Sharif Digital Repository

In the present study, a preconditioned Euler flow solver is developed to accurately and efficiently compute the inviscid flowfield around hovering helicopter rotor. The preconditioning method proposed by Eriksson is applied. The three-dimensional preconditioned Euler equations written in a rotating coordinate frame are solved by using a cell-centered finite volume Roe's method on unstructured meshes. High-order accuracy is achieved via the reconstruction of flow variables using the MUSCL interpolation technique. Calculations are carried out for an isolated rotor in hover for different conditions and the computed surface pressure distributions are compared with the experimental data. The... 

Conventional buffet onset methods for a 2D supercritical airfoil, SC0410, in transonic regime for various Mach number and various angles of attack have been surveyed. The existing methods give good results for high subsonic and transonic regimes, but demand a computational procedure to detect the buffet onset. One of these methods, trailing edge pressure divergence, that have been recognized inappropriate in other studies for supercritical airfoils, shows acceptable result at least for the present supercritical airfoil. A new method has been proposed by the authors for transonic regime that is based on the physical definition of the buffet onset from the surface pressure distribution... 

A numerical solution technique is developed to simulate the flowfield of hovering helicopter rotor. The present method uses a coupled prescribed wake-Euler solver to efficiently allow the vortical wake effects. The threedimensional Euler equations written in a rotating coordinate frame are solved by using a cell-centered finite volume scheme that is based on the Roe's flux-difference splitting on unstructured meshes. High-order accuracy is achieved via the reconstruction of flow variables using the MUSCL interpolation technique. Calculations are carried out for an isolated rotor in hover for two operating conditions of subsonic and transonic tip Mach numbers, Mtip = 0.44 and Mtip = 0.877 ,... 

Extensive experiments were conducted to study the effect of various parameters on the surface pressure distribution and transition point of an aerofoil section used in a wind turbine blade. In this paper details of the variation of transition point on the aforementioned aerofoil are presented. The aerofoil spanned the wind-tunnel test section and was oscillated sinusoidally in pitch about the quarter chord. The imposed variables of the experiments were free stream velocity, amplitude of motion, mean angle-of-attack, and oscillation frequency. The spatial-temporal progressions of the leading-edge transition point and the state of the unsteady boundary-layer were measured using eight... 

A series of low speed wind tunnel tests were conducted to study the unsteady aerodynamic behavior of an airfoil sinusoidally oscillating in plunge. The experiments included measuring the surface pressure distribution over a range of reduced frequencies, k = 0.03 - 0.06. In addition, steady state data were acquired and were used to furnish a baseline for further analysis and comparison. The model was oscillated with amplitude of ±15 cm and at three different mean angles of attack of 0, 10° and 18°. The unsteady aerodynamic loads were calculated from the surface pressure measurements, 64 ports, along the chord for both upper and lower surfaces. The plunging displacements were transformed into... 

Wind tunnel experiments were conducted to evaluate surface pressure distribution over a semi span swept wing with a sweep angle of 33°. The wing section has a laminar flow airfoil similar to that of the NACA 6-series. The tests were conducted at speeds ranging from 50 to 70 m/s with and without surface roughness. Surface static pressure was measured on the wing upper surface at three different chordwise rows located at the inboard, middle, and outboard stations. The differences between pressure distributions on the three sections of the wing were studied and the experimental results showed that roughness elements do not influence the pressure distribution significantly, except at the inboard... 

Extensive wind tunnel tests were performed on several wing- body-tail combinations in subsonic flow to study the effects of wing geometric parameters on the flow field over the tail. For each configuration, tail surface pressure distribution, as well as the velocity contour at a plane perpendicular to the flow direction behind the wing was measured. The results show a strong effect of wing to tail span ratio, as well as wing aspect ratio, on the flowfield downstream of the wing. For low sweep wings, as those considered here, wing and body interference effects on the tail are associated with the wing tip vortex and nose-body vortex  

Numerous experiments were conducted on the section of a 660 kw wind turbine blade in a subsonic wind tunnel. The selected airfoil was tested with a clean and distributed contamination roughness surface, with a high and low tunnel turbulence intensity. Surface contamination was simulated by applying 0.5 mm height roughness over the entire upper surface of the airfoil. The surface pressure distribution is measured under a steady and unsteady condition, at three Reynolds numbers; 0.43, 0.85, and 1.3 million, and over a range of angles of attack, AOA=7°-19°. Unsteady data were acquired by both pitch and plunge-type oscillation of the model about its quarter chord at a reduced frequency of 0.07.... 

A series of wind tunnel tests was conducted to examine how an end plate affects the pressure distributions of two wings with leading edge (LE) sweep angles of 23° and 40°. All the experiments were carried out at a midchord Reynolds number of 8×105, covering an angle of attack (AOA) range from −2° to 14°. Static pressure distribution measurements were acquired over the upper surfaces of the wings along three chordwise rows and one spanwise direction at the wing quarter-chord line. The results of the tests confirm that at a particular AOA, increasing the sweep angle causes a noticeable decrease in the upper-surface suction pressure. Furthermore, as the sweep angle increases, the development of... 

In this paper, an extensive experimental tests were performed to determine the aerodynamic characteristics of an airfoil undergoing static, dynamic pitch, dynamic plunge and dynamic combined pitch and plunge motions for various cases. This paper, however, focuses on the effects of reduced frequencies and mean angles of attack on the surface pressure distribution and on the corresponding lift of the airfoil oscillating in either pure pitch or in combined pitch–plunge motions. The angles of attack variations were set such that the model motion would be ceased lower than the static stall, near the static stall and beyond the static stall angles of attack. All tests were conducted at a constant... 

Wind turbines are of the most promising devices to cut down carbon emissions. However, some phenomena that adversely affect their performance are inevitable. The aim of the present paper is to investigate flow separation prevention exploiting leading edge (LE) single dielectric barrier discharge plasma actuator (SDBD-PA) on an airfoil belonging to a section of a locally developed wind turbine. The numerical results of the surface pressure distribution over the airfoil were compared with the experimental measurements carried out by the authors on the same blade section, and good agreement was found between numerical and experimental data for both plasma-OFF and plasma-ON cases. An in-depth... 

An extensive experimental investigation was conducted to study the effects of Dielectric Barrier Discharge (DBD), on the flow field of an airfoil at low Reynolds number. The DBD was mounted near the leading edge of a section of a wind turbine blade. It is believed that DBD can postpone the separation point on the airfoil by injecting momentum to the flow. The effects of steady actuations on the velocity profiles in the wake region have been investigated. The tests were performed at α = 4 to 36 degrees i.e. from low to deep stall angles of attack regions. Both surface pressure distribution and wake profile show remarkable improvement at high angles of attack, beyond the static stall angle of... 

Experimental results of surface pressure distribution over a thin supercritical airfoil and its wake are presented. All tests were conducted at free stream Mach numbers ranging from 0.27 to 0.85 and at di erent angles of attacks in a transonic wind tunnel. The model was equipped with static pressure ori ces connected to high-frequency pressure transducers. The present paper evaluates variations of shock wave location with both Mach number and angle of attack variation, as well as its interaction with the boundary layer, leading to the bu et phenomenon. Note that, for this thin supercritical airfoil, there exist only a few experimental results regarding surface pressure distributions,... 

Numerous experiments were conducted on an oscillating airfoil in a subsonic wind tunnel. The experiments involved measuring the surface pressure distribution when the model oscillated in two types of motion, pitch and plunge, at three different Reynolds numbers, 0.42, 0.63 and 0.84 million, and over a range of reduced frequencies, k = 0.03 0.09. The unsteady aerodynamic loads were calculated from the surface pressure measurements, 64 ports, along the chord for both upper and lower surfaces of the model. Particular emphasis was placed on the effects of different types of motion on the unsteady pressure distribution of the airfoil at pre-stall, near-stall and post-stall conditions. It was...