Compressible Flow: F-35

Description

The Lockheed Martin F-35 Lightning II is an American family of single-seat, single-engine, all-weather stealth multirole combat aircraft, also capable of electronic warfare and intelligence, surveillance, and reconnaissance missions.

Notably, the F-35 can reach speeds of around 500 m/s, placing the surrounding airflow firmly in the supersonic regime. Since full-scale wind tunnel experiments at these conditions are costly in both time and money, CFD solvers are frequently used for initial evaluation. This project studies the supersonic, compressible flow around an F-35 aircraft. The geometry consists of a 20 m F-35 positioned inside a 150 m wind tunnel.

The mesh contains 7,182,542 elements. In terms of quality, a maximum skewness of 0.79 with an average of 0.22 is satisfactory for this problem. To resolve the boundary layer accurately, 25 prism layers were added adjacent to both the wind tunnel walls and the aircraft body. The mesh was generated in ANSYS Meshing and subsequently converted to a polyhedral mesh within ANSYS Fluent, reducing the count to 1,845,364 elements while preserving the same quality. As with any numerical study, the first step in the modeling was the creation of the CAD geometry.

Methodology

Air is treated as a compressible ideal gas, and a Mach number of 2.0 is reached at the maximum speed of 544 m/s. Solving this problem requires the flow equations to be handled in their differential form.

A non-isothermal, compressible ideal-gas condition was assumed inside the wind tunnel; consequently, the energy equation was solved alongside the flow and turbulence equations. The governing mass and momentum equations are written in their standard conservative form for compressible flow.

Conclusion

The drag force and the shock profiles were obtained over the course of the study. After the solution converged, the results were examined through post-processing. As a check on convergence, the drag value was monitored throughout the solution iterations: the solution was deemed converged once the drag force settled to a constant value and the residuals dropped below 10⁻⁶.

The results are then presented for the pressure and velocity fields. The shock profile is visible in both the pressure and Mach number contours, while the velocity field is shown through both contours and streamlines to give deeper insight into the flow. The temperature gradient and its variation across different locations are also presented, since the temperature rise is an important factor in compressible aerodynamic calculations. Finally, the drag force was calculated at 181.66 kN, a reasonable value for a 20 m aircraft with the stated specifications.