Flow characterisation for a validation study in high-speed aerodynamics

Validation studies are becoming increasingly relevant when investigating complex flow problems in high-speed aerodynamics. These investigations require calibration of numerical models with accurate data from the physical wind tunnel being studied. This paper presents the characterisation process for a joint experimental-computational study to investigate the streamwise corners of a Mach 2.5 channel flow. As well as checks of flow quality typically performed for phenomenological investigations, additional quantitative tests are conducted. The extra care to obtain high quality data and eliminate any systematic errors reveal useful information about the wind tunnel flow. Further important physical insights are gained from designing and conducting wind tunnel tests in conjunction with numerical simulations. Crucially, the close experimental-computational collaboration enabled the identification of secondary flows in the sidewall boundary-layers; these strongly influence the flow in the corner regions, the target of the validation study

Experimental Stage Separation Tool Development in NASA Langley's Aerothermodynamics Laboratory

As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools

The influence of nozzle geometry on corner flows in supersonic wind tunnels

In supersonic flows, the separation in streamwise corners is a significant and widely encountered problem which can not be reliably predicted with the numerical methods commonly used in industry. The few previous studies on this topic have suggested conflicting corner flow topologies. Experiments of supersonic flow are typically conducted in wind tunnels with rectangular cross-sections, which use either a symmetric (full) or asymmetric (half-liner) nozzle configuration. However, the effect of the nozzle arrangement on the corner flow itself is not known. This paper examines the influence of nozzle geometry on the corner regions of a Mach 2.5 flow using a joint experimental-computational approach. The full setup and half-liner configuration are shown to produce different corner flow structures. The corner regions of the full setup and top corners of the half-liner exhibit thin sidewall boundary layers and a single primary vortex on the floor or ceiling. Meanwhile, the bottom corners of the half-liner configuration contain thick sidewall boundary layers and a counter-rotating vortex pair. Considerable vertical velocities are measured within the sidewall boundary layers. These are directed towards the tunnel centre-height for the full setup and downwards with the half-liner. The differences in sidewall cross flows between the two nozzle arrangements are likely due to distinct pressure distributions in the nozzle, where the secondary flows are set up. Measurements suggest that these nozzle-dependent transverse flows are responsible for the differences in corner flowfield between the two configurations. The proposed mechanism also explains observed differences in corner flow topology between previous studies in the literature; nozzle geometry therefore appears to be the dominant influence on corner flows in supersonic wind tunnels

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

This paper will investigate the validation of the NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as a baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet experiment conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a free-stream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the fanface diameter. The numerical simulations with and without tunnel walls are performed, quantifying tunnel wall effects on the BLI inlet flow. A comparison is made between the numerical simulations and the BLI inlet experiment for the baseline and VG vane cases at various inlet mass flow rates. A comparison is also made to a BLI inlet jet configuration for varying actuator mass flow rates at a fixed inlet mass flow rate. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCP avg, very well within the designed operating range of the BLI inlet. A comparison of the average total pressure recovery showed that the simulations were able to predict the trends but had a negative 0.01 offset when compared to the experimental levels. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion levels for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a misalignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery offset that was 0.01 lower than the experiment as was seen in the baseline. Comparisons of the flow features for the jet cases revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment