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 poi
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 XH≥1.5 for reducing foreign object debris (FOD) intake into the compressor section.
대표청구항▼
1. A gas turbine engine comprising: a nacelle, the nacelle including:i. a nacelle inlet;ii. a nacelle outlet aft of the nacelle inlet; andiii. 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
1. A gas turbine engine comprising: a nacelle, the nacelle including:i. a nacelle inlet;ii. a nacelle outlet aft of the nacelle inlet; andiii. 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; anda 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≤distanceXdistanceH≤4for reducing foreign object debris (FOD) intake into the compressor section; andwherein 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 Mc is defined from the point T to the point U, and wherein a distance Mp is defined from the point T to the point V, wherein distanceMcdistanceMp≤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 ravg 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≤distanceraveragedistanceRavg≤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; andiii. 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; anda 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≤distanceXdistanceH≤4for 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 Ravg 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≤distanceraveragedistanceRavg≤0.325Also for reducing FOD intake into the compressor section, andwherein 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 Mc is defined from the point T to the point U, and wherein a distance Mp is defined from the point T to the point V, wherein distanceMcdistanceMp≤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.
연구과제 타임라인
LOADING...
LOADING...
LOADING...
LOADING...
LOADING...
이 특허에 인용된 특허 (3)
Gilchrist Alan R. (Fairfield OH) Sullivan Thomas J. (Fairfield OH) Walker Roger C. (Middletown OH), Hybrid spinner nose configuration in a gas turbine engine having a bypass duct.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.