The mixing of fuel and air is a key factor in determining NOx emissions during combustion. Lean-premixed burning strategies allow to control the flame temperature and therefore NOx emissions. However, for highly reactive fuels like hydrogen, the high flame speed makes full premixing dangerous due to the increased risk of flashback. In these cases, current combustor geometries are often operated in partially premixed modes with the fuel injected as close as possible to the combustion chamber. This highlights the need for effective mixing strategies to achieve a high degree of mixing over a short distance. This is even more critical in fuel-flexible combustion systems (e.g., combustors capable of burning both CH4 and H2), as the mixing process is heavily influenced by the varying properties of the fuel mixture. In such cases, a comprehensive understanding of the mixing process is required to minimize NOx emissions under all fuel blends conditions. This paper investigates the mixing of fuel jets into a swirling air cross-flow of a partially-premixed, swirl stabilized combustor using a combined experimental and numerical approach. The injector features an axial swirler and a mixing tube where the air and the fuel jets mix before entering the combustion chamber. The experiments are performed in cold flow conditions. A variable mixture of helium–air is used to represent different blends of CH4-H2 fuel, and the mixing process is visualized by seeding the fuel stream with DEHS droplets. Large-Eddy Simulations (LES) confirm the suitability of helium as a surrogate for H2 by demonstrating similar macro-mixing behavior for the two gases. This study examines the impact of varying fuel composition and momentum flux ratio (Jswirl) between the fuel jet and the swirling cross-flow on mixing performance. The results indicate that fuel with lower density achieve better mixing with the air at the mixing tube outlet. A numerical analysis of the radial transport terms reveals that higher H2 content in the fuel makes it less subject to outward convection which causes stratification close to the mixing tube outlet. Furthermore, the contribution of the molecular diffusion term increases with higher levels of H2, resulting in improved mixing. When increasing Jswirl (up to Jswirl = 10) increases the penetration of the fuel jet into the swirling flow. Above a critical value of Jswirl, the mixture homogeneity at the mixing tube outlet becomes insensitive to Jswirl for the investigated geometry. Overall, the fuel composition was found to have a greater influence on the level of mixing close to the mixing tube outlet than variations in Jswirl.

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.

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.

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.

The AeroCity is a new form of transportation concept that has been developed to provide high-speed ground transportation at a much lower cost than the existing high-speed railway. Utilizing the Wing-in-Ground (WIG) effect, the AeroCity vehicle does not require complex infrastructures like other contemporary concepts, such as the Hyperloop or Maglev trains. In the current work, the aerodynamic characteristics of the AeroCity vehicle are examined through a Computational Fluid Dynamics (CFD) analysis. The results from the CFD analysis qualitatively match with the findings of wind tunnel experiments. Surface streamlines and boundary layer measurements correspond well with the numerical data. However, the force measurements show a discrepancy. It is found that the separation bubble over the side plates is not captured by the CFD, and this is responsible for an under-prediction of the drag at higher free-stream velocities. The Transition SST model improved the matching between the experiments and numerical simulations. The influence of the moving ground is numerically investigated, and the effect of non-moving ground on the vehicle aerodynamics was found not to be significant. Finally, the inclusion of the track wall is examined. It is found that the merging of the wingtip vortices is responsible for a significant drag increase and, therefore, an alternative track geometry should be investigated.

This experimental study focuses on the aerodynamic interaction between an over-the-wing (OTW) propeller and a wing boundary layer. An OTW propeller is positioned above the hinge line of a wing with a trailing-edge flap. Measurements are carried out with and without axial pressure gradients by deflecting the flap and by extending the flat upper surface of the wing in the streamwise direction, respectively. Surface-pressure taps, microphones, and particle image velocimetry are combined to quantify both the time-averaged and unsteady interaction effects. Results show that the propeller generates an adverse pressure gradient on the wing surface that scales linearly with thrust and decreases with increasing blade-tip clearance. The pressure gradient is partially caused by slipstream contraction, which decelerates the flow near the wall. Additionally, the surface-pressure fluctuations generated beneath the propeller blades and slipstream are stronger than the time-averaged pressure increase due to flow deceleration. Consequently, the propeller triggers flow separation over the hinge line when the flap is deflected. A parametric study of different propeller locations indicates that increasing the tip clearance is not an effective way to mitigate flow separation. However, displacing the propeller half a radius upstream of the hinge line creates a Coandă effect, which allows the flow to remain attached.

The increasing demand for short-range passenger air transport and the strong push for aircraft with electric propulsion has renewed research interest in propellers. Despite the unmatched aerodynamic efficiency of propellers, their relatively high noise emissions limit widespread application on aircraft. Previous research has not systematically addressed the tradeoff between aerodynamic and aeroacoustic performance. This paper presents the results of an optimization study aimed at minimizing propeller noise without compromising aerodynamic efficiency. In the optimization, a blade-element-momentum-theory (BEMT) model is utilized which accounts for the effects of blade sweep on the blade loading. This BEMT model is coupled to a frequency-domain code for tonal noise prediction. A novel scaling approach is presented to directly relate the propeller noise emissions to the propeller thrust. Dedicated wind-tunnel experiments were performed to validate the analysis models. Good agreement between numerical and experimental results is obtained at low to moderate blade loading conditions. The optimization study shows that the blade sweep is an important design parameter to simultaneously maximize aerodynamic and acoustic performance. Compared to a modern baseline design, a noise reduction of 2.9 dB is achieved without reduction in propeller efficiency.

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.

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.

This paper presents two novel data assimilation (DA) techniques for reconstructing steady turbulent flows at high Reynolds numbers by introducing perturbations to the Reynolds stress tensor computed by the turbulence model of a Favre-averaged Navier–Stokes (FANS) code. These techniques minimize the least-squares difference between an experimentally measured mean flow quantity and the corresponding quantity as computed by the FANS code. The two DA methods differ from each other in the choice of the control parameters: one perturbs the eigenvalues and eigenvectors of a baseline Reynolds stress, whereas the other perturbs the components of a baseline realizable Reynolds stress such that the perturbed result is still realizable. For the optimization procedure, a gradient-based algorithm is used in combination with a discrete adjoint methodology. The DA methods are applied to high-Reynolds-number problems, and their results compared with a reference technique. The results show that the approaches developed in this work are more effective at reconstructing the turbulent flowfield than standard techniques, but are more computationally expensive due to the high dimensionality of the optimization problem. Furthermore, it appears that only small perturbations to the control parameters are necessary to obtain significant improvements over the baseline results.

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.

A numerical investigation of a propeller with swirl recovery vanes, for which experimental data exist, is performed. A second swirl recovery vane geometry, with shorter vanes to avoid the impingement of the propeller tip vortices, is also investigated. For the baseline swirl recovery vanes, the efficiency of the propulsive system increases by 2.4% with respect to the isolated propeller. This is obtained by converting angular momentum in axial momentum. A reduction of the swirl angle in the near wake by 48% is found. Most of the thrust is generated at the root of the vanes. Leading-edge impingement noise is the dominant source. The vanes cause noise to increase by 20 dB with respect to the isolated propeller in the axial direction, where noise from the propeller vanishes. In the axial direction, sound pressure level spectra show tonal peaks at harmonics of the second blade passing frequency, while in the other directions, peaks are present at harmonics of the first blade passing frequency. However, the overall isolated propeller noise is 23 dB higher than the noise generated by the swirl recovery vanes. Shortening the vane length causes a 13% reduction of the thrust generated by the vanes with respect the baseline case but no variation of the far-field noise.

Among all the interference effects contaminating wind-tunnel experiments, those due to the presence of the walls are particularly challenging to account for. Their influence becomes hard to quantify in the non-linear flow regime, for which classical linear correction methodologies are inadequate. This work proposes to use a variational data assimilation framework to tackle the problem of wall interference corrections in the presence of non-linear flow phenomena. The main idea is to use a gradient-based optimization technique to minimize the difference between an experimentally measured quantity of interest and the same quantity obtained by means of a RANS simulation. The control parameters contain the free-stream angle of attack and Mach number, and terms correcting the approximations introduced by eddy viscosity turbulence models. The former return the corrections to the angle of attack and Mach number, while the latter treat the main source of error in RANS simulations. Two parameterizations of the turbulence model are presented in this paper, each one correcting the error introduced by the turbulence model in its own way. One re-calibrates the balance of terms within the turbulent transport equations, while the other directly assimilates the Reynolds stress anisotropy tensor, thus bypassing the limitations introduced by the Boussinesq hypothesis. The two methodologies are tested on transonic 2D cases, and are shown to yield more accurate corrections than traditional methods.

Despite a steady decrease in the overall usage of wind tunnels due to the increasing reliability and speed of computational fluid dynamics (CFD), ground testing remains a fundamental part of the development of an air vehicle. In this context, one of the new trends is represented by data assimilation (DA), whereby experimental and numerical measurements are combined together in an attempt to extract more complete and accurate information from the available data. This work presents an assessment of a variational data assimilation framework for the problem of wind tunnel wall interference corrections. The methodology minimizes the discrepancy between experimental measurements and the corresponding values from a Reynolds-averaged Navier-Stokes (RANS) simulation by optimally tuning the free-stream angle of attack and Mach number, as well as a corrective field for the turbulence model. The framework is first validated on synthetic data, and then tested on two cases at high Mach number and low angle of attack using pressure coefficients on the model’s surface as the reference measurements. The assimilated pressure coefficients match the reference ones better than those obtained from a linear correction technique. Furthermore, assimilating also a corrective term in the turbulence model improves the quality of the results with respect to assimilating only the angle of attack and Mach number.

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.

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.

This experimental study focuses on the aerodynamic interaction effects that occur between an over-the-wing (OTW) propeller and a wing boundary-layer. An OTW propeller is positioned above the hinge-line of a wing featuring a plain flap. The measurements are carried out with and without axial pressure gradients by deflecting the flap and by extending the wing in streamwise direction to simulate a flat-plate configuration, respectively. Wing pressure taps and phase-free particle-image-velocimetry (PIV) are used to quantify the time-averaged interaction effects, while embedded microphones and phase-locked PIV are used to analyze unsteady interaction effects. Results show that the propeller generates an adverse pressure gradient on the wing surface which increases linearly with thrust and decreases as the blade tipclearance is increased. The pressure gradient is partially caused by the slipstream contraction, which creates a streamwise velocity deficit near the wall immediately behind the propeller disk. Moreover, the rotation of the propeller blades generates pressure fluctuations on the surface, the amplitude of which exceeds both the pressure fluctuations produced by the tip-vortices and the time-averaged pressure effect of the slipstream. Consequently, the propeller triggers flow separation when the flap is deflected. A parametric study of different propeller locations indicates that increasing the tip-clearance is not an effective way to mitigate flow separation. However, displacing the propeller half a radius upstream induces a Coanda effect which allows the flow to remain attached.