epicyclic gear train, wherein the gear
reduction effects a reduction in the speed of the fan relative to a
speed of the low pressure turbine, the epicyclic gear train having
a gear reduction ratio of greater than 2.5:1; wherein each of the
low pressure compressor and the low pressure turbine includes a
number of blades in each of a plurality of blade rows, the number
of blades rotatable at least some of the time at a rotational speed
in operation, and the number of blades and the rotational speed
being such that the following formula holds true for at least one
of the blade rows of the low pressure compressor: 5500
Hz.ltoreq.(number of blades.times.rotational speed)/60.ltoreq.10000
Hz, the rotational speed being an approach speed in revolutions per
minute, taken at an approach certification point as defined in Part
36 of the Federal Airworthiness Regulations; and the following
formula holds true for at least one of the blade rows of the low
pressure turbine: (number of blades.times.rotational
speed)/60<10000 Hz, the rotational speed being an approach speed
in revolutions per minute, taken at an approach certification point
as defined in Part 36 of the Federal Airworthiness Regulations; and
wherein the engine is rated to produce 15,000 pounds of thrust or
more.
2. The gas turbine engine as set forth in claim 1, wherein the
formula results in a number less than or equal to 10000 for at
least a plurality of the blade rows of the low pressure
compressor.
3. The gas turbine engine as set forth in claim 2, wherein the
formula results in a number greater than or equal to 5500 for at
least half of the blade rows of the low pressure compressor.
4. The gas turbine engine as set forth in claim 3, wherein the
formula results in a number greater than or equal to 5500 for at
least one of the blade rows of the low pressure turbine.
5. The gas turbine engine as set forth in claim 4, wherein the
formula results in a number greater than or equal to 6000 for at
least one of the blade rows of the low pressure compressor.
6. The gas turbine engine as set forth in claim 5, wherein the
formula results in a number less than or equal to 10000 for at
least a plurality of the blade rows of the low pressure
turbine.
7. The gas turbine engine as set forth in claim 6, wherein the
formula results in a number greater than or equal to 6000 for at
least half of the blade rows of the low pressure compressor.
8. The gas turbine engine as set forth in claim 7, wherein the
formula results in a number greater than or equal to 6000 for at
least half of the blade rows of the low pressure turbine.
9. The gas turbine engine as set forth in claim 8, wherein the
formula results in a number greater than or equal to 6000 for three
of the blade rows of the low pressure turbine.
10. The gas turbine engine as set forth in claim 9, wherein the
formula results in a number less than or equal to 10000 for at
least a majority of the blade rows of the low pressure turbine.
11. The gas turbine engine as set forth in claim 10, wherein the
formula results in a number less than or equal to 10000 for at
least a majority of the blade rows of the low pressure
compressor.
12. The gas turbine engine as set forth in claim 11, further
comprising: a core flowpath and a mid-turbine frame arranged
between the low pressure turbine and the high pressure turbine, the
mid-turbine frame having airfoils positioned in the core flowpath,
and the mid-turbine frame supporting at least one bearing system;
wherein the low pressure compressor includes three stages, the low
pressure turbine includes a greater number of stages than the low
pressure compressor, the high pressure turbine includes two stages,
and the high pressure compressor includes a greater number of
stages than the low pressure turbine; and wherein the fan has a low
corrected fan tip speed of less than 1150 ft/second.
13. The gas turbine engine as set forth in claim 12, wherein the
gear reduction is a planetary gear system.
14. A gas turbine engine comprising: a fan section including a fan,
and a low fan pressure ratio of less than 1.45, wherein the low fan
pressure ratio is measured across a fan blade alone; a compressor
section including a low pressure compressor and a high pressure
compressor; wherein the fan delivers air into a bypass duct defined
within a nacelle, and a portion of air into the compressor section,
with a bypass ratio defined as the volume of air delivered into the
bypass duct compared to the volume of air delivered into the
compressor section, and the bypass ratio being greater than 10; a
turbine section including a low pressure turbine and a high
pressure turbine; wherein the low pressure turbine includes an
inlet, an outlet and a pressure ratio of greater than 5, the
pressure ratio being pressure measured prior to the inlet as
related to pressure at the outlet prior to an exhaust nozzle; a
gear reduction including an epicyclic gear train, wherein the gear
reduction is positioned intermediate the low pressure compressor
and a shaft driven by the low pressure turbine such that a fan
rotor of the fan section and the low pressure compressor are
rotatable at a common speed in operation, the epicyclic gear train
having a gear reduction ratio of greater than 2.5:1; wherein each
of the low pressure compressor and the low pressure turbine
includes a number of blades in each of a plurality of blade rows,
the number of blades rotatable at least some of the time at a
rotational speed in operation, and the number of blades and the
rotational speed being such that the following formula holds true
for at least one of the blade rows of the low pressure compressor:
5500 Hz.ltoreq.(number of blades.times.rotational
speed)/60.ltoreq.10000 Hz, the rotational speed being an approach
speed in revolutions per minute, taken at an approach certification
point as defined in Part 36 of the Federal Airworthiness
Regulations; and the following formula holds true for at least one
of the blade rows of the low pressure turbine: (number of
blades.times.rotational speed)/60.ltoreq.10000 Hz, the rotational
speed being an approach speed in revolutions per minute, taken at
an approach certification point as defined in Part 36 of the
Federal Airworthiness Regulations; and wherein the engine is rated
to produce 15,000 pounds of thrust or more.
15. The gas turbine engine as set forth in claim 14, wherein the
formula results in a number less than or equal to 10000 for at
least a plurality of the blade rows of the low pressure
compressor.
16. The gas turbine engine as set forth in claim 15, wherein the
formula results in a number greater than or equal to 5500 for at
least one of the blade rows of the low pressure turbine.
17. The gas turbine engine as set forth in claim 16, wherein the
formula results in a number greater than or equal to 6000 for at
least one of the blade rows of the low pressure compressor and for
at least one of the blade rows of the low pressure turbine.
18. The gas turbine engine as set forth in claim 17, wherein the
formula results in a number greater than or equal to 6000 for at
least half of the blade rows of the low pressure compressor.
19. The gas turbine engine as set forth in claim 18, wherein the
formula results in a number greater than or equal to 6000 for at
least half of the blade rows of the low pressure turbine.
20. The gas turbine engine as set forth in claim 19, wherein the
formula results in a number less than or equal to 10000 for at
least a plurality of the blade rows of the low pressure
turbine.
21. The gas turbine engine as set forth in claim 20, wherein the
formula results in a number less than or equal to 7000 for only a
majority of the blade rows of the low pressure compressor and for
only a majority of the blade rows of the low pressure turbine.
22. The gas turbine engine as set forth in claim 20, wherein the
formula results in a number less than or equal to 10000 for at
least a majority of the blade rows of the low pressure compressor
and for at least a majority of the blade rows of the low pressure
turbine.
23. The gas turbine engine as set forth in claim 22, wherein the
formula results in a number less than or equal to 7000 for at least
a majority of the blade rows of the low pressure compressor and for
at least a majority of the blade rows of the low pressure turbine,
and the formula results in a number greater than or equal to 6000
for at least a majority of the blade rows of the low pressure
compressor and for at least a majority of the blade rows of the low
pressure turbine.
24. The gas turbine engine as set forth in claim 23, further
comprising: a core flowpath and a mid-turbine frame arranged
between the low pressure turbine and the high pressure turbine, the
mid-turbine frame having airfoils positioned in the core flowpath,
and the mid-turbine frame supporting at least one bearing system;
wherein the low pressure compressor includes three stages, the low
pressure turbine includes five stages, the high pressure turbine
includes two stages, and the high pressure compressor includes
eight stages; and wherein the fan has a low corrected fan tip speed
of less than 1150 ft/second.
25. A gas turbine engine comprising: a fan section including a fan,
and a low fan pressure ratio of less than 1.45, wherein the low fan
pressure ratio is measured across a fan blade alone; a compressor
section including a low pressure compressor, an intermediate
pressure compressor and a high pressure compressor; wherein the fan
delivers air into a bypass duct defined within a nacelle, and a
portion of air into the compressor section, with a bypass ratio
defined as the volume of air delivered into the bypass duct
compared to the volume of air delivered into the compressor
section, and the bypass ratio being greater than 10; a turbine
section including a low pressure turbine, an intermediate pressure
turbine and a high pressure turbine, the intermediate pressure
turbine driving the intermediate pressure compressor, and the high
pressure turbine driving the high pressure compressor; wherein the
low pressure turbine includes an inlet, an outlet and a pressure
ratio of greater than 5, the pressure ratio being pressure measured
prior to the inlet as related to pressure at the outlet prior to an
exhaust nozzle; a gear reduction including an epicyclic gear train,
wherein the gear reduction effects a reduction in the speed of the
fan relative to a speed of the low pressure turbine, the epicyclic
gear train having a gear reduction ratio of greater than 2.5:1;
wherein each of the low pressure compressor and the low pressure
turbine includes a number of blades in each of a plurality of blade
rows, the number of blades rotatable at least some of the time at a
rotational speed in operation, and the number of blades and the
rotational speed being such that the following formula holds true
for at least one of the blade rows of the low pressure compressor:
6000 Hz.ltoreq.(number of blades.times.rotational
speed)/60.ltoreq.10000 Hz, the rotational speed being an approach
speed in revolutions per minute, taken at an approach certification
point as defined in Part 36 of the Federal Airworthiness
Regulations; and the following formula holds true for at least one
of the blade rows of the low pressure turbine: (number of
blades.times.rotational speed)/60.ltoreq.10000 Hz, the rotational
speed being an approach speed in revolutions per minute, taken at
an approach certification point as defined in Part 36 of the
Federal Airworthiness Regulations; and wherein the engine is rated
to produce 15,000 pounds of thrust or more.
26. The gas turbine engine as set forth in claim 25, wherein the
formula results in a number less than or equal to 10000 for at
least a plurality of the blade rows of the low pressure compressor
and for at least a plurality of the blade rows of the low pressure
turbine.
27. The gas turbine engine as set forth in claim 26, wherein the
formula results in a number greater than or equal to 5500 for at
least half of the blade rows of the low pressure compressor and for
at least half of the blade rows of the low pressure turbine.
28. The gas turbine engine as set forth in claim 27, wherein the
formula results in a number less than or equal to 7000 for at least
a majority of the blade rows of the low pressure compressor and for
at least a majority of the blade rows of the low pressure
turbine.
29. The gas turbine engine as set forth in claim 28, wherein the
formula results in a number greater than or equal to 6000 for at
least a majority of the blade rows of the low pressure compressor
and for at least a majority of the blade rows of the low pressure
turbine.
30. The gas turbine engine as set forth in claim 29, further
comprising: a core flowpath and a mid-turbine frame arranged
between the low pressure turbine and the high pressure turbine, the
mid-turbine frame having airfoils positioned in the core flowpath,
and the mid-turbine frame supporting at least one bearing system;
wherein the gear reduction is a planetary gear system, and the gear
reduction is positioned intermediate the low pressure compressor
and a shaft driven by the low pressure turbine such that a fan
rotor of the fan section and the low pressure compressor are
rotatable at a common speed in operation; wherein the low pressure
compressor includes three stages, the low pressure turbine includes
five stages, the high pressure turbine includes two stages, and the
high pressure compressor includes eight stages; and wherein the fan
has a low corrected fan tip speed of less than 1150 ft/second.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 15/662,528, filed Jul. 28, 2017, which is a
continuation of U.S. patent application Ser. No. 15/270,027, filed
Sep. 20, 2016, which is a continuation of U.S. patent application
Ser. No. 15/014,363, filed Feb. 3, 2016, which is a continuation of
U.S. patent application Ser. No. 14/967,478, filed Dec. 14, 2015,
which is a continuation-in-part of U.S. patent application Ser. No.
14/591,975, filed Jan. 8, 2015, which is a continuation-in-part of
U.S. patent application Ser. No. 14/144,710, filed Dec. 31, 2013,
which is a continuation of U.S. patent application Ser. No.
14/016,436, filed Sep. 3, 2013, now U.S. Pat. No. 8,714,913, issued
May 6, 2014, which is a continuation of U.S. patent application
Ser. No. 13/630,276, filed Sep. 28, 2012, now U.S. Pat. No.
8,632,301, issued Jan. 21, 2014.
BACKGROUND
[0002] This application relates to the design of a gas turbine
engine rotor which can be operated to produce noise that is less
sensitive to human hearing.
[0003] Gas turbine engines are known, and typically include a fan
delivering air into a compressor. The air is compressed in the
compressor and delivered downstream into a combustor section where
it was mixed with fuel and ignited. Products of this combustion
pass downstream over turbine rotors, driving the turbine rotors to
rotate.
[0004] Typically, there is a high pressure turbine rotor, and a low
pressure turbine rotor. Each of the turbine rotors include a number
of rows of turbine blades which rotate with the rotor. Interspersed
between the rows of turbine blades are vanes.
[0005] The high pressure turbine rotor has typically driven a high
pressure compressor rotor, and the low pressure turbine rotor has
typically driven a low pressure compressor rotor. Each of the
compressor rotors also include a number of compressor blades which
rotate with the rotors. There are also vanes interspersed between
the rows of compressor blades.
[0006] The low pressure turbine or compressor can be a significant
noise source, as noise is produced by fluid dynamic interaction
between the blade rows and the vane rows. These interactions
produce tones at a blade passage frequency of each of the low
pressure turbine rotors, the low pressure compressor rotors, and
their harmonics.
[0007] The noise can often be in a frequency range that is very
sensitive to humans. To mitigate this problem, in the past, a
vane-to-blade ratio has been controlled to be above a certain
number. As an example, a vane-to-blade ratio may be selected to be
1.5 or greater, to prevent a fundamental blade passage tone from
propagating to the far field. This is known as "cut-off."
[0008] However, acoustically cut-off designs may come at the
expense of increased weight and reduced aerodynamic efficiency.
Stated another way, by limiting the designer to a particular vane
to blade ratio, the designer may be restricted from selecting such
a ratio based upon other characteristics of the intended
engine.
[0009] Historically, the low pressure turbine has driven both a low
pressure compressor section and a fan section. More recently, a
gear reduction has been provided such that the fan and low pressure
compressor can be driven at distinct speeds.
SUMMARY
[0010] In a featured embodiment, a gas turbine engine has a fan, a
compressor section having a low pressure portion and a high
pressure portion, a combustor section, and a turbine having a first
turbine rotor. The first turbine rotor drives the fan. A gear
reduction effects a reduction in the speed of the fan relative to a
speed of the first turbine rotor. Each of the compressor rotor and
the first turbine rotor includes a number of blades in each of a
plurality of rows. The blades operate at least some of the time at
a rotational speed. The number of blades and the rotational speed
are such that the following formula holds true for at least one of
the blade rows of the first turbine rotor and/or the compressor
rotor: (number of blades.times.rotational speed)/60.gtoreq.5500.
The rotational speed is an approach speed in revolutions per
minute.
[0011] These and other features of this application will be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] FIG. 1 shows a gas turbine engine.
[0013] FIG. 2 shows another embodiment.
[0014] FIG. 3 shows yet another embodiment.
DETAILED DESCRIPTION
[0015] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown),
or an intermediate spool, among other systems or features. The fan
section 22 drives air along a bypass flowpath B in a bypass duct
defined within a nacelle 15, while the compressor section 24 drives
air along a core flowpath C for compression and communication into
the combustor section 26 then expansion through the turbine section
28. Although depicted as a turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with turbofans as
the teachings may be applied to other types of turbine engines
including three-spool architectures.
[0016] The engine 20 generally includes a low speed spool 30 and a
high speed spool 32 mounted for rotation about an engine central
longitudinal axis A relative to an engine static structure 36 via
several bearing systems 38. It should be understood that various
bearing systems 38 at various locations may alternatively or
additionally be provided.
[0017] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 is connected to the fan 42
through a geared architecture 48 to drive the fan 42 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 50 that interconnects a high pressure compressor 52
and high pressure turbine 54. A combustor 56 is arranged between
the high pressure compressor 52 and the high pressure turbine 54. A
mid-turbine frame 57 of the engine static structure 36 is arranged
generally between the high pressure turbine 54 and the low pressure
turbine 46. The mid-turbine frame 57 further supports bearing
systems 38 in the turbine section 28. The inner shaft 40 and the
outer shaft 50 are concentric and rotate via bearing systems 38
about the engine central longitudinal axis A which is collinear
with their longitudinal axes.
[0018] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion.
[0019] The terms "low" and "high" as applied to speed or pressure
for the spools, compressors and turbines are of course relative to
each other. That is, the low speed spool operates at a lower speed
than the high speed spool, and the low pressure sections operate at
lower pressure than the high pressures sections.
[0020] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than ten (10), the geared architecture 48 is an epicyclic
gear train, such as a planetary gear system or other gear system,
with a gear reduction ratio of greater than about 2.3 and the low
pressure turbine 46 has a pressure ratio that is greater than about
5. In one disclosed embodiment, the engine 20 bypass ratio is
greater than about ten (10:1), the fan diameter is significantly
larger than that of the low pressure compressor 44, and the low
pressure turbine 46 has a pressure ratio that is greater than about
5:1. Low pressure turbine 46 pressure ratio is pressure measured
prior to inlet of low pressure turbine 46 as related to the
pressure at the outlet of the low pressure turbine 46 prior to an
exhaust nozzle. The geared architecture 48 may be an epicycle gear
train, such as a planetary gear system or other gear system, with a
gear reduction ratio of greater than about 2.5:1. In some
embodiments, the bypass ratio is less than about thirty (30), or
more narrowly less than about twenty (20). In embodiments, the gear
reduction ratio is less than about 5.0, or less than about 4.0. It
should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and
that the present invention is applicable to other gas turbine
engines including direct drive turbofans.
[0021] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of 1 bm
of fuel being burned divided by 1 bf of thrust the engine produces
at that minimum point. "Low fan pressure ratio" is the pressure
ratio across the fan blade alone, without a Fan Exit Guide Vane
("FEGV") system. The low fan pressure ratio as disclosed herein
according to one non-limiting embodiment is less than about 1.45.
"Low corrected fan tip speed" is the actual fan tip speed in ft/sec
divided by an industry standard temperature correction of
[(Tambient .degree. R)/(518.7).degree. R].sup.0.5. The "Low
corrected fan tip speed" as disclosed herein according to one
non-limiting embodiment is less than about 1150 ft/second.
[0022] The use of the gear reduction between the low speed spool
and the fan allows an increase of speed to the low pressure
compressor. In the past, the speed of the low pressure turbine and
compressor has been somewhat limited in that the fan speed cannot
be unduly large. The maximum fan speed is at its outer tip, and in
larger engines, the fan diameter is much larger than it may be in
smaller power engines. However, the use of the gear reduction has
freed the designer from limitation on the low pressure turbine and
compressor speeds caused by a desire to not have unduly high fan
speeds.
[0023] It has been discovered that a careful design between the
number of rotating blades, and the rotational speed of the low
pressure turbine can be selected to result in noise frequencies
that are less sensitive to human hearing. The same is true for the
low pressure compressor 44.
[0024] A formula has been developed as follows:
(blade count.times.rotational speed)/60 sec.gtoreq.5500 Hz.
[0025] That is, the number of rotating blades in any low pressure
turbine stage, multiplied by the rotational speed of the low
pressure turbine 46 (in revolutions per minute), divided by 60 sec
should be greater than or equal to about 5500 Hz. The same holds
true for the low pressure compressor stages. More narrowly, the
amounts should be greater than or equal to about 6000 Hz. In
embodiments, the amount is less than or equal to about 10000 Hz, or
more narrowly less than or equal to about 7000 Hz. A worker of
ordinary skill in the art would recognize that the 60 sec factor is
to change revolutions per minute to Hertz, or revolutions per one
second. For the purposes of this disclosure, the term "about"
means.+-.3% of the respective quantity unless otherwise
disclosed.
[0026] The operational speed of the low pressure turbine 46 and low
pressure compressor 44 as utilized in the formula should correspond
to the engine operating conditions at each noise certification
point defined in Part 36 or the Federal Airworthiness Regulations.
More particularly, the rotational speed may be taken as an approach
certification point as defined in Part 36 of the Federal
Airworthiness Regulations. For purposes of this application and its
claims, the term "approach speed" equates to this certification
point. In other embodiments, the above formula results in a number
that is less than or equal to about 7000 Hz at approach speed.
[0027] It is envisioned that all of the rows in the low pressure
turbine 46 meet the above formula. However, this application may
also extend to low pressure turbines wherein the majority of the
blade rows, or at least half of the blade rows, in the low pressure
turbine meet the above formula, but perhaps some may not. By
implication at least one, or less than half, of the rows meet the
formula. The same is true for low pressure compressors, wherein all
of the rows in the low pressure compressor 44 would meet the above
formula. However, the application may extend to low pressure
compressors wherein only the majority of the blade rows, or at
least half of the blade rows, in the low pressure compressor meet
the above formula, but some perhaps may not. Of course, by
implication the formula may be true for at least some of the
turbine rows but no compressor rows. In some cases, only one row of
the low pressure turbine and/or low pressure compressor may meet
the formula. Also, the formula may apply to at least some
compressor rows, but no row in the turbine meets the formula.
[0028] This will result in operational noise that would be less
sensitive to human hearing.
[0029] In embodiments, it may be that the formula can result in a
range of greater than or equal to 5500 Hz, and moving higher. Thus,
by carefully designing the number of blades and controlling the
operational speed of the low pressure turbine 46 (and a worker of
ordinary skill in the art would recognize how to control this
speed) one can assure that the noise frequencies produced by the
low pressure turbine are of less concern to humans.
[0030] The same holds true for designing the number of blades and
controlling the speed of the low pressure compressor 44. Again, a
worker of ordinary skill in the art would recognize how to control
the speed.
[0031] In embodiments, it may be only the low pressure turbine
rotor 46, or the low pressure compressor rotor 44 which is designed
to meet the meet the above formula. On the other hand, it is also
possible to ensure that both the low pressure turbine 46 and low
pressure compressor 44 meet the above formula.
[0032] This invention is most applicable to jet engines rated to
produce 15,000 pounds of thrust or more. In this thrust range,
prior art jet engines have typically had frequency ranges of about
4000 hertz. Thus, the noise problems as mentioned above have
existed.
[0033] Lower thrust engines (<15,000 pounds) may have operated
under conditions that sometimes passed above the 4000 Hz number,
and even approached 6000 Hz, however, this has not been in
combination with the geared architecture, nor in the higher powered
engines which have the larger fans, and thus the greater
limitations on low pressure turbine or low pressure compressor
speed.
[0034] FIG. 2 shows an embodiment 200, wherein there is a fan drive
turbine 208 driving a shaft 206 to in turn drive a fan rotor 202. A
gear reduction 204 may be positioned between the fan drive turbine
208 and the fan rotor 202. This gear reduction 204 may be
structured and operate like the gear reduction disclosed above. A
compressor rotor 210 is driven by an intermediate pressure turbine
212, and a second stage compressor rotor 214 is driven by a turbine
rotor 216. A combustion section 218 is positioned intermediate the
compressor rotor 214 and the turbine section 216.
[0035] FIG. 3 shows yet another embodiment 300 wherein a fan rotor
302 and a first stage compressor 304 rotate at a common speed. The
gear reduction 306 (which may be structured as disclosed above) is
intermediate the compressor rotor 304 and a shaft 308 which is
driven by a low pressure turbine section.
[0036] The FIGS. 2 and 3 engines may be utilized with the speed and
blade features disclosed above.
[0037] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
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