| 摘要 |
The invention concerns a supporting structure (37, 37') for a gas turbine engine (1, 100), said structure (37, 37') comprising an inner ring (10), an outer ring (11), and a plurality of circumferentially spaced, load carrying radial elements (12) connecting the inner and outer rings (10, 11), said radial elements (12) being configured to transfer loads between the inner ring (10) and the outer ring (11), wherein a gas channel for a primary axial gas flow is defined between the inner and outer rings (10, 11), wherein the supporting structure (37, 37') has an inlet side for primary gas flow entrance and an outlet side for primary gas outflow, wherein the radial elements (12) have an airfoil shape with a leading edge (121) directed towards the inlet side, a trailing edge (122) directed towards the outlet side, and two opposite sides (123, 124) connecting the leading edge (121) and the trailing edge (122), wherein the locus of points midway between the two opposite sides (123, 124) forms a mean camber line of each radial element (12), wherein at least one of the radial elements (12) comprises a gas passage arrangement configured to lead a separate bleeding gas flow from the supporting structure (37, 37'). The invention is characterized in that the gas passage arrangement comprises a radially extending gas channel (32) arranged inside the radial element (12) and at least one opening (34) in communication with said gas channel (32), wherein the at least one opening (34) is arranged at one of the sides (123, 124) of the radial element (12). The invention also concerns a gas turbine engine (1, 100) comprising a supporting structure (37, 37') of the above type. |
