Abstract: In order to improve the accuracy of simulation and obtain the aerodynamic heating consistency of the prediction method for the typical integrated design of the aircraft forebody model, experiments were carried out at FD−20a shock tunnel under the condition of Ma = 6, Re = 1.14 × 10⁷～2.98 × 10⁷ m⁻¹, α = 0°～8°. Numerical simulations were applied through the compressible Navier−Stokes equation implemented with the finite volume method, Roe’s flux difference splitting scheme, LU−SGS spatial method and Spalart–Allmaras (SA) turbulence model. Simulation results were compared with the experimental data to validate the prediction methods. Results show that with the increase of the complexity and disturbance intensity, the heat flux consistency decreases. The compression surface flow was dominated by the attached flow and small separation, where relative good consistency was found between the simulation and experimental data with the average difference being about 22.3%. The throat boundary layer was interfered by shock waves, and 43.5% difference was found between the heat flux simulation and experimental data. The heat flux difference in the isolation section increased to 31.8%. When the oblique shock impinged on the subsonic part of the bow shock and three dimensional flow patterns were obvious, the difference of heat flux reached the maximum of around 100% in three dimensional areas. Mesh, simulation methods and unsteady characteristics are concluded as the reasons for heat flux consistency decreasing along the airflow.