U.S. patent application number 15/599993 was filed with the patent office on 2017-09-07 for geared turbofan gas turbine engine architecture.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Daniel Bernard Kupratis, Frederick M. Schwarz.
Application Number | 20170254273 15/599993 |
Document ID | / |
Family ID | 48869040 |
Filed Date | 2017-09-07 |
United States Patent
Application |
20170254273 |
Kind Code |
A1 |
Kupratis; Daniel Bernard ;
et al. |
September 7, 2017 |
GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE
Abstract
A gas turbine engine typically includes a fan section, a
compressor section, a combustor section and a turbine section. A
speed reduction device such as an epicyclical gear assembly may be
utilized to drive the fan section such that the fan section may
rotate at a speed different than the turbine section so as to
increase the overall propulsive efficiency of the engine. In such
engine architectures, a shaft driven by one of the turbine sections
provides an input to the epicyclical gear assembly that drives the
fan section at a speed different than the turbine section such that
both the turbine section and the fan section can rotate at closer
to optimal speeds providing increased performance attributes and
performance by desirable combinations of the disclosed features of
the various components of the described and disclosed gas turbine
engine.
Inventors: |
Kupratis; Daniel Bernard;
(Wallingford, CT) ; Schwarz; Frederick M.;
(Glastonbury, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
48869040 |
Appl. No.: |
15/599993 |
Filed: |
May 19, 2017 |
Related U.S. Patent Documents
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
|
|
15395277 |
Dec 30, 2016 |
9695751 |
|
|
15599993 |
|
|
|
|
14789300 |
Jul 1, 2015 |
|
|
|
15395277 |
|
|
|
|
PCT/US13/23559 |
Jan 29, 2013 |
|
|
|
14789300 |
|
|
|
|
13645606 |
Oct 5, 2012 |
8935913 |
|
|
PCT/US13/23559 |
|
|
|
|
13363154 |
Jan 31, 2012 |
|
|
|
13645606 |
|
|
|
|
61653745 |
May 31, 2012 |
|
|
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2240/60 20130101;
F05D 2240/12 20130101; F01D 9/065 20130101; F05D 2240/35 20130101;
F04D 27/009 20130101; Y02T 50/60 20130101; F01D 9/02 20130101; F02C
9/18 20130101; F05D 2260/4031 20130101; F02C 7/36 20130101; F02K
3/06 20130101; Y02T 50/671 20130101; F02K 3/04 20130101; F02C 3/36
20130101; F02K 3/072 20130101; F05D 2220/323 20130101; F02C 7/06
20130101; F02C 3/04 20130101; F04D 29/325 20130101; F05D 2260/40311
20130101; F05D 2220/32 20130101; F01D 5/06 20130101; F02C 3/107
20130101; F02K 1/78 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02C 3/04 20060101 F02C003/04; F02C 9/18 20060101
F02C009/18; F04D 27/00 20060101 F04D027/00; F04D 29/32 20060101
F04D029/32; F02C 3/107 20060101 F02C003/107; F02K 1/78 20060101
F02K001/78; F01D 9/02 20060101 F01D009/02; F02K 3/06 20060101
F02K003/06; F01D 5/06 20060101 F01D005/06 |
Claims
1. A gas turbine engine comprising: a fan including a plurality of
fan blades rotatable about an axis; a compressor section; a
combustor in fluid communication with the compressor section; a
turbine section in fluid communication with the combustor, the
turbine section including a fan drive turbine and a second turbine,
wherein the second turbine is disposed forward of the fan drive
turbine and the fan drive turbine includes a plurality of turbine
rotors with a ratio between the number of fan blades and the number
of fan drive turbine rotors is between 2.5 and 8.5; and a speed
change system configured to be driven by the fan drive turbine to
rotate the fan about the axis; and a power density at Sea Level
Takeoff greater than 1.5 lbf/in.sup.3 and less than or equal to 5.5
lbf/in.sup.3 and defined as thrust in lbf measured by a volume of
the turbine section in in.sup.3 measured between an inlet of a
first turbine vane in said second turbine to an exit of a last
rotating airfoil stage in said fan drive turbine.
2. The engine as recited in claim 1, wherein the speed change
system comprises a gear reduction having a gear ratio greater than
about 2.3.
3. The engine as recited in claim 2, wherein the fan drive turbine
has from three to six stages.
4. The engine as recited in claim 3, wherein the fan has less than
18 of said fan blades and the second turbine has two stages.
5. The engine as recited in claim 4, further comprising a frame
structure positioned between the fan drive turbine and the second
turbine, and a plurality of vanes associated with the frame
structure, and a flow path through said frame structure being part
of the volume of the turbine section.
6. The engine as recited in claim 3, wherein the fan delivers a
portion of air into a bypass duct, and a bypass ratio being defined
as the portion of air delivered into the bypass duct divided by an
amount of air delivered into the compressor section, with the
bypass ratio being greater than 10.0.
7. The engine as recited in claim 6, further comprising a low fan
pressure ratio across the plurality of fan blades alone of less
than 1.45.
8. The engine as recited in claim 7, wherein the fan drive turbine
includes an inlet having an inlet pressure, an outlet that is prior
to any exhaust nozzle and having an outlet pressure, and a pressure
ratio defined as a ratio of the inlet pressure to the outlet
pressure, and wherein the pressure ratio of the fan drive turbine
is greater than about 5.
9. The engine as recited in claim 8, wherein the fan has less than
18 of said fan blades and the second turbine has two stages.
10. The engine as recited in claim 9, further comprising 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 at an
ambient temperature divided by [(Tram .degree. R)/518.7).sup.0.5],
where T represents the ambient temperature in degrees Rankine.
11. The engine as recited in claim 8, wherein the fan drive turbine
has a first exit area and rotates at a first speed, the second
turbine section has a second exit area and rotates at a second
speed, which is faster than the first speed, the first and second
speeds being redline speeds, a first performance quantity is
defined as the product of the first speed squared and the first
area, a second performance quantity is defined as the product of
the second speed squared and the second area, and a performance
ratio of the first performance quantity to the second performance
quantity is greater than about 0.5.
12. The engine as recited in claim 11, wherein the performance
ratio is above or equal to about 0.8.
13. The engine as recited in claim 12, wherein the performance
ratio is above or equal to about 1.0.
14. The engine as recited in claim 13, wherein the first
performance quantity is above or equal to about 4.
15. The engine as recited in claim 13, wherein the performance
ratio is less than or equal to 1.5.
16. The engine as recited in claim 13, wherein the power density is
greater than or equal to 3.0 lbf/in.sup.3.
17. The engine as recited in claim 13, wherein the fan has less
than 18 fan blades, and the second turbine has two stages.
18. The engine as recited in claim 17, further comprising 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 at an
ambient temperature divided by [(Tram .degree. R)/518.7).sup.0.5],
where T represents the ambient temperature in degrees Rankine.
19. A gas turbine engine comprising: a fan including a plurality of
fan blades rotatable about an axis; a compressor section; a
combustor in fluid communication with the compressor section; a
turbine section in fluid communication with the combustor, the
turbine section including a fan drive turbine and a second turbine,
wherein the second turbine is disposed forward of the fan drive
turbine and the fan drive turbine includes a plurality of turbine
rotors with a ratio between the number of fan blades and the number
of fan drive turbine rotors is between 2.5 and 8.5; and a speed
change system having a gear reduction with a gear ratio greater
than 2.3, the speed change system configured to be driven by the
fan drive turbine to rotate the fan about the axis; and a power
density at Sea Level Takeoff greater than or equal to 1.5
lbf/in.sup.3 and defined as thrust in lbf divided by a volume of
the turbine section in inch.sup.3 measured between an inlet of a
first turbine vane in said second turbine to an exit of a last
rotating airfoil stage in said fan drive turbine.
20. The engine as recited in claim 19, wherein the fan drive
turbine has from three to six stages.
21. The engine as recited in claim 20, further comprising a low fan
pressure ratio across a fan blade alone of less than 1.45, and
wherein the fan delivers a portion of air into a bypass duct, and a
bypass ratio being defined as the portion of air delivered into the
bypass duct divided by an amount of air delivered into the
compressor section, with the bypass ratio being greater than
10.0.
22. The engine as recited in claim 21, wherein the fan drive
turbine includes an inlet having an inlet pressure, an outlet that
is prior to any exhaust nozzle and having an outlet pressure, and a
pressure ratio defined as a ratio of the inlet pressure to the
outlet pressure, and wherein the pressure ratio of the fan drive
turbine is greater than about 5.
23. The engine as recited in claim 22, further comprising 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 at an
ambient temperature divided by [(Tram .degree. R)/518.7).sup.0.5],
where T represents the ambient temperature in degrees Rankine.
24. The engine as recited in claim 23, wherein the fan drive
turbine has a first exit area and rotates at a first speed, the
second turbine section has a second exit area and rotates at a
second speed, which is faster than the first speed, the first and
second speeds being redline speeds, a first performance quantity is
defined as the product of the first speed squared and the first
area, a second performance quantity is defined as the product of
the second speed squared and the second area, and a performance
ratio of the first performance quantity to the second performance
quantity is greater than about 0.8.
25. The engine as recited in claim 24, wherein the fan has less
than 18 fan blades and the second turbine has two stages.
26. A gas turbine engine comprising: a fan including a plurality of
fan blades rotatable about an axis, wherein the plurality of fan
blades is less than 18 fan blades; a low fan pressure ratio across
a fan blade alone of less than 1.45, and wherein the fan is
configured to deliver a portion of air into a bypass duct, and a
bypass ratio being defined as the portion of air delivered into the
bypass duct divided by the amount of air delivered into the
compressor section, with the bypass ratio being greater than 10.0;
a compressor section; a combustor in fluid communication with the
compressor section; a turbine section in fluid communication with
the combustor, the turbine section including a fan drive turbine
and a second turbine, wherein the second turbine is a two stage
turbine and is disposed forward of the fan drive turbine, and the
fan drive turbine includes a plurality of turbine rotors with a
ratio between the number of fan blades and the number of fan drive
turbine rotors is between 2.5 and 8.5; a low speed spool associated
with the fan drive turbine and including an inner shaft, and a high
speed spool associated with the second turbine and including an
outer shaft, the inner shaft and outer shaft being concentric; a
planetary gearbox having a sun gear, a plurality of planet gears
configured to rotate and spaced apart by a carrier configured to
rotate in a direction common to the sun gear, a non-rotating ring
gear, and a gear reduction with a gear ratio greater than 2.3, the
gearbox configured to be driven by the fan drive turbine to rotate
the fan about the axis; and a power density at Sea Level Takeoff
greater than or equal to 1.5 lbf/in.sup.3 and defined as thrust in
lbf and a volume of the turbine section in inch.sup.3 and defined
as thrust in lbf divided by a volume of the turbine section in
inch.sup.3 measured between an inlet of a first turbine vane in
said second turbine to an exit of a last rotating airfoil stage in
said fan drive turbine.
27. The engine as recited in claim 26, further comprising 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 at an
ambient temperature divided by [(Tram .degree. R)/518.7).sup.0.5],
where T represents the ambient temperature in degrees Rankine, and
wherein the fan drive turbine includes an inlet having an inlet
pressure, an outlet that is prior to any exhaust nozzle and having
an outlet pressure, and a pressure ratio defined as a ratio of the
inlet pressure to the outlet pressure, and wherein the pressure
ratio of the fan drive turbine is greater than about 5.
28. The engine as recited in claim 26, wherein the fan drive
turbine has a first exit area and rotates at a first speed, the
second turbine section has a second exit area and rotates at a
second speed, which is faster than the first speed, the first and
second speeds being redline speeds, a first performance quantity is
defined as the product of the first speed squared and the first
area, a second performance quantity is defined as the product of
the second speed squared and the second area, and a performance
ratio of the first performance quantity to the second performance
quantity is greater than about 0.8.
29. The engine as recited in claim 28, further comprising 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 at an
ambient temperature divided by [(Tram .degree. R)/518.7).sup.0.5],
where T represents the ambient temperature in degrees Rankine, and
wherein the fan drive turbine includes an inlet having an inlet
pressure, an outlet that is prior to any exhaust nozzle and having
an outlet pressure, and a pressure ratio defined as a ratio of the
inlet pressure to the outlet pressure, and wherein the pressure
ratio of the fan drive turbine is greater than about 5.
30. The engine as recited in claim 26, wherein the fan drive
turbine includes an inlet having an inlet pressure, an outlet that
is prior to any exhaust nozzle and having an outlet pressure, and a
pressure ratio defined as a ratio of the inlet pressure to the
outlet pressure, and wherein the pressure ratio of the fan drive
turbine is greater than about 5.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. application Ser.
No. 15/395,277, filed Dec. 30, 2016, which is a continuation of
U.S. application Ser. No. 14/789,300, filed Jul. 1, 2015, which is
a continuation-in-part of International Application No.
PCT/US13/23559 filed Jan. 29, 2013, which claims priority to U.S.
application Ser. No. 13/645,606 filed Oct. 5, 2012, now U.S. Pat.
No. 8,935,913 granted Jan. 20, 2015, which was a continuation in
part of U.S. application Ser. No. 13/363,154 filed on Jan. 31, 2012
and claims priority to U.S. Provisional Application No. 61/653,745
filed on May 31, 2012.
BACKGROUND
[0002] A gas turbine engine typically includes a fan section, a
compressor section, a combustor section and a turbine section. Air
entering the compressor section is compressed and delivered into
the combustion section where it is mixed with fuel and ignited to
generate a high-speed exhaust gas flow. The high-speed exhaust gas
flow expands through the turbine section to drive the compressor
and the fan section. The compressor section typically includes low
and high pressure compressors, and the turbine section includes low
and high pressure turbines.
[0003] The high pressure turbine drives the high pressure
compressor through an outer shaft to form a high spool, and the low
pressure turbine drives the low pressure compressor through an
inner shaft to form a low spool. The inner shaft may also drive the
fan section. A direct drive gas turbine engine includes a fan
section driven by the inner shaft such that the low pressure
compressor, low pressure turbine and fan section rotate at a common
speed in a common direction.
[0004] A speed reduction device such as an epicyclical gear
assembly may be utilized to drive the fan section such that the fan
section may rotate at a speed different than the turbine section so
as to increase the overall propulsive efficiency of the engine. In
such engine architectures, a shaft driven by one of the turbine
sections provides an input to the epicyclical gear assembly that
drives the fan section at a speed different than the turbine
section such that both the turbine section and the fan section can
rotate at closer to optimal speeds.
[0005] Although geared architectures have improved propulsive
efficiency, turbine engine manufacturers continue to seek further
improvements to engine performance including improvements to
thermal, transfer and propulsive efficiencies.
SUMMARY
[0006] A gas turbine engine according to an exemplary embodiment of
this disclosure, among other possible things includes a fan
including a plurality of fan blades rotatable about an axis, a
compressor section, a combustor in fluid communication with the
compressor section, and a turbine section in fluid communication
with the combustor. The turbine section includes a fan turbine and
a second turbine. The second turbine is disposed forward of the fan
drive turbine. The fan drive turbine includes a plurality of
turbine rotors with a ratio between the number of fan blades and
the number of fan drive turbine rotors is greater than about 2.5. A
speed change system is driven by the fan drive turbine for rotating
the fan about the axis. The fan drive turbine has a first exit area
and rotates at a first speed. The second turbine section has a
second exit area and rotates at a second speed, which is faster
than the first speed. A first performance quantity is defined as
the product of the first speed squared and the first area. A second
performance quantity is defined as the product of the second speed
squared and the second area. A performance ratio of the first
performance quantity to the second performance quantity is between
about 0.5 and about 1.5.
[0007] In a further embodiment of the foregoing engine, the
performance ratio is above or equal to about 0.8.
[0008] In a further embodiment of any of the foregoing engines, the
first performance quantity is above or equal to about 4.
[0009] In a further embodiment of any of the foregoing engines, the
speed change system includes a gearbox. The fan and the fan drive
turbine both rotate in a first direction about the axis and the
second turbine section rotates in a second direction opposite the
first direction.
[0010] In a further embodiment of any of the foregoing engines, the
speed change system includes a gearbox. The fan, the fan drive
turbine, and the second turbine section all rotate in a first
direction about the axis.
[0011] In a further embodiment of any of the foregoing engines, the
speed change system includes a gearbox. The fan and the second
turbine both rotate in a first direction about the axis and the fan
drive turbine rotates in a second direction opposite the first
direction.
[0012] In a further embodiment of any of the foregoing engines, the
speed change system includes a gearbox. The fan is rotatable in a
first direction and the fan drive turbine, and the second turbine
section rotate in a second direction opposite the first direction
about the axis.
[0013] In a further embodiment of any of the foregoing engines, the
speed change system includes a gear reduction having a gear ratio
greater than about 2.3.
[0014] In a further embodiment of any of the foregoing engines, the
fan delivers a portion of air into a bypass duct. A bypass ratio
being defined as the portion of air delivered into the bypass duct
divided by the amount of air delivered into the compressor section,
with the bypass ratio being greater than about 6.0.
[0015] In a further embodiment of any of the foregoing engines, the
bypass ratio is greater than about 10.0.
[0016] In a further embodiment of any of the foregoing engines, a
fan pressure ratio across the fan is less than about 1.5.
[0017] In a further embodiment of any of the foregoing engines, the
fan has about 26 or fewer blades.
[0018] In a further embodiment of any of the foregoing engines, the
fan drive turbine section has up to 6 stages.
[0019] In a further embodiment of any of the foregoing engines, the
ratio between the number of fan blades and the number of fan drive
turbine rotors is less than about 8.5.
[0020] In a further embodiment of any of the foregoing engines, a
pressure ratio across the fan drive turbine is greater than about
5:1.
[0021] In a further embodiment of any of the foregoing engines,
includes a power density greater than about 1.5 lbf/in.sup.3 and
less than or equal to about 5.5 lbf/in.sup.3.
[0022] In a further embodiment of any of the foregoing engines, the
fan drive turbine includes a first aft rotor attached to a first
shaft. The second turbine includes a second aft rotor attached to a
second shaft. A first bearing assembly is disposed axially aft of a
first connection between the first aft rotor and the first shaft. A
second bearing assembly is disposed axially aft of a second
connection between the second aft rotor and the second shaft.
[0023] In a further embodiment of any of the foregoing engines, the
fan drive turbine includes a first aft rotor attached to a first
shaft. The second turbine includes a second aft rotor attached to a
second shaft. A first bearing assembly is disposed axially aft of a
first connection between the first aft rotor and the first shaft. A
second bearing assembly is disposed axially forward of a second
connection between the second aft rotor and the second shaft.
[0024] In a further embodiment of any of the foregoing engines, the
fan drive turbine includes a first aft rotor attached to a first
shaft. The second turbine includes a second aft rotor attached to a
second shaft. A first bearing assembly is disposed axially aft of a
first connection between the first aft rotor and the first shaft. A
second bearing assembly is disposed within the annular space
defined between the first shaft and the second shaft.
[0025] In a further embodiment of any of the foregoing engines, the
fan drive turbine includes a first aft rotor attached to a first
shaft. The second turbine includes a second aft rotor attached to a
second shaft. A first bearing assembly is disposed axially forward
of a first connection between the first aft rotor and the first
shaft. A second bearing assembly is disposed axially aft of a
second connection between the second aft rotor and the second
shaft.
[0026] 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.
[0027] 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
[0028] FIG. 1 is a schematic view of an example gas turbine
engine.
[0029] FIG. 2 is a schematic view indicating relative rotation
between sections of an example gas turbine engine.
[0030] FIG. 3 is another schematic view indicating relative
rotation between sections of an example gas turbine engine.
[0031] FIG. 4 is another schematic view indicating relative
rotation between sections of an example gas turbine engine.
[0032] FIG. 5 is another a schematic view indicating relative
rotation between sections of an example gas turbine engine.
[0033] FIG. 6 is a schematic view of a bearing configuration
supporting rotation of example high and low spools of the example
gas turbine engine.
[0034] FIG. 7 is another schematic view of a bearing configuration
supporting rotation of example high and low spools of the example
gas turbine engine.
[0035] FIG. 8A is another schematic view of a bearing configuration
supporting rotation of example high and low spools of the example
gas turbine engine.
[0036] FIG. 8B is an enlarged view of the example bearing
configuration shown in FIG. 8A.
[0037] FIG. 9 is another schematic view of a bearing configuration
supporting rotation of example high and low spools of the example
gas turbine engine.
[0038] FIG. 10 is a schematic view of an example compact turbine
section.
[0039] FIG. 11 is a schematic cross-section of example stages for
the disclosed example gas turbine engine.
[0040] FIG. 12 is a schematic view an example turbine rotor
perpendicular to the axis or rotation.
[0041] FIG. 13 is another embodiment of an example gas turbine
engine for use with the present invention.
[0042] FIG. 14 is yet another embodiment of an example gas turbine
engine for use with the present invention.
DETAILED DESCRIPTION
[0043] FIG. 1 schematically illustrates an example gas turbine
engine 20 that includes a fan section 22, a compressor section 24,
a combustor section 26 and a turbine section 28. Alternative
engines might include an augmenter section (not shown) among other
systems or features. The fan section 22 drives air along a bypass
flow path B while the compressor section 24 draws air in along a
core flow path C where air is compressed and communicated to a
combustor section 26. In the combustor section 26, air is mixed
with fuel and ignited to generate a high pressure exhaust gas
stream that expands through the turbine section 28 where energy is
extracted and utilized to drive the fan section 22 and the
compressor section 24.
[0044] Although the disclosed non-limiting embodiment depicts a
turbofan gas turbine engine, 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; for
example a turbine engine including a three-spool architecture in
which three spools concentrically rotate about a common axis such
that a low spool enables a low pressure turbine to drive a fan via
a gearbox, an intermediate spool enables an intermediate pressure
turbine to drive a first compressor of the compressor section, and
a high spool enables a high pressure turbine to drive a high
pressure compressor of the compressor section.
[0045] The example 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 connects a fan 42 and a low pressure (or first) compressor
section 44 to a low pressure (or first) turbine section 46. The
inner shaft 40 drives the fan 42 through a speed change device,
such as 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 (or second)
compressor section 52 and a high pressure (or second) turbine
section 54. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via the bearing systems 38 about the engine
central longitudinal axis A.
[0047] A combustor 56 is arranged between the high pressure
compressor 52 and the high pressure turbine 54. In one example, the
high pressure turbine 54 includes at least two stages to provide a
double stage high pressure turbine 54. In another example, the high
pressure turbine 54 includes only a single stage. As used herein, a
"high pressure" compressor or turbine experiences a higher pressure
than a corresponding "low pressure" compressor or turbine.
[0048] The example low pressure turbine 46 has a pressure ratio
that is greater than about 5. The pressure ratio of the example low
pressure turbine 46 is measured prior to an inlet of the low
pressure turbine 46 as related to the pressure measured at the
outlet of the low pressure turbine 46 prior to an exhaust
nozzle.
[0049] A mid-turbine frame 58 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 58 further supports
bearing systems 38 in the turbine section 28 as well as setting
airflow entering the low pressure turbine 46.
[0050] The core airflow C is compressed by the low pressure
compressor 44 then by the high pressure compressor 52 mixed with
fuel and ignited in the combustor 56 to produce high speed exhaust
gases that are then expanded through the high pressure turbine 54
and low pressure turbine 46. The mid-turbine frame 58 includes
vanes 60, which are in the core airflow path and function as an
inlet guide vane for the low pressure turbine 46. Utilizing the
vane 60 of the mid-turbine frame 58 as the inlet guide vane for low
pressure turbine 46 decreases the length of the low pressure
turbine 46 without increasing the axial length of the mid-turbine
frame 58. Reducing or eliminating the number of vanes in the low
pressure turbine 46 shortens the axial length of the turbine
section 28. Thus, the compactness of the gas turbine engine 20 is
increased and a higher power density may be achieved.
[0051] The disclosed gas turbine engine 20 in one example is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 includes a bypass ratio greater than about six
(6), with an example embodiment being greater than about ten (10).
The example geared architecture 48 is an epicyclical gear train,
such as a planetary gear system, star gear system or other known
gear system, with a gear reduction ratio of greater than about
2.3.
[0052] In one disclosed embodiment, the gas turbine engine 20
includes a bypass ratio greater than about ten (10:1) and the fan
diameter is significantly larger than an outer diameter of the low
pressure compressor 44. It should be understood, however, that the
above parameters are only exemplary of one embodiment of a gas
turbine engine including a geared architecture and that the present
disclosure is applicable to other gas turbine engines.
[0053] 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
cruise fuel consumption relative to the thrust it produces--also
known as "bucket cruise Thrust Specific Fuel Consumption
('TSFC)"--is the industry standard parameter of pound-mass (lbm) of
fuel per hour being burned divided by pound-force (lbf) of thrust
the engine produces at that minimum bucket cruise point.
[0054] "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. In another
non-limiting embodiment the low fan pressure ratio is less than
about 1.45.
[0055] "Low corrected fan tip speed" is the actual fan tip speed in
ft/sec divided by an industry standard temperature correction of
[(Tram .degree. R)/518.71).sup.0.5]. The "Low corrected fan tip
speed", as disclosed herein according to one non-limiting
embodiment, is less than about 1150 ft/second.
[0056] The example gas turbine engine includes the fan 42 that
comprises in one non-limiting embodiment less than about 26 fan
blades. In another non-limiting embodiment, the fan section 22
includes less than about 18 fan blades. Moreover, in one disclosed
embodiment the low pressure turbine 46 includes no more than about
6 turbine stages schematically indicated at 34. In another
non-limiting example embodiment the low pressure turbine 46
includes about 3 or more turbine stages. A ratio between the number
of fan blades 42 and the number of low pressure turbine stages is
between about 2.5 and about 8.5. The example low pressure turbine
46 provides the driving power to rotate the fan section 22 and
therefore the relationship between the number of turbine stages 34
in the low pressure turbine 46 and the number of blades 42 in the
fan section 22 disclose an example gas turbine engine 20 with
increased power transfer efficiency.
[0057] Increased power transfer efficiency is provided due in part
to the increased use of improved turbine blade materials and
manufacturing methods such as directionally solidified castings,
and single crystal materials that enable increased turbine speed
and a reduced number of stages. Moreover, the example low pressure
turbine 46 includes improved turbine disks configurations that
further enable desired durability at the higher turbine speeds.
[0058] Referring to FIGS. 2 and 3, an example disclosed speed
change device is an epicyclical gearbox of a planet type, where the
input is to the center "sun" gear 62. Planet gears 64 (only one
shown) around the sun gear 62 rotate and are spaced apart by a
carrier 68 that rotates in a direction common to the sun gear 62. A
ring gear 66, which is non-rotatably fixed to the engine static
casing 36 (shown in FIG. 1), contains the entire gear assembly. The
fan 42 is attached to and driven by the carrier 68 such that the
direction of rotation of the fan 42 is the same as the direction of
rotation of the carrier 68 that, in turn, is the same as the
direction of rotation of the input sun gear 62.
[0059] In the following figures nomenclature is utilized to define
the relative rotations between the various sections of the gas
turbine engine 20. The fan section is shown with a "+" sign
indicating rotation in a first direction. Rotations relative to the
fan section 22 of other features of the gas turbine engine are
further indicated by the use of either a "+" sign or a "-" sign.
The "-" sign indicates a rotation that is counter to that of any
component indicated with a "+" sign.
[0060] Moreover, the term fan drive turbine is utilized to indicate
the turbine that provides the driving power for rotating the blades
42 of the fan section 22. Further, the term "second turbine" is
utilized to indicate the turbine before the fan drive turbine that
is not utilized to drive the fan 42. In this disclosed example, the
fan drive turbine is the low pressure turbine 46, and the second
turbine is the high pressure turbine 54. However, it should be
understood that other turbine section configurations that include
more than the shown high and low pressure turbines 54, 46 are
within the contemplation of this disclosure. For example, a three
spool engine configuration may include an intermediate turbine (not
shown) utilized to drive the fan section 22 and is within the
contemplation of this disclosure.
[0061] In one disclosed example embodiment (FIG. 2) the fan drive
turbine is the low pressure turbine 46 and therefore the fan
section 22 and low pressure turbine 46 rotate in a common direction
as indicated by the common "+" sign indicating rotation of both the
fan 42 and the low pressure turbine 46. Moreover in this example,
the high pressure turbine 54 or second turbine rotates in a
direction common with the fan drive turbine 46. In another example
shown in FIG. 3, the high pressure turbine 54 or second turbine
rotates in a direction opposite the fan drive turbine (low pressure
turbine 46) and the fan 42.
[0062] Counter rotating the low pressure compressor 44 and the low
pressure turbine 46 relative to the high pressure compressor 52 and
the high pressure turbine 54 provides certain efficient aerodynamic
conditions in the turbine section 28 as the generated high speed
exhaust gas flow moves from the high pressure turbine 54 to the low
pressure turbine 46. The relative rotations in the compressor and
turbine sections provide approximately the desired airflow angles
between the sections, which improves overall efficiency in the
turbine section 28, and provides a reduction in overall weight of
the turbine section 28 by reducing or eliminating airfoils or an
entire row of vanes.
[0063] Referring to FIGS. 4 and 5, another example disclosed speed
change device is an epicyclical gearbox referred to as a star type
gearbox, where the input is to the center "sun" gear 62. Star gears
65 (only one shown) around the sun gear 62 rotate in a fixed
position around the sun gear and are spaced apart by a carrier 68
that is fixed to a static casing 36 (best shown in FIG. 1). A ring
gear 66 that is free to rotate contains the entire gear assembly.
The fan 42 is attached to and driven by the ring gear 66 such that
the direction of rotation of the fan 42 is opposite the direction
of rotation of the input sun gear 62. Accordingly, the low pressure
compressor 44 and the low pressure turbine 46 rotate in a direction
opposite rotation of the fan 42.
[0064] In one disclosed example embodiment shown in FIG. 4, the fan
drive turbine is the low pressure turbine 46 and therefore the fan
42 rotates in a direction opposite that of the low pressure turbine
46 and the low pressure compressor 44. Moreover in this example the
high spool 32 including the high pressure turbine 54 and the high
pressure compressor 52 rotate in a direction counter to the fan 42
and common with the low spool 30 including the low pressure
compressor 44 and the fan drive turbine 46.
[0065] In another example gas turbine engine shown in FIG. 5, the
high pressure or second turbine 54 rotates in a direction common
with the fan 42 and counter to the low spool 30 including the low
pressure compressor 44 and the fan drive turbine 46.
[0066] Referring to FIG. 6, the bearing assemblies near the forward
end of the shafts in the engine at locations 70 and 72, which
bearings support rotation of the inner shaft 40 and the outer shaft
50, counter net thrust forces in a direction parallel to the axis A
that are generated by the rearward load of low pressure turbine 46
and the high pressure turbine 54, minus the high pressure
compressor 52 and the low pressure compressor 44, which also
contribute to the thrust forces acting on the corresponding low
spool 30 and the high spool 32.
[0067] In this example embodiment, a first forward bearing assembly
70 is supported on a portion of the static structure schematically
shown at 36 and supports a forward end of the inner shaft 40. The
example first forward bearing assembly 70 is a thrust bearing and
controls movement of the inner shaft 40 and thereby the low spool
30 in an axial direction. A second forward bearing assembly 72 is
supported by the static structure 36 to support rotation of the
high spool 32 and substantially prevent movement along in an axial
direction of the outer shaft 50. The first forward bearing assembly
70 is mounted to support the inner shaft 40 at a point forward of a
connection 88 of a low pressure compressor rotor 90. The second
forward bearing assembly 72 is mounted forward of a connection
referred to as a hub 92 between a high pressure compressor rotor 94
and the outer shaft 50. A first aft bearing assembly 74 supports
the aft portion of the inner shaft 40. The first aft bearing
assembly 74 is a roller bearing and supports rotation, but does not
provide resistance to movement of the shaft 40 in the axial
direction. Instead, the aft bearing 74 allows the shaft 40 to
expand thermally between its location and the bearing 72. The
example first aft bearing assembly 74 is disposed aft of a
connection hub 80 between a low pressure turbine rotor 78 and the
inner shaft 40. A second aft bearing assembly 76 supports the aft
portion of the outer shaft 50. The example second aft bearing
assembly 76 is a roller bearing and is supported by a corresponding
static structure 36 through the mid turbine frame 58 which
transfers the radial load of the shaft across the turbine flow path
to ground 36. The second aft bearing assembly 76 supports the outer
shaft 50 and thereby the high spool 32 at a point aft of a
connection hub 84 between a high pressure turbine rotor 82 and the
outer shaft 50.
[0068] In this disclosed example, the first and second forward
bearing assemblies 70, 72 and the first and second aft bearing
assemblies 74, 76 are supported to the outside of either the
corresponding compressor or turbine connection hubs 80, 88 to
provide a straddle support configuration of the corresponding inner
shaft 40 and outer shaft 50. The straddle support of the inner
shaft 40 and the outer shaft 50 provide a support and stiffness
desired for operation of the gas turbine engine 20.
[0069] Referring to FIG. 7, another example shaft support
configuration includes the first and second forward bearing
assemblies 70, 72 disposed to support the forward portion of the
corresponding inner shaft 40 and outer shaft 50. The first aft
bearing 74 is disposed aft of the connection 80 between the rotor
78 and the inner shaft 40. The first aft bearing 74 is a roller
bearing and supports the inner shaft 40 in a straddle
configuration. The straddle configuration can require additional
length of the inner shaft 40 and therefore an alternate
configuration referred to as an overhung configuration can be
utilized. In this example the outer shaft 50 is supported by the
second aft bearing assembly 76 that is disposed forward of the
connection 84 between the high pressure turbine rotor 82 and the
outer shaft 50. Accordingly, the connection hub 84 of the high
pressure turbine rotor 82 to the outer shaft 50 is overhung aft of
the bearing assembly 76. This positioning of the second aft bearing
76 in an overhung orientation potentially provides for a reduced
length of the outer shaft 50.
[0070] Moreover the positioning of the aft bearing 76 may also
eliminate the need for other support structures such as the mid
turbine frame 58 as both the high pressure turbine 54 is supported
at the bearing assembly 76 and the low pressure turbine 46 is
supported by the bearing assembly 74. Optionally the mid turbine
frame strut 58 can provide an optional roller bearing 74A which can
be added to reduce vibratory modes of the inner shaft 40.
[0071] Referring to FIG. 8A and 8B, another example shaft support
configuration includes the first and second forward bearing
assemblies 70, 72 disposed to support corresponding forward
portions of each of the inner shaft 40 and the outer shaft 50. The
first aft bearing 74 provides support of the outer shaft 40 at a
location aft of the connection 80 in a straddle mount
configuration. In this example, the aft portion of the outer shaft
50 is supported by a roller bearing assembly 86 supported within a
space 96 defined between an outer surface of the inner shaft 40 and
an inner surface of the outer shaft 50.
[0072] The roller bearing assembly 86 supports the aft portion of
the outer shaft 50 on the inner shaft 40. The use of the roller
bearing assembly 86 to support the outer shaft 50 eliminates the
requirements for support structures that lead back to the static
structure 36 through the mid turbine frame 58. Moreover, the
example bearing assembly 86 can provide both a reduced shaft
length, and support of the outer shaft 50 at a position
substantially in axial alignment with the connection hub 84 for the
high pressure turbine rotor 82 and the outer shaft 50. As
appreciated, the bearing assembly 86 is positioned aft of the hub
82 and is supported through the rearmost section of shaft 50.
Referring to FIG. 9, another example shaft support configuration
includes the first and second forward bearing assemblies 70, 72
disposed to support corresponding forward portions of each of the
inner shaft 40 and the outer shaft 50. The first aft bearing
assembly 74 is supported at a point along the inner shaft 40
forward of the connection 80 between the low pressure turbine rotor
78 and the inner shaft 40.
[0073] Positioning of the first aft bearing 74 forward of the
connection 80 can be utilized to reduce the overall length of the
engine 20. Moreover, positioning of the first aft bearing assembly
74 forward of the connection 80 provides for support through the
mid turbine frame 58 to the static structure 36. Furthermore, in
this example the second aft bearing assembly 76 is deployed in a
straddle mount configuration aft of the connection 84 between the
outer shaft 50 and the rotor 82. Accordingly, in this example, both
the first and second aft bearing assemblies 74, 76 share a common
support structure to the static outer structure 36. As appreciated,
such a common support feature provides for a less complex engine
construction along with reducing the overall length of the engine.
Moreover, the reduction or required support structures will reduce
overall weight to provide a further improvement in aircraft fuel
burn efficiency.
[0074] Referring to FIG. 10, a portion of the example turbine
section 28 is shown and includes the low pressure turbine 46 and
the high pressure turbine 54 with the mid turbine frame 58 disposed
between an outlet of the high pressure turbine and the low pressure
turbine. The mid turbine frame 58 and vane 60 are positioned to be
upstream of the first stage 98 of the low pressure turbine 46.
While a single vane 60 is illustrated, it should be understood
these would be plural vanes 60 spaced circumferentially. The vane
60 redirects the flow downstream of the high pressure turbine 54 as
it approaches the first stage 98 of the low pressure turbine 46. As
can be appreciated, it is desirable to improve efficiency to have
flow between the high pressure turbine 54 and the low pressure
turbine 46 redirected by the vane 60 such that the flow of
expanding gases is aligned as desired when entering the low
pressure turbine 46. Therefore vane 60 may be an actual airfoil
with camber and turning, that aligns the airflow as desired into
the low pressure turbine 46.
[0075] By incorporating a true air-turning vane 60 into the mid
turbine frame 58, rather than a streamlined strut and a stator vane
row after the strut, the overall length and volume of the combined
turbine sections 46, 54 is reduced because the vane 60 serves
several functions including streamlining the mid turbine frame 58,
protecting any static structure and any oil tubes servicing a
bearing assembly from exposure to heat, and turning the flow
entering the low pressure turbine 46 such that it enters the
rotating airfoil 100 at a desired flow angle. Further, by
incorporating these features together, the overall assembly and
arrangement of the turbine section 28 is reduced in volume.
[0076] The above features achieve a more or less compact turbine
section volume relative to the prior art including both high and
low pressure turbines 54, 46. Moreover, in one example, the
materials for forming the low pressure turbine 46 can be improved
to provide for a reduced volume. Such materials may include, for
example, materials with increased thermal and mechanical
capabilities to accommodate potentially increased stresses induced
by operating the low pressure turbine 46 at the increased speed.
Furthermore, the elevated speeds and increased operating
temperatures at the entrance to the low pressure turbine 46 enables
the low pressure turbine 46 to transfer a greater amount of energy,
more efficiently to drive both a larger diameter fan 42 through the
geared architecture 48 and an increase in compressor work performed
by the low pressure compressor 44.
[0077] Alternatively, lower priced materials can be utilized in
combination with cooling features that compensate for increased
temperatures within the low pressure turbine 46. In three exemplary
embodiments a first rotating blade 100 of the low pressure turbine
46 can be a directionally solidified casting blade, a single
crystal casting blade or a hollow, internally cooled blade. The
improved material and thermal properties of the example turbine
blade material provide for operation at increased temperatures and
speeds, that in turn provide increased efficiencies at each stage
that thereby provide for use of a reduced number of low pressure
turbine stages. The reduced number of low pressure turbine stages
in turn provide for an overall turbine volume that is reduced, and
that accommodates desired increases in low pressure turbine
speed.
[0078] The reduced stages and reduced volume provide improve engine
efficiency and aircraft fuel burn because overall weight is less.
In addition, as there are fewer blade rows, there are: fewer
leakage paths at the tips of the blades; fewer leakage paths at the
inner air seals of vanes; and reduced losses through the rotor
stages.
[0079] The example disclosed compact turbine section includes a
power density, which may be defined as thrust in pounds force (lbf)
produced divided by the volume of the entire turbine section 28.
The volume of the turbine section 28 may be defined by an inlet 102
of a first turbine vane 104 in the high pressure turbine 54 to the
exit 106 of the last rotating airfoil 108 in the low pressure
turbine 46, and may be expressed in cubic inches. The static thrust
at the engine's flat rated Sea Level Takeoff condition divided by a
turbine section volume is defined as power density and a greater
power density may be desirable for reduced engine weight. The sea
level take-off flat-rated static thrust may be defined in
pounds-force (lbf), while the volume may be the volume from the
annular inlet 102 of the first turbine vane 104 in the high
pressure turbine 54 to the annular exit 106 of the downstream end
of the last airfoil 108 in the low pressure turbine 46. The maximum
thrust may be Sea Level Takeoff Thrust "SLTO thrust" which is
commonly defined as the flat-rated static thrust produced by the
turbofan at sea-level.
[0080] The volume V of the turbine section may be best understood
from FIG. 10. As shown, the mid turbine frame 58 is disposed
between the high pressure turbine 54, and the low pressure turbine
46. The volume V is illustrated by a dashed line, and extends from
an inner periphery I to an outer periphery O. The inner periphery
is defined by the flow path of rotors, but also by an inner
platform flow paths of vanes. The outer periphery is defined by the
stator vanes and outer air seal structures along the flowpath. The
volume extends from a most upstream end of the vane 104, typically
its leading edge, and to the most downstream edge of the last
rotating airfoil 108 in the low pressure turbine section 46.
Typically this will be the trailing edge of the airfoil 108.
[0081] The power density in the disclosed gas turbine engine is
much higher than in the prior art. Eight exemplary engines are
shown below which incorporate turbine sections and overall engine
drive systems and architectures as set forth in this application,
and can be found in Table I as follows:
TABLE-US-00001 TABLE 1 Turbine section Thrust volume Thrust/turbine
SLTO from the section volume Engine (lbf) Inlet (lbf/in.sup.3) 1
17,000 3,859 4.40 2 23,300 5,330 4.37 3 29,500 6,745 4.37 4 33,000
6,745 4.84 5 96,500 31,086 3.10 6 96,500 62,172 1.55 7 96,500
46,629 2.07 8 37,098 6,745 5.50
[0082] Thus, in example embodiments, the power density would be
greater than or equal to about 1.5 lbf/in.sup.3. More narrowly, the
power density would be greater than or equal to about 2.0 lbf
/in.sup.3. Even more narrowly, the power density would be greater
than or equal to about 3.0 lbf/in.sup.3. More narrowly, the power
density is greater than or equal to about 4.0 lbf/in.sup.3. Also,
in embodiments, the power density is less than or equal to about
5.5 lbf/in.sup.3.
[0083] Engines made with the disclosed architecture, and including
turbine sections as set forth in this application, and with
modifications within the scope of this disclosure, thus provide
very high efficient operation, and increased fuel efficiency and
lightweight relative to their thrust capability.
[0084] An exit area 112 is defined at the exit location for the
high pressure turbine 54 and an exit area 110 is defined at the
outlet 106 of the low pressure turbine 46. The gear reduction 48
(shown in FIG. 1) provides for a range of different rotational
speeds of the fan drive turbine, which in this example embodiment
is the low pressure turbine 46, and the fan 42 (FIG. 1).
Accordingly, the low pressure turbine 46, and thereby the low spool
30 including the low pressure compressor 44 may rotate at a very
high speed. Low pressure turbine 46 and high pressure turbine 54
operation may be evaluated looking at a performance quantity which
is the exit area for the respective turbine section multiplied by
its respective speed squared. This performance quantity ("PQ") is
defined as:
PQ.sub.ltp=(A.sub.lps.times.V.sub.lpt.sup.2) Equation 1
PQ.sub.hpt=(A.sub.hpt.times.V.sub.hpt.sup.2) Equation 2
[0085] where A.sub.lpt is the area 110 of the low pressure turbine
46 at the exit 106, V.sub.lpt is the speed of the low pressure
turbine section; A.sub.hpt is the area of the high pressure turbine
54 at the exit 114, and where V.sub.hpt is the speed of the high
pressure turbine 54. As known, one would evaluate this performance
quantity at the redline speed for each turbine section.
[0086] Thus, a ratio of the performance quantity for the low
pressure turbine 46 compared to the performance quantify for the
high pressure turbine 54 is:
(A.sub.lpt.times.V.sub.lpt.sup.2)/(A.sub.hpt.times.V.sub.hpt.sup.2)=PQ.s-
ub.ltp/PQ.sub.hpt Equation 3
[0087] In one turbine embodiment made according to the above
design, the areas of the low and high pressure turbines 46, 54 are
557.9 in.sup.2 and 90.67 in.sup.2, respectively. Further, the
redline speeds of the low and high pressure turbine 46, 54 are
10179 rpm and 24346 rpm, respectively. Thus, using Equations 1 and
2 above, the performance quantities for the example low and high
pressure turbines 46,54 are:
PQ.sub.ltp=(A.sub.lpt.times.V.sub.lpt.sup.2)=(557.9 in.sup.2)(10179
rpm).sup.2=57805157673.9 in.sup.2 rpm.sup.2 Equation 1
PQ.sub.hpt=(A.sub.hpt.times.V.sub.hpt.sup.2)=(90.67 in.sup.2)(24346
rpm).sup.2=53742622009.72 in.sup.2 rpm.sup.2 Equation 2
[0088] and using Equation 3 above, the ratio for the low pressure
turbine section to the high pressure turbine section is:
Ratio=PQ.sub.hpt/PQ.sub.hpt=57805157673.9 in.sup.2
rpm.sup.2/53742622009.72 in.sup.2 rpm.sup.2=1.075
[0089] In another embodiment, the ratio is greater than about 0.5
and in another embodiment the ratio is greater than about 0.8. With
PQ.sub.ltp/PQ.sub.hpt ratios in the 0.5 to 1.5 range, a very
efficient overall gas turbine engine is achieved. More narrowly,
PQ.sub.ltp/PQ.sub.hpt ratios of above or equal to about 0.8
provides increased overall gas turbine efficiency. Even more
narrowly, PQ.sub.ltp/PQ.sub.hpt ratios above or equal to 1.0 are
even more efficient thermodynamically and from an enable a
reduction in weight that improves aircraft fuel burn efficiency. As
a result of these PQ.sub.ltp/PQ.sub.hpt ratios, in particular, the
turbine section 28 can be made much smaller than in the prior art,
both in diameter and axial length. In addition, the efficiency of
the overall engine is greatly increased.
[0090] Referring to FIG. 11, portions of the low pressure
compressor 44 and the low pressure turbine 46 of the low spool 30
are schematically shown and include rotors 116 of the low pressure
turbine 46 and rotors 132 of the low pressure compressor 44. Each
of the rotors 116 includes a bore radius 122, a live disk radius
124 and a bore width 126 in a direction parallel to the axis A. The
rotor 116 supports turbine blades 118 that rotate relative to the
turbine vanes 120. The low pressure compressor 44 includes rotors
132 including a bore radius 134, a live disk radius 136 and a bore
width 138. The rotor 132 supports compressor blades 128 that rotate
relative to vanes 130.
[0091] The bore radius 122 is that radius between an inner most
surface of the bore and the axis. The live disk radius 124 is the
radial distance from the axis of rotation A and a portion of the
rotor supporting airfoil blades. The bore width 126 of the rotor in
this example is the greatest width of the rotor and is disposed at
a radial distance spaced apart from the axis A determined to
provide desired physical performance properties.
[0092] The rotors for each of the low compressor 44 and the low
pressure turbine 46 rotate at an increased speed compared to prior
art low spool configurations. The geometric shape including the
bore radius, live disk radius and the bore width are determined to
provide the desired rotor performance in view of the mechanical and
thermal stresses selected to be imposed during operation. Referring
to FIG. 12, with continued reference to FIG. 11, a turbine rotor
116 is shown to further illustrate the relationship between the
bore radius 126 and the live disk radius 124. Moreover, the
relationships disclosed are provided within a known range of
materials commonly utilized for construction of each of the
rotors.
[0093] Accordingly, the increased performance attributes and
performance are provided by desirable combinations of the disclosed
features of the various components of the described and disclosed
gas turbine engine embodiments.
[0094] FIG. 13 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.
[0095] FIG. 14 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.
[0096] The embodiments 200, 300 of FIG. 13 or 14 may be utilized
with the features disclosed above.
[0097] Although an example embodiment has been disclosed, a worker
of ordinary skill in this art would recognize that certain
modifications would come within the scope of this disclosure. For
that reason, the following claims should be studied to determine
the scope and content of this disclosure.
* * * * *