This thesis presents a comprehensive numerical investigation of wall interference effects in transonic wind tunnel testing of swept-wing configurations, by the example of the Cryogenic Ludwieg Tube Göttingen (KRG) at the German Aerospace Center (DLR). By employing Reynolds-Averaged Navier–Stokes (RANS) simulations using the DLR TAU code and the Spalart–Allmaras turbulence model, the study quantifies how test section walls affect the spanwise flow homogeneity around a swept wing at high Reynolds numbers. A detailed analysis was conducted to isolate the influence of key flow parameters, like angle of attack, sweep
angle, chord length, and Mach number, on wall-induced distortions.

The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.

Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition.

The Delft Laminar Hump (DeLaH), discovered by the Aerodynamics Department of the Faculty of Aerospace Engineering at the Delft University of Technology, is a symmetrical smooth hump that is placed at a set distance parallel to the leading edge of the wing that reduces skin friction drag as it attenuates the growth of the crossflow instabilities (CFI). With the hump, transition can be delayed up to 14%. The only requirements are that the hump needs an natural laminar flow (NLF) airfoil and a condition where CFI is the dominant transition mechanism to be effective. Generally, CFI dominates when the wing sweep is greater than 30 − 35 deg, making the vertical stabilizer the best candidate for implementing the hump.

This research implements the Delft Laminar Hump on the vertical stabilizer of subsonic transport aircraft by modeling the effect of the hump as a shift in transition location. By using a Quasi-3D aerodynamic analysis in combination with a transition location database, the effect of the hump on the lift and drag coefficient of the vertical stabilizer is analyzed. The transition location database is constructed by using the external velocity of airfoil sections of the vertical tailplane and the boundary layer solver and stability analysis developed by the Group of Flow Control and Stability within the Delft University of Technology. With this, the N-factor curves along the chord can be calculated. Knowing the respective N-factors of the clean and hump configuration, the associated transition locations can be determined, which are then used to calculate the lift and drag coefficients of the vertical tailplane. To evaluate the aerodynamic effect of the hump on the full aircraft directional and lateral stability a stability analysis based on a method by Fokker / Obert is performed and checked against the CS-25 for Large Aeroplanes regulations by European Aviation Safety Authority (EASA).

It was found that the Delft Laminar Hump (DeLaH) on the vertical stabilizer of a subsonic transport aircraft does not affect the vertical tailplane lift curve slope, thus not affecting the stability of the aircraft. In contrast, the vertical tailplane drag coefficient is reduced by the hump. Retrofitting the hump on Airbus A320 (conventional tail) and the Fokker F-28 Mk1000 (T-tail) results in a reduction of the vertical tailplane drag coefficient of 6.73% and 8.72%, respectively. Translating this vertical tail drag reduction to the full aircraft drag coefficient results in a reduction of 0.17% and 0.34%. To evaluate the effect of the hump on weight and fuel consumption, additional weight and mission analyses are performed. Evaluating the harmonic range, an fuel reduction due to the hump of 0.16% and 0.32% is established, for the Airbus A320 and Fokker F-28 Mk1000, respectively. The aircraft weight is reduced by the same percentage through the fuel reduction, as it is assumed that the added weight due to the hump itself is negligible.

Additionally, two sensitivity analyses were performed, namely, sweep angle variation and surface area scaling to analyze the effectiveness of the hump for different vertical tailplane geometries. The effect of sweep angle on hump effectiveness does not affect the vertical tailplane lift curve slope and thus also not the stability coefficients. For the full aircraft drag coefficient, the hump effectiveness has an exponential relation with sweep angle and is most effective at lower sweep angles. A maximum full aircraft drag reduction was found of 0.41% at 30 deg sweep with an equivalent fuel reduction of 0.39%. Overall it is concluded that lower sweep angles are beneficial as the hump is most effective and results in reduced vertical tailplane weight, less fuel weight as well as an increased stability margin. Analyzing the effect of surface area scaling, the hump has no effect on the vertical tailplane lift curve slope regardless of surface area, again retaining the aircraft’s stability. For the full aircraft drag coefficient, the hump effectiveness increases linearly for increasing surface area up to 0.37% at a surface scaling factor of 1.2 times the original vertical tailplane surface area. In terms of fuel reduction, a maximum value of 0.35% was found. There will be an optimal vertical tailplane surface area, since the hump effectiveness increases for increasing surface area, whilst for the full aircraft drag, vertical tailplane weight, and fuel weight a smaller surface area is preferred. The stability margin becomes the limiting factor as a minimum surface area is required for sufficient stability. Comparing the baseline aircraft, the hump is more effective for the Fokker F-28 Mk1000 over the entire range of scaling factors and sweep angles. This leads to the suspicion that taper- and aspect ratio, and cruise speed play an important role in the effectiveness of the hump, but more research is required.

This research shows that the Delft Laminar Hump has a significant drag-reducing effect on the overall aircraft, whilst not affecting the aircraft’s stability. Even though the fuel savings for an individual aircraft are not very large, on a fleet level this would be significant. The hump can be retrofitted on existing aircraft by gluing it on the outer skin, making it a relatively simple and cheap way to improve efficiency for aircraft manufacturers. Nevertheless, more research is required before implementing the hump on commercial aircraft as it is still unknown whether the hump also works on the suction side of the wing as well as whether the hump causes a shock at cruise Mach. Also, interaction effects with the horizontal stabilizer and fuselage need to be taken into account, to fully quantify the effectiveness of the hump. Dedicated wind tunnel experiments, flight tests, and/or CFD simulations are necessary.

Hot-wire anemometry (HWA) remains an essential diagnostic tool for high–frequency velocity measurements in aerodynamic research, yet its practical accuracy strongly depends on robust temperature correction and reliable multi–component probe calibration. Despite the maturity of the technique, the literature provides no consensus on how best to correct for fluid–temperature variations or how to calibrate X-wire probes with consistent accuracy across wide velocity and angular ranges. The situation is further complicated at the TU Delft Low-Speed Wind Tunnel Laboratory (LSL), where the absence of an in-situ multi-wire calibration system has made efficient and repeatable probe preparation difficult. This thesis addresses these issues by developing a new automated calibration framework for multi-wire CTA probes and performing a comprehensive evaluation of existing temperature-correction and X-wire calibration schemes.
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface

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%.