Symmetrical Airfoil Section Research Articles

Experimental and computational investigations of aerodynamic characteristics of a finite rectangular wing-in-ground effect

Wind tunnel experiments are carried out on a 3D wing-in-ground effect at pre- and post-stall angles of attack at a relatively low Reynolds number, Re ≈ 8.8 × 104. The rectangular wing, aspect ratio, AR = 6.4 used in the present work has a symmetric airfoil section, NACA0012 and is studied at different ranges of ground proximity and also compared to a wing of cambered airfoil section, NACA4415 not in ground proximity. Unsteady and time-averaged six-component aerodynamic characteristics, coefficients of Lift, Drag, Side forces, Pitch, Roll, Yaw moments at several angles of attack in the range – 8° < α < 18° are reported. Aerodynamic efficiency and drag polars are studied in lieu of ground proximity at several angles of attack. Effect of ground proximity on the occurrence of stall and [Formula: see text] is studied. Numerical calculations using steady-state CFD are also made for quantitative comparison. The effect of ground proximity on span-wise pressure distribution and induced velocity in the vicinity of the wing is also discussed in detail.

Modification of turbulent airfoil section flow using a combined control action

Both numerically and experimentally, the possibility of using a combined method to control the incompressible turbulent boundary layer on a symmetric airfoil section of 12-% relative thickness implemented via the blowing/suction of air through a finely perforated section provided on the wing surface part was examined. The study was performed at Reynolds number Rec = 0.7·106 in the range of angles of attack α = −6 ÷ 6°. The mechanism of action of this flow control method on the aerodynamic characteristics of the wing was identified. An ambiguous pattern of the effect due to blowing/suction from the viewpoint of ensuring a maximum lift, a gain in wing lift-to-drag ratio, and a reduction of wing aerodynamic drag was revealed.

Passive flow control for aerodynamic performance enhancement of airfoil with its application in Wells turbine – Under oscillating flow condition

In this work, the passive flow control method was applied to improve the performance of symmetrical airfoil section in the stall regime. In addition to the commonly used first law analysis, the present study utilized an entropy generation minimization method to examine the impact of the flow control method on the entropy generation characteristics around the turbine blade. This work is performed using a time-dependent CFD model of isolated NACA airfoil, which refers to the turbine blade, under sinusoidal flow boundary conditions, which emulates the actual operating conditions. Wells turbine is one of the most proper applications that can be applied by passive flow control method because it is subjected to early stall. Additionally, it consists of a number of blades that have a symmetrical airfoil section subject to the wave condition. It is deduced that with the use of passive flow control, torque coefficient of blade increases by more than 40% within stall regime and by more than 17% before the stall happens. A significantly delayed stall is also observed.

Numerical investigation of a hovering micro rotor in close proximity to a ceiling plane

This paper presents a numerical investigation into fluid interactions between a hovering micro rotor and a ceiling plane. The model consists of a two-bladed rotor with a symmetric NACA0009 airfoil section, an angle of attack of 12° and a blade tip Reynolds number of 30,000. The rotor–ceiling fluid interactions are modelled using a second order implicit finite volume solution to the incompressible Reynolds-averaged Navier–Stokes equations. Viscous effects are modelled using a Spalart–Allmaras turbulence model and an immersed boundary is used to model the ceiling plane. The methodology is validated by comparison against a two-dimensional free-vortex model as well as experimental results. The results show that the ceiling plane forces the flow upstream of the rotor to move predominantly parallel to the ceiling plane, resulting in an increase in the radial contraction of the wake downstream of the rotor. The increase in the radial component of fluid momentum causes the vertical component of fluid momentum across the rotor disk to decrease, causing the induced velocity across the blade span to decrease as the rotor–ceiling gap decreases. The change in induced velocity, which is shown to be inversely proportional to the rotor–ceiling gap, causes the local effective angle of attack along the blade span to increase with decreasing rotor–ceiling gap. This causes the sectional lift across the blade span to increase at a faster rate than sectional drag, driving an improvement in rotor efficiency as rotor–ceiling gap decreases. In principle, the improvement in rotor efficiency means that the power required to maintain a constant thrust decreases as rotor–ceiling gap decreases.

Numerical Study of Rotating Cylinder Effects on the Performance of a Symmetrical Airfoil Section

Numerical studies of a NACA 0024 airfoil section with a leading edge rotating cylinder are made using ANSYS CFX 14. Three different turbulence models, SST k-ω, Standard k-ω and k-e were utilized to investigate the flow field of the airfoil section. The lift, drag and lift to drag ratios are obtained for each turbulence model. The time dependent simulations gave the mean values of lift and drag coefficients in addition to the pressure distributions and the velocity profiles on the airfoil surfaces. The lift increased approximately 77% and the stall angle increased from 10 to 18 degrees with the use of a rotating cylinder. This is a proof that the rotating cylinder had a desirable effect on the performance of this airfoil section.

Numerical Investigation on a New Type of Fully Reversible Axial Fan with Two Blade Rows on Both Sides of the Rotor

The fully reversible axial flow fans,which can rotate reversibly to ventilate in the opposite direction with the same performance curves,are widely demanded in subways,tunnels,underground warehouses and many other fields. A new configuration method,which can obviously improve the efficiency of reversible axial flow fans,is proposed to meet this special requirement. The new approach is to install two rows of blade with reverse symmetrical airfoil section at stagger angle of 90° on both sides of the rotor of reversible axial fans. The same computational domain is used for reversible flow fans with and without the two-row guide vanes,and the detailed numerical studies reveal that the upstream blade row has little influence on the direction of the incoming flow while the downstream blade row working as a straightener reduces the circumferential velocity of the rotor-exit air remarkably and makes the outlet air flow in nearly axial direction in the diffuser. This leads to a greatly decreased friction loss,and the greater the rate of divergence in the diffuser is,the more obvious the effect will be. In addition,part of the swirl dynamic energy of the rotor-exit air is recovered and changed into static pressure rise. The above two mechanisms contribute to the evident increase in both total pressure rise and efficiency of the new type of reversible axial fans with additional blade rows than those without any stators. Furthermore,the installation of the blade rows on both sides of the fully reversible axial fan guarantees a symmetric configuration and consequently guarantees the fully reversible performance also.

Vortex Cell with a Cylinder Rotating inside it on the Surface of an Airfoil Section at High Reynolds Numbers

An algorithm for eliminating flow separation in the nose region of a symmetric airfoil section at nonzero angle of incidence is proposed. Below the separation-prone region of the section, to increase the critical angle of incidence, a vortex cell with a cylinder rotating inside is located. The shape of the vortex cell and the precise location of the cylinder are so chosen as to create the same pressure gradient in the mixing layer as in the external flow.

内圧により大きく変形する翼型の形状解析

A preliminary study is presented to realize an idea to optimize the shape of an airfoil section in accordance with airplane flight conditions, deforming it elastically by internal pressure control. Here is calculated by analytical method rather than by finite element method the shape of an airfoil section after deformed largely as an elastic thin shell by applying internal pressure. In the analysis are treated the two types of airfoil sections; a symmetric airfoil section and a cambered airfoil section. In the former case, the good agreement is obtained between the analytical results and the experimental results. Moreover, it is confirmed that this method can be applied to any cylindrical shells with a uniform section besides airfoil sections, if the ordinate of the initial shape is expressed in a polynomial of the abscissa.

Investigation of the experimental applicability of the Kirchhoff formulation for sound radiation from an oscillating airfoil

The Kirchhoff integral formulation, which is often applied in aeroacoustic problems involving radiation from surfaces, is quantitatively compared to experimental results for the case of an oscillating airfoil. A rigid symmetrical airfoil section was oscillated at small amplitude in a stationary medium (no airflow) to produce the sound field. This pure-tone acoustic source was noncompact in that the compressible frequency parameter kb where k is the wave number and b (22.9 cm) is the semichord, was 1.6 and 2.6, respectively, for the cases considered. Simultaneous amplitude and phase measurements were made of surface velocity and pressure distributions and the acoustic free field. The theory is applied using measured surface pressure and motion, taking into account airfoil thickness and contour. The result of the investigation was that theory overpredicted the sound pressure level by 2–5 dB, depending on direction. Differences were also noted in the sound field phase behavior.

Aeroelastic Stability and Control of an Oblique Wing: Wind Tunnel Experiments

Results are presented for wind tunnel tests of an elastic wing model to verify the theoretical predictions for the aeroelastic instability of an oblique wing. The model wing has an elliptic planform of 10 to 1 axis ratio and a symmetrical airfoil section of 7-1/2% thickness/chord ratio. The wing is of wood and as may be seen in the photographs presented, slack wires are used to limit the amplitude of unstable motions. The fuselage is mounted on bearings permitting freedom of roll, but provision is made to clamp the fuselage for some of the tests. It is found that freedom in roll increases the dynamic pressure at which aeroelastic instability first appears. With the model free in roll, the effectiveness of the ailerons in maintaining trim is not noticeably affected by passage through the speed at which the wing would become unstable if clamped.

Computation of Transonic Flow about Finite Lifting Wings

The results of Murman and Krupp (1971) are used to develop a procedure for computing a transonic flow about a finite lifting wing. A small disturbance equation describing the velocity potential of three-dimensional wings is solved in the process. The procedure is applied to a wing with a symmetric biconvex airfoil section in a zero-incidence subsonic flow, to a supersonic rectangular lifting wing with a sharp leading edge and to a subcritical nonlifting rectangular wing with a blunt leading edge.