Granted Patent B2

US 12,385,449 · App. 18/461,709 · Granted Aug 12, 2025

Inventors: Daniel Bernard Kupratis (Wallingford, CT); Paul R. Hanrahan (Farmington, CT)

Assignee: RTX CORPORATION

F02K3/077F02K3/075F01D5/145F02K1/72F02K3/06F04D29/542F04D29/661

Abstract

Claims (42)

1. A section for a gas turbine engine comprising:

a rotor including a row of blades extending in a radial direction outwardly from a hub, wherein the hub is rotatable about a longitudinal axis such that the row of blades deliver flow to a bypass flow path, an intermediate flow path and a core flow path;

a first case surrounding the row of blades to establish the bypass flow path through a first duct;

a first flow splitter that divides flow between the bypass flow path and a second duct;

an aftmost row of guide vanes extending in the radial direction across the bypass flow path;

a second case including a second flow splitter radially inboard of the first flow splitter, wherein the second flow splitter is axially forward of an axially forwardmost compressor blade row of the gas turbine engine relative to the longitudinal axis that is surrounded by the second case, the second flow splitter divides flow from the second duct between the intermediate flow path and the core flow path, a forwardmost edge of the second flow splitter is axially aft of a forwardmost edge of the first flow splitter with respect to the longitudinal axis, and the forwardmost edges of the first and second flow splitters are axially forward of the aftmost row of guide vanes with respect to the longitudinal axis;

a bypass port that interconnects the intermediate and bypass flow paths at a position axially aft of a position of the aftmost row of guide vanes relative to the longitudinal axis, and the bypass port configured to direct flow from the intermediate flow path to the bypass flow path;

wherein each of the blades extends in the radial direction between a 0% span position at the hub and a 100% span position at a tip, leading edges of the blades at the 100% span position are established along a first reference plane, leading edges of the guide vanes at a 100% span position of the respective guide vanes are established along a second reference plane, the first and second reference planes are perpendicular to the longitudinal axis, a first axial length is established between the first and second reference planes, a tip radius is established between the tip of one of the respective blades and the longitudinal axis, and a ratio of the first axial length divided by the tip radius is greater than or equal to 0.5; and

wherein the forwardmost edge of the first flow splitter is situated in the radial direction at a first splitter position, the forwardmost edge of the second flow splitter is situated in the radial direction at a second splitter position, the first splitter position is radially aligned with or radially inward of a 25% span position of the blades, and the second splitter position being radially aligned with or radially outward of a 5% span position of the blades.

2. The section as recited in claim 1 , wherein the hub is driven by a turbine through a geared architecture.

3. The section as recited in claim 1 , wherein each of the blades extends in the radial direction between the 0% span position at the hub and the 100% span position at the tip to establish a hub-to-tip ratio, and the hub-to-tip ratio of each one of the blades is less than or equal to 0.4, the hub-to-tip ratio of each one of the blades measured relative to the forwardmost portion of the leading edge of the respective blade.

4. The section as recited in claim 1 , wherein the first duct establishes the bypass flow path, the second duct branches between the intermediate and core flow paths at the second flow splitter, a first annulus area is established by the first duct at the forwardmost edge of the first flow splitter, a second annulus area is established by the second duct at the forwardmost edge of the first flow splitter, a bypass area ratio is defined as the first annulus area divided by the second annulus area, and the bypass area ratio is greater than or equal to 10, and less than or equal to 35.

5. The section as recited in claim 4 , wherein a maximum value of a pressure ratio across the blades alone is less than 1.45 at cruise at 0.8 Mach and 35,000 feet.

6. The section as recited in claim 4 , wherein only one row of guide vanes extends in the radial direction across the bypass flow path and the row of guide vanes comprises the aftmost row of guide vanes.

7. The section as recited in claim 4 , wherein each of the blades extends in the radial direction between the 0% span position at the hub and the 100% span position at the tip to establish a hub-to-tip ratio, and the hub-to-tip ratio of each one of the respective blades is between 0.16-0.36 measured relative to the forwardmost portion of the leading edge of the respective blade.

8. The section as recited in claim 7 , wherein the row of blades establishes a blade quantity (BQ), the blade quantity (BQ) is at least 12 but not more than 20 blades, the aftmost row of guide vanes establishes a vane quantity (VQ), and the vane quantity (VQ) is at least 20 but not more than 40 guide vanes.

9. The section as set forth in claim 8 , wherein a ratio of VQ/BQ is between 2.0 and 2.6.

10. The section as recited in claim 1 , further comprising:

one or more blocker doors situated in the intermediate flow path downstream of the bypass port, and wherein the one or more blocker doors are moveable to direct flow from the intermediate flow path to the bypass port.

11. The section as set forth in claim 10 , wherein:

the one or more blocker doors are configured to pivot from the second case towards the bypass flow path to direct flow from the intermediate flow path to the bypass port; and

louvers are situated in the bypass port, the louvers have an airfoil-shaped geometry and are arranged to direct flow from the intermediate flow path to the bypass flow path.

12. A gas turbine engine comprising:

a fan section including a fan having a row of blades, each extending in a radial direction between a 0% span position at a hub and a 100% span position at a respective tip, wherein the hub is rotatable about an engine longitudinal axis such that the row of blades deliver flow to a bypass flow path, an intermediate flow path and a core flow path;

a compressor section establishing the core flow path;

a turbine section that drives the fan section and the compressor section;

a fan case including a bypass duct surrounding the row of blades to establish the bypass flow path;

a housing including a first flow splitter that divides flow between the bypass flow path and a second duct;

an aftmost row of guide vanes in the bypass duct that extend in the radial direction across the bypass flow path;

an engine case including a second flow splitter radially inboard of the first flow splitter and that divides flow from the second duct between the intermediate flow path and the core flow path, wherein the second flow splitter is axially forward of an axially forwardmost compressor blade row of the compressor section that is surrounded by the engine case relative to the engine longitudinal axis, a forwardmost edge of the second flow splitter is axially aft of a forwardmost edge of the first flow splitter with respect to the engine longitudinal axis, and the forwardmost edges of the first and second flow splitters are axially forward of the row of aftmost guide vanes with respect to the engine longitudinal axis;

a bypass port that interconnects the intermediate and bypass flow paths at a position downstream of the aftmost row of guide vanes, and the bypass port configured to direct flow from the intermediate flow path to the bypass flow path;

wherein each of the blades extends in the radial direction between the 0% span position and 100% span position to establish a hub-to-tip ratio, and the hub-to-tip ratio of each one of the blades is less than or equal to 0.4 measured relative to a forwardmost portion of a leading edge of the respective blade;

the leading edges of the blades at the 100% span position are established along a first reference plane, leading edges of the guide vanes at a 100% span position of the respective guide vanes are established along a second reference plane, the first and second reference planes are perpendicular to the engine longitudinal axis, a first axial length is established between the first and second reference planes, a tip radius is established between the tip of one of the respective blades and the longitudinal axis, and a ratio of the first axial length divided by the tip radius is greater than or equal to 0.5; and

wherein the forwardmost edge of the first flow splitter is situated in the radial direction at a first splitter position, the forwardmost edge of the second flow splitter is situated in the radial direction at a second splitter position, the first splitter position is radially aligned with or radially inward of a 25% span position of the blades relative to the engine longitudinal axis, and the second splitter position is radially aligned with or radially outward of a 5% span position of the blades relative to the engine longitudinal axis.

13. The gas turbine engine as recited in claim 12 , further comprising a geared architecture, wherein the turbine section includes a fan drive turbine, and the geared architecture drives the fan at a different speed than a speed of the fan drive turbine.

14. The gas turbine engine as recited in claim 12 , further comprising one or more blocker doors situated in the intermediate flow path downstream of the bypass port, wherein the one or more blocker doors are moveable between a first position and a second position, wherein the one or more blocker doors are configured to communicate flow in the intermediate flow path to a vent port in the first position but are configured to block flow in the intermediate flow path to the vent port in the second position.

15. The gas turbine engine as recited in claim 14 , wherein the vent port is axially aft of an exit of the bypass flow path with respect to the engine longitudinal axis.

16. The gas turbine engine as recited in claim 14 , further comprising a thrust reverser including a cascade that selectively communicates airflow from the bypass duct, wherein the cascade extends axially aft of both the aftmost row of guide vanes and the bypass port with respect to the engine longitudinal axis.

17. The gas turbine engine as recited in claim 12 , wherein:

the compressor section includes a first compressor and a second compressor, the first compressor comprises the axially forwardmost compressor blade row of the compressor section, and both the second flow splitter and an entrance of the intermediate flow path are axially forward of the compressor section relative to the engine longitudinal axis.

18. The gas turbine engine as recited in claim 17 , wherein the first compressor includes three compressor blade rows.

19. The gas turbine engine as recited in claim 17 , wherein the aftmost row of guide vanes extend in the radial direction across the intermediate flow path relative to the engine longitudinal axis.

Continuity (2)

References Cited (400)