Lift Distribution Research Articles

Simultaneous shape and topology optimization of wings

This paper presents a method for simultaneous optimization of the outer shape and internal topology of aircraft wings, with the objective of minimizing drag subject to lift and compliance constraints for multiple load cases. The physics are evaluated by the means of a source-doublet panel method for the aerodynamic response and linear elastic finite elements for the structural response, which are one-way coupled. At each design iteration, a mapping procedure is applied to map the current wing shape and corresponding pressure loads to the unfitted finite element mesh covering the design domain. Wings of small fixed-wing airplanes both with and without a stiffening strut are optimized. The resulting wings show internal topologies with struts and wall-truss combinations, depending on the design freedom of the shape optimization. The lift distributions of the optimized wings show patterns like the ones obtained when performing optimization of wing shapes with constraints on the bending moment at the root.

Over-wing integration of ultra-high bypass ratio engines: A coupled wing redesign and engine position study

The integration of next-generation high-bypass turbofan engines poses a major challenge to the aeronautical industry due to the larger fans necessary to achieve more fuel-efficient engines. The limited space underneath the wings and the strict ground clearance constraints bring the necessity to investigate solutions other than the conventional under-wing mounted engines. Over-wing installed nacelles have the potential to solve the ground clearance issue and, in addition, might reduce ground noise due to acoustic shielding from the wing. Nevertheless, a strong and complex coupling between aerodynamics and propulsion is the result of such integration choice, and traditional design practices may result in configurations with prohibitively high drag penalties. This paper presents a novel wing redesign method, specifically developed for over-wing mounted engines. The wing is reshaped to recover the spanwise lift distribution of the clean airframe (wing-body) configuration, for a single aisle airliner at a cruise condition. The wing redesign is conducted along with an engine position sensitivity study, in which the wing is reshaped for different engine axial and vertical positions. The coupling between propulsion and aerodynamics is thoroughly investigated, as well as the interaction and interference effects between the wing, pylon, and nacelle. Moreover, the best over-wing solution is compared to a baseline under-wing mounted nacelle. Results show that, by applying the developed method, an overall drag reduction of 17.65 counts, or 6.4%, was obtained, compared to the initial over-wing configuration, comprising the original wing and baseline engine position. Nonetheless, the best over-wing nacelle design is still 5.58 counts, or 2%, higher in overall drag compared to the baseline under-wing mounted nacelle case.

Numerical analysis of a biomimetic UAV with variable length grids wingtips

Birds can reduce the induced drag using an elongation of the wings called primary feathers. In this paper, a numerical analysis of a biomimetic Unmanned Aerial Vehicle (UAV) that imitates that elongation at its wing tips is presented. Specifically, the UAV has a rectangular wing and three grids at the tip that changes the lift distribution over the wing. By using Computational Fluid Dynamics (CFD) the effect of the grids span has been compared for the UAV without grids and for 1/3, 2/3, and fully extended grids. Grids with a variable size along the chord are also tested. The aerodynamic performance (lift, drag, and efficiency), and the rolling capabilities have been obtained under different flight conditions. The results have demonstrated that an increase of grids span translates into a 15–20% increase in aerodynamic efficiency of the UAV during cruise and climb phases of the flight. The grids also provide the ability for rolling under positives angles of attack.

Two-Way Coupled Aero-Structural Optimization of Stable Flying Wings

An aero-structural algorithm to optimize a flying wing in cruise conditions for preliminary design is developed using two-way interaction between the structure and aerodynamics. A particle swarm routine is employed to solve the multi-objective optimization, aiming to reduce the weight of the structure and the aerodynamic drag at the design point. Different shapes are evaluated during the optimization process until the algorithm reaches the optimal wing aspect ratio, taper ratio, angle of incidence, twist angle, swept angle, and airfoil shape, where a six-parameters method is employed to allow reflex airfoils. A main isotropic I-beam models the wing structure. An extended vortex lattice model is employed to model the aerodynamics, along with a high-order panel method with fully coupled viscous interaction. The finite element method is used to solve the flying-wing structure under static loads. An algorithm is developed to iterate between the deflection of the wing and its impact on the aerodynamics until convergence is reached. Different constraints are implemented into the objective function to fulfil the structural criteria and the longitudinal static stability. A comparison against a baseline optimization is carried out, achieving higher efficiency and promising results in elliptical lift distribution, and a high static margin, without the use of non-constant twist. The results suggest that combining both reflex airfoils and sweep with washout is the optimal solution to reduce the drag and weight, keeping the longitudinal static stability criteria for tailless aircraft in the lower end of the transonic regime.

Effect of longitudinal lift distribution on sonic boom of a canard-wing-stabilator-body configuration

After the last flight of the Concorde in 2003, sonic boom has been one of the obstacles to the return of a supersonic transport aircraft to service. To reduce the sonic boom intensity to an acceptable level, it is of great significance to study the effect of lift distribution on far-field sonic boom, since lift is one of the most important contributors to an intense sonic boom. Existing studies on the longitudinal lift distribution used low-fidelity methods, such as Whitham theory, and in turn, only preliminary conclusions were obtained, such as that extending the lift distribution can reduce sonic boom. This paper uses a newly developed high-fidelity prediction method to quantitatively study the effect of longitudinal lift distribution on the sonic boom of a Canard-Wing-Stabilator-Body (CWSB) configuration. This high-fidelity prediction method combines near-field CFD simulation with far-field propagation by solving the augmented Burgers equation. A multipole analysis method is employed for the extraction of near-field waveform in order to reduce computational cost. Seven configurations with the same total lift but different distributions are studied, and the quantitative relationship between the longitudinal lift distribution and far-field sonic boom intensity is investigated. It is observed that a small lift generated by the stabilator can prevent aft-stabilator and aft-fuselage shocks from merging, while the balanced lift generated by the canard and wing can effectively keep the corresponding shocks further apart, which is beneficial for reducing both the on-track and off-track sonic boom. In turn, the acoustic level perceived at the ground can be reduced by 5.9 PLdB on-track and 5.4 PLdB off-track, on average.

Study on the Performance of a Cycloidal Rotor in Hover at Different Reynolds Numbers

AbstractThis paper studied the influence of the Reynolds number on lift, power, and efficiency of a three-bladed cycloidal rotor in hover. In a parameter study, Reynolds numbers from 10,000 to 600,000 were investigated. The fluid mechanics were solved using incompressible 2D unsteady Reynolds-averaged Navier–Stokes (URANS) computational fluid dynamics (CFD). The CFD model was first carefully validated using experimental data for a pitching airfoil undergoing dynamic stall. The model was then verified to reproduce the shapes of both the force and power curves and the wake flow of a landmark cyclorotor experimental study. Two different flow regimes were identified: for the first regime Re ranged from 10,000 to 100,000 and two dynamic flow separations occurred, the first at maximum pitch angle and bottommost position of the blade and the second shortly afterwards. The second flow regime had Re from 200,000 to 600,000 and avoided the first separation due to the increased Reynolds number. Both separations were visible in the flow as well as in the azimuthal lift distribution. Regardless of the flow regime, rotation averaged lift and power followed the predictions of momentum theory, except for Re of 500,000. At this Re, a significant drop in power was observed and corresponds to findings of a 20-year old experimental study which were at the time classified as possibly wrong. This particular behavior at Re of 500,000 is explained by a change in shape of the downwash, which avoids flow separation as the blade starts to travel upwards.

A LAMINAR FLOW, PROPULSIVE, JET-FLAPPED CONCEPT FOR ELECTRICALLY POWERED TRANSPORT AIRCRAFT

Friction drag constitutes approximately half of the total drag of subsonic civil transport aircraft at cruise conditions. Several means were examined to control the flow over an aircraft and achieve laminar flow. Here, a new concept for friction drag reduction in the form of an integration of the aerodynamics and propulsion of the aircraft is put forward. Engines buried in the wing and at the rear of the fuselage suck the boundary layer of the entire wing and fuselage surface, and then, they used it as intake air and exhaust through ducts. At the wings, the engines exhaust in the form of a jet flap at the trailing edge providing distributed propulsion. By this laminar flow, propulsive concept laminar flow is established over the entire aircraft, resulting in substantial drag reduction. The analysis showed that out of the four electrically powered aircraft versions considered only the combined lift distribution with tailless fuselage is about to be feasible. It was also found that the example aircraft design is inappropriate. It is expected that a design purposely based on the proposed concept would bring electrically powered transport aircraft within the specific energy levels of present batteries.

Numerical Investigation of Particle Focusing Patterns in Laminar Pipe Flow With High Reynolds Numbers

The inertial focusing characteristics of particles in laminar flow pipes with high <i>Re</i> numbers were studied based on the “relative motion model”. In order to solve the problem of long pipes with high <i>Re</i> number flow, periodic boundary conditions were imposed on the inlet and outlet of the pipe. The research results show that the use of periodic boundary conditions can effectively reduce the computational, and the mechanical properties of particles in high <i>Re</i> flow can be calculated by using <i>L</i>=4<i>D</i> pipe. The difference from the low <i>Re</i> number is that as the <i>Re</i> number continues to increase,the lift force of the particles in the radial direction is no longer distributed as a parabola. The lift curve has a concave area between <i>r</i>⁺ =0.5 ~ 0.7, and there is a tendency for a new inertial focus point to appear in this section. By means of particles of <i>a</i>⁺ =1/17 for <i>Re</i> &gt; 1 000, this new focus point position is solvable. In addition, in the analysis of the flow field, a secondary flow occurs around the particle, and its intensity gradually increases with the <i>Re</i> number and the closeness of the particle to the wall. The generation of the secondary flow affects the spatial distribution of the particle lift.

Coordinated Roll Control of Conformal Finless Flying Wing Aircraft

Attitude control of tailless aircraft is typically achieved through a system of trailing-edge flaps that produce coupled control effects in pitch, roll and yaw due to changes in lift and drag distributions. This paper presents a novel control allocation method which allows conformal trailing-edge controls to generate yawing moments with no coupling, thereby simplifying the required control suite. The method is demonstrated through modelling of a coordinated roll manoeuvre of a tailless aircraft in which yaw control is provided by trailing-edge flaps only. For a typical tailless aircraft configuration, it is shown that a coordinated rolling manoeuvre can be achieved with a reduction of at least 30% of the required control deflection compared to solutions using non-conformal controls such as split drag rudders.

Vision-Based Nonlinear Incremental Control for a Morphing Wing With Mechanical Imperfections

Morphing structures have acquired much attention in the aerospace community because they enable an aircraft to actively adapt its shape during flight, leading to fewer emissions and fuel consumption. Researchers have designed, manufactured, and tested a morphing wing named SmartX-Alpha, which can actively alleviate loads while achieving the optimal lift distribution. However, the widely existing mechanical imperfections can degrade the performance of the morphing wing and even lead to instabilities. To tackle these issues, this article proposes a vision-based adaptive control approach to actively compensate for mechanical imperfections. In this approach, an incremental model is constructed online to identify the system dynamics using servo commands and vision measurements, and then, nonlinear dynamic inversion control is applied based on the identified model. This data-driven control approach with visual feedback has been validated by real-world experiments on the SmartX-Alpha. The results demonstrate that the vision-based system combined with the proposed control methodology can actively compensate for mechanical imperfections with minimal adjustments to the actual system design. Compared to a controller that only uses a feedforward input-output mapping, this proposed approach improves the system performance and decreases the tracking errors by more than 62% despite disturbances. The results collectively demonstrate the effectiveness of the proposed control system, which sets a foundation for realizing morphing in next-generation aircraft.

Morphing wing design using integrated and distributed trailing edge morphing

The presented study investigates the design and development of an autonomous morphing wing concept developed in the scope of the SmartX project, which aims to demonstrate in-flight performance optimisation with active morphing. To progress this goal, a novel distributed morphing concept with six translation induced camber morphing trailing edge modules is proposed in this study. The modules are interconnected using elastomeric skin segments to allow seamless variation of local lift distribution along the wingspan. A fluid-structure interaction optimisation tool is developed to produce an optimised laminate design considering the ply orientation, laminate thickness, laminate properties and actuation loads of the module. Analysis of the kinematic model of the integrated actuator system is performed, and a design is achieved, which meets the required continuous load and fulfils both static and dynamic requirements in terms of bandwidth and peak actuator torque with conventional actuators. The morphing design is validated using digital image correlation measurements of the morphing modules. Characterisation of mechanical losses in the actuator mechanism is performed. Out-of-plane deformations in the bottom skin and added stiffness of the elastomer are identified as the impacting factors of the reduced tip deflection.

Aerodynamic Shape Optimisation Using Parametric CAD and Discrete Adjoint

This paper presents an optimisation framework based on an open-source CAD system and CFD solver. In this work, the high-fidelity flow solutions and surface sensitivities are obtained using the primal and discrete adjoint formulations of SU2. This paper shows the direct use of CAD models for optimisation by developing a CAD system application programming interface and creating a link between the CAD-MESH-CFD analysis. A methodology to obtain geometric sensitivities is introduced, enabling the calculation of accurate gradients with respect to CAD variables and the deformation of the analysis mesh during the optimisation process. This methodology guarantees that the new surface mesh lies exactly on the CAD geometry. The optimisation framework is applied to a rectangular wing and a three section high-lift aerofoil configuration derived from the NASA CRM-HL configuration. Both geometries are created using FreeCAD. The performance objectives are to decrease the drag while constraining the lift to be above a desired value. The twist distribution of the wing was parameterised within the CAD system, allowing the minimisation of the induced drag by obtaining a nearly elliptical lift distribution. For the high-lift configuration, the position and rotation of the flap and slat were parameterised with respect to the original section; the final optimised positions yield a drag reduction of approximately 16.5%. These results show that the CAD parameterisation can be reliably used to obtain efficient optimums while operating directly on the CAD geometries.

Modern Implementation and Evaluation of Lifting-Line Theory for Complex Geometries

A numerical lifting-line method (implemented in an open-source software package) is presented that can accurately estimate the aerodynamics of wings with arbitrary sweep, dihedral, and twist. Previous numerical lifting-line methods have suffered from grid convergence challenges and limitations in accurately modeling the effects of sweep, or have relied on empirical relations for swept-wing parameters and have been limited in their application to typical wing geometries. This work presents novel improvements in accuracy, flexibility, and speed for complex geometries over previous methods. In the current work, thin-airfoil theory is used to correct section lift coefficients for sweep, providing a more general closure to the lifting-line problem. A linearized solution is presented, which can be used as a rapid approximation for the full solution, or as an initial guess for the nonlinear system of equations to speed convergence. Sensitivities to model parameters are investigated, and appropriate recommendations for these parameters are given. Agreement with Prandtl’s classical lifting-line method is excellent in the case of straight wings. Comparison with experimental data shows that this method can reasonably predict lift, drag, and lift distribution for a range of wing configurations. The speed and accuracy of this method make it well-suited for preliminary design and optimization.

Bell Shaped Lift Distribution, Aerodynamic Soaring and Control Surface Dynamics of Birds

Abstract: The optimal aerodynamic traits of the natural flyers have always been a foundation for the designing of bionic drones. In this paper, the aerodynamic soaring capabilities of the albatross and eagle are discussed in detail along with the eagle’s control surface aerodynamics. The bell shaped lift distribution characteristics exhibited by the albatross are also discussed along with its subsequent sub traits such as proverse yaw, wing-tip vortices and so on. The main soaring dynamics discussed are the dynamic soaring exhibited by the albatross and the utilization of thermal updrafts for thermal soaring by the eagle. Apart from these, the different control surfaces and their actuations along with wing morphing techniques are also highlighted. A proper and near perfect manifestation of the traits of natural flyers could be a harbinger to near perpetual modern aircrafts. Aerodynamics and propulsion system are the areas of research that are of concern in this manuscript.

Analysis of flow-induced performance change of cycloidal rotors: Influence of pitching kinematic and chord-to-radius ratio

As a new type of the propulsive devices, cycloidal rotor has been attracting more attention recently. The present work concentrates on the analysis of the performance and internal flow field of a two-bladed cycloidal rotor operating at low Reynolds numbers, with special emphasis on understanding how the flow structure affects the performance at different working conditions. The symmetrical/asymmetrical pitching motions are evaluated initially and the primary results show that the asymmetrical pitching with a mean pitching angle of 5° improves the efficiency of the rotating system. Then, the effect of the chord-to-radius ratio at three different conditions is discussed, which shows that the chord-to-radius ratio of 0.45 is the best value to maximise the performance. The flow pattern, involving the blade-wake interaction, wake-wake interaction, three vortex structures on the blade surface, roll-up vortices inside the laminar boundary layer, separation bubble, leading-edge vortex, trailing-edge vortex and flow separation vortex, are discussed in detail under various operating conditions. The study on the performance of a single blade, the forces (lift and drag) distributions, the blade loadings and the near-wall flows on two blades, provide a new understanding of the global performance change of the propeller.

Solution of a Nonplanar Lifting-Line Model Using the Finite Element Method

A reduction in the structural wing weight is generally followed by an increase in the structural flexibility, leading to a significantly deformed wing configuration. A modified Prandtl’s lifting-line model proposed herein aims to approximate the spanwise lift distribution and the induced drag using a finite element method approach. A numerical validation considering planar elliptical wings is carried out for several mesh configurations. The proposed approach has an optimal agreement to the analytical solution, showing a global error inferior to in terms of the spanwise circulation distribution. Next, analyses of nonplanar wings are made compared to the available literature data. For the first case, the relative difference regarding the lift coefficient is less than 0.3%, whereas for the induced drag, it is less than 2.2%. For the second case, a maximum difference for the lift coefficient is less than 0.5%. The numerical strategy proposed in this paper approximates the solution of the aerodynamic problem to the approach via finite element methods used in structural analysis. The presented approach can later lead to a hybrid finite element with aerodynamic and structural degrees of freedom, which can be helpful for the development of a monolithic approach.

Simplified vortex methods to model wake vortex roll-up in real-time simulations for fuel-saving formation flight

One possibility for reducing fuel consumption is to fly in the upwind field of the wake vortex generated by an aircraft that is flying ahead. Migratory birds use this principle. Manually flying an aircraft at the point of optimal fuel reduction is not suited for routine flight operations as the pilot workload is excessively high. Hence, an autopilot function has to carry out this task. For designing the autopilot, a flight mechanical simulation with a wake vortex velocity model is required that has the ability to calculate the vortex-induced velocity fields. This paper contributes to the choice of a real-time simulation method for modelling vortex-induced velocities that the wake vortex of a leading aircraft generates and that the trailing aircraft shall use during fuel-saving formation flight.Two different wake vortex velocity models are introduced and compared during steady, horizontal flight. One model is based on the Lifting Line Method (LLM) and the other on the unsteady Vortex Lattice Method (VLM). Both models are able to calculate the wake vortex roll-up phase for arbitrary lift distributions, whereas the commonly used Single Horseshoe Vortex Model (SHVM) ignores the near-field roll up. The differences in the induced upwind distribution and vortex filament position are analysed for coarse spatial and temporal discretisation that the real-time constraint requires. Despite the more stringent simplifications of LLM, both methods yield similar filament positions and similar velocity fields for the same discretisation of the lifting surfaces. Finally, the influence of the discretisation parameters is discussed and parameter values are recommended for using VLM and LLM in real-time flight simulations.

Numerical investigation of the aerodynamic interference in 2-aircraft formation flight

Due to obvious reduction in drag, formation flight attracts wide attention of major powers in the world. Based on wild geese flying in a ‘V’ formation, a formation composed of two flying-wing aircrafts is designed. The aerodynamic interference in 2-aircraft formation flight is studied by CFD. Results show that 2-aircraft formation flight has little impact on leading aircraft, but has an evident impact on following aircraft. At the same AoA, the effect of increment in lift and reduction in drag of following aircraft is large. As considering the balance of gravity and lift, the reduction in drag of following aircraft is much larger, that is up to 25.1%. Under the aerodynamic interference effect, the following aircraft has a nosedown pitching moment increment, and becomes lateral static instability, which is opposite to the characteristic of solo flying-wing aircraft. The aerodynamic interference mainly comes from upwash flow induced by wingtip vortex of leading aircraft, which reduces the actual AoA and induced drag of following aircraft, and changes the lift distribution of both sides of following aircraft.

Numerical Assessment of a BAY-Type Model for different Vortex Generator Shapes Applied on a Wind Turbine Airfoil

Modern wind turbines are increasingly equipped with vortex generators in the inner part of the rotor blade. On large blades it is usual to install hundreds of these small passive devices to delay flow separation. Therefore, affordable numerical methods to include vortex generators in the design process are required. One very promising method based on the addition of source terms in the flow equations is the BAY model. In this work a BAY-type model is presented and assessed for three different vortex generator shapes (rectangular, delta and cropped-delta) placed on a DU 97-W-300 airfoil and computed with steady-state Reynolds-averaged Navier-Stokes numerical methods. In order to quantify the circulation of the main streamwise vortex, a method to identify the vortex core based on the λ2-criterion is presented and shows good results. The computations reveal very good agreement with the experimental data for attached flow in terms of lift and pressure distribution while drag is inherently underestimated by the BAY model due to the inviscid modelling of the vortex generators. It is displayed that this underestimation is proportional to the vortex generator surface area. Concerning the different shapes, it is shown that the delta shape reveals larger deviations from the fully resolved case than the other geometries and therefore requires higher mesh resolution.

Aerodynamic Analysis and Optimization of Wings for the Jain University Sailplane Using XFLR5

The wing of a sailplane must be designed to conform to specified mission and design requirements. High aerodynamic efficiency and good stall characteristics are some of the design requirements, while achieving level flight at a particular altitude could be a mission requirement. The aerospace engineering students of Jain University set out the task of designing a sailplane capable of flying at an altitude of 7,000 feet. The design task was undertaken using the computational code XFLR5. In the first step, the aerodynamic characteristics of five different airfoils are analyzed considering the effect of altitude. Secondly, the effect of chord length and span on rectangular wing aerodynamics is investigated. The design convergence is obtained for a wing of chord length 1 m and span 20 m. The effect of wing twist on the initially converged design to obtain a more elliptical lift distribution is then performed to obtain optimized designs.