VORTEX CONTROL BY THE SPANWISE SUCTION FLOW ON THE UPPER SURFACE OF DELTA WING

The numerical investigation has been performed to explore the feasibility of vortex control by leading edge sucking excitation on a delta wing. The results reveal that the flow on the upper surface of the delta wing changes significantly in a wide range of the angle of attack. For the vortical flow at moderate angle of attack, the secondary and tertiary vortices are weakened or suppressed, and the total lift is almost unchanged. For the stalled flow at high angle of attack, the leading edge concentrated vortex is recovered, and the lift is enhanced with increasing suction rate. For the bluff-body flow at even high angles of attack, the lift can still be improved. The concentrated vortex disappears on the upper surface, and the load increment is nearly unchanged along the chordwise direction.

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 toad ceil 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 Study on Aerodynamic Noise Radiated from Delta Wing

Aerodynamic noise has been impairing the comfort of passengers in automobiles.Studies have shown that the aerodynamic noise is generated by the separation of the flow and the generation of the longitudinal vortex at the front pillar(A-pillar)and the door mirror.To remove the effects of the door mirror and extract the longitudinal vortex from A-pillar,studies employ the delta wing model.This research also employed the model and observed relations between the generated sound from the vortex at the A-pillar and the surface pressure fluctuation of the wing.The experiment was carried out in a wind tunnel of the Japan Aerospace Exploration Agency(JAXA)wind tunnel using the delta wing model.The radiated sound was measured using a far-field microphone to characterize the sound,and microphone array to conduct sound source exploration.Distribution of surface pressure fluctuation was measured using electret condenser microphones.Results showed that the radiated sound has a characteristic of dipole sound,and broadband sound from 1 kHz is radiated from the apex of the wing.Those indicate that sound generated from the apex of the delta wing was scattered at the surface of the delta wing,which follows the Lighthill-Curle theory.Surface pressure fluctuation with high fluctuation was distributed following the cone-like shape of the longitudinal vortex.Their peaks moved to the apex with the frequency increase.Coherence between far-field sound and surface pressure fluctuation was calculated.The point which is 70 mm inward from the apex showed higher value than those at the apex.As the diameter of the longitudinal vortex grows at the downstream,it is considered that a certain vortex scale radiates the most noise.

Aerodynamic Sound Radiated from Longitudinal and Transverse Vortex Systems Generated around the Leading Edge of Delta Wings

Flow around the front pillar of an automobile is typical of a flow field with separated and reattached flow by a vortex system. It is known that the vortex system causes the greatest aerodynamic sound around a vehicle. The objective of the present study is to clarify the relationship between vortical structures and aerodynamic sound by the vortex system generated around the front pillar. The vortex system consists of the longitudinal and the transverse system. The characteristics of the longitudinal vortex system were investigated in comparison with the transverse one. Two vortex systems were reproduced by three-dimensional delta wings. The flow visualization experiment and the computational fluid dynamics (CFD) captured well the characteristics of the flow structure of the two vortex systems. These results showed that the longitudinal with the rotating axis along mean flow direction had cone-shaped configuration whereas the transverse with the rotating axis vertical to mean flow direction had elliptic one. Increasing the tip angles of the wings from 40 to 140 degrees, there first exists the longitudinal vortex system less than 110 degrees, with the transition region ranging from 110 to 120 degrees, and finally over 120 degrees the transverse appears. The characteristics of aerodynamic sound radiated from the two vortex systems were investigated in low Mach numbers, high Reynolds number turbulent flows in the lownoise wind tunnel. As a result, it was found that the aerodynamic sound radiated from both the longitudinal and the transverse vortex system was proportional to the fifth from sixth power of mean flow velocity, and that the longitudinal vortex generated the aerodynamic sound larger than the transverse.

EXPERIMENTS AND ANALYSIS OF HYSTERESIS OF VORTEX BREAKDOWN OVER PITCHING-UP DELTA WINGS

An investigation of the hysteresis of vortex breakdown over pitching-up deltawings is presented. Based on experiments, there are two main reasons which can be usedto explain the hysteresis of vortex breakdown. One is the time or phase lag due to the ini-tial suPerPOsition of linearized small disturbance, which enlarges the range of the incidenceof keeping attached flow and postpones the initial incidence of vortex breakdown. Theother is the reduction of adverse pressure gradient due to the periphery centrifugal instabil-ity after the concentrated vortex has come into being, which is the key reason of hystere-sis of vortex breakdown- The structure of vortex over a pitching up delta wing can be di-vided into three layers, which is not a significant a1teration compared with that of vortexover a static delta wing.

Piezoelectric energy harvesting from flexible delta wings

The potential for harvesting energy from a flexible delta wing using a piezoelectric bimorph is experimentally investigated.Different configurations of the proposed harvesting system were tested in a wind tunnel over a broad range of airspeeds.In addition to evaluating the level of harvested power,an analysis is performed to extract critical aspects for the relation between speed,flexibility,geometry and the potential power that can be harvested from a clamped,cantilevered flexible delta wing at low angles of attack and low speeds.This analysis provides an insight into parameters that impact energy harvesting from flexible membranes or elements.

Mechanism of Generation and Collapse of a Longitudinal Vortex System Induced around the Leading Edge of a Delta Wing

The purpose of the paper is to clarify the mechanism of generation and collapse of a longitudinal vortex system induced around the leading edge of a delta wing. CFD captured well characteristics of flow structure of the vortex system. It is found that the vortex system has a cone-shaped configuration, and both rotational velocity and vorticity have their largest values at the tip of the vortex and reduce downstream along the vortical axis. This resulted in inducing the largest negative pressure at the tip of the delta wing surface. The collapse of the vortex system was also studied. The system can still remain until the tip angle of 110 degrees. However, between 110 degrees and 120 degrees, the system becomes unstable. Over 120 degrees, the characteristics of the vortex are considered to have converted from the longitudinal vortex to the transverse one.

Aerodynamics of non-slender delta and reverse delta wings:Wing thickness,anhedral angle and cropping ratio

The effects of thickness-to-chord(t=c)ratio,anhedral angle(d),and cropping ratio from trailing-edge(Cr%)on the aerodynamics of non-slender reverse delta wings in comparison to non-slender delta wings with sweep angle of 45°were characterized in a low-speed wind tunnel using force and pressure measurements.The measurements were conducted for total of 8 different delta and reverse delta wings.Two different t/c ratios of 5.9%and 1.1%,and two different anhedral angles ofd=15°and 30°for non-cropped and cropped at Cr=30%conditions were tested.The results indicate that the reverse delta wings generate higher lift-to-drag ratio and have better longitudinal static stability characteristics compared to the delta wings.The wing thickness has favorable effect on longitudinal static stability for the reverse delta wing whereas longitudinal static stability is not influenced by wing thickness for the delta wing.For reverse delta wings,the anhe-draled wing without cropping has adverse effect on aerodynamic performance and decreases the lift-to-drag ratio.Cropping in anhedraled wing causes significant improvement in lift-to-drag ratio,shift in aerodynamic and pressure centers towards the trailing-edge,and enhancement in longitudi-nal static stability.

Experimental investigation of boundary layer transition over a delta wing at Mach number 6

An experimental study on the boundary layer transition over a delta wing was carried out at Mach number 6 in a quiet wind tunnel.The Nano-tracer-based Planar Laser Scattering(NPLS)and Temperature-Sensitive Paints(TSP)techniques were used to measure the fine flow field structure and the wall Stanton number of the delta wing.The influence of factors such as the angle of attack and the Reynolds number was studied.The following results were obtained.The boundary layer transition between the leading edge and the centerline was dominated by the crossflow instability.At the location of the initial appearance of the traveling crossflow waves,the Stanton number began to rise.The Stanton number reached a maximum when the crossflow waves were broken up to turbulence.Increasing the angle of attack increased the spanwise pressure gradient at the windward side of the delta wing,thereby increasing the crossflow instability and advancing the boundary layer transition front.However,increasing the angle of attack caused the transition front to move backward at the leeward side.In addition,the sensitivity of the boundary layer transition to the Reynolds number varied with the angle of attack and the region.

Supersonic flow over a pitching delta wing using surface pressure measurements and numerical simulations

Experimental and numerical methods were applied to investigating high subsonic and supersonic flows over a 60°swept delta wing in fixed state and pitching oscillation. Static pressure coefficient distributions over the wing leeward surface and the hysteresis loops of pressure coefficient versus angle of attack at the sensor locations were obtained by wind tunnel tests. Similar results were obtained by numerical simulations which agreed well with the experiments. Flow structure around the wing was also demonstrated by the numerical simulation. Effects of Mach number and angle of attack on pressure distribution curves in static tests were investigated. Effects of various oscillation parameters including Mach number, mean angle of attack, pitching amplitude and frequency on hysteresis loops were investigated in dynamic tests and the associated physical mechanisms were discussed. Vortex breakdown phenomenon over the wing was identified at high angles of attack using the pressure coefficient curves and hysteresis loops, and its effects on the flow features were discussed.

Numerical investigation of unsteady vortex breakdown past 80°/65° double-delta wing

An improved delayed detached eddy simulation （IDDES） method based on the k-x-SST （shear stress transport） turbulence model was applied to predict the unsteady vortex breakdown past an 80o/65o double-delta wing （DDW）, where the angles of attack （AOAs） range from 30° to 40°. Firstly, the IDDES model and the relative numerical methods were validated by simulating the massively separated flow around an NACA0021 straight wing at the AOA of 60°. The fluctuation properties of the lift and pressure coefficients were analyzed and compared with the available measurements. For the DDW case, the computations were compared with such mea-surements as the mean lift, drag, pitching moment, pressure coefficients and breakdown locations. Furthermore, the unsteady properties were investigated in detail, such as the frequencies of force and moments, pressure fluctuation on the upper surface, typical vortex breakdown patterns at three moments, and the distributions of kinetic turbulence energy at a stream wise section. Two dominated modes are observed, in which their Strouhal numbers are 1.0 at the AOAs of 30°, 32° and 34° and 0.7 at the AOAs of 36o, 38° and 40°. The breakdown vortex always moves upstream and downstream and its types change alternatively. Furthermore, the vortex can be identified as breakdown or not through the mean pressure, root mean square of pressure, or even through correlation analysis.

Experimental study of flow field distribution over a generic cranked double delta wing

The flow fields over a generic cranked double delta wing were investigated. Pressure and velocity distributions were obtained using a Pitot tube and a hot wire anemometer. Two different leading edge shapes, namely ＂sharp＂ and ＂round＂, were applied to the wing. The wing had two sweep angles of 55° and 30°. The experiments were conducted in a closed circuit wind tunnel at velocity 20 m/s and angles of attack of 5°- 20° with the step of 5°. The Reynolds number of the model was about 2 - 105 according to the root chord. A dual vortex structure was formed above the wing surface. A pressure drop occurred at the vortex core and the root mean square of the measured velocity increased at the core of the vortices, reflecting the instability of the flow in that region. The magnitude of power spectral density increased strongly in spanwise direction and had the maximum value at the vortex core. By increasing the angle of attack, the pressure drop increased and the vortices became wider; the vortices moved inboard along the wing, and away from the surface; the flow separation was initiated from the outer portion of the wing and developed to its inner part. The vortices of the wing of the sharp leading edge were stronger than those of the round one.

Supersonic flow of a Chaplygin gas past a delta wing

We consider the problem of supersonic flow of a Chaplygin gas past a delta wing with a shock or a rarefaction wave attached to the leading edges.The flow under study is described by the three-dimensional steady Euler system.In conical coordinates,this problem can be reformulated as a boundary value problem for a nonlinear equation of mixed type.The type of this equation depends fully on the solutions of the problem itself,and thus it cannot be determined in advance.We overcome the difficulty by establishing a crucial Lipschitz estimate,and finally prove the unique existence of the solution via the method of continuity.

Effects of sinusoidal leading edge on delta wing performance and mechanism

Lift and drag characteristics of delta wings with low swept angle and various sinusoidal leading edges(SLE) were investigated in a wind tunnel.Three amplitudes and three wavelengths of SLE were tested.It is revealed that,in comparison with the baseline case,when the leading-edge amplitude A?5%C(root chord length of a delta wing),the stall of the delta wing can be delayed without penalty on the maximum lift coefficient;meanwhile,the lift-to-drag ratio was kept nearly unchanged.These are beneficial to aircraft maneuverability and agility.Surface oil and hydrogen-bubble flow visualization experiments were further conducted to provide a general view of the underlying flow mechanism of SLE on delta wings.It was found that,for the flow over delta wing with SLE,vortices were generated from every crest of SLE,in contrast to the dual leading-edge vortex structure generated from the apex of the base wing.At high angle of attack,the breakdown of those vortices originating from the crests of SLE may provide additional turbulent kinetic energy to the flow,resulting in the increase of the flow reattachment region on the leeward side,therefore the stall can be delayed.

Detached-eddy simulations and analyses on new vortical flows over a 76/40° double delta wing

Numerical simulations were performed on the massively separated flows of a 76/40° double delta wing using detached-eddy simulation(DES).A new type of cross-flow vortex is suggested.A vortex was initially generated near the junction of the strake and wing,which then moved towards the wing tip at certain wavelength and speed.Analyses were made in detail on the mechanism of the generation of the cross-flow vortex,that is,the inviscid cross-flow instability which differs from that of the swept blunt wing.Cross-section topology of the cross-flow vortex is also investigated,and the wavelength of the vortex array and the characteristic frequency are given.The analyses showed that the cross-flow vortices have an influence on the pressure distribution,which can cause a 10%-20% deviation from the averaged distribution.

Numerical simulations of leading-edge vortex core axial velocity for flow over delta wings

Numerical simulations have been performed to investigate the characteristics of leading-edge vortex core axial velocity over two delta wings with leading edge swept angles Λ =50°and 76°, respectively. It is obtained that Reynolds number has the most important effect on the axial velocity of the primary leading-edge vortex core. At Reynolds numbers larger than 105, the jet-like flow of the vortex core is the most common type for both the large and the moderate swept delta wings. While if Reynolds number decreases to 103―104, the core axial velocity distributions for these two delta wings present the wake-like profile for all angles of attack considered in the present investigation.

Experimental and computational investigation of novel vertical tail buffet suppression method for high sweep delta wing

Our research aim is to investigate the buffet alleviation effect of static or vibrating bulges attached on the forebody surface of the model.Experiments and numerical simulations on a model consisting of a sharp-edged,70°-leading edge sweep delta wing and twin swept back vertical tails were conducted.Models with different bulges were tested and computed at 10 and 20 m/s of free stream velocity at angles of attack ranging from 20°–50°.Dynamic strain gauge and multichannel data acquisition and analysis system were employed for the measurement of unsteady root strain on the vertical tails.Experimental and computational results show that both static and vibrating bulges behave effectively as a novel tool to alleviate tail buffet,and the alleviation effect depends largely on the vibrating frequency.Besides,one-sided bulge can only alleviate the buffeting response for the tail of the same side,and it has no obvious alleviation effect for the opposite tail.Results of the spectral analysis reveal that there are generally three peaks of spectral density for aircrafts of this configuration,and bulges used in this paper could alleviate tail buffeting,but the total lift and drag of the whole model show no obvious deviation compared to the base model and the dominant frequency of the vibration of the tails has not shifted.

Effects of flexibility on aerodynamic performance of delta wings with different sweep angles

Force measurement and surface oil flow visualization experiments were conducted in a wind tunnel to investigate the effects of flexibility on aerodynamic performance of delta wings with different sweep angles.The experimental results indicate that the maximum lift coefficient is increased and the stall angle is delayed as the sweep angle increases for both rigid and flexible wings.It is also found that the maximum lift coefficients of the flexible wings with a sweep angle from 35° to 50° are higher than those of the rigid ones.The increment of the maximum lift coefficient in particular achieves 32.9% compared with the case without lift enhancement for the 40° flexible delta wing.Moreover,the surface oil flow visualization experiments show that the stall of the flexible wing of the moderate low sweep angle is accompanied by helical flow structure,while the vortex bursting appears on the corresponding rigid wing.

Effects of coarse riblets on air flow structures over a slender delta wing using particle image velocimetry

This research investigates the aerodynamic performance and flow characteristics of a delta wing with 65° sweep angle and with coarse axial riblets,and then compares with that of a smooth-surface delta wing.Particle Image Velocimetry(PIV)were utilized to visualize the flow over the wing at 6 cross-sections upright to the wing surface and parallel to the wing span,as well as 3 longitudinal sections on the leading edge,symmetry plane,and a plane between them at Angles of Attack(AOA)=20°and 30°and Re=1.2×10~5,2.4×10~5,and 3.6×10~5.The effects of the riblets were studied on the vortices diameter,vortex breakdown location,vortices distance from the wing surface,flow lines pattern nearby the wing,circulation distribution,and separation.The results show that the textured model has a positive effect on some of the parameters related to drag reduction and lift increase.The riblets increase the flow momentum near the wing’s upper surface except near the apex.They also increase the flow momentum behind the wing.

The Mechanism of Aerodynamic Hysteresis for Sinusoidally Oscillating Delta Wings

An unsteady model of vortex system is developed to simulate the phenomena of aerodynamic hysteresis of sinusoidally oscillating delta wings.The dynamic behavior of leading-edge separation vortices simulated by the present method is in qualitative agreement with that of flow visualization by Gad-el-Hak and Ho.The calculated lift hysteresis loops are in quantitative agreement with the force measurements in the tunnel.The aerodynamic mechanism of the hysteresis phenomena is further investigated by the present method.