Deterministic Optimization of Single-Slotted Flaps Using an Automated CFD Workflow

Pulsed jet actuators (PJAs) are one of the candidate technologies to be integrated in Fowler flaps to increase the maximum lift coefficient of transport aircraft in the landing configuration. The total system consists of the actuators plus sensors, a piping system to supply pressurized air and a (redundant) power and communication system to provide actuator control. In this paper, it is investigated what increase in the maximum lift coefficient is required to justify the added weight and power off-takes that accompany the integration of pulsed jet actuators. This is done by making an automated design process for the overall aircraft, the piping assembly system, and the electrical wiring interconnection system. These last two sub-systems rely on KBE techniques that automate dimensioning and performance evaluation. A test case is specified that encompasses the design of a typical single-aisle mid-range aircraft with and without the PJA system installed. It is concluded that the introduction of the PJA system requires at least an increase in maximum lift coefficient of 0.2 to justify the increase in system mass and power off-takes. Furthermore, it is shown that if the maximum lift coefficient increases with 0.4, only small reductions in maximum take-off weight (−0.3 %) and operating empty weight (−0.6 %) can be expected, while the total fuel burn remains virtually constant.

This study presents a surrogate-model-assisted Quasi-Newton optimisation framework for simultaneously improving the aerodynamic performance and radar stealth characteristics of an unmanned aerial vehicle (UAV). High-fidelity computational fluid dynamics (CFD) and computational electromagnetics (CEM) simulations are integrated through surrogate models generated via a face-centred central composite design within a design of experiments framework. Quadratic polynomial response surface equations are constructed for key aerodynamic and radar cross-section (RCS) metrics, enabling analytical gradient evaluation. A gradient-based quasi-Newton method with Broyden–Fletcher–Goldfarb–Shanno Hessian updates is employed to minimise a scalarised objective function combining normalised maximum lift coefficient, overall RCS and frontal RCS. Constraints are imposed on the lift-to-drag ratio ( $L/D \geq 10$ ) and static longitudinal stability ( ${C_{m0}} \geq 0$ ). Analytical derivatives from the response surface equations (RSEs) eliminate the need for direct numerical differentiation of CFD/CEM outputs, reducing computational cost and eliminating simulation noise. An interior-point sequential quadratic programming strategy is used to ensure satisfaction of nonlinear constraints during the optimisation process. The optimised UAV design demonstrates a $12{\rm{\% }}$ increase in maximum lift coefficient and a $30{\rm{\% }}$ reduction in both overall and frontal RCS compared to the baseline configuration. The results are confirmed through high-fidelity CFD and RCS simulations and are further validated experimentally in an anechoic chamber, with close agreement across all measured frequencies. The proposed methodology provides an efficient and experimentally verified approach for integrated aerodynamic and stealth optimisation in UAV design.

<p>Flow control has significant technological importance as it can manipulate the flow field in a desired way either actively or passively. This wide research area has remained the point of attention for many years as it is applicable to various applications. Blowing as a flow control method, among other methods, is more technically feasible and market ready technique. A brief review from the existing literature on various studies on blowing has been presented along with their outcome. Then, studies were conducted to investigate the performance variation (in terms of lift coefficient, drag coefficient) of different airfoils with respect to various blowing parameters. It was observed that for NACA 0012 airfoil the maximum lift coefficient peaks at a blowing ratio 0.2 and then it decreases whereas the stalling angle increases with rise of blowing ratio. For LA203A both maximum lift coefficient and stalling angle increases with blowing ratio. For thick airfoil, it was found that lift increases with the rise of moment coefficient and blowing ratio and mid chord slots gives better performance at lower angles of attack whereas leading edge slots exhibits better performance at higher angles of attack. In case of thick elliptical airfoil, increase in lift coefficient was noticed with the increase in moment coefficient and blowing ratio but an optimum jet width to chord ratio (0.41) was found beyond which increase in jet width causes drop in lift coefficient. Additionally, it was found that at lower blowing angle thick elliptical airfoil performs better compared to that of higher angles of attack. Study on NACA 0012 and Aerospatiale A proved the fact that lifts increases with rise in blowing ratio at three different jet diameters to chord ratios. This study also found that higher blowing ratio is also advantageous for the turbo machinery in increasing the pressure difference. Lastly, the study on low Re airfoil flow found that increasing blowing ratio has a deleterious effect on aerodynamics performance (in terms of lift and drag coefficient).</p>

In this study, Reynolds-averaged Navier-Stokes simulation (RANS) of the uniform blowing and suction (UB/US) control on a Clark-Y airfoil was performed aiming at improving an airfoil performance by friction drag reduction. First, the control effect when only the uniform blowing control or uniform suction control is applied on the airfoil surface was investigated by changing the control locations. The blowing or suction velocity was 0.14% of the free-stream velocity and the blowing/suction area was set at four different locations from the leading edge to the trailing edge on both the upper and lower surfaces. The Reynolds number based on the chord length is 1.5 × 106. The angle of attack is set to 0°. It was found that friction drag is decreased/increased by single UB/US control. It was also found that the lift-to-drag ratio improved with UB on the lower surface or US on the upper surface, and decreased with UB on the upper surface or US on the lower surface. In the combined control of UB and US, the blowing and suction velocity was 0.14% or 0.26% of the free-stream velocity and the locations of blowing/suction and flow conditions were the same as those in the cases with either UB or US. It seemed that the lift-to-drag ratio was improved by the combined control of UB on the lower surface and US on the upper surface. In particular, the lift-to-drag ratio was most improved by US on the lower rear surface and UB on the upper rear surface.

In this study, we determined an aerodynamic configuration to design structures applying composites for large-scale horizontal-axis wind turbine blades. A new aerodynamic and structural design method for large wind turbine blades is presented. The rated power of the wind turbine blade is 25 MW class. The tip speed ratio is 7. The diameter of the designed blade is 260 m. Therefore, thick airfoils were selected to design large-scale wind turbine blades considering structural stiffness and maximum lift coefficients. For the aerodynamic design method, it was designed with the optimal angle of attack having the maximum lift-to-drag ratio. The blade element theory and vortex theory were applied to aerodynamic design. For the aerodynamic design results, its validity was investigated via aerodynamic performance analysis. As a result of analyzing aerodynamic performance, it was confirmed that higher power was generated. At 12.5 m/s of rated wind speed, electrical power was 28.32 MW. The structural design considering the aerodynamic design results was carried out. The composite laminate theory was adopted. Structural safety was evaluated for the designed blades. Finally, the structural design results were analyzed as sufficiently valid.

In this research, the flow physics and aerodynamic performance of dragonfly cross sections, used in Micro Aerial Vehicles (MAVs), in low Reynolds are investigated. The main objective of the research is to study the performance of dragonfly wing cross-sections flapping motion in Reynolds 5000 and 10,000. Pitching motion is one of the most important mechanisms in force lifting generation, and the effects of Reynolds number and mean angle of attack on aerodynamic coefficients have been extensively investigated for the pitching motion. In the present study, the geometry of two cross sections of dragonfly was extracted. Incompressible, two-dimensional and unsteady Navier–Stokes equations have been used to simulate the flow. k − ɛ RNG model was used for turbulence modeling. To simulate the wing pitching motion, the dynamic mesh method was used. The results showed that in flapping motion, pitching-up rotation has caused a rapid increase in lift coefficient. Furthermore, it was found that the absence of stall does not increase the lift and drag coefficients, while formation of new strong vorticity layers have caused an increase in lift coefficient. On the other hand, corrugations on the cross sections of the dragonfly in the pitching motion cause the delay of separation and increasing the lift coefficient. In flapping motion and the pitching motion, the lift coefficients of three cross sections were increased due to stronger vorticity layers by reducing the Reynolds number. Due to the existence of corrugations, the first and the second cross sections have good aerodynamic performance, compared to the flat plate. The comparison carried out in the current research showed that the second cross section is a proper replacement for the flat plate in MAVs.

The present work has as objective to do a comple te aerodynamic analysis of the Unmanned Aerial Vehicle for Ecological Conservation. The panel method code PAN AIR is used to compute the inviscid flowfield. The viscous effects in drag are estimated by the classic Hoerner method, and the maximum lift coeff icient via the classic Valarezo and Chin method. The numerical aerodynamic forces of the complete airplane are compared to experimental data for validation. The spanload and wing pressure distribution are estimated for four configurations: wing, wing -body, wing -body -tail, and wing -body -tail with wing twist. The sources of induced drag for all configurations are achieved graphically via Trefftz plane. All the data were estimated at cruise flight, Reynolds number equal to 1.413×10 6

The purpose of this research is to study the lift and drag characteristics of a NACA 0015 airfoil with a backward facing step. A backward facing step on the upper surface has the potential to improve the aerodynamic characteristics by reducing the drag coefficients and increasing the lift-to-drag ratios considerably. For a free stream velocity, the point of separation is determined by CFD analysis with increasing the angle of attack. A backward facing step is introduced just prior to the point of separation. The goal of improving the aerodynamic performance of the airfoil is inserted by searching the best configuration through gradually increasing the depth of facing step. A wind tunnel test on a NACA 0015 has been also conducted to verify the results from CFD analysis. The results show that reduction in drag is possible by separation delay with introducing the step. Though the reduction in lift has been observed, the lift to drag ratio also increases which is an important parameter in aerodynamic characteristics.

PurposeSome recent effort showed that usage of Krueger flaps helps to maintain laminar flow in cruise flight. Such flaps are positioned higher relative to the chord to shield the leading edge from the insect contamination during take-off. The flap passes several through critical intermediate position during the deployment to its design position. The purpose of this paper is to analyse the aerodynamics.Design/methodology/approachTo better understand such flow phenomena, the combined approach of computational fluid dynamics and experimental methods were used. Flow simulation was performed with in-house finite volume Navier–Stokes solver in fully turbulent unsteady RANS regime. The experimental data were obtained by means of force and pressure measurements and some areas of the flow field were examined with 2 C particle image velocimetry.FindingsThe airfoil with flap in critical position has a very limited maximum lift coefficient. The maximum achievable lift coefficient during the deployment is significantly affected by the vertical position of the trailing edge of the flap. The most unfavourable position during the deployment is not the flap perpendicular to the chord, but the flap inclined closer to it is the retracted position.Research limitations/implicationsThe flap movement was not simulated either in the simulation or in the experiment. Only intermediate static positions were examined.Practical implicationsA better understanding of aerodynamic phenomena connected with the deployment of a Krueger flap can contribute to the simpler and lighter of kinematics and also to decrease time-to-market.Originality/valueLimited experimental and computational results of Krueger flap in critical positions during the deployment are published in the literature.

THE possibility of applying boundary layer control by suction at the trailing edge to reduce the profile drag of a helicopter rotor blade using centrifugal forces to provide the suction was investigated. It was shown theoretically that it is possible to use this method of suction, and that it is advantageous to suck at the trailing edge if the drag reduction obtained was greater than a certain value which depended on the amount of suction. Experimental work was carried out on a model wing with different types of slots to determine the profile drag reduction possible using theoretically predicted suction flow rates. The experiments were carried out at an effective Reynolds Number based on wing chord of 0.62 x 108. It was found that a single long slot of 2 per cent chord width spanning the whole blade, as close as possible to the trailing edge on the upper surface, gave the best results. The maximum drag reduction occurred at high values of lift coefficient and was of the order of 20 per cent. Using the experimental results it was shown that a helicopter employing this type of boundary layer control would require less engine power.

Aerodynamic shape optimisation technology is presented, comprising an efficient variable fidelity shape parameterisation method, an efficient andhigh quality mesh deformation scheme, and a parallel optimisation algorithm. The objective of the research presented here is the comparison of truly three-dimensional optimisations of aircraft wings in both aerodynamic and aeroelastic environments. The novel shape parameterisation technique allows various fidelities of design parameters, ranging from detailed surface changes to novel truly three-dimensional planform adjustments. An efficient interpolation scheme, using radial basis functions, transfers domain element movements into direct deformations of the design surface and corresponding CFD mesh, thus allowing total independence from the grid generation package and type (structured or unstructured). Optimisation is independent from the CFD flow solver by obtaining sensitivity information for an advanced parallel gradient-based optimiser by finite-differences. This ‘wrap-around’ optimisation technique is applied to a modern large transport aircraft wing in the cruise flight condition for minimum drag with stringent constraints in lift, volume, and two root moments. The objective of all optimisations is aerodynamic, however the static aeroelastic deflection provided by an aeroelastic solver will give that particular optimisation increased accuracy and real world relevance. The result of a constrained inviscid aerodynamic optimisation is presented and has a significant reduction in drag when compared to the initial wing with no violation of any constraints. The shape parameterisation method demonstrates that only a low number of design variables are necessary to achieve innovative planform and surface geometries with dramatically improved performance. Computational fluid dynamics (CFD) methods are now commonplace in aerospace industries, and at the forefront of analysis capabilities, providing a fast and effective method of predicting a design’s aerodynamic performance. However, with ever increasing complexity of designs, engineers can often struggle to interpret the intricacies of the CFD results sufficiently to be able to manually alter the geometry to improve performance. Hence, there has been an increase in demand for intelligent and automatic shape optimisation schemes. This requires combining geometry control methods with numerical optimisation algorithms, to provide a mechanism to mathematically seek improved and optimum designs, using CFD as the analysis tool. Optimisation requires consideration of three issues, each of which have numerous solutions: shape parameterisation including CFD surface and volume mesh deformation, computation of design variable derivatives, and effective use of these derivatives to improve design. Geometry parameterisation is critical for effective shape optimisation. This is the method of representing the design surface, and defines the degrees of freedom in which the geometry can be altered and, ideally, this should be linked with an effective method of deforming the CFD surface and volume mesh in a corresponding fashion. Parameterising complex shapes is a problem that remains a serious obstacle to both manual and automatic CFD-based optimisation. A wide variety of shape control and morphing methods have been developed, but many do not allow sufficiently

Computational fluid dynamics simulation of aerodynamic performance and flow separation by single element and slatted airfoils under rainfall conditions

Considerable drag reduction and fuel saving of a tractor–trailer using additive aerodynamic devices

Aerodynamic drag of fast-moving train has significant impact on its fuel consumption and design safety. To improve aerodynamic performance, the drag forces on the train surfaces must be reduced. Train’s front-end nose design has boundless geometrical variations that can be applied for improved performance. Three geometric characteristics were used in present research: A-pillar roundness, nose length that controls its shrinking, and nose bluntness. Latin-Hypercube based random sampling method was used to determine appropriate values of these parameters for specific range. Using Computational Fluid Dynamics (CFD), a numerical approach was used to analyze train aerodynamic performance based on selected factors and operating conditions. Comparisons based on analysis of the base design and proposed design models were made and evaluated for enhanced performance. CFD analysis of the base design and iterative improvements in modified designs indicate drag reduction through change in A-pillar roundness and bluntness of nose to be 10% and 22%, respectively. The increase in nose length which controls nose shrinkage causes the drag to increase by 35%. Similarly, the boundary layer and pressure distribution on the front end of train were also considered and analyzed for performance improvement. It was concluded that the length controlling nose shrinkage of proposed train design is the key factor among the selected geometric parameters that has more influence on drag on high-speed train surface in comparison to other parameters considered. This research effort offers a modification, analysis and comparison of front-end nose geometry of high-speed train to improve aerodynamic performance and consequently fuel consumption.

An experimental study was conducted to assess the benefits of using flexible-membrane airfoils/wings at low Reynolds numbers for micro air vehicle applications compared with using a conventional rigid airfoil/wing. In addition to measuring aerodynamic forces acting on flexible-membrane airfoils/wings, a high-resolution particle image velocimetry system was used to conduct flowfield measurements to quantify the transient behavior of vortex and turbulent flow structures around the flexible-membrane airfoils/wings to elucidate the associated underlying fundamental physics. The aerodynamic force measurements revealed that flexible-membrane airfoils could provide better aerodynamic performance compared with their rigid counterpart at low Reynolds numbers. The flexibility (or rigidity) of the membrane skins of the airfoils was found to greatly affect their aerodynamic performance. Particle image velocimetry measurements elucidated that flexible-membrane airfoils could change their camber (i.e., crosssectional shape) automatically to adapt incoming flows to balance the pressure differences on the upper and lower surfaces of the airfoils, therefore suppressing flow separation on the airfoil upper surfaces. Meanwhile, deformation of the flexible-membrane skins was found to cause significant airfoil trailing-edge deflection (i.e., lift the airfoil trailing edge up from its original designed position), which resulted in a reduction of the effective angles of attack of the flexible-membrane airfoils, thereby delaying airfoil stall at high angles of attack. The nonuniform spanwise deformation of the flexible-membrane skins of the flexible-membrane airfoils was found to significantly affect the characteristics of vortex and turbulent flow structures around the flexible-membrane airfoils.