Note: Descriptions are shown in the official language in which they were submitted.
CA 02863620 2014-10-29
PPH
LOW NOISE COMPRESSOR ROTOR FOR GEARED TURBOFAN ENGINE
BA CKGROUND
[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
1
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
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 OF THE INVENTION
[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 low pressure portion. The low pressure turbine portion drives
the low pressure
compressor portion and the fan. A gear reduction effects a reduction in the
speed of the fan
relative to a speed of the low pressure turbine and the low pressure
compressor portion. At least
one of the low pressure turbine portion and low pressure compressor portion
has 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 at least one of the low
pressure turbine portion
and/or the low pressure compressor sections: (number of blades x rotational
speed)/60 > 5500.
The rotational speed is an approach speed in revolutions per minute.
[0011] In another embodiment according to the previous embodiments,
the formula
results in a number greater than or equal to 6000 Hz.
[0012] In another embodiment according to any of the previous
embodiments, the gas
turbine engine is rated to produce 15,000 pounds of thrust or more.
2
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
[0013] In another embodiment according to any of the previous
embodiments, the at
least one of the low pressure turbine portion and the low pressure compressor
portion is the low
pressure compressor portion.
[0014] In another embodiment according to any of the previous
embodiments, the
formula holds true for a plurality of the blade rows of the low pressure
compressor portion.
[0015] In another embodiment according to any of the previous
embodiments, the
formula results in a number greater than or equal to 6000.
[0016] In another embodiment according to any of the previous
embodiments, the gas
turbine engine is rated to produce 15,000 pounds of thrust or more.
[0017] In another embodiment according to any of the previous
embodiments, the
gear reduction has a gear ratio of greater than about 2.3.
[0018] In another embodiment according to any of the previous
embodiments, the
gear reduction has a gear ratio of greater than about 2.5.
[0019] In another embodiment according to any of the previous
embodiments, the fan
delivers air into a bypass duct, and a portion of air into the compressor
section. A bypass ratio is
defined as the volume of air delivered into the bypass duct compared to the
volume of air
delivered into the compressor section. The bypass ratio is greater than about
6.
[0020] In another embodiment according to any of the previous
embodiments, the
bypass ratio is greater than about 10.
[0021] In another embodiment according to any of the previous
embodiments, the
gear reduction has a gear ratio of greater than about 2.3.
[0022] In another embodiment according to any of the previous
embodiments, the fan
delivers air into a bypass duct, and a portion of air into the compressor
section. A bypass ratio is
defined as the volume of air delivered into the bypass duct compared to the
volume of air
delivered into the compressor section. The bypass ratio is greater than about
6.
[0023] In another embodiment according to any of the previous
embodiments, the
bypass ratio is greater than about 10.
[0024] In another featured embodiment, a compressor module has a low
pressure
portion having a number of blades in each of a plurality of rows of the low
pressure portion. The
3
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
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 low pressure portion: (number of blades x rotational speed)/60 > 5500
Hz. The rotational
speed is an approach speed in revolutions per minute.
[0025] In another embodiment according to the previous embodiment, the
formula
results in a number greater than or equal to 6000 Hz.
[0026] In another embodiment according to any of the previous
embodiments, the gas
turbine engine is rated to produce 15,000 pounds of thrust or more.
[0027] In another embodiment according to any of the previous
embodiments, the
formula holds true for a plurality of the blade rows of the low pressure
portion.
[0028] In another embodiment according to any of the previous
embodiments, the
formula holds true for all of the blade rows of the low pressure portion.
[0029] In another embodiment according to any of the previous
embodiments, the
formula holds true for a plurality of the blade rows of the low pressure
portion.
[0030] 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
[0031] Figure 1 shows a gas turbine engine.
DETAILED DESCRIPTION
[0032] Figure 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-
4
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
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.
[0033]
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.
[0034]
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.
[0035]
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.
[0036]
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.
[0037]
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
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
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.
[0038] 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 (`TSFCT - 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)]" .
The "Low corrected fan tip speed" as disclosed herein according to one non-
limiting
embodiment is less than about 1150 ft / second.
[0039] 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
6
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
designer from limitation on the low pressure turbine and compressor speeds
caused by a desire to
not have unduly high fan speeds.
[0040] 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.
[0041] A formula has been developed as follows:
(blade count x rotational speed)/60> 5500.
[0042] 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 should be greater than or equal to 5500. The same holds true for
the low pressure
compressor stages. More narrowly, the amounts should be above 6000. A worker
of ordinary
skill in the art would recognize that the 60 factor is to change revolutions
per minute to Hertz, or
revolutions per one second.
[0043] 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.
[0044] It is envisioned that all of the rows in the low pressure
turbine 46 may meet
the above formula. However, this application may also extend to low pressure
turbines wherein
only one of the blade rows in the low pressure turbine meet the above formula.
In other
embodiments, plural rows meet the formula and in other embodiments, the
majority 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 may meet the above formula. However, the
application may extend
to low pressure compressors wherein only one of the blade rows in the low
pressure compressor
meet the above formula. In other embodiments, plural rows meet the formula and
in other
embodiments, the majority of the rows meet the formula.
7
CA 02863620 2014-07-31
WO 2013/122713 PCT/US2013/022035
[0045] This will result in operational noise that would be less
sensitive to human
hearing.
[0046] 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 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.
[0047] 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.
[0048] 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.
[0049] 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.
[0050] 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 or
low pressure
compressor speed.
[0051] 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.
8
