Low Sweep Angle Research Articles

Characterization of Out-of-Plane Curved Fluidic Oscillators

A novel fluidic oscillator design, in which the oscillator is curved along the primary flow direction, was evaluated. The effects of mass flow rate, hydraulic diameter, aspect ratio, curvature radius, surface roughness, inlet orientation, and nonconstant curvatures were experimentally investigated. Measurements of the oscillation Strouhal number, discharge coefficient, spreading angle, and sweeping angle were used to characterize the jet. The resulting oscillators were found to fall into one of three categories: oscillated at the same frequency and sweep angle as conventional flat oscillators; oscillated at a slightly higher frequency and lower sweep angle than flat oscillators; or no dominant oscillation frequency detected and with no sweeping action. An unsteady Reynolds-averaged Navier–Stokes computational fluid dynamics simulation revealed fundamental differences in the internal flow mechanisms between flat and curved oscillators that drive the sweeping jet. The curvature between 37.5 and 62.5% of the total length, or the region from the inlet nozzle to halfway through the main chamber, was a primary factor influencing the response type of a design. Due to the curvature of these oscillators, they have the ability to be used in geometrically constrained spaces, such as the leading edge of wings and turbine vanes.

Duality of the Stream Pattern of Supersonic Viscous Gas Flow past a Blunt-Fin Junction: the Effect of a Low Sweep Angle

Duality of the Stream Pattern of Supersonic Viscous Gas Flow past a Blunt-Fin Junction: the Effect of a Low Sweep Angle

Duality of the stream pattern of supersonic viscous gas flow past a blunt-fin junction: effect of a low sweep angle

Results of numerical solution of the problem of supersonic flow past a symmetric blunt fin mounted on a plate with a developing boundary layer are presented. Generally, the case considered corresponds to the flow configuration used in the experimental and computational study by Tutty et al. (2013), in which the laminar air flow with the freestream Mach number of 6.7 is considered. Previously, the authors have shown (2020) that for these conditions two stable solutions corresponding to metastable flow states with different configurations of the vortex structure and the pattern of local heat transfer are predicted. In present paper, the influence of a low sweep angle of a blunt leading edge on the vortex structure in the separation region, local heat transfer, and the possibility of obtaining a dual solution are investigated. The bifurcation diagrams showing for two solutions the main horseshoe vortex center location and the length of separation region versus the skew angle are presented.

Balancing Laminar Extension and Wave Drag for Transonic Swept Wings

At a transonic condition, the design of a natural laminar flow (NLF) wing is challenging because the extension of the laminar flow needs to be finely balanced with the potential wave drag increase. To achieve this balance, it is proposed to unlock the wing sweep and introduce a three-dimensional (3-D) contour shock-control bump (SCB) in the optimization of the NLF infinite swept wing aiming at total drag reduction. A 3-D Reynolds-averaged Navier–Stokes flow solver is extended to incorporate laminar–turbulent transition prediction due to streamwise and crossflow instabilities. The flow solver is integrated in a gradient-based optimization framework. The transition criteria including streamwise and crossflow instabilities are coupled in the sensitivity calculation using a discrete-adjoint solver. Transonic design optimization of the sectional profile and wing sweep angle is first conducted at a Mach number of 0.78. The optimization managed to alleviate the shock wave, reducing the pressure drag while it failed to extend the laminar flow. Furthermore, a combined optimization of the wing with a parameterized SCB at the same condition is carried out. The optimized wing with the SCB features a long favorable pressure gradient region and a low-sweep angle, which allows for a large proportion of laminar flow without significant pressure drag. The shock wave is controlled by the 3-D bump, and the SCB has little effect on the upstream flow.

Shock Control of a Low-Sweep Transonic Laminar Flow Wing

This paper presents a combined experimental and computational study of a low-sweep transonic natural laminar flow (NLF) wing with shock-control bumps (SCBs). A transonic NLF wing with a relatively low sweep angle of 20 deg was chosen for this study. To avoid the complexity of the flow introduced by perforated/slotted walls commonly used for transonic wind-tunnel tests for reducing the wall interference, both experimental tests and computational simulations were conducted with solid wind-tunnel wall conditions. This allows for like-to-like validation of the computational simulation. Optimization of the shock-control bumps was first conducted to design the wind-tunnel test model with bumps. Two critical parameters of the three-dimensional SCBs for shock control (i.e., bump crest position and bump height) were optimized in terms of total drag reduction at the given design point in the wind tunnel. We show that the strong shock wave on the low-sweep NLF wing can be effective controlled by well-designed SCBs deployed along the wing span. The optimized SCBs result in 18.5% pressure drag reduction with 5% viscous drag penalty, and the SCBs also bring some benefits at off-design conditions. The wind-tunnel tests include pressure measurement, particle image velocimetry, and temperature-sensitive paint to provide detailed insight into the shock-control flowfield and to validate the computational simulations. Comparisons include surface pressure profile, velocity distribution, and transition location.

Shape Optimization and Sensitivity Analysis of a Morphing-Wing Aircraft

Morphing-wing concepts have received growing interests in the recent years for enhancing the multi-mission performance of unmanned aerial vehicle. However, to obtain a feasible morphing strategy, the optimization design studies are required. In this paper, an aerodynamic optimization study for a morphing-wing aircraft with variable sweep, span, and chord length was conducted to obtain the optimum configurations at subsonic, transonic and supersonic conditions. The optimization objective is to obtain maximum lift-to-drag ratios subject to lift coefficients and static stability constraints at each flight condition. A genetic algorithm in conjunction with surrogate models was employed to search optimum solutions in the entire design space. The aerodynamic forces are calculated by an Euler-based solver and friction drag estimation code. The optimum configuration corresponding to any flight condition can be determined through the optimization. A global sensitivity analysis based on the surrogate model was also carried out, hence the contribution of each design variable to the optimization objective can be analyzed. The results indicated that the optimum wing at subsonic speeds have a lower sweep angle and high aspect ratio, and the wing sweep has a primary contribution to lift-to-drag ratios, and the span is secondary; at transonic conditions, the medium-sweep wing is the optimum, and the contribution of the span is far more than other variables; at supersonic conditions, the optimum configuration becomes a high-sweep cropped delta wing, and the span has a dominant contribution, and the sweep is secondary.

Finite Element Analysis of Carbon Fibre Reinforced Aircraft Wing

Subsonic aircraft is used as business jets and commercial airplanes. The wings are usually straight or with low sweep angle. The main group of materials used in aircraft construction has been wood, steel, aluminum alloys and more recently, titanium alloys and fiber reinforced composites. Several factors influence the selection of material of which strength allied to lightness is the most important. Composite materials are well known for their excellent combination of high structural stiffness and low weight.CFC is seen to have a modulus twice &strength three times that of aluminum alloy, the conventional material used in aircraft construction. In the present work the aircraft wing components skin is analyzed considering both isotropic and composite materials. The parametric study conducted using MSc Nastran finite element package. From the studies conducted regarding the weight reduction, it is estimated that replacement of Al. alloy by CFC results in 52.96% saving in the total structural weight of the aircraft wing.

Shock control bump optimization for a low sweep supercritical wing

Shock control bumps are a promising technique in reducing wave drag of civil transport aircraft flying at transonic speeds. This paper investigates the optimization of 3D shock control bumps on a supercritical wing with a sweep angle of 16° at the 1/4 chord. A similar supercritical wing with a higher sweep angle of 24.5° at the 1/4 chord has been adopted as a baseline for the study. Numerical results show that the drag coefficient of the low sweep wing with the optimized 3D shock control bumps is reduced below that for the high sweep wing, indicating shock control bumps can be used as an effective means to reduce the wave drag caused by reducing the wing sweep angle. From the point of view of the wing structure design, lower sweep angle will also bring the benefits of weight reduction, resulting in further fuel reduction.

블레이드 후향각이 원심압축기의 성능과 유동에 미치는 영향

This paper presents a numerical investigation of the influence of the blade back sweep angle on the performance and flow characteristics in a centrifugal compressor with a vaneless diffuser. Five impellers with different back sweep angles were tested in the flow simulations. It was found that a low back sweep angle could improve the total-to-total pressure ratio and the work coefficient over whole operating ranges. However, the flow field in an impeller with a low back sweep angle produced a more non-uniform velocity distribution at the impeller exit because the wake region was significantly increased. As a consequence, the impeller with a low back sweep angle caused a low diffuser performance.

Yaw Angle Effect on Flow Structure over the Nonslender Diamond Wing

Yaw Angle Effect on Flow Structure over the Nonslender Diamond Wing

Effects of flexibility on aerodynamic performance of delta wings with different sweep angles

Force measurement and surface oil flow visualization experiments were conducted in a wind tunnel to investigate the effects of flexibility on aerodynamic performance of delta wings with different sweep angles. The experimental results indicate that the maximum lift coefficient is increased and the stall angle is delayed as the sweep angle increases for both rigid and flexible wings. It is also found that the maximum lift coefficients of the flexible wings with a sweep angle from 35° to 50° are higher than those of the rigid ones. The increment of the maximum lift coefficient in particular achieves 32.9% compared with the case without lift enhancement for the 40° flexible delta wing. Moreover, the surface oil flow visualization experiments show that the stall of the flexible wing of the moderate low sweep angle is accompanied by helical flow structure, while the vortex bursting appears on the corresponding rigid wing.

Flow past a delta wing with a sinusoidal leading edge: near-surface topology and flow structure

The near-surface flow structure and topology on a delta wing of low sweep angle, having sinusoidal leading edges of varying amplitude and wavelength, are investigated using a stereoscopic technique of high-image-density particle image velocimetry at a Reynolds number of 15,000. Identification of critical points, in conjunction with surface-normal vorticity and velocity, provides a basis for determining the effectiveness of a given leading edge. At high angle of attack, where large-scale three-dimensional separation occurs from the wing with a straight leading edge, an amplitude of the leading-edge protuberance as small as one-half of one percent of the chord of the wing can substantially alter the near-surface topology. When the amplitude reaches a value of four percent of the chord, it is possible to completely eradicate the negative focus of large-scale, three-dimensional separation, in favor of a positive focus of attachment. Moreover, alteration of the near-surface topology is most effective when the ratio of the wavelength to amplitude of the sinusoidal leading edge is maintained at a small value.

Control of Flow Structure on Delta Wing with Steady Trailing-Edge Blow

The near-surface flow structure and topology are characterized on a delta wing of low sweep angle, which is subjected to trailing-edge blowing. A technique of high-image-density particle image velocimetry is employed to determine the topological critical points adjacent to the surface and in the near wake of the wing, in relation to the dimensionless magnitude of the blowing coefficient. These topological features are, in turn, interpreted in conjunction with patterns of surface-normal vorticity and near-surface velocity fluctuations, which provide an indication of buffet loading. Even though the jet blowing is at the trailing edge, it has a remarkable, global influence on the surface patterns located well upstream; at high angle of attack, it leads to eradication of large-scale, three-dimensional separation in the vicinity of the apex.

Flow Structure on a Delta Wing of Low Sweep Angle

The instantaneous and averaged flow structure past a delta wing of low sweep angle is investigated using a technique of high-image-density particle image velocimetry. Emphasis is on crossflow planes, where vortex breakdown and stall occur, and the identification of buffeting mechanisms in these regions. At all values of angle of attack up to the fully stalled condition, the averaged vorticity layer exhibits an elongated form; the classical (single) large-scale concentration of vorticity within the leading-edge vortex of a highly swept wing is not present. At low angle of attack α, this elongated, averaged layer can exhibit, however, well-defined concentrations of vorticity. These elongated vorticity layers are accompanied by narrow recirculation zones adjacent to the wing surface. Furthermore, the averaged streamline topology exhibits, at lower α, a saddle point located slightly outboard of the leading edge, in contrast to a saddle point located on the plane of the symmetry of a highly swept wing. Patterns of velocity fluctuation and Reynolds stress show peaks that are generally coincident with large values of averaged vorticity, which indicates that they arise from unsteady events in regions of high shear

Characteristics of fin buffeting over delta wings

Characteristics of fin buffeting over various delta wings was investigated by measurements of fin tip acceleration, particle image velocimetry measurements, and flow visualization. For slender wings, there is a strong asymmetry with regard to the effect of fin location on the buffeting response, and largest response occurs for the inboard fin locations. For outboard fin locations, the buffeting is small, even at high angles of attack when vortex breakdown is well upstream. For inboard fin locations, the largest fin vibrations were found for breakdown locations well upstream of the fin, and when the shear layer of the breakdown region impinges on the fin. The results show that single-fin buffeting may be as important as twin-fin buffeting. For a wing with low sweep angle, vortex shedding rather than vortex breakdown may become the responsible mechanism, and large buffeting excitation for outboard fin locations at a specific incidence is possible. For a double-delta wing, the interaction of strake vortices with the wing flow has a major influence on the buffeting characteristics. Finally, it was found that a nonzero fin angle did not have much effect on the conclusions arrived in this study.

Transonic aeroelastic analysis of the B-1 wing

The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low- and high-sweep cases, at 25.0 and 67.5 deg, respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg sweep case and also for small angles of attack at 67.5 deg sweep case. The aeroelastic response results show that the wing is stable at the low-sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher-sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading-edge separation vortices and not to shock wave motion, as was previously proposed.