US2023228218A1PendingUtilityA1
Reliable gearbox for gas turbine engine
Est. expiryDec 5, 2039(~13.4 yrs left)· nominal 20-yr term from priority
Inventors:Mark Spruce
F02C 7/36F16H 1/28F16H 57/082F05D 2220/323F16H 1/32F16H 1/2809F16H 3/70F16H 48/10F16H 57/023F16H 57/08F16H 57/025F02C 3/04F16H 2057/085F05D 2260/40311Y02T50/60
73
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0
Cited by
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References
0
Claims
Abstract
A gas turbine engine configured with an engine core. A fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear. Additionally, the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×1018 N2m−2.
Claims
exact text as granted — not AI-modified1 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear, and wherein: a gear ratio of the gearbox is in a range from 3 to 3.9; a specific thrust of the engine at cruise is in a range from 70 to 100 NKg −1 s; a fan tip loading defined as dH/U tip 2 , where dH is an enthalpy rise across the fan and U tip is a velocity of the fan tip is in a range from 0.28 to 0.36 at cruise conditions; an overall pressure ratio defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of the compressor is in a range from 40 to 65 at cruise conditions; and a first relationship is defined as:
(
gear
ratio
×
fan
tip
loading
×
overall
pressure
ratio
)
specific
thrust
the first relationship being in a range 0.40 to 0.99 KgN −1 s −1 .
2 . The gas turbine engine of claim 1 , wherein the first relationship is in a range 0.40 to 0.72 KgN −1 s −1 .
3 . The gas turbine engine of claim 1 , wherein the first relationship is in a range 0.40 to 0.51 KgN −1 s −1 .
4 . The gas turbine engine of claim 1 , wherein the first relationship is in a range 0.40 to 0.48 KgN −1 s −1 .
5 . The gas turbine engine of claim 1 , wherein the first relationship is in a range 0.42 to 0.56 KgN −1 s −1 .
6 . The gas turbine engine of claim 5 , wherein:
the gear ratio of the gearbox is in a range from 3 to 3.4; the specific thrust of the engine at cruise is in a range from 85 to 95 NKg −1 s; the fan tip loading is in a range from 0.28 to 0.31 at cruise conditions; and the overall pressure ratio is in a range from 40 to 50 at cruise conditions.
7 . The gas turbine engine of claim 1 , wherein the gear ratio of the gearbox is in a range from 3.2 to 3.4.
8 . The gas turbine engine of claim 1 , wherein specific thrust of the engine at cruise is in a range from 80 to 90 NKg −1 s.
9 . The gas turbine engine of claim 1 , wherein the fan tip loading is in a range from 0.28 to 0.34 at cruise conditions.
10 . The gas turbine engine of claim 1 , wherein the overall pressure ratio is in a range from 45 to 55 at cruise conditions.
11 . The gas turbine engine of claim 1 , wherein:
the gear ratio of the gearbox is in a range from 3 to 3.4; the specific thrust of the engine at cruise is in a range from 80 to 95 NKg −1 s; the fan tip loading is in a range from 0.28 to 0.31 at cruise conditions; and the overall pressure ratio is in a range from 40 to 50 at cruise conditions.
12 . The gas turbine engine of claim 1 , wherein:
(i) a bypass ratio at cruise is in a range from 12.5 to 18; and/or (ii) each fan blade has a radial span extending from a hub at a 0% span position to a tip at a 100% span position, and the ratio of a radius of the fan blade at the hub to a radius of the fan blade at the tip is in a range from 0.27 to 0.30.
13 . The gas turbine engine of claim 1 , wherein the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×10 18 N 2 m 2 and less than 2.6×10 22 N 2 m −2 .
14 . The gas turbine engine of claim 1 , wherein a gear mesh stiffness between the planet gears and the sun gear is greater than or equal to 1.20×10 9 N/m.
15 . The gas turbine engine of claim 1 , wherein the effective linear torsional stiffness of the planet carrier is greater than or equal to 7.00×10 9 N/m.
16 . The gas turbine engine of claim 1 , wherein
a carrier to ring mesh ratio of:
the
effective
linear
torsional
stiffness
of
the
planet
carrier
(
34
)
gear
mesh
stiffness
between
the
planet
gears
(
32
)
and
the
ring
gear
(
38
)
is greater than or equal to 0.2.
17 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear, and wherein: a gear ratio of the gearbox is in a range from 3 to 3.9; a fan tip loading defined as dH/U tip 2 , where dH is an enthalpy rise across the fan and U tip is a velocity of the fan tip is in a range from 0.28 to 0.36 at cruise conditions; an overall pressure ratio defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of the compressor is in a range from 40 to 65 at cruise conditions; and a second relationship is defined as:
(
gear
ratio
×
overall
pressure
ratio
)
fan
tip
loading
the second relationship being in a range 428.6 and 682.8.
18 . The gas turbine engine of claim 17 , wherein:
the gear ratio of the gearbox is in a range from 3 to 3.4; the fan tip loading is in a range from 0.28 to 0.34 at cruise conditions; the overall pressure ratio defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of the compressor is in a range from 40 to 55 at cruise conditions; and the second relationship is in a range of 443.5 and 512.1.
19 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear, and wherein: a gear ratio of the gearbox is in a range from 3 to 3.9; an overall pressure ratio defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of the compressor is in a range from 40 to 65 at cruise conditions; a specific thrust of the engine at cruise is in a range from 70 to 100 NKg −1 s; and a third relationship is defined as:
(
gear
ratio
×
overall
pressure
ratio
)
specific
thrust
the third relationship being in a range 1.38 and 3.29 KgN −1 s −1 .
20 . The gas turbine engine according to claim 19 , wherein:
the gear ratio of the gearbox is in a range from 3 to 3.4; the overall pressure ratio is in a range from 40 to 50 at cruise conditions; the specific thrust of the engine at cruise is in a range from 85 to 95 NKg −1 s; and the third relationship is in a range 1.39 and 2.00 KgN −1 s −1 .Join the waitlist — get patent alerts
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