Explore the vast range of titles available.

Forward-swept wings can be expected to be lower-boom planforms with similar amount of drag as backward-swept wings because of their good lift distributions. In this study, the equivalent area distribution of a ten-seater supersonic forward-swept wing aircraft with a canard was designed to obtain design knowledge for leading boom reduction. The equivalent area distribution of the aircraft was calculated by solving the compressible Euler equation. A feasible target equivalent area distribution was generated based on Darden’s method and compared with the equivalent area distribution. To achieve a closer match in terms of lift and geometry with the target, the main wing planform and the position of the main wing along the body and vertical axes were modified. The low-boom performances were evaluated using the extended Burgers equation. The design results indicated that the forward-swept wing configuration with a canard could divide the single peak of the leading boom into two peaks. Thus, the sonic boom strength of the canard configuration was 2.5 PLdB lower than that of the configuration without the canard wing.

The current study investigates the natural laminar-to-turbulent transition in the boundary layers of straight (STR) and swept (SWE) wings. The experiments are conducted in a subsonic low-speed wind tunnel using two wings configurations employing infrared thermography and pressure measurements. More specifically, image processing is used to analyze the flow pattern and infer the effects of Tollmien–Schlichting (T-S) and crossflow (CF) instabilities at low to moderate Re numbers (Re < 10 6 ), which is one of the main contributions of this research. Additionally, using the pressure distribution data as a reference, point analyses are conducted to explore the physical mechanisms of transition. The results of this study provide insight into the boundary layer transition region in straight and swept wings, concerning the effects of low Re numbers and their viscous associated implications, such as laminar separation bubbles (LSB) and adverse pressure gradient, phenomena that are encountered in many aerospace applications such as small UAVs with fixed wings.

Lift alleviation by a mini-spoiler on aerofoils, unswept and swept wings encountering an isolated counter-clockwise vortical gust was investigated by means of force and velocity measurements. The flow separation region behind the spoiler remains little affected during the gust encounter. The maximum lift reduction is found for the static stall angle of attack. The change in the maximum lift during the gust encounter is approximately equal to that in steady freestream. The comparison with plunging aerofoils reveals that, for the same maximum gust and plunge velocity, the effectiveness of the mini-spoiler is much better in travelling gusts. This reveals the importance of the streamwise length scale of the incident gust. For the unswept wing, there is some three-dimensionality of the flow separation induced by the mini-spoiler near the wing tip. The magnitude of the lift reduction can be estimated using the aerofoil data and by making an aspect ratio correction for the reduced effective angle of attack. For the swept wing, the mini-spoiler can disrupt the formation of a leading-edge vortex induced by the incident vortex on the clean wing and can still reduce the maximum lift.

Theoretical assessments of the crossflow (CF) stabilization due to flow slip produced by small grooves on a swept supersonic wing are performed using the linear theory for inviscid flow, the local similar approximation of the boundary layer flow, the slip boundary conditions on the grooved surface and the linear stability theory. The eN computations for stationary CF mode predict that spanwise-invariant grooves with their half-period equal to 0.25 of the boundary-layer displacement thickness can delay the CF-induced transition onset by about 10% on a 30∘ swept wing having a parabolic airfoil of 5% thickness ratio, at freestream Mach number 2. It is concluded that the groove laminarization concept deserves further studies.Graphic abstract

In this paper the design optimization-under-uncertainty of a forward swept wing (FSW) blended wing body (BWB) unmanned aerial vehicle (UAV) is examined. Conventional BWBs are often tailless, which leads to a backward swept wing to ensure longitudinal static stability. This in turn can induce flow separation at the tip, leading to a loss of lift, controllability and the appearance of a nose-up pitching moment. A possible solution to this problem is a conceptual redesign by introducing a forward swept wing, which is inherently free of tip-stall, but needs a careful design in order to be controllable. However, fixed wing UAVs are often produced by means of direct injection expanded foam moulding, which is characterized by not negligible production tolerances. This lead to a reliability-based robust design optimization problem, for which a novel framework is employed: SAMURAI. Firstly, the method accounts for computational cost by means of surrogate modelling, an analytical treatment of the problem and an asynchronous updating scheme that balances design space exploration and objective exploitation. Secondly, the method treats the problem as a multi-objective problem, which leads to a Pareto front of robust and reliable designs. The result is a novel series of UAV designs that are inherently free of tip stall, perform robustly and meet the stability requirements with the target reliability obtained with a computationally feasible budget.

Flow over a moderately swept wing is characterised by complex localised flow vortex topologies such as ‘closed’ separation bubbles or ‘open’ separation structures. A model of a complex cambered, twisted, tapered wing with 40° leading edge sweep, representative of those designed for manoeuvre at high subsonic Mach numbers, was investigated using the oil-film visualisation, stereo particle image velocimetry and force moment measurements. Wind-tunnel tests were conducted at a range of Reynolds number from 2.1×105 to 8.4×105 and at angles of incidence from −1° to 22°. Still images combined with video clips enable flow patterns over the wing model to be interpreted more clearly and accurately. Using successive images extracted from the video of flow visualisation, the movement of the oil pigment has been estimated. The influence of the Reynolds number and incidence angle was discussed through analysing the flow pattern over the wing surface. Additionally, the link between the flow structures present and the wing aerodynamic performance was studied.

In this work, a study has been conducted to observe the influence of fuel weight distribution on the flutter characteristics of a high aspect ratio swept wing. In this paper, B777-200 wing model having a 34º sweptback angle is used as a baseline and two other models with the swept angles of 0º (straight wing) and 30º (forward swept) are considered in this study. Aeroelastic analysis is performed ranging from sea level up to 35,000 ft and the influence of in-flight fuel management in several flight altitudes is also investigated. There are four different fuel distribution model are investigated and it was found that some fuel distribution configuration are more critical which may lead to flutter. Moreover it was found that the flight altitude significantly affects the aeroelastic instability.

Variable forward-swept wing (VFSW) aircraft is a kind of typical morphing aircraft. With the change of forward-swept angle, the airfoil of Variable forward-swept wing which is based on incoming flow will also be changed, and this change has a very important influence on flight performance. In order to explore this impact, this paper firstly build nine airfoil models under different forward-swept angle, through the calculation by fluent, the impact on airfoil aerodynamic characteristics from angle of attack, mach number and forward-swept angle is analyzed. The results show that, while flying at the sea level in a low angle of attack, the lift coefficient, resistance coefficient and moment coefficient will increase with the rise of forward-swept angle, but while flying in a high angle of attack, these coefficients will decrease with the rise of forward-swept angle. The results also can provide references for further research of aerodynamic configuration of variable forward-swept wing aircraft.

Promising low-noise aircraft architectures have been identified over the last few years at DLR. A set of DLR aircraft concepts was selected for further assessment in the context of sustainable and energy-efficient aviation and was established at the TU Braunschweig in 2019, the Cluster of Excellence for Sustainable and Energy-Efficient Aviation (SE2A). Specific Top-Level aircraft requirements were defined by the cluster and the selected DLR aircraft designs were improved with focus on aircraft noise, emissions, and contrail generation. The presented paper specifically addresses the reduction of aviation noise with focus on noise shielding and modifications to the flight performance. This article presents the state of the art of the simulation process at DLR and demonstrates that the novel aircraft concepts can reduce the noise impact by up to 50% in terms of sound exposure level isocontour area while reducing the fuel burn by 6%, respective to a conventional aircraft for the same mission. The study shows that a tube-wing architecture with a top-mounted, forward-swept wing and low fan pressure ratio propulsors installed above the fuselage at the wing junction can yield significant noise shielding at improved low-speed performance and reduce critical fuel burn and emissions.

In order to study the optimal sweepback angle when a variant unmanned aerial vehicle (UAV) exhibits a low radar cross-section (RCS) indicator during phase flight, an auto sweep scheme based on electromagnetic scattering evaluation and an improved particle swarm optimization algorithm was presented in this article. An aircraft model with variable swept wings was built, and high-precision grids were used to discretize the target surface. The results showed that the optimal sweep angle did not change with the increase in the initial azimuth angle when the observation field was horizontal and the ending azimuth was 90°. While the increase in the elevation angle affected the optimal sweepback angle of the aircraft under the given conditions, when the observation initial azimuth angle was 90°, the auto sweep scheme could reduce the mean and some minima of the RCS indicator curve of the aircraft and could provide the aircraft with an optimal sweep angle under different observation conditions. The presented method was effective in learning the optimal sweep angle of the aircraft when low scattering characteristics were required during the phase flight.