Thin Wing Research Articles

3723 有限インピーダンスを用いた翼騒音の低減 : 効果的なインピーダンス分布(J13-4 流動励起騒音(1)(翼&エッジ形状関連),J13 流体関連の騒音と振動)

This paper describes optimal distribution of surface impedance for aerodynamic noise from a thin wing. Incidence turbulence noise generated by flow-wing interaction and its reduction by utilizing finite impedance are discussed in this paper. Green's function is introduced to formulate the dipole sound field of wing with perforated surface. After the effect of finite impedance is formulated for the wing noise generated by incident turbulence flow, numerical examples of the effect are shown finally, which prove the effectiveness of the optimization of impedance distribution for reducing the wing noise.

Self-consistent solution for a collisionless plasma slab in motion across a magnetic field

The problem of the dynamics of a plasma slab moving across a magnetic field is treated in the framework of the kinetic theory. A velocity distribution function (VDF) is found for each plasma species, electrons and protons, in terms of the constants of motion defined by the geometry of the problem. The zero- and first-order moments of the VDF are introduced into the right-hand side term of Maxwell’s equations to compute the electric and magnetic vector potentials and corresponding fields. The solutions are found numerically. We obtain a region of plasma convection—the slab proper—where the plasma moves with a uniform velocity, Vx=V0=(E×B∕B2)x. At the core margins two plasma “wings” are formed, each being the result of a pair of interpenetrated boundary layers with different transition lengths. Inside these wings, the plasma velocity is not uniform, Vx≠(E×B∕B2)x. It decreases from the maximum value obtained in the core to a minimum value in the central region of the wings where a flow reversal is found with the plasma convecting in the opposite direction to the core motion. There is also an asymmetry of the velocity gradient at the borders of the core, which results in a corresponding asymmetry in the thickness of the wings. Furthermore, it is found that the reversed plasma flow in the thinner wing is larger than that in the broader wing. For a fixed direction of the magnetic field the two plasma wings interchange position with respect to the center of the slab when the plasma bulk velocity reverses sign.

Processing of Gasdynamic Numerical Solutions in Similarity Parameters

The distinctive features of the application of similarity parameters (those of Tsien and Hayes, Sychev, Messiter, the linear theory, and hypersonic theory of the flow past slender bodies) are considered with reference to examples of the supersonic flow past sharp bodies of revolution, thin delta wings, and blunt bodies (ellipsoids of revolution and elliptical cylinders). Approaches to the construction of generalizing dependences on the basis of numerical solutions are discussed. These dependences can simplify the overview of numerous solutions and the search for optimal shapes in certain classes of bodies; they give preliminary design estimates and can also be used in complicated problems in which the gasdynamic results are only part of a more general, for example, dynamic problem.

On numerical modelling of a three-dimensional flow past a wing with external flow suction and on the effect of flow suction on trailing vortices

In this paper, we propose a numerical method of calculating a three-dimensional unsteady flow past a thin wing in the presence of flow suction on the wing surface through a point hole or a narrow slot. The proposed method is a further development of the discrete vortex frame method that allows one to model a flow past bodies and the evolution of a vortex wake behind them in the framework of an ideal incompressible fluid model.

On numerical modelling of a three-dimensional flow past a wing with external flow suction and on the effect of flow suction on trailing vortices

In this paper, we propose a numerical method of calculating a three-dimensional unsteady flow past a thin wing in the presence of flow suction on the wing surface through a point hole or a narrow slot. The proposed method is a further development of the discrete vortex frame method that allows one to model a flow past bodies and the evolution of a vortex wake behind them in the framework of an ideal incompressible fluid model.

Study on the Transonic Flows around a Thin Wing with an Aileron

Study on the Transonic Flows around a Thin Wing with an Aileron

Aerodynamics of delta wings with flaps at hypersonic speeds

AbstractControl surface effectiveness is an important parameter for any aeroplane. For a hypersonic aircraft, though the power required to operate the flaps is determined by low speed flying conditions, it is imperative to know the effect of flaps at hypersonic speeds. Hence, studies have been done on this topic by aerodynamicists for over 40 years. In spite of this, only a limited data is available in the literature on this subject.This paper discusses the experimental study of the effect of sweep on the aerodynamic characteristics of thin slab delta wings with flaps at hypersonic speeds. For the purpose of this investigation, a novel special thin six-component balance, which has a thickness of 4mm and can be housed inside wings with 8mm thickness, has been designed. The wings had a sweep of 76°, 70° and 65°, t/c of 0.053 and flaps with 12% of wing area and 12% of wing chord. Testing were done at Mach 8.2, Re number of 2.13 x 106 (based on chord), from α = –12° to 12° and flap angle of 20°, 30° and 40°. Separation lengths, measured from Schlieren pictures, clearly show that there is ‘no appreciable’ effect of sweep on them. Also, using a simple local flow field calculation, the separation has been identified to be transitional in nature. These features of separation reflect in the force data. Because of the small separation length, the flaps (inspite of their small size) were very effective in generating additional CN, CM and Cl, which increased with increase in flap angle. In general, the CN, CM and XCP were unaffected by sweep for symmetric flap deflection at positive incidences and asymmetric flap case. For symmetric flap case at negative incidences, only CN was not influenced by the sweep but CM decreased and XCP moved upstream as the sweep is decreased. The wing with lower sweep produces higher CA and lower (L/D)max for both symmetric and asymmetric flaps. The rolling moment and adverse yaw increased with decrease in sweep for asymmetric flap deflection. Newtonian theory is shown to be incapable of predicting the effect of sweep on Cl, Cn and on the incremental values of CN, CM and CA.In conclusion, it can be said that a small flap is generally adequate for hypersonic aeroplanes provided they operate at altitudes where transitional and turbulent separation can be expected to occur. This would make the flaps effective and thus enable ample control authority.

ACTIVE WING FLUTTER SUPPRESSION USING A TRAILING EDGE FLAP

The aeroservoelastic behaviour of a thin rectangular wing with a controllable trailing edge flap is investigated. A rather high aspect ratio motivates a numerical model based on linear beam theory for the structural dynamics and strip theory for the unsteady aerodynamic loads. Experimental flutter testing shows good agreement with the numerical stability analysis, and the impact of the trailing edge flap on the dynamics is verified by open-loop testing. The problem of stabilizing the wing utilizing the trailing edge flap is posed, and the design of a fixed-structure feedback controller is performed using numerical optimization. The problem of maximizing closed-loop modal damping with constraints on actuator performance is solved for a sequence of flow speeds and the obtained controller is synthesized using gain scheduling. The fairly large predicted increase in critical speed is experimentally verified with satisfactory accuracy.

Instationary aeroelastic computation of yacht sails

AbstractEffective schemes exist to calculate aerodynamic forces for thin bodies and structural dynamics of flexible membranes. The fluid dynamic of thin wings in a irrotational flow leads to the lifting surface theory. Neglecting the inertia of the membrane the structural dynamics are solved by the non‐linear (FEM).But the interaction of flexible membranes and an irrotational flow causes problems due to the different nature of the mathematical equations. On the one hand, there is a partial differential equation for the structural dynamics and on the other hand, there is a singular integral equation for the aerodynamics. The numerical discretization scheme has to fit these different types of equation.Our work introduces a new interaction scheme to couple the singular integral equation of the lifting surface theory with the non‐linear FEM of the membrane static. The fundamental examinations, showed by Schoop et al. (International Journal for Numerical Methods in Engineering 1998; 41: 217–219), are applied to realistic sail geometries and the aerodynamics is extended to instationary flow conditions. Copyright © 2001 John Wiley & Sons, Ltd.

Thin flat balances for testing thin delta wings at hypersonicspeeds

This paper concerns the design anddevelopment of two special novel thin internal strain gaugebalances for testing thin slab delta wings at hypersonicspeeds. Of the two balances, one is a thin three-componentbalance and the other is a thin six-component balance. Unlikethe conventional balances, which are circular in cross section,the balances discussed here are flat and have rectangular crosssections. Here, the design philosophy and the metallic architectureof these balances are discussed in detail. This paper is dividedinto three parts wherein the first part deals with the thinthree-component balance the second part addresses the thinsix-component balance and, in the third part, typical resultsobtained using these two balances are presented.The thin three-component balance has a thickness of 2.5 mm andcan be housed inside a wing as thin as 6 mm. Because of itsthinness, the axial output exhibited non-linear interaction from the normalforce loading. Hence, combined loading is performed. In order todeduce the second-order coefficients, a new methodhas been proposed and a comparison with the existing method showsthat this is better than the previous method. This balance has beenused to generate aerodynamic data on thin slab delta wings withand without lee-side balance housing bodies. In this way, it is demonstratedthat the lee-side bodies do interfere even when they come into theaerodynamic shadow of the wing.The thin six-component balance has a thickness of 4 mm and can be housedinside an 8 mm thick wing. Although this balance is also thin, it exhibitedexcellent linearity in all the bridges. In spite of this, it is subjected toseveral possible combinations of combined loading and the `work back' loadsare calculated using the linear coefficients. The work back loads agree verywell with the applied loads. Hence, linear coefficients have been used for thepurpose of data reduction. The effects of flap deflection on the aerodynamiccharacteristics of thin slab delta wings have been studied using this balance.All the experiments were conducted at Mach 8.2 and a Reynolds number of2.13×106.

Notes on gas–dynamic design of supersonic flying vehicles

It is shown that the lift–to–drag ratio of a thin delta wing is significantly lower than the lift–to–drag ratio of an infinitely long swept plate with an identical lift force. The effect of sweep on a finite wing may be used by excluding disturbances from the leading edge of the wing via introducing a “hardened” stream surface (wedge) and increasing the wing length. A three–shock waverider is proposed for choosing the optimal parameters. The sharp wedge may be avoided by replacing planar shock waves by a cylindrical shock wave upstream of the blunted wedge. If the leading edge of the wedge is not parallel to the rib that is a source of the expansion wave, a plate with zero wave drag, generating a lift force, may be obtained behind this rib. The system of regularly intersecting shock waves may be applied to design a forward–swept wing.

Swept thin wings in unsteady sonic flow

Swept thin wings in unsteady sonic flow

On the aerodynamic characteristics of thin slab delta wings at hypersonic speeds - part 2: effect of wing geometry

AbstractWe report our findings on the aerodynamic characteristics of thin slab delta wings at hypersonic speeds. We have tested wings with three sweeps (76°,70° and 65°) and four t/c ratios (0·04, 0·053, 0·066 and 0·10) using a special thin three-component balance. To date, the wing with tic of 0·040 is the thinnest wing ever tested without lee-side body at any hypersonic Mach number. The study was carried out at Mach number of 8·2, Reynolds number of 2·13 x 106 and incidence of -4° to 12°. In general, for any given t/c ratio, the increase in sweep decreased the CN whereas increased CM and CA. A possible explanation, for the effect of sweep, is given using the pressure data available in the literature. We also observed that for the sweeps considered, the t/c does not have any effect on CN and CL but increasing t/c increased CM and CA, reduced (L/D)max and moved the Xcp upstream. Blunt slab delta wings are shown to possess a unique property, wherein the CD increases with increase in sweep at higher sweep angles. This property also reflects in the (L/D) characteristics. Furthermore, the slab delta wings does not exhibit the classical CLα behaviour of sharp delta wings given in the literature. An empirical correlation has been developed to correlate CAo over a wide range of Mach number, t/c and sweep. Linearised theory has been shown to be useful in predicting the (L/D)max at Mach numbers as high as 20, for a wide range of t/c and sweep back angle.

On the aerodynamic characteristics of thin slab delta wings at hypersonic speeds — part 1: interference of lee side bodies

AbstractAn experimental investigation on the interference effect of lee-side balance housing bodies on the longitudinal aerodynamic characteristics of thin slab delta wings has been made using a special and novel thin three-component balance. This study has been conducted using wings with a t/c of 0.053 and sweep of 76°, 70° and 65° for three different lee-side bodies viz:Body (A), Body (B) and Body (C) of which Body (C) is the smallest. The testing was carried out at Mach number of 8·2, Reynolds number of 2·13 ˘ 106 and incidence of –4° to 12°. For the smallest body (Body (C)) tested, at the incidence at which this body first goes into the aerodynamic shadow of the wing (α = 6·5°), the difference in CN, CA, CM and L/D between the wing-body and the plain wing is around by 22%, 12%, 75% and 35% respectively. This difference decreases with increase in incidence and at the highest angle-of-attack up to which testing was done (α = 12°), the. corresponding values of CN, CA, CM and L/D are around 9%, 12%, 15% and 4% respectively. Contrary to the assumption in the literature pertinent to force measurements on thin slab delta wings, the interference of the lee-side bodies are found to be appreciable even after they come into the aerodynamic shadow of the wing.

Conceptual Design Studies of a Strut-Braced Wing Transonic Transport

Recent transonic airliner designs have generally converged upon a common cantilever low-wing configuration. It is unlikely that further large strides in performance are possible without a significant departure from the present design paradigm. One such alternative configuration is the strut-braced wing (SBW), which uses a strut for wing-bending load alleviation, allowing increased aspect ratio and reduced wing thickness to increase the lift to drag ratio. The thinner wing has less transonic wave drag, permitting the wing to unsweep for increased areas of natural laminar flow and further structural weight savings. High aerodynamic efficiency translates into smaller, quieter, less expensive engines and less pollution. A multidisciplinary design optimization (MDO) approach is essential to realize the full potential of this synergistic configuration caused by the strong interdependence of structures, aerodynamics, and propulsion. NASA defined a need for a 325-passenger transport capable of flying 7500 n miles at Mach 0.85 for a 2010 service entry date

Low Reynolds Number Aerodynamics of Low-Aspect-Ratio, Thin/Flat/Cambered-Plate Wings

The design of micro aerial vehicles requires a better understanding of the aerodynamics of small low-aspect-ratio wings. An experimental investigation has focused on measuring the lift, drag, and pitching moment about the quarter chord on a series of thin flat plates and cambered plates at chord Reynolds numbers varying between 60,000 and 200,000. Results show that the cambered plates offer better aerodynamic characteristics and performance. It also appears that the trailing-edge geometry of the wings and the turbulence intensity in the wind tunnel do not have a strong effect on the lift and drag for thin wings at low Reynolds numbers. Moreover, the results did not show the presence of any hysteresis, which is usually observed with thick airfoils/wings

A Cartesian Grid Generation Method Considering a Complicated Cell Geometry at the Body Surface

A cell-splitting method for Cartesian grid generation that has the capability of taking into account the cases of thin body and sharp edge is proposed in this paper. Such cases are frequently found when solving the flow around a very thin wing, such as that of a supersonic transport (SST). The method has also been extended to treat the problem of multiple solid regions within a cell, which is sometimes encountered at a highly curved body surface. Validation of the method proposed here is carried out on a sharp, thin double wedge in a supersonic flow, where significant improvements in accuracy are achieved at the cost of a small increase in the number of cells. Furthermore, application of the present method to a model of SST shows its effectiveness on a three-dimensional, realistic geometry. As a result of making a pseudo-planar approximation for body surface elements, the total number of body surface elements was reduced by a factor of about 3.2 in this application. Local grid refinement by relocating grid cells to a curved surface is also proposed, so that a more accurate solution is obtained with a reasonable number of cells.

The temperature dependence of the cross section for the energy pooling process Na(3P)+Na(3P) Na(4D)+Na(3S)

We report the measurements of the temperature dependence of the cross section 4D for the energy pooling process Na(3P)+Na(3P) Na(4D)+Na(3S). The latest two, as yet undisputed, results for 4D obtained by different authors at T = 597 K and T = 483 K suggest that this cross section decreases with increasing T, which contradicts the theory and other experiments on similar processes. To resolve this controversy and to examine the temperature trend of the cross section, we have measured the 4D in the temperature range 567-705 K, covering the high-temperature region that has not yet been investigated experimentally. To determine 4D we have excited sodium atoms in the quasistatic wing of the D1 line using a cw dye laser and measured the fluorescence intensity for the 4D3P3/2 transition, relative to the intensity of the optically thin quasistatic wing of the D2 line. The spatial distribution of the number density of the sodium atoms in the 3P3/2 state and the sodium ground-state number density were measured too. The method used for the determination of the cross section is advantageous since it entirely circumvents the need to account for the radiation trapping of 3P level radiation, which was substantial under experimental conditions of the ground-state densities being 1014-1016 cm-3. The measurements of the cross section 4D in the investigated temperature range have shown that it increases as ~exp(- E/kT). From the experiment we obtained E = (608±95) cm-1, which is in excellent agreement with the energy defect (613 cm-1) for the considered process, and in fair agreement with the values which follow from recent theoretical calculations.

Incompressible flow past oscillatory wings of low aspect ratio by the integral equations method

In the framework of the linearized perturbation theory, the pressure jump over the oscillating wing is the solution of a two-dimensional integral equation. Performing an asymptotic expansion with respect to the aspect ratio and keeping the leading terms, we reduce the integral equation to a one-dimensional one, obtaining a simplified method of solving the lifting surface integral equation for a class of thin wings of low aspect ratio with a straight trailing edge. The one-dimensional integral equation is solved for the delta flat plate and the pressure coefficient field and the lift and moment coefficients are calculated. The range of validity of the new method is discussed in the final part of the paper. Copyright © 1999 John Wiley & Sons, Ltd.

Fast nonlinear stratified flow over a thin wing

Fast 2-D stratified flow over a thin wing is considered. The problem is reduced to the Fredholm integral equation on the wing, which is uniquely solvable. In this way the solution of the nonlinear problem is obtained and can be computed by standard codes.