The 'hammerhead fairing' (HHF) launcher configuration uses a payload fairing that has a larger diameter than the launch vehicle. This allows for larger payloads to be carried using the same launcher components. However, the HHF leads to high aerodynamic forcing in the transonic regime called buffeting. If unaccounted for, buffeting can lead to launcher failure. The flow phenomena associated with buffeting were investigated on the Vega-E launcher, which is currently being developed by the ESA, using Pressure Sensitive Paint (PSP) in the TST-27 tunnel.

PSPs are a relatively new flow measurement technique that allows for quantitative measurements of surface pressure levels in a non-intrusive manner. High spatial resolution, which is equivalent to using millions of pressure taps, can be achieved using PSP. This allows for CFD-like surface pressure resolution to be obtained from wind tunnel experiments. However, PSPs had previously never been successfully implemented in Delft. Therefore, a test set up was devised for experiments in the TST-27. The final implementation was found to have a pressure uncertainty of at most 2.5 kPa, with a spatial resolution of 0.088mm.

First, validation measurements were done on the Coe and Nute Model 11. The Model 11 is a common HHF test case in literature that is used to study the buffeting phenomenon. Results obtained in the TST-27 were compared to historical data and showed good matching when accounting for blockage. After the validation, measurements were taken on the Vega-E launcher at different Mach numbers and angles of attack to quantify the effects of each variable on the pressure field. Lastly, the impact of protuberances on the model was studied. The protuberances that were used mimic those that will be present in the full-scale launcher.

The aerospace industry’s push for faster, lighter, and more durable designs highlights the critical challenge of panel flutter in supersonic flight, a poorly understood phenomenon intensified by shock-wave/boundary-layer interactions (SWBLI). This study experimentally investigates the effects of static and dynamic panel deformations on shock-induced panel flutter at Mach 2.0, using open, closed, and ventilated cavity conditions. Results show that cavity closure significantly influences panel response, with the open cavity causing greater oscillations due to strong pressure waves entering it, causing resonance behaviour; closing the cavity seems to filter out this resonance behaviour. Additionally, temperature changes and wind tunnel confinement were identified as factors affecting panel and flow behaviour. The results underscore the importance of understanding experimental conditions to accurately interpret fluid-structure interactions in supersonic conditions.

As the space industry continues to grow rapidly, the development of reusable launch vehicles has become crucial in achieving cost-effective and sustainable access to space. Payload capacity optimisation has been at the forefront of this effort, leading to a renewed interest in hammerhead or bulbous payload fairing (PLF) configurations. These designs feature a larger diameter in the PLF than in the rest of the launch vehicle, enabling the same structure to be used for large payloads. However, the transonic regime poses unique challenges for these PLFs due to their susceptibility to flow separation and strong pressure fluctuations.

This master thesis investigates the influence of nose and boat-tail geometries on flow phenomena, particularly shock wave generation, around hammerhead configurations in transonic conditions. The research was conducted through three distinct experimental campaigns in the transonic wind tunnel TST-27 at the high-speed laboratory of TU Delft, employing schlieren, oil flow, and Particle Image Velocimetry (PIV) techniques.

The study revealed notable insights into the aerodynamics of the hammerhead PLF configurations. Boat-tails set at five and 15-degree angles were observed to broaden the range of shock wave oscillations, introducing an additional shock wave that could occasionally merge with existing ones. Moreover, the conic nose design induced higher shock wave oscillations, while the bi-conic nose introduced an extra shock wave compared to the conic and ogive noses. The study also found that altering the nose shape while keeping the boat-tail constant, or viceversa, resulted in similar effects on flow dynamics.

These findings underscore the critical role that nose and boat-tail geometries play in shaping the aerodynamic behaviour of hammerhead PLF configurations. The results have implications for such configurations’ design and stability considerations in transonic conditions. It was observed that raising the Mach number heightened shock wave oscillations and flow detachment. The angle of attack disrupted model symmetry, primarily impacting reattachment patterns. Furthermore, the conic nose exhibited greater unsteadiness due to oscillations in the shock waves compared to the other geometries. The study also provided detailed insights into shock wave spectral characteristics and identified potential influences of pressure wave oscillations on shock wave behaviour. The conclusions lead to recommend designs with bi-conic nose, and avoid boat-tail angles around 15 degrees.

Supersonic panel flutter is a self-excited vibration, where energy is transferred into thin flexible panels, such as skin panels in supersonic vehicles, from the (supersonic) flow around them, which can lead to their rapid fatigue failure. Relevant examples of this include the failure of skin panels in the X-15 vehicle, or violent oscillations in non-adaptable nozzles during start-up, such as the space shuttle main booster. With the contemporary trends of decreasing weight and increasing operational speeds in the design of next generation aircraft and launch vehicles, such structures will become more common, and flutter needs to be inherently accounted for in the design.

In recent years, advancements in numerical techniques, such as Computational Fluid Dynamics (CFD) and the Finite Element Method (FEM), have led to numerous studies that address the effect of non-linearities, such as flutter behaviour in the transonic and hypersonic regime. In spite of that, experimental data needed to validate these studies is limited, based on coarse pointwise measurement techniques, and only focused on the structural response of the panel rather than on the entire Fluid Structure Interaction (FSI). The use of simultaneous, full-field and non-intrusive measurements can be used to address these issues. To this extent, panel designs were developed and an experimental campaign was proposed where Digital Image Correlation (DIC) and Schlieren are simultaneously used by means of high speed cameras to study classical panel flutter at Mach 2 inside the ST-15 wind tunnel at the Delft University of Technology. Additionally, a Laser Doppler Vibrometer (LDV) was used to validate the DIC measurements. The experiment allowed for a comparison with well understood theory, such that the setup can be proven to be robust, and can therefore later be used to assess the effect of panel flutter non-linearities or other supersonic FSIs.

Synchronisation between DIC and schlieren was demonstrated through a matching frequency spectrum, and maximum correlation inside corresponding pixel windows between panel displacement and schlieren grey-scale fluctuations with a zero time-lag. One finding was that during the panel down-stroke phase, a travelling wave character was observed which has previously been found only under transonic conditions. Furthermore, at moderate dynamic flutter pressures, a typical second bending panel flutter mode shape was observed, which grew into a single bending mode again with increasing higher dynamic flutter pressures. From the spectral analysis it was found that all configurations fluttered at the same frequency of 770 [Hz], independent of dynamic flutter pressure or edge conditions. This was higher than expected from analysis. These combined observations led to the conclusion that, in the current facility, a frequency lock-in occurs due to merging of the flutter phenomenon with a wind tunnel resonance vibration.

These observations provide an insight into combining these experimental techniques to study supersonic FSI behaviour, and particularly into certain unexpected behaviour encountered at the TU Delft ST-15 wind tunnel. This can be used to further tune future experiments and guide researchers to obtain a better understanding in the non-linearities found in panel flutter.

The vibration of panel structural elements immersed in a supersonic flow is a poorly understood fluid-structure interaction (FSI) that can affect the performance and structural integrity of supersonic aircraft and spacecraft systems. These adverse effects are further amplified when a shock-wave/boundary-layer interaction (SWBLI) is formed over the panel. A better understanding of this phenomenon—referred to as shock-induced panel flutter—is therefore crucial for the design of future high-speed vehicles.
An experimental method is developed to study shock-induced panel flutter at Mach 2 using planar particle image velocimetry (PIV) and stereographic digital image correlation (DIC) to obtain simultaneous, full-field structural displacement and flow velocity measurements. High-speed cameras are employed to conduct spectral analyses of the panel’s motion and the low-frequency dynamics of the SWBLI during the interaction. To avoid optical interference between the PIV and DIC systems, an optical isolation system is devised using fluorescent paint, dedicated light sources, and color lens filters.
The devised experimental setup is used to study the effects of an impinging oblique shock on the dynamics of a flexible panel during flutter, and the effects on the mean flow separation and interaction length of a SWBLI when a rigid wall is substituted by a compliant plate. The coupling of the panel and SWBLI is also studied, identifying the regions in the flow of maximum correlation between the panel’s motion and the flow velocity fluctuations. The obtained results suggest that the inviscid flow region upstream of the SWBLI may play a significant role in the fluid-structure interaction. In addition, a parametric study is conducted to determine the effects of panel aspect ratio and edge boundary conditions on the three-dimensionality of the panel dynamics. The reported findings serve as a reference for future researchers when designing both experimental and numerical simulations of shock induced panel flutter, particularly if two-dimensional flutter is to be recreated.

Transonic buffet is an aerodynamic instability that appears on aerofoils and wings at certain combinations of Mach number and angle of attack, and which limits the aircraft flight envelope. This thesis project presents an experimental
investigation of buffet employing stereoscopic PIV and Background Oriented Schlieren. By using four aerodynamic models, a 2D aerofoil and three 3D wings with sweep angles of 0^o, 15^o and 30^o, the differences between the 2D buffet typical of aerofoils and the 3D buffet typical of swept wings is studied, scrutinising the influence of the sweep angle. The results confirm that for the same aerodynamic conditions the 2D buffet experiences larger shock oscillations than the 3D case. A frequency analysis reveals a clear peak at 160 Hz for the 2D case, which is also visible in the 3D case for low sweep angles. As the sweep angle increases, this peak is replaced by a bump at higher frequencies.