Abstract: Injection element of coaxial design for rocket combustion chambers with a combustion chamber pressure of 3 to 30 bar and for operation with two hypergolically reacting propellants, with a central body for the oxidant, whose flow channel has a twisting insert, a cross section contraction and a sharp opening edge, and with a sleeve concentrically surrounding the central body to limit an annular flow channel for the propellant, wherein the flow channel has a circular cylindrical outlet with a sharp opening edge. The flow channel of the central body has a circular cylindrical outlet. The opening edges of the central body and of the sleeve are located approximately in the plane of the front surface of the combustion chamber. A flow divider, which divides the propellant flow into five to twenty equal individual flows is arranged in the flow channel for the propellant. The velocities of flow of the oxidant and of the propellant are between 5 m/sec and 20 m/sec at the point of entry into the combustion chamber.