Airfoil Surface Research Articles

Experimental Study on the Effect of Porous Media on the Aerodynamic Performance of Airfoils

Porous media has potential applications in fluid machinery and in aerospace science and engineering due to its excellent drag-reduction properties. We carried out experimental time-resolved particle image velocimetry (TR-PIV) research, laying porous media with different pore densities on the suction side of an airfoil in the low-turbulence recirculation wind tunnel of Nanjing University of Aeronautics and Astronautics to study the effects and mechanisms of porous media on airfoil aerodynamic performance. We also used a smooth airfoil model in the experiment for comparison. Comparing the aerodynamic forces, pressure distributions, and the airfoil’s suction side flow field, we found that the porous media with different pore densities had different effects on the airfoil’s aerodynamic performance. Although the porous media with 20PPI (pores per inch) increased the pressure drag and reduced the airfoil lift, it considerably reduced the friction drag, thus significantly improving the airfoil’s aerodynamic force. The flow visualization results indicated that, although the porous media with 20PPI reduced the circulation of flow velocity around the suction side of airfoil, it also destroyed the vortex structure, broke the low-frequency large-scale vortex into a high-frequency granular vortex, inhibited the amplitude of vortex fluctuation, reduced the shear stress on the airfoil surface, weakened the vortex energy of different modes, and accelerated the vortex’s spatio-temporal evolution.

Numerical Analysis of Slotted Wings Using Fluid-Structure Interaction

For shorter landing and take-off path in airports, the aircrafts should reduce their speed with keeping high lifting force. This paper is to identify solutions to increase the lift force of the wing significantly under several flight scenarios (such as takeoff and landing) using leading-edge slats and their relationship with the dynamic parameters of the aerodynamic wing. The study is performed by the use of ABAQUS 2016 software. The problem is solved for turbulent flow and 2-dimensional composite wing at constant Reynolds’s number of (6.49 × 105) and constant boundary conditions. Various depths have been used for the auxiliary airfoil at constant width and gap. All stresses at the wing base were obtained. The pressure distribution on the airfoil surface was determined, air velocity distribution was tracked over the surface, lift and drag forces and their coefficients were computed. The results show that the highest value of the lift coefficient is 0.489 at the depth (-3 %) of the wing chord, it decreases when the depth of the slat becomes zero %, and the rise returns with increasing depth to (4 %), but it does not reach the maximum value, while the highest drag coefficient was (1.89) at depth (4 %) of the wing chord. The maximum value of Von Mises stress was found at depth of 4 % with value of 1.605 × 105 Pa.

Calibration and error compensation of scanner-based robotic belt grinding system

Turbine blade is one of the key components of turbomachinery, and the machining quality of its airfoil surface has an important impact on the operating efficiency, service life, and safety of the turbine. Robotic abrasive belt grinding has become an important means of blade airfoil surface machining. Based on the self-developed robot grinding system, this paper proposes a scanner-based system calibration method to determine the relationships among coordinate systems of measurement, robot base, tool, and workpiece. An accurate and efficent calibration method of workpiece coordinate system based on rapid surface segmentation of point cloud using CAD model is established. An blade error surface is constructed by B-spline interpolation to compensate system errors such as absolute positioning error of robot, calibration error, and blade shape error. The calibration process is illustrated at the end of this paper. The effectiveness of the calibration and error compensation methods proposed in this paper are verified by the simulation results.

Vortex evolution near the Coanda surface of a circulation control airfoil with pulsed actuation

A special vortex evolution process was found near the Coanda surface for circulation control airfoil with pulsed actuation. For the jet-on process, two successive shedding vortices from the lower lip edge and the separation vortex on the Coanda surface collide and interact with each other, and periodic vortex shedding forms after several vortex shedding and collision cycles. For the jet-off process, the periodic shedding vortices were replaced by a row of vortices with reverse rotation direction from the upper lip edge, and then reverse vortices broke into small-scale vortices.

Pitch–Plunge Equivalence in Dynamic Stall of Ramp Motion Airfoils

Large-eddy simulations (LES) are employed to investigate the pitch–plunge equivalence of an SD7003 airfoil undergoing constant ramp motions at Reynolds number . The equivalence is constructed based on the geometric effective angle of attack according to the quasi-steady thin-airfoil theory. Two rates of descent (or pitch up) are analyzed for different Mach numbers in order to investigate the effects of compressibility on the evolution of the dynamic stall vortex (DSV). During the onset of the DSV and its transport along the airfoil surface, remarkable similarities are found between pitch and plunge in terms of flow topology, aerodynamic loads, and signatures of wall pressure and friction coefficients. However, these flow similarities cease at high-load conditions as the DSV becomes more susceptible to the peculiarities of the airfoil motion, manifested here by different trailing-edge vortices. Employing a correction for the rotation-induced apparent camber effect present in the pitching case, which results from the quasi-steady thin-airfoil theory, improves the agreement between pitch and plunge. However, it is not sufficient to assimilate their disparate trailing-edge systems. Results also demonstrate that the limit angle at which pitch–plunge equivalence remains valid decreases for higher Mach numbers.

Flow and heat transfer studies of multijet impingement cooling for different configurations: A review

AbstractIncreasing the gas turbine engine's turbine intake temperature has long been a potential strategy for increasing the specific work output of the engine. However, the melting temperature of the turbine blades and vane material limits the maximum intake temperature. As a result, internal and external cooling techniques are commonly used to maintain the vane material in a safe condition. This study provided an overview of internal impingement cooling to highlight the significance of geometrical variations, such as flat plate, curve plate, and actual vanes. It was observed that flat and curved plate impingement heat transfer studies were reported extensively, whereas limited studies were found on the conjugate effects on airfoil surfaces. The importance of conjugate heat transfer studies and their impact has recently been described in the literature. In most of the literature, a wide range of instruments, such as Laser Doppler Velocimeter, Particle Image Velocimeter, liquid crystal sheets, and so forth, were used for experimental investigations. According to most studies, the local value of internal surface temperature and heat transfer coefficient are vital factors of local flow behavior. Jet‐to‐jet spacing, jet‐to‐plate spacing, jet hole diameter, and jet Reynolds numbers played a crucial role in both numerical and experimental analyses. Different geometric variations strongly influence flow behavior. Therefore, the usual method for determining interior temperature distributions and heat transfer coefficients by considering generalized geometries like the flat and curved plate may not produce accurate conjugate solutions. Most of the computational studies on the flat and curved plate indicate the usage of κ−ω shear stress transport and κ–ε realizable model to predict the heat transfer coefficient.

A new airfoil based on banana leaf midrib morphology for horizontal axis wind turbine

The airfoil profile of this study is based on plant morphology, which is rarely modeled in airfoil profiling. Designing, modeling, and analyzing the aerodynamics of one of the projected morphological parts of the banana leaf midrib that forms a 45-degree angle with the longitudinal axis as a distinct new airfoil. The profile of the LM-45 airfoil has a thickness of 16.28%, a chord length of 1500mm, a Reynolds number between 3E5 and 7E5, and a wind speed between 3 and 7 meters per second. An aerodynamic analysis using QBLADE and ANSYS FLUENT determined the glide ratio and pressure distribution across the airfoil surface. LM-45 has a maximum glide ratio of 123.07 at an angle of attack of 2 degrees and a Reynolds number of 7E5, and a minimum glide ratio of 70.02 at an angle of attack of 4 degrees and a Reynolds number of 3E5. As indicated by the glide ratio, the banana leaf midrib profile can be applied to the airfoil profile for wind speeds between 3 and 7 m/s at an attack angle between 2 and 4 degrees.

Airfoil trailing-edge noise and drag reduction at a moderate Reynolds number using wavy geometries

This study utilizes a hybrid aeroacoustic model to investigate how airfoils with spanwise wavy geometries can be used to reduce trailing-edge noise alongside improving the aerodynamic performance. A smooth airfoil is compared to four variants, which have spanwise surface waves of different wavelengths, at a Reynolds number of Re = 64 000 and an angle-of-attack of 1°. The first three variants have a geometry modified by a single wavelength, whereas the fourth has a surface composed of two wavelengths, which creates a more irregular surface variation. The results show that modest noise reductions of around 4 dB are achieved for the first three variants, but a much larger reduction of 17.7 dB is achieved for the fourth variant. The mechanisms behind the noise reduction are explored, and it is shown that the geometry reduces the spanwise correlation of the pressure fluctuations and also modifies the boundary layer dynamics, which contributes to the large reduction. It is further shown that a wavy geometry can reduce the drag force by reducing the shear stress over parts of the airfoil surface and by limiting the flow separation on the suction side. The fourth variant is also assessed across a wider range of angles (0°≤α≤4°) and is shown to produce less noise than the smooth wing across the entire range as well as a drag reduction for α≥1°.

Enhancing Performance of Small Capacity Horizontal Axis Wind Turbine Using Grooved Linearized-Chord Blades

Abstract In the current article, E216 airfoil blades with linearized chord of a small-scale horizontal axis wind turbine are numerically investigated to enhance the performance of the turbine. The blade is modified by including grooves on its suction side. Three-dimensional Reynold’s averaged Navier–Stokes (RANS) simulations are performed with Shear Stress Transport k–ω as a turbulence model. The computed power-coefficient results are first validated with previous measurements by the present authors on a wind turbine of 1 m rotor diameter at two wind speeds of 6 and 8 m/s. Another validation with measurements from the literature is performed via comparison of the pressure-coefficient distribution along surfaces of E216 airfoil at an angle of attack of 6 deg. Based on the successful computation, six different rotor models are numerically investigated with different numbers and locations of grooves created along the blade length, from the hub to the tip region, on the blade suction side. The lift and drag coefficients are compared for the different simulated rotor models. Furthermore, the torque coefficient, thrust coefficient, and static torque coefficient are obtained for the simulated models. It is found that the best-simulated configuration is Model 1 having one groove at 80% blade-chord length from the leading edge, with a maximum power coefficient of 0.429 at a tip-speed ratio of 4.75. The aerodynamic performance of the blades is strongly affected by the groove on the suction surface. The results show that Model 1 blade reduces the generated mean-axial force and its vibration frequency on the rotor.

Numerical and experimental study on power extraction performance of a semi-active flapping airfoil with bioinspired dimple

Inspired from the special structure of dragonfly wings (cambered and corrugated groove), in this work, the bioinspired dimple structure is introduced into the design of the airfoil to improve the power extraction performance of a semi-active flapping airfoil, and the characteristic parameters of the bioinspired dimple structure are optimized by the combination of numerical analysis and Taguchi method. The results show that the bioinspired dimple can effectively improve the performance of the flapping airfoil. The average power coefficient and power extraction efficiency of the optimized dimple flapping airfoil are 31.73% and 31.23% higher than that of the baseline smooth airfoil respectively, and the influence of the three considered characteristic parameters on the power extraction performance is ranked as: lx (the location of bioinspired dimple) >ly (the depth of bioinspired dimple)>w (the inlet width of bioinspired dimple). Further analysis the flow field around the flapping airfoil shows that the bioinspired dimple can promote the separation of the leading edge vortex from the airfoil surface during stroke reversal, meanwhile at the process of upstroke and downstroke, the bioinspired dimple can also delay the leading edge vortex to separate, which is the reason for the dimple airfoil extract more power from the fluid. Based on the numerical results, experiments test are also carried out to evaluate the power extraction performance of the optimized bioinspired dimple airfoil, and the results show that at the same experiment condition, more 12.81% average power coefficient is extracted by the optimized dimple flapping airfoil, which again verifies the effectiveness of the numerical analysis conclusion obtained in this work.

Experimental study on airfoil flow separation control via an air-supplement plasma synthetic jet

An air-supplement plasma synthetic jet (PSJ) actuator increases the air supplemental volume in the recovery stage and improves the jet energy by attaching a check valve to the chamber of a conventional actuator. To explore the flow control effect and mechanism of the air-supplement actuator, via particle image velocimetry experiments in a low-speed wind tunnel, the flow field and boundary layer characteristics of a two-dimensional airfoil surface under different actuation states were compared for different attack angles and jet orifices. The experimental results show that, compared with the conventional actuation state, the jet energy of the air-supplement PSJ is higher and the indirect mixing effect of the counter-vortex sequence produced by the jet-mainstream interaction is stronger. Furthermore, the boundary layer mixing effect is better, which can further suppress flow separation and improve the critical flow separation attack angle. Moreover, increasing the jet momentum coefficient can enhance the flow control effect. The findings of this study could provide guidance for the flow control application of air-supplement PSJs.

Analysis of stagger effects on Busemann supersonic biplane airfoil in shock tube tests by point diffraction interferometer method

The Busemann biplane concept is proposed for low-boom low-drag supersonic aircraft. Studies on the supersonic biplane have been focused on supersonic flow, whereas the understanding of transonic aerodynamic characteristics is essential for aircraft design. Significantly, the complicated flow field with a shock and boundary layer interaction between Busemann biplane elements in transonic flow remains unclear. In the present study, the transonic aerodynamic characteristics of the two-dimensional Busemann biplane are investigated by shock tube tests and numerical simulations, focusing on the flow field between the biplane elements. In addition, the stagger effects, which change the relative position of the lower and upper elements in the freestream direction to avoid the choked flow between the biplane elements, are also analyzed. A shock tube system creates a test flow at Mach numbers of 0.6 to 0.8, and the Reynolds number is set at 2.85 × 105. The Point Diffraction Interferometer (PDI) method is applied to determine the density distribution around the model and the pressure coefficient on the surfaces of the biplane elements. Symmetrical density distribution is confirmed in the Busemann biplane (the baseline model, or the biplane with no stagger). A normal shock wave is generated between the biplane elements. When the Mach number increases, the normal shock wave moves downstream. The normal shock position is quantitatively stable with small fluctuation amplitudes. The fringes are strongly curved near the surfaces of the biplane elements due to the flow separation near the airfoil surfaces. When the Mach number increases, the separation area increases. When the Mach number increases, the peak value of the pressure coefficient on the surfaces of the biplane elements increases. On the other hand, the normal shock wave between the biplane elements is eliminated in the staggered models. The changes in density distribution between the biplane elements are minor compared to the case of the baseline model, which indicates that the stagger reduces the choked flow between the biplane elements. The peak value of the pressure coefficient on the wing surface increases when the stagger value increases. The stagger effects on reducing the total drag coefficient of the biplane models are confirmed.

Numerical and experimental investigation of bionic airfoils with leading-edge tubercles at a low-Re in considering stall delay

The hydrodynamic characteristics and suction side flow behaviour of bionic airfoils are investigated numerically and experimentally at a low Reynolds number (Re). Hydrodynamic coefficient curves indicate that the hydrodynamic performance of the bionic airfoils is sensitive to the amplitude but not to the wavelength. When the amplitude is large enough, the bionic airfoils do not stall. As determined from numerical results and dye visualization, the bionic airfoils exhibit a double periodic flow pattern in the wingspan direction, and the trough-1 cross-section exhibits prominent separation at the leading edge, while the inflow flow attaches to the airfoil surface until the trailing-edge at the trough-2 cross-section. Because the vortex pairs formed by each group of adjacent tubercles interact with each other, the effective angle of attack (AOA) at the trough-2 cross-section is decreased, delaying the stall. In the post-stall region, the lift coefficient (CL) of each bionic airfoil is larger than that of the reference airfoil. The vortex pairs generated by the tubercles are completely mixed downstream, bringing more energy from the far wall to the near wall. The flow separation is delayed on both sides, and the effective AOA is decreased.

Effects of trailing-edge serration shape on airfoil noise reduction with zero incidence angle

When controlling the trailing-edge (TE) interference noise of airfoil, the design of the TE serration shape is still an open issue. To this end, the flow and noise generation for different TE serration shapes are explored by the wall-resolved implicit large-eddy simulation and acoustic analogy. The feather-like serrations are found to achieve the most prominent noise reduction among the four types of curved serrations, especially in the low-frequency range. With the aid of acoustic analogy, the coherence analysis of far-field noise produced by the dipole sources on the airfoil surface is performed. The results show that destructive interference is still the critical mechanism responsible for noise reduction. Considering only the dipole sources, we find that the feather-like serrated TE shape can obtain the best noise reduction performance among all the serrated cases. Furthermore, since flow structures are reorganized near the TE serrations, we investigated the flow noise sources quantitatively in the near field. In these cases, the noise source due to flow structures is suppressed to a greater extent in the feather-like serrated case near the TE serration roots. Consequently, the above findings indicate that the feather-like serration favors suppressing dipole and flow noise sources in the near field, which makes it an efficient configuration for reducing airfoil noise.

Successive Complementary Expansion Method for Solving Supersonic Laminar Flow Over Thin Airfoils& Numerical Simulation

The aim of the present study is two-fold. Firstly, it attempts to solve the asymptotic model of the 2D stationary potential flow of a supersonic viscous fluid over thin airfoils. In the case of a high Re number: Re=O(10SUP5/SUP), and a free stream Mach number of MSUB∞/SUB = 1.45, the asymptotic solution is built using the new successive complementary expansion method. The theoretical Mach number of the flow is estimated in the outer inviscid zone along the airfoil surface, and a correction term is added to account for the viscous boundary layer effect in order to get an accurate approximation throughout the whole domain.BR Secondly, for comparison purposes, numerical simulations of the NACA 43013 airfoil were run for the Mach number distribution in each area. On the one hand, when compared to numerical simulations, the asymptotic supersonic laminar viscous flow solution shows good agreement for the laminar boundary layer but begins to diverge slightly from 𝒙 = 𝟎, 𝟑.At this point, the laminar-turbulent transition starts, then the turbulent boundary layer develops towards the trailing edge. On the other hand, the Mach number distribution for the outer inviscid region matches the simulation and illustrates the change in velocity upon interaction with both the detached shock wave and the expansion waves.

Simulation of and Experimental Research on Rivulet Model on Airfoil Surface

The occurrence of aircraft icing can significantly affect flight performance. One of the most important aspects in the study of anti-icing technology for aircraft is the distribution of overflow water. Owing to the external airflow pressure, shear stress, and surface tension, the water film breaks up to form steady rivulets. Experiments on NACA0012 airfoil surfaces were conducted based on an open straight-flow and low-speed wind tunnel. Simultaneously, an engineered three-dimensional rivulet model considering the surface roughness was established based on the energy-minimum principle. A comparison between the simulation and experimental results shows that the errors in the water film breakup location and the flow velocity of rivulets are less than 20%, and the errors in the spacing and width of rivulets are less than 40%. In addition, the effects of surface temperature and uniform roughness on water film breakup were investigated. Furthermore, the rivulet model was applied to the numerical calculation of the thermal performance of hot-air anti-icing systems. The simulations reveal that the uniform roughness of the wing surface causes the water film to break earlier. As the surface roughness increases, the thickness, spacing, and width of the rivulets increase, and the rivulet flow velocity decreases.

Computational Study of the Plasma Actuator Flow Control for an Airfoil at Pre-Stall Angles of Attack

Large-eddy simulations of the flows over an NACA0015 airfoil were conducted to investigate a flow control authority of a dielectric barrier discharge plasma actuator at pre-stall angles of attack. The Reynolds number was set to 63,000, and angles of attack were set to 4, 6, 8, and 10 degrees. The plasma actuator was installed at 5% chord length from the airfoil’s leading edge. Good flow control authority was confirmed in terms of lift-to-drag ratio increase and drag reduction. These improvements mainly result from the reduction of the pressure drug, which is due to the change in pressure distribution accompanying the movement and shrink of the laminar separation bubble on the airfoil surface. Additionally, although flow control using a burst drive with a nondimensional burst frequency of six improves the lift-to-drag ratio at all angles of attack, the phenomena leading to the improvement differ between near-stall angles (10 and 12 degrees) and the other lower angles. At near-stall angles, the turbulent transition is rapidly promoted by PA, and the flow is reattached. Whereas, at the lower angles, the transport of two-dimensional vortex structures, which maintain their structures up to downstream and suppress the turbulent transition, makes the flow reattachment.

Numerical Investigation on Flow Control of a Hypersonic Airfoil by Plasma Synthetic Jet

Plasma synthetic jets (PSJs) have received widespread attention because of their rapid response capability. The use of PSJ shock and thermal effects for active flow control has been investigated in many recent studies. Through numerical simulation, this study found that the geometric position of a PSJ actuator on an airfoil surface is the key factor affecting flow control effects. Different geometric positions have different effects on the lift-drag characteristics of an airfoil, and the flow control mechanism has different components. The flow control mechanism can be divided into three types: the thermal effect, the coupling of the thermal effect and single shock, and the coupling of thermal effect and multiple-shock rebound. The concept of distance lplacement provides a reference for determining the geometric arrangement of a PSJ actuator. This study showed that in addition to improving the lift-drag characteristics of an airfoil by reducing drag, directly increasing lift is also a new choice. In addition, in a confirmatory experiment, the nanosecond pulse PSJ actuator pushed the bow shock in a Mach 5 flow; the longest distance was about 1% of the characteristic length of the airfoil.

Mechanisms of trace water vapor desublimation over airfoil in transonic cryogenic wind tunnels

Cryogenic wind tunnels can provide a larger operating Reynolds number compared to conventional ones. However, the internationally known “moisture contamination problem” resulting from the residual trace water vapor desublimation in the cryogenic wind tunnel may affect the accuracy of the aerodynamic data. Due to extreme difficulties in experiments, the detailed trace water desublimation characteristics remain unrevealed. An Euler–Euler two-phase flow model based on classical nucleation theory and a droplet growth model were established to predict the trace water vapor spontaneous desublimation in the nitrogen flow over airfoil in a transonic cryogenic wind tunnel. The proposed model was validated by experimental data obtained from the literature and showed good agreements. The verified model was applied to 0.152-m National Advisory Committee for Aeronautics 0012 airfoil in a Langley 0.3-m transonic cryogenic wind tunnel under a series of operating conditions. The simulated results reveal that the water vapor desublimation process in a transonic cryogenic wind tunnel can be divided into two patterns by a critical region. In pattern I, over the upper critical total temperature, the water vapor desublimates quickly during the rapid expansion of gas flow over the airfoil surface. In pattern II, below the lower critical total temperature, the water vapor is sufficiently supercooled and desublimates completely into small ice particles in the free-stream flow. For pattern I, the desublimation characteristics were analyzed in detail, and the influences of the desublimation process on the aerodynamic data were quantitatively evaluated, which can provide theoretical guidance for the practical operation of transonic cryogenic wind tunnels.

Numerical Simulation of Ground Effect on Circulation Control Airfoil

Circulation control (CC) is used extensively to control the attitude of rudderless aircraft experiencing ground effect in take-off and landing phases. The investigation of ground effect on airfoil CC is necessary to improve flight performance and quality in proximity to the ground. The aerodynamics and flow field of a modified NACA0012 airfoil with CC in ground effect are investigated with numerical simulations. The compressible Reynolds-averaged Navier–Stokes equations with the shear-stress transport k − ω turbulence model equations are solved using the Finite Volume Method (FVM). Simulation results show that the ground effect changes the lift increment per unit jet momentum coefficient, and CC can reverse the polarity ground effect. The effective angle of attack α E and the downwash space downstream airfoil are altered by the ground effect resulting in variations of airfoil surface pressure and lift. Unlike the unbounded flow field, the jet attachment distance is not only determined by the jet momentum coefficient but it can change with the ride height, which is the distance from the ground to the center of the semicircular Coanda surface, for the same jet momentum coefficient.