Nacelle and compressor inlet arrangements
A gas turbine engine includes a nacelle defining a centerline axis and an annular splitter radially inward from the nacelle. A spinner is radially inward of the nacelle forward of a compressor section. A fan blade extends from a fan blade platform. A distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform. A distance H is the radial distance from the first point to the second point. The relative position of the first point and the second point is governed by the ratio of X H ≥ 1.5 for reducing foreign object debris (FOD) intake into the compressor section.
1. A gas turbine engine comprising:
a nacelle, the nacelle including:
i. a nacelle inlet;
ii. a nacelle outlet aft of the nacelle inlet; and
iii. a bypass duct therebetween;
a compressor section aft of the nacelle inlet;
an annular splitter separating the bypass duct from the compressor section;
a spinner radially inward of the nacelle forward of the compressor section;
a fan blade platform defined in a fan section aft of the spinner and radially inward of the nacelle; and
a fan blade extending from the fan blade platform toward the nacelle,
wherein a distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform, and wherein a distance H is the radial distance from the first point to the second point, wherein
1.5
≤
distance
X
distance
H
≤
4
for reducing foreign object debris (FOD) intake into the compressor section; and
wherein a point Z is defined at an intersection of the centerline axis and a line C normal to the centerline axis extending radially inward from the leading edge of the fan blade where the fan blade meets the fan blade platform, wherein a point W is defined at the intersection of the line C and the leading edge of the fan blade where the fan blade meets the fan blade platform, a distance L is defined from the point Z to a tip of the spinner, wherein a point V is defined along the centerline axis at a distance 0.25 times the distance L aft of the tip of the spinner, wherein a point U is defined at an intersection of a line E normal to the centerline axis extending radially outward from the point V and a line F extending from the point W to the tip of the spinner, wherein a distance M c is defined from the point T to the point U, and wherein a distance M p is defined from the point T to the point V, wherein
distance
Mc
distance
Mp
≤
1
/
2.
2. A gas turbine as recited in claim 1 , wherein a distance r is defined radially from the centerline axis to the first point, and an average distance r avg is defined radially from the centerline axis to a leading edge of the nacelle inlet taken over a section of the nacelle ranging from a first position to a second position, wherein
0.245
≤
distance
r
average
distance
R
avg
≤
0.325
.
3. A gas turbine engine as recited in claim 2 , wherein the first position is defined on the leading edge of the nacelle inlet at a 3 o'clock position and the second position is defined on an opposing side of the leading edge of the nacelle at a 9 o'clock position.
4. A gas turbine engine as recited in claim 1 , wherein the bypass duct and the compressor section define a bypass ratio ranging from 10 to 16.
5. A gas turbine engine as recited in claim 1 , further comprising a combustor section and a turbine section wherein the fan section, the compressor section, the combustor section and the turbine section are configured to produce a thrust ranging from 24,000 to 36,000 pounds.
6. A gas turbine engine as recited in claim 1 , further comprising a combustor section and a turbine section wherein the fan section, the compressor section, the combustor section and the turbine section are configured to produce a thrust ranging from 24,000 to 36,000 pounds and wherein the fan section includes a geared fan.
7. A gas turbine engine comprising:
a nacelle defining a centerline axis, the nacelle including:
i. a nacelle inlet;
ii. a nacelle outlet aft of the nacelle inlet; and
iii. a bypass duct therebetween;
a compressor section aft of the nacelle inlet;
a combustor section aft of the compressor section;
a turbine section aft of the combustor section, wherein a fan section, the compressor section, the combustor section and the turbine section are configured to produce a thrust ranging from 24,000 to 36,000 pounds;
an annular splitter separating the bypass duct from the compressor section;
a spinner radially inward of the nacelle forward of the compressor section;
a fan blade platform defined in the fan section aft of the spinner and radially inward of the nacelle; and
a fan blade extending from the fan blade platform toward the nacelle,
wherein a distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform, and wherein a distance H is the radial distance from the first point to the second point, wherein
1.5
≤
distance
X
distance
H
≤
4
for reducing foreign object debris (FOD) intake into the compressor section;
and wherein a distance r is defined radially from the centerline axis to the first point, and an average distance R avg is defined radially from the centerline axis to a leading edge of the nacelle inlet taken over a section of the nacelle ranging from a first position to a second position wherein
0.245
≤
distance
r
average
distance
R
avg
≤
0.325
Also for reducing FOD intake into the compressor section, and
wherein a point Z is defined at an intersection of the centerline axis and a line C normal to the centerline axis extending radially inward from the leading edge of the fan blade where the fan blade meets the fan blade platform, wherein a point W is defined at the intersection of the line C and the leading edge of the fan blade where the fan blade meets the fan blade platform, a distance L is defined from the point Z to a tip of the spinner, wherein a point V is defined along the centerline axis at a distance 0.25 times the distance L aft of the tip of the spinner, wherein a point U is defined at an intersection of a line E normal to the centerline axis extending radially outward from the point V and a line F extending from the point W to the tip of the spinner, wherein a distance M c is defined from the point T to the point U, and wherein a distance M p is defined from the point T to the point V, wherein
distance
Mc
distance
Mp
≤
1
/
2.
8. A gas turbine engine as recited in claim 7 , wherein the first position is defined on the leading edge of the nacelle inlet at a 3 o'clock position and the second position is defined on an opposing side of the leading edge of the nacelle at a 9 o'clock position.
9. A gas turbine engine as recited in claim 7 , wherein the bypass duct and the compressor section define a bypass ratio ranging from 10 to 16.
10. A gas turbine engine as recited in claim 7 , wherein the fan section includes a geared fan.