An important challenge in the electrification of aircraft propulsion systems is the design of thermal management systems because of an increased heat load that needs to be dissipated. As an alternative to high-drag external heat exchangers, one can consider surface heat exchangers to dissipate the extra thermal energy. This reduces the size of the external heat exchangers and consequently reduces drag. However, a non-adiabatic wall can affect skin friction through a movement of boundary layer transition. Limited studies are available on the effect of non-adiabatic surfaces on laminar-to-turbulent transition in swept wing boundary layers dominated by crossflow instability (CFI). Therefore, the current work experimentally investigates the effect of surface heating on the stability and breakdown of the stationary crossflow instability. The experimental work is supported by Compressible Linear Stability Theory (CLST) computations.

Hot-Wire Anemometry (HWA) and Cold-Wire Anemometry (CWA) measurements of the boundary layer are performed on the STEP model, which features a 45 degree swept flat plate, for both adiabatic and heated surface conditions. Both the experimental and CLST results show a destabilisation of the primary instability linked to the increase in the growth rate of the stationary crossflow (CF) mode. The experimental results show that the type-I secondary CF instability exhibits a larger magnitude in the presence of wall heating and the mode emerges upstream compared to the adiabatic wall condition. The type-III mode displays a significant increase in magnitude in the presence of wall heating, thereby indicating a considerable destabilisation. The effect on laminar breakdown is identified by analysing velocity fluctuations in the 12-17 kHz frequency band in planes parallel to the surface for two different wall distances. A temperature ratio of 1.035 is found to advance breakdown by 5.7%.

Maintaining laminar flow on large swept surfaces of subsonic transport aircraft, i.e. the wings and the stabilisers, is currently posing a considerable challenge for aerodynamic design. Improving the efficiency of aircraft by delaying or removing the laminar-to-turbulent transition process over the wing and tail parts can substantially reduce contaminant emissions. The dominant flow instability causing laminar-turbulent transition of swept-wing flow is the so-called crossflow instability (CFI). Ongoing research at TU Delft has shown potential to delay transition by use of passive mechanisms. As such, a framework has been designed to numerically compute crossflow development and transition to turbulence on swept wings. Through the use of experimental data acquired in wind-tunnel measurements at TU Delft, the CFI development and transition process on swept wings has been modelled numerically by means of Direct Numerical Simulation (DNS). Based on a DNS laminar flow field generated from the pressure distribution along the model surface, a numerical primary CFI mode in good agreement with the experiment was obtained through Non-linear Parabolized Stability Equations (NPSE). Following this steady flow field analysis, the simulation was made unsteady by the implementation of numerical free-stream turbulence. This novel method resulted in unprecedented modelling of the receptivity mechanisms of transition in three-dimensional crossflow cases, overcoming ad-hoc treatments. Both experimental and numerical flow fields indicated a Type-I dominant secondary CFI (i.e. K-H type of response in the laterally inclined shear layer of the stationary crossflow vortex), which consequently carries the formation of near-wall hairpins and ultimately turbulence. Crossflow vortex frequency content also agrees well in the low-frequency band (450Hz ≤ 𝑓 ≤ 3000Hz), whilst the numerical high-frequency content (3500Hz ≤ 𝑓 ≤ 9000𝐻𝑧) does show a distinct delay in amplitude growth throughout the majority of the transition region. Contradicting the promising qualitative analysis of the free-stream turbulence methodology, this discrepancy in the frequency spectrum indicates the major shortcoming in the numerical setup, which was shown to be biased towards introducing more low-frequency disturbances at the inflow boundary.

Hybrid laminar flow control (HLFC) reduces skin friction drag, by combining boundary layer suction with pressure gradient tailoring to delay boundary layer transition. This research addresses two key developments required to perform the aerodynamic design of suction-type HLFC components. First, an existing numerical transition prediction tool was made compatible with boundary layer suction, by incorporating suction in its boundary layer solver and by ensuring that an appropriate grid of disturbance frequencies is evaluated with linear stability theory. Second, the scalability of the pressure losses associated to the perpendicular flow through large-scale and actual-scale perforated sheets was investigated experimentally. This development is required to predict the pressure losses of generic HLFC surfaces. It was found that the scalability of these pressure losses is limited, because the frictional losses inside holes decrease continually with decreasing hole diameter. In contrast, the inertial losses across holes reach a steady value with sufficiently many holes.