Aeropropulsive flows may contain localized separated flow regions whose analysis using RANS methodology may be inadequate, particularly so for massively separated flows having unsteady/oscillatory features. DES-type methods provide for the improved analysis of these local regions, and can often demarcate the zones to be analyzed using RANS or LES methodology in an automated fashion. The DDES implementation described utilizes an extended k-ε model in RANS regions containing a number of unique features catered to analyzing aeropropulsive flows. Several aeropropulsive cases of interest are considered for which test data is available. Cases include: a supersonic base flow calculation; a transverse supersonic air jet interacting with a subsonic airtstream; and two angled fuel simulant jets (M j = 2.5 helium jet and M j = 1.0 ethylene jet) injected into a M ∞ = 2.0 airstream. The results obtained for the supersonic base flow case as well for the transverse air jet in a crossflow case, both compare favorably with the experimental measurements. The predictions for the helium injection case also agree reasonably well with the experimental data. The ethylene case case was not as well predicted which may be due to the RANS representation of the approach boundary layer for this low momentum ratio case (as also pointed out by other investigators).

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Supersonic impinging jet flowfields contain self-sustaining acoustic feedback features that create high levels of discrete frequency tonal noise. These types of flowfields are typically found with short takeoff and landing military aircraft as well as jet blast deflector operations on aircraft carrier decks. The US Navy has a goal to reduce the noise generated by these impinging jet configurations and is investing in computational aeroacoustics to aid in the development of noise reduction concepts. In this paper, implicit Large Eddy Simulation (LES) of impinging jet flow-fields are coupled with a far-field acoustic transformation using the Ffowcs Williams and Hawkings (FW-H) equation method. The LES solves the noise generating regions of the flow in the nearfield, and the FW-H transformation is used to predict the far-field noise. The noise prediction methodology is applied to a Mach 1.5 vertically impinging jet at a stand-off distance of five nozzle throat diameters. Both the LES and FW-H acoustic predictions compare favorably with experimental measurements. Time averaged and instantaneous flowfields are shown. A calculation performed previously at a stand-off distance of four nozzle throat diameters is revisited with adjustments to the methodology including a new grid, time integrator, and longer simulation runtime. The calculation exhibited various feedback loops which were not present before and can be attributed to an explicit time marching scheme. In addition, an instability analysis of two heated jets is performed. Tonal frequencies and instability modes are identified for the sample problems.

Current CFD models fail to accurately predict boundary layer asymmetry on spin-stabilized projectiles, particularly in the transonic and subsonic flow regimes. Consequently, these models cannot accurately characterize the Magnus moment, a key component in aerodynamic behavior. This work seeks to capture boundary layer thickness asymmetry, an indicator of Magnus effects, around a spinning projectile using Magnetic Resonance Velocimetry (MRV). The MRV technique allows for collection of three-dimensional, non-intrusive, high-resolution velocity field measurements that can be used for comparison to and validation of current computational models. In this experiment, a modified M80 projectile was designed to thicken the hydrodynamic boundary layer for technique validation. The scaled projectile was mounted in a custom-designed test rig at a 2° nosedown angle of attack. The apparatus rotated the projectile at various spin rates in a constant flow of copper-sulfate solution. Initial results revealed azimuthal differences in boundary layer thickness for three different cases — no spin, nominal spin, and double spin — particularly in the tapered rear (boattail) region of the projectile. The introduction of spin shifted the boundary layer thickness in the spin direction resulting in lateral boundary layer asymmetry in the boattail region, a phenomenon that likely affects the stability of spin-stabilized projectiles.

Film cooling holes with a compound angle are commonly used on high pressure turbine components in lieu of axial holes to improve effectiveness or as a result of manufacturing constraints. Whereas large eddy simulation (LES) of axial holes is becoming more common place, assessment of LES predictive ability for compound angle hole has been limited. For this study, the selected compound angle round (CAR) hole configuration has a 30 degree injection angle, a 45 degree compound angle, and a density ratio of 1.5. The geometry, flow conditions, and experimental adiabatic effectiveness validation data are from McClintic et al. [28]. The low free stream Mach number of the experiment puts the flow in the incompressible regime. Two LES solvers are evaluated, Fluent and FDL3Di, on structured meshes with a range of blowing ratios simulated for plenum, inline coolant crossflow, and counter coolant crossflow feed holes. When a steady inlet profile is used for the main flow, LES agreement with the data is poor. The inclusion of a resolved turbulent boundary layer significantly improves the predictive quality for both solvers; consequently, resolved inflow turbulence is a required aspect for CAR hole LES. The remaining differences between the simulations and IR data are partly attributed to the steady coolant inlet profiles used for the counter and inline cross feeds.