Moving Wall Effect Research Articles

Moving wall effect on normal shock wave–turbulent boundary layer interaction on an airfoil

PurposeThis paper aims to describe a proposal for an innovative method of normal shock wave–turbulent boundary layer interaction (SBLI) and shock-induced separation control.Design/methodology/approachThe concept is based on the introduction of a tangentially moving wall upstream of the shock wave and in the interaction region. The SBLI control mechanism may be implemented as a closed belt floating on an air cushion, sliding over two cylinders and forming the outer skin of the suction side of the airfoil. The presented exploratory numerical study is conducted with SPARC solver (steady 2D RANS). The effect of the moving wall is presented for the NACA 0012 airfoil operating in transonic conditions.FindingsTo assess the accuracy of obtained solutions, validation of the computational model is demonstrated against the experimental data of Harris, Ladson & Hill and Mineck & Hartwich (NASA Langley). The comparison is conducted not only for the reference (impermeable) but also for the perforated (permeable) surface NACA 0012 airfoils. Subsequent numerical analysis of SBLI control by moving wall confirms that for the selected velocity ratios, the method is able to improve the shock-upstream boundary layer and counteract flow separation, significantly increasing the airfoil aerodynamic performance.Originality/valueThe moving wall concept as a means of normal shock wave–turbulent boundary layer interaction and shock-induced separation control has been investigated in detail for the first time. The study quantified the necessary operational requirements of such a system and practicable aerodynamic efficiency gains and simultaneously revealed the considerable potential of this promising idea, stimulating a new direction for future investigations regarding SBLI control.

Comment on "Moving-Wall Effect on Unsteady Boundary Layer"

I NRef. 1 an elegantanalysis is presentedwhich demonstratestheoretically that the moving-wall effect in uences the separation of the boundary layer on a stalling airfoil in the manner described in the discussion of Fig. 14 in Ref. 2. That is, the moving-wall effect will amplify the accelerated-ow effect on a pitching airfoil but counteract it on an airfoil describing a plunging motion. It would be of signiŽ cant interest if the authors were to extend their theoretical analysis to the one-degree-of-freedom three-dimensional  ow on a coning body of revolutionillustratedby Fig. 6 in Ref. 4. As has been demonstrated analytically, the self-induced coning of a body of revolution6 is generatedby a  ow mechanism fundamentally the same as that giving rise to dynamic lift overshoot on an oscillating airfoil. This is important in view of the fundamental role that this  ow phenomenon plays on a combat aircraft maneuvering at high angles of attack.7;8

Universality of the Moving-Wall Effect

An analysis of published experimental results on the unsteady aerodynamics of separated e ow has established that the importance of the moving-wall effect is not limited to two-dimensional dynamic airfoil stall. It plays an equally prominent role in the case of three-dimensional unsteady separated e ow on a maneuvering combat aircraft. The moving-wall effect can remain effective in the presence of e ow unsteadiness. However, when the e ow unsteadiness causes bifurcation of the time-averaged e ow, the moving-wall effect can be suppressed.

Effect of Deep-Stall Wing Dynamics on Forebody-Induced Wing Rock

Introduction A RECENT review 1 of wing rock of advanced aircraft, such as X-29A, X-31, and F-18 HARV, showed that wing rock occurs Ž rst at moderate angles of attack through dynamic stall of outer wing panels with moderately swept leading edges. This is illustrated by the experimental results for the X-29A aircraft (Fig. 1). When this type of wing rock was suppressed by the use of the  aperons, forebody-induced wing rock occurred at higher angles of attack, a > 35 deg (Fig. 1b). The driving  ow mechanism for the wing rock at these high angles of attack is the moving wall effect on the forebody cross ow separation at a > uA, where uA is the forebody apex half-angle. The wing rock reaches its maximum amplitude at an angle of attack just below that for which static asymmetric cross ow separation occurs, i.e., at a 2uA, the moving wall effect has to overcome the static cross ow asymmetry. This is the reason for the rapid decrease of the limit cycle amplitude for a > 55 deg in Fig. 1b. In the analysis in Ref. 1 of the experimentally observed wing rock of a generic aircraft model (Fig. 2), a certain roll-damping value had to be assumed to obtain a limit-cycle oscillation. In the wind-tunnel test bearing friction could have supplied the needed roll damping. However, in free  ight, the roll damping has to be generated by other means to obtain an oscillation of the limit-cycle type. As discussed in Ref. 1, the only source of this roll damping at high angles of attack is the deep-stall damping-in-plunge of the outboard wing sections. The wing on the generic aircraft model (Fig. 2) has a  atplate airfoil section. It and the inverted 7.4% Clark Y airfoil have very similar deep-stall cl(a) characteristics 9 (Fig. 3). Judging by the deep-stall characteristics for the NACA 0012 and NACA 0015 airfoil sections (Fig. 4), the mean cl(a) slope at a > 8 deg in Fig. 3 should give a conservative estimate for the  at-plate airfoil section, i.e., cla ’ 0.95. As the tangential force is negligibly small, cna ’ 0.95/cos a. For the plunging wing section during wing rock, the sectional normal force at the spanwise location y = hb/2, where b is the wing span, is

Moving Wall Effects Causing Fluidelastic Instability

In a marine application a two-cylinder configuration was used that created a bistable jet in the gap between the two cylinders. A literature search was performed to gain an understanding of the unsteady fluid dynamics leading to the observed violent structural response to this flow phenomenon. As no satisfactory explanation could be found in the available literature, a study was performed. It showed that such fluidelastic instability couls result from the coupling between flow separation and the motion of the tandem-cylinder configuration generated by moving-wall effects.

Flat spin of a circular cylinder

Experimental results are analyzed to develop an understanding of the flow physics causing flat spin of circular cylinders in the critical Reynolds number region. The analysis shows that the flow phenomenon playing the key role in producing flat spin is a three-dimensional moving wall effect, which becomes especially powerful in the critical flow region by controlling flow separation through its effect on boundary-layer transition. Three-dimensional flow phenomena appear to establish the spanwise venting needed to sustain the asymmetric flow separation. It is also found that three-dimensional nonuniform roughness as well as body microasymmetry play important roles in initiating the flat-spin motion.

Dynamic airfoil flow separation and reattachment

An analysis has been performed of existing theoretical and experimental results to determine the flow physics responsible for the observed delay of flow reattachment to a negative flow incidence in a modestly rapid pitchdown motion of an airfoil. It is found that the responsible flow physics are essentially the same as those causing dynamic overshoot of static lift maximum in pitch-up motions, i.e., the accelerated flow and moving wall effects. The analysis shows that the dominant role played by the moving wall effect during pitch-down motions is in agreement with existing experimental results for a rotating circular cylinder. Applying the moving wall effect, measured through the Magnus lift on a rotating circular cylinder, to an airfoil describing pitch-down motions produces predictions that are in good agreement with experimental results. A thorough understanding of this moving wall effect is required for prediction of the high-alpha unsteady aerodynamics of advanced aircraft.

Moving wall effect in relation to other dynamic stall flow mechanisms

An analysis is performed to determine the importance of the so-called moving wall effect relative to other unsteady flow mechanisms present in the dynamic stall process. Analysis of existing theoretical and experimental results indicates that the tangential moving wall effect on the initial boundary-layer development, close to the flow stagnation point, is not satisfactorily represented in current numerical methods. The analysis shows that this moving wall effect plays a dominant role in inducing self-excited oscillations, such as the wing rock of advanced aircraft, and also plays a significant role in so-called supermaneuvers of these aircraft.

Experimental Study on the Aerodynamic Performance of a Savonius Rotor

The aerodynamic performance of a Savonius rotor has been studied by measuring the pressure distributions on the blade surfaces at various rotor angles and tip-speed ratios. It is found that the pressure distributions on the rotating rotor differ remarkably from those on the still rotor especially on the convex side of the advancing blade, where a low pressure region is formed by the moving wall effect of the blade. The torque and power performances, evaluated by integrating the pressure, are in close agreement with those by the direct torque measurement. The drag and side force performance is also studied.

Fickle effect of nose microasymmetry on the high-alpha aerodynamics

Uose ?~~croasymmetry has been shown to determine the direction and magnitude of the large sideforce generated by a slender forebody at high angles of attack. It has also been shown that this force can cause the yawing moment characteristics of an advanced aircraft to vary between several highly nonlinear types, requiring extensive testing to define. The present paper shows that such extensive tests would he of little help, and that the coupling between vehicle motion and forebody flow separation, the so called moving-wall effect, will determine the characteristics in free flight.

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.

Moving wall effects in unsteady flow

L'effet de paroi mobile est present aussi bien dans les ecoulements bidimensionnels que tridimensionnels. Influence sur le decollement des couches limites laminaires et turbulentes et sur la transition des couches limites

Fluid mechanics of dynamic stall part I. Unsteady flow concepts

Advanced military aircraft 'supermaneuverability' requirements entail the sustained operation of airfoils at stalled flow conditions. The present work addresses the effects of separated flow on vehicle dynamics; an analytic method is presented which employs static experimental data to predict the separated flow effect on incompressible unsteady aerodynamics. The key parameters in the analytic relationship between steady and nonsteady aerodynamics are the time-lag before a change of flow conditions can affect the separation-induced aerodynamic loads, the accelerated flow effect, and the moving wall effect.

Is any free flight/wind tunnel equivalence concept valid for unsteady viscous flow?

An analysis is presented showing that the present practice of using a free flight/wind tunnel equivalence concept derived for steady flow will not correctly simulate the unsteady flow effects when estimating the impact of dynamic stall on the performance of axial flow compressors and helicopter rotors. Furthermore, not even a free flight/wind tunnel equivalence concept derived for unsteady inviscid flow will correctly simulate the dynamic stall characteristics because of the unsteady viscous boundary condition at the airfoil surface, the so-called moving wall effect. c / k Ka

Aerodynamic characteristics of noncircular bodies in flat spin and coning motions

Experimental results for bodies of square cross-section coning at angles up to 90 deg show that the measured side force characteristics are extremely nonlinear, exhibiting both discontinuities and hysteresis effects. The present paper analyzes these results to determine to what extent the nonsteady wall boundary condition has influenced the flow separation and associated lateral characteristics of the coning body. It is shown that the socalled moving wall effects have a dominant influence on the flow separation and can explain the unusual side force characteristics measured in the experiments. Nomenclature c = cross-sectional chord (diameter for a circular cross section) h = cross-sectional maximum height L1 = sectional lift; coefficient, cl=LI/(pw U2^ /2)c M = freestream Mach number TV = coning rate N1 = sectional normal force coefficient, cn=N1/(PooUl/2)c p = roll rate P = static pressure; coefficient, Cp = (p—p(X)/(p(X>t/i/2) r = corner radius Re = Reynolds number, ^U^c/v^ U = velocity

Dynamics of forebody flow separation and associated vortices

It is well established that there is a strong coupling between body motion and boundary layer separation with attendant vortex shedding. In the present paper this coupling is studied for the particular case of a missile or an aircraft fuselage at very high angles of attack. It is shown that the unusual results obtained in recent tests can be explained by considering the so-called moving-wall effect on boundary layer transition and/or separation.

Unsteady flow concepts for dynamic stall analysis

It is well established that there is a strong coupling between airfoil motion and boundary layer separation with attendant vortex shedding Unsteady flow mechanisms that influence this dynamic stall event have been described previously Until now sufficient information has not been available to determine the relative im portance of various unsteady flow effects, such as the time varying inviscid pressure gradient and the unsteady viscous boundary condition at the wall, the moving wall effect Recent experimental results provide the needed information, revealing how the mode of oscillation for the airfoil determines which unsteady flow effect will dominate Nomenclature chord length frequency dynamic overshoot parameter, Eqs. (5 7) section lift, coefficient c, = l/(p^ Ui/2)c Mach number section pitching moment, coefficient cm = mp c / Ka / M m q -pitch rate Re = Reynolds number based on chord length, =