F-35 Considering Compressible Flow, ANSYS Fluent CFD Simulation Training

Description

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

Interestingly, the F-35 speed could reach 500 m/s, which results in supersonic airflow. Therefore, actual case wind tunnel experiments are expensive in terms of both costs and time, so CFD solvers are often employed to evaluate for initial tests.

This project is going to study a Supersonic Compressible Flow adjacent to an F-35 plane. The geometry is a 20-meter F-35 plane inside a 150-meter wind tunnel.

For the current problem, the element number is equal to 7,182,542. Regarding the quality of the mesh, the maximum skewness of 0.79 with an average of 0.22 is a satisfactory mesh for the current problem.

Also, 25 prism layers were added adjacent to both water tunnel walls and the turbine’s body to accurately calculate the boundary layer.

Finally, the mesh is generated through ANSYS-Meshing. Finally, the mesh was converted to Polyhedral in the ANSYS FLUENT software, and it was reduced to 1,845,364 elements with the same quality. As a numerical study, the initial step towards the modeling is the production of the CAD geometry.

F-35 Methodology

The air is considered a compressible ideal gas, and the Mach number of 2.0 was achieved at the maximum speed of 544 m/s. To study the current problem, one must solve the flow equations in the differential form.

Also, the non-isothermal, compressible ideal gas condition was assumed inside the wind tunnel. Therefore, the energy equation was also solved in addition to the flow and the turbulent equations. Briefly, the governing mass and momentum equations are written as follows:

F-35 Conclusion

Also, the drag force and the shock profiles were obtained during the study. After the solution convergence, we observe the results through post-processing. Meanwhile, as an assurance of a valid convergence, we monitor the drag value during the solution iterations.

In this study, the solution was a converged one when the drag force reached a constant rate, and the residuals were below 10-6 values. Afterward, we present the results regarding the pressure and the velocity field in the figures.

Furthermore, we can observe the shock profile from both the pressure and the Mach number contours. For the velocity field, we represent both contour and streamlines to give much insight into the problem.

Also, we present the temperature gradient and its variations in different locations because the increase in temperature is important in aerodynamic compressible flow calculations. Finally, we calculate the drag force at 181.66 (kN), which is accurate for a 20-meter airplane with the noted specifications.