Skin friction fields on delta wings

The normalized skin friction fields on a 65° delta wing and a 76°/40° double-delta wing are measured by using a global luminescent oil-film skin friction meter. The detailed topological structures of skin friction fields on the wings are revealed for different angles of attack and the important features are detected such as reattachment lines, secondary separation lines, vortex bursting and vortex interaction. The comparisons with the existing flow visualization results are discussed.     

Automatic liquid crystal thermography for transient heat transfer measurements in hypersonic flow

A technique has been developed to measure surface heat transfer on windtunnel models in hypersonic flow based on the colour response of encapsulated thermochromic liquid crystals. The method supplies results of a superior spatial resolution at experimental uncertainties comparable to traditional methods. The approach is different from other liquid crystal applications in several key areas. It combines the calibration of the liquid crystal coating with the actual mesurement and therefore allows for an efficient experiment. The method is automated in most steps involved. Results are shown for the flow over an axisymmetric compression corner at Mach 5 and compared with surface thermocouple measurements.     

Computation of three-dimensional flow using the Euler equations and a multiple-grid scheme

The unsteady Euler equations are numerically solved using the finite volume one-step scheme recently developed by Ron-Ho Ni. The multiple-grid procedure of Ni is also implemented. The flows are assumed to be homo-enthalpic; the energy equation is eliminated and the static pressure is determined by the steady Bernoulli equation; a local time-step technique is used. Inflow and outflow boundaries are treated with the compatibility relations method of ONERA. The efficiency of the multiple-grid scheme is demonstrated by a two-dimensional calculation (transonic flow past the NACA 12 aerofoil) and also by a three-dimensional one (transonic lifting flow past the M6 wing). The third application presented shows the ability of the method to compute the vortical flow around a delta wing with leading-edge separation. No condition is applied at the leading-edge; the vortex sheets are captured in the same sense as shock waves. Results indicate that the Euler equations method is well suited for the prediction of flows with shock waves and contact discontinuities, the multiple-grid procedure allowing a substantial reduction of the computational time.     

A parametric study of LES on laminar-turbulent transitional flows past an airfoil

Low-Reynolds-number aerodynamic performance of small-sized air vehicles is an area of increasing interest. In this study, low-Reynolds-number flows past an SD7003 airfoil are investigated to understand important viscous features of laminar separation and transitional flow followed by the complicated behavior of the flow reattachment process. In order to satisfy the three-dimensional (3D) requirement of the code, a simple “3D wing” is constructed from a two-dimensional (2D) airfoil. A parametric study of large eddy simulation (LES) on the airfoil flows at Re = 60,000 is performed. Effects of grid resolution and sub-grid scale (SGS) models are investigated. Although 3D effects cannot be accurately captured owing to the limitation of the grid resolution in the spanwise direction, the preliminary LES calculations do reveal some important flow characteristics such as leading-edge laminar separation and vortex shedding from the primary laminar separation bubble on the low-Reynolds-number airfoil.     

Measurement of unsteady boundary layer developed on an oscillating airfoil using multiple hot-film sensors

 The spatial-temporal progressions of the leading-edge stagnation, separation and reattachment points, and the state of the unsteady boundary layer developed on the upper surface of a 6 in. chord NACA 0012 airfoil model, oscillated sinusoidally within and beyond the static-stall angle, were measured using 140 closely-spaced, multiple hot-film sensors (MHFS). The MHFS measurements show that (i) the laminar separation point and transition were delayed with increasing α and the reattachment and relaminarization were promoted with decreasing α, relative to the static case, (ii) the pitchup motion helped to keep the boundary layer attached to higher angles of attack over that could be obtained statically, (iii) the dynamic stall process was initiated by the turbulent flow separation in the leading-edge region as well as by the onset of flow reversal in the trailing-edge region, and (iv) the dynamic stall process was found not to originate with the bursting of a laminar separation bubble, but with a breakdown of the turbulent boundary layer. The MHFS measurements also show that the flow unsteadiness caused by airfoil motion as well as by the flow disturbances can be detected simultaneously and nonintrusively. The MHFS characterizations of the unsteady boundary layers are useful in the study of unsteady separated flowfields generated by rapidly maneuvering aircraft, helicopter rotor blades, and wing energy machines. Received: 17 June 1997 / Accepted: 10 December 1997     

Visual study of the three-dimensional flow diagram around a delta wing in a subsonic stream

Experimental methods, particularly visualization methods, permit a sufficiently detailed representation of the flow around bodies of complex shape, whose analysis meets with a considerable number of difficulties. The flow around a delta wing in the 1–90-m/sec free-stream velocity range is studied in this paper by using three-dimensional visual methods. Since stream separation and vortex-system formation play the main role in the flow formation over a wing surface, the main purpose of the experiment was to trace the physical process of dynamic development of the flow resulting in separation and vortex formation.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 2, pp. 190–194, March–April, 1976.     

Investigation of Advanced Turbulence Models for the Flow in a Generic Wing-Body Junction

The predictive performance of several turbulence models, among them formulations based on non-linear stress-strain relationships and on stress-transport equations, is examined in a collaborative university-industry study directed towards a generic wing-body junction. The geometry consists of a variation of the symmetric NACA 0020 aerofoil mounted on a flat plate, with the oncoming stream aligned with the aerofoil's symmetry plane. The dominant feature of this flow is a pronounced horseshoe vortex evolving in the junction region following separation ahead of the aerofoil's leading edge. This case is one of 6 forming a broad programme of turbulence-model validation by UMIST, Loughborough University, BAE Systems, Aircraft Research Association, Rolls-Royce plc and DERA. Key aspects of this collaboration were a high level of interaction between the partners, the use of common grids and boundary conditions, and numerical verifications aimed at maximizing confidence in the validity of the computational solutions. In total, 12 turbulence models were studied by four partners. Model performance is judged by comparing solutions with experimental data for pressure fields on the plane wall and around the aerofoil; for velocity, turbulence energy, shear stress and streamwise normal stress in the upstream symmetry plane; and for velocity, turbulence energy and shear stress in cross-flow planes downstream of the aerofoil leading edge. The emphasis of the study is on the structure of the horseshoe vortex and its effects on the forward flow. The main finding of the study is that, for this particular 3D flow, second-moment closure offers predictive advantages over the other models examined, especially in terms of the far-field structure of the horse-shoe vortex, although no model achieves close agreement with the experimental data in respect of both mean flow and turbulence quantities. This revised version was published online in July 2006 with corrections to the Cover Date.     

Experimental Study on the Behaviour of Local Laminar Separation Bubble at a Rectangular Wing Section

This paper identifies laminar separation bubbles at the root or span-wise midsection of a rectangular wing using direct surface pressure measurements in the wind tunnel and analyses their behavior at different Reynolds numbers and angles of attack. The separation, transition, and reattachment locations are determined as functions of the angles of attack and the Reynolds number. The transition structure and turbulence characteristics in the separated shear layer are studied using laser Doppler velocimetry. Surface pressure data and simultaneously acquired velocity signals are correlated to show the pattern of growing disturbances in the shear layer. Surface oil flow visualizations clarified the wingtip and separation bubble’s interactions near the leading edge of the wing at the higher angles of attack. Turbulence statistics are also calculated from the streamwise velocity distributions, and an apparent deviation is observed for the skewness and flatness values from the normal distributions in the near-wall region. The separation bubble effect on aerodynamic coefficients of a 3D rectangular wing root section is studied and reported.

     

Application of a vortex lattice numerical model in the calculation of inviscid incompressible flow around delta wings

The flow over a flat plate delta wing at incidence and in sideslip is studied using vortex lattice models based on streamwise penelling. For the attached flow problem the effect of sideslip is simulated by modifying the standard vortex lattice model for zero sideslip by aligning the trailing vortices aft of the wing along the resultant flow direction. For the separated flow problem a non-linear vortex lattice model is developed for both zero and non-zero sideslip angles in which the shape and position of the leading edge separation vortices are calculated by an iterative procedure starting from an assumed initial shape. The theoretical values are compared with available theoretical and experimental results.     

Experimental analysis of the flow field over a novel owl based airfoil

The aerodynamics of a newly constructed wing model the geometry of which is related to the wing of a barn owl is experimentally investigated. Several barn owl wings are scanned to obtain three-dimensional surface models of natural wings. A rectangular wing model with the general geometry of the barn owl but without any owl-specific structure being the reference case for all subsequent measurements is investigated using pressure tabs, oil flow pattern technique, and particle-image velocimetry. The main flow feature of the clean wing is a transitional separation bubble on the suction side. The size of the bubble depends on the Reynolds number and the angle of attack, whereas the location is mainly influenced by the angle of attack. Next, a second model with a modified surface is considered and its influence on the flow field is analyzed. Applying a velvet onto the suction side drastically reduces the size of this separation at moderate angles of attack and higher Reynolds numbers.     

Numerical Simulation of a Passive Control of the Flow Around an Aerofoil Using a Flexible,Self Adaptive Flaplet

Self-activated feathers are used by almost all birds to adapt their wing characteristics to delay stall or to moderate its adverse effects (e.g., during landing or sudden increase in angle of attack due to gusts). Some of the feathers are believed to pop up as a consequence of flow separation and to interact with the flow and produce beneficial modifications of the unsteady vorticity field. The use of self adaptive flaplets in aircrafts, inspired by birds feathers, requires the understanding of the physical mechanisms leading to the mentioned aerodynamic benefits and the determination of the characteristics of optimal flaps including their size, positioning and ideal fabrication material. In this framework, this numerical study is divided in two parts. Firstly, in a simplified scenario, we determine the main characteristics that render a flap mounted on an aerofoil at high angle of attack able to deliver increased lift and improved aerodynamic efficiency, by varying its length, position and its natural frequency. Later on, a detailed direct numerical simulation analysis is used to understand the origin of the aerodynamic benefits introduced by the flaplet movement induced by the interaction with the flow field. The parametric study that has been carried out, reveals that an optimal flap can deliver a mean lift increase of about 20% on a NACA0020 aerofoil at an incidence of 20 ^o degrees. The results obtained from the direct numerical simulation of the flow field around the aerofoil equipped with the optimal flap at a chord Reynolds number of 2 × 10⁴ shows that the flaplet movement is mainly induced by a cyclic passage of a large recirculation bubble on the aerofoil suction side. In turns, when the flap is pushed downward, the induced plane jet displaces the trailing edge vortices further downstream, away from the wing, moderating the downforce generated by those vortices and regularising the shedding cycle that appears to be much more organised when the optimal flaplet configuration is selected.     

Interactive effects between aerofoils and time independent flow perturbations

A technique is developed to assess the interdependent effects of an aerofoil in the vicinity of, immersed in or passing through, a shear flow. The two former cases in which there is no time-dependency are analogous to an aircraft wing and slipstream combination, while the latter case may be likened to a compressor rotor blade responding to a circumferential distortion. In the model, however, no time-dependent terms are included.A case study is made of an aerofoil of known geometry passing through a combined pressure/temperature distortion of stated initial geometry. The changes both in the streamline patterns of the distortion and of the pressure distribution are indicated and it is seen that the resulting lift response of the aerofoil bears some resemblance to that measured on rotating blades in compressors with circumferential flow conditions.     

Measurement of surface shear stress vector distribution using shear-sensitive liquid crystal coatings

The global wall shear stress measurement technique using shear-sensitive liquid crystal (SSLC) is extended to wind tunnel measurements. Simple and common everyday equipment is used in the measurement; in particular a tungsten-halogen light bulb provides illumination and a saturation of SSLC coating color change with time is found. Spatial wall shear stress distributions of several typical flows are obtained using this technique, including wall-jet flow, vortex flow generated by a delta wing and junction flow behind a thin cylinder, although the magnitudes are not fully calibrated. The results demonstrate that SSLC technique can be extended to wind tunnel measurements with no complicated facilities used.     

基于雨燕翅膀的仿生三角翼气动特性计算研究

针对低雷诺数微型飞行器的气动布局,设计出类似雨燕翅膀的一组具有不同前缘钝度的中等后掠(Λ=50?)仿生三角翼.为了定量对比研究三角翼后缘收缩产生的气动效应,设计了一组具有同等后掠的普通三角翼.为了深入研究仿生三角翼布局的前缘涡演化特性以及总体气动特性,采用数值模拟方法详细地探索了低雷诺数(Re=1.58×104)流动条...     

The Effect of the Mach Number on a Turbulent Backward-Facing Step Flow

The flow around a backward-facing step in the sub-, trans- and supersonic regimes was investigated at the Trisonic Wind Tunnel Munich with particle image velocimetry and dynamic pressure measurements. These two techniques were combined to simultaneously measure and correlate the velocity fluctuations in a streamwise vertical plane with the pressure fluctuations on the reattachment surface. The results show that the dynamic loads on the reattachment surface increase from subsonic up to the transonic regime while the mean reattachment location moves downstream. As soon as the flow becomes locally supersonic aft of the backward-facing step, the mean reattachment location suddenly moves upstream while the normalized dynamic loads drastically decrease. By correlating the velocity and the dynamic pressure data, it was shown that a clear separation between outer flow and the flow close to the surface aft of the step is responsible for the drastic load reduction. Due to the large difference in pressure/density, the disturbances from the locally supersonic flow do not have an effect on the flow close to the surface. This is also reflected in the power spectral densities of the pressure fluctuations on the surface, showing that at supersonic free-stream Mach numbers a low-frequency pumping motion of the locally subsonic flow is the dominant mode, while in sub-/transonic flow Kelvin-Helmholtz instabilities and a cross-pumping motion of the shear layer dominate the dynamic loads.     

A computational study on robust prediction of transition point over NACA0012 aerofoil surfaces from laminar to turbulent flows

Flowtransition from laminar toturbulent is prerequisite todecide whereabouts to apply surface flowcontrol techniques. This appears missing in a number of works in which thecontrol effects were merelyinvestigated without getting insight into alteration of transition position. The aim of this study is to capture the correctposition of transition overNACA0012 aerofoil at different angles of attack. Firstly, an implicit, time marching, highresolution total variation diminishing (TVD) scheme was developed to solve the governingNavier—Stokes equations forcompressible fluid flows around aerofoil sections to obtain velocity profiles around the aerofoilsurfaces. Secondly, the linear instability solver based on the Orr—Sommerfeld equations and the e^N methods were developed to calculate the onset of transition over the aerofoil surfaces. Forthe low subsonic Mach number of 0.16, the accuracy of the compressible solutions was assessed bysome available experimental results of low speed incompressible flows. In allcases, transition positionswere accurately predicted which shows applicability and superiority of the present work to beextended for higher Mach number compressible flows. Here, transition prediction methodology is described and the results of this analysiswithout active flow controlor separation are presented.     

Numerical simulation of an axisymmetric separated and reattached flow over a longitudinal blunt circular cylinder

This article presents a numerical investigation of turbulent flow in an axisymmetric separated and reattached flow over a longitudinal blunt circular cylinder. The governing equations were discretized by the finite-volume method and SIMPLER method was applied to solve the equations on a staggered grid. The turbulent flow was numerically simulated using the standard k–ε, Abe–Kondoh–Nagano (AKN) and Shear Stress Transport (SST) turbulence models. The comparisons made between numerical results and experimental measurements showed that the SST model is superior to other models in the present calculation.Computations were performed for three different Reynolds numbers of 6000, 10 000 and 20 000 based on the cylinder diameter. To our knowledge, this study represents the first numerical investigation of the present flow configuration. The computational results were validated with the available experimental data of reattachment length, mean velocity distribution and wall static pressure coefficient in the turbulent blunt circular cylinder flows. Further, other characteristics of the flow, such as turbulent kinetic energy, pressure, streamlines, and the velocity vectors are discussed.The results show that the main characteristics of the turbulence flow in the separation region, such as reattachment length or velocity profiles, are nearly independent of the Reynolds number. The obtained results showed that a secondary separation bubble may appear in the main separation bubble near the leading edge. Furthermore, it was found that the turbulent kinetic energy has a large effect on the formation of the secondary bubble.     

Supersonic flow past a wedge on a flat plate. Laminar boundary layer separation and reattachment

Supersonic laminar flow past a two-dimensional “flat-plate/wedge“ configuration is numerically investigated. The pressures at the boundary layer separation and reattachment points are calculated over wide Mach and Reynolds number ranges. The minimum angles of the wedge surface inclination at which a return flow occurs are determined. The results are presented in the form of generalized Mach-number-dependences of the theoretical pressure on the wedge surface initiating boundary layer separation and the pressure at the boundary layer reattachment point.     

Aerodynamic improvement of a delta wing in combination with leading edge flaps

Recently, various studies of micro air vehicle(MAV) and unmanned air vehicle(UAV) have been reported from wide range points of view. The aim of this study is to research the aerodynamic improvement of delta wing in low Reynold's number region to develop an applicative these air vehicle. As an attractive tool in delta wing, leading edge flap(LEF) is employed to directly modify the strength and structure of vortices originating from the separation point along the leading edge. Various configurations of LEF such as drooping apex flap and upward deflected flap are used in combination to enhance the aerodynamic characteristics in the delta wing. The fluid force measurement by six component load cell and particle image velocimetry(PIV) analysis are performed as the experimental method. The relations between the aerodynamic superiority and the vortex behavior around the models are demonstrated.     

Experimental investigation to study flow and heat transfer characteristics over a NACA0018 aerofoil for different height ratios

Experiments are carried out to study flow and heat transfer characteristics over NACA0018 aerofoil when the body approaches the wall of a wind tunnel. Investigations have been done to study the effect of wall proximity due to flow separation around the body at Reynolds number 2.5 × 10⁵, different height ratios and various angles of attack. The static pressure distribution has been measured on upper and lower surfaces of the aerofoil. The results have been presented in the form of pressure coefficient, drag coefficient for different height ratios. Pressure coefficient values are decreased and then increased on the lower surface of the aerofoil and decreased on the upper surface of the aerofoil at all angles of attack. The negative pressure coefficient and drag coefficient decreases as the body approaches the upper wall of wind tunnel. The maximum value of drag coefficient has been observed at an angle of attack 30° for the aerofoil at all height ratios. The Heat transfer experiments have been carried out under constant heat flux condition. Heat transfer coefficients are determined from the measured wall temperature and ambient temperature and presented in the form of Nusselt number. The variation of local as well as average Nusselt number has been shown with non dimensional distance for different angles of attack and for various height ratios. The local as well as average Nusselt number decreases as the height ratio decreases for all non-dimensional distance and angles of attack respectively. Maximum value of average Nusselt number has been observed at an angle of attack 40°.