We carried out high-resolution large-eddy simulations (LESs) to investigate the effects of several separation-control approaches on a NACA4412 wing section with spanwise width of Lz=0.6 at an angle of attack of AoA=11° at a Reynolds number of Rec=200,000 based on chord length c and free-stream velocity U∞. Two control strategies were considered: (1) steady uniform blowing and/or suction on the suction and/or pressure sides, and (2) periodic control on the suction side. A wide range of control configurations were evaluated in terms of aerodynamic efficiency (i.e., lift-to-drag ratio) and separation delay. Uniform blowing and/or suction effectively delayed flow separation, leading to a lift increase of up to 11%, but yielded only marginal improvements in aerodynamic efficiency. In contrast, periodic control neither enhanced separation delay nor improved efficiency. A detailed analysis of the interaction between uniform blowing and/or suction and turbulent boundary layers (TBLs) over the wing was performed, including assessments of (1) integral boundary-layer quantities, (2) turbulence statistics, and (3) power-spectral densities. Significant modifications in Reynolds stresses and spectral characteristics were observed. To the authors’ best knowledge, this is the first numerical study utilizing high-resolution LESs to provide comprehensive assessments on separation control.

The purpose of this work is to design an aerodynamically shaped vortex generator set-up to delay flow separation on a swept wing with dihedral of an unmanned aerial vehicle being designed at KTH Royal Institute of Technology. Therefore, a 2.5D CFD study of the wing was performed using the Spalart-Allmaras turbulence model. The optimization of the vortex generator set-up followed a multipoint Pareto strategy to establish an optimum design of the vortex generator vanes including its airfoil cross-section. The resulting vortex generator setup achieved a respective improvement of the maximum lift coefficient and stall angle of attack with respect to the baseline wing of 26.34% and 3 deg as a counter-rotating arrangement and 24.02% and 2 deg as a co-rotating set-up. The optimization procedure showed that the optimum cant angle of the vanes, a geometric parameter not tested in the available literature, contributed to 2.45% of the overall improvement of the maximum lift coefficient. The optimization procedure also showed that the flow separation control performance of the vortex generators is sensitive to its airfoil cross-section, and among all the airfoils tested, the S1223 cross-section showed a superior performance. Finally, the optimum height-toboundary-layer-thickness ratio obtained was 1.301 and a further numerical flow visualization demonstrated that the aerodynamically shaped vortex generators produced a vortex system similar to that of a delta wing, with the difference of the influence of the wing’s wall on the axial flow, that generated a primary vortex submerged in the boundary layer. Because of the resulting leading-edge separation vortex system, the penalty drag of the optimized aerodynamically shaped vortex generators was comparable to that of a conventional, flat-plate vortex generator. Nonetheless, the airfoil-shaped vanes produced higher maximum lift coefficients than the flat-plate vanes configurations.

The purpose of this work is to design an aerodynamically shaped vortex generator set-up to delay flow separation on a swept wing with dihedral of an unmanned aerial vehicle being designed at KTH Royal Institute of Technology. Therefore, a 2.5D CFD study of the wing was performed using the Spalart-Allmaras turbulence model. The optimization of the vortex generator set-up followed a multipoint Pareto strategy to establish an optimum design of the vortex generator vanes including its airfoil cross-section. The resulting vortex generator setup achieved a respective improvement of the maximum lift coefficient and stall angle of attack with respect to the baseline wing of 26.34% and 3 deg as a counter-rotating arrangement and 24.02% and 2 deg as a co-rotating set-up. The optimization procedure showed that the optimum cant angle of the vanes, a geometric parameter not tested in the available literature, contributed to 2.45% of the overall improvement of the maximum lift coefficient. The optimization procedure also showed that the flow separation control performance of the vortex generators is sensitive to its airfoil cross-section, and among all the airfoils tested, the S1223 cross-section showed a superior performance. Finally, the optimum height-to-boundary-layer-thickness ratio obtained was 1.301 and a further numerical flow visualization demonstrated that the aerodynamically shaped vortex generators produced a vortex system similar to that of a delta wing, with the difference of the influence of the wing's wall on the axial flow, that generated a primary vortex submerged in the boundary layer. Because of the resulting leading-edge separation vortex system, the penalty drag of the optimized aerodynamically shaped vortex generators was comparable to that of a conventional, flat-plate vortex generator. Nonetheless, the airfoil-shaped vanes produced higher maximum lift coefficients than the flat-plate vanes configurations.

This paper presents a numerical aerodynamic study of the stall characteristics and flow separation mechanismsof a blended wing body unmanned aerial vehicle being developed at KTH Royal Institute of Technologyand proposes a bio-inspired leading-edge modification to control the separation mechanism andimprove the aerodynamic performance of the aircraft at high angles of attack. A numerical study of theaircraft was performed at cruise speed, corresponding to Reynold’s number of 1.3x106, employing an UnsteadyReynolds Averaged Navier-Stokes solver with the Spalart-Allmaras turbulence model. Numericalresults indicated that the aircraft is characterized by the presence of an unstable longitudinal vortex – visibleat the stall angle of 9 deg – which breaks up at an angle of attack of 10 deg, resulting in an unsteady,full-chord stall cell in the mid-span region of the wing section. To mitigate this phenomenon, a modificationto the leading edge between 0.4 m and 1.8 m wing spans was implemented inspired by the geometryof the nose of a porpoise whale, effectively generating a porpoise (hump) leading-edge inboard section.Preliminary numerical results indicate an increase in stall angle of attack to ∼13 deg and an in maximumlift coefficient to ∼1.0. Furthermore, the porpoise hump allowed controlling the stall behavior of the aircraftby enforcing a wing tip, trailing-edge separation stall achieved by the generation of an extended flowacceleration region at the leading edge.

The manuscript presents the conceptual design phase of an unmanned aerial vehicle, withthe objective of a systems approach towards the integration of a hydrogen fuel-cell system and Li-ionbatteries into an aerodynamically efficient platform representative of future aircraft configurations.Using a classical approach to aircraft design and a combination of low- and high-resolution computationalsimulations, a final blended wing body UAV was designed with a maximum take-off weightof 25 kg and 4 m wingspan. Preliminary aerodynamic and propulsion sizing demonstrated that theaircraft is capable of completing a 2 h long mission powered by a 650Wfuel cell, hybridized with a100 Wh battery pack, and with a fuel quantity of 80 g of compressed hydrogen.

Optical phenomena in the upper atmosphere, such as northern lights, airglow, noctilucent clouds and thunderstorm-related transient luminous phenomena reveal the complex processes coupling different layers of the atmosphere and the near earth space. Bad weather and lighting conditions, as well as geographical constraints, limit the possibilities of ground based imaging. Therefore, an autonomous high altitude long endurance (HALE) fixed-wing unmanned aerial vehicle (UAV) is proposed for atmospheric imaging, as a joint student-driven research project between the Aeronautics and Vehicle Engineering- and the Space and Plasma Physics departments at KTH Royal Institute of Technology. The Autonomous Light Platform for High Altitude atmospheric imaging (ALPHA) is specifically designed for operations in the environmentally harsh conditions found in Arctic nighttime. This paper presents the conceptual design phase of the aircraft, as well as the initial manufacturing and flight testing methodology of a half-scale prototype.

An experimental study was conducted to investigate and compare the effects of beveled and double-beveled circular nozzles on supersonic jet screech at a nozzle pressure ratio of 5. In particular, near- and far-field microphone measurements and schlieren visualizations were utilized to look into selected acoustic and flow features associated with jet screech radiation. Results show that beveled nozzles eliminate jet screech by producing asymmetric shock structures and instability waves that are mismatched in phase and amplitude. In contrast, double-beveled nozzles produce symmetric shock structures and amplify screech intensity, even when jet mixing effects have been significantly enhanced. It is further observed that amplified jet screeches produced by double-beveled nozzles are highly unsteady and undergo nonperiodic and stochastic temporal variations. Last but not least, double-beveled nozzles also significantly impact screech peak noise locations and confer different changes along different measurement planes. The present study demonstrates that not all beveled-type nozzles are able to mitigate jet screech, with nonoptimal designs amplifying it instead.

Wing-in-Ground (WIG) effect aircraft have become an interesting concept in the context of reducing the environmental footprint and increasing the speed of coastal transport. However, obtaining an efficient wing shape in a cost-effective manner is an elusive goal in the early stages of any aircraft design. In this work, a multi-objective wing planform optimisation methodology is proposed by combining a parametric shape modeller OpenVSP, a low fidelity solver VSPAERO and Non-dominated Sorting Genetic Algorithm (NSGA-II) to support the preliminary design of wing-in-ground effect aircraft. Methodology is demonstrated by performing three different wing planform optimisations ranging from planar wing optimisation to nonplanar wingtip optimisation by improving both lift to drag ratio and static height stability characteristics of a wing planform in ground effect. The analysis of the Pareto optimal solutions suggests that when employing the Vortex Lattice Method (VLM) for aerodynamics and stability derivatives computation, the optimiser converged to drooped wing type configuration which enhances both aerodynamic efficiency and static height stability characteristics of wing-alone configuration.

This paper discusses the wind tunnel test of a 37.5% scale model of a blended wing body unmanned aerial vehicle. The model was mounted on a fixed belly-fairing sting, and the pitch variation was facilitated via a linear actuator. The model and set-up were directly mounted on a force plate placed underneath the tunnel floor. Aerodynamic loads were acquired at a range of wind tunnel velocities between 10m/s and 30m/s, corresponding to a Reynolds number range between ∼246 000 and ∼760 000, and results show good repeatability throughout the experimental campaign. A virtual wing tunnel was set up using computational fluid dynamics representative of the real-life facility, and the model was simulated inside the tunnel for a range of angles of attack between -4deg and 10deg without support, and at a reduced range between 0deg and 6deg with the support “installed.” Classical analytical methods for boundary and support interference corrections were implemented, and corrections factors were estimated by numerical simulation of the model with and without the support installed. Preliminary comparison between experimental measurements from force balance and numerical data show satisfactory agreement in the prediction of the lift curve slope, with some disagreement in the prediction of the stall condition. Prediction of drag and pitching moment are significantly impacted by the presence of the support and show poor agreement with numerical data.

Detailed near- and far-field acoustic measurements were conducted for two circularstepped nozzles with 30 deg and 60 deg design inclinations at over- and perfectlyexpandedsupersonic jet flow conditions and compared to those for a circular nonsteppednozzle. Far-field acoustic results show that stepped nozzles play an insignificant role inaltering noise emissions at perfectly expanded condition. At an over-expanded condition,however, the longer stepped nozzle produces significant noise reductions at the sidelineand upstream quadrants, while the shorter stepped nozzle does not. Noise spectra analysisand Schlieren visualizations show that noise reduction can be primarily attributed tomitigations in the broadband shock-associated noise (BSAN), due to the ability of the longerstepped nozzle in suppressing shock strengths at downstream region. Near-fieldacoustic measurements reveal that the source region, as well as the intensity of turbulentand shock noises, are highly sensitive to the stepped nozzle configuration. Furthermore,BSAN seems to be eliminated by the longer stepped nozzle in near-field region due to theshock structure modifications.