U.S. patent application number 15/292438 was filed with the patent office on 2017-10-19 for turbine section of high bypass turbofan.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Paul R. Adams, Wesley K. Lord, Shankar S. Magge, Frederick M. Schwarz, Joseph Brent Staubach, Gabriel L. Suciu.
Application Number | 20170298832 15/292438 |
Document ID | / |
Family ID | 54929995 |
Filed Date | 2017-10-19 |
United States Patent
Application |
20170298832 |
Kind Code |
A1 |
Adams; Paul R. ; et
al. |
October 19, 2017 |
TURBINE SECTION OF HIGH BYPASS TURBOFAN
Abstract
A turbofan engine has an engine case and a gaspath through the
engine case. A fan has a circumferential array of fan blades. The
engine further has a compressor, a combustor, a gas generating
turbine, and a low pressure turbine section. A speed reduction
mechanism couples the low pressure turbine section to the fan. A
bypass area ratio is greater than about 6.0. The low pressure
turbine section airfoil count to bypass area ratio is below about
170.
Inventors: |
Adams; Paul R.;
(Glastonbury, CT) ; Magge; Shankar S.; (South
Windsor, CT) ; Staubach; Joseph Brent; (Colchester,
CT) ; Lord; Wesley K.; (South Glastonbury, CT)
; Schwarz; Frederick M.; (Glastonbury, CT) ;
Suciu; Gabriel L.; (Glastonbury, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
54929995 |
Appl. No.: |
15/292438 |
Filed: |
October 13, 2016 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14793785 |
Jul 8, 2015 |
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15292438 |
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14692090 |
Apr 21, 2015 |
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14793785 |
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13599175 |
Aug 30, 2012 |
9010085 |
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14692090 |
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13475252 |
May 18, 2012 |
8844265 |
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13599175 |
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11832107 |
Aug 1, 2007 |
8256707 |
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13475252 |
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61593190 |
Jan 31, 2012 |
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61498516 |
Jun 17, 2011 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 11/122 20130101;
F05D 2220/32 20130101; F02C 9/18 20130101; F02C 3/107 20130101;
F02K 3/06 20130101; F02K 3/075 20130101; F05D 2220/323 20130101;
F02C 7/20 20130101; F05D 2240/35 20130101; F01D 25/24 20130101;
F04D 19/02 20130101; F01D 5/06 20130101; F02C 7/36 20130101; F05B
2250/283 20130101; F05D 2240/60 20130101; F05D 2260/40311 20130101;
Y02T 50/60 20130101; F02C 3/04 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F01D 11/12 20060101 F01D011/12; F01D 25/24 20060101
F01D025/24; F02C 3/04 20060101 F02C003/04; F02C 9/18 20060101
F02C009/18; F02C 7/20 20060101 F02C007/20; F04D 19/02 20060101
F04D019/02; F01D 5/06 20060101 F01D005/06; F02K 3/06 20060101
F02K003/06; F02K 3/075 20060101 F02K003/075; F02C 3/107 20060101
F02C003/107 |
Claims
1. A turbofan engine comprising: a fan including a circumferential
array of fan blades; a compressor in fluid communication with the
fan, the compressor including a second compressor section and a
first compressor section, the second compressor section including a
second compressor section inlet with a second compressor section
inlet annulus area; a fan duct including a fan duct annulus area
outboard of the second compressor section inlet, wherein the ratio
of the fan duct annulus area to the second compressor section inlet
annulus area defines a bypass area ratio that is greater than 6.0;
a combustor in fluid communication with the compressor; a shaft
assembly having a first portion and a second portion; a turbine in
fluid communication with the combustor, the turbine having a
two-stage first turbine section coupled to the first portion of the
shaft assembly to drive the first compressor section, and a
four-stage second turbine section coupled to the second portion of
the shaft assembly to drive the fan, each of the second turbine
section including blades and vanes, and a second turbine airfoil
count defined as the numerical count of all of the blades and vanes
in the second turbine section; wherein a ratio of the second
turbine airfoil count to the bypass area ratio is less than 150;
wherein the second turbine section further includes a maximum gas
path radius and the fan blades include a maximum radius, and a
ratio of the maximum gas path radius to the maximum radius of the
fan blades is less than 0.55; and wherein a hub-to-tip ratio
(Ri:Ro) of the second turbine section is between 0.4 and 0.5
measured at the maximum Ro axial location in the second turbine
section; and a planetary gearbox coupled to the fan and rotatable
by the second turbine section through the second portion of the
shaft assembly to allow the second turbine to turn faster than the
fan, the gearbox having a speed reduction ratio between 2:1 and
13:1 determined by the ratio of diameters within the gearbox.
2. The turbofan engine as recited in claim 1, wherein the bypass
area ratio is greater than 8.0 and the sum of the plurality of
stages of the second compressor section, first compressor section,
second turbine section and first turbine section is no less than
seventeen and no more than twenty one.
3. The turbofan engine as recited in claim 2, wherein the second
compressor section is a four-stage second compressor.
4. The turbofan engine as recited in claim 3, wherein the first
compressor section is a nine-stage first compressor.
5. The turbofan engine as recited in claim 1, wherein the bypass
area ratio is greater than 8.0.
6. The turbofan engine as recited in claim 5, further comprising a
fan case and vanes, the fan case encircling the fan and supported
by the vanes, and wherein the second turbine section includes a
plurality of blade stages interspersed with a plurality of vane
stages, and each stage of the second turbine section includes a
disk with a circumferential array of blades, each blade including
an airfoil extending from an inner diameter to an outer diameter,
wherein the inner diameter is associated with a platform and the
outer diameter is associated with a shroud.
7. The turbofan engine as recited in claim 1, wherein the bypass
area ratio is greater than 8.0 and further comprising: an engine
aft mount location configured to support an engine mount when the
engine is mounted and to react at least a thrust load of the
engine; and an engine forward mount location configured to support
an engine mount when the engine is mounted.
8. The turbofan engine as recited in claim 7, wherein the engine
aft mount location engages with an engine thrust case.
9. The turbofan engine as recited in claim 8, wherein the engine
aft mount location is located between the second turbine section
and the first turbine section.
10. The turbofan engine as recited in claim 9, wherein the engine
forward mount location engages with an intermediate case.
11. A turbofan engine comprising: a fan including a circumferential
array of fan blades; a compressor in fluid communication with the
fan, the compressor including a second compressor section and a
first compressor section, the second compressor section including a
second compressor section inlet with a second compressor section
inlet annulus area; a fan duct including a fan duct annulus area
outboard of the second compressor section inlet, wherein the ratio
of the fan duct annulus area to the second compressor section inlet
annulus area defines a bypass area ratio that is greater than 6.0;
a combustor in fluid communication with the compressor; a turbine
in fluid communication with the combustor, the turbine having a
first turbine section and a second turbine section, wherein a
hub-to-tip ratio (Ri:Ro) of the second turbine section is between
0.4 and 0.5 measured at the maximum Ro axial location in the second
turbine section; and an epicyclic gearbox coupled to the fan and
rotatable by the second turbine section through the second portion
of the shaft assembly to allow the second turbine section to turn
faster than the fan, the gearbox having a speed reduction ratio
between 2:1 and 13:1.
12. The turbofan engine as recited in claim 11, wherein the bypass
area ratio is greater than 8.0 and wherein the sum of the plurality
of stages of the second compressor section, first compressor
section, second turbine section and first turbine section is no
less than seventeen and no more than twenty one.
13. The turbofan engine as recited in claim 12, wherein the first
turbine section is a two-stage first turbine.
14. The turbofan engine as recited in claim 13, wherein the second
turbine section is a four-stage second turbine.
15. The turbofan engine as recited in claim 11, wherein the second
compressor section is a four-stage second compressor.
16. The turbofan engine as recited in claim 11, wherein the first
compressor section is a nine-stage first compressor.
17. The turbofan engine as recited in claim 11, wherein the bypass
area ratio is greater than 8.0, the second turbine section is a
four-stage second turbine and the first turbine section is a
two-stage first turbine.
18. The turbofan engine as recited in claim 17, wherein the gearbox
is a planetary gearbox, and further comprising a fan case and
vanes, the fan case encircling the fan and supported by the
vanes.
19. The turbofan engine as recited in claim 11, wherein the bypass
area ratio is greater than 8.0 and the second turbine section
includes a second turbine airfoil count defined as the numerical
count of all of the blades and vanes in the second turbine section,
wherein a ratio of the second turbine airfoil count to the bypass
area ratio is less than 150.
20. The turbofan engine as recited in claim 11, wherein the bypass
area ratio is greater than 8.0, the second turbine section further
includes a maximum gas path radius, and the fan blades include a
maximum radius, wherein a ratio of the maximum gas path radius to
the maximum radius of the fan blades is less than 0.55.
21. The turbofan engine as recited in claim 11, wherein the bypass
area ratio is greater than 8.0.
22. The turbofan engine as recited in claim 11, further comprising
an engine intermediate case, including an engine forward mount
location configured to support an engine mount when the engine is
mounted, and an engine thrust case including an engine aft mount
location configured to support an engine mount and react at least a
thrust load when the engine is mounted.
23. The turbofan engine as recited in claim 22, wherein the engine
aft mount location is located between the second turbine section
and the first turbine section.
24. A turbofan engine comprising: a fan including a circumferential
array of fan blades; a compressor in fluid communication with the
fan, the compressor including a first compressor section and a
second compressor section; a combustor in fluid communication with
the compressor; a turbine in fluid communication with the
combustor, the turbine having a first turbine section and a second
turbine section, the second turbine section driving the first
compressor section and the fan, wherein: the second turbine section
includes no less than three stages and no more than five stages; a
hub-to-tip ratio (Ri:Ro) of the second turbine section is between
0.4 and 0.5 measured at the maximum Ro axial location in the second
turbine section; and a planetary gearbox coupled to the fan and
rotatable by the second turbine section through the second portion
of the shaft assembly to allow the second turbine to turn faster
than the fan, the gearbox having a speed reduction ratio between
2:1 and 13:1.
25. The turbofan engine as recited in claim 24, wherein the second
turbine section includes a second turbine airfoil count defined as
the numerical count of all of the blades and vanes in the second
turbine section, wherein a ratio of the second turbine airfoil
count to the bypass area ratio is less than 150.
26. The turbofan engine as recited in claim 24, wherein the second
turbine section further includes a maximum gas path radius and the
fan blades include a maximum radius, and a ratio of the maximum gas
path radius to the maximum radius of the fan blades is less than
0.55.
27. The turbofan engine as recited in claim 24, further comprising:
the second compressor section including a second compressor section
inlet with a second compressor section inlet annulus area; and a
fan duct including a fan duct annulus area outboard of the second
compressor section inlet, wherein the ratio of the fan duct annulus
area to the second compressor section inlet annulus area defines a
bypass area ratio that is greater than 6.0.
28. The turbofan engine as recited in claim 24, wherein the second
turbine section is a four-stage second turbine and the first
turbine is a two-stage first turbine.
29. A turbine module for a geared turbofan engine comprising: a
plurality of stages configured to drive a compressor and cooperate
with a gearbox to drive a fan upon connection of the turbine module
with other sections of the geared turbofan engine; and a core
gaspath defined between an inboard boundary characterized by a
radius Ri and an outboard boundary characterized by a radius Ro,
wherein a ratio of the radius Ri to the radius Ro (Ri:Ro) is
between 0.4 and 0.5 measured at the maximum Ro axial location.
30. The turbine module of claim 29, wherein the plurality of stages
is no less than three stages and no more than four stages.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 14/793,785, filed Jul. 8, 2015, which is a
continuation-in-part of U.S. patent application Ser. No.
14/692,090, filed Apr. 21, 2015, which was a continuation of U.S.
patent application Ser. No. 13/599,175, filed Aug. 30, 2012, which
was a continuation of U.S. patent application Ser. No. 13/475,252,
filed May 18, 2012, now U.S. Pat. No. 8,844,265, issued Sep. 30,
2014, which was a continuation-in-part of U.S. patent application
Ser. No. 11/832,107, filed Aug. 1, 2007, and claimed the benefit of
U.S. Patent Provisional Application No. 61/593,190, filed Jan. 31,
2012, and U.S. Provisional Application No. 61/498,516, filed Jun.
17, 2011.
BACKGROUND
[0002] The disclosure relates to turbofan engines. More
particularly, the disclosure relates to low pressure turbine
sections of turbofan engines which power the fans via a speed
reduction mechanism.
[0003] There has been a trend toward increasing bypass ratio in gas
turbine engines. This is discussed further below. There has
generally been a correlation between certain characteristics of
bypass and the diameter of the low pressure turbine section
sections of turbofan engines.
SUMMARY
[0004] One aspect of the disclosure involves a turbofan engine
having an engine case and a gaspath through the engine case. A fan
has a circumferential array of fan blades. The engine further has a
compressor in fluid communication with the fan, a combustor in
fluid communication with the compressor, a turbine in fluid
communication with the combustor, wherein the turbine includes a
low pressure turbine section having 3 to 6 blade stages. A speed
reduction mechanism couples the low pressure turbine section to the
fan. A bypass area ratio is greater than about 6.0. A ratio of the
total number of airfoils in the low pressure turbine section
divided by the bypass area ratio is less than about 170, said low
pressure turbine section airfoil count being the total number of
blade airfoils and vane airfoils of the low pressure turbine
section.
[0005] In additional or alternative embodiments of any of the
foregoing embodiments, the bypass area ratio may be greater than
about 8.0 or may be between about 8.0 and about 20.0.
[0006] In additional or alternative embodiments of any of the
foregoing embodiments, a fan case may encircle the fan blades
radially outboard of the engine case.
[0007] In additional or alternative embodiments of any of the
foregoing embodiments, the compressor may comprise a low pressure
compressor section and a high pressure compressor section.
[0008] In additional or alternative embodiments of any of the
foregoing embodiments, the blades of the low pressure compressor
section and low pressure turbine section may share a low shaft.
[0009] In additional or alternative embodiments of any of the
foregoing embodiments, the high pressure compressor section and a
high pressure turbine section of the turbine may share a high
shaft.
[0010] In additional or alternative embodiments of any of the
foregoing embodiments, there are no additional compressor or
turbine sections.
[0011] In additional or alternative embodiments of any of the
foregoing embodiments, the speed reduction mechanism may comprise
an epicyclic transmission coupling the low speed shaft to a fan
shaft to drive the fan with a speed reduction.
[0012] In additional or alternative embodiments of any of the
foregoing embodiments, the low pressure turbine section may have an
exemplary 2 to 6 blade stages or 2 to 3 blade stages.
[0013] In additional or alternative embodiments of any of the
foregoing embodiments, a hub-to-tip ratio (Ri:Ro) of the low
pressure turbine section may be between about 0.4 and about 0.5
measured at the maximum Ro axial location in the low pressure
turbine section.
[0014] In additional or alternative embodiments of any of the
foregoing embodiments, a ratio of maximum gaspath radius along the
low pressure turbine section to maximum radius of the fan may be
less than about 0.55, or less than about 0.50, or between about
0.35 and about 0.50.
[0015] In additional or alternative embodiments of any of the
foregoing embodiments, the ratio of low pressure turbine section
airfoil count to bypass area ratio may be between about 10 and
about 150.
[0016] In additional or alternative embodiments of any of the
foregoing embodiments, the airfoil count of the low pressure
turbine section may be below about 1600.
[0017] In additional or alternative embodiments of any of the
foregoing embodiments, the engine may be in combination with a
mounting arrangement (e.g., of an engine pylon) wherein an aft
mount reacts at least a thrust load.
[0018] The details of one or more embodiments are set forth in the
accompanying drawings and the description below. Other features,
objects, and advantages will be apparent from the description and
drawings, and from the claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0019] FIG. 1 is an axial sectional view of a turbofan engine.
[0020] FIG. 2 is an axial sectional view of a low pressure turbine
section of the engine of FIG. 1.
[0021] FIG. 3 is transverse sectional view of transmission of the
engine of FIG. 1.
[0022] FIG. 4 shows another embodiment.
[0023] FIG. 5 shows yet another embodiment.
[0024] Like reference numbers and designations in the various
drawings indicate like elements.
DETAILED DESCRIPTION
[0025] FIG. 1 shows a turbofan engine 20 having a main housing
(engine case) 22 containing a rotor shaft assembly 23. An exemplary
engine is a high-bypass turbofan. In such an engine, the normal
cruise condition bypass area ratio of air mass flowing outside the
case 22 (e.g., the compressor sections and combustor) to air mass
passing through the case 22 is typically in excess of about 4.0
and, more narrowly, typically between about 4.0 and about 12.0. Via
high 24 and low 25 shaft portions of the shaft assembly 23, a high
pressure turbine section (gas generating turbine) 26 and a low
pressure turbine section 27 respectively drive a high pressure
compressor section 28 and a low pressure compressor section 30. As
used herein, the high pressure turbine section experiences higher
pressures that the low pressure turbine section. A low pressure
turbine section is a section that powers a fan 42. Although a
two-spool (plus fan) engine is shown, one of many alternative
variations involves a three-spool (plus fan) engine wherein an
intermediate spool comprises an intermediate pressure compressor
between the low fan and high pressure compressor section and an
intermediate pressure turbine between the high pressure turbine
section and low pressure turbine section.
[0026] The engine extends along a longitudinal axis 500 from a fore
end to an aft end. Adjacent the fore end, a shroud (fan case) 40
encircles the fan 42 and is supported by vanes 44. An aerodynamic
nacelle around the fan case is shown and an aerodynamic nacelle 45
around the engine case is shown.
[0027] The low shaft portion 25 of the rotor shaft assembly 23
drives the fan 42 through a speed reduction mechanism 46. An
exemplary speed reduction mechanism is an epicyclic transmission,
namely a star or planetary gear system. As is discussed further
below, an inlet airflow 520 entering the nacelle is divided into a
portion 522 passing along a core flowpath 524 and a bypass portion
526 passing along a bypass flowpath 528. With the exception of
diversions such as cooling air, etc., flow along the core flowpath
sequentially passes through the low pressure compressor section,
high pressure compressor section, a combustor 48, the high pressure
turbine section, and the low pressure turbine section before
exiting from an outlet 530.
[0028] FIG. 3 schematically shows details of the transmission 46. A
forward end of the low shaft 25 is coupled to a sun gear 52 (or
other high speed input to the speed reduction mechanism). The
externally-toothed sun gear 52 is encircled by a number of
externally-toothed star gears 56 and an internally-toothed ring
gear 54. The exemplary ring gear is coupled to the fan to rotate
with the fan as a unit.
[0029] The star gears 56 are positioned between and enmeshed with
the sun gear and ring gear. A cage or star carrier assembly 60
carries the star gears via associated journals 62. The exemplary
star carrier is substantially irrotatably mounted relative via
fingers 404 to the case 22.
[0030] Another transmission/gearbox combination has the star
carrier connected to the fan and the ring is fixed to the fixed
structure (case) is possible and such is commonly referred to as a
planetary gearbox.
[0031] The speed reduction ratio is determined by the ratio of
diameters within the gearbox. An exemplary reduction is between
about 2:1 and about 13:1.
[0032] The exemplary fan (FIG. 1) comprises a circumferential array
of blades 70. Each blade comprises an airfoil 72 having a leading
edge 74 and a trailing edge 76 and extending from an inboard end 78
at a platform to an outboard end 80 (i.e., a free tip). The
outboard end 80 is in close facing proximity to a rub strip 82
along an interior surface 84 of the nacelle and fan case.
[0033] To mount the engine to the aircraft wing 92, a pylon 94 is
mounted to the fan case and/or to the other engine cases. The
exemplary pylon 94 may be as disclosed in U.S. patent application
Ser. No. 11/832,107 (US2009/0056343A1). The pylon comprises a
forward mount 100 and an aft/rear mount 102. The forward mount may
engage the engine intermediate case (IMC) and the aft mount may
engage the engine thrust case. The aft mount reacts at least a
thrust load of the engine.
[0034] To reduce aircraft fuel burn with turbofans, it is desirable
to produce a low pressure turbine with the highest efficiency and
lowest weight possible. Further, there are considerations of small
size (especially radial size) that benefit the aerodynamic shape of
the engine cowling and allow room for packaging engine
subsystems.
[0035] FIG. 2 shows the low pressure turbine section 27 as
comprising an exemplary three blade stages 200, 202, 204. An
exemplary blade stage count is 2-6, more narrowly, 2-4, or 2-3,
3-5, or 3-4. Interspersed between the blade stages are vane stages
206 and 208. Each exemplary blade stage comprises a disk 210, 212,
and 214, respectively. A circumferential array of blades extends
from peripheries of each of the disks. Each exemplary blade
comprises an airfoil 220 extending from an inner diameter (ID)
platform 222 to an outer diameter (OD) shroud 224 (shown integral
with the airfoil
[0036] An alternative may be an unshrouded blade with a rotational
gap between the tip of the blade and a stationary blade outer air
seal (BOAS). Each exemplary shroud 224 has outboard sealing ridges
which seal with abradable seals (e.g., honeycomb) fixed to the
case. The exemplary vanes in stages 206 and 208 include airfoils
230 extending from ID platforms 232 to OD shrouds 234. The
exemplary OD shrouds 234 are directly mounted to the case. The
exemplary platforms 232 carry seals for sealing with inter-disk
knife edges protruding outwardly from inter-disk spacers which may
be separate from the adjacent disks or unitarily formed with one of
the adjacent disks.
[0037] Each exemplary disk 210, 212, 214 comprises an enlarged
central annular protuberance or "bore" 240, 242, 244 and a thinner
radial web 246, 248, 250 extending radially outboard from the bore.
The bore imparts structural strength allowing the disk to withstand
centrifugal loading which the disk would otherwise be unable to
withstand.
[0038] A turbofan engine is characterized by its bypass ratio (mass
flow ratio of air bypassing the core to air passing through the
core) and the geometric bypass area ratio (ratio of fan duct
annulus area outside/outboard of the low pressure compressor
section inlet (i.e., at location 260 in FIG. 1) to low pressure
compressor section inlet annulus area (i.e., at location 262 in
FIG. 2). High bypass engines typically have bypass area ratio of at
least four. There has been a correlation between increased bypass
area ratio and increased low pressure turbine section radius and
low pressure turbine section airfoil count. As is discussed below,
this correlation may be broken by having an engine with relatively
high bypass area ratio and relatively low turbine size.
[0039] By employing a speed reduction mechanism (e.g., a
transmission) to allow the low pressure turbine section to turn
very fast relative to the fan and by employing low pressure turbine
section design features for high speed, it is possible to create a
compact turbine module (e.g., while producing the same amount of
thrust and increasing bypass area ratio). The exemplary
transmission is a epicyclic transmission. Alternative transmissions
include composite belt transmissions, metal chain belt
transmissions, fluidic transmissions, and electric means (e.g., a
motor/generator set where the turbine turns a generator providing
electricity to an electric motor which drives the fan).
[0040] Compactness of the turbine is characterized in several ways.
Along the compressor and turbine sections, the core gaspath extends
from an inboard boundary (e.g., at blade hubs or outboard surfaces
of platforms of associated blades and vanes) to an outboard
boundary (e.g., at blade tips and inboard surfaces of blade outer
air seals for unshrouded blade tips and at inboard surfaces of OD
shrouds of shrouded blade tips and at inboard surfaces of OD
shrouds of the vanes). These boundaries may be characterized by
radii R.sub.I and R.sub.O, respectively, which vary along the
length of the engine.
[0041] For low pressure turbine radial compactness, there may be a
relatively high ratio of radial span (R.sub.O-R.sub.I) to radius
(R.sub.O or R.sub.I). Radial compactness may also be expressed in
the hub-to-tip ratio (R.sub.I:R.sub.O). These may be measured at
the maximum R.sub.O location in the low pressure turbine section.
The exemplary compact low pressure turbine section has a hub-to-tip
ratio close to about 0.5 (e.g., about 0.4-0.5 or about 0.42-0.48,
with an exemplary about 0.46).
[0042] Another characteristic of low pressure turbine radial
compactness is relative to the fan size. An exemplary fan size
measurement is the maximum tip radius R.sub.Tmax. of the fan
blades. An exemplary ratio is the maximum R.sub.O along the low
pressure turbine section to R.sub.Tmax. of the fan blades.
Exemplary values for this ratio are less than about 0.55 (e.g.,
about 0.35-55), more narrowly, less than about 0.50, or about
0.35-0.50.
[0043] To achieve compactness the designer may balance multiple
physical phenomena to arrive at a system solution as defined by the
low pressure turbine hub-to-tip ratio, the fan maximum tip radius
to low pressure turbine maximum Ro ratio, the bypass area ratio,
and the bypass area ratio to low pressure turbine airfoil count
ratio. These concerns include, but are not limited to: a)
aerodynamics within the low pressure turbine, b) low pressure
turbine blade structural design, c) low pressure turbine disk
structural design, and d) the shaft connecting the low pressure
turbine to the low pressure compressor and speed reduction device
between the low pressure compressor and fan. These physical
phenomena may be balanced in order to achieve desirable
performance, weight, and cost characteristics.
[0044] The addition of a speed reduction device between the fan and
the low pressure compressor creates a larger design space because
the speed of the low pressure turbine is decoupled from the fan.
This design space provides great design variables and new
constraints that limit feasibility of a design with respect to
physical phenomena. For example the designer can independently
change the speed and flow area of the low pressure turbine to
achieve optimal aerodynamic parameters defined by flow coefficient
(axial flow velocity/wheel speed) and work coefficient (wheel
speed/square root of work). However, this introduces structural
constraints with respect blade stresses, disk size, material
selection, etc.
[0045] In some examples, the designer can choose to make low
pressure turbine section disk bores much thicker relative to prior
art turbine bores and the bores may be at a much smaller radius
R.sub.B. This increases the amount of mass at less than a "self
sustaining radius". Another means is to choose disk materials of
greater strength than prior art such as the use of wrought powdered
metal disks to allow for extremely high centrifugal blade pulls
associated with the compactness.
[0046] Another variable in achieving compactness is to increase the
structural parameter AN.sup.2 which is the annulus area of the exit
of the low pressure turbine divided by the low pressure turbine rpm
squared at its redline or maximum speed. Relative to prior art
turbines, which are greatly constrained by fan blade tip mach
number, a very wide range of AN.sup.2 values can be selected and
optimized while accommodating such constraints as cost or a
countering, unfavorable trend in low pressure turbine section shaft
dynamics. In selecting the turbine speed (and thereby selecting the
transmission speed ratio, one has to be mindful that at too high a
gear ratio the low pressure turbine section shaft (low shaft) will
become dynamically unstable.
[0047] The higher the design speed, the higher the gear ratio will
be and the more massive the disks will become and the stronger the
low pressure turbine section disk and blade material will have to
be. All of these parameters can be varied simultaneously to change
the weight of the turbine, its efficiency, its manufacturing cost,
the degree of difficulty in packaging the low pressure turbine
section in the core cowling and its durability. This is
distinguished from a prior art direct drive configuration, where
the high bypass area ratio can only be achieved by a large low
pressure turbine section radius. Because that radius is so very
large and, although the same variables (airfoil turning, disk size,
blade materials, disk shape and materials, etc.) are theoretically
available, as a practical matter economics and engine fuel burn
considerations severely limit the designer's choice in these
parameters.
[0048] Another characteristic of low pressure turbine section size
is airfoil count (numerical count of all of the blades and vanes in
the low pressure turbine). Airfoil metal angles can be selected
such that airfoil count is low or extremely low relative to a
direct drive turbine. In known prior art engines having bypass area
ratio above 6.0 (e.g. 8.0-20), low pressure turbine sections
involve ratios of airfoil count to bypass area ratio above 190.
[0049] With the full range of selection of parameters discussed
above including, disk bore thickness, disk material, hub to tip
ratio, and R.sub.O/R.sub.Tmax., the ratio of airfoil count to
bypass area ratio may be below about 170 to as low as 10. (e.g.,
below about 150 or an exemplary about 10-170, more narrowly about
10-150). Further, in such embodiments the airfoil count may be
below about 1700, or below about 1600.
[0050] FIG. 4 shows an embodiment 600, wherein there is a fan drive
turbine 608 driving a shaft 606 to in turn drive a fan rotor 602. A
gear reduction 604 may be positioned between the fan drive turbine
608 and the fan rotor 602. This gear reduction 604 may be
structured and operate like the gear reduction disclosed above. A
compressor rotor 610 is driven by an intermediate pressure turbine
612, and a second stage compressor rotor 614 is driven by a turbine
rotor 216. A combustion section 618 is positioned intermediate the
compressor rotor 614 and the turbine section 616.
[0051] FIG. 5 shows yet another embodiment 700 wherein a fan rotor
702 and a first stage compressor 704 rotate at a common speed. The
gear reduction 706 (which may be structured as disclosed above) is
intermediate the compressor rotor 704 and a shaft 708 which is
driven by a low pressure turbine section.
[0052] One or more embodiments have been described. Nevertheless,
it will be understood that various modifications may be made. For
example, when reengineering from a baseline engine configuration,
details of the baseline may influence details of any particular
implementation. Accordingly, other embodiments are within the scope
of the following claims.
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