The impact of surface roughness geometry on aero-engine intakes at incidence

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Shock Wave-Boundary-Layer Interactions, or SBLI’s, are known to form on engine inlets within a complex transonic flow-field during typical take-off and climb configurations. On the engine inlet, there are a number of potential sources of surface roughness, such as novel de-icing and acoustic systems, or surface contamination. The impact on the flow-field structure, as a result of this roughness, may lead to detrimental side effects, such as losses in engine efficiency or intake flow stability. Previous research into two-dimensional roughness shapes demonstrated flow-field changes, for example a thicker downstream-boundary layer compared to a smooth surface. This paper compares the impact of a two-dimensional ridge roughness to a three-dimensional cubed roughness on the inlet flow-field. The effect of these rough surfaces is examined with schlieren photography and Laser Doppler Velocime-try (LDV) techniques. At an on-design condition, a rough surface promotes a smaller supersonic region, and a thicker boundary-layer downstream of the interaction compared to a smooth surface. At off-design upper surface mass flow rate conditions, modelling a higher mass flow engine demand, the supersonic region grows, leading to a shock location further downstream. Under these conditions, roughness also promotes a thicker downstream boundary-layer. However, comparing the two-dimensional with three-dimensional roughness at an approximate fan-face location, shows that three-dimensional roughness is more benign for all off-design cases. This suggests that the topology of the roughness is influencing the condition of the boundary-layer at this location

Normal-shock/boundary-layer interactions in transonic intakes at high incidence

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

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

Normal shock wave-turbulent boundary layer interactions in transonic intakes at incidence

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. 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.EPSR

Influence of near-leading edge curvature on the performance of aero-engine intake lips at high-incidence

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 shock- induced separation reported under on-design conditions, which correspond to typical take- off 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

Effect of lip shape on shock wave-boundary layer interactions in transonic intakes at incidence

The flow field around five transonic inlet lips at high incidence is investigated for a variety of flow conditions around a design point representative of high incidence manoeuvring. Changes to the operating point are simulated by varying the angle of incidence as well as changing the mass flow rate over the lip, intended to mimic the effect of an increase in engine flow. For these inflow conditions, the flow on the surface of the lip is characterised by a supersonic region, terminated by a near-normal shock wave. Of particular interest is the effect of lip geometry and operating point on the boundary layer at the equivalent fan location. The parametric investigation revealed a significant effect of lip shape on the position and severity of the shock wave-boundary layer interaction. From correlation studies based on the parametric investigation, it appears that the extent of shock-induced separation is the main factor affecting the boundary layer state downstream of the normal shock wave-boundary layer interaction. Somewhat surprisingly, this was found to be independent of shock strength but potentially related to the severity of the diffusion downstream of the shock. Alongside delaying flowreattachment, this diffusion is also likely to have a direct detrimental effect on the boundary layer development close to the engine fan

Onset of unsteadiness in aero-engine intakes at incidence

© 2017 by Andrea Coschignano. 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 flow field unsteadiness. Hflowever, a trend of increasing separation with greater mass flow was observed

Pedagogical principles and methods underpinning education of health and social care practitioners on experiences and needs of older LGBT+ people: findings from a systematic review

There is a growing awareness of the need for LGBT + competency training to ensure that the health and social care services offered to older LGBT + people is affirmative and gender sensitive. To conduct a synthesis of the literature that describes the pedagogical principles, curriculum content and methods (teaching and assessment) used to educate health and social care practitioners on the experiences and needs of older LGBT + people. Systematic thematic review of literature. MEDLINE, CINAHL, PsycINFO, EMBASE, Web of Science, Social Sciences Index, ERIC. In accordance with the Preferred Reporting Items for Systematic Reviews and Meta-Analyses (PRISMA) Statement, this review examined peer-reviewed papers published in English, prior to April 2018 that addressed pedagogical and curriculum issues on the inclusion of needs and experiences of older LGBT + people. The combined searches yielded 2214 papers of which 17 papers were eligible for inclusion, 10 discussion papers and 7 evaluation studies. Analysis identified the following themes: i) Acknowledging the wider historical context of older LGBT + people's lives; ii) Recognising that older LGBT + people are not a homogenous group; iii) Incorporating a multitude of theories and models from different perspectives; iv) Alerting practitioners to the health issues and disparities facing older LGBT + people; v) Including content that supports inclusive care for older LGBT + people; vi) Addressing barriers to older LGBT + people accessing health care; vii) Interactive activities are the preferred pedagogical strategy; viii) Involving older LGBT + people in curriculum development is a core principle; and ix) Mandatory education is not always the solution. As the field matures there is a need for more exploration of curriculum principles, assessment strategies and strategies to overcome barriers to the inclusion of issues experienced by older LGBT + people within curricula. [Abstract copyright: Copyright © 2019 Elsevier Ltd. All rights reserved.

The impact of roughness size on the shock wave-boundary-layer interaction on aero-engine intakes at incidence

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Shock Wave-Boundary-Layer Interactions, or SBLIs, are known to form in engine inlets within a complex transonic flow-field during typical take-off and climb conditions. In the engine inlet, there are a number of potential sources of surface roughness, such as novel de-icing and acoustic systems, or surface contamination. The impact on the flow-field structure, as a result of this roughness, may lead to detrimental side effects, such as losses in engine efficiency or intake flow stability. Previous research into the effect of large-scale two-dimensional, and three-dimensional roughness elements of similar size, demonstrated flow-field changes such as a thicker downstream boundary-layer, compared to a smooth surface. This paper compares the impact of two groups of rough surfaces on the inlet flow-field. The first is a smaller group which has a ratio of average roughness element height to maximum lip thickness less than 0.3. The second is a group with larger scale roughness element size, where this ratio is greater than 0.3. The effect of these rough surfaces is examined with Schlieren photography and Laser Doppler Velocimetry (LDV) techniques. At an on-design condition, modelling typical take-off conditions, the smaller roughnesses promote a larger supersonic region, and a thinner boundary-layer downstream of the interaction, compared to the larger sized roughnesses. At off-design conditions, which model an increase in the mass flow engine demand, the supersonic region grows, leading to a shock location further downstream for all surfaces. Although the larger roughnesses are seen to have a weaker shock strength than the smaller roughnesses, the larger surface roughness size also promotes a less full, but thicker downstream boundary-layer, with a greater separation, indicated by the larger λ-foot in the Schlieren images. This suggests that the size of the roughness is influencing the condition of the boundary-layer at the downstream location and the overall flow-structure

Experimental investigations of shock-wave/boundary-layer interactions in transonic aircraft engine intakes at high incidence

This paper presents results from a simplified experimental rig that aims to replicate the key physics of a transonic aircraft intake at high incidence for realistic altitudes. The equivalent flight conditions replicated are for free stream Mach numbers M∞ = 0: 25 - 0: 45 and incidence angles in the range α = 30 ± 5°. Measurement techniques from the simplified two-dimensional geometric setup include; schlieren imaging, surface oilflows, pressure sensitive paint, pressure and temperature measurements. CFD models that look to predict this behaviour have a limited accuracy largely due to a lack of experimental validation data - this lends itself as one of the key motivations for the work presented in this paper. Currently there is good qualitative agreement between the experimental results and the initial computation(s)