Effect of vortex generator on the flow field over a conventional delta wing in subsonic flow condition at higher angles of attack

Flow over a conventional delta wing has been studied experimentally at a subsonic flow of 20 m/sec and the flow field developed at higher angle of attack varying from 10° to 20° has been captured. A vortex generator is mounted on the leeward surface of the delta wing and its effect on the flow field is studied. The set of wing tip vortices generated over the delta wing is captured by the oil flow visualization and the streamline over the delta wing surface captured with and without a vortex generator are compared. Based on the qualitative results, the effect of the vortex generator on the lift coefficient is anticipated. Further, force measurement is carried out to quantitatively analyze the effect of vortex generator on the lift and drag coefficient experienced by the delta wing and justify the anticipation made out of the qualitative oil flow visualization tests. In the present study, the effect of mounting of a vortex generator is found to be minimal on the lift coefficient experienced by the delta wing. However, a significant reduction in the drag coefficient with increase in angle of attack was observed by mounting a typical vortex generator.

Characteristics of open cavity flow with floor inclinations at M = 2.0

Experimental studies on open cavity flows at supersonic speed of M = 2.0 were carried out. Oil flow visualization tests were made to understand the steady features of the surface flow field. Unsteady pressure measurements were done at five locations inside the cavity and pressure spectrum of these measurements were obtained. Cavity floor was made inclined to influence the flow directing towards the cavity leading edge with both, a favourable and adverse slope, by giving a positive and negative inclination angles to the floor, respectively. It is observed that the negative inclinations to the cavity floor behaves in a similar way to the base cavity, but a positive inclination helps to reduce the fluctuating pressures by 80% and reduce OASPL to the order of 14 dB and more.

Visualization and measurement of supersonic jet flow field evolution issuing from a rectangular nozzle with aft-deck

An experimental study is reported of supersonic jet surface flow structure visualization and wall shear stress field measurement issuing from a rectangular nozzle with aft-deck. The near-field surface flow structures evolution from over-expansion to under-expansion with the increase of nozzle pressure ratio (NPR) are successfully captured by surface oil flow visualization and shear sensitive liquid crystal coating (SSLCC) technique. The quantitative measurement result of shear stress vector field obtained by SSLCC shows that shear stress directions change significantly across the shock wave and expansion fans, while the magnitudes of shear stress have no obvious changes. Surface streamlines calculated by SSLCC image keep great consistency with the streamlines visualized using oil flow technique, which demonstrates the accuracy and potential application of SSLCC in supersonic jet surface flow visualization.

STUDI EKSPERIMEN OIL FLOW VISUALIZATION PADA AIRFOIL NACA 0012 DENGAN TRAPEZOIDAL VORTEX GENERATOR MENGGUNAKAN OPEN CIRCUIT SUBSONIC WIND TUNNEL

Pada pembuatan suatu pesawat terbang, suatu analisis sebelum terbang terhadap kinerja aerodinamika dari pesawat tersebut sangat diperlukan, terutama untuk daerah dimana metoda analitik/empirik tidak dapat menjangkaunya seperti perkiraan CLmax, karakteristik stall dan lain-lainnya. Satu hal penting yang harus diperhatikan dalam pendesainan suatu pesawat yaitu pemilihan airfoil dan modifikasinya. Modifikasi airfoil dilakukan untuk menunda separasi aliran dan meningkatkan performa airfoil, salah satunya dengan vortex generator. Hal ini dapat diindikasikan dengan tertundanya separasi aliran yang melintasi permukaan atas dari airfoil. Topik yang dikaji dalam penelitian ini adalah aliran melintasi airfoil NACA 0012 dengan penambahan vortex generator. Tujuan penelitian ini adalah untuk membandingkan karakteristik aliran fluida dengan dan tanpa penambahan vortex generator. Profil vortex generator yang digunakan adalah trapezoidal vortex generator dengan konfigurasi straight dan ditempatkan pada x/c = 20% arah chord line dari leading edge. Variasi yang digunakan adalah bilangan Reynolds (Re) dan sudut serang (α) pada airfoil. Kecepatan freestream yang digunakan yaitu kecepatan 10 m/s dan 20 m/s, pada sudut serang (α) 0°, 4°, 10°, 12°, 15°,dan 17°. Dari penelitian ini didapatkan performa aerodinamika dan fenomena aliran di sekitar airfoil. Perihal ini ada peningkatan performa aerodinamika pesawat dari sudut serang 0° sampai sudut serang 12° terbukti dengan meningkatnya kecepatan transisi dari laminar boundary layer menjadi turbulent boundary layer.

Experimental investigation of plasma vortex generator in flow control

Purpose The purpose of this paper is to compare the effects of two different configurations of plasma streamwise vortex generators (PSVG), including comb-type and mesh-type in controlling flow. This is demonstrated on the NACA 0012 airfoil. Design/methodology/approach The investigations have been done experimentally at the various electric and aerodynamic conditions. The surface oil flow visualization method has been used to the better understanding of the flow physics and the interaction of the oncoming flow passing over the airfoil and the vortex generated by comb-type PSVG. Findings This paper demonstrates the potential capabilities of the mesh-type and comb-type PSVGs in controlling flow in unsteady operation. It was found that the vortex generated by the mesh-type PSVG in unsteady operation was an order of magnitude stronger than comb-type PSVG. The flow visualisation technic proved that only a part of the plasma actuator is effective in the condition that the actuator is installed only on a portion of the upper surface of the airfoil. Originality/value This paper experimentally confirms the capabilities of the mesh-type PSVG unsteady operation in compare with comb-type PSVG in controlling flow, whereby recommends using mesh-type PSVG in the leading edge in front of comb-type PSVG on the entire wingspan to prevent the stall.

On the origin and propagation of perturbations that cause shock train inherent unsteadiness

In constant area back pressured ducts, shock trains exhibit inherent unsteadiness where the shock system fluctuates about its time-averaged position despite constant bulk inflow and outflow conditions. In this work, the underlying causes of inherent unsteadiness are identified and the flow dynamics of the system is studied for a shock train in a Mach 2.0 ducted flow that is mechanically back pressured. High-speed schlieren movies and pressure measurements are collected to quantify the shock system fluctuations. Cross-spectral analysis of this data is used to identify specific perturbations, i.e. the fluid phenomena that impact the shock train motion. Key information about each perturbation is also obtained, including where it originates, what direction it travels and how it impacts each shock. Oil flow visualization and particle image velocimetry are then used to gain insight into the physical structure of perturbations and the flow phenomena that generate them. The results identify a complex, frequency-dependent dynamical system that is influenced by (i) upstream propagating acoustic waves that emanate from separation bubbles, (ii) vortices that shed from separation bubbles and convect downstream and (iii) upstream propagating acoustic waves generated in the diffuser. With this information, a scaling argument for the shock train inherent unsteadiness is presented.

Influence of multi-wall separation control on normal-shock-induced separation in supersonic duct flows

In this work, separation due to the interaction of normal shock with wall boundary layers in a supersonic duct flow is studied using schlieren, oil flow visualization, and pressure measurements. This study uses separation control devices on three walls of a rectangular duct and investigates the influence of adding control on one wall and on the other walls. It was found that control on any wall has effect on separation on other walls. This effect was found to be due to the change in the size of corner separations. Pressure recovery was found to be steeper with the introduction of more control devices in the duct.

Flow Characteristics Over a Missile at Higher Angle of Attacks using Experiments and Simulation

The need to determine forces and moments acting on bodies of revolution at high angles of attack originally arose in connection with airships. In general missile aerodynamics is dominated with vertical and separated flows. The advent of highly manoeuvrable missile design concepts has required a major effort in understanding the problem of slender bodies of revolution at high angles of attack. An extensive experimental investigation was conducted to determine the aerodynamic forces and moments. The aerodynamic data are needed to establish the structural and control system requirements. Tests were conducted on a smooth missile model with several interchangeable nose parts in a wind tunnel at subsonic level at high angle of attack. The yaw moments and side forces are determined using force measuring systems by which maximum induced side force was found. Using oil flow visualization technique, vortex shedding location and vortex strength were traced at various angles of attack. Computational analysis was carried out for the cone, elliptical and ogive nose shapes. It is then compared with the experimental results to know where the least flow separation occurs and proper reasons for the side forces.

A Low Speed Compressor Test Rig for Flutter Investigations

A test facility for aereolastic investigations has been installed at the chair of Aero Engines at the Technische Universität Berlin. The test rig provides data for tool and code validation and is used for basic aeroelastic experiments. It is a low speed wind tunnel which allows free and controlled flutter testing. The test section contains a linear cascade with eleven compressor blades. Nine of them are elastically suspended. The paper presents a detailed description of the test facility, results to evaluate the overall flow quality and an aeroelastic model to predict the flutter velocity and critical interblade phase angles. Hot-wire anemometry has been applied to examine the inlet flow for several Mach- and Reynolds numbers. The results show small turbulence intensities. The blade surface pressure distribution and the flow field of the blade’s suction and pressure side has been accessed by oil flow visualization.