U.S. patent application number 13/590328 was filed with the patent office on 2013-08-08 for low noise turbine for geared turbofan engine.
The applicant listed for this patent is Detlef Korte, Bruce L. Morin. Invention is credited to Detlef Korte, Bruce L. Morin.
Application Number | 20130202403 13/590328 |
Document ID | / |
Family ID | 46641531 |
Filed Date | 2013-08-08 |
United States Patent
Application |
20130202403 |
Kind Code |
A1 |
Morin; Bruce L. ; et
al. |
August 8, 2013 |
LOW NOISE TURBINE FOR GEARED TURBOFAN ENGINE
Abstract
A gas turbine engine is utilized in combination with a gear
reduction to reduce the speed of a fan relative to a low pressure
turbine speed. The gas turbine engine is designed such that a blade
count in the low pressure turbine multiplied by the speed of the
low pressure turbine will result in operational noise that is above
a sensitive range for human hearing. A method and turbine module
are also disclosed.
Inventors: |
Morin; Bruce L.;
(Springfield, MA) ; Korte; Detlef; (Munich,
DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Morin; Bruce L.
Korte; Detlef |
Springfield
Munich |
MA |
US
DE |
|
|
Family ID: |
46641531 |
Appl. No.: |
13/590328 |
Filed: |
August 21, 2012 |
Related U.S. Patent Documents
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
|
|
13403005 |
Feb 23, 2012 |
8246292 |
|
|
13590328 |
|
|
|
|
61592643 |
Jan 31, 2012 |
|
|
|
Current U.S.
Class: |
415/1 ;
415/122.1 |
Current CPC
Class: |
F02C 7/20 20130101; F01D
5/02 20130101; Y02T 50/60 20130101; F05D 2260/961 20130101; F05D
2270/333 20130101; F02K 3/06 20130101; F02C 7/06 20130101; F04D
29/056 20130101; Y02T 50/671 20130101; F02C 3/107 20130101 |
Class at
Publication: |
415/1 ;
415/122.1 |
International
Class: |
F01D 5/02 20060101
F01D005/02 |
Claims
1. A gas turbine engine comprising: a fan and a turbine having a
fan drive rotor; a gear reduction effecting a reduction in the
speed of said fan relative to an input speed from said fan drive
rotor; said fan drive rotor having a number of turbine blades in at
least one of a plurality of rows of said fan drive rotor, and said
turbine blades operating at least some of the time at a rotational
speed, and said number of turbine blades in said at least one row
and said rotational speed being such that the following formula
holds true for said at least one row of the fan drive turbine
(number of blades.times.speed)/60.gtoreq.5500 Hz; and said
rotational speed being in revolutions per minute.
2. The gas turbine engine as set forth in claim 1, wherein the
formula results in a number greater than or equal to 6000 Hz.
3. The gas turbine engine as set forth in claim 2, wherein said gas
turbine engine is rated to produce 15,000 pounds of thrust or
more.
4. The gas turbine engine as set forth in claim 1, wherein the
formula holds true for the majority of blade rows of the fan drive
rotor.
5.-7. (canceled)
8. A method of designing a gas turbine engine comprising the steps
of: including a gear reduction between a fan drive turbine rotor
and a fan, and selecting a number of blades in at least one row of
the fan drive turbine rotor, in combination with a rotational speed
of the fan drive turbine rotor, such that the following formula
holds true for said at least one row of the fan drive turbine
rotor: (number of blades.times.speed)/60.gtoreq.5500 Hz; and said
rotational speed being in revolutions per minute.
9. The method of designing a gas turbine engine as set forth in
claim 8, wherein the formula results in a number greater than or
equal to 6000.
10. The method of designing a gas turbine engine as set forth in
claim 9, wherein said gas turbine engine is rated to produce 15,000
pounds of thrust or more.
11. The method as set forth in claim 8, wherein the formula holds
true for the majority of the blade rows of the fan drive
turbine.
12.-15. (canceled)
16. The turbine module as set forth in claim 27, wherein the
formula results in a number greater than or equal to 6000.
17. (canceled)
18. The turbine module as set forth in claim 27, wherein the
formula holds true for the majority of blade rows of the fan drive
rotor.
19.-20. (canceled)
21. The gas turbine engine as set forth in claim 1, wherein said
rotational speed being an approach speed.
22. The method as set forth in claim 8, wherein said rotational
speed is an approach speed.
23. The turbine module as set forth in claim 27, wherein the
rotational speed is an approach speed.
24. The gas turbine engine as set forth in claim 1, wherein said
turbine section having a higher pressure turbine rotor and a lower
pressure turbine rotor, with said fan drive rotor being said lower
pressure turbine rotor.
25. The method as set forth in claim 8, wherein a turbine section
including a higher pressure turbine rotor and a lower pressure
turbine rotor, and said fan drive turbine rotor being said lower
pressure turbine rotor.
26. The turbine module as set forth in claim 27, wherein there
being a higher pressure turbine rotor and a lower pressure turbine
rotor, and said fan drive rotor being said lower pressure turbine
rotor.
27. A turbine module comprising: a fan drive rotor having a first
blade row that includes a number of blades, the first blade row
being capable of rotating at a rotational speed, so that when
measuring said rotational speed in revolutions per minute: (number
of blades.times.said rotational speed)/60.gtoreq.5500 Hz.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 13/403,005, filed Feb. 23, 2012, which claims
priority to U.S. Provisional Application No. 61/592,643, filed Jan.
31, 2012.
BACKGROUND
[0002] This application relates to the design of a turbine 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 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 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."
[0007] 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.
[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 distinct speeds.
SUMMARY OF THE INVENTION
[0009] In a featured embodiment, a gas turbine engine has a fan, a
compressor section including a low pressure compressor and a high
pressure compressor, a combustor section, a high pressure turbine
and a low pressure turbine. The low pressure turbine drives the low
pressure compressor and fan. A gear reduction is utilized to reduce
the speed of the fan from an input speed from the low pressure
turbine. The low pressure turbine has a number of turbine blades in
each of a plurality of rows of the turbine. The low pressure
turbine blades operate at a rotational speed. The number of blades
and the rotational speed are selected 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.
[0010] In another embodiment according to the foregoing embodiment,
the formula results in a number greater than or equal to 6000.
[0011] In another embodiment according to the foregoing embodiment,
the gas turbine engine is rated to produce 15,000 pounds of thrust
or more.
[0012] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for the majority of the blade
rows of the low pressure turbine.
[0013] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for all of the blade rows of the
low pressure turbine.
[0014] In another embodiment according to the featured embodiment,
the formula holds true for the majority of the blade rows of the
low pressure turbine.
[0015] In another embodiment according to the featured embodiment,
the formula holds true for all of the blade rows of the low
pressure turbine.
[0016] In another featured embodiment, a method of designing a gas
turbine engine includes the steps of including a gear reduction
between a low pressure turbine and a fan, and selecting the number
of blades of the low pressure turbine rotors, in combination with
the rotational speed of the low pressure turbine, 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.
[0017] In another embodiment according to the foregoing embodiment,
the formula results in a number greater than or equal to 6000.
[0018] In another embodiment according to the foregoing embodiment,
the gas turbine engine is rated to produce 15,000 pounds of thrust
or more.
[0019] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for the majority of the blade
rows of the low pressure turbine.
[0020] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for all of the blade rows of the
low pressure turbine.
[0021] In another embodiment according to the featured embodiment,
the formula holds true for the majority of the blade rows of the
low pressure turbine.
[0022] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for all of the blade rows of the
low pressure turbine.
[0023] In another featured embodiment, a turbine module for a gas
turbine engine has a low pressure turbine with a number of turbine
blades in each of a plurality of rows of the turbine. The low
pressure turbine blades operate at a rotational speed. The number
of blades and the rotational speed are selected 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.
[0024] In another embodiment according to the foregoing embodiment,
the formula results in a number greater than or equal to 6000.
[0025] In another embodiment according to the foregoing embodiment,
the gas turbine engine is rated to produce 15,000 pounds of thrust
or more.
[0026] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for the majority of the blade
rows of the low pressure turbine.
[0027] In another embodiment according to the immediately foregoing
embodiment, the formula holds true for all of the blade rows of the
low pressure turbine.
[0028] In another embodiment according to the featured embodiment,
the formula holds true for the majority of the blade rows of the
low pressure turbine.
[0029] In another embodiment according to the featured embodiment,
the formula holds true for all of the blade rows of the low
pressure turbine. These and other features of the invention would
be better understood from the following specifications and
drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0030] FIG. 1 shows a gas turbine engine.
DETAILED DESCRIPTION
[0031] 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 while the compressor
section 24 drives air along a core flowpath 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.
[0032] 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.
[0033] 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.
[0034] 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.
[0035] 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.
[0036] 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. 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.
[0037] 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 1.45. "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.
[0038] 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 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
speed caused by a desire to not have unduly high fan speeds.
[0039] 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.
[0040] A formula has been developed as follows:
(blade count.times.rotational speed)/60.gtoreq.5500.
[0041] 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 should
be greater than or equal to 5500. More narrowly, the amount should
be above 6000.
[0042] 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 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.
[0043] It is envisioned that all of the rows in the low pressure
turbine meet the above formula. However, this application may also
extend to low pressure turbines wherein the majority of the blade
rows in the low pressure turbine meet the above formula, but
perhaps some may not.
[0044] This will result in operational noise that would be less
sensitive to human hearing.
[0045] In embodiments, it may be that the formula can result in a
range of greater than or equal to 5500, 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.
[0046] 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.
[0047] Lower thrust engines (<15,000 pounds) may have operated
under conditions that sometimes passed above the 4000 number, and
even approached 6000, 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 speed.
[0048] 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.
* * * * *