High-lift Configuration Research Articles

The mitigation of airframe noise using a Krueger flap as a leading-edge device in a high-lift configuration

High-lift device is a potent airframe noise contributor. The conventional slat, commonly used in a high-lift device, is known as one of the dominant noise sources. Here, we combine an experimental and a numerical approach to investigate the noise generation of a conventional slat and two Krueger flaps as leading-edge devices in a high-lift configuration. To reduce the computational cost, a short span of the entire high-lift device is selected as the focusing region for scale-resolving and acoustic computations using Improved Delayed Detached Eddy Simulation (IDDES) coupled with the Ffowcs-Williams and Hawkings (FWH) analogy. Simulation results show that turbulent cross flows induce slightly lower-level noise spectra at frequencies less than 300 Hz when the two spanwise sides of the FWH permeable integral surface are closed. Wind-tunnel test results show that the reference and optimal Krueger configurations effectively attenuate the dominant tone of the conventional slat by 4 and 6 dB, respectively. Besides, the optimal configuration is very effective in noise reduction in a wide frequency range. However, the reference Krueger configuration increases the noise level by 0 − 4 dB at frequencies exceeding 5, 500 Hz. Overall, the optimal Krueger configuration is a good choice for high-lift device noise mitigation.

Wall-Modeled Large-Eddy Simulation Method for Unstructured-Grid Navier–Stokes Solvers

This paper reports on the implementation and assessment of a wall-modeled large-eddy simulation (WMLES) methodology in an unstructured-grid, node-centered flow solver, FUN3D, that is developed and supported at the NASA Langley Research Center. Finite-volume (FV) and finite-element (FE) discretization schemes considered in the study provide formal second-order spatial accuracy. Large-eddy simulations (LES) resolve large-scale turbulent-flow features, and small-scale effects are modeled using the Vreman subgrid-scale model. At solid-wall boundaries, a shear-stress model is employed to provide a proper boundary-flux closure. The nonlinear equations are integrated in time using either an optimized backward difference formula or an implicit multistage Runge–Kutta temporal scheme. The implicit equations at each time step are solved by strong nonlinear iteration schemes. WMLES demonstrations are shown for two high-lift configurations, namely, the McDonnell Douglas 30P30N multi-element airfoil and a NASA High-Lift Common Research Model. Results show that the WMLES approaches implemented in the FV and FE discretization methods produce consistent solutions and are capable of capturing key aerodynamic characteristics and flow structures for high-lift configurations at a wide range of angles of attack, including maximum-lift conditions. In the 30P30N example, correct trends in the variations of integrated aerodynamic forces and moments, surface pressure distributions, and boundary-layer profiles are captured as the Reynolds number is increased.

A critical assessment of Navier–Stokes and lattice Boltzmann frameworks applied to high-lift configuration through a multiresolution approach

The present work focuses on a thorough assessment of the influence of two very different numerical approaches, namely, Navier–Stokes (NS) and the lattice Boltzmann method (LBM), to simulate the flow past a three-element airfoil through zonal detached eddy simulation (ZDES). Both computations (ZDES-NS and ZDES-LBM) are compared to the reference results, namely, a wall-resolved large eddy simulation (WRLES) as well as the experimental data. It is shown that despite very different numerical modeling, the two ZDES provide very consistent results, with the first- and second-order statistics obtained with equivalent accuracy in the impingement region. In light of present results, the ZDES mode 2 (2020) turbulence model within an LBM framework appears as a judicious combination for high-lift flow applications owing to its robustness regarding the use of very fine isotropic Cartesian grids. In addition, ZDES-NS exhibits a very good agreement with both references, especially WRLES despite having 40 times less nodes.

Aerodynamic Prediction of the High-Lift Common Research Model with Modified Turbulence Model

Aerodynamic simulations were carried out in the study presented in this paper focusing on the stall performance of the High-Lift Common Research Model obtained from the fourth AIAA High-Lift Prediction Workshop. Various turbulence models of Reynolds-averaged Navier–Stokes simulations are analyzed. A modified version of the transitional k−v2¯−ω model was developed to enhance stall prediction accuracy for high-lift configurations with a nacelle chine. The vortex generator, three-element airfoil, and high-lift model are numerically simulated. The results reveal that implementing a k−v2¯−ω model with the separation shear layer fixed notably enhances the stall prediction behavior for both the three-element airfoil and high-lift configuration without affecting the prediction of the vortex strength of a vortex generator. Moreover, incorporating rotation correction into the SPF k−v2¯−ω model improves the prediction of vortex strength and further enhances stall prediction for the high-lift configuration. The relative error in predicting the maximum lift coefficient is less than 5% of the experimental data. The study also investigated the impact of the nacelle chine on the stall behavior of the high-lift configuration. The results demonstrate that the chine vortex can mitigate the adverse effects of the nacelle/pylon vortex system and increase the maximum lift coefficient.

Prediction of aerodynamic characteristics of high-lift Common Research Model in ground effect

Abstract Reynolds Averaged Navier-Stokes (RANS) simulations are performed to investigate the aerodynamic characteristics of the NASA Common Research Model (CRM) in its high-lift (HL) configuration in close proximity to the ground. The RANS simulations are conducted at a moderate Reynolds number of $Re = 5.49 \times {10^6}$ and $M = 0.2$ with the use of the Spalart-Allmaras (SA) turbulence model. out of ground effect (OGE) simulation results are validated against available wind tunnel data before proceeding to in ground effect (IGE) simulations. The obtained computational results in the immediate vicinity of the ground with asymmetric aircraft attitudes demonstrate significant changes in the longitudinal and lateral-directional aerodynamic characteristics, which should be taken into account in flight dynamics analysis of aircraft during take-off and landing in crosswind conditions.

Advances in Wall Interference Corrections for Two-Dimensional Lifting Bodies

A novel approach to correcting the effects of wall interference has been developed for low-speed closed-wall wind tunnels using a nonlinear least-squares optimization process independent of the model geometry. Potential flow singularities are used to represent the test object, and the method of images is used to represent the tunnel walls. Trust-region reflective optimization enables the locations and strengths of singularities to be determined iteratively with only wall pressure measurements as independent inputs. Results are obtained from a range of two-dimensional and three-dimensional airfoil test cases using both computational fluid dynamics simulations and wind tunnel experiments, including a NACA 0012 with deployed spoiler, a NACA 0015, a NACA 4415, and a McDonnell Douglas 30P30N high-lift configuration. The nonlinear least-squares wall-correction method is shown to demonstrate a considerable improvement over the conventional two-variable technique.

Numerical Study of Flow Control on Simplified High-Lift Configurations

Enhancement in the aerodynamic performance of wings and airfoils is very notable when Active Flow Control (AFC) is applied to Short Take-off and Landing aircraft (STOL). The present numerical study shows the application of steady, pulsed and synthetic tangential jets applied to the plain flap shoulder of a modified NASA Trapezoidal Wing. Pulsed jets are modeled by sinusoidal and square waveforms while synthetic jets are modeled only by pure sine waveform. The freestream airflow conditions are Mach number equal to 0.2 and Reynolds number equal to 4.3 million based on the mean aerodynamic chord. The presented simulations are two-dimensional and based on RANS for steady jet cases and URANS for pulsed and synthetic cases, compiled with the open-source suite SU2 and adapted for time varying boundary conditions. Numerical results for modified configurations based on the same baseline wing profile considering different leading edges, jet slot height, flap position, blowing mass flow, type and frequency of the jets are presented. Curves of pressure coefficient distribution revealed a substantial influence upstream of the AFC, around the slat and main element. The jet slot height analysis showed that the lift gain is also influenced by the slot size due to the change of the local flow velocity considering the same blowing momentum coefficient. Regarding the jet frequency, no significant differences on the lift coefficients were found between the reduced frequencies F+ equal to 1 and 2. Results of aerodynamic loads showed an improved lift coefficient in relation to the baseline airfoil when pulsed and steady jets are employed. Pulsed jets under square waveform were effective even at high deflected flap condition at 50°, with a significant gain in the lift coefficient of 36%, in relation to the uncontrolled case, combined with a drag reduction of 20%, and a decrease in mass flow up to 49% in relation to the steady jet for the same lift gain. Although sine and square waveform results presented similar improvements for lift, the drag is around 15% higher for the former. When compared with the steady jet case, the mass flow reduction is 36% for the sinewave. Synthetic jets with zero-net-mass-flux proved superior to the baseline conventional multi-element airfoil only with deployed flap at 50°, where a modest lift improvement of 5% was observed.

HLPW-4: Wall-Modeled Large-Eddy Simulation and Lattice–Boltzmann Technology Focus Group Workshop Summary

A summary of the nine submissions to the Wall-Modeled Large-Eddy Simulation and Lattice–Boltzmann (WMLESLB) Technical Focus Group (TFG) at the 4th High lift Prediction Workshop is provided. The focus of this TFG was to assess the current capabilities of WMLES and LB methods on a complex high-lift configuration across a wide range of angles of attack. Analysis of the submitted data suggests that >250 M spatial degrees of freedom are needed to accurately predict pitching moments at high angles of attack due to large pressure gradients present on the outboard slat and main element for α>17° (corrected for free-air). While some scatter is reported in pitching moment coefficient at the low-angles of attack (α<11°), excellent agreement is observed between submissions near the CL,max state. Objective superiority of WMLES methods over Reynolds-averaged Navier–Stokes (RANS) can be seen in terms of lack of excess outboard separation; a majority of the WMLES and LB submissions predict wedge-shaped separation patterns consistent with the experimental oil flow. The in-tunnel simulations show excellent agreement with the experiment in terms of a) integrated loads, b) surface flow-topology, and c) mechanism for the onset of inboard stall. Further evidence is provided to demonstrate both qualitative and quantitative superiority of the WMLES submissions over RANS.

Fourth High-Lift Prediction/Third Geometry and Mesh Generation Workshops: Overview and Summary

The Fourth AIAA Computational Fluid Dynamics (CFD) High Lift Prediction Workshop and the Third Geometry and Mesh Generation Workshop were held collaboratively with the common goal of assessing the numerical prediction capability of current-generation computational fluid dynamics technology for swept medium-/high-aspect-ratio wings in high-lift configurations. A key aspect of this joint endeavor was the use of technology focus groups: an innovative new approach for workshops involving close collaboration between participants. These groups, which included both mesh-generation and flow-solver experts, worked to accelerate advancements for their particular methodologies by addressing key questions of importance before the workshop. The high-lift version of the NASA Common Research Model (CRM-HL) configuration was the focus of this workshop. Measured experimental wind-tunnel data were available for comparison. The workshop also included a two-dimensional turbulence model verification exercise based on the CRM-HL wing shape. Altogether, 44 participants submitted a total of 184 datasets of CFD results. This paper provides a high-level summary of the results and conclusions from the workshop. Like at past workshops, fixed-grid Reynolds-averaged Navier–Stokes methods continued to be inaccurate and inconsistent for predicting high-lift flow physics near maximum lift conditions. However, mesh adaptation definitively brought more consistency. Scale-resolving methods appeared most promising for predicting high-lift flow physics near maximum lift.

Approaches to the Numerical Simulation of the Acoustic Field Generated by a Multi-Element Aircraft Wing in High-Lift Configuration

Approaches to the Numerical Simulation of the Acoustic Field Generated by a Multi-Element Aircraft Wing in High-Lift Configuration

Wake Vortex Far-Field of a Generic Transport Aircraft Affected by Oscillating Flaps

The influence of oscillating trailing-edge flaps of a generic transport aircraft on the middle and far fields of the wake vortex system is investigated. The aim is to be able to reduce separation distances between aircraft during approach in order to increase capacity at airports. By introducing a certain disturbance by means of oscillating flaps in the near field, the long-wavelength Crow instability should be excited in the further downstream development. The evolution of the wake vortex system is investigated using a large-eddy simulation approach. The wake vortex system is examined up to 133 spans downstream for two high-lift configurations. A nonactuated configuration with statically deflected flaps is compared to a configuration with actuated flaps. The results show that the long-wavelength Crow instability cannot be excited by actuated trailing-edge flaps. However, a faster decrease of the dimensionless circulation over the downstream position can be observed for the actuated configuration compared to the nonactuated configuration, due to a faster diffusion of the rolled-up vortex pair. Furthermore, the averaged induced rolling moment on an aircraft [with the International Civil Aviation Organization’s (ICAO’s) classification of “light”] encountering the wake vortex of the actuated configuration (ICAO’s classification of “heavy”) is up to 32.5% lower than for a follower aircraft, which is entering the wake vortex of the nonactuated configuration.

Investigation of Improvement Design on Aileron Surface Flow State of High Lift Configuration in BWB

The aileron is one of the most important tools for adjusting the roll attitude of the aircraft, but the surface flow state of the aileron is likely to be affected by high-lift devices. In this paper, by using computational fluid dynamics (CFD) simulations and wind tunnel tests, the Krueger flap effects on the surface flow state of ailerons in a typical blended wing body civil aircraft were investigated. In order to increase the lift, deflecting the Krueger flap makes the flow separation occur on the aileron surface of BWB civil aircraft. This way of the surface stall that the flow separation in the aileron zone first appears at the wing tip rather than at the wing root is unreasonable for civil aircraft. For the above problem, a sensitivity analysis of the design parameters of the Krueger flaps was carried out. The results indicate that the angle of the outboard Krueger flap mainly affects the flow separation of the ailerons. Its length affects the pitch moment tremendously, while its width slot affects the pitch moment slightly. Finally, the design principles of the BWB Krueger flap for the improvement aileron surface flow state were proposed, and the redesign of the BWB high lift configuration significantly improved the flow state of the aileron zone at a minimal cost of aerodynamic characteristics without losing the existing great aerodynamic performance.

Disturbance region update method with preconditioning for steady compressible and incompressible flows

The disturbance region update method (DRUM) was an acceleration methodology devoted to steady compressible flows. In this work, we develop DRUM for steady incompressible flows, named the disturbance region update method with preconditioning (DRUM-P). Compared with the conventional global update method (GUM), DRUM-P adopts the advective and the viscous dynamic computational domains (DCDs) updated with the time-marching process to achieve a computational speedup for compressible and incompressible flows. When DRUM is extended to incompressible flows, the stiffness of the governing equations would slow down the convergence to the steady state and thus delay the contraction of the DCDs. By employing the local preconditioning technique, DRUM-P overcomes those difficulties and exploits the full convergence potential of DRUM for incompressible flows. Besides, the operation of contracting the advective DCD is modified in subsonic flows, where the advective DCD is allowed to contract from the downstream before the upstream. A new viscous-dominated criterion of the viscous DCD based on the total pressure loss and the total enthalpy loss is proposed to improve accuracy in practical cases with complex geometric shapes and flow features. Eight test cases demonstrate the efficiency and accuracy of DRUM-P. Firstly, DRUM-P achieves remarkable convergence efficiency for solving compressible and incompressible flows, especially the compressible and incompressible mixed flows, e.g., the flow over high-lift configuration. The acceleration ability benefits from the preconditioning technique, the two high-flexible DCDs, and the modified operation of the advective DCD. Secondly, the numerical results obtained by DRUM-P are in excellent agreement with those obtained by GUM. Moreover, the numerical results verify that the proposed viscous-dominated criterion is grid-independent and more accurate in practical cases.

Aerodynamic Shape Optimisation Using Parametric CAD and Discrete Adjoint

This paper presents an optimisation framework based on an open-source CAD system and CFD solver. In this work, the high-fidelity flow solutions and surface sensitivities are obtained using the primal and discrete adjoint formulations of SU2. This paper shows the direct use of CAD models for optimisation by developing a CAD system application programming interface and creating a link between the CAD-MESH-CFD analysis. A methodology to obtain geometric sensitivities is introduced, enabling the calculation of accurate gradients with respect to CAD variables and the deformation of the analysis mesh during the optimisation process. This methodology guarantees that the new surface mesh lies exactly on the CAD geometry. The optimisation framework is applied to a rectangular wing and a three section high-lift aerofoil configuration derived from the NASA CRM-HL configuration. Both geometries are created using FreeCAD. The performance objectives are to decrease the drag while constraining the lift to be above a desired value. The twist distribution of the wing was parameterised within the CAD system, allowing the minimisation of the induced drag by obtaining a nearly elliptical lift distribution. For the high-lift configuration, the position and rotation of the flap and slat were parameterised with respect to the original section; the final optimised positions yield a drag reduction of approximately 16.5%. These results show that the CAD parameterisation can be reliably used to obtain efficient optimums while operating directly on the CAD geometries.

Higher-order aerodynamic numerical simulations in compressible RANS framework with inverse-ω scale variable

The use of higher-order methods to robustly compute turbulent flows governed by the multi-equation turbulence model in Reynolds-averaged Navier-Stokes (RANS) framework is a challenging topic in the computational fluid dynamics (CFD) community. In this paper, we propose two aspects of measures involving the physical model and the numerical scheme. Firstly, a λ-scale (=1/ωn) is derived instead of the well-known ω in SSG/LRR Reynolds-stress model (RSM) and SST eddy viscosity model (EVM) due to its natural boundary conditions for viscous walls and help to the positivity preservation of dissipation rate ε in turbulence equations when n=1/(2m),m=1,2,3.... Secondly, a geometric-conservation-preservation negativity-correction (GCPNC) method is proposed on the implicit time integration procedure with right-hand-side (RHS) term discretized by high-order weighted compact nonlinear schemes (WCNSs). It can guarantee the positivity of normal Reynolds stresses or turbulent kinetic energy. Numerical tests with real-life compressible flows are reported to demonstrate the robustness and accuracy of the present methods. Relevant cases include the supersonic square duct flow, the flows over delta wing, wing-body configuration and high-lift configuration.

Numerical Simulation of the Flow in the ONERA F1 Wind Tunnel

This paper presents the outcome of numerical simulations of the flowfield inside the ONERA F1 wind tunnel, which is a large-scale pressurized low-speed wind tunnel mostly used to investigate the takeoff and landing performance of aircrafts in high-lift configuration. The simulations were carried out under Reynolds-averaged Navier–Stokes modeling assumptions. Results were compared to available experimental data characterizing the tunnel flowfield, especially in the test section. The overall flow physics is well captured by the simulations. However, when looking at the result with a more demanding level of accuracy, some important characteristics of the velocity distribution in the test section, such as flow upwash and total pressure distribution, are not replicated. Reasons are discussed in the paper, and some recommendations are derived for future similar simulations.

A confidence-based aerospace design approach robust to structural turbulence closure uncertainty

We propose a confidence-based design approach robust to turbulence closures model-form uncertainty in Reynolds-Averaged Navier–Stokes computational models. The Eigenspace Perturbation Method is employed to compute turbulence closure uncertainty estimates of the performance targeted by the optimizer. The magnitude of the uncertainty estimates is exploited to establish an indicator parameter associated to the credibility of numerical prediction. The proposed approach restricts the optimum search only to design space regions for which the credibility indicator suggests trustworthy RANS model predictions. In this way, we improve the efficiency of the design process, potentially avoiding designs for which the computational model is unreliable. The reference test case consists in a two-dimensional single element airfoil resembling a morphing wing section in a high-lift configuration. Results show that the prediction credibility constraint has a non negligible impact on the definition of the optimal design.

Effect of the trailing-edge flap on tones due to self-excited oscillation within the leading-edge slat cove

To study the tonal noise characteristics of the leading-edge slat of high-lift configurations with and without a deployed trailing-edge flap, experiments are conducted in the D5 aero-acoustic wind tunnel at Beihang University. The numerical simulation method is used to obtain the necessary flow information. The experimental results show that low to mid frequency tonal noise generated by self-excited oscillation within the slat cove is dominant and its corresponding frequencies are basically unchanged whether the flap is deployed or not. However, the primary mode of self-excited oscillation within the slat cove switches to a higher one when the flap is deployed. Further analysis results demonstrate that variation of the primary mode is found to be closely related to the flow characteristics in the self-excited oscillation feedback loop. The number of the primary mode is generally proportional to the ratio between the vortex shedding frequency and the self-excited oscillation frequency. The flap being deployed results in an increase in the effective angle of attack of both the main wing and slat, which leads to a thinner separating boundary layer, thus increasing further the vortex shedding frequency and this ratio.

Ffowcs Williams – Hawkings analogy for near-field acoustic sources analysis

The paper expands the scope of applying the Ffowcs Williams – Hawkings integration method. We propose using the acoustic field generated from time-dependent data stored on the FW-H control surface as the same common field for computational acoustic beamforming and dynamic mode decomposition methods to analyze the aerodynamic noise sources. We exemplify that it leads to obtaining mutually consistent and complementary information for reliable prediction of acoustic sources characteristics in the process of inverting data produced by a CFD simulation. Moreover, as the results of applying computational acoustic beamforming and dynamic mode decomposition methods depend on many geometric and algorithmic inputs, the proposed approach makes it possible to use various sets of the latter for a comprehensive analysis of obtained inversions and to form the final answer by an averaging procedure. We illustrate this by taking advantage of fast generating the examined acoustic field snapshots in any required region by the FW-H integration method for the recently developed new inverse computational acoustic beamforming algorithm and the standard dynamic mode decomposition method when carrying out a sensitivity study of the predicted acoustic source. The capabilities of the developed approach are demonstrated on the data of CFD scale-resolving simulation of turbulent flow over the 30P30N high-lift configuration.

Wind Tunnel Installation Effects on Japan Aerospace Exploration Agency’s Standard Model

This paper investigates the interference effects of wind tunnel floor and walls on the Japan Aerospace Exploration Agency’s High-Lift Configuration Standard Model (JSM) as an extension to a paper published in the Journal of Aircraft, Vol. 56, No. 3, May–June 2019, pp. 1080–1098. An unstructured hybrid grid generator, MEGG3D, was applied to JSM wing–body and wing–body–nacelle–pylon configurations to generate grids for free-air flow conditions. These grids were then reused in an investigation of the interference effects of Japan Aerospace Exploration Agency’s low-speed wind tunnel walls to better compare the computational results, both with each other and with experiment. The geometry of the wind tunnel used in our computational simulations is described in detail so that they can be reproduced. A Reynolds-averaged Navier–Stokes flow solver for unstructured hybrid grids, the Tohoku University Aerodynamic Simulation (TAS) code, was used for the simulations. The computational results indicated that the standoff and test section floor boundary layer significantly affected the aerodynamics of the JSM at high angles of attack. Better agreement with experiment was observed for the wing–body configuration in the aerodynamic characteristics near maximum lift if the effects of the standoff and floor boundary layer were considered.