Wave Drag Reduction Research Articles

Steady Aerodynamics of Miniature Trailing-Edge Devices in Transonic Flows

The steady aerodynamic effect of miniature trailing-edge devices on a supercritical airfoil was investigated in transonic flow both experimentally and numerically. The investigations were performed for the Gurney flap, the split flap, and the divergent trailing edge, and they showed a strong influence on the supersonic flow regime, influencing the lift and drag characteristics. At constant lift, a redistribution of the lift generation from the front to the rear part of the airfoil takes place. Total-drag reductions can be achieved by the reduction of wave drag. The influences of the different types of miniature trailing-edge devices are shown. The investigation of the influence of geometry parameters showed that the variation of the height of the Gurney flap and of the deflection angle of the split flap have similar effects. Perforation of the miniature trailing-edge devices showed only a small influence. The investigation of the influence of the freestream parameters indicated that increases in Mach number and Reynolds number increase the miniature trailing-edge device effect. A drag-optimal application is possible with adaptive miniature trailing-edge devices, and the same performance is achieved by an adaptive Gurney flap and an adaptive split flap. A comparison with a flexible trailing edge could identify performance advantages of the miniature trailing-edge devices.

Pathways for Communicating the Effects of Stratospheric Ozone to the Polar Vortex: Role of Zonally Asymmetric Ozone

Abstract A mechanistic model that couples quasigeostrophic dynamics, radiative transfer, ozone transport, and ozone photochemistry is used to study the effects of zonal asymmetries in ozone (ZAO) on the model’s polar vortex. The ZAO affect the vortex via two pathways. The first pathway (P1) hinges on modulation of the propagation and damping of a planetary wave by ZAO; the second pathway (P2) hinges on modulation of the wave–ozone flux convergences by ZAO. In the steady state, both P1 and P2 play important roles in modulating the zonal-mean circulation. The relative importance of wave propagation versus wave damping in P1 is diagnosed using an ozone-modified refractive index and an ozone-modified vertical energy flux. In the lower stratosphere, ZAO cause wave propagation and wave damping to oppose each other. The result is a small change in planetary wave drag but a large reduction in wave amplitude. Thus in the lower stratosphere, ZAO “precondition” the wave before it propagates into the upper stratosphere, where damping due to photochemically accelerated cooling dominates, causing a large reduction in planetary wave drag and thus a colder polar vortex. The ability of ZAO within the lower stratosphere to affect the upper stratosphere and lower mesosphere is discussed in light of secular and episodic changes in stratospheric ozone.

Wave Drag Reduction with Acting Spike Induced by Laser-Pulse Energy Depositions

Wave Drag Reduction with Acting Spike Induced by Laser-Pulse Energy Depositions

Active Transonic Aerofoil Design Optimization Using Robust Multiobjective Evolutionary Algorithms

The use of adaptive wing/aerofoil designs is being considered, as they are promising techniques in aeronautic/ aerospace since they can reduce aircraft emissions and improve aerodynamic performance of manned or unmanned aircraft. This paper investigates the robust design and optimization for one type of adaptive techniques: active flow control bump at transonic flow conditions on a natural laminar flow aerofoil. The concept of using shock control bump is to control supersonic flow on the suction/pressure side of natural laminar flow aerofoil that leads to delaying shock occurrence (weakening its strength) or boundary layer separation. Such an active flow control technique reduces total drag at transonic speeds due to reduction of wave drag. The location of boundary-layer transition can influence the position and structure of the supersonic shock on the suction/pressure side of aerofoil. The boundarylayer transition position is considered as an uncertainty design parameter in aerodynamic design due to the many factors, such as surface contamination or surface erosion. This paper studies the shock-control-bump shape design optimization using robust evolutionary algorithms with uncertainty in boundary-layer transition locations. The optimization method is based on a canonical evolution strategy and incorporates the concepts of hierarchical topology, parallel computing, and asynchronous evaluation. The use of adaptive wing/aerofoil designs is being considered, as they are promising techniques in aeronautic/ aerospace since they can reduce aircraft emissions and improve aerodynamic performance of manned or unmanned aircraft. This paper investigates the robust design and optimization for one type of adaptive techniques: active flow control bump at transonic flow conditions on a natural laminar flow aerofoil. The concept of using shock control bump is to control supersonic flow on the suction/pressure side of natural laminar flow aerofoil that leads to delaying shock occurrence (weakening its strength) or boundary-layer separation. Such an active flow control technique reduces total drag at transonic speeds due to reduction of wave drag. The location of boundary-layer transition can influence the position and structure of the supersonic shock on the suction/pressure side of aerofoil. The boundarylayer transition position is considered as an uncertainty design parameter in aerodynamic design due to the many factors, such as surface contamination or surface erosion. This paper studies the shock-control-bump shape design optimization using robust evolutionary algorithms with uncertainty in boundary-layer transition locations. The optimization method is based on a canonical evolution strategy and incorporates the concepts of hierarchical topology, parallel computing, and asynchronous evaluation. Two test cases are conducted: the first test assumes the boundary-layer transition position is at 45% of chord from the leading edge, and the second test considers robust design optimization for the shock control bump at the variability of boundary-layer transition positions. The numerical result shows that the optimization method coupled to uncertainty design techniques produces Pareto optimal shock-control-bump shapes, which have low sensitivity and high aerodynamic performance while having significant total drag reduction.

Hypersonic wave drag reduction performance of cylinders with repetitive laser energy depositions

It has been widely research that wave drag reduction on hypersonic vehicle by laser energy depositions. Using laser energy to reduce wave drag can improve vehicle performance. A second order accurate scheme based on finite-difference method and domain decomposition of structural grid is used to compute the drag performance of cylinders in a hypersonic flow of Mach number 2 at altitude of 15km with repetitive energy depositions. The effects of frequency on drag reduction are studied. The calculated results show: the recirculation zone is generated due to the interaction between bow shock over the cylinder and blast wave produced by energy deposition, and a virtual spike which is supported by an axis-symmetric recirculation, is formed in front of the cylinder. By increasing the repetitive frequency, the drag is reduced and the oscillation of the drag is decreased; however, the energy efficiency decreases by increasing the frequency.

Baroclinic vortex influence on wave drag reduction induced by pulse energy deposition

We present the results of numerical analysis of wave drag reduction by a single-pulse energy deposition in a supersonic flow field around a sphere. The wave drag for the sphere was reduced as a result of the interaction between a low-density core following the blast wave produced by the energy deposition and the bow shock developed in front of the sphere. We investigated the drag reduction mechanism in terms of the unsteady flow field induced by the interaction. The effects of deposited energy and deposition location on energy reduction were examined by parametric study. From the obtained results, we refined the parameters, utilizing the baroclinic source term that produced vorticity in the vortex equation when the gradients of density and pressure were not parallel. The baroclinic vortex driven by Richtmyer–Meshkov-like instability was strong enough to contribute to the temporary low-entropy shock formation that caused low wave drag for the supersonic object. We determined that the reduced energy had a linear dependence on the radius of the low-density core formed in the blast wave and was proportional to the square of the freestream Mach number. Such dependencies could be predicted with the assumption that the energy was consumed by the baroclinic vortex generation and advected downward without thermalization in an inviscid shock layer.

A kinematic and dynamic comparison of surface and underwater displacement in high level monofin swimming

Fin swimming performance can be divided into underwater and surface water races. World records are about 10% faster for underwater swimming vs. surface swimming, but little is known about the advantage of underwater swimming for monofin swimming. Some authors reported that the air–water interface influences the kinematics and leads to a narrow vertical amplitude of the fin. On the one hand, surface swimming is expected to affect drag parameters (cross-sectional area ( S) and active drag (AD)) when compared to underwater swimming. On the other hand, the surface swimming technique may also affect efficiency ( η F ). The aim of this study is therefore to evaluate and compare drag parameters and efficiency during underwater and surface swimming. To this end, 12 international level monofin swimmers were measured during both underwater and surface swimming. Kinematic parameters (both dimensional and non-dimensional), η F (calculated according to the Elongated-Body Theory), and AD (computed with Velocity Perturbation Method) were calculated for an underwater and a surface fin swimming trial, performed at maximal speed. As expected, results showed significantly lower velocities during surface swimming vs. underwater V 1 ,under ¯ = 2.5 m s −1 vs. V 1 ,surf ¯ = 2.36 m s −1, p < .01). Velocities during underwater and surface swimming were strongly correlated ( r = .97, p < .01). Underwater swimming was also associated with higher vertical amplitudes of the fin compared to surface swimming ( V under ¯ = 0.55 m vs. V surf ¯ ) = 0.46 m, p < .01). Length-specific amplitudes ( A under/ L b ) were in the order of 20% during underwater swimming as for undulating fish, and significantly higher than during surface swimming ( A surf/ L b = 17%, p < .01). Efficiency for surface swimming was about 6% lower than for underwater swimming ( η F ,under ¯ = 0.79 vs. η F ,surf ¯ = 0.74, p < .01). This decrease could be associated with an increase in swimming frequency for surface swimming ( f ¯ surf = 2.15 Hz vs. f ¯ under = 2.08 Hz, p < .01). Active drag during surface swimming was about 7% higher than for underwater swimming ( AD under ¯ = 78.9 N vs. AD surf ¯ = 84.7 N, p < .01). A significantly smaller cross-sectional area for surface swimming ( S under ¯ = 0.053 m 2 vs. S surf ¯ = 0.044 m 2, p < .01) and higher drag coefficient for surface swimming ( C d ,under ¯ = 0.47 vs. C d ,surf ¯ = 0.69, p < .01) were measured. Finally, correlation between cross-sectional area and vertical amplitude of the fin was reported for both underwater and surface swimming. These results suggest that the performance improvement during underwater swimming is not only linked to a wave drag reduction effect but also to a specific swimming technique due to the free surface.

超音速複葉翼形状及び翼胴形状の空力特性に関する基礎研究

In order to develop a quiet supersonic transport, it is necessary to reduce shock waves around the transport. Shock waves, in general, are the cause of the airplane's sonic boom. Authors have been studying an aerodynamic feasibility of supersonic biplanes based on the concept of the Busemann biplane. In this paper, the three dimensional effect of wing geometries on their wave drags, including wing tip effects and the interference effects between the wing and a body (Wing-Body configurations) are investigated, using CFD code in Euler (inviscid) mode. As a result, we can conclude that the supersonic biplane wings at their design Mach number (M∞=1.7) are still capable of reducing wave drag significantly similar to that of the 2-D supersonic biplane.

Shock-Wave/Boundary-Layer Interaction Control Using Three-Dimensional Bumps for Transonic Wings

Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established X-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30%, especially when significant wave drag is present. Counter-rotating streamwise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.

Three-dimensional contour bumps for transonic wing drag reduction

Two-dimensional and three-dimensional contour bumps are designed and optimized for substantial wave drag reduction for an un-swept natural laminar flow (NLF) wing (RAE5243 aerofoil section) at transonic speeds. An NLF aerofoil wing is chosen in this study, as shock control is more crucial for such wings due to the requirement of favourable pressure gradients on a substantial part of the wing. For the validation purpose and to focus on the wave drag issues, the boundary layer is assumed to be fully turbulent from the leading edge. Key bump geometrical parameters including the maximum height, the length, and the crest position have been chosen for the parameterization of the two-dimensional and three-dimensional shock control bumps. For the three-dimensional bumps, an array of the contour bumps is installed spanwise on the transonic wing and their width and spanwise spacing are chosen as additional design parameters. Both the two-dimensional and the three-dimensional bump shapes are optimized using a discrete adjoint-based optimization method. The performance of the three-dimensional contour bumps are compared in detail with the similarly optimized two-dimensional bumps both at and around the design point. The results show that, for the NLF wing studied, the optimized three-dimensional bumps are as effective as the optimized two-dimensional bump in terms of total drag reduction at the given design point, despite the significant difference in their geometrical shapes. More importantly, in terms of the operational range for varying lift conditions for practical applications, the three-dimensional bumps outperform the two-dimensional bump by a substantial margin.

Design Space Exploration of Supersonic Formation Flying Focusing on Drag Minimization

This study investigates the drag characteristics of a supersonic formation-flying concept that aims to reduce wave drag and sonic boom via shock-wave and expansion-fan interaction. Because of the complex interaction patterns seen in three-aircraft formations, optimization was applied as a rational means for design. To consider both the cruise efficiency and the safety of the aircraft, the objective functions are chosen to be the total LID of the formation and the minimum separation distance among the aircraft. The design variables define the relative positions of the aircraft and, as for the coordinate definition, the skewed cylindrical coordinate system that has been proven to be very effective in extracting the physics of the flowfield was chosen. Optimization results show a good correlation with results from previous studies. However, optimization arrived at a formation that unexpectedly achieved high performances in regions of the design space in which the previous study suggested bad performance. These solutions took advantage of the design space of three-aircraft formations (i.e., the synergistic effects of the flowfield of the two leading aircraft) and exploited the difference between the design space of two-aircraft formations.

A combined experimental and numerical study of flow structures over three-dimensional shock control bumps

A detailed study of transonic flow over three-dimensional bumps has been conducted using experimental measurements and computational simulation. The aim of the investigation is to determine the flow characteristics over these potentially effective flow control devices for wave drag reduction for transonic aircraft wings. Careful qualitative and quantitative matching of the simulation and experimental conditions has considerably improved the agreement between them. In both the experiments and the computation, the shock position in the working section was found to be sensitive not only to the back pressure and the sidewall effects, but also to the incoming boundary layer characteristics. Two turbulence models, a modified Baldwin–Lomax model with enhanced performance in separated flow (curvature model) and a two-equation model ( k – ω ) have been implemented in the numerical simulation. Overall, both turbulence models gave reasonable results for the uncontrolled and controlled cases regarding the inviscid flowfield and the shock structure. In particular, a pair of streamwise vortices embedded in the boundary layer was captured by the two models. The traces in the surface oil flow pictures from the experiment also suggested the existence of the streamwise vortices. The combined surface and flowfield data provide some further insight into the flow physics on the shock control ramp bump, which is discussed in the paper. In addition, the study also demonstrates the enhanced capability of the algebraic model in capturing separated flow features with lower computational cost as compared to the k – ω model.

Plasma mitigation of shock wave: experiments and theory

Two types of plasma spikes, generated by on-board 60 Hz periodic and pulsed dc electric discharges in front of two slightly different wind tunnel models, were used to demonstrate the non-thermal plasma techniques for shock wave mitigation. The experiments were conducted in a Mach 2.5 wind tunnel. (1) In the periodic discharge case, the results show a transformation of the shock from a well-defined attached shock into a highly curved shock structure, which has increased shock angle and also appears in diffused form. As shown in a sequence with increasing discharge intensity, the shock in front of the model moves upstream to become detached with increasing standoff distance from the model and is eliminated near the peak of the discharge. The power measurements exclude the heating effect as a possible cause of the observed shock wave modification. A theory using a cone model as the shock wave generator is presented to explain the observed plasma effect on shock wave. The analysis shows that the plasma generated in front of the model can effectively deflect the incoming flow; such a flow deflection modifies the structure of the shock wave generated by the cone model, as shown by the numerical results, from a conic shape to a curved one. The shock front moves upstream with a larger shock angle, matching well with that observed in the experiment. (2) In the pulsed dc discharge case, hollow cone-shaped plasma that envelops the physical spike of a truncated cone model is produced in the discharge; consequently, the original bow shock is modified to a conical shock, equivalent to reinstating the model into a perfect cone and to increase the body aspect ratio by 70%. A significant wave drag reduction in each discharge is inferred from the pressure measurements; at the discharge maximum, the pressure on the frontal surface of the body decreases by more than 30%, the pressure on the cone surface increases by about 5%, whereas the pressure on the cylinder surface remains unchanged. The energy saving from drag reduction is estimated to make up two-thirds of the energy consumed in the electric discharge for the plasma generation. The measurements also show that the plasma effect on the shock structure lasts much longer than the discharge period.

Role of on-board discharge in shock wave drag reduction and plasma cloaking

In the present paper, a physical model is proposed for reducing the problem of the drag reduction of an attached bow shock around the nose of a high-speed vehicle with on-board discharge, to the problem of a balance between the magnetic pressure and gas pressure of plane shock of a partially ionized gas consisting of the environmental gas around the nose of the vehicle and the on-board discharge-produced plasma. The relation between the shock strength and the discharge-induced magnetic pressure is studied by means of a set of one-fluid, hydromagnetic equations reformed for the present purpose, where the discharge-induced magnetic field consists of the electron current (produced by the discharge)-induced magnetic field and the partially ionized gas flow-induced one. A formula for the relation between the above parameters is derived. It shows that the discharge-induced magnetic pressure can minimize the shock strength, successfully explaining the two recent experimental observations on attached bow shock mitigation and elimination in a supersonic flow during on-board discharge [Phys. Plasmas 9 (2002) 721 and Phys. Plasmas 7 (2000) 1345]. In addition, the formula implies that the shock elimination leaves room for a layer of higher-density plasma rampart moving around the nose of the vehicle, being favourable to the plasma radar cloaking of the vehicle. The reason for it is expounded.

Nonlinear effects of pulsed periodic energy supply on shock wave structure of transonic flow around airfoils

The work deals with an investigation of possibilities of controlling the aerodynamic characteristics of airfoils with the aid of external local pulsed-periodic energy supply at transonic flight regimes. The alteration of flow structure near a symmetric airfoil and its wave drag has been studied on the basis of a numerical solution of two-dimensional nonstationary gas dynamics equations versus the energy supply period, the localization and shape of the energy supply zone. The energy supply upstream of the closing shock in the immediate proximity of the contour in the zones extended along it is found to result in a considerable reduction of the profile wave drag. The nature of such a drag reduction is elucidated. The existence of a limiting frequency of energy supply is established.

Normal shock wave/turbulent boundary-layer interaction control using ‘smart’ piezoelectric actuators

AbstractThis paper looks at active control of the normal shock wave/turbulent boundary layer interaction (SBLI) using smart flap actuators. The actuators are manufactured by bonding piezoelectric material to an inert substrate to control the bleed/suction rate through a plenum chamber. The cavity allows rapid thickening of the boundary-layer approaching the shock, which splits into a series of weaker shocks forming a lambda shock foot, thus reducing wave drag. Active control allows optimisation of the interaction, as it would be capable of either positioning the control region around the original shock position using a series of unimorph flaps or fixing the shock position by controlling the rate of mass transfer.The level of control achieved by unimorph piezoelectric actuators is not large because of small amounts of deflection possible. It is believed that to provide optimal control a piezoelectric material, which can provide greater strain and hence higher amounts of deflection is needed. However, currently such a piezoelectric material is not commercially available.

Evaluation of wave drag reduction by flow control

An analytical expression is proposed to estimate the wave drag of an aerofoil equipped with shock control. The analysis extends the conventional approach for a single normal shock wave in the absence of control, based on the knowledge that all types of successful shock control on transonic aerofoils cause bifurcated λ-shock structures. The influence of surface curvature on the λ-shock structure has been taken into account. The extended method has been found to produce fairly good agreement with the results obtained by CFD methods while requiring negligible computational effort. This new formulation is expected to be beneficial in the industrial design process of transonic aerofoils and wings where a large number of computational simulations have to be performed.

Drag Reduction of a Near-Sonic Airplane by Using Computational Fluid Dynamics

Aerodynamic shape optimization for an airplane cruising at a near-sonic regime is discussed based on computational fluid dynamics. Japan Aerospace Exploration Agency's scaled experimental supersonic airplane model, NEXST-1, was employed as the baseline model of the optimization. NEXST-1 was accepted as a candidate for a near-sonic airplane because of the existence of a drag bucket at a near-sonic regime. In the present optimization, the section airfoil shape and the planform shape were optimized independently of each other in the near-sonic regime of Mach number 0.98. For the optimization, a genetic algorithm was used with unstructured mesh Euler simulations. The results of the optimizations showed considerable improvement in the lift-to-drag ratio in the near-sonic regime. The optimization of the section airfoil shape and that of the planform shape yielded to the reduction of wave drag and induced drag, respectively.

Passive Shock Control Concept for Drag Reduction in Transonic Flow

The operating principle of a novel passive shock control concept for drag reduction on swept wings called D-Strips is introduced, and results of a first experimental proof of concept are presented. Wind-tunnel experiments are conducted at transonic airspeeds using an airfoil VC-Opt (9.2% relative thickness) with forced boundary-layer transition. Schlieren pictures confirm the operating principle. Wake measurements demonstrate that, locally, in the wake of D-Strips the total drag increases. However, more globally, at spanwise locations away from these positions the total drag is measurably reduced if a sufficiently strong shockwave is present above the wing-suction-side surface. In contrast to passive control by ventilation, outside of the wake of the D-Strips a wave and viscous drag reduction and observed. Furthermore, the maximum lift tends to be increased by D-Strips suggesting that the buffet-onset limit is delayed. An application of D-Strips to a swept wing at cruise conditions can yield net aircraft drag reduction because the drag rise is limited to the D-Strips and its wake, but the shock-weakening effect and the drag reduction are distributed in the wide-splayed characteristics upstream of the shock front. The effectiveness of D-Strips is less sensitive to the location of the shock wave than control concepts like two-dimensional bumps, which are only effective at the design conditions. Another application of D-Strips that can be considered is the regions with a distinct shock, for example, between engines or in the wing-root region or at the inboard engine. However, there is a need for further investigations in order to optimize D-Strips and to quantify their impact on aircraft performance. Recent experimental results indicate that airscoop-type configurations might perform better as D-Strips than Velcro-type strips. Because D-Strips are easy to apply on existing wings, an application of D-Strips for retrofit is possible.

Investigation of active control of swept shock wave/turbulent boundary-layer interactions – PSP results

AbstractAn investigation of active control of the swept shock wave/boundary-layer interaction using ‘smart’ flap actuators is presented. The actuators are manufactured by bonding piezoelectric material to an inert substrate to control the bLeed/suction rate through a plenum chamber. The cavity provides communication of signals across the shock, allowing rapid thickening of the boundary layer approaching the shock. This splits the shock foot into a series of weaker shocks forming a lambda structure, thus reducing wave drag. Active control allows optimisation of the unimorph deflection, hence rate of mass transfer.In this paper, results of the interaction using pressure sensitive paint (PSP) are emphasised. It is shown that the use of PSP, in conjunction with discrete pressure data, enables the main features of the interaction to be observed when the actuators are subject to different deflections.