Surface imperfections on laminar aircraft surfaces, such as panel joints and discontinuities, can strongly modify boundary-layer stability. Sharp-edged rectangular bumps (SERBs) combine two widely studied surface features, the Forward-Facing Step (FFS) and a Backward-Facing Step (BFS), and provide a useful geometry for examining how step-induced mean-flow distortion and separation affect the development of Tollmien-Schlichting (TS) waves. However, previous studies on similar bump shapes have mostly considered a single incoming TS-wave amplitude, leaving the dependence of the SERB-induced boundary-layer response on incoming TS-wave amplitude insufficiently resolved.

To investigate how the boundary-layer response to a SERB depends on the amplitude of an incoming forced TS wave, this work combined numerical stability analysis with experiments performed on an unswept flat plate in the low-turbulence environment of the anechoic wind tunnel (A-tunnel) at TU Delft. Dielectric-Barrier Discharge (DBD) plasma actuators were used to force single-frequency TS-wave disturbances while allowing for easy amplitude variation. Surface microphones and Hot-Wire Anemometry (HWA) measurement systems were used to collect data to analyse the harmonic content of the disturbance, its wall-normal structure at the forcing frequency, and its downstream growth, comparing across the clean flat plate with and without the SERB installed over a range of forcing amplitudes. Independent monitoring of forcing consistency was shown to be important in long-duration plasma-forced experiments, where maintaining a consistent forcing amplitude is required. Simultaneous microphone measurements were necessary to distinguish genuine flow-induced amplification from transient variations in actuator output.

The results show that the influence of the SERB is strongly region-dependent: the forced TS-wave behaves like the clean case upstream of the bump and undergoes local amplification in the regions immediately upstream of the FFS and in the region just downstream of the FFS lip, a reduction in growth rate and stabilisation over the bump, and greater amplification downstream of the BFS. Greater forcing amplitudes are associated with a broader redistribution of spectral energy in the wall-normal direction and an earlier loss of a distinct shear-layer instability signature introduced in the recirculation region behind the BFS. The SERB, therefore, does not behave as a simple amplifier of TS waves, but instead produces a region-dependent response, with local amplification ahead of the FFS, reduced growth over the bump, and stronger re-amplification downstream of the BFS, where the response becomes increasingly sensitive to the amplitude of the incoming disturbance. ... Read more

This study aims to characterize the aerodynamics of rotor sails with the focus on the effects of end plate configuration. This is achieved through an experimental campaign on the wake development behind a scaled rotor sail (1:100) with varying end plate sizes in the SLT wind tunnel at TU Delft. Measurements were conducted for Re=15000, spin ratio up to 3, using a rotor of aspect ratio AR=5.2 and end plate diameters of De/D=1.0, 1.6 and 3.0. To omit heavy vibrations, a modal analysis was performed to determine the eigen frequency of the model.

Wake measurements were obtained using both a pressure rake and a 7-hole probe, making this the first experimental study to report results on all three velocity components in the wake of a scaled rotor sail. Three main wake characteristics were analyzed: 1) flow direction of the wake, 2) vortical topology of the wake flow and 3) shedding of the flow.

Results demonstrate that the end plate configuration has a large influence on the development and structure of tip vortices near the rotor’s free ends. A key limitation of this research was the increase in wake blockage for higher spin ratios affecting the pressure measurements. These findings provide new insights into the role of end plate geometry in rotor sail performance and wake dynamics. ... Read more

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%. ... Read more

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. KH-type 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 (450 Hz ≤ f ≤ 3000 Hz), whilst the numerical high-frequency content (3500 Hz ≤ f ≤ 9000 Hz) 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 a major shortcoming in the numerical setup, which was shown to be biased towards introducing more low-frequency disturbances at the inflow boundary. ... Read more

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. ... Read more