Wing Section Research Articles

Experimental Flowfield Analyses of Installed Subsonic Jet–Pylon Interaction

This paper presents an experimental investigation into the flowfield of single-stream jets blocked by internal and external rigid nozzle structures. The jets studied include axisymmetric round and annular jets and asymmetric blocked pylon jets. The pylon nozzles tested have three different thicknesses, reducing the exit flow area by 5, 10, and 20%. Hot-wire anemometry is used to record the unsteady velocity field both at the nozzle exit and up to 10 jet diameters downstream. Installation effects, created by mounting a wing section close to the jet, and forward-flight effects are also considered. The nominal Mach numbers studied are and 0.6 for the jet flow, and , 0.1, and 0.2 for the flight stream. Results show that the nozzle blockages 1) redirect the jet toward the wing, 2) increase the entrainment rate in locations far downstream of the nozzle exit, 3) produce dominant effects compared to the change of entrainment induced by the presence of the wing, and 4) considerably increase the normalized turbulence intensity generated in in-flight conditions.

Techniques for Turbulence Tripping of Boundary Layers in RANS Simulations

The exact placement of the laminar–turbulent transition has a significant effect on relevant characteristics of the boundary layer and aerodynamics, such as drag, heat transfer and flow separation on e.g. wings and turbine blades. Tripping, which fixes the transition position, has been a valuable aid to wind-tunnel testing during the past 70 years, because it makes the transition independent of the local condition of the free-stream. Tripping helps to obey flow similarity for scaled models and serves as a passive control mechanism. Fundamental fluid-mechanics studies and many engineering developments are based on tripped cases. Therefore, it is essential for computational fluid dynamics (CFD) simulations to replicate the same forced transition, in spite of the advanced improvements in transition modelling. In the last decade, both direct numerical simulation (DNS) and large-eddy simulations (LES) include tripping methods in an effort to avoid the need for modeling the complex mechanisms associated with the natural transition process, which we would like to bring over to Reynolds-averaged Navier–Stokes (RANS) turbulence models. This paper investigates the implementation and performance of such a technique in RANS and specifically in the k-omega SST model. This study assesses RANS tripping with three alternatives: First, a recent approach of turbulence generation, denoted as turbulence-injection method (kI), is evaluated and investigated through different test cases; second, a predefined transition point is used in a traditional transition model (denoted as IM method); and third a novel formulation combining the two previous methods is proposed, denoted gamma -kI. The model is compared with DNS, LES and experimental data in a variety of test cases ranging from a turbulent boundary layer on a flat plate to the three-dimensional (3D) flow over a wing section. The desired tripping is achieved at the target location and the simulation results compare very well with the reference results. With the application of the novel model, the challenging transition region can be excluded from a simulation, and consequently more reliable results can be provided.

Stability Analysis and Augmentation Design of a Bionic Multi-Section Variable-Sweep-Wing UAV Based on the Centroid Self-Trim Compensation Morphing

In this study, a design scheme for a high-aspect-ratio bionic multi-section variable-sweep wing unmanned aerial vehicle (UAV) that utilizes the reverse coordinated change in the sweep angle of the inner and outer wing sections is proposed, which improves the aerodynamic performance and realizes the self-trim compensation of the wing’s centroid. According to the layout characteristics of this type of UAV, a reasonable distribution design of the wingspan ratio of the inner and outer sections is explored, to reduce the impact of aerodynamic center movement and moment of inertia change. The calculation and analysis results show that the coordinated variable-sweep scheme can significantly improve the influence of sweep angle change on the longitudinal static stability margin of UAVs with a high aspect ratio. The coordinated sweep angle change in the inner and outer wing sections can not only reduce the drag during high-speed flight, but also play a significant role in improving the performance of the aircraft at different stages in the mission profile. Appropriately increasing the wingspan proportion of the inner section can reduce the trim resistance of the V-tail, reduce the thrust of the engine, and increase the range and duration of the UAV. From the perspective of stability change, the multi-section variable-sweep wing UAV with a wingspan ratio of the inner and outer sections that is between 1.41 and 1.78 has better dynamic stability performance. Among them, the UAV with a wingspan ratio of the inner and outer sections that is equal to 1.41 has better longitudinal stability performance, while the UAV with a wingspan ratio of the inner and outer sections that is equal to 1.78 has better lateral/directional stability performance.

Imaging of Organic Samples with Megaelectron Volt Time-of-Flight Secondary Ion Mass Spectrometry Capillary Microprobe.

Time-of-flight Secondary Ion Mass Spectrometry (TOF SIMS) with MeV primary ions offers a fine balance between secondary ion yield for molecules in the mass range from 100 to 1000 Da and beam spot size, both of which are critical for imaging applications of organic samples. Using conically shaped glass capillaries with an exit diameter of a few micrometers, a high energy heavy primary beam can be collimated to less than 10 μm. In this work, imaging capabilities of such a setup are presented for some organic samples (leucine-evaporated mesh, fly wing section, ink deposited on paper). Lateral resolution measurement and molecular distributions of selected mass peaks are shown. The negative influence of the beam halo, an unavoidable characteristic of primary beam collimation with a conical capillary, is also discussed. A new start trigger for TOF measurements based on the detection of secondary electrons released by the primary ion is presented. This method is applicable for a continuous primary ion beam, and for thick targets that are not transparent to the primary ion beam. The solution preserves the good mass resolution of the thin target setup, where the detection of primary ions with a PIN diode is used for a start trigger, reduces the background, and enables a wide range of samples to be analyzed.

Study of the Aerodynamic Characteristics on the ‎Computed Flowfield During Thrust Reversers Operation

With the increasing of the bypass ratio of modern aero-engines, the problem of wing-engine interference is more prominent, especially for the layout design of engine nacelle. The reverse thrust cascade is widely used in turbofan engines with high bypass ratio. In order to meet the requirements of the integrated analysis of airframe, wing, engine nacelle and hanger, the aerodynamic characteristics of aircraft/engine integration configuration with reverse cascade is numerically studied via CFD (Computational Fluid Dynamics) method. The streamline distribution, iso-surface of total temperature, vorticity distribution and total reverse thrust efficiency on different engine nacelle layouts are compared and analyzed in details. The results show that the lift coefficient of the wing decreases by 21.2% and 45.02% respectively as the engine moves forwards horizontally by 11%L and 21.2%L. The lift coefficient of the wing decreases by 2.4% and 4.82% respectively as the engine moves subsidence by -3.5%L and 3.5%L. The influential region of the reverser airflow in the radial and circumferential direction without airframe interference is significantly larger than that in the case with aircraft/engine integration. The reverser flow is susceptible to be interfered by the adjacent fuselage and wing sections, and the development of reverse thrust flow is significantly limited. Compared to the baseline nacelle location, the reverse thrust performs badly as engine nacelle moves backward and lateral horizontally. The total reverse thrust efficiency decreases gradually as the engine nacelle moves forward horizontally.

Wing Flutter Suppression under the Influence of Fuel Sloshing in External Tank

The control of aeroelastic response of a wing section with influence of the fuel sloshing in external tank is examined through analytical studies. Based on the simplified mass-spring-damper equivalent mechanical model of the liquid sloshing and the dynamic modeling method of the wing/store system, the aeroelastic model of the two-dimensional wing/external tank/fuel sloshing system is established in this paper. The aeroelastic instability of the system is studied, and a flutter suppression controller is designed based on linear active disturbance rejection control (LADRC). The stability analysis shows that the tracking error and the estimation error of the linear state observer (ESO) are convergent. It is shown through simulation results that wing flutter suppression under the interference of fuel sloshing and external tank vibrating can be achieved by using two control surfaces.

Topology Optimization of Large-Scale 3D Morphing Wing Structures

This work proposes a systematic topology optimization approach for simultaneously designing the morphing functionality and actuation in three-dimensional wing structures. The actuation was modeled by a linear-strain-based expansion in the actuation material. A three-phase material model was employed to represent structural and actuating materials and voids. To ensure both structural stiffness with respect to aerodynamic loading and morphing capabilities, the optimization problem was formulated to minimize structural compliance, while the morphing functionality was enforced by constraining a morphing error between the actual and target wing shape. Moreover, a feature-mapping approach was utilized to constrain and simplify the actuator geometries. A trailing edge wing section was designed to validate the proposed optimization approach. Numerical results demonstrated that three-dimensional optimized wing sections utilize a more advanced structural layout to enhance structural performance while keeping the morphing functionality better than two-dimensional wing ribs. The work presents the first step towards the systematic design of three-dimensional morphing wing sections.

Stall Cell Prediction Using a Lifting-Surface Model

This paper investigates the use of a lifting surface method coupled with a nonlinear lift curve, the nonlinear vortex lattice method (NL-VLM), toward the study of stall cells on wings. First, Spalart’s spectral lifting line model is thoroughly investigated. It is found that the Gaussian filter leads to the existence of two different solutions, one directly interpolated in the lift polar and the other retrieved from the circulation distribution. Then the NL-VLM method is applied to the canonical cases of the infinite and the elliptic wing. In this model, the stall cells is represented as a sharp transition between stalled and prestalled wing section, meaning that no wing section effectively sees a negative slope of the lift against angle-of-attack relation. This sharp transition can be smoothed by introducing an artificial dissipation that also results in two separated solutions, without changing the underlying characteristics of the model. The effect of several parameters, like the grid spacing and the relaxation factor, is studied, with little effect on the wavelength of the predicted stall cells. The effect of specific parameters of the lift polar like the magnitude of the negative slope in the stall regime and the trend in the deep stall range is also investigated. It is observed that the slope of the lift versus angle of attack must not stay negative when the angle of attack goes to infinity for the system of equations to have a steady solution. The amplitude of the negative lift-curve slope has an effect on the wavelength of the cells explaining why the stall cells would not be observed for a leading-edge-type stall. Finally rectangular wings are studied for comparison to experimental data and higher-fidelity numerical simulations, showing the predictive capability of the proposed low-order model.

Numerical study on the influence of wall temperature gradient on aerodynamic characteristics of low aspect ratio flying wing configuration

With the aim for a low-aspect-ratio flying wing configuration, this study explores the influence of wall temperature gradient on the laminar and turbulent boundary layers of aircraft surface and determines the effect on the transition Reynolds number and wall friction drag. A four-equation turbulence model with transition mode is used to numerically simulate the flow around the model. The variation of wall friction coefficient, transition Reynolds number, and turbulent boundary layer flow with wall temperature are emphatically investigated. Results show that when the wall temperature increases from 288 to 500 K, the boundary layer transition Reynolds number for the wing section increased by approximately 28% and the surface friction drags decreases by approximately 10.7%. The hot wall enhances the viscous effects of the laminar temperature boundary layer, reduces the Reynolds shear stress and turbulent kinetic energy, and increases the flow stability. However, the velocity gradient and shear stress in the bottom of the turbulent boundary layer decreases, which leads to reduced friction shear stress on the wall surface. Therefore, for the low-aspect-ratio flying wing model, the hot wall can delay the boundary layer transition and reduce the friction drag coefficient in the turbulent region.

Use of the parachute system for the collective rescue of passengers and aircraft crew

This study covers the issues of the security system in civil aviation. The relevance of the topic is conditioned by a qualitatively new approach to the design of civil aviation, since the modern equipment has the means intended for rescue directly in the area of contact of an emergency aircraft with the ground. When hitting the ground, as a rule, the lower part of the airframe is crushed, the fuselages break in the centre wing section, and the interior equipment breaking off is very traumatic to passengers, the engines are extremely fire and explosion dangerous, which leads to an emergency situation developing into a catastrophic one. The purpose of the study is to investigate the existing security system of a civil aircraft and develop the modification of aviation equipment with the help of collective rescue systems. Two main areas of application of collective rescue systems in aviation technology are analysed. As a solution to the problem at hand, a modification in the form of a parachute system of collective rescue is proposed. A detailed description of the system operation with the use of the necessary calculations for weight, aerodynamic and similar costs is given. The system under study relates to aviation technology and can be used in the development of transport (passenger, cargo, business class, etc.) aircraft. The proposed modification would significantly increase the level of flight safety due to a preventive measure that would exclude the contact area of the emergency aircraft with the ground.

Trailing-edge noise reduction of a wing by a surface modification

There is an increased emphasis on reducing airframe noise in the last decades. Airframe noise is sound generated by the interaction of a turbulent flow with the aircraft geometry, and significantly contributes to the overall noise production during the landing phase. One examples of airframe noise is the noise generated at a wing's trailing edge, i.e., trailing-edge noise. In this contribution, we numerically explore the local application of riblets for the purpose of trailing-edge noise reduction. Two configurations are studied: i) a clean NACA0012 wing section as a reference, and ii) the same configuration with riblets installed at the wing's aft part. The numerical investigation follows a hybrid computational aeroacoustics approach, where the time-average flow is studied by means of RANS. Noise sources are generated by means of a stochastic approach called Fast Random Particle Mesh method. The results show a deceleration of the flow behind the riblets. Furthermore, the turbulent kinetic energy indicates increased unsteadiness behind the riblets which is shifted away from the wall due to the presence of the riblets. Lastly, the sound sources are investigated by means of the 3D Lamb-vector, which indicates a slight reduction in magnitude near the trailing edge.

Dynamic flight load measurements with MEMS pressure sensors

The determination of structural loads plays an important role in the certification process of new aircraft. Strain gauges are usually used to measure and monitor the structural loads encountered during the flight test program. However, a time-consuming wiring and calibration process is required to determine the forces and moments from the measured strains. Sensors based on MEMS provide an alternative way to determine loads from the measured aerodynamic pressure distribution around the structural component. Flight tests were performed with a research glider aircraft to investigate the flight loads determined with the strain based and the pressure based measurement technology. A wing glove equipped with 64 MEMS pressure sensors was developed for measuring the pressure distribution around a selected wing section. The wing shear force determined with both load determination methods were compared to each other. Several flight maneuvers with varying loads were performed during the flight test program. This paper concentrates on the evaluation of dynamic flight maneuvers including Stalls and Pull-Up Push-Over maneuvers. The effects of changes in the aerodynamic flow characteristics during the maneuver could be detected directly with the pressure sensors based on MEMS. Time histories of the measured pressure distributions and the wing shear forces are presented and discussed.

Polycarbonate based multifunctional self-sensing 2D and 3D printed structures for aeronautic applications

Different polycarbonate materials have been reinforced with carbon nanotubes to tune electrical conductivity and to induce piezoresistive self-sensing capabilities. Further, the composites were processed by solvent casting and fused deposition modelling (filament for 3D printing) methods. An electrical conductivity percolation threshold of ≈0.3 wt.% has been found for solvent casted films, whereas samples processed by 3D printing show higher thresholds, ≈2 wt.%, presenting all samples a similar maximum electrical conductivity (σ ≈ 1 × 10−3 S m−1), thermal and chemical properties. Overall mechanical properties are larger for the solvent cast films concerning the 3D printed ones, particularly the elongation at break. The piezoresistive sensitivity, obtained after four-point-bending and uniaxial strain experiments, shows gauge factors up to 1.7, independently of the processing method. The functionality of the materials has been demonstrated by the implementation of an airplane wing section model with self-sensing capabilities. Two implemented strategies showed the suitability of the developed materials for real-time monitoring of the wing mechanical deformation.

A BRIEF REVIEW ON VARIOUS NOISE REDUCTION TECHNIQUES EMANATING FROM WINGS OF AN AIRCRAFT

— One of the challenges and a point of need in the aviation industry is the problem of high level of noise emanation. There are various sources of noise emanating from aircraft which are the noises from engine and airframe including the wings. Noise from the wing is an important phenomenon to be put into consideration as it contributes significantly to the noise emanating from aircraft especially during take-off and landing. Airfoil noise emission emanate on the wing section of an aircraft due to the interaction of the airfoil turbulent boundary layer with the sharp leading edge. Noise reduction then becomes a thing of concern for many engineers and researchers who took it upon their shoulder the responsibility of coming up with design that can help reduce noise emanating from aircrafts to mitigate the noise. Due to this, many researchers have focused on reducing the noise emanating from the aircraft through various means and approaches which include theoretical, computational, and experimental analysis. In a means of compiling and making available a brief review of such work, for easy and quick access necessitate the need of this work. This paper focused on giving a summary and review of different work done by past and current researchers in the field. The review was based on selected current papers on the cutting-edge technology towards enhancing aerodynamic performance of the wing of aircraft and reduction of noise emanating from it. Although fruitful results have been achieved in the aerodynamic noise control, but only little work has been done to summarize the main findings and progresses in this area. This work focused on review of work carried out by many researchers in this area which involves various techniques and attempts ranging from leading edge serration, trailing edge serration, active flow actuation, flexible wing using theoretical, computational, and experimental approach etc. Additionally, some suggestions were made towards further study and improv

Revisit of the variable stiffness method for aeroelastic computations with/without thermal effects based on computational fluid dynamics

In the present study, a variable stiffness method (VSM) is revisited for aeroelastic computations with/without thermal effects. Unlike the traditional CFD/CSM coupling method (TM), which predicts aeroelastic responses by varying freestream conditions (e.g. freestream density, velocity), the freestream conditions in VSM can be fixed. The VSM is first verified theoretically and adopted to predict nonlinear aeroelastic responses. For aeroelastic computations, the method is applied to predict flutter onset of Isogai wing section and limit cycle oscillation (LCO) of Goland+ wing at transonic conditions. The structural free-play nonlinearity is included in the Isogai wing section aeroelastic system to further demonstrate the method. The aeroelastic computation with thermal effects is considered as the flutter onset prediction of a simply supported panel in supersonic flow. It is shown that the VSM method can replicate the nonlinear aeroelastic responses predicted by its traditional CFD/CSM coupling method counterpart under the same flow similarity parameters (e.g. Mach number, Reynolds number), whereby the freestream conditions need to be adjusted. Limitations of the VSM are also pointed out and discussed. To further assess the VSM, an ARMA model is constructed under the framework of the VSM and demonstrated by flutter onset prediction of the Isogai wing section at transonic regime.

Dynamic response of the wing during UAV collision

An airborne collision between civil Unmanned Aerial Vehicles (UAVs) and manned aircraft is a potential threat to aircraft operation safety. In this paper, the UAV collision test was performed on a commuter aircraft wing section to investigate the dynamic response of this primary operation component. The projectile was a 735 g drone named ‘Mavic’. Explicit dynamic code ABAQUS was employed to simulate the collision process based on the difference of collision scenarios to assess the hazard. The results showed that 735 g drone impact at the aircraft maximum approach flap and cruising speed could cause some damage on the wing front spar and the situation is more serious than 910 g bird strike in which the hardness of drone components rather than kinetic energy is a decisive factor. The lithium-ion battery penetrated the airframe which may be a potential source of ignition.

АНАЛІЗ НАПРУЖЕНО-ДЕФОРМОВАНОГО СТАНУ ФІТИНГОВОГО СТИКУ ПАНЕЛЕЙ ВІД'ЄМНОЇ ЧАСТИНИ КРИЛА ТА ЦЕНТРОПЛАНА ТРАНСПОРТНОГО ЛІТАКА

Fitting joints are widely used in aircraft structures, and they are responsible for the interconnection of important components. The stress-strain state analysis of the fitting joint must be carried out before the performance analysis of the fitting joint. With the help of 3D modeling software (CATIA) and finite element analysis software (ANSYS), the stress-strain state of each component in the fitting joint of outer wing section was calculated in this paper. In the CATIA, the solid model is simplified and segmented according to the size of the cross section and the height of the center of gravity of the model. In the ANSYS, the beam elements are used to replace the simplified segmented model to obtain the internal force distribution of the solid model and to determine the magnitude and change law of the stress applied to the end of the solid model. When calculating the force transmitted by the fastener, the pre-tightening force of the bolt and the interaction between the surfaces of the component are taken into account, so as to simulate the real force situation well. Therefore, it is a very feasible method to use the CATIA and ANSYS to obtain the stress-strain state of components in the fitting joint of center wing section and outer wing section.The results show that under the working conditions of the fitting joint (130Mpa), the fitting of outer wing section with center section has a maximum stress of 245.79Mpa and a maximum strain of 0.0035, the stringer of outer wing section has a maximum stress of 293.17Mpa and a maximum strain of 0.0047, the lower panel of outer wing section has a maximum stress of 289.53Mpa and a maximum strain of 0.0042. The connecting bolts (M8 and M6) have a maximum stress of 686.81Mpa and a maximum strain of 0.0063, which meets the design requirements. In addition, according to the analysis results of the stress-strain state of the fitting joint of outer wing section, the force distribution of the bolts in the fitting joint of outer wing section with center section was obtained in this paper. It has been confirmed that due to the different positions and force areas of the bolts, the force distribution between rows of bolts is uneven, and the first row of bolts has a more force.

A new simplified numerical approach for shape design of underwater wings

ABSTRACT This paper presents a new simplified numerical approach (SNA) for predicting hydrodynamic data of underwater wings (UWs). Instead of full three-dimensional (3D) flow field analyses, SNA combines Vortex Lattice Method (VLM) and several two-dimensional Reynolds-Average Navier-Stokes (2D RANS) simulations of wing sections to gain results. Hence, SNA shows a remarkably lower computational cost than higher-fidelity methods such as 3D RANS. Moreover, SNA shows good agreements with 3D RANS and experimental data. Thereafter, SNA is applied to shape optimizations of UWs for efficiency and accuracy. Two kinds of optimisation algorithms that include local and global finders are adopted in the optimiser. Searching behaviours of the two algorithms indicate that the problem is not proper to be solved by the local optimiser for its multimodality, while the global optimiser is more efficient. Furthermore, the final optimised wing has few wing-tip vortices and a smooth pressure distribution.

Enhancement of airfoil performance by rotating cylinder

This study used a NACA 6409 smooth wing section. Its performance is compared to an enhanced design incorporating a revolving cylinder on the upper airfoil surface. The boundary conditions for the present work were Reynolds (105), and the attack angles were (0,2,4,6,8,10,12,14,16). A turbulence model was developed (SST K-), one of the aerodynamics models. This model is well-suited for sensing flow near a wall. The cylinder was evaluated at three positions (25,50,75) percent of the chord's length. The results indicated that position (25 percent C) was optimal, as it resulted in a 24.45 percent increase in lift coefficient at the angle of attack (12). And when the results were compared to those of other studies, they revealed a significant agreement.

A repairable carbon nanotube web-based electro-thermal heater and damage sensor for aerospace applications

ABSTRACTWe previously described an efficient, lightweight and flexible electro-thermal system, based on directly drawn carbon nanotube web (CNT web), as part of an icing protection system for carbon fibre reinforced polymer (CFRP) composite aircraft structures. The location of the heating elements on critical lifting surface leading edges or nacelle intake lips makes them particularly susceptible to impact damage, which may leave no visible mark. This makes it desirable to have both a mechanism for identifying the location of damage to the CNT structure (and by inference, potential damage to the underlying CFRP) and a process for restoring the CNT heater to full operation. With the CNT web acting as a sensor, impact damage is identified by an increase in electrical resistance and, particularly, by infrared imaging, which reveals a cold spot or zone depending upon the CNT web layup. Whereas a unidirectional CNT web layup exhibits a large increase in resistance and loss of a full width band of operation, a cross ply quasi-isotropic CNT web arrangement suffers only a small increase in resistance and a loss of function that is highly localised to the damaged area. A novel methodology, based on dispersed CNT in resin, is described for repairing and reconnecting the CNT structure and restoring functionality. A CNT web-based electro-thermal element was applied to the leading edge of a representative carbon-fibre composite wing section to demonstrate the flexibility of this system.