The results of the flow investigation at the supersonic combustor entrance at tests in the regime of an attached pipe-line are presented. As a source of working gas, a discharge pre-chamber of an impulse wind tunnel was used. The three-dimensional numerical simulation was performed using the software package ANSYS Fluent from the nozzle critical section to the combustor entrance with experimental initial data. The experiments were carried out with the following airflow parameters: M = 4, total pressure P0 = 5.7÷2.0 MPa and total temperature T0=1750÷1200 K. Air flow characteristics and their distribution at the combustor entrance were obtained. It is established that the Mach number field non-uniformity in the flow core at the combustor entrance is 2.3% with a maximum thickness of the boundary layer of 10.7 mm. A good agreement between the calculated and experimental data was obtained.