Quantification of added mass effects using PIV data for a translating and rotating flat plate

Added mass characterises the additional force required to accelerate a body when immersed in an ideal fluid. It originates from an asymmetric change to the surrounding pressure field so the fluid velocity satisfies the no through flow condition. This is intrinsically linked with the production of boundary vorticity. A body in potential flow may be represented by an inviscid vortex sheet and added mass forces determined using impulse methods. However, most fluids are not inviscid. It has been theorised that viscosity causes the ‘added mass vorticity’ to form in an intensely concentrated boundary layer region, equivalent to the inviscid distribution. Experimentally this is difficult to confirm due to limited measurement resolution and the presence of additional boundary layer vorticity, some the result of induced velocities from free vorticity in the flow field. The aim of this paper is to propose a methodology to isolate the added mass vorticity experimentally with Particle Image Velocimetry, and confirm that it agrees with potential flow theory even in separated flows. Experiments on a flat plate wing undergoing linear and angular acceleration show close agreement between the theoretical and measured added mass vorticity distributions. This is demonstrated to be independent of changes to flow topology due to flow separation. Flow field impulse and net force are also consistent with theory. This paper provides missing experimental evidence coupling added mass and the production of boundary layer vorticity, as well as confirmation that inviscid unsteady flow theory describes the added mass effect correctly even in well-developed viscous flows.Cambridge Commonwealth European and International Trust, Schlumberger Limite

Vortex Generators to Control Boundary Layer Interactions

Devices for generating streamwise vorticity in a boundary includes various forms of vortex generators. One form of a split-ramp vortex generator includes a first ramp element and a second ramp element with front ends and back ends, ramp surfaces extending between the front ends and the back ends, and vertical surfaces extending between the front ends and the back ends adjacent the ramp surfaces. A flow channel is between the first ramp element and the second ramp element. The back ends of the ramp elements have a height greater than a height of the front ends, and the front ends of the ramp elements have a width greater than a width of the back ends

Vortical structures on three-dimensional shock control bumps

Three-dimensional shock control bumps have long been investigated for their promising wave drag reduction capability. However, a recently emerging application has been their deployment as “smart” vortex generators, which offset the parasitic drag of their vortices against their wave drag reduction. It is known that three-dimensional shock control bumps produce streamwise vortices under most operating conditions; however, there have been very few investigations that have aimed to specifically examine the relevant flow structures. In particular, the strength of the vortices produced as well as the factors influencing their production are not well known. This paper uses a joint experimental and computational approach to test three different shock control bump shapes, categorizing their flow structures. Four common key vortical structures are observed, predominantly shear flows, although all bumps also produce a streamwise vortex pair. Both cases with and without flow separation on the bump tails are scrutinized. Finally, correlations between the strength of the main wake vortices and pressure gradients at various locations on the bumps are assessed to investigate which parts of the flow control the vortex formation. Spanwise flows on the bump ramp are seen to be the most influential factor in vortex strength.The research leading to these results has received funding from the European Union’s Seventh Framework Programme (FP7/2007-2013) for the Clean Sky Joint Technology Initiative as part of the NextWing program under grant agreement no. 271843.This is the author accepted manuscript. The final version is available from the American Institute of Aeronautics and Astronautics via http://dx.doi.org/10.2514/1.J05466

Assessment of Computational Fluid Dynamics (CFD) Models for Shock Boundary-Layer Interaction

A workshop on the computational fluid dynamics (CFD) prediction of shock boundary-layer interactions (SBLIs) was held at the 48th AIAA Aerospace Sciences Meeting. As part of the workshop numerous CFD analysts submitted solutions to four experimentally measured SBLIs. This paper describes the assessment of the CFD predictions. The assessment includes an uncertainty analysis of the experimental data, the definition of an error metric and the application of that metric to the CFD solutions. The CFD solutions provided very similar levels of error and in general it was difficult to discern clear trends in the data. For the Reynolds Averaged Navier-Stokes methods the choice of turbulence model appeared to be the largest factor in solution accuracy. Large-eddy simulation methods produced error levels similar to RANS methods but provided superior predictions of normal stresses

Experimental investigation of transonic external fan cowl separation

When a civil aircraft engine is shut down during the cruise phase of flight and thus begins to windmill, a supersonic region forms on the external surface of the fan cowl. The terminating normal shock can separate the turbulent boundary layer developing on this external surface. A series of experiments are performed in a quasitwo-dimensional wind tunnel rig to investigate the influence of various parameters on this flow problem. As the engine mass-flow rate is reduced, an increase in normal shock strength results in the onset of flow separation which thickens the boundary layer developing on the external fan cowl surface by a factor of three. A reduction in incoming Mach number from the nominal value of 0.65 to 0.60 weakens the shock wave and thus delays flow separation. If the incoming boundary layer is laminar rather than turbulent, the normal shock Mach number is observed to increased by 10%. Despite the stronger shock, no significant flow separation can be detected even for the lowest engine mass-flow rates studied and the external nacelle surface boundary layer is measured to be thinner than for the turbulent case.This project has received funding from the Clean Sky 2 Joint Undertaking (JU) under grant agreement No 101007598. The JU receives support from the European Union’s Horizon 2020 research and innovation programme and the Clean Sky 2 JU members other than the Union

Influence of boundary-layer state on development downstream of normal shock interactions

Flat plate experiments investigating flow development downstream of normal shock wave–laminar boundary layer interactions are reported. Laminarity is desirable for low drag, but problems are expected at shock interactions due to separation, leading to un- steadiness and large boundary layer growth. A flat plate with elliptical leading edge profile is used to generate a laminar boundary layer, which encounters a Mach 1.2 normal shock at a Reynolds number of 1.6x10^6. Upstream of the interaction, the boundary layer is too thin for detailed measurement, but schlieren images are presented and laser Doppler ve- locimetry results reported downstream. For comparison, similar measurements are then made with trips located upstream. Far downstream, shape factor and skin friction are seen to be independent of incoming boundary layer state, contrary to expectations: although separation does occur in the laminar case, it is very thin and the anticipated impact on the boundary layer health is not seen. In fact, an equilibrium turbulent boundary layer is recovered far sooner downstream and is thinner than in the tripped cases, which exhibit attached turbulent interactions. In summary, normal shock wave–laminar boundary layer interactions did not exhibit the anticipated detrimental effects, and instead appear quite benign

Experimental investigation of external fan cowl separation for compact nacelles in windmilling scenarios

The slim fan cowl profiles used for ultra-high bypass ratio aircraft engines are designed considering off-design operating conditions, such as engine windmilling during take-off climb out or during cruise. The current paper describes wind tunnel experiments studying how incoming Mach number and engine mass-flow rate influence the aerodynamics governing external fan cowl flow separation in both these windmilling scenarios. A transonic region may form on the forebody surface if the engine becomes inoperative during take-off climb out, with peak Mach number up to 1.2. The subsequent adverse pressure gradient can separate the local boundary layer, resulting in flow separation which originates near the highlight and a more uniform fan cowl pressure distribution. Meanwhile, engine shut down during cruise results in a large supersonic region on the external fan cowl surface which terminates in a normal shock wave. When the Mach number of this shock exceeds about 1.35, a closed separation bubble develops, which causes up to a four-fold increase in the boundary-layer thickness downstream of the shock wave.European Union funding: 10100759

Design of a new test rig to investigate transonic external fan cowl separation

Ultra high-bypass ratio engines, which show considerable promise in reducing the environmental impact of commercial aviation, generally adopt slim fan cowl profiles. These geometries can be more sensitive to separation on the external surfaces in engine windmilling conditions during take-off climb out or during cruise. This paper describes the development of a two-dimensional wind tunnel rig which can accurately replicate the separation mechanisms experienced by real aero-engine nacelles. This design process highlights the importance of considering factors such as Reynolds-number effects, tunnel-wall effects, the two-dimensional nature of the rig, and the tunnel boundary layers.This project has received funding from the Clean Sky 2 Joint Undertaking (JU) under grant agreement No 101007598. The JU receives support from the European Union’s Horizon 2020 research and innovation programme and the Clean Sky 2 JU members other than the Union

Characteristics of shock-induced boundary layer separation on nacelles under windmilling diversion conditions

The boundary layer on the external cowl of an aero-engine nacelle under windmilling diversion conditions is subjected to a notable adverse pressure gradient due to the interaction with a near-normal shock wave. Within the context of Computational Fluid Dynamics (CFD) methods, the correct representation of the characteristics of the boundary layer is a major challenge to capture the onset of the separation. This is important for the aerodynamic design of the nacelle as it may assist in the characterization of candidate designs. This work uses experimental data obtained from a quasi-2D rig configuration to provide an assessment of the CFD methods typically used within an industrial context. A range of operating conditions is investigated to assess the sensitivity of the boundary layer to changes in inlet Mach number and mass flow through a notional windmilling engine. Fully turbulent and transitional boundary layer computations are used to determine the characteristics of the boundary layer and the interaction with the shock on the nacelle cowl. The correlation between the onset of shock induced boundary layer separation and pre-shock Mach number is assessed and the boundary layer integral characteristics ahead of the shock and the post-shock recovery evaluated and quantified. Overall, it was found that the CFD is able to discern the onset of boundary layer separation for a nacelle under windmilling conditions