U.S. patent application number 15/245383 was filed with the patent office on 2016-12-15 for low noise 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.
Application Number | 20160362983 15/245383 |
Document ID | / |
Family ID | 55911852 |
Filed Date | 2016-12-15 |
United States Patent
Application |
20160362983 |
Kind Code |
A1 |
Morin; Bruce L. ; et
al. |
December 15, 2016 |
LOW NOISE TURBINE FOR GEARED TURBOFAN ENGINE
Abstract
A method of designing a gas turbine engine comprises the steps
of including a fan section with a fan. A turbine section is
included having a first turbine and a second turbine. A gear
reduction is included between the fan and the first turbine, the
gear reduction being configured to receive an input from the first
turbine and to turn the fan at a lower speed than the first turbine
in operation. The first turbine is designed to include a number of
turbine blades in each of a plurality of rows of the first turbine,
the first turbine blades operating at least some of the time at a
rotational speed, 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 first turbine: (number of
blades.times.speed)/60.gtoreq.5500.
Inventors: |
Morin; Bruce L.;
(Longmeadow, MA) ; Korte; Detlef; (Karlsfeld,
DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation
MTU Aero Engines AG |
Farmington
Munich |
CT |
US
DE |
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|
Family ID: |
55911852 |
Appl. No.: |
15/245383 |
Filed: |
August 24, 2016 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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15007784 |
Jan 27, 2016 |
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15245383 |
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14996544 |
Jan 15, 2016 |
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15007784 |
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14795931 |
Jul 10, 2015 |
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14996544 |
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14248386 |
Apr 9, 2014 |
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14795931 |
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PCT/US2013/020724 |
Jan 9, 2013 |
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14248386 |
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61592643 |
Jan 31, 2012 |
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61884660 |
Sep 30, 2013 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F04D 25/045 20130101;
F01D 5/14 20130101; F05D 2260/40311 20130101; F01D 15/12 20130101;
F01D 25/24 20130101; F02C 3/04 20130101; F05D 2270/333 20130101;
F04D 29/325 20130101; F04D 29/321 20130101; F02K 3/06 20130101;
F04D 29/053 20130101; F05D 2220/323 20130101; F05D 2260/96
20130101; F05D 2270/304 20130101; F02C 7/36 20130101; F05D 2230/50
20130101; F01D 5/06 20130101; F02C 3/107 20130101; G06F 30/17
20200101; F02K 3/04 20130101; F04D 29/663 20130101 |
International
Class: |
F01D 5/14 20060101
F01D005/14; F02C 3/04 20060101 F02C003/04; F02K 3/06 20060101
F02K003/06; F01D 15/12 20060101 F01D015/12; F04D 29/32 20060101
F04D029/32 |
Claims
1. A method of designing a gas turbine engine comprising the steps
of: including a fan section with a fan, the fan including at least
one fan blade; the fan section designed to achieve a low fan
pressure ratio less than about 1.45, wherein the low fan pressure
ratio is measured across a fan blade alone; further including a
turbine section having a first turbine and a second turbine;
further including a gear reduction between the fan and the first
turbine, the gear reduction including an epicycle gear train having
a gear reduction ratio of greater than about 2.5:1, and the gear
reduction being configured to receive an input from the first
turbine and to turn the fan at a lower speed than the first turbine
in operation; the first turbine designed to achieve a pressure
ratio greater than about 5:1, the first turbine including an inlet
having an inlet pressure, and an outlet that is prior to any
exhaust nozzle and having an outlet pressure, and the pressure
ratio of the first turbine being a ratio of the inlet pressure to
the outlet pressure; and the first turbine designed to further
include a number of turbine blades in each of a plurality of rows
of the first turbine, the first turbine blades operating at least
some of the time at a rotational speed, 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 first turbine:
(number of blades.times.speed)/60.gtoreq.5500; wherein the
rotational speed is 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 gas turbine
engine is designed to produce 15,000 pounds of thrust or more.
2. The method as recited in claim 1, wherein the formula results in
a number greater than 6000.
3. The method as recited in claim 2, wherein the formula results in
a number less than or equal to about 10000.
4. The method as recited in claim 3, wherein the formula results in
a number less than 7000.
5. The method as recited in claim 1, wherein the formula holds true
for a majority of the blade rows of the first turbine.
6. The method as recited in claim 5, wherein the formula results in
a number greater than 6000.
7. The method as recited in claim 6, wherein the formula results in
a number less than or equal to about 10000.
8. The method as recited in claim 7, wherein the formula results in
a number less than 7000.
9. The method as recited in claim 5, further including a
mid-turbine frame arranged between the second turbine and the first
turbine.
10. The method as recited in claim 9, further including a
compressor section configured to drive air along core flowpath, and
a plurality of bearing systems configured to support the first
turbine and the second turbine, wherein the mid-turbine frame
includes airfoils positioned in the core flowpath and is configured
to support at least one of the bearing systems.
11. The method as recited in claim 10, wherein the second turbine
has two stages.
12. The method as recited in claim 5, further including a first
compressor, and a shaft configured to be driven by the first
turbine, the gear reduction arranged intermediate the first
compressor and the shaft.
13. The method as recited in claim 12, wherein the second turbine
has two stages.
14. The method as recited in claim 5, further designing the engine
to achieve a bypass ratio greater than ten (10), and wherein the
fan is designed to have a low corrected fan tip speed less than
about 1150 ft/second, wherein the low corrected fan tip speed is an
actual fan tip speed in ft/second at an ambient temperature divided
by [(Tambient .degree. R)/(518.7 .degree. R)].sup.0.5.
15. The method as recited in claim 14, wherein the fan section is
designed for cruise.
16. The method as recited in claim 1, wherein the formula holds
true for all of the blade rows of the first turbine.
17. The method as recited in claim 16, wherein the formula results
in a number greater than 6000.
18. The method as recited in claim 17, wherein the formula results
in a number less than or equal to about 10000.
19. The method as recited in claim 18, wherein the formula results
in a number less than 7000.
20. The method as recited in claim 1, wherein the formula does not
hold true for all of the blade rows of the first turbine.
21. A method of designing a gas turbine engine comprising the steps
of: including a fan section with a fan, the fan including at least
one fan blade; the fan section designed to achieve a low fan
pressure ratio less than about 1.45, wherein the low fan pressure
ratio is measured across a fan blade alone; further including a
turbine section having a first turbine and a second turbine;
further including a gear reduction between the fan and the first
turbine, the gear reduction including an epicycle gear train having
a gear reduction ratio of greater than about 2.5:1, and the gear
reduction being configured to receive an input from the first
turbine and to turn the fan at a lower speed than the first turbine
in operation; the first turbine designed to achieve a pressure
ratio greater than about 5:1, the first turbine including an inlet
having an inlet pressure, and an outlet that is prior to any
exhaust nozzle and having an outlet pressure, and the pressure
ratio of the first turbine being a ratio of the inlet pressure to
the outlet pressure; the first turbine designed to further include
a number of turbine blades in each of a plurality of rows of the
first turbine, and the turbine blades of the first turbine
operating at least some of the time at a rotational speed, 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 first turbine: (number of blades.times.speed)/60.gtoreq.5500;
wherein the rotational speed is a cruise speed in revolutions per
minute, taken at a cruise certification point; and wherein the gas
turbine engine is designed to produce 15,000 pounds of thrust or
more.
22. The method as recited in claim 21, wherein the formula results
in a number less than 7000.
23. The method as recited in claim 21, wherein the formula holds
true for a majority of the blade rows of the first turbine.
24. The method as recited in claim 23, wherein the formula results
in a number less than 7000.
25. The method as recited in claim 23, further comprising a
mid-turbine frame arranged between the second turbine and the first
turbine.
26. The method as recited in claim 25, further including a
compressor section configured to drive air along core flowpath, and
a plurality of bearing systems configured to support the first
turbine and the second turbine, wherein the mid-turbine frame
includes airfoils positioned in the core flowpath and is configured
to support at least one of the bearing systems.
27. The method as recited in claim 23, further including a first
compressor, and a shaft configured to be driven by the first
turbine, the gear reduction arranged intermediate the first
compressor and the shaft.
28. The method as recited in claim 23, further designing the engine
to achieve a bypass ratio greater than ten (10), wherein the fan
section is designed for cruise, and wherein the fan is designed to
achieve a low corrected fan tip speed less than about 1150
ft/second, wherein the low corrected fan tip speed is an actual fan
tip speed in ft/second at an ambient temperature divided by
[(Tambient .degree. R)/(518.7 .degree. R)].sup.0.5.
29. The method as recited in claim 21, wherein the formula holds
true for all of the blade rows of the first turbine.
30. A method of designing a turbine section comprising the steps
of: including a low pressure turbine designed to achieve a pressure
ratio greater than about 5:1, the low pressure turbine including an
inlet having an inlet pressure, and an outlet that is prior to any
exhaust nozzle and having an outlet pressure, and the pressure
ratio of the low pressure turbine being a ratio of the inlet
pressure to the outlet pressure; and the low pressure turbine
further designed to include a number of turbine blades in each of a
plurality of rows of the low pressure turbine, a majority of the
turbine blades of the low pressure turbine operating at least some
of the time at a rotational speed, 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 turbine:
(number of blades.times.speed)/60.gtoreq.5500; wherein the
rotational speed is an approach speed in revolutions per minute,
taken at an approach certification point as defined in Part 36 of
the Federal Airworthiness Regulations.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 15/007,784, filed Jan. 27, 2016, which is a
continuation of U.S. patent application Ser. No. 14/996,544 filed
Jan. 15, 2016, which is a continuation-in-part of U.S. patent
application Ser. No. 14/795931, filed Jul. 10, 2015, which is a
continuation-in-part of U.S. patent application Ser. No.
14/248,386, filed Apr. 4, 2014, which is a continuation-in-part of
International Application No. PCT/US2013/020724 filed Jan. 9, 2013
which claims priority to U.S. Provisional Application No.
61/592,643, filed Jan. 31, 2012. U.S. patent application Ser. No.
14/248,386 further claims priority to U.S. Provisional Application
No. 61/884,660 filed Sep. 30, 2013.
BACKGROUND
[0002] This application relates to the design of a turbine which
can be operated to produce noise to which human hearing is less
sensitive.
[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 includes a
number of rows of turbine blades which rotate with the rotor.
Typically interspersed between the rows of turbine blades are
vanes.
[0005] The low pressure turbine 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 stages, and
their harmonics.
[0006] The noise can often be in a frequency range to which humans
are very sensitive. To mitigate this problem, in the past, a
vane-to-blade ratio of the fan drive turbine 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
acoustic "cut-off."
[0007] However, acoustically cut-off designs may come at the
expense of increased weight and reduced aerodynamic efficiency.
Stated another way, if limited 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.
[0008] 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 different speeds.
SUMMARY
[0009] In a featured embodiment, a method of designing a gas
turbine engine comprises the steps of including a fan section with
a fan, the fan including at least one fan blade. The fan section is
designed to achieve a low fan pressure ratio less than about 1.45,
wherein the low fan pressure ratio is measured across a fan blade
alone. A turbine section has a first turbine and a second turbine.
A gear reduction is included between the fan and the first turbine,
and includes an epicycle gear train having a gear reduction ratio
of greater than about 2.5:1. The gear reduction is configured to
receive an input from the first turbine and to turn the fan at a
lower speed than the first turbine in operation. The first turbine
is designed to achieve a pressure ratio greater than about 5:1. The
first turbine includes an inlet having an inlet pressure, and an
outlet that is prior to any exhaust nozzle and having an outlet
pressure. The pressure ratio of the first turbine is a ratio of the
inlet pressure to the outlet pressure. The first turbine is
designed to further include a number of turbine blades in each of a
plurality of rows of the first turbine, the first turbine blades
operating at least some of the time at a rotational speed, 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 first turbine: (number of blades.times.speed)/60.gtoreq.5500.
The rotational speed is an approach speed in revolutions per
minute, taken at an approach certification point as defined in Part
36 of the Federal Airworthiness Regulations. The gas turbine engine
is designed to produce 15,000 pounds of thrust or more.
[0010] In another embodiment according to the previous embodiment,
the formula results in a number greater than 6000.
[0011] In another embodiment according to any of the previous
embodiments, the formula results in a number less than or equal to
about 10000.
[0012] In another embodiment according to any of the previous
embodiments, the formula results in a number less than 7000.
[0013] In another embodiment according to any of the previous
embodiments, the formula holds true for a majority of the blade
rows of the first turbine.
[0014] In another embodiment according to any of the previous
embodiments, the formula results in a number greater than 6000.
[0015] In another embodiment according to any of the previous
embodiments, the formula results in a number less than or equal to
about 10000.
[0016] In another embodiment according to any of the previous
embodiments, the formula results in a number less than 7000.
[0017] In another embodiment according to any of the previous
embodiments, a mid-turbine frame is arranged between the second
turbine and the first turbine.
[0018] In another embodiment according to any of the previous
embodiments, a compressor section is configured to drive air along
core flowpath, and a plurality of bearing systems is configured to
support the first turbine and the second turbine. The mid-turbine
frame includes airfoils positioned in the core flowpath and is
configured to support at least one of the bearing systems.
[0019] In another embodiment according to any of the previous
embodiments, the second turbine has two stages.
[0020] In another embodiment according to any of the previous
embodiments, a first compressor is included, and a shaft is
configured to be driven by the first turbine. The gear reduction is
arranged intermediate the first compressor and the shaft.
[0021] In another embodiment according to any of the previous
embodiments, the second turbine has two stages.
[0022] In another embodiment according to any of the previous
embodiments, the engine is designed to achieve a bypass ratio
greater than ten (10). The fan is designed to have a low corrected
fan tip speed less than about 1150 ft/second, wherein the low
corrected fan tip speed is an actual fan tip speed in ft/second at
an ambient temperature divided by [(Tambient .degree.
R)/(518.7.degree. R)].sup.0.5.
[0023] In another embodiment according to any of the previous
embodiments, the fan section is designed for cruise.
[0024] In another embodiment according to any of the previous
embodiments, the formula holds true for all of the blade rows of
the first turbine.
[0025] In another embodiment according to any of the previous
embodiments, the formula results in a number greater than 6000.
[0026] In another embodiment according to any of the previous
embodiments, the formula results in a number less than or equal to
about 10000.
[0027] In another embodiment according to any of the previous
embodiments, the formula results in a number less than 7000.
[0028] In another embodiment according to any of the previous
embodiments, the formula does not hold true for all of the blade
rows of the first turbine.
[0029] In another featured embodiment, a method of designing a gas
turbine engine comprises the steps of including a fan section with
a fan, the fan having at least one fan blade. The fan section is
designed to achieve a low fan pressure ratio less than about 1.45,
wherein the low fan pressure ratio is measured across a fan blade
alone. A turbine section has a first turbine and a second turbine.
A gear reduction is between the fan and the first turbine and
includes an epicycle gear train having a gear reduction ratio of
greater than about 2.5:1. The gear reduction is configured to
receive an input from the first turbine and to turn the fan at a
lower speed than the first turbine in operation. The first turbine
is designed to achieve a pressure ratio greater than about 5:1, the
first turbine including an inlet having an inlet pressure, and an
outlet that is prior to any exhaust nozzle and having an outlet
pressure. The pressure ratio of the first turbine is a ratio of the
inlet pressure to the outlet pressure. The first turbine is
designed to further include a number of turbine blades in each of a
plurality of rows of the first turbine, and the turbine blades of
the first turbine operating at least some of the time at a
rotational speed, 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 first turbine: (number of
blades.times.speed)/60.gtoreq.5500. The rotational speed is a
cruise speed in revolutions per minute, taken at a cruise
certification point. The gas turbine engine is designed to produce
15,000 pounds of thrust or more.
[0030] In another embodiment according to the previous embodiment,
the formula results in a number less than 7000.
[0031] In another embodiment according to any of the previous
embodiments, the formula holds true for a majority of the blade
rows of the first turbine.
[0032] In another embodiment according to any of the previous
embodiments, the formula results in a number less than 7000.
[0033] In another embodiment according to any of the previous
embodiments, a mid-turbine frame is arranged between the second
turbine and the first turbine.
[0034] In another embodiment according to any of the previous
embodiments, a compressor section is configured to drive air along
core flowpath, and a plurality of bearing systems is configured to
support the first turbine and the second turbine, wherein the
mid-turbine frame includes airfoils positioned in the core flowpath
and is configured to support at least one of the bearing
systems.
[0035] In another embodiment according to any of the previous
embodiments, a first compressor is included, and a shaft is
configured to be driven by the first turbine. The gear reduction is
arranged intermediate the first compressor and the shaft.
[0036] In another embodiment according to any of the previous
embodiments, the engine is designed to achieve a bypass ratio
greater than ten (10), wherein the fan section is designed for
cruise, and wherein the fan is designed to achieve a low corrected
fan tip speed less than about 1150 ft/second, wherein the low
corrected fan tip speed is an actual fan tip speed in ft/second at
an ambient temperature divided by [(Tambient .degree.
R)/(518.7.degree. R)].sup.0.5.
[0037] In another embodiment according to any of the previous
embodiments, the formula holds true for all of the blade rows of
the first turbine.
[0038] In another featured embodiment, a method of designing a
turbine section comprises the steps of including a low pressure
turbine designed to achieve a pressure ratio greater than about
5:1. The low pressure turbine includes an inlet having an inlet
pressure, and an outlet that is prior to any exhaust nozzle and
having an outlet pressure. The pressure ratio of the low pressure
turbine is a ratio of the inlet pressure to the outlet pressure.
The low pressure turbine is further designed to include a number of
turbine blades in each of a plurality of rows of the low pressure
turbine, a majority of the turbine blades of the low pressure
turbine operating at least some of the time at a rotational speed,
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 turbine: (number of
blades.times.speed)/60.gtoreq.5500. The rotational speed is an
approach speed in revolutions per minute, taken at an approach
certification point as defined in Part 36 of the Federal
Airworthiness Regulations.
[0039] Although the different examples have the specific components
shown in the illustrations, embodiments of this disclosure are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0040] These and other features disclosed herein can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0041] FIG. 1 shows a gas turbine engine.
[0042] FIG. 2 shows another embodiment.
[0043] FIG. 3 shows yet another embodiment.
DETAILED DESCRIPTION
[0044] 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 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
two-spool 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 two-spool turbofans as the
teachings may be applied to other types of turbine engines
including three-spool architectures.
[0045] 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.
[0046] 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.
[0047] 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.
[0048] 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.
[0049] 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 star system, a planetary gear system or other
gear system, with a gear reduction ratio of greater than about
2.3:1 or greater than about 2.5:1. 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. The low pressure turbine 46
pressure ratio is a ratio of the pressure measured at inlet of low
pressure turbine 46 to the pressure at the outlet of the low
pressure turbine 46 (prior to an exhaust nozzle). 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.
[0050] 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 lbm of
fuel being burned divided by lbf 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 1.50 and, in some
embodiments, 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 deg R)/518.7) 0.5].
The "Low corrected fan tip speed" as disclosed herein according to
one non-limiting embodiment is less than about 1150 ft/second.
[0051] The use of the gear reduction between the low pressure
turbine spool and the fan allows an increase of speed to the low
pressure compressor. In the past, the speed of the low pressure
turbine has been somewhat limited in that the fan speed cannot be
unduly high. The maximum fan speed is at its outer tip, and in
larger engines, the fan diameter is much larger than it may be in
lower power engines. However, a gear reduction may be used to free
the designer from compromising low pressure turbine speed in order
not to have unduly high fan speeds.
[0052] 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.
[0053] A formula has been developed as follows:
(blade count.times.rotational speed)/(60
seconds/minute).gtoreq.4000 Hz.
[0054] That is, the number of rotating blades in any low pressure
turbine stage, multiplied by the rotational speed of the low
pressure turbine (in revolutions per minute), divided by 60 seconds
per minute (to put the amount per second, or Hertz) should be
greater than or equal to 4000 Hz. In one embodiment, the amount is
above 5500 Hz. And, in another embodiment, the amount is above
about 6000 Hz.
[0055] The operational speed of the low pressure turbine as
utilized in the formula should correspond to the engine operating
conditions at each noise certification point currently defined in
Part 36 or the Federal Airworthiness Regulations. More
particularly, the rotational speed may be taken as an approach
certification point as currently 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.
[0056] Although the above formula only needs to apply to one row of
blades in the low pressure turbine 26, in one embodiment, all of
the rows in the low pressure turbine meet the above formula. In
another embodiment, the majority of the blade rows in the low
pressure turbine meet the above formula.
[0057] This will result in operational noise to which human hearing
will be less sensitive.
[0058] In embodiments, it may be that the formula can result in a
range of greater than or equal to 4000 Hz, and moving higher. Thus,
by carefully designing the number of blades and controlling the
operational speed of the low pressure turbine (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.
[0059] This invention is most applicable to jet engines rated to
produce 15,000 pounds of thrust or more and with bypass ratios
greater than about 8.0.
[0060] 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.
[0061] 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.
[0062] 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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