Axial compressor integrated bladed rotor airfoil design scheme

1 . An integrally bladed rotor (IBR) of an aircraft engine, comprising:

a hub, the hub including:

a blade platform; and

a hub web extending radially inwardly from the blade platform; and

blades, each blade comprising:

an airfoil section extending from the blade platform to a blade tip defining a radially outboard extent of the rotor blade to aerodynamically interact with a flow of air for compressing the air; and

a root fillet integrally formed with the blade platform and from which the airfoil section integrally extends;

wherein the blade platform includes:

a radially outer surface from which each blade extends; and

a radially inner surface radially offset from the radially outer surface defining a platform thickness between the radially outer surface and the radially inner surface;

wherein the platform thickness is locally increased at a hub web intersection with the blade platform, defining an increased platform thickness portion, a platform fillet defined at an end of the increased platform thickness portion;

wherein a hub web axial thickness is locally increased at the hub web intersection with the blade platform, defining an increased web thickness portion, a web fillet defined at an end of the increased web thickness portion;

wherein the local increase in platform thickness equals the local increase in hub web axial thickness between the platform fillet and the web fillet.

2 . The IBR of claim 1 , wherein the localized increased platform thickness is centered on and extends axially from a web centerline of the hub web.

3 . The IBR of claim 1 , wherein the localized increased platform thickness extends axially in a range of 50% to 70% of an axial chord length of the blades.

4 . The IBR of claim 1 , wherein the platform thickness is locally increased by 50% to 100% compared to one of a leading edge platform thickness and a trailing edge platform thickness.

5 . The IBR of claim 1 , wherein the localized increase of hub web axial thickness extends radially inwardly along a range of 10% to 30% of a radial hub web length.

6 . The IBR of claim 1 , wherein the localized increase of the hub web axial thickness is axially centered on a web centerline of the hub web.

7 . A compressor assembly of a gas turbine engine of an aircraft, comprising one or more compressor rotors arrayed along an engine central longitudinal axis, at least one compressor rotor of the one or more compressor rotors is an integrally bladed rotor (IBR) including:

a hub, the hub including:

a blade platform; and

a hub web extending radially inwardly from the blade platform; and

blades, each blade comprising:

an airfoil section extending from the blade platform to a blade tip defining a radially outboard extent of the rotor blade to aerodynamically interact with a flow of air for compressing the air; and

a root fillet integrally formed with the blade platform and from which the airfoil section integrally extends;

wherein the blade platform includes:

a radially outer surface from which each blade extends; and

a radially inner surface radially offset from the radially outer surface defining a platform thickness between the radially outer surface and the radially inner surface;

wherein the platform thickness is locally increased at a hub web intersection with the blade platform, defining an increased platform thickness portion, a platform fillet defined at an end of the increased platform thickness portion;

wherein a hub web axial thickness is locally increased at the hub web intersection with the blade platform, defining an increased web thickness portion, a web fillet defined at an end of the increased web thickness portion;

wherein the local increase in platform thickness equals the local increase in hub web axial thickness between the platform fillet and the web fillet.

8 . The compressor assembly of claim 7 , wherein the localized increased platform thickness is centered on and extends axially from a web centerline of the hub web.

9 . The compressor assembly of claim 7 , wherein the localized increased platform thickness extends axially in a range of 50% to 70% of an axial chord length of the blades.

10 . The compressor assembly of claim 7 , wherein the platform thickness is locally increased by 50% to 100% compared to one of a leading edge platform thickness and a trailing edge platform thickness.

11 . The compressor assembly of claim 7 , wherein the localized increase of hub web axial thickness extends radially inwardly along a range of 10% to 30% of a radial hub web length.

12 . The compressor assembly of claim 7 , wherein the localized increase of the hub web axial thickness is axially centered on a web centerline of the hub web.

13 . A gas turbine engine of an aircraft, comprising:

a turbine driven by products of a combustor; and

a compressor operably connected to the turbine and driven by rotation of the turbine, the compressor including an integrally bladed rotor (IBR), the IBR including:

a hub, the hub including:

a blade platform; and

a hub web extending radially inwardly from the blade platform; and

blades, each blade comprising:

an airfoil section extending from the blade platform to a blade tip defining a radially outboard extent of the rotor blade to aerodynamically interact with a flow of air for compressing the air; and

a root fillet integrally formed with the blade platform and from which the airfoil section integrally extends;

wherein the blade platform includes:

a radially outer surface from which each blade extends; and

a radially inner surface radially offset from the radially outer surface defining a platform thickness between the radially outer surface and the radially inner surface;

wherein the platform thickness is locally increased at a hub web intersection with the blade platform, defining an increased platform thickness portion, a platform fillet defined at an end of the increased platform thickness portion;

wherein a hub web axial thickness is locally increased at the hub web intersection with the blade platform, defining an increased web thickness portion, a web fillet defined at an end of the increased web thickness portion;

wherein the local increase in platform thickness equals the local increase in hub web axial thickness between the platform fillet and the web fillet.

14 . The gas turbine engine of claim 13 , wherein the localized increased platform thickness is centered on and extends axially from a web centerline of the hub web.

15 . The gas turbine engine of claim 13 , wherein the localized increased platform thickness extends axially in a range of 50% to 70% of an axial chord length of the blades.

16 . The gas turbine engine of claim 13 , wherein the platform thickness is locally increased by 50% to 100% compared to one of a leading edge platform thickness and a trailing edge platform thickness.

17 . The gas turbine engine of claim 13 , wherein the localized increase of the hub web axial thickness is axially centered on a web centerline of the hub web.