Development and Verification of Acoustic Metamaterial Liner for Noise Suppression in Slat

The noise sources of commercial aircrafts may be split in two categories: the first category covers all the noise generated by the propulsion unit while the second category is associated with the airframe itself. Based on the development of high-bypass low noise turbofan engines the contribution of airframe noise in approach and landing configuration is enhanced, thus playing a growing role in aircraft design. The high-lift devices, especially the upper slat trailing edge portion and the flap sidewalls as well as the landing gear are the main contributors to the airframe noise during take-off and landing. This paper presents the results of a Computational Aeroacoustics (CAA) analysis for a 2D high-lift profile as part of the Airbus strategy for low-noise high-lift design. Previous, primarily experimental [1,2,3], research activities have shown that the slat (also in conjunction with the main wing leading edge) is one of the major contributors to high-lift airframe noise. With focus on the slat, especially the pressure side (cove area) and the upper trailing edge are to be identified as being the primary source of slat noise. Thus, the current activities concentrate on investigating slat noise by injecting a single test-vortex upstream of the slat hook (i.e. the lower slat trailing edge) for different Re-numbers, Ma-numbers and angles of attack into a steady state flow field coming from CFD.

*† ‡ Slat noise originates due to unsteady flow within the slat cove and in the trailing edge wake of the slat. An aeroacoustic study was conducted in the 5½-by-4-foot Cambridge University Markham wind tunnel in order to further understand and treat slat noise. A twopart wing and slat model was used for this research. Separate low and high frequency phased microphone arrays consisting of 48 channels each were used to determine the source strengths associated with the noise within the slat cove. The aerodynamic forces on the models were determined separately using a three component overhead force-balance. Modifications to the airfoil geometry were made in order to analyze their effectiveness in reducing the overall leading edge noise. The alterations reported here are filling in the slat cove and drooping the leading edge geometry. The slat brackets were identified as significant contributors to the cumulative noise from the slat region and thus a supplementary study on a simple plate bracket, a cylindrical and an I-beam shaped bracket. The models were tested at angles of attack from six to sixteen degrees, and at flow speeds of 20, 30, and 40 m/s. Boundary layer trips were used to simulate the full-scale, high Reynolds number flow. The tests confirmed that the slat is a significant source of noise. A drooped profile is successful in removing the slat noise with a small reduction in aerodynamic performance. A filled cove on the other hand reduces the sources of noise by a few decibels. However, the concept is less effective than hoped due to the enhancement of other sources of noise. The force balance results show that, after filling in the slat cove, the model stalls prematurely. Thus, for modifications reported here, any acoustic benefits are likely to be accompanied by aerodynamic performance penalties. The I-beam supports induce a greater acoustic intensity than a cylindrical support, and both are noisier than a plate support. Additionally, the contribution to the overall noise in the slat region, from supports, is greater at the lower freestream velocities.

The turbulent flowfield above the wing of a delta—canard—configuration at moderate (α = 15°) and high (α = 24°) angle of attack was measured at a Re—number of 0.97.106 in a wind tunnel by hotwire anemometry. Leading edge flap settings of ηl.e. = 0° and ηl.e. = −20° were used. At moderate angle of attack and deflected leading edge flap a strong vortex originates from the side edge of the non-deflected inboard wing leading edge part. This inboard wing vortex is located close to the fuselage. It is a dominant flow feature and forms the center of the vortical flow separating from the wing surface. The separation line is clearly different from the leading edge flap hinge line. At high angle of attack the flow separates at the leading edge for both the non—deflected and deflected leading edge ease. The resulting leading edge vortex is subject to breakdown dose to the apex. In case of the deflected leading edge, the interaction of inboard wing vortex and leading edge vortex remits in decreased downstream expansion of the burst vortex, also reducing turbulence intensity levels.

This paper presents an analysis of the slat noise for Hybrid Wing Body (HWB) aircraft, based on a database from a 3% scale wind tunnel test. It is shown that the HWB slats are one of the dominant noise components, characterized by its broad spectral shape with a peak frequency that depends on both the mean flow velocity and the aircraft angle of attack, the former following the conventional Strouhal number scaling and the latter explainable by the dependence of the coherence length of the slat unsteady flows on the aircraft angle of attack. While the overall levels of the slat noise are shown to approximately follow the fifth power law in the flow Mach number, the effects of the Mach number manifest themselves in the noise spectra in both the amplitude and the spectral shape. The slat noise amplitude is shown to also depend on the angle of attack, assuming a minimum in the range of 3 to 5 degrees and increasing when the angle of attack moves away from this range. These features are all modeled and incorporated in slat noise prediction methodologies, extending the prediction capability from conventional aircraft designs to HWB configurations. Comparisons between predictions and data show very good agreements both in various parametric trends and in the absolute levels. The HWB aircraft is designed to operate at angles of attack much higher than those of conventional aircraft. This is shown to significantly increase the HWB slat noise. To further illustrate, the test data are extrapolated to full scale and compared with the slat noise of the Boeing 777 aircraft, showing that the former is higher the latter.

This study investigates the slat noise of a two-dimensional scaled, unswept, and untapered MD30P30N high-lift model. The experimental data refer to aeroacoustic and aerodynamic measurements in a closed-section wind tunnel for a wide range of angles of attack (from up to the stall; approximately at 18 deg) and Mach numbers between 0.07 and 0.1. Three slat configurations (the original MD30P30N, another with a higher slat deflection, and one with smaller slat gap and overlap) are studied experimentally. The signal processing applied to the acoustic data involves conventional beamforming enhanced by two deconvolution algorithms, namely, DAMAS and CLEAN-SC. An original variation of the beamforming cluster approach that is based on the coherence level among microphone pairs is introduced, and it improves the results obtained by DAMAS. Below and above 12 deg angles of attack, the slat noise is very small and mostly below the wind-tunnel background noise for all configurations. Between and 12 deg angles of attack, the slat noise spectra are substantially affected by the slat configuration, although it always contains a dominant low-frequency content, a midfrequency broadband noise, and a single high-frequency broad peak. Within this range, the lower angles of attack display the strongest low-frequency narrowband peaks. In fact, at lower angles of attack, the low-frequency narrowband peaks scale with a Mach power above 10.

Use of oscillatory actuation of the leading edge of a thin, flat, rigid airfoil, as a potential mechanism for control or improved performance of a micro-air vehicle (MAV), was investigated by performing direct numerical simulations at low Reynolds numbers. The leading edge of the airfoil is hinged at one-third chord length allowing dynamic variations in the effective angle of attack through specified oscillations (flapping). This leading edge actuation results in transient variations in the effective camber and angle of attack that can be used to alleviate the strength of the leading edge vortex at high angles of attack. A fictitious-domain based finite volume approach [Apte et al., JCP 2009] was used to compute the moving boundary problem on a fixed background mesh. The flow solver is three-dimensional, parallel, secondorder accurate, capable of using structured or arbitrarily shaped unstructured meshes and has been validated for a range of canonical test cases including flow over cylinder and sphere at different Reynolds numbers, and flow-induced by inline oscillation of a cylinder. Flow over a plunging SD7003 airfoil at two Reynolds numbers (1000 and 10,000) was computed and results compared with those obtained using AFRL’s high-fidelity solver [Visbal, AIAA J. (2009)] to show good predictive capability. To assess the effect of an actuated leading edge on the flow field and aerodynamic loads, two-dimensional parametric studies were performed on a thin, flat airfoil at 20 degrees angle of attack and Reynolds number of 14,700 (based on the chord length) with sinusoidal actuation of the leading edge over a range of reduced frequencies (k=0.57-11.4) and actuation amplitudes. It was found that high-frequency, low-amplitude actuation of the leading edge significantly alters the leading edge boundary-layer and vortex shedding and increases the mean lift-to-drag ratio. This study indicates that the concept of an actuated leading-edge has potential for development of control techniques to stabilize and maneuver MAVs in response to unsteady perturbations at low Reynolds numbers. The summer research at AFRL’s computational sciences division has resulted in several opportunities for future collaborations with AFRL scientists and researchers. At Oregon State, new projects for senior students are initiated to build and modify the existing physical setup and measure lift and drag coefficients.

: This thesis investigated the feasibility of using a portion of the leading edge (LE) strake hinged along the longitudinal axis as a roll control device for a high performance aircraft at high angles of attack (AOA). A wind tunnel test was conducted to gather static force and moment data for use in a six degree of freedom computer simulation. Asymmetric strake deflections, both dihedral and anhedral, were investigated. The longitudinal coefficients were little affected by strake deflection, but the lateral directional coefficients showed a nonlinear, but repeatable, behavior with strake deflection. Comparisons to be published data indicate that the strakes produce similar behavior for different aircraft designs. Simulations of the aircraft response to the strakes showed that an improvement over current roll performance could be obtained by combining the positive strake deflection with the ailerons up to 38 deg. AOA, after which the strakes alone produced the best roll performance. Sideslip and AOA must be closely controlled or the air-craft will either not roll, or will depart during the roll. The roll performance using hinged strakes at high AOA is compared to roll performance using differential (LE) flaps. The diffential LE flaps produce comparable roll rates with less sideslip than produced by the hinged strakes. However, the possibility exists of combining hinged strakes with differential (LE) flaps for improved roll performance.

It is well known that aircraft undergoing high angle of attack excursions (e.g., the Cobra maneuver) may experience asymmetric forebody vortex shedding under certain conditions. The detachment of these vortices from the forebody is similar to that observed from von Karman vortex shedding from cylinders. A series of experiments, where a von Karman vortex street wake was made to impinge upon a 70-degree delta wing, was conducted at Wichita State University. The aim is to better understand the mixing mechanism that occurs at these high angle of attack flight regimes in a less Reynolds number sensitive environment. A von Karman wake having a frequency similar to the forebody shedding process is used. As the von Karman vortex filaments are entrained by the shear layer, they appear to wrap themselves around the core. There is a temporal correlation to the burst location of the delta wing’s leading-edge vortices, indicating modulation of the burst location. Additionally, the core becomes distorted. It has been found that the vortex burst location is moved forward towards the apex when subjected to the von Karman wake. Nomenclature c Wing root chord d Cylinder diameter FOV Field of View LE Leading Edge Rec Chord Reynolds number, U∞c/ν ReD Diameter Reynolds number, U∞d/ν s Coordinate along root chord U∞ Freestream Velocity x Coordinate along horizontal direction α Angle of attack (deg) Introduction Delta wings have evolved over the years and are now used primarily in the form of leading edge extensions on many fighter aircraft. As these aircraft become more and more maneuverable, the understanding of the physics of time-dependent unsteady flows is becoming more important. In particular, if accurate computational models of these maneuvers are to be developed, it is necessary to understand the mechanisms involved in features such as vortex bursting and mixing on the lee side of the delta wing. This is true of maneuvers such as the Cobra and other “hyper-agile” maneuvers. To understand the situation, it is important to briefly discuss some of the experimental findings in the areas of vortex bursting behavior, the Cobra maneuver physics, and forebody shedding. On the Impingement of a von Karman Vortex Street on a Delta Wing Ismael Heron and Roy Y. Myose Department of Aerospace Engineering Wichita State University, Wichita KS 67260-0044 It is well documented that delta wings at a fixed angle of attack generate lift by separating a shear layer of air (or fluid, such as water) at the leading edge, and this shear layer forms two strong counter-rotating vortices on either side of the wing. These leadingedge (LE) vortices undergo small fluctuations in space, but remain relatively fixed over the suction side of the delta wing, and are critical to the generation of lift, as they produce a large suction peak on the surface. Two much smaller vortices, the secondary vortices, are also formed, as seen in Figure 1. In other words, LE vortices are the result of a balance between vorticity being generated at the leading edge, and the ability of the flow field to convect said vorticity along the vortex core. The LE vortices are not stable, and at some point their coherent structure will undergo a dramatic change, expanding around the core, slowing down axially, and either forming a bubble or a spiral, with the spiral form being more predominant at Reynolds numbers of interest to delta wing designers. This change, called vortex burst or breakdown, is dependent on the aspect ratio of the wing, angle of attack, pressure gradients, yaw angle, and swirl angle of the vortex, among others. The exact reason for this bursting is not known, but research has focused on two general areas. (a) The flows upstream and downstream of the vortex burst are two separate and very different flows, and the vortex burst is a necessary feature, similar to a hydraulic jump. (b) The core of the LE vortex serves as a mechanical waveguide for longitudinal waves; these waves either coalesce, or they become critical, thereby triggering the burst. Regardless of the burst triggering mechanism, the effect of the vortex burst is to reduce the lift generated by the delta wing. If the delta wing is pitched to a given angle of attack (α) and then maintained at that angle until the transient flow features die down, it is said to be tested under “static” conditions. As this 22nd Applied Aerodynamics Conference and Exhibit 16 19 August 2004, Providence, Rhode Island AIAA 2004-4731 Copyright © 20 * Graduate Research Assistant, Student Member AIAA. † Associate Professor, Associate Fellow AIAA. Copyright ©2004 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.

Observing the progress in the technology of unmanned combat aerial vehicles (UCAVs) than it can be foreseen that in the future the role of manned combat aircraft will be taken over more and more by unmanned systems. The abandonment of pilots allows for more freedom in the aerodynamic design of the vehicle in regard to weight and acceleration, however, new stealth constraints have a severe impact. The design of UCAV configurations is driven by the special requirements of upcoming missions, as for instance the capabilities of long endurance fl
\nights joined with a low observability. The new demands have a crucial impact on the aerodynamic shape, and, hence, require new solutions for maneuver con-
\ntrol in respect to integration of engine in- and outlets, actuators and other devices. Additionally the new capabilities of long endurance 
\nights has to be joined with low observability. In particular UCAVs are suited to the exploitation of non-conventional control technologies, such as aerodynamic morphing, 
\nflow control, or thrust vectoring. This paper presents a numerical investigation of innovative morphing technologies for aircraft applications and explores the feasibility for such technologies to enhance the maneuverability of unmanned combat aerial vehicles. In Aerodynamics the morphing is understood as a smooth, continuous change in geometry of the outer surface, e.g. the twist of complete wing can be changed in order control role moments. In the considered case morphing is used to provide additional 
\now control mechanisms taking into account the constraint of fllow radar signature. In particular morphing can be used to change the geometry of the airfoil leading edge to control the 
\now around the wings in order to generate additional lift or induce role moments. This paper is concerned with the investigation of the feasibility and effectiveness of morphing devices for the aerodynamic control of UCAVs. Focal point of this work is the morphing of leading edges in order to generate additional lift or role moments for UCAVs 
\nflying in high angle of attack mode. Therefore, the objective of the numerical investigations is the evaluation of the potential of morphing leading edges especially for the enhancement of the maneuverability of Delta- and Lambda-wing configurations with their vortex-dominated 
\now field. Of special interest is 
\now control at high angles of attack (AoA) by targeting an minimized shape adaptation in order to generate a needed additional lift, to induce role moments or to reduce the risk of the potential deleterious effects of vortex breakdown. The morphing of the leading edge must not confused with classical Kruger leading edge 
\nflaps. Although very effective in regard to the increase of lift such fl
\naps can't be used for UCAVs since gaps between fl
\naps and the wing have to be avoided in order to fulll the hard conditions of radar camouflage
\n.
\nThe presented work is part of the DLR Project FaUSST which is the successor of the DLRProject UCAV-2010. The UCAV-2010 project was set up to identify and assess UCAV relevant technologies. The investigations have covered the pre-design process with fast,
\nlow fidelity methods as well as the detailed examination using high fidelity methods like URANS simulations. The verification of the promissing technologies and tools have been done by virtual and experimental prototypes. The combined research has lead to a Lambda wing type UCAV configuration with a medium sweep of the leading edge. The outcoming
\nDLR-F17 conguration has been mainly derived from the so called Saccon geometry developed by EADS-MAS for the RTO/AVT-161 task group.
\nSince the capability of medium to high AoA maneuverability will be investigated the angle of attack has been varied from 6 to 16 degrees.
\nAfter presentation of the technical part, the results of leading edge manipulation are discussed. One focal point is the comparison of the impact of differently morphed leading edge shapes on the fl
\night performance at specific 
\nflight situations. The aerodynamical coefficients, in particular Cl, Cd and Cm, are related to morphing parameters of the changed shape of the leading edges. Another key aspect is the varied effective angle of attack of
\nthe middle part of the wings, see Figure 2. In case of a basic high angle of attack the effective AoA (due to morphing) can change the vortical 
\nflow above the deformed wings enormously, even when the geometrical angle of the airfoil has been morphed only slightly.
\nThis work discusses the effectiveness of counter-rotating morphing of leading edges, since the requirements in regard to radar signature do not allow large geometry changes. This point reduces the degree of freedom of any shape morphing drastically. Thereto, only proper combinations of left and associated right wing deformations have been analyzed and discussed. After the discussion of the most interesting deformations an evaluation of the presented test cases in regard to the maneuverability enhancement of UCAVs concludes the presentation.

The Roe format and LU-SGS method are used to discretize the NS equation, and the second-order precision unilateral differential discrete rigid body dynamic equations are coupled to solve the NS equation and the rigid body dynamics equations at Ma=0.2, and numerically simulate the 80°swept delta wing. In the rock history under different leading edge shapes, the influence of the leading edge shape on the rock characteristics of the delta wing is studied. The calculation model consists of three different leading edge shape delta wings with a lower sharpened, a sharpened tip and a double-sided sharpened tip. The results show that the shape of the leading edge significantly affects the bifurcation and amplitude characteristics of the delta wing rock. The taper angle of attack of the lower sharpened leading edge delta wing is the largest, and the splitting angle of the upper sharpened leading edge delta wing is the smallest. When the angle of attack is 25°, all three wings form equal-amplitude oscillations in the form of limit cycles. The amplitude of the lower sharpening leading edge delta wing is the smallest, and the amplitude of the upper sharpening leading edge delta wing is the largest; when the angle of attack increases At 30°, the three wings eventually form a constant-amplitude oscillation in the form of a limit ring, but the double-sided sharpened, upper-curved leading edge delta wing will flip and the other side will swing in equal amplitude.

In previous work a RANS based simulation technique for the simulation of broadband slat noise was established. Good agreement was found between predicted and measured slat noise spectra. These predictions were based on 2D CAA computations and a connection to 3D measured data is only possible assuming a certain functional behavior of the spanwise coherence of the essential slat noise source. For this purpose, results from trailing edge noise measurements were used. In this work the simulation strategy is extended to 3D CAA computations, resolving the spanwise slat noise coherence as part of the CAA computations. The considered wing span is one main-chord, which is large enough to establish a realistic 3D problem for the turbulence as well as for the sound radiation. The Fast Random Particle Mesh (FRPM) method is applied for this study to generate fluctuating sound sources from steady RANS turbulence statistics. The study is conducted for the 30P30N airfoil with 0.457m main chord. The Mach number is 0.17 and the angle of attack is 4∘. Good agreement is found between the previous 2D and the 3D results as well as with unsteady simulations published in the literature. The influence of sweep on slat noise generation is studied.

Experimental investigation on the effect of slat geometrical configurations on aerodynamic noise

The aerodynamic forces and moments of a flexible delta wing in pitching motion were experimentally studied in a low-speed wind tunnel. Three types of flexible delta wing were investigated, the flexible parts of which were 44, 70 and 99% of the delta wing. Aerodynamic characteristics were different among the three types of flexible and completely hard delta wing, and it was found that the winding-up of the leading edge of the delta wing is the key factor for determining the leading edge vortex on the upper side of the wing and the pressure distribution on the windward side. Lift, drag, and pitching moment formed a hysteresis loop with an angle of attack in pitching motion, particularly in a region with a large attack angle, accompanied by leading edge vortex breakdown. The flow visualization of leading edge vortices was also carried out to explain the dynamic characteristics of the delta wings.

Leading-edge erosion of wind turbines is caused by particles in wind. Impacts at a high velocity damage the blades severely and lower the aerodynamic efficiency. The aim of this paper is to estimate the flow rate hitting on the leading edge of the airfoil NACA4412, which forms the tip of a wind turbine blade, with the geometry referred to NREL offshore 5-MW baseline wind turbine. The methodology is to first compute the coordinates of the points on the boundary of the airfoil that is defined to be the leading edge, then with trigonometry relationships of the airfoil, the length that is passed through by wind can be found. Finally, relating them together, a differential equation is established. It reveals how the angle of attack and flow rate through the leading edge are correlated. With graphs plotted, the equation states that changing the angle of attack from 6 to 7 has 6.2% less erosion while Cl/Cd is dropped by 4.6% only. Hence, for situations where leading edge erosion and aerodynamic efficiency are considered equally important, it is better to vary the angle of attack to 7 degrees rather than the optimal 6 degrees.

A turbulence model study was performed to analyze the flow around the Tubercle Leading Edge (TLE) wing. Five turbulence models were selected to evaluate aerodynamic force coefficients and flow mechanism by comparing with existing literature results. The selected models are realizable k-ε, k-ω Shear Stress Transport (SST), (γ−Reθ) SST model, Transition k-kl-ω model and Stress- ω Reynolds Stress Model (RSM). For that purpose, the TLE wing model was developed by using the NACA0021 airfoil profile. The wing model is designed with tubercle wavelength of 0.11c and amplitude of 0.03c. Numerical simulation was performed at chord-based Reynolds number of Rec = 120,000. The Computational Fluid Dynamic (CFD) simulation reveals that among the selected turbulence models, Stress- ω RSM estimated aerodynamic forces (i.e. lift and drag) coefficients closest to that of the experimental values followed by realizable k-ε, (γ−Reθ) SST model, k-ω SST model and k-kl-ω model. However, at a higher angle of attacks i.e. at 16° & 20° k-ω SST model predicted closest drag and lift coefficient to that of the experimental values. Additionally, the critical observation of pressure contour confirmed that at the lower angle of attack Stress- ω RSM predicted strong Leading Edge (LE) suction followed by realizable k-ε, (γ−Reθ)SST model, k-ω SST model and k-kl-ω model. Thus, the superiority of Stress- ω RSM in predicting the aerodynamic force coefficients is shown by the flow behavior. In addition to this pressure contours also confirmed that k-kl-ω model failed to predict tubercled wing aerodynamic performance. At higher angles of attacks k-ω SST model estimated aerodynamic force coefficients closest to that of the experimental values, thus k-ω SST model is used at 16° & 20° AoAs. The observed streamline behavior for different turbulence models showed that the Stress- ω RSM model and k-kl-ω model failed to model flow behavior at higher AoAs, whereas k-ω SST model is a better approach to model separated flows that experience strong flow recirculation zone.