Computational Fluid Dynamics Validation of a Single Central Nozzle Supersonic Retropropulsion Configuration

Supersonic retropropulsion provides an option that can potentially enhance drag characteristics of high mass entry, descent, and landing systems. Preliminary flow field and vehicle aerodynamic characteristics have been found in wind tunnel experiments; however, these only cover specific vehicle configurations and freestream conditions. In order to generate useful aerodynamic data that can be used in a trajectory simulation, a quicker method of determining vehicle aerodynamics is required to model supersonic retropropulsion effects. Using computational fluid dynamics, flow solutions can be determined which yield the desired aerodynamic information. The flow field generated in a supersonic retropropulsion scenario is complex, which increases the difficulty of generating an accurate computational solution. By validating the computational solutions against available wind tunnel data, the confidence in accurately capturing the flow field is increased, and methods to reduce the time required to generate a solution can be determined. Fun3D, a computational fluid dynamics code developed at NASA Langley Research Center, is capable of modeling the flow field structure and vehicle aerodynamics seen in previous wind tunnel experiments. Axial locations of the jet terminal shock, stagnation point, and bow shock show the same trends which were found in the wind tunnel, and the surface pressure distribution and drag coefficient are also consistent with available data. The flow solution is dependent on the computational grid used, where a grid which is too coarse does not resolve all of the flow features correctly. Refining the grid will increase the fidelity of the solution; however, the calculations will take longer if there are more cells in the computational grid. Nomenclature A = model base area P exit = nozzle exit pressure A exit = nozzle exit area P ∞ = freestream pressure C D = drag coefficient P t,jet = nozzle total pressure C P = pressure coefficient q ∞ = freestream dynamic pressure C T = thrust coefficient R = specific gas constant γ = ratio of specific heats T = thrust M exit = nozzle exit Mach number T ∞ = freestream temperature ρ = density T jet = nozzle total temperature ρ ∞ = freestream density u = nozzle exit axial velocity component P = pressure V ∞ = freestream velocity P 0,jet = nozzle total pressure y = nozzle exit radial position component BFI = blunt flow interaction CFD = computational fluid dynamics EDL = entry, descent, and landing LJP = long jet penetration NASA = …