Gas turbine engines and epicyclic gearboxes with planet gear clearances

1 . A gas turbine engine comprising:

a gearbox assembly mechanically coupled to an LP shaft of the gas turbine engine, wherein the gearbox assembly comprises a sun gear, a ring gear, a carrier, and a plurality of planet gears arranged in a planetary configuration, wherein each of the plurality of planet gears comprises:

a bearing pin comprising a pin outer surface;

an annular planet gear rim comprising an inner surface, wherein the inner surface and the pin outer surface define a clearance, wherein the clearance is greater than zero when a radial component force, a pinch component force, a tangential component force, and a centrifugal component force are applied to the planet gear;

a planet gear bending stress neutral axis radius, wherein the planet gear bending stress neutral axis radius is a radius where stresses and strains within the annular planet gear rim are zero when the radial component force, the pinch component force, the tangential component force, and the centrifugal component force are applied to the planet gear; and

a pin clearance parameter defined by:

PCP

=

K

1

c

r

⁢

GR

GR

-

2

⁢

r

p

2

[

K

2

⁢

r

p

3

⁢

Ω

fan

3

-

HP

fan

N

p

⁢

(

GR

-

2

GR

)

2

]

wherein “PCP” is the pin clearance parameter in rpm, “c r ” is the clearance in inches, “GR” is a gear ratio of the gearbox assembly, “r p ” is the planet gear bending stress neutral axis radius in inches, “N p ” is a number of the plurality of planet gears, “HP fan ” is a fan power of the gas turbine engine in horsepower at takeoff conditions, “Ω fan ” is a fan speed of the gas turbine engine in rpm at takeoff conditions, K 1 is a first constant of 1.96×10 −5 per horsepower-minute-inch, and K 2 is a second constant of 4.91×10 −9 horsepower-minutes cubed per cubic inch, and

wherein the pin clearance parameter is greater than or equal to zero rpm and less than or equal to 3,334 rpm; and

a lubricant system configured to supply a lubricant to the gearbox assembly, the lubricant system comprising:

a supply line;

a first sensing component configured to measure a lubricant temperature at a first sensing position upstream of the gearbox assembly; and

a second sensing component configured to measure a lubricant temperature at a second sensing position downstream of the gearbox assembly;

wherein the lubricant system is configured to determine a difference of the lubricant temperature measured at the first sensing position and the lubricant temperature measured at the second sensing position, and

wherein the lubricant system is configured to increase a flow rate of the lubricant through a component supply line if the difference of the lubricant temperature measured at the first sensing position and the lubricant temperature measured at the second sensing position exceeds a predetermined threshold lubricant temperature difference.

2 . The gas turbine engine of claim 1 , wherein:

the sun gear further comprises a plurality of sun gear teeth,

the ring gear further comprises a plurality of ring gear teeth,

the gear ratio of the gearbox assembly is a sum of the number of the plurality of ring gear teeth and the number of the plurality of sun gear teeth divided by the number of the plurality of sun gear teeth, and

wherein the gear ratio is in a range from 2.5 to 5.

3 . The gas turbine engine of claim 1 , wherein the pin clearance parameter includes values in a range from zero rpm to 3,000 rpm,

wherein the lubricant system further comprises a bypass valve fluidly coupled to the supply line, the component supply line, and a bypass line, and

wherein the bypass valve is configured to control a flow of lubricant through at least one of the component supply line or the bypass line.

4 . The gas turbine engine of claim 1 , wherein the pin clearance parameter includes values in a range from 48 rpm to 1,334 rpm,

wherein the lubricant system further comprises a bypass valve fluidly coupled to the supply line, the component supply line, and a bypass line, and

wherein the bypass valve is configured to control a flow of lubricant through at least one of the component supply line or the bypass line.

5 . The gas turbine engine of claim 1 , wherein the pin clearance parameter includes values in a range from 80 rpm to 1,300 rpm,

wherein the lubricant system further comprises a bypass valve fluidly coupled to the supply line, the component supply line, and a bypass line, and

wherein the bypass valve is configured to control a flow of lubricant through at least one of the component supply line or the bypass line.

6 . The gas turbine engine of claim 1 , wherein the gas turbine engine is configured to produce the fan power in a range from 7,000 horsepower to 80,000 horsepower at takeoff conditions.

7 . The gas turbine engine of claim 1 , wherein the gas turbine engine is configured to produce the fan speed in a range from 1,600 rpm to 3,334 rpm at takeoff conditions.

8 . The gas turbine engine of claim 1 , wherein the gas turbine engine further comprises an HP compressor disposed aft of the gearbox assembly, wherein the HP compressor comprises eight, nine, or ten HP compressor stages.

9 . The gas turbine engine of claim 1 , wherein the gas turbine engine further comprises an LP turbine coupled to the LP shaft and comprising a plurality of LP turbine stages, wherein the number of the plurality of LP turbine stages is three, four, five, or six.

10 . The gas turbine engine of claim 1 , wherein the gas turbine engine further comprises:

a fan shaft coupled to the carrier of the gearbox assembly; and

a fan coupled to the fan shaft, wherein the fan comprises a fan diameter within a range from 80 inches to 95 inches.

11 . The gas turbine engine of claim 10 , wherein the fan diameter is within a range from 85 inches to 90 inches.

12 . The gas turbine engine of claim 1 , wherein each of the plurality of planet gears further comprises a journal bearing around which the annular planet gear rim is disposed.

13 . A gas turbine engine comprising:

a gearbox assembly mechanically coupled to an LP shaft of the gas turbine engine, wherein the gearbox assembly comprises a sun gear, a ring gear, a carrier, and a plurality of planet gears arranged in a planetary configuration, wherein each of the plurality of planet gears comprises:

a bearing pin comprising a pin outer surface;

an annular planet gear rim comprising an inner surface, wherein the inner surface and the pin outer surface define a clearance, wherein the clearance is greater than zero when a radial component force, a pinch component force, a tangential component force, and a centrifugal component force are applied to the planet gear;

a planet gear bending stress neutral axis radius, wherein the planet gear bending stress neutral axis radius is a radius where stresses and strains within the annular planet gear rim are zero when the radial component force, the pinch component force, the tangential component force, and the centrifugal component force are applied to the planet gear; and

a pin clearance parameter defined by:

PCP

=

K

1

c

r

⁢

GR

GR

-

2

⁢

r

p

2

[

K

2

⁢

r

p

3

⁢

Ω

fan

3

-

HP

fan

N

p

⁢

(

GR

-

2

GR

)

2

]

wherein “PCP” is the pin clearance parameter in rpm, “c r ” is the clearance in inches, “GR” is a gear ratio of the gearbox assembly, “r p ” is the planet gear bending stress neutral axis radius in inches, “N p ” is a number of the plurality of planet gears, “HP fan ” is a fan power of the gas turbine engine in horsepower at takeoff conditions, “Ω fan ” is a fan speed of the gas turbine engine in rpm at takeoff conditions, K 1 is a first constant of 1.96×10 −5 per horsepower-minute-inch, and K 2 is a second constant of 4.91×10-9 horsepower-minutes cubed per cubic inch, and

wherein the pin clearance parameter is greater than or equal to zero rpm and less than or equal to 3,334 rpm; and

a lubricant system configured to supply a lubricant to the gearbox assembly, the lubricant system comprising:

a reservoir;

a supply line fluidly coupled to the reservoir;

a bypass valve fluidly coupled to the supply line;

a component supply line fluidly coupled to the bypass valve and fluidly coupled to the gearbox assembly;

a scavenge line fluidly coupled to the gearbox assembly and fluidly coupled to the reservoir; and

a bypass line fluidly coupled to the bypass valve and fluidly coupled to the scavenge line.

14 . The gas turbine engine of claim 13 , wherein the lubricant system is configured to supply lubricant to the gearbox assembly at a rate in a range from 5 gallons per minute to 55 gallons per minute, as measured at a max takeoff condition, and wherein the gas turbine engine comprises a fan having a fan diameter ranging from 80 inches to 95 inches.

15 . The gas turbine engine of claim 14 , wherein the lubricant system is configured to supply lubricant to the gearbox assembly at a rate in a range from 5.25 gallons per minute to 25 gallons per minute, and wherein the fan diameter ranges from 85 inches to 95 inches.

16 . The gas turbine engine of claim 13 , wherein the lubricant system further comprises:

a first sensing component configured to measure a lubricant temperature at a first sensing position upstream of the gearbox assembly; and

a second sensing component configured to measure a lubricant temperature at a second sensing position downstream of the gearbox assembly.

17 . The gas turbine engine of claim 16 , wherein the gearbox assembly has a gear ratio in a range from 2.5 to 5.

18 . The gas turbine engine of claim 16 , wherein the lubricant system is configured to determine a difference of the lubricant temperature measured at the first sensing position and the lubricant temperature measured at the second sensing position.

19 . The gas turbine engine of claim 18 , wherein the lubricant system is configured to increase a flow rate of the lubricant through the component supply line if the difference of the lubricant temperature measured at the first sensing position and the lubricant temperature measured at the second sensing position exceeds a predetermined threshold lubricant temperature difference.

20 . The gas turbine engine of claim 13 , wherein the lubricant system further comprises a heat exchanger fluidly coupled to the scavenge line.