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A gas turbine engine disk that includes a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end. The disk hub has a radially-displaced annular hub surface exposed to high pressure, high temperature discharge gases during engine operatio

A gas turbine engine disk that includes a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end. The disk hub has a radially-displaced annular hub surface exposed to high pressure, high temperature discharge gases during engine operation. The radially-displaced annular hub surface acts as an axial free surface mitigating the formation undesirable axial stress in the disk hub.

대표청구항 ▼

What is claimed is: 1. A gas turbine engine, comprising a high pressure turbine section, including a gas turbine engine disk defining a disk hub surface having an annular cylindrical surface section and a radially-displaced annular hub surface section comprising an axially-extending and radially-di

What is claimed is: 1. A gas turbine engine, comprising a high pressure turbine section, including a gas turbine engine disk defining a disk hub surface having an annular cylindrical surface section and a radially-displaced annular hub surface section comprising an axially-extending and radially-displaced chamfer for disrupting an axial path from the disk hub to the disk hub surface, the chamfer formed in a forward side of the disk hub and extending forwardly from the annular cylindrical surface section at a position below and in vertical alignment with a geometrically-centered, lowest internal temperature portion of the disk hub, the annular cylindrical surface section extending axially to a rearward side of a rearward shaft for shifting the lowest internal temperature portion and the maximum axial stress out of vertical alignment with the geometrically-centered portion of the disk hub, thereby reducing the magnitude of the axial stress on the disk hub. 2. A gas turbine engine disk according to claim 1, wherein the chamfer of the radially-displaced annular hub surface section extends axially at least one-half of the disk hub surface, and wherein the chamfer is planar and between 0 and 50 degrees. 3. A gas turbine engine disk, comprising: (a) a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end; and (b) the disk hub defining a disk hub surface having an annular cylindrical surface section and a radially-displaced annular hub surface section exposed to high pressure, high temperature discharge gases during engine operation, the radially-displaced annular hub surface section comprising a first concave annular recess having a shape defined by a first radius and a second concave annular recess axially spaced-apart from the first concave annular recess and having a shape defined by a second radius different than the first radius, wherein the first concave annular recess and the second concave annular recess define a convex annular member formed in the hub surface that joins the first concave annular recess and the second concave annular recess. 4. A method of reducing axial stress in a disk hub of a gas turbine disk, comprising the steps of: (a) determining the location of stress gradients in a gas turbine disk hub; and (b) forming a radially-displaced, axially-extending annular surface section in a disk hub surface of the disk hub that axially displaces the axial stress relative to with a geometrically-centered, lowest internal temperature portion of the disk hub for shifting maximum axial stress on the disk hub out of vertical alignment with a geometrically-centered portion of the disk hub, thereby reducing the magnitude of the axial stress on the disk hub, the radially-displaced, axially-extending annular surface section comprising a first concave annular recess having a shape defined by a first radius and a second concave annular recess axially spaced-apart from the first concave annular recess and having a shape defined by a second radius different than the first radius, wherein the first concave annular recess and the second concave annular recess define a convex annular member formed in the hub surface that joins the first concave annular recess and the second concave annular recess. 5. A gas turbine engine disk comprising: (a) a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end; and (b) the disk hub defining a disk hub surface having an annular cylindrical surface section and a radially-displaced annular hub surface section exposed to high pressure, high temperature discharge gases during engine operation, the radially-displaced annular hub surface section comprising at least one radially-extending annular continuously curved convex ring and acting as an axial free surface mitigating the formation of axial stress in the disk hub. 6. A method of reducing axial stress in a disk hub of a gas turbine disk, comprising the steps of: (a) determining the location of stress gradients in a gas turbine disk hub; and (b) forming a radially-displaced, axially-extending annular surface section in a disk hub surface of the disk hub that axially displaces the axial stress relative to with a geometrically-centered, lowest internal temperature portion of the disk hub for shifting maximum axial stress on the disk hub out of vertical alignment with a geometrically-centered portion of the disk hub, thereby reducing the magnitude of the axial stress on the disk hub, the radially-displaced, axially-extending annular surface section comprising at least one radially-extending annular continuously curved convex ring and acting as an axial free surface mitigating the formation of axial stress in the disk hub.