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.

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.

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.

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.

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.