During high-incidence manoeuvres, shock-wave boundary layer interactions can develop over transonic inlet lower lips, significantly impacting aerodynamic performance. Here, a novel experimental rig is used to investigate the nature and severity of these interactions for a typical high incidence scenario. Furthermore, we explore the sensitivity to changes in angle of incidence and mass flow rate, as potentially experienced across off-design operations. The reference flow-field, informed by typical climb conditions, is defined by an incidence of 23 • and a free stream Mach number M=0.435. The lower lip flow is characterised by a rapid acceleration around the leading-edge and a M≈1.4 shock ahead of the intake diffuser. Overall, this flow-field is found to be relatively benign, with minimal shock-induced separation. Downstream of the interaction, the boundary layer recovers a healthy profile ahead of the nominal fan location. Increasing incidence by 2 • , the separation becomes noticeably larger and unsteadiness develops. Detrimental effects are exacerbated at an even higher incidence of 26 •. Increasing the mass flow rate over the lip by up to 15% of the initial value has minor effects on performance and is not found to inhibit the boundary layer profile recovery. Nomenclature α Angle of incidence δ Boundary layer thickness δ * Boundary layer displacement thickness θ Boundary layer momentum thickness c Intake chord length H Shape factor L * Interaction length

This paper describes the investigation into the flow over the lip of subsonic engine intakes at incidence, focusing on the shock wave-boundary layer interaction occurring over the inner lip. A baseline geometry is considered along with two variations, characterised by a sharper and a blunter intake highlight (i.e.: nacelle leading edge) respectively. Results to date reveal a relatively benign interaction for the baseline model, with small or no shockinduced separation reported under on-design conditions, which correspond to typical takeoff or climb circumstances. The alternative geometries reveal a considerable influence of near-highlight curvature on the flow development. In particular, a blunter nose leads to the formation of a larger supersonic region, terminated by a consequently stronger shock, which shows a greater degree of shock-induced separation and increased total pressure losses and unsteadiness. The sharp nose, on the other hand, resulted in the compression occurring via three separate shock-waves, all of which weak. Overall, none of the three intake geometries showed inherently unsteady behaviour. However, this is expected to occur as the engine flow demand increases. Further testing is in progress to assess off-design performance and to produce a complete operational envelope for intakes at incidence.

The interaction between a normal shock wave and a boundary layer is investigated over a curved surface for a Reynolds number range, based on boundary-layer growing length x, of 0.44 × 10 6 ≤ Re x ≤ 1.09 × 10 6. The upstream boundary layer develops around the leading edge of the model before encountering a M ∼1.4 normal shock. This is followed by adverse pressure gradients. The shock position and strength are kept constant as Re is progressively varied. Infra-red thermography is used to determine the nature of the upstream boundary layer. Across the Re range, this is observed to vary from fully laminar to fully turbulent across the entire span. Regardless of the boundary-layer state, the interaction remains benign in nature, without large scale shock-induced separation or unsteadiness. Schlieren images show a pronounced oblique wave developing upstream of the main shock for the laminar cases, this is believed to correspond to the separation and subsequent transition of the laminar shear layer. Downstream of the shock, in the presence of adverse pressure gradients, the boundarylayer growth rate is inversely proportional to Re. Nonetheless, across the entire range of inflow conditions the boundary layer recovers quickly to a healthy turbulent boundary layer. This suggests the upstream boundary-layer state, and its transition mechanism, to have little effect on the outcome of its interaction with a normal shock wave.

The flow field around a transonic engine inlet lip at high incidence is investigated for a variety of flow conditions around the design point. Generally, the flow on the upper surface of the lip is characterised by a supersonic region, terminated by a near-normal shock wave. At the nominal design point, the shock is not strong enough to cause significant flow separation, resulting only in marginal losses in pressure recovery. Off-design conditions were explored by altering the angle of attack as well as changing the mass flow rate over the upper lip, intended to mimic the effect of an increase in engine flow. The results suggest that angle of attack has the greatest effect on the flow field. In particular, even a relatively small increase of 2 • can lead to large and highly unsteady flow separation with an associated shock oscillation. Both qualitative and quantitative measurements suggest a noticeably reduced aerodynamic performance resulting from higher incidence operation. In contrast, an increase of up to 5.2% in mass flow over the upper part of the intake lip did not result in large separated regions or flow-field unsteadiness.

The flow field around a transonic engine inlet lip at high incidence is investigated for a variety of flow conditions around the design point. Generally, the flow on the upper surface of the lip is characterised by a supersonic region, terminated by a near-normal shock wave. At the nominal design point, the shock is not strong enough to cause significant flow separation for each of the shapes investigated. Off-design conditions were explored by altering the angle of attack as well as changing the mass flow rate over the upper lip, intended to mimic a greater mass flow demand by a turbofan engine. The results suggest that angle of attack is the dominant parameter, where an even relatively small increase of 2 • can lead to large and highly unsteady flow separation with an associated shock oscillation. This is a consequence of the significantly stronger shock compared to the on design case. Both qualitative and quantitative measurements suggest a noticeably reduced aerodynamic performance resulting from higher incidence operation. In contrast, an increase of up to 5.2% in mass flow did not result in large separated regions or flowfield unsteadiness. However, a trend of increasing separation with greater mass flow was observed.