This section includes elementary examples of case studies beginning with 2D cases, as shown in Figure 14.5. The first diagram represents an aerofoil showing the grid layout.
The domain of analysis is large with the anisotropic adaptive grid (Figure 14.5a), which is more dense the closer it is to the LE and trailing edge matching the surface grids and where shocks are present. When the solver has been run, the results can be seen in the postprocessor showing the Mach number isolines (Figure 14.5b). In another run with a different setup, the results are shown in a color spectrum (i. e., the gray-scale version in Figure 14.5c).
The next example is a simple, isolated 3D wing, as shown in Figure 14.6a. Half is shown with a simple grid and the other half is shown in shaded geometry. The drag polar from the CFD analysis is compared with results in Figure 14.6b.
CFD analysis of an isolated fuselage should be easy but internal and external flow through the nacelle (Figure 14.7) can prove to be difficult.
Whereas CFD studies on aerofoils exist, flow-field analysis on nacelles is rare. Chen  et al. presented a flow-field analysis over a symmetric isolated nacelle
(a) Isolated Wing Geometry (b) Comparison of the CFD Analysis
Figure 14.6. CFD analysis of a wing ( – Henley innovations)
Figure 14.7. Nacelle grids for internal and external flow analysis
using a Euler solver (Figure 14.8). Subsequent studies by Uanishi  et al. showed confirmation of the velocity field obtained by Chen. No work has been found for velocity fields over the nacelle using Navier-Stokes solvers.
In a more recent analysis , it is stated that “…the observed scatter in the absolute CFD-based drag estimates is still larger than the desired single drag count error margin that is defined for drag prediction work. Yet, the majority of activities conducted during an aircraft development program are incremental in nature, i. e., testing/computing a number of options and looking for the best relative performance.”