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.

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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.

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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.

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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. @en

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.@en

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.@en

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.

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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.

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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.@en

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.

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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.

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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.

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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.

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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. @en

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. @en

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.

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This paper addresses the design of swirl recovery vanes for propeller propulsion in tractor configuration at cruise conditions using numerical tools.Amultifidelity optimization framework is formulated for the design purpose, which exploits low-fidelity potential flow-based analysis results as input for high-fidelity Euler equation-based simulations. Furthermore, a model alignment procedure between low- and high-fidelity models is established based on a shapepreserving response prediction algorithm. Two cases of swirl recovery are examined. The first is the swirl recovery by the trailing wing, which leads to a reduction of the lift-induced drag. This is achieved by the optimization of the wing twist distribution. The second case is swirl recovery by a set of stationary vanes, which leads to production of additional thrust. In the latter case, four configurations are evaluated by locating the vanes at different azimuthal and axial positions relative to the wing. An optimum configuration is identified where the vanes are positioned on the blade-downgoing side downstream of the wing. For the configuration and conditions examined, the wing twist optimization reduces the induced drag by 3.9 counts (5.9% of wing-induced drag), whereas the optimized 4-bladed SRVs lead to an induced-drag reduction of 6.1 counts (9.2% of wing-induced drag). @en

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. @en

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.@en

The swirl recovery vane (SRV) oriented in the slipstream of the propeller can in principle recover the swirl effect and thus would improve the propulsion performance in terms of thrust production and propulsive efficiency. The present study employs the design of experiments (DoEs) method to optimize the geometry of the specific SRV for Fokker 29 propeller for the sake of further enhancing the thrust generation and swirling recovery. First, orthogonal experiment was employed to identify the most significant factors, which directly influence the thrust production. Second, steepest ascent method was used to search the optimum range of target factors through climbing and factorial experiments. The resulting optimal solution was evaluated by the center composite experiment. Results show that the thrust generated by the SRV has been increased significantly (11.78%) after optimization at the design point, and a 0.66% increment in the total efficiency of the propeller-SRV system has been obtained. For the off-design point, an increment of the total efficiency (2.10%) can be observed at low rotating speed. Additionally, the optimized SRV is able to correct the out-flow behavior at the tip region of the vane, where the tip vortex and swirl kinetic energy loss is weaken, and the thrust distribution along the spanwise direction tends to be more uniform.@en