A gas turbine is provided including a compressor, a combustion chamber, a guide vane row and a rotor airfoil row. The guide vane row includes a plurality of guide vane airfoils having a blade and an inner platform. The ratio between the pitch and the leading edge diameter of the guide vane airfoils is between 6.3-7.6 and the ratio between the platform length and the leading edge diameter of the guide vane airfoils is between 4.0-5.5.
1. Gas turbine comprising at least a combustion chamber ( 2 ), a guide vane row ( 3 ) and a rotor airfoil row, said guide vane row ( 3 ) comprising a plurality of guide vane airfoils comprising a blade ( 7 ) and an inner platform ( 8 ), wherein a ratio between a pitch (P) and a leading edge diameter (D) of the guide vane airfoils is between 6.3-7.6 and a ratio between a platform length (L) and the leading edge diameter (D) of the guide vane airfoils is between 4.0-5.5, the platform length (L) is defined by the axial distance between a leading edge ( 15 ) of a guide vane blade ( 7 ) and an inner guide vane platform inlet ( 16 ) measured at half height of the guide vane blade ( 7 ).
2. Gas turbine as claimed in claim 1 , wherein the ratio between the pitch (P) and the leading edge diameter (D) of the guide vane airfoils is between 6.8-7.0.
3. Gas turbine as claimed in claim 1 , wherein the ratio between the pitch (P) and the leading edge diameter (D) of the guide vane airfoils is between 6.7-7.1.
4. Gas turbine as claimed in claim 1 , wherein the ratio between the platform length (L) and the leading edge diameter (D) of the guide vane airfoils is between 4.5-5.0.
5. Gas turbine as claimed in claim 1 , wherein the ratio between the platform length (L) and the leading edge diameter (D) of the guide vane airfoils is between 4.6-4.8.
6. Gas turbine as claimed in claim 1 , wherein the area of the gases path in the zone of the first guide vane row ( 3 ) continuously decreases.
7. Gas turbine as claimed in claim 6 , wherein an annulus constriction in the zone of the first guide vane row ( 3 ) is between 1.0-1.5, wherein the annulus constriction is defined by a ratio between a hot gases path area at a cross section of the platform inlet ( 16 ) and a hot gases path area at the leading edges ( 15 ) of the guide vane blades ( 7 ).
8. Gas turbine as claimed in claim 6 , wherein an annulus constriction in the zone of the first guide vane row ( 3 ) is between 1.1-1.4, wherein the annulus constriction is defined by a ratio between a hot gases path area at a cross section of the platform inlet ( 16 ) and a hot gases path area at the leading edges ( 15 ) of the guide vane blades ( 7 ).
9. Gas turbine as claimed in claim 6 , wherein an annulus constriction in the zone of the first guide vane row ( 3 ) is between 1.2-1.3, wherein the annulus constriction is defined by a ratio between a hot gases path area at a cross section of the platform inlet ( 16 ) and a hot gases path area at the leading edges ( 15 ) of the guide vane blades ( 7 ).
10. Gas turbine as claimed in claim 1 , wherein the inner platform ( 8 ) of said guide vane airfoils define with an inner wall ( 4 ) of the combustion chamber ( 2 ) an inner gap ( 11 ), wherein the guide vane airfoils include an outer platform ( 9 ) defining, with an outer wall ( 5 ) of the combustion chamber ( 2 ), an outer gap ( 12 ), the inner gap ( 11 ) and the outer gap ( 12 ) are aligned with each other with respect to a plane ( 20 ) perpendicular to the gas turbine axis ( 21 ).