US2022136434A1PendingUtilityA1
Turbine positioning in a gas turbine engine
Est. expiryDec 19, 2039(~13.4 yrs left)· nominal 20-yr term from priority
B64D 27/402F02C 3/113F02C 7/06F02K 3/06F05D 2260/40F05D 2240/54F01D 25/16F05D 2260/96Y02T50/60F05D 2260/40311F02C 3/06F02C 7/36F01D 15/12B64D 27/26
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Claims
Abstract
A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core. The engine core further includes three bearings arranged to support the core shaft including a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings. A minor span to turbine length ratio is equal to or less than 1.05.
Claims
exact text as granted — not AI-modified1 . A gas turbine engine for an aircraft comprising:
an engine core comprising:
a turbine that is a lowest pressure turbine of the gas turbine engine and has a turbine length defined as a distance between a root of a most upstream blade of the turbine at a leading edge thereof and a root of a most downstream blade of the turbine at a trailing edge thereof, the turbine comprising two to five sets of turbine blades,
a compressor that is a lowest pressure compressor of the gas turbine engine,
a core shaft connecting the turbine to the compressor, and
three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span defined as a distance between the two rearward bearings; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; wherein a minor span to turbine length ratio of:
the
minor
span
the
turbine
length
is equal to or less than 1.05.
2 . The gas turbine engine of claim 1 , wherein the minor span to turbine length ratio is in a range from 0.70 to 1.05.
3 . The gas turbine engine of claim 1 , wherein the minor span to turbine length ratio is in a range from 0.85 to 0.95.
4 . The gas turbine engine of claim 1 , wherein the turbine comprises a total of four sets of the turbine blades, and the two rearward bearings are located downstream of a trailing edge of a turbine blade of a third set of the turbine blades from a front of the turbine, at a root of the turbine blade.
5 . The gas turbine engine of claim 1 , wherein the two rearward bearings are located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the lowest pressure turbine blade.
6 . The gas turbine engine of claim 1 , wherein the turbine comprises a total of three sets of the turbine blades.
7 . The gas turbine engine of claim 1 , wherein the minor span is less than or equal to 350 mm.
8 . The gas turbine engine of claim 1 , wherein a length of the core shaft is in a range from 1800 to 2900 mm.
9 . The gas turbine engine of claim 1 , wherein a length of the core shaft is in a range from 2300 to 2800 mm.
10 . The gas turbine engine of claim 1 , wherein the core shaft is configured to have a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm.
11 . The gas turbine engine of claim 10 , wherein the upper bound on the running speed range of the core shaft is an upper bound on a maximum take-off running speed and the lower bound on the running speed range of the core shaft is a minimum running speed under ground idle conditions.
12 . The gas turbine engine of claim 1 , wherein the core shaft has a length between the forward bearing and a rearmost rearward bearing, and the three bearings are arranged such that a length ratio of the minor span to the length of the core shaft is equal to or less than 0.14.
13 . The gas turbine engine of claim 12 , wherein the length ratio of the minor span to the length of the core shaft is in a range from 0.08 to 0.13.
14 . The gas turbine engine of claim 1 , wherein a forwardmost rearward bearing of the two rearward bearings has a bearing stiffness in a range of 30 kN/mm to 100 kN/mm.
15 . The gas turbine engine of claim 14 , wherein a stiffness ratio of the bearing stiffness at the forwardmost rearward bearing to the minor span is in a range from 0.08 to 0.5 kN/mm 2 .
16 . The gas turbine engine of claim 1 , wherein the fan has a fan diameter in a range from 330 cm to 380 cm.
17 . The gas turbine engine of claim 1 , further comprising a tail bearing housing located rearward of the turbine and comprising two bearing discs, each bearing disc being arranged to support one of the two rearward bearings.
18 . The gas turbine engine of claim 17 , wherein the two bearing discs are oriented at least substantially perpendicular to an axis of the gas turbine engine.
19 . The gas turbine engine of claim 1 , wherein:
the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.
20 . A gas turbine engine for an aircraft comprising:
an engine core comprising:
a turbine that is a lowest pressure turbine of the gas turbine engine and has a turbine length defined as a distance between a root of a most upstream blade of the turbine at a leading edge thereof and a root of a most downstream blade of the turbine at a trailing edge thereof,
a compressor that is a lowest pressure compressor of the gas turbine engine,
a core shaft connecting the turbine to the compressor, and
three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span defined as a distance between the two rearward bearings;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that is configured to receive an input from the core shaft, and output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
a minor span to turbine length ratio of:
the
minor
span
the
turbine
length
is equal to or less than 1.05; and
the core shaft has a length, L, between the forward bearing and a rearmost rearward bearing, in a range from 1800 mm to 2900, and
the gearbox has a gear ratio in a range of 3.1 to 3.8.Join the waitlist — get patent alerts
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