Target Airfoil Research Articles

Comparative Analysis of Aerodynamic Lift and Drag of Commercially Used Air-foil Incorporated with Co-flow Jet

In this research, a baseline airfoil is modified by incorporating a Co-flow Jet into the airfoil. Co-flow jet is also called zero net mass flux as the same amount of highly pressurized air is injected towards the streamlines of an airfoil from the leading edge and sucks this air without changing the amount in the trailing edge using an air pump. A performance boost is expected for the use of the Co-flow Jet. An airplane’s performance is based on a few parameters such as the coefficient of lift, coefficient of drag, boundary layer separation, and many others. In this research, a comparative analysis is run between a baseline airfoil and a CFJ airfoil. NACA-4412 is chosen as the target airfoil. The research shows how the incorporation of CFJ in an airfoil increases the coefficient of lift (CL) and decreases the coefficient of drag (CD). Thus, it is possible to maintain a better CL/CL ratio by using CFJ. The research also mentions the increase in stall angle due to the use of CFJ. It is a computational analysis performed in the software Ansys. For simplicity, the analysis is limited to two-dimensional analysis. This research may bring a better improvement in the aircraft and aerodynamic sectors. Though the results obtained from the research may not fully be similar to the hypothesis, it has been tried to maintain that the results of the simulation hardly differ from the theoretical analysis.

Parallel vortex body interaction enabled by active flow control

An experimental investigation has been conducted to demonstrate the utility of active flow control as a disturbance generator for vortex body interaction studies. The technique is used to explore the flow physics of parallel vortex body interaction between two NACA 0012 airfoils in series. Experiments were carried out at a chord-based Reynolds number of 740,000 relative to the first airfoil. Active flow control in the form of nanosecond pulse-driven dielectric barrier discharge plasma actuation, originating close to the leading edge, was used to produce vortex shedding from the upstream (disturbance) airfoil at various frequencies ($$0.038 \le F^+ \le 0.762$$). These vortices were characterized, showing reduced circulation and diameter with increasing frequency, before examining the downstream wake-airfoil interactions. Time-resolved pressure and phase-locked PIV measurements were taken on the downstream (target) airfoil for multiple angles of attack. For $$F^+ \le 0.5$$, the target airfoil is subject to strong oscillations from the wake of the disturbance airfoil that lead to large fluctuations in lift and pitching moment. However, a further increase in $$F^+$$ reattaches the flow over the disturbance airfoil and no major vortex body interactions are observed on the target. Governing parameters for this type of vortex body interaction are explored, and differences between isolated and non-isolated encounters as well as the presence of a viscous response are examined. Finally, means to alleviate loads caused by the incident vortex are explored.

Aerodynamic Performance Estimation of Camber Morphing Airfoils for Small Unmanned Aerial Vehicle

This paper proposes a methodology to harvest the benefits of camber morphing airfoils for small unmanned aerial vehicle (SUAV) applications. Camber morphing using discrete elements was used to morph the base airfoil, which was split into two, three, and four elements, respectively, to achieve new configurations, into the target one. . In total, thirty morphed airfoil configurations were generated and tested for aerodynamic efficiency at the Reynolds numbers of 2.5 × 105 and 4.8 × 105, corresponding to loiter and cruise Reynolds numbers of a typical SUAV. The target airfoil performance could be closely achieved by combinations of 5 to 8 morphed configurations, the best of which were selected from a pool of thirty morphed airfoil configurations for the typical design specifications of SUAV. Interestingly, some morphed airfoil configurations show a reduction in drag coefficient of 1.21 to 15.17% compared to the target airfoil over a range of flight altitudes for cruise and loiter phases. Inspired by the drag reductions observed, a case study is presented for resizing a SUAV accounting for the mass addition due to the morphing system retaining the benefits of drag reduction.

Aerodynamics of Discrete Location Camber Morphing Airfoils in Low Reynolds Number Flows

Objectives: This paper focuses on morphing the base airfoil similar into that of the target airfoil for the application of small unmanned aerial vehicle in low Reynolds number regime. Methods/Statistical Analysis: In this study, discrete location camber morphing approach was used to achieve the morphing configurations of base airfoil. Discrete location camber morphing method was classified into single, two and three location morphing configuration based on the number of morphing locations. The aerodynamic performance of morphed airfoil configurations were studied at the low Reynolds numbers of 2.5 × 10 5 and 3.9 × 10 5 using XFLR5 - e N method. Findings: The base airfoil for this study was selected as NACA0012. The E 207 airfoil which has better aerodynamic performance in low Reynolds number regime was selected as the target airfoil. Eleven different morphing configurations of base airfoil were developed for this study, which falls under these three classifications. Two out of eleven morphed configurations have similar geometric features and equivalent performance as that of E 207 for different range of angles of attack. These two morphed configurations showed a rise of about 3% in maximum aerodynamic efficiency compared to the target airfoil for the tested Reynolds number. Out of these two morphed configurations, one belongs to two location morphing method and another belongs to three location morphing method. This study also reveals that atleast one morphing location has to be closer to maximum camber position of the target airfoil to achieve an effective morphing. There is a possibility for switching between these two morphed configurations as they have two common morphing locations during the flight. Application/Improvements: This type of camber morphing can be positively applied in small unmanned aerial vehicles to achieve better aerodynamic performance over the entire flight mission.

Asymmetric distributions in pressure/load fluctuation levels during blade-vortex interactions

The separation distance (h/c) between vortex and blade is a critical parameter in blade-vortex interaction, which strongly affects the magnitudes of induced pressure fluctuations (peak-to-peak values) and noise on helicopters. Parallel blade-vortex interaction (BVI) was studied in a subsonic wind tunnel to better understand the flow physics of BVIs with small h/c values and further improve the current BVI alleviation schemes. A vortex was generated by a sudden pitch motion applied to an airfoil through a pneumatic system. As the vortex interacted with a downstream target airfoil in a parallel manner, the vortex dynamics and induced loads were studied. The induced pressure/load fluctuations were obtained from the unsteady pressure data on the target airfoil, and the flow fields during the interaction were visualized using particle image velocimetry (PIV). Despite the fact that a symmetric airfoil was studied, the pressure data clearly displayed an asymmetric distribution in fluctuation levels above and below the chord line (h/c=0). Maximum pressure and load fluctuations occurred when the vortex passed on one side of the symmetric airfoil with small h/c, rather than in the head-on case (h/c=0). This side of the airfoil is defined as the strong fluctuation side, with the opposite side defined as the weak fluctuation side. PIV data revealed that this phenomenon was directly related to the asymmetry in velocity field induced by the vortex. The load fluctuation level also exhibited weak dependence on the decay in vortex strength due to both viscous- and inviscid-type interaction with the airfoil. However, the role of vortex decay became insignificant at higher Reynolds number due to lower decay rate. These findings cleared some confusion in previous studies and suggested that the magnitude of BVI fluctuations could be reduced by passing the airfoil below a clockwise vortex or above a counter-clockwise vortex.

Vortex dynamics during blade-vortex interactions

Vortex dynamics during parallel blade-vortex interactions (BVIs) were investigated in a subsonic wind tunnel using particle image velocimetry (PIV). Vortices were generated by applying a rapid pitch-up motion to an airfoil through a pneumatic system, and the subsequent interactions with a downstream, unloaded target airfoil were studied. The blade-vortex interactions may be classified into three categories in terms of vortex behavior: close interaction, very close interaction, and collision. For each type of interaction, the vortex trajectory and strength variation were obtained from phase-averaged PIV data. The PIV results revealed the mechanisms of vortex decay and the effects of several key parameters on vortex dynamics, including separation distance (h/c), Reynolds number, and vortex sense. Generally, BVI has two main stages: interaction between vortex and leading edge (vortex-LE interaction) and interaction between vortex and boundary layer (vortex-BL interaction). Vortex-LE interaction, with its small separation distance, is dominated by inviscid decay of vortex strength due to pressure gradients near the leading edge. Therefore, the decay rate is determined by separation distance and vortex strength, but it is relatively insensitive to Reynolds number. Vortex-LE interaction will become a viscous-type interaction if there is enough separation distance. Vortex-BL interaction is inherently dominated by viscous effects, so the decay rate is dependent on Reynolds number. Vortex sense also has great impact on vortex-BL interaction because it changes the velocity field and shear stress near the surface.

Research on Correction to Fitting Factors of Shape Function and Convergence of Wind Turbine Airfoils

Based on the functional expression methods of wind turbine airfoils, the method of the correction to parameter factors of shape function by iterative calculation in the principle of making the residual error minimum between the fitting airfoil and the target airfoil is presented in this paper, which makes the fitting precision improved compared with the parametric representation of original airfoils. The method of the correction to parameter factors of shape function proposed in this paper is used for parametric representation of more than 20 kinds of typical airfoils and then the geometric and aerodynamic convergence are intensive studied. The results show that the minimal order of the integrated expression of airfoils is decreased by the proposed method in this paper and the mathematical models of airfoils which facilitate the unification of optimal design are established.

Validation of a smart structural concept for wing-flap camber morphing

The study is aimed at investigating the feasibility of a high TRL solution for a wing flap segment characterized by morphable camber airfoil and properly tailored to be implemented on a real-scale regional transportation aircraft. On the base of specific aerodynamic requirements in terms of target airfoil shapes and related external loads, the structural layout of the device was preliminarily defined. Advanced FE analyses were then carried out in order to properly size the load-carrying structure and the embedded actuation system. A full scale limited span prototype was finally manufactured and tested to: • demonstrate the morphing capability of the conceived structural layout; • demonstrate the capability of the morphing structure to withstand static loads representative of the limit aerodynamic pressures expected in service; • characterize the dynamic behavior of the morphing structure through the identification of the most significant normal modes. Obtained results showed high correlation levels with respect to numerical expectations thus proving the compliance of the device with the design requirements as well as the goodness of modeling approaches implemented during the design phase.

Multidisciplinary design optimization of adaptive wing leading edge

Adaptive wing can significantly enhance aircraft aerodynamic performance, which refers to aerodynamic and structural optimization designs. This paper introduces a two-step approach to solve the interrelated problems of the adaptive leading edge. In the first step, the procedure of airfoil optimization is carried out with an initial configuration of NACA 0006. On the basis of the combination of design of experiment (DOE), response surface method (RSM) and genetic algorithm (GA), an adaptive airfoil can be obtained whose lift-to-drag ratio is larger than the baseline airfoil’s at the given angle of attack and subsonic speed. The next step is to design a compliant structure to achieve the target airfoil shape, which is the optimization result of the previous step. In order to minimize the deviation of the deformed shape from the target shape, the load path representation topology method is presented. This method is developed by way of GA, with size and shape optimization incorporated in it simultaneously. Finally, a comparison study with the Solid Isotropic Material with Penalization (SIMP) method in Altair OptiStruct is conducted, and the results demonstrate the validity and effectiveness of the proposed approach.

Airfoil design via optimal control theory

In airfoil design, one problem of greatinterest is to find the target airfoil profile to achieve a giventarget velocity distribution. It can be formulated as an optimalcontrol problem, with the control being the airfoil profile andthe governing equation being the full potential equation in thetransonic regime. To discretize the problem, one approach is toemploy the finite element method. In the discretized space, adirect relationship between the objective function and the unknownprofile co-ordinates can be defined via the finite element basisfunctions. Moreover, it is advantageous to derive the gradient inthe discretized space rather than the continuous space to avoidcontamination by discretization errors. In this paper, thisapproach is studied. In particular, a new formulation is proposed.A novel decomposition of the discrete space for the potentialfunction, the gradient is derived and an efficient algorithm usingthe quasi-Newton method is described. In generating and adjustingthe mesh during iterations, the elliptic mesh generation techniqueis used.

PARALLEL BLADE–VORTEX INTERACTION

Laser–Doppler velocimetry was employed to experimentally investigate two-dimensional airfoil–vortex interaction. Vortices were generated by sinusoidally oscillating a NACA 0012 airfoil about its quarter-chord at a reduced frequency ofk=2·05 and an amplitude of ±10° in angle of attack. The target airfoil, a NACA 632A015 was immersed in the wake, two chord lengths downstream of the vortex generator's trailing edge. Phase-averaged velocity measurements of the flow around the target airfoil were made with the latter at angles of attack ofα=0° (unloaded blade) andα=10° (loaded blade). A close encounter with a counterclockwise rotating vortex was studied for both angles of attack and a head-on collision which splits the counterclockwise rotating vortex in two was investigated for collision which splits the counterclockwise rotating vortex in two was investigated forα=10°. Vorticity fields were constructed and surface pressure fluctuations on the airfoil were determinded. It was found that the most violent interaction occurs when the disturbing vortex passes very near over the loaded airfoil. On the other hand, with head-on collision, the disturbing vortex is split, the pressure spikes are drastically reduced and the vortex disintegrates. The opportunity arises therefore to unload a blade during the short interval of vortex interaction, in order to reduce the severity of blade–vortex interaction.

Unsteady Pressure Measurements for Parallel Vortex-Airfoil Interaction at Low Speed

Two-dimensional vortex-airfoil interaction at airfoil chord Reynolds number 6 x 10 4 is studied experimentally by immersing an instrumented NACA 0012 airfoil downstream of a vortex generator airfoil that pitches in a tailored nonsinusoidal form. A clockwise rotating vortex is generated to approach the target airfoil from left to right. When the vortex interacts with a nonlifting airfoil, it creates rather sharp and local variations of surface pressure in the vicinity of the leading edge.

Airfoil design via transpiration and optimal control theory

In airfoil design, one of the problems of great interest is to find the target airfoil profile to achieve a given target velocity distribution. This problem can be formulated as an optimal control problem, and there are different ways of formulating and solving it numerically [14]. The transpiration technique provides a very powerful tool to simulate airfoil profile geometry movement without the need for mesh regeneration. On the other hand, the least-squares technique proposed in [8] provides a robust and efficient algorithm for solving the full potential equation. Here, these two techniques will be employed to give a new design method. Both the analysis and the design problem are formulated as a single optimal control problem and the conjugate gradient method is applied to solve the problem iteratively.

Head-on parallel blade-vortex interaction

An experimental and computational study was carried out to investigate the parallel head-on blade-vortex interaction (BVI) and its noise generation mechanism. A shock tube, with an enlarged test section, was used to generate a compressible starting vortex which interacted with a target airfoil. The dual-pulsed holographic interferometry (DPHI) technique and airfoil surface pressure measurements were employed to obtain quantitative flow data during the BVI. A thin-layer Navier-Stokes code (BV12D), with a high-order upwind-biased scheme and a multizonal grid, was also used to simulate numerically the phenomena occurring in the head-on BVI. The detailed structure of a convecting vortex was studied through independent measurements of density and pressure distributions across the vortex center. Results indicate that, in a strong head-on BVI, the opposite pressure peaks are generated on both sides of the leading edge as the vortex approaches. Then, as soon as the vortex passes by the leading edge, the high-pressure peak suddenly moves toward the low-peak-reducing in magnitude as it moves--simultaneously giving rise to the initial sound wave. In both experiment and computation, it is shown that the viscous effect plays a significant role in head-on BVIs.