Lobe lance for a gas turbine combustor

The invention relates to a fuel lance (21) for injecting a gaseous and/or liquid fuel mixed with air into an axial hot gas flow flowing through a sequential combustor of a gas turbine, said fuel lance (21) comprising at least one finger (22) extending in a longitudinal direction into said axial hot gas flow of said gas turbine essentially perpendicular to said hot gas flow; wherein said at least one finger (22) is configured as a streamlined body which has a streamlined cross-sectional profile; wherein said body has two lateral surfaces essentially parallel to said axial hot gas flow, joined at their upstream side by a leading edge (23) and joined at their downstream side forming a trailing edge (24);wherein a plurality of nozzles (27) for injecting a gaseous and/or liquid fuel mixed with air are distributed along said trailing edge (24);wherein means (28) for improving the mixing quality and reducing pressure loss in said sequential combustor are provided in the trailing edge region of said body; wherein said body comprises an enclosing outer wall (75) defining said streamlined cross-sectional profile, wherein within said outer wall (75) there is provided a longitudinally extending air plenum (31) for the distributed introduction of air into said at least one finger (22), said air plenum (31) having a distance from said outer wall (75), defining a first interspace; and wherein said air plenum (31) is provided with a plurality of distributed impingement cooling holes (66), such that air exiting through said impingement cooling holes (66) impinges on the inner side of the leading edge region of said body. The cooling is improved by said impingement cooling holes (66) each having a hole diameter between 1.2 and 1.8mm; and said impingement cooling holes (66) being arranged at said air plenum (31) with a pitch ratio, i.e. the ratio of the distance between said holes and the hole diameter, between 3 and 10.