Experimental study on the flow field inside the nozzle for radial turbine was performed. At design point, the pressure is high and the Mach number is low at the pressure side of the nozzle inlet semi-vaneless space as the flow turns through the nozzle vanes. As the flow accelerates through the nozzle passage to the throat the pressure level at the pressure and suction sides becomes similar. The flow continued accelerating from the throat to the inlet of turbine wheel and the pressure field became uniform in the circumferential direction in the vaneless space. In high expansion ratio condition, strong favorable pressure gradient band region occurred just after the throat in the semi-vaneless space in the circumferential direction and the pressure became uniform in the circumferential direction after this band. In low expansion ratio condition, core flow acceleration is dominant after the throat and this non-uniform pressure field reached to the inlet of turbine wheel.