Unsteady pressure fluctuations in transonic launcher configurations represent a major challenge, driven by shock oscillations, flow separation, and compressibility effects. To address these phenomena, this work presents and validates a methodology for reconstructing instantaneous pressure fields from planar (2D2C) and stereo particle image velocimetry (2D3C PIV) data using Taylor’s hypothesis. The approach is first assessed using a simulated PIV dataset of an axisymmetric backward-facing step, representing a launcher base flow configuration. In this case, the Taylor-based pressure reconstruction is validated against reference data and compared with results from a hypothetical time-resolved pressure reconstruction, demonstrating the accuracy of the method. The methodology is then applied to wind tunnel PIV data of a VEGA-like hammerhead launcher model at Mach 0.8 and zero angle of attack. Hammerhead configurations, characterized by payload fairings with a larger diameter than the main body, are particularly prone to separation and intense pressure fluctuations in the transonic ascent phase. The reconstructed pressure fields are analyzed together with unsteady transducer data, providing a general characterization of the flow features (oscillating shock, separation, and reattachment) and showing very good agreement for instantaneous, mean, and fluctuating components (ΔC_{p avg} ~ 0.01–0.02; ΔC_{p std} ~ 0.02). Finally, the analysis highlights the impact of neglecting out-of-plane velocity components, which introduces noticeable discrepancies in separated regions.

In this study, plume experiments were conducted to mimic the thermodynamic conditions on Saturn's moon, Enceladus. The icy moon's subsurface ocean and cracks in the surface have been simulated using a liquid water reservoir and a narrow channel, while the low-pressure environment at Enceladus’ surface was achieved with a vacuum chamber. We aimed to examine how channel temperature affected the plume's temperature, solid fraction and velocity, testing two models with differing wall temperatures: room temperature and near 0 °C. The colder setup better replicated Enceladus’ plume, producing a saturated flow in which nucleation of icy particles is possible. A conservative 1.5%–3% minimum solid fraction is estimated from measurements and modelling. Pitot-tube measurements indicated velocities around 400–500 m/s at the channel outlet. Flow temperature and velocity are closely correlated with wall temperature, indicating effective heat transfer. With a plume model based on the energy conservation law, we concluded that supersonic plume velocities observed on Enceladus cannot be achieved with straight channels, i.e. without requiring extreme expansion ratios. Additionally, the research provides evidence of the relationship between the crevasse's expansion ratio and the temperatures of flow and crevasse walls.

Oscillating loads from shock-wave/turbulent boundary-layer interactions (STBLIs) can cause structural components to resonate and potentially fail due to high-cycle fatigue.

An experimental investigation of separation bubble shaped control bumps for oblique shock wave–boundary-layer interactions was performed in two supersonic wind tunnel facilities at Mach 2.5 and 2, with incident shock deflection angles of 8° and 12°, respectively, and momentum thickness Reynolds numbers of approximately 1.5 × 10⁴. Shock control bumps were designed to replicate the time-averaged separation bubble shape, and were placed onto the floor in the separation location. This resulted in almost complete elimination of flow separation. There was also a marked improvement in the downstream boundary-layer state. A low-frequency bubble breathing oscillation was identified in the baseline interaction using high-speed shadowgraphy and particle image velocimetry measurements. This oscillation was strongly suppressed in the controlled interactions. Velocity fluctuations in the downstream boundary layer were also significantly reduced. We propose that the key mechanism by which flow separation is eliminated is by breaking down the overall pressure rise into smaller steps, each of which is below the separation threshold. A key feature is the bump crest expansion fan, located near to where the incident shock terminates, which negates the shock induced pressure jump. Thus, the precise bump geometry is critical for control efficacy and should be designed to manage these pressure rise steps as well as the expansion fan strength and location with respect to the incident shock wave. The length of the bump faces must also be sufficiently long for the boundary layer to recover between successive adverse pressure jumps.

Low emissions and fuel flexibility are two important criteria required for gas turbine combustors to facilitate the energy transition to low-carbon fuels for propulsion and power applications. A jet-stabilized combustor, having both these characteristics, was operated with CH ₄–H ₂ fuel mixtures with H ₂ varying from 0 to 100 % and with varying equivalence ratios (ϕ). Comprehensive measurements were carried out of the velocity field using Particle Image Velocimetry (PIV), temperature and gas composition by traversing probes in the chamber, and flame topology using chemiluminescence imaging. The flow field in this combustor consists of a jet that undergoes recirculation, generating Central and Peripheral Recirculation Zones (CRZ and PRZ). The recirculation ratio in the PRZ is found to be twice that of the CRZ. Increasing H ₂ % for the same ϕ leads to higher NO _x. Ultra-low ϕ flames could be stabilized only at H ₂≥50 %, which in turn leads to low NO _x due to low adiabatic flame temperatures. The combination of temperature, gas composition (CO/NO), and chemiluminescence images is used to identify the extent and location of the reaction zone. Distributed reaction zones, stabilizing at around 30 % of the length of the chamber, are achieved at lean conditions, whereas an increase in H ₂ % makes the reaction zone more compact and shifts upstream towards the burner head. Flame kernels are extracted from the instantaneous chemiluminescence images, and probability distribution functions for their aspect ratio and axial location are constructed. It is seen that reducing ϕ leads to low aspect ratio kernels that tend to occur further downstream, whereas increasing H ₂ % leads to higher aspect ratio kernels, stabilizing upstream. These flame kernel statistics are also used to identify ignition modes (autoignition/flame propagation) for varying fuel H ₂ % and inlet ϕ based on a hypothesis of flame stabilization mechanisms.

Reynolds numbers at cruise altitude can be such that a laminar boundary layer persists on the suction side of a transonic fan blade up to the shock-wave/boundary-layer interaction (SBLI). In a transitional SBLI which exhibits sufficiently large shock-induced separation, a shock oscillation mechanism characterized by growth and natural suppression of the upstream laminar section of the separation bubble occurs. To validate the shock oscillation mechanism observed in large eddy simulations (LES), the shock oscillation mechanism is studied experimentally using high-speed Schlieren and spark-light shadowgraphy. A characteristic length based on the distance of laminar separation shock travel is proposed. Strouhal numbers from LES and the experiment collapse at around 0.075. A strong dependency of the oscillation mechanism on free-stream turbulence and boundary-layer state is shown. Dominant oscillation frequencies are an order of magnitude lower for the turbulent interaction as opposed to the laminar case. For the laminar case, dynamic mode decomposition showed a strong relationship of the laminar separation shock with the separation bubble and reflected shock movement. The turbulent interaction shows a significantly lower reflected shock travel distance. The findings experimentally confirm that stabilization of the shock is achieved by tripping the boundary layer.

The design of transonic compressors increasingly focuses on higher blade loading, sparking interest in shock oscillation mechanisms in highly loaded transonic fans operating at cruise altitude. At such conditions, low chord Reynolds numbers (1.4 Mio.) may sustain a laminar boundary layer on the suction side of the blade up to the shock-wave/boundary-layer interaction (SBLI). The resulting interaction with large separation (pre-shock Mach number of 1.6) cause shock oscillations and structural excitation. In this study, we demonstrate that a canonical research configuration enables the experimental investigation of a specific shock oscillation mechanism relevant to transonic fans at altitude, providing a basis for validation. Using Large Eddy Simulations and experimental data, we show that the oscillation mechanism depends on the conditions at the SBLI rather than the geometry. The oscillation arises from the growth and self-suppression of the upstream laminar section of the separation bubble. Periodic collapse of this laminar section generates turbulence that entrains the separation bubble, influencing the dynamics of the reflected shock. The reflected shock movement resembles the cascade passage shock behavior, driven by blockage variations from the separation bubble. Additionally, we examine the numerical requirements to resolve this mechanism. These findings provide insights to advance compressor designs and hypersonic applications featuring similar mechanisms.

In this experimental study the effect of three-dimensional shock control bumps (SCB) on oblique shock wave/boundary layer interactions is investigated as a passive control method. It aims to develop an understanding of the effect of such devices on the interaction structure by means of studying the influence of the position of the bump with respect of the shock-impingement location, as well as the effect of the bump geometry (more in particular, the ramp section of the bump). The experiments were conducted in the ST-15 wind-tunnel at the Delft University of Technology for fully developed turbulent boundary layer conditions with Re_θ of 21.8·10³ and freestream Mach number of 2.0. The control effectiveness is assessed from the size of the separated flow region, as well as the downstream boundary layer velocity profile. For this, PIV is employed as the main diagnostic method to characterise the flow field. In addition to this, high-speed Schlieren and oil flow measurements were performed to asses the effect of the SCB on the overall interaction structure.

The jet-in-coflow is a two-stream configuration having engineering applications in combustors and gas turbine engine exhausts. In practical systems, the coflow generates a boundary layer of the outer wall of the jet pipe and may also have a certain level of turbulence. In the current work, the evolution of this flow configuration is studied using an air-air turbulent jet in a low turbulence coflow (turbulence intensity < 6%), and the 2D velocity field is measured by planar particle image velocimetry. Cases of varying coflow ratio (ratio of coflow velocity to jet velocity) of 0 (turbulent free jet), 0.09, 0.15, and 0.33 are generated by keeping a constant velocity jet (Re = 14000) and varying the coflow velocity. The trends of jet centerline properties such as velocity decay, jet spread, and jet momentum of jet-in-coflow cases, scaled to represent an equivalent free jet, show deviations from that of the turbulent free jet. The radial profile of mean velocity shows a region of velocity deficit, compared to a turbulent free jet, on the coflow side in the jet-in-coflow cases. In contrast, the turbulence intensity and Reynolds shear stress profiles show an enhanced peak near the interface for the jet-in-coflow cases. Further, conditional statistics were extracted by detecting the interface between the jet and the surroundings, wherein the same trends are observed. The low turbulence levels of the coflow have little effect on the jet/coflow interface, as seen by the conditional enstrophy diffusion and tortuosity compared to a turbulent free jet. The differences at the jet/coflow interface of a jet-in-coflow with respect to a turbulent free jet are attributed to the boundary layer initially developed by the turbulent coflow over the pipe generating the jet, and these are seen throughout the near-to-intermediate field (0≤x/D≤40).

For the largest wind turbines currently being designed, operation close to cut-out conditions can lead to the tip airfoil experiencing transonic flow conditions. To date, this phenomenon has been explored primarily through numerical simulations, but modelling uncertainties limit the reliability of these predictions. In response to this challenge, our study marks the first experimental investigation of a wind turbine airfoil under transonic conditions, for which we selected the FFA-W3-211 airfoil. Measurements were carried out in the high-subsonic range (Mach 0.5 and 0.6), utilizing schlieren visualization and particle image velocimetry (PIV) to characterize the airfoil across a range of angles of attack (AoAs) expected to be close to the boundary of transonic flow occurrence. Unsteady shock wave formation was observed for the higher Mach number, with the shock oscillation range increasing with steeper angles of attack. In addition, it was confirmed that the presence of a local supersonic flow region does not necessarily result in a shock wave. For cases with shock waves and trailing-edge separation, a buffet cycle was identified that is similar to, but distinct from, those seen in aviation applications. Our findings highlight the need for unsteady analyses even in steady operating conditions and call for dedicated research on wind turbine tip airfoils in transonic flow.

Hammerhead launcher configurations, characterized by a larger diameter in the payload fairing than the rest of the launch vehicle, face substantial challenges during transonic operations due to their susceptibility to flow separation. This experimental study investigates the influence of the nose and boat tail geometry on the flow around hammerhead configurations in the transonic regime (Ma = 0.7–0.8) and for various angles of attack (α = 0–4°). To gain a general understanding of the shockwave structures, flow separation and reattachment, oil flow and schlieren visualizations were employed. Schlieren visualizations were also utilized to characterize the level of unsteadiness in these regions. Additionally, particle image velocimetry was employed to quantify variations in the velocity field. The study’s findings reveal an optimization of flow performance in the presence of a bi-conic nose, attributed to the creation of two-shockwave structures with relatively low intensity. This is in contrast to the ogive and conic noses, which exhibit a single, more detrimental shockwave structure. The investigation into different boat tail angles indicates that adopting low-angle boat tails (5° and 15° compared to 34°) leads to a noticeable reduction in the separated area, albeit associated with an increase in the range of oscillation of the shockwave structures.

In this paper, a non-intrusive pressure measurement scheme based on particle image velocimetry (PIV) is presented for the complex supersonic flows with intense shock systems, by elaborately combining the MacCormack method, the streamline-based method, and the spatial integration in conservative form. According to the detailed analyses of flow structures, the pressure fields are well reconstructed by the proposed scheme for the two typical shock-wave/boundary-layer interactions containing regular and Mach reflections, which are induced by the relatively strong oblique shock waves generated by the wedges of 21° and 17° in the freestreams of Mach 2.5 and 2.0, respectively. Based on the theoretical solutions by oblique shock relationship, free interaction theory, and shock polar analysis, this pressure reconstruction scheme is completely validated to effectively suppress the propagation of PIV velocity error to the pressure field and the accumulation of reconstructed pressure error behind the strong shock wave. Compared with the literature presently, this work would be the most challenging application of PIV-based pressure measurement to such complex supersonic flows with intense shock reflections, large oscillations, wide speed ranges, and various compressible flow structures. These good results could confirm the feasibility and high accuracy of the proposed reconstruction scheme and may greatly promote its applications in academic research and engineering test for supersonic flows in the future.

This study investigates the spatial evolution of a zero pressure gradient turbulent boundary layer (TBL) imposed by a square-wave (SqW) of steady spanwise wall-forcing, which varies along the streamwise direction (x). The SqW wall-forcing is imposed experimentally via a series of streamwise periodic belts running in opposite spanwise directions, following the methodology of Knoop et al. [Exp. Fluids 65, 65 (2024)]0723-486410.1007/s00348-024-03799-9, with the streamwise extent increased to beyond ∼11 times the boundary layer thickness (δo) in the present study. This unique setup is leveraged to investigate the influence of viscous-scaled wavelength of SqW wall-forcing on the turbulent drag reduction efficacy for λx+=471 (suboptimal), 942 (near-optimal), and 1884 (postoptimal conditions), at fixed viscous-scaled wall-forcing amplitude, A+=12, and friction Reynolds number, Reτ=960. The TBL's response to this wall-forcing is elucidated by drawing inspiration from established knowledge on traditionally studied sinusoidal forcing, based on analysis of the streamwise-phase variation of the Stokes strain rate (SSR). The analysis reveals the SqW forcing to be characterized by a combination of two markedly different SSR regimes whose influence on the overlying turbulence is found to depend on the forcing waveform: subphase I of local and strong impulses of SSR downstream of the half- (λx/2) and full-phase (λx) locations, associated with a reversal in spanwise forcing directions, leading to significant turbulence attenuation, and subphase II of near-zero SSR over the remainder of forcing phase that enables turbulence recovery (when wall-forcing magnitudes and direction remain constant). Upon the initial imposition of the SqW forcing, the Reynolds stresses are strongly attenuated over the short streamwise extent of x/δ0<0.5 for all wavelengths, whereas the skin-friction transient is more gradual. Thereafter, once the forcing is ultimately established, the suboptimum and optimum wavelength regimes display no distinctive responses to the individual SSR subphases; rather, the drag-reduced TBL response is quasi-streamwise homogeneous. In contrast, an SSR-related phenomenology establishes itself clearly for the postoptimal case, in which a local attenuation of near-wall turbulence characterizes subphase I, while the turbulent energy recovers in subphase II owing to the extended region of near-zero SSR.

For the largest wind turbines currently designed, when operating at rated power and at high wind speeds, the tip airfoils can experience large negative angles of attack. For these conditions and in combination with turbulence, the airfoils are at risk of reaching locally supersonic flow, even at low free-stream Mach numbers. The possibility of shock wave formation and its consequences endangers the lifetime of these largest rotating machines ever built. So far only numerical analyses of this challenge have been attempted with significant modelling uncertainty. Here, for the first time, a wind turbine airfoil (the FFA-W3-211, used at the blade tip of the IEA 15MW reference wind turbine) is studied under transonic conditions using experimental techniques. Schlieren visualization and Particle Image Velocimetry were employed for free-stream Mach numbers of 0.5 and 0.6 and various angles of attack. It was shown that calculations based on isentropic flow theory and compressibility corrections were able to predict the situations where supersonic flow occurred. However, they could not predict the frequency of occurrence and whether shock waves were formed. In conclusion, an unsteady characterization of such airfoil behavior in transonic flow seems to be warranted.

Hammerhead launcher configurations, characterized by a larger diameter in the payload fairing than the rest of the launch vehicle, face substantial challenges during transonic operations due to their susceptibility to flow separation and intense pressure fluctuations. This experimental study investigates the influence of the nose and boat-tail geometry on the flow around hammerhead configurations in the transonic regime (Ma=0.7-0.8) and for various angles of attack (α=0-4°). To gain a general understanding of the main flow features, such as shockwave formation, separated flow in the boat tail region, and flow reattachment, oil flow and schlieren visualizations were employed. Schlieren visualizations were also utilized to characterize the level of unsteadiness in these regions. Additionally, particle image velocimetry was employed to quantify variations in the velocity field. The study's findings reveal an optimization of flow performance in the presence of a bi-conic nose, attributed to the creation of two-shockwave structures with relatively low intensity. This is in contrast to the ogive and conic noses, which exhibit a single, more detrimental shockwave structure (with the conic nose being the least favorable configuration). The investigation into different boat tail angles indicates that adopting low-angle boat tails (5° and 15° compared to 34°) leads to a noticeable reduction in the separated area, albeit associated with an increase in the range of oscillation of the shockwave structures.