Introducing H ₂ as fuel in gas turbines is a promising step towards decarbonizing the energy sector. However, the future availability of H ₂ in large quantities remains uncertain. Consequently, designing fuel flexible (CH ₄/H ₂) combustion chambers for various fuel blends is necessary. The distinct combustion characteristics of H ₂, such as high flame speeds and high adiabatic flame temperatures, pose challenges when designing systems that can operate in a stable manner and with low emissions across a wide range of fuel mixtures. This paper investigates the fuel-flexibility of an atmospheric laboratory scale, partially premixed swirl stabilized combustor. By deploying a non-rotating axial air jet (AAI) in the center-line of the swirling flow, the flashback risk for high H ₂ content fuels is minimized. This study provides detailed insights into AAI's interaction with CH ₄/H ₂ fuel blends, analyzing the resulting flow field from Particle Image Velocimetry, emissions from exhaust gas analyser measurements, and flame structures from OH* chemiluminescence and OH Planar Laser Induced Fluorescence. The results show that AAI enables flame stabilization across the full range from 100% CH ₄ to 100% H ₂ in the same injector geometry. However, a high portion of the total airflow must be injected axially to stabilize H ₂ flames. Increasing the level of AAI increases NO emissions and alters flame stabilization mechanisms. This is likely due to a decrease in mixing quality, resulting in the fuel staying close to the periphery of the mixing tube. Switching the fuel from 100% CH ₄ to 100% H ₂ leads to an increase in NO emission, despite lower adiabatic flame temperatures for the perfectly premixed case. This indicates that the mixing process and flame location within the combustion chamber are essential in controlling NO emissions. Moreover, the flow field transforms significantly from a swirl-stabilized flow field featuring an inner recirculation zone to one resembling the one of a jet flame.

Large eddy simulations (LES) with flamelet and presumed filtered density function closure are used to simulate turbulent premixed and partially premixed hydrogen flames. Different approaches to model differential diffusion are investigated and compared. In particular, two existing models are extended to the LES framework to correct the resolved diffusive flux of the controlling variables due to differential diffusion. A lean premixed turbulent hydrogen flame in a slotted burner configuration is simulated first to compare the capability of the considered models in capturing local mixture fraction redistribution, super-adiabatic temperatures and thermo-diffusive instabilities. Results show that both models describe the formation of cellular burning structures. Next, a partially premixed lifted hydrogen flame in vitiated hot coflow is simulated to gain insight on the relevance of differential diffusion modelling at a higher turbulence level, a different combustion mode and in the presence of a complex stabilisation mechanism. Good predictions of the turbulent mixing and temperature fields are observed. Moreover, results show that the flame lift-off height has an appreciable sensitivity to the differential diffusion model. When differential diffusion is included only in the thermochemistry database, only mild effects on the predicted temperature fields, mixing and flame height are observed. On the contrary, a considerable shift of the flame base is observed when corrections are applied in the LES at the resolved level, depending on what controlling variables are considered. Further analyses reveal how the corrections of diffusive fluxes in the thermochemistry and at the LES level affect differently the flame burning mode, whose details are given throughout the paper.

Lean-premixed swirl-stabilized combustion is a successful strategy to reduce pollutant emissions. However, these combustion systems are especially prone to thermoacoustic instabilities. The precessing vortex core (PVC) plays a significant role in suppressing or exciting those instabilities. Therefore, it is necessary to predict the PVC dynamics in different operating conditions. The introduction of alternative aviation fuels like hydrogen in fuel-flexible gas turbines might require changes in the combustor geometry. However, the influence of particular geometric parameters on the PVC dynamics in less conventional combustion chamber configurations is not yet clear. To contribute to the knowledge of PVC dynamics in different combustor geometries, this paper presents an experimental study of the PVC dynamics in isothermal conditions in a counter-rotating dual swirler configuration in different confinement ratios. Additionally, a non-rotating axial air jet can be injected on the center line of the primary swirler as a provision for increased flashback resistance in the reacting case with H ₂. PVC frequencies and amplitudes are obtained by spectral proper orthogonal decomposition (SPOD) of time-resolved PIV measurements, and by time-resolved pressure measurements. The study shows that the frequency of the PVC scales with StPVC = 0.78, based on the diameter and the bulk velocity of the mixing tube. The PVC frequency is only determined by the conditions in the primary swirler and is fully independent of the amount of airflow going through the secondary counter-rotating swirler. Introducing an axial air jet on the center line decreases the PVC frequency significantly, which can be related to the change in effective swirl number. It is also shown that the smallest combustion chamber diameter results in the highest spectral energy for the PVC mode for all investigated points, hence the PVC motion is the strongest. Meanwhile, the biggest combustion chamber diameter shows the weakest pressure fluctuations. The results obtained in this study provide evidence that the periodic oscillations arising in the swirling flow field can be predicted and follow a Strouhal scaling independent of the geometry, even for more unconventional configurations.

Growing interest in unconventional aircraft designs coupled with miniaturization of electronics and advancements in manufacturing techniques have revived the interest in the use of Sub-scale Flight Testing (SFT) to study the flight behaviour of full-scale aircraft in the early stages of design process by means of free-flying sub-scale models. SFT is particularly useful in the study of unconventional aircraft configurations as their behaviour cannot be reliably predicted based on legacy aircraft designs. In this paper, we survey the evolution of various design approaches (from 1848 to 2021) used to ensure similitude between a sub-scale model and its full-scale counterpart, which is an essential requirement to effectively perform SFT. Next, we present an exhaustive list of existing sub-scale models used in SFT and analyse the key trends in their design approaches, test-objectives, and applications. From this review, we conclude that the state-of-the-art sub-scale model design methods available in literature have not been used extensively in practice. Furthermore, we argue that one sub-scale model is not sufficient to predict the complete flight behaviour of a full-scale aircraft, but a catalog of tailored sub-scale models is needed to predict full-scale behaviour. An introduction to the development of such a catalog is presented in this paper, but the development of a formal methodology remains an open challenge. Establishing an approach to develop and use a SFT catalog of models to predict full-scale aircraft behaviour will help engineers enhance confidence on their designs and make SFT a viable and attractive testing method in the early stages of design.

Novel aircraft designs with (distributed) propellers often feature a close coupling between propellers and air¬frame, leading to unsteady blade loading which impacts propeller efficiency, noise emissions, and vibrations. The goal of this paper is to study the impact of such installation effects on propeller design optimization. A combination of existing, rapid analysis models is used to compute the installed aerodynamic and acoustic propeller performance. These analysis models are coupled to a gradient-based optimization scheme for the design studies. Comparisons are made between optimizations performed with and without taking installation effects into account, analyzing a 5-deg angle-of-attack case and a boundary-layer-ingestion case. The results show that increasing the blade count improves aerodynamic and acoustic performance both for isolated and installed configurations. Furthermore, the acoustic performance is improved significantly by decreasing the blade tip Mach number, albeit with associated efficiency penalty. For the nonuniform inflow fields considered, accounting for installation effects inside the optimization procedure did not lead to large benefits in terms of aerodynamic and acoustic performance compared to the isolated design. For the 5-deg angle-of-attack case, the installed design was similar to the design for symmetric inflow, with negligible change in propeller efficiency and at most 0.5 dB noise reduction. For the boundary-layer-ingestion case, the installed design featured in¬creased solidity and decreased twist, resulting in 0.4% lower energy consumption compared to the design for symmetric inflow. Acoustic performance was not evaluated for this case. Future work will focus on the sensi¬tivity of the results to the blade tip Mach number, the impact of sweep and airfoil design on installed propeller performance, and acoustic optimization of installed configurations with more complex nonuniform inflow fields.

This paper presents a novel approach for correcting wind-tunnel wall interference in the nonlinear flow regime, that is, in the presence of phenomena such as flow separation and shocks. The methodology uses a gradient-based optimization to minimize the difference between experimental measurements and a Favre-averaged Navier–Stokes (FANS) simulation. The aim is to exploit the high-fidelity experimental data to correct turbulence-modeling errors in the FANS simulations, as well as to use the accurate angle of attack and Mach number from the FANS simulations to correct the in-tunnel flow conditions. The optimization is carried out directly in free air, thus avoiding the requirement to mesh the wind-tunnel walls and/or to model the ventilated-wall boundary condition. A byproduct of this method is the availability of flow information everywhere around the test object, which augments and complements the experimental data. The methodology is tested on two-dimensional and three-dimensional flow cases, demonstrating a significant improvement in the agreement between experimental and numerical data.

The unsteady aerodynamic characteristics of a 50° sweep delta wing performing pitching oscillations at angle of attack between 0° and 60° were tested in a water channel facility. Both force and velocity measurement results were analyzed and compared with the numerical simulation results for the pitching reduced frequency of 0.069 and 0.55. Both lift hysteresis and an unsteady phenomenon of leading-edge vortex (LEV) evolution were observed, which were significantly influenced by the pitch rate of the wing. As the wing pitch rate became sufficiently high, instead of being dissipated and convecting downstream, the LEV remained over the suction surface of the wing, which provide additional lift overshoot at a high angle of attack. In addition, as the wing pitched downstroke due to the negative induced camber effect, the lower surface of the wing turned from a pressure surface to a suction one and hence, there were significant lift losses.

Advances in aerodynamic and propulsive efficiency of future aircraft can be achieved by strategic installation of propellers near the airframe. This paper presents a robust and computationally efficient engineering method to estimate the load distribution of a propeller operating in arbitrary nonuniform flow that is induced by the airframe and by different flight conditions. The time-resolved loading distribution is computed by determining the local blade section advance ratio and using the sensitivity distribution along the blade, which is a property of the propeller in isolated conditions. The method is applied to four representative validation cases by comparing to full-blade computational fluid dynamics (CFD) simulations and experimental data. For the evaluated cases, it is shown that the changes in the propeller loads due to the nonuniform inflow are predicted with errors ranging from 0.5 up to 12% compared to the validation data. By extending the quasi-steady approach with a correction to account for unsteady effects, the time-resolved blade loading is also well approximated, without adding computational cost. The proposed method provided a time-resolved solution within several central processing unit seconds, which is seven orders of magnitude faster compared to full-blade CFD computations.

An anti-fairing is a concave deformation of the wall around a wing-body junction that can decrease the aerodynamic drag through the activation of a propulsive force generated by the interaction of the curved concave shape and the high-pressure region in proximity of the wing leading-edge. Although this mechanism is well understood, the dynamics of the interaction between the anti-fairing and the junction flow remain largely unexplored. This work brings together all the numerical and experimental studies of the anti-fairing to investigate its effect on turbulent quantities and the robustness of its design to changes to the incoming flow parameters, and to estimate the drag change with respect to a normal wing/flat-plate configuration. It is found that the interaction of the streamwise pressure gradient generated by the anti-fairing with the incoming boundary layer substantially reduces the shear responsible for viscous drag. Furthermore, no significant influence of the incoming boundary layer thickness on the anti-fairing performance is observed. However, a direct drag measurement with a force balance casts some doubts on the possibility to achieve large drag reductions.

The impingement of a propeller slipstream on a downstream surface causes unsteady loading, which may lead to vibrations responsible for structure-borne noise. A low-speed wind-tunnel experiment was performed to quantify the potential of a flow-permeable leading edge to alleviate the slipstream-induced unsteady loading. For this purpose, a tractor propeller was installed at the tip of a pylon featuring a replaceable leading-edge insert in the region of slipstream impingement. Tests were carried out with four flow-permeable inserts, with different hole diameters and cavity depths, and a baseline solid insert. Particle-image-velocimetry measurements showed that the flow through the permeable surface caused an increase in boundary-layer thickness on the pylon's suction side. This led to a local drag increase and reduced lift, especially for angles of attack above 6 deg. Furthermore, it amplified the viscous interaction with the propeller tip-vortex cores, reducing the velocity fluctuations near the pylon surface by up to 35%. Consequently, lower tonal noise emissions from the pylon were measured in the far field. This suggests that the desired reduction in surface pressure fluctuations was achieved by application of the flow-permeable leading edge.

Junction flows occur when a boundary layer develops on a wall and encounters an obstacle protruding from this surface. When the obstacle generates enough of an adverse pressure gradient to separate the flow, the aerodynamic drag is increased. In this paper, aerodynamic shape optimization (ASO) is employed to optimize a wing/body junction geometry at a chordReynolds number ofReC = 9.7 105,where thewing is theprotrusionandthebodyis representedby a flat plate. In contrast to conventional ASOs, thewing shape is kept fixed and only deformations of the body are allowed in order to study its influence on the junction drag. The obtained optimized design is characterized by a concave shape similar to a dent in the junction area and differentiates itself from the traditional convex fairings. For this reason, it is named the anti-fairing. Wind-tunnel experiments using stereoscopic particle image velocimetry in the wake of the junction area and a new set of Reynolds-averaged Navier-Stokes simulations with a finermesh than that used in the optimization are performed in order to validate the optimization, estimate the drag reduction with respect to the baseline geometry and two different leading-edge fairings, and investigate the mechanism by which drag is reduced. The anti-fairing is shown to systematically reduce drag and outperform leading-edge fairings thanks to the interaction between the wing and the front part of the concavity, generating a pressure force in the direction opposite to the drag force.

Wingtip-mounted propellers installed in a tractor configuration can decrease the wing induced drag by attenuating the wingtip vortex by the propeller slipstream. This paper presents an aerodynamic analysis of the propeller-wing interaction effects for the wingtip-mounted propeller configuration, including a comparison with a conventional configuration with the propeller mounted on the inboard part of the wing. Measurements were taken in a low-speed wind tunnel at Delft University of Technology, with two wing models and a low-speed propeller. Particle-image-velocimetry measurements downstream of a symmetric wing with integrated flap highlighted the swirl reductions characteristic of the wingtip-mounted propeller due to wingtip-vortex attenuation and swirl recovery. External-balance and surface-pressure measurements confirmed that this led to an induced-drag reduction with inboard-up propeller rotation. In a direct comparison with a conventional propeller-wing layout, the wingtip-mounted configuration showed a drag reduction of around 15% at a lift coefficient of 0.5 and a thrust coefficient of 0.12. This aerodynamic benefit increased upon increasing the wing lift coefficient and propeller thrust setting. An analysis of the wing performance showed that the aerodynamic benefit of the wingtip-mounted propeller was due to an increase of the wing's effective span-efficiency parameter.