Ogive-cylinder Body Research Articles

Investigation of Flow Around a Slender Body at High Angles of Attack

A numerical investigation of flow around a slender body at high angles of attack is presented. Large eddy simulation of the flow around an ogive-cylinder body at high angles of attack is carried out. Asymmetric vortex flow was observed at angles of attack of α = 55◦ and 65◦ . The results showed that the phenomenon is present in the absence of artificial geometrical or flow perturbation. Contrary to the accepted notion that flow asymmetry is due to a convective instability, the development of vortex asymmetry in the absence of perturbations indicates the existence of absolute instability. An investigation of the unsteady flow field was carried out using dynamic mode decomposition. The analysis identified two distinct unsteady modes; high-frequency mode and low-frequency mode. At angle of attack 45◦ the high-frequency mode is dominant in the frontal part of the body and the low-frequency mode is dominant at the rear part. At α = 55◦ , the highfrequency mode is dominant downstream of vortex breakdown. At α = 65◦ , the spectrum shows a wide range of modes. Reconstruction of the dynamical modes shows that the low-frequency mode is associated with the unsteady wake and the high-frequency mode is associated with unsteady shear layer.

Effect of multiple rings on side force over an ogive-cylinder body at subsonic speed

Experiments and computations were performed over an ogive-cylinder body having an lift-to-drag ratio of 16 at a diameter Reynolds number of 29000. The side force on the slender body augments with increasing angles of attack for the case without a ring. This increase was mainly due to the increase in the asymmetry of the existing vortex pair in the wake of the body. Attempts were made to completely reduce the existing side force at the angle of attack ranging from 35° to 45°. Three rectangular cross-sectioned circumferential rings having a height of 3% of the local diameter were placed at axial distances of 2.5, 3.5 and 4.5 times the base diameter from the tip of the body so as to reduce the side force. The results obtained indicate that inclusion of three rings completely alleviated the side force on the slender body at the angle of attack ranging from 0° to 45°. The presence of rings was found to alter the growth of the vortices that helped in the reduction of the side force. Computations performed were in reasonable agreement with the experiments.

Simulation of flow around a slender body at high angles of attack

LES of the flow around an ogive-cylinder body at high angles of attack were carried out to investigate the possibility of the development of asymmetric wake-vortex without the introduction of artificial perturbations. The study investigated the effect of grid resolution and scheme bias on the solution. The numerical solution was found to be sensitive to the bias in the numerical scheme. The simulation was carried for angles of attack α = 30°, 40°, 50°, 55°, and 60°. The simulation at α = 30° − 40° produced symmetric wake-vortex. At α = 50°, the wake-vortex is also symmetric but with vortex separation. At α = 60°, the wake-vortex becomes asymmetric. At 60°, the wake-vortex is highly asymmetric with vortex separation and breakdown. It was concluded that asymmetric flow around slender bodies at high angles of attack can be simulated in the absence geometrical or flow perturbations.

Self-excited dynamics of an elastically restrained slender rigid body in uniform compressible laminar flow

In this research, we employ a computational methodology to investigate the self-excited angular dynamics of an elastically restrained rigid body in uniform compressible laminar flow. We investigate the onset of vortex-induced vibration of a two-dimensional slender rigid body configuration with a torsion spring at the leading edge and investigate the fluid–structure interaction of a three-dimensional ogive–cylinder body restrained by three torsion springs in the pitch, roll, and yaw directions. The methodology combines the implicit finite-difference Beam and Warming algorithm for the Navier–Stokes equations and a fourth-order Runge–Kutta solver for the rigid body equations of motion. The results for a fixed body reveal a critical angle-of-attack beyond which the two-dimensional flow becomes unsteady culminating with periodic drag and lift forces due to an asymmetric vortex pair shed at the body free edge. Numerical analysis of a two-dimensional pitch motion yields periodic limit cycles for a low Reynolds number flow. These self-excited oscillations evolve to ultrasubharmonic, quasiperiodic and non-stationary chaotic-like dynamics with increasing Reynolds number. Numerical analysis of three-dimensional coupled angular dynamics reveals a similar bifurcation structure for finite Reynolds number flow which includes quasiperiodic and non-stationary dynamics with increasing angles-of-attack.

Effect of ring size on the side force over ogive-cylinder body at subsonic speed

ABSTRACTExperiments and computations have been made to obtain the details of the flow field over a slender body at high angles of attack at a freestream velocity of 17 m/s corresponding to a Reynolds number of 2.9×104based on the base diameter. Experiments indicated that the existence of side force at higher angles of attack is mainly due to the presence of asymmetric vortices in the leeward side. A rectangular cross-section circular ring placed at an axial distance of 3.5 times the base diameter reduced the side force at all the angles of attack. Investigations were made to obtain the effect of the height of the ring at an angle-of-attack of 50° where the side force experienced is relatively large. A ring placed at a distance of 3.5 times the base diameter alters the initial vortices and hence helps in substantial reduction of the side force. Studies with rings of different heights indicate that a ring having a height of 3% of the local diameter reduced the side force at almost all the angles of attack for the present flow conditions and provided the least disturbance to the lift and drag of the body.

Fluid–Structure Interaction of an Elastically Mounted Slender Body at High Incidence

A numerical study of the fluid–structure interaction of an inclined tangent ogive-cylinder rigid body, mounted on torsion springs that can rotate in the pitching, yawing, and rolling directions, is conducted. The body is subjected to three-dimensional compressible laminar flow at a Reynolds number of 30,000 and a Mach number of 0.2. A second-order implicit finite difference scheme is employed for the flow equations, adapted to a curvilinear coordinate system, whereas the coupled structural equations, written using an Euler angle notation, are solved by a fourth-order Runge–Kutta method. A detailed investigation of two angles of attack, and , is described. For a fixed body at , the yaw moment changed smoothly with the location of the circumferential disturbance, whereas at the yaw moment was characterized by an almost square-wave variation behavior with bifurcation points at disturbance circumferential roll angles , , and . For an elastically mounted body, the response at for all disturbance locations resulted in large-amplitude periodic yawing and pitching moments. The response at resulted in large-amplitude moments, which occurred only for disturbance locations that coincided with the bifurcation points of the fixed body. The response at the bifurcation points was found to be quasi-periodic for , periodic for , and nonstationary for .

Low-frequency vortex oscillation around slender bodies at high angles-of-attack

Low-frequency unsteadiness of vortices over an ogive-cylinder body was studied using numerical simulation and Dynamic Mode Decomposition (DMD). The results revealed that the vortices over the fore-body of the slender body can exhibit vortex oscillation with a non-dimensional frequency of 0.054 at 70° angle-of-attack, but the flow pattern over the aft-body downstream is Kármán vortex shedding. The vortex oscillation induces larger sectional side-forces on the body than vortex shedding. The DMD analysis for the sectional flow fields at various sections shows that the most energetic mode corresponds to the vortex oscillation at the fore-body and the vortex shedding at the aft-body except the mode for a mean flow, and the associated frequencies are identical to the ones of the sectional side-forces obtained by Fourier transformation.

고받음각 오자이브의 비대칭 와류에 작용하는 구동기 효과 분석

The effect of the flow actuator on the asymmetric vortex structure around the ogive-cylinder body with fineness ratio of 4 flying at the speed of Mach 0.1 at angle of attack of 50 degree is studied. The ogive-cylinder model is developed with the actuator placed near the nose tip and numerically simulated using the in-house CFD code named KFLOW. The numerical simulation employs two different actuator modeling: one is the boundary condition given by blowing normal to the surface and another shearing on the surface. The numerical simulation reveals that response of the vortex structure to the actuation is dependent on the type of modeling as well as the strength and direction of the actuation.Keywords : Plasma Actuator(플라즈마 구동기), Asymmetric Vortex(비대칭 와류), Ogive(오자이브), High Angle of Attackt(고받음각), Flow Control(유동 제어)

고받음각 동체에 발생하는 측력의 실험적 재현 및 수치적 분석

Behavior of the side force generated at high angles of attack by two ogive-cylinder bodies of revolution with nose fineness ratio of 2.3 (B1) and 3.5 (B2) and the effect of a strip placed close the nose tip of each body (B1S and B2S) are analyzed through the wind tunnel test at ReD=200,000 and a=42~60 deg. The side force generated by B1 is increased by placing a strip. The side force generated by B2 is in the starboard direction and its magnitude is higher than that of the B1S. The effect of the strips with various dimensions placed on B2 is investigated. It is found that the 4-layer strip placed on the starboard reversed the direction of the side force into port direction. It is confirmed by numerical simulations that the strip promotes the flow separation and increases the average pressure on the side where it is placed and consequently produces the side force in the corresponding direction.

Computational and Experimental Investigations of Boundary Layer Tripping

Supersonic flow over a tapered body of revolution has been investigated both experimentally and numerically. The experimental study consisted of a series of wind tunnel tests on an ogive-cylinder body. Static pressure distributions on the body surfaces at several longitudinal cross sections, as well as the boundary layer profiles at various angles of attack have been measured. Further, the flow around the model was visualized using Schlieren technique. Tests with a natural development of the boundary layer and with tripping were also carried out. All tests were conducted in the trisonic wind tunnel of Qadr Research Center. Our results show that artificial boundary layer tripping has minor effect on the static surface pressure distribution (depending on its diameter and installation location), while the changes in total pressure around the body were significant. Tripping the boundary layer increased its thickness, changed its profile particularly near the body surface. Two oblique shock waves were formed in the front and behind the trip wire. In this study, using multi-block grid, the thin layer Navier-Stokes (TLNS) equations were solved around the above models. Also patched method was used near the interfaces. Good agreements were achieved when the numerical results were compared with the corresponding experimental data.

Side Force Suppression by Dimples on Ogive-Cylinder Body

Side Force Suppression by Dimples on Ogive-Cylinder Body

Flow Complexities of Slender Wing Rock

Introduction M ILITARY design trends toward the use of unmanned, Ž nless aircraft have renewed interest in the unsteady aerodynamics of slender wing rock. One important aspect of the problem of wing rock control is the potential impact of typical conŽ gurational details on delta-wing aircraft, such as the presence of a centerbody. Early experimental results for a sharp-edged 80-deg delta wing indicated that breakdownof the leading-edgevorticesplayed a role2,3 (Fig. 1). It could be shown that rather than being the cause of the wing rock problem, the breakdownphenomenon limited the growth of the limit-cycle amplitude. By the consideration of the effect of the roll-inducedsideslip on the effective leading-edge sweep of the windward (dipping) wing-half, an upper limit for the limit-cycle amplitude could be determined. The measured start of wing rock (Fig. 1a) is in good agreement with the prediction that, for zero bearing friction, loss of roll damping should occur at a 20 deg (Fig. 2). Bearing friction caused the delay to a = 25 deg of the loss of roll damping in the test of a pure delta-wingmodel (Fig. 1a). It will be shown that the earlier start of wing rock for the othermodel3 was caused by the presence of a centerbody or fuselage. As even unmannedcombat aircraftare likely to have a fuselageor centerbodyof some kind, it is important to know that the effect of a centerbodyondelta-wingaerodynamicscanbe large.8 Experimental results for a 69.33-degdelta-wing–bodyconŽ gurationdemonstrated that the centerbody promoted vortex breakdown.9 This could be explained by the body-induced camber effect, which according to experimental results for the effect of static camber,11 would have promoted breakdown, in agreement with the experimental results. Figure 3 shows conŽ gurational details of the models giving the results in Fig. 1. Whereas the Langley model behaves essentially as a pure,sharp-edgeddeltawing, theothermodel3 hasa centerbody. Becauseof its bluntness, it behavesas a cylindricalcenterbody, promoting vortex breakdown, thereby causing a reduction of the maximum wing rock amplitude. The earlier loss of roll damping, before the predictedvalue7 a 1⁄4 20 deg (Fig. 2), could also havebeen causedby the presenceof the centerbody.By limiting thewing area, the centerbody acts to increase the effective leading-edge sweep, thereby causing earlier loss of roll damping. The experimental results13 in Fig. 4 show that when a pointed ogive-cylinder centerbody, similar to the one used in Ref. 9, is moved aft to start behind the wing apex, as in the case of the extensively tested 65-degdelta-wing–body conŽ guration, insteadof promotingbreakdown, the body delayedvortex breakdownto occur 30% or more aft of the measured position for a pure 65-deg deltawing.14 It is described in Ref. 5 how this is the expected resultwhen the pointed ogive-cylinder body is moved aft, as shown in Fig. 4,

Experimental Investigation of Super- and Hypersonic Jet Interaction on Missile Configurations

An experimental investigation of jet interaction on conŽ gurations with lifting surfaces is presented. The experiments were carried out at supersonic Mach numbers of 2, 3.3, and 4.5 and at the hypersonic Mach number of 8. The test model was an ogive-cylinder body with interchangeable lifting surfaces that can be mounted at various locationsalong the body.Several body–surface combinationswere tested representing wing, tail, wing plus tail, and strake conŽ gurations. In some cases, the presence of the surfaces resulted in large force and/or moment ampliŽ cation due to the interaction of the jet-induced  owŽ eld with the planar surfaces. Large ampliŽ cation factors were obtained at zero angle of attack for some conŽ gurations even at the lowMach number of 2. In contrast, for similar tests using body of revolutionwithout additional surfaces, only small force ampliŽ cation, at best, could be obtained at small angles of attack. The force ampliŽ cationwas also shown to vary inversely with jet pressure. For low values of the jet stagnation pressure, ampliŽ cation factors in excess of 2 were obtained. The interaction moments could be used to obtain an additional, aerodynamic, lifting force, which for some conŽ gurations can further increase the control force.

Experimental study of controlled tip disturbance effect on flow asymmetry

The effect on the asymmetric mean flow observed on pointed bodies of revolution at incidence of changing the size and location of a controlled disturbance as well as changes in angle of attack and flow conditions are evaluated experimentally. Flow visualization and side-force measurements are carried out for a generic ogive-cylinder body inclined at high angle of attack in a low-speed wind tunnel. For all angles of attack tested (30°–60°), minute changes in the size or location of the controlled disturbance result in finite changes in the asymmetric flow field, even to the extent of reversing the sign of the side force or becoming almost symmetric. The process is reversible; returning the wire to an original position likewise restores the corresponding flow field and mean side force. The variation of side force with continuous variation of a perturbation’s size or location remains continuous and single valued, even in the incidence range of 50° to 60°, where ‘‘bistable’’ behavior of the asymmetric flow field is observed.

Asymmetric turbulent vortical flows over slender bodies

Time-accurate numerical solutions have been obtained of equations modeling turbulent subsonic flows over a slender ogive-cylinder body of revolution in the high-angle-of-attack regime where a large asymmetry in the mean flow has been observed experimentally. A modified algebraic eddy-viscosity turbulence model was utilized to correctly compute the effects of the asymmetric vortices on the underlying viscous layers. In order to reproduce any one of the experimentally observed asymmetric flow-fields, it was found necessary to add a small geometrical disturbance near the body apex. By determining an appropriate size of the disturbance, it was possible to obtain excellent agreement between numerical results and experimental data for angles of attack of 30 and 40 deg, Reynolds numbers of 3.0 x 10 to the 6th and = 4.0 x 10 to the 6th, and several roll angles. When the disturbance was removed, the flow field returned to its original symmetric shape. These results are similar in behavior to solutions obtained previously for laminar flows. Just as in the laminar case, results suggest that the origin of the asymmetry is a convective-type instability of an originally symmetric flow.

Numerical prediction of subsonic turbulent flows over slender bodiesat high incidence

The physical aspects governing accurate numerical simulation of turbulent flows having large regions of crossflow separation are re-examined. Time-accurate, three-dimensional fine-grid Navier-Stokes solutions were obtained for turbulent subsonic flows over a slender ogive-cylinder body of revolution at large angles of attack. These flowfields are complex and contain regions of crossflow separation and an organized leeward-side vortex structure. An algebraic eddy-viscosity turbulence model has been modified to correctly account for the effects of the vortices on the underlying viscous layers

Effect of splitter plate on unsteady flows around a body of revolution at incidence

Numerical solutions of the thin-layer approximation of the three-dimensional Navier–Stokes equations have been obtained for flows around an ogive-cylinder body with and without a splitter plate. The numerical results were compared qualitatively with experimental surface pressure, hot-wire anemometer measurements, and smoke visualization of the leeside vortex flow field at high angles of attack. Both experimental and numerical results show that the presence of a splitter plate in the leeward plane of symmetry suppresses the low-frequency, large-scale von Kármán vortex shedding from the body and leaves the two primary vortices symmetric and almost parallel to the upper surface of the body. It is suggested that the presence of the splitter plate prevents the interaction between flows on either side of the symmetry plane. As a result of the enforced symmetry (a) the antisymmetric mode of the convective instability near the apex of the body cannot be excited and therefore the vortices remain symmetric, staying low and parallel to the upper body surface; and (b) the antisymmetric mode of the absolute instability mechanism cannot be initiated, which suppresses the alternate shedding of vortices from the cylindrical portion of the body. On the other hand, high-frequency fluctuations of the shear layer (which could be the result of a local instability not subject to symmetry conditions) remain virtually unaffected by the presence of the splitter plate.

The effect of flow curvature on the aerodynamic characteristics of an ogive-cylinder body

SummaryThe pressure distributions over a 6:1 fineness ratio ogive cylinder model have been obtained over a wide pitch range in the rectilinear flow provided by the CoA 8 ft x 6 ft low speed wind tunnel and the curvilinear motion provided by the CoA Whirling Arm facility. The pressure distributions were integrated first to obtain the local normal-force loading distribution along the body and then the overall normal-force and pitching moment and hence the centre of normal-force.A comparison of the results showed that the main difference between the aerodynamic characteristics was a considerable positive increase in normal-force loading over the whole of the afterbody in curvilinear motion which varied little with the magnitude or sign of the pitch angle. Some smaller changes were also apparent in the forebody loading characteristics. These changes resulted in the body developing considerably more normal-force and nose-down pitching moment in curvilinear motion than in rectilinear flow with the resultant large rearward movement of the centre of normalforce.It was found possible to estimate the main features of the loadings, but the theoretical methods available did not predict very well the variations in loading obtained experimentally at the extreme beginning and end of the parallel afterbody.

Wing rock generated by forebody vortices

An analysis is performed of experimental results showing that at high angles of attack wing rock occurs for a tapered wing of aspect ratio 4, provided it is preceded by a slender ogive-cylinder body. In spite of the rather large wing span and associated roll damping, the wing rock is more violent than for a delta wing with highly swept leading edges. This surprising result is found to be caused by the critical flow conditions existing on the slender forebody, allowing the moving-wall effect on boundary-layer transition to generate the switching from symmetric to asymmetric separation and thus producing the vortex asymmetries needed to cause the observed wing rock.

Computation of turbulent supersonic flows around pointed bodies having crossflow separation

The numerical method developed by Schiff and Sturek (1980) on the basis of the thin-layer parabolized Navier-Stokes equations of Schiff and Steger (1980) is extended to the case of turbulent supersonic flows on pointed bodies at high angles of attack. The governing equations, the numerical scheme, and modifications to the algebraic eddy-viscosity turbulence model are described; and results for three cones and one ogive-cylinder body (obtained using grids of 50 nonuniformly spaced points in the radial direction between the body and the outer boundary) are presented graphically and compared with published experimental data. The grids employed are found to provide sufficient spatial resolution of the leeward-side vortices; when combined with the modified turbulence model, they are shown to permit accurate treatment of flows with large regions of crossflow separation.