Gas turbine axial compressor

An axial compressor for a gas turbine, in particular an aircraft engine, having at least one rotor blade or guide vane having a blade or vane element, which is arranged in the flow duct, and a leading edge and a trailing edge, which are joined to each other through a pressure side and a suction side, wherein,

• in each profile section of the blade or vane element in a range between 5% and 95% of a radial blade or vane element height from a blade or vane element root to a blade or vane tip,
• a profile skeletal line has at most two points of inflection;
• a metal angle at the leading edge lies in a range between 20° and 75°;
• a metal angle at the trailing edge lies in a range between −20° and 70° and is at most equal to the metal angle at the leading edge;
• and
• a stacking angle lies in a range between 0° and 70°;
• and wherein
• a local or absolute minimum of a front load angle, which is equal to the difference obtained from the arithmetic mean of the metal angles at the leading edge and trailing edge as minuend and the stacking angle as subtrahend
• (
• λ
• =
• α
• le
• +
• α
• te
• 2
• -
• β
• )
• lies in a range between 55% and 85% of the radial blade or vane element height;
• and a local or absolute maximum of the front load angle lies in a range between 95% and 100% of the radial blade or vane element height at a blade or vane tip.