U.S. patent application number 15/921934 was filed with the patent office on 2018-07-19 for gas turbine engine for long range aircraft.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Frederick M. Schwarz, Gabriel L. Suciu.
Application Number | 20180202369 15/921934 |
Document ID | / |
Family ID | 53367863 |
Filed Date | 2018-07-19 |
United States Patent
Application |
20180202369 |
Kind Code |
A1 |
Schwarz; Frederick M. ; et
al. |
July 19, 2018 |
GAS TURBINE ENGINE FOR LONG RANGE AIRCRAFT
Abstract
A gas turbine engine comprises a fan for delivering air into a
bypass duct as bypass flow, into a core housing as core flow, with
the core housing containing an upstream compressor rotor and a
downstream compressor rotor. An overall pressure ratio is defined
across the upstream and downstream compressor rotors. A bypass
ratio is defined as a volume of air delivered as bypass flow
compared to a volume of air delivered into the core housing. The
overall pressure ratio is greater than or equal to about 45.0, and
the bypass ratio is greater than or equal to about 11.0.
Inventors: |
Schwarz; Frederick M.;
(Glastonbury, CT) ; Suciu; Gabriel L.;
(Glastonbury, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
53367863 |
Appl. No.: |
15/921934 |
Filed: |
March 15, 2018 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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14107273 |
Dec 16, 2013 |
9976489 |
|
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15921934 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 7/36 20130101; F02K
3/06 20130101; F05D 2260/40311 20130101; F05D 2250/30 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02K 3/06 20060101 F02K003/06 |
Claims
1. A gas turbine engine comprising: a fan for delivering air into a
bypass duct as bypass flow and into a core housing as core flow; an
upstream compressor rotor and a downstream compressor rotor; an
overall pressure ratio greater than or equal to 45.0; a bypass
ratio defined as a volume of air delivered as bypass flow compared
to a volume of air delivered into said core housing, and said
bypass ratio greater than or equal to 11; a fan drive turbine
driving said upstream compressor rotor and driving said fan through
a gear reduction, said fan drive turbine having at least three
stages; and wherein a ratio of a tip speed at said downstream
compressor rotor compared to a tip speed at said upstream
compressor rotor is less than or equal to 1.18.
2. The gas turbine engine as set forth in claim 1, further
comprising two additional turbine rotors upstream of said fan drive
turbine.
3. The gas turbine engine as set forth in claim 2, wherein said fan
drive turbine has six or fewer stages.
4. The gas turbine engine as set forth in claim 3, wherein said
ratio of said tip speed at said downstream compressor rotor
compared to said tip speed at said upstream compressor rotor is
greater than or equal to 1.0.
5. The gas turbine engine as set forth in claim 4, wherein said fan
drive turbine having fewer stages than said upstream compressor
rotor.
6. The gas turbine engine as set forth in claim 4, wherein said fan
has a low fan pressure ratio of less than 1.45.
7. The gas turbine engine as set forth in claim 6, wherein said gas
turbine engine being designed for use on a long range aircraft,
said long range aircraft being defined at least by one of an
aircraft with at least two passenger aisles, or an aircraft with a
flight length equal to, or greater than 6.0 hours.
8. The gas turbine engine as set forth in claim 7, wherein said fan
drive turbine has a turbine pressure ratio that is greater than
five 5:1, said turbine pressure ratio defined as a pressure
measured prior to an inlet of said fan drive turbine as related to
a pressure at an outlet of said fan drive turbine prior to an
exhaust nozzle.
9. The gas turbine engine as set forth in claim 6, wherein said fan
drive turbine has a turbine pressure ratio that is greater than
five 5:1, said turbine pressure ratio defined as a pressure
measured prior to an inlet of said fan drive turbine as related to
a pressure at an outlet of said fan drive turbine prior to an
exhaust nozzle.
10. The gas turbine engine as set forth in claim 9, wherein: said
upstream compressor rotor has seven or fewer stages; a gear ratio
of said gear reduction is greater than or equal 2.6; and said
bypass ratio is greater than or equal to 14.0.
11. The gas turbine engine as set forth in claim 10, wherein said
gear reduction is a planetary gear system and said overall pressure
ratio is equal to or greater than 60.
12. The gas turbine engine as set forth in claim 11, wherein said
upstream compressor rotor has at least three stages.
13. The gas turbine engine as set forth in claim 12, further
comprising two turbine rotors upstream of said fan drive
turbine.
14. The gas turbine engine as set forth in claim 3, wherein said
fan has a low fan pressure ratio of less than 1.45.
15. The gas turbine engine as set forth in claim 14, wherein said
fan drive turbine has a turbine pressure ratio that is greater than
five 5:1, said turbine pressure ratio defined as a pressure
measured prior to an inlet of said fan drive turbine as related to
a pressure at an outlet of said fan drive turbine prior to an
exhaust nozzle.
16. The gas turbine engine as set forth in claim 15, wherein: said
fan drive turbine has six or fewer stages; and said upstream
compressor rotor has seven or fewer stages.
17. The gas turbine engine as set forth in claim 16, wherein a gear
ratio of said gear reduction is greater than or equal 2.6, and said
bypass ratio is greater than or equal to 14.0.
18. The gas turbine engine as set forth in claim 17, wherein said
ratio of a tip speed at said downstream compressor rotor compared
to said tip speed at said upstream compressor rotor is greater than
or equal to 1.0.
19. The gas turbine engine as set forth in claim 18, wherein said
gas turbine engine being designed for use on a long range aircraft,
said long range aircraft being defined at least by one of an
aircraft with at least two passenger aisles, or an aircraft with a
flight length equal to, or greater than 6.0 hours.
20. The gas turbine engine as set forth in claim 18, wherein: said
gear reduction is a planetary gear system; a gear ratio of said
gear reduction is greater than or equal 3.6; and said overall
pressure ratio is equal to or greater than 65.
21. The gas turbine engine as set forth in claim 20, wherein said
gas turbine engine being designed for use on a long range aircraft,
said long range aircraft being defined at least by one of an
aircraft with at least two passenger aisles, or an aircraft with a
flight length equal to, or greater than 6.0 hours.
22. A method of designing a gas turbine engine for a long range
aircraft comprising: driving a fan through a gear reduction with a
fan drive turbine, with said fan delivering air into a bypass duct
as bypass flow and into a core housing as core flow, with said core
flow flowing into an upstream compressor rotor and a downstream
compressor rotor, wherein a bypass ratio is defined as a volume of
air delivered as bypass flow compared to a volume of air delivered
into said core housing, said bypass ratio greater than or equal to
11, said fan drive turbine having at least three stages but no more
than six stages, and two additional turbine rotors upstream of said
fan drive turbine; driving said upstream and downstream compressor
rotors, with an overall pressure ratio greater than or equal to
45.0; and wherein said gas turbine engine being mounted on said
long range aircraft during operation, with said long range aircraft
being defined at least by one of an aircraft with at least two
passenger aisles, or an aircraft with a flight length equal to, or
greater than 6.0 hours.
23. The method as set forth in claim 22, wherein: the step of
driving said upstream and downstream compressor rotors establishes
a ratio of a tip speed at said downstream compressor rotor compared
to a tip speed at said upstream compressor rotor that is less than
or equal to 1.18 and is greater than or equal to 1.0; and the step
of driving said fan occurs for a duration of said flight length,
said flight length including a first time at a low power cruise
condition and a second time at a climb condition and a take-off
condition, and a ratio of said first time to said second time being
at least 15:1.
24. The method as set forth in claim 23, wherein: said fan drive
turbine has fewer than six stages; and said upstream compressor
rotor has at least three stages, but fewer than seven stages.
25. The method as set forth in claim 24, wherein said long range
aircraft is defined as an aircraft with a flight length of 3,000 to
8,000 miles.
26. The method as set forth in claim 25, wherein said gas turbine
engine is a first engine and a second engine, each of said first
and second engines mounted on said long range aircraft during
operation such that a take-off thrust of said long range aircraft
is greater than 50,000 lbf at static CSLTO 86.degree. F.
27. The method as set forth in claim 22, wherein: the step of
driving said upstream and downstream compressor rotors includes
driving said upstream compressor rotor with said fan drive turbine;
said bypass ratio is greater than or equal to 14.0; said fan has a
low fan pressure ratio of less than 1.45; and said fan drive
turbine has a turbine pressure ratio that is greater than five 5:1,
said turbine pressure ratio defined as a pressure measured prior to
an inlet of said fan drive turbine as related to a pressure at an
outlet of said fan drive turbine prior to an exhaust nozzle.
28. The method as set forth in claim 27, wherein said overall
pressure ratio equal to or greater than 60.
29. The method as set forth in claim 28, wherein said gas turbine
engine is a first engine and a second engine, each of said first
and second engines mounted on said long range aircraft during
operation such that a take-off thrust of said long range aircraft
is greater than 50,000 lbf at static CSLTO 86.degree. F., and said
step of driving said upstream and downstream compressor rotors
establishes a ratio of a tip speed at said downstream compressor
rotor compared to a tip speed at said upstream compressor rotor
that is less than or equal to 1.18.
30. The method as set forth in claim 27, wherein the step of
driving said upstream and downstream compressor rotors establishes
a ratio of a tip speed at said downstream compressor rotor compared
to a tip speed at said upstream compressor rotor that is greater
than or equal to 1.0.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 14/107,273, filed Dec. 16, 2013.
BACKGROUND OF THE INVENTION
[0002] This application relates to a gas turbine engine designed
for use on longer range aircraft.
[0003] Gas turbine engines are known and may include a fan
delivering air into a bypass duct as propulsion air. In addition,
the fan typically delivers air into a core housing and to a
compressor. There are, typically, at least two compressor rotors
with an upstream or lower pressure rotor compressing the air and
then delivering it into a downstream or higher pressure rotor. The
compressed air from the downstream compressor rotor is typically
delivered into a combustion section where it is mixed with fuel and
ignited. Products of this combustion may pass downstream over
turbine rotors including an upstream turbine rotor that drives the
downstream compressor rotor and a downstream turbine rotor that
drives the upstream compressor rotor.
[0004] In one type engine, the downstream turbine rotor also drove
a fan rotor, such that the fan rotor, the upstream compressor
rotor, and the downstream turbine rotor all rotated at a single
speed. More recently, a gear reduction has been placed between the
fan rotor and the downstream turbine rotor or the fan drive
turbine.
[0005] It is desirable to increase the compression ratio or the
amount of compression done to air across the two compressor rotors.
However, there has been a significant limitation in that the stress
and temperature at the downstream end of the downstream compressor
rotor limits how high the overall compression ratio may reach.
[0006] This area must be designed to withstand the repeated
application of the highest stress situations for the gas turbine
engine which typically occurs during take-off.
SUMMARY OF THE INVENTION
[0007] In a featured embodiment, a gas turbine engine comprises a
fan for delivering air into a bypass duct as bypass flow, and into
a core housing as core flow, with the core housing containing an
upstream compressor rotor and a downstream compressor rotor. An
overall pressure ratio is defined across the upstream and
downstream compressor rotors. A bypass ratio is defined as a volume
of air delivered as bypass flow compared to a volume of air
delivered into the core housing. The overall pressure ratio is
greater than or equal to about 45.0, and the bypass ratio is
greater than or equal to about 11.0
[0008] In another embodiment according to the previous embodiment,
the gas turbine engine is designed for use on long range aircraft
defined as aircraft with at least two passenger aisles.
[0009] In another embodiment according to any of the previous
embodiments, a fan drive turbine is configured to drive the
upstream compressor rotor and the fan rotor through a gear
reduction, with the fan drive turbine having at least three
stages.
[0010] In another embodiment according to any of the previous
embodiments, the fan drive turbine has six or fewer stages.
[0011] In another embodiment according to any of the previous
embodiments, a ratio of a tip speed at the downstream compressor
rotor compared to a tip speed at the upstream compressor rotor is
less than or equal to about 1.18 and greater than or equal to about
1.0.
[0012] In another embodiment according to any of the previous
embodiments, the bypass ratio is greater than or equal to about
14.0.
[0013] In another embodiment according to any of the previous
embodiments, the overall pressure ratio across the upstream and
downstream compressor rotors is equal to or greater than about
60.0.
[0014] In another embodiment according to any of the previous
embodiments, a gear ratio of the gear reduction is greater than or
equal about 2.6.
[0015] In another embodiment according to any of the previous
embodiments, the gas turbine engine is designed for use on long
range aircraft defined as aircraft with a flight length equal to,
or greater than about 6.0 hours.
[0016] In another embodiment according to any of the previous
embodiments, the upstream compressor rotor has at least three
stages.
[0017] In another embodiment according to any of the previous
embodiments, the upstream compressor rotor has seven or fewer
stages.
[0018] In another embodiment according to any of the previous
embodiments, a fan drive turbine is configured to drive the
upstream compressor rotor and the fan rotor through a gear
reduction, with the fan drive turbine having at least three
stages.
[0019] In another embodiment according to any of the previous
embodiments, the fan drive turbine has six or fewer stages.
[0020] In another embodiment according to any of the previous
embodiments, a ratio of a tip speed at the downstream compressor
rotor compared to a tip speed at the upstream compressor rotor is
less than or equal to about 1.18 and greater than or equal to about
1.0.
[0021] In another embodiment according to any of the previous
embodiments, the bypass ratio is greater than or equal to about
14.0.
[0022] In another embodiment according to any of the previous
embodiments, the overall pressure ratio across the upstream and
downstream compressor rotors is equal to or greater than about
60.0.
[0023] In another embodiment according to any of the previous
embodiments, a gear ratio of the gear reduction is greater than or
equal about 2.6.
[0024] In another embodiment according to any of the previous
embodiments, a fan drive turbine is configured to drive the fan
rotor through a gear reduction.
[0025] In another embodiment according to any of the previous
embodiments, the fan drive turbine also is configured to drive the
upstream compressor rotor, along with the fan rotor.
[0026] In another embodiment according to any of the previous
embodiments, there are two additional turbine rotors upstream of
the fan drive turbine for respectively driving the upstream and
downstream compressor rotors.
[0027] In another embodiment according to any of the previous
embodiments, a ratio of a tip speed at the downstream compressor
rotor compared to a tip speed at the upstream compressor rotor is
less than or equal to about 1.18 and greater than or equal to about
1.0.
[0028] In another embodiment according to any of the previous
embodiments, the overall pressure ratio across the upstream and
downstream compressor rotors is equal to or greater than about
60.0.
[0029] In another embodiment according to any of the previous
embodiments, a gear ratio of the gear reduction is greater than or
equal about 2.6.
[0030] These and other features may be best understood from the
following drawing and specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0031] FIG. 1A schematically shows a gas turbine engine wherein the
fan-driving turbine also drives the upstream compressor.
[0032] FIG. 1B schematically shows an aircraft that may incorporate
the engine of FIG. 1A or FIG. 2.
[0033] FIG. 2 schematically shows another gas turbine engine.
DETAILED DESCRIPTION
[0034] A gas turbine engine 20 shown schematically in FIG. 1A, is
designed for use on long range aircraft.
[0035] As known, a fan rotor 22 delivers bypass air B within a
nacelle 24. A core engine housing 26 receives core air flow C from
the fan rotor 22. The core air flow C initially reaches an upstream
compressor rotor 28, which compresses the air to a first lower
level and then delivers that air into a downstream compressor rotor
38, where additional compression occurs.
[0036] The air from the compressor rotor 38 is delivered into a
combustion section 42, mixed with fuel and ignited. Products of
this combustion pass downstream over an upstream turbine rotor 40,
which operates at a higher pressure and speed and drives a shaft 36
to drive the downstream compressor rotor 38. Downstream of the
turbine rotor 40, the products of combustion drive a fan drive
turbine 34 which is a downstream turbine rotor and operates at a
lower pressure and speed than does the turbine rotor 40. The fan
drive turbine 34 drives the upstream compressor rotor 28 through a
shaft 30 and drives the fan rotor 22 through a gear reduction
32.
[0037] Applicant has recognized that on longer range aircraft the
percentage of engine operation time at high stress level, such as
take-off, becomes a very small percentage of the overall operation
time. Long range aircraft may be defined as traveling 3,000 to
8,000 miles or more, and from about 6 to about 16 hours of flying
time during a typical flight.
[0038] Often these aircraft can also be described as "twin aisle"
aircraft because they are wide body aircraft as opposed to single
aisle aircraft which are used on inter-continental flights or to
feed airports that are used as hub-and-spoke airports for
connecting flights.
[0039] As shown in FIG. 1B, an aircraft 100 schematically includes
two engines 20. There are aisles 104 and 106 separating seating
areas 102 and 108.
[0040] Take-off and climb will typically occur for only 45 seconds
at take-off power followed by perhaps 20 minutes at climb. In such
an example, the ratio of time at low power cruise to climb and
take-off is at least 15 to 1 and can become as high as 40 to 1.
[0041] With such systems, it is possible to achieve higher pressure
ratios because the high stress situations on the downstream end of
the downstream compressor rotor 38 will occur much less frequently
and over a small percentage of the overall operational time across
the life of the engine.
[0042] In disclosed embodiments, the overall pressure ratio or
amount of compression by the combined rotors 28 and 38 to the core
air flow C may be equal to or above about 45.0, above about 60.0,
and above about 65.0.
[0043] In addition, a bypass ratio or ratio of the volume of air
delivered as bypass flow B to the volume of air delivered as core
air flow C may be equal to or greater than about 11.0, and greater
than or equal to 12.0, or greater than or equal to 14.0.
[0044] A gear ratio of the gear reduction 32 may be greater than or
equal to about 2.6, greater than or equal to about 2.9, or greater
than or equal to about 3.6.
[0045] The upstream compressor rotor 28 may have less than or equal
to seven stages, or from three to seven stages. The fan drive
turbine 34 may have three or more stages, or from three to six
stages.
[0046] A tip speed at the tip 39 of the downstream compressor rotor
38 compared to the tip speed at tip 29 of the upstream compressor
rotor 28 is less than or equal to about 1.8 and greater than or
equal to about 1.0. Tip speed is defined as the tangential velocity
of the leading edge of the longest blade.
[0047] Finally, the take-off thrust for the long range aircraft is
greater than about 50,000 lbf at static CSLTO 86.degree. F. The
take-off thrust may be as high as 124,000 lbf at sea-level static
condition.
[0048] Another factor for peak efficiency which this compression
section is designed to produce is that there be at least two
turbine rotors (40) in front of the fan-driving turbine. This
allows for reasonable mach numbers through the turbine section
ahead of the fan drive turbine which improves overall engine
efficiency and reduces temperatures into the fan drive turbine to a
manageable level. In contrast, a single turbine rotor ahead of the
fan drive turbine may lower overall engine efficiency and drive up
temperatures into the fan drive turbine to such an extent that at
least one of the fan drive turbine stages may have to be cooled,
owing to the stresses there. Cooling the fan drive turbine also
lowers the efficiency of the overall engine cycle.
[0049] The fan diameter is significantly larger than that of the
low pressure compressor 44, and the low pressure turbine 46 has a
pressure ratio that is greater than about five (5:1). The low
pressure turbine 46 pressure ratio is pressure measured prior to
inlet of low pressure turbine 46 as related to the pressure at the
outlet of the low pressure turbine 46 prior to an exhaust nozzle.
The geared architecture 48 may be an epicyclic gear train, such as
a planetary gear system or other gear system, with a gear reduction
ratio of greater than about 2.3:1. It should be understood,
however, that the above parameters are only exemplary of one
embodiment of a geared architecture engine and that the present
invention is applicable to other gas turbine engines including
direct drive turbofans.
[0050] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of lbm of
fuel being burned divided by lbf of thrust the engine produces at
that minimum point. "Low fan pressure ratio" is the pressure ratio
across the fan blade alone, without a Fan Exit Guide Vane ("FEGV")
system. The low fan pressure ratio as disclosed herein according to
one non-limiting embodiment is less than about 1.45. "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)/(518.7.degree. R)].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.
[0051] Another embodiment engine 120 is illustrated in FIG. 2. Many
components are effectively identical to those shown in FIG. 1A, and
carry the same numbers, simply increased by 100.
[0052] Whereas the FIG. 1A engine 20 has a fan drive turbine 34
driving a compressor 28, and a fan rotor 22 through the gear
reduction 32, the FIG. 2 engine includes a higher pressure first
turbine 140 driving a higher pressure downstream compressor section
138. An intermediate turbine section 141 drives the first
compressor rotor, or the upstream compressor rotor 128. The fan
drive turbine 134 only drives the fan rotor 122 through a gear
reduction 132.
[0053] The quantities as described above with regard to the FIG. 1A
embodiment, would also be true of the FIG. 2 embodiment.
[0054] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *