Wing Model Research Articles

Sweep Angles Influence on the Aerodynamic Characteristics of NACA 2412 Wing with Supersonic Flow

A sweep process of a wing is very useful technique to enhance an aircraft performance especially the aircraft which fly in the sonic and supersonic areas. In this research, the influence of swept wing on aerodynamic characteristics generation of aircraft at supersonic speeds has been studied and analyzed using CFD technique. The 3D wing model analyzed and presented in this paper of NACA 2412 profile. The governing equations such as continuity, Energy Equation, and Reynolds Averaging Never-Stokes were solved in conjunction with an analysis of dynamic meshing technique to predict the aerodynamic characteristics and pressure distribution for a wing with a sweep angle ranging from 0° to 60° and a wide range of angle of attack (AOA). In conclusion, depending on the analysis result, increasing of sweep angle can increase the ratio of lift to drag coefficient by about 15%. Moreover, the sweep angle tends to change the load distribution on wing span. Generally, the swept wing has good performance coefficient through the designing an aircraft envelope by optimizing the sweep to get an optimum value of lift/drag ratio. The efficacy and authenticity of the CFD simulations model are shown by comparing the aerodynamic properties findings from the most recent wind tunnel test.

Optimization Efficiency of the Aircraft Wing of Cessna 172 Skyhawk by Absorbent Adverse Pressure Using Tangential Suction Slot Without Vacuum Device

Efficiency of a wing is sufficient at stall angle of attack and breakdown by effecting turbulent flow intensity on upper wing surface which is destroyed the airflow pattern and generated adverse pressure, backflow airstream and forward transition separation point movement closely to the leading edge of the wing. To enhancement a wing efficiency factor at high angle of attacks in this research is using partially sucking adverse pressure to delay the separation point to control the backflow rate and expelling it through the wing tips ends using tangential suction slot channel without using any mechanical means depending on differential static pressure between adverse pressure and the static pressure generated on both win tips ends. Numerical CFD analysis has been managed the case study of the effect suction slot channel located on the upper wing surface at (70%) of the wing chord from the leading edge as a proposed modified wing model of the Cessna 172 Skyhawk half wing at angle of attacks (10°, 12°, 14°, 16°, 18°, 20°, 22°, 24°). The analysis study was conducted at stall speed (90 km/hr.) without flaps where it is equal (25 m/sec). The results of the analytical shows at the range of angle of attacks were chosen between (10° to 24°), lift coefficient CL increased (3.757%), drag coefficient CD is decreased (0.530%) and the wing efficiency lift to drag ratio Δ(CL/CD) is increased to (5.712%) while Δ(CD/CL) is decreased to (5.231%) which these parameters reflected to enhancement the wing to reduce required power at the minimum speed, increase angle of climb and decrease angle of descent reduce landing distance with high angle of attacks.

Aeroelastic Shape Control Using Fiber-Optic-Measured Strain Data and Multiple Control Surfaces

The paper presents an experimental study of aeroelastic shape sensing and control based on strain data measured in fiber-optic sensors (FOSs). Past studies demonstrated how FOS strain data could be used to reconstruct the static and dynamic deformed shapes of a flexible wing. The current study further develops these capabilities and demonstrates wing shape control. Specifically, FOS-measured strain data are used in an optimization scheme, with a target of keeping the wing’s elastic deformations small, below a user-defined threshold, while maintaining a constant lift value (as required for trimmed flight) and with minimal usage of the control surfaces. The method is demonstrated computationally and experimentally on a wing model with four control surfaces. The wing was designed for this study, fabricated using additive manufacturing, and wind-tunnel tested. The trim optimization is performed both computationally and experimentally, following the same procedures but using strain and strain-modes data from finite element analysis and experiments. The wind-tunnel study demonstrated how a flexible wing’s shape could be controlled based on experimental strain data only, without resorting to a computational model. In both the computational and experimental parts of the study, considerable control-surface deformations were required. This primarily indicates that large control surfaces are not the optimal means for wing shape control. However, the same strain-based technique can be used with other effectors, such as distributed control surfaces, strain actuators, etcetera. The study opens a path to using FOSs to optimize winged unmanned-aerial-vehicle performance, and thus design lighter and more efficient platforms.

Flutter Tests of the Pazy Wing

The paper presents the flutter and limit-cycle oscillation (LCO) tests of the Pazy wing, a very flexible wing model designed to study aeroelastic phenomena associated with nonlinear geometry and provide data for nonlinear aeroelastic model validation. The Pazy wing model, analyses, and data from earlier aeroelastic tests are provided by Avin et al. (“Experimental Aeroelastic Benchmark of a Very Flexible Wing,” AIAA Journal, 2022, pp. 1–24). In the current study, three velocity sweeps were performed at different angles of attack (AoA), during which the wing entered and exited limit-cycle oscillations (LCO). A motion-recovery camera system and fiber-optics Bragg grating sensors were used to track the deformations and strains over the wingspan. The wing encountered LCO when the wingtip deformation was approximately 25% of the span. The onset velocity, frequency, and mode shape varied, depending on the test’s AoA. The paper first focuses on the onset of the oscillations, where the onset velocity and oscillation mode-shapes are analyzed. The paper then continues to investigate the LCO characteristics and the nonlinear behavior of the wing. The wing models and experimental data are publicly available through the Third Aeroelastic Prediction Workshop.**

The Drag of an Elliptical Airfoil at a Reynolds Number of 1000

There have been many studies on the mechanisms of unsteady aerodynamics, such as leading-edge vortex (LEV) formation, wing-wake interaction, and spanwise flow. Spanwise flow can only be observed on three-dimensional wing models; however other phenomena such as LEV and wing-wake interaction can be captured using two-dimensional airfoil models. This study focuses on two-dimensional elliptical airfoil because this profile can generate counter-rotating vortices used by insects to generate aerodynamic forces. This research aims to analyze the drag production of two-dimensional elliptical airfoils flapping with bumblebee-inspired kinematics in asymmetrical normal-hovering mode at a typical Reynolds number range of . It is found that drag is generated during the downstroke while thrust during the upstroke. It is also found that the creation and shedding of counter-rotating vortices are closely related to the generation of thrust. The results also indicate that asymmetrical strokes can be used in normal hovering to minimize drag or produce thrust.

Control of flow separation over a wing model with plasma synthetic jets

An array of 30 plasma synthetic jet actuators (PSJAs) is deployed using a modified multichannel discharge circuit to suppress the flow separation over a straight-wing model. The lift and drag of the wing model are measured by a force balance, and the velocity fields over the suction surface are captured by a particle imaging velocimetry system. Results show that the flow separation of the straight wing originates from the middle of the model and expands towards the wingtips as the angle of attack increases. The flow separation can be suppressed effectively by the PSJAs array. The best flow control effect is achieved at a dimensionless discharge frequency of F + = 1, with the peak lift coefficient increased by 10.5% and the stall angle postponed by 2°. To further optimize the power consumption of the PSJAs, the influence of the density of PSJAs on the flow control effect is investigated. A threshold of the density exits (with the spanwise spacing of PSJAs being 0.2 times of the chord length in the current research), below which the flow control effect starts to deteriorate remarkably. In addition, for comparison purposes, a dielectric barrier discharge (DBD) plasma actuator is installed at the same location of the PSJAs. At the same power consumption, 4.9% increase of the peak lift coefficient is achieved by DBD, while that achieved by PSJAs reaches 5.6%.

A Taguchi based iterative wing structural design for a low speed, hybrid UAV

An iterative structural design process for the conventional, composite wing of a low speed, hybrid Unmanned Aerial Vehicle (UAV) is presented in this paper. The relevant design goals are light weight, strength and stiffness derived from the customer specifications, conceptual design and STANAG-4671 airworthiness standards. To achieve the design goals, a modified Taguchi model was applied in conducting iterative Finite Element Analysis of the loads and stresses on the wing model, using ABAQUS CAE. In this analysis, the pitch of the rib and thicknesses of the spar and skin were applied as control factors in three levels, leading to an L9 orthogonal array. Mass, maximum deflection and Tsai-Hill failure index of the wing structure were measured as responses. The result shows that varying the skin thickness had the most impact on the wing mass, failure index and maximum deflection. The design goal of wing mass- less than 2.5kg, deflection of 10% and Tsai-Hill failure index value- less than 1 were achieved after 9 iterations.

An efficient reduced-order framework for active/passive hybrid flutter suppression

AbstractFlutter suppression is an important measure to improve fatigue life and enhance the performance of aircraft in modern aircraft design. In order to design more effective controllers for flutter suppression with high efficiency, an efficient reduced-order framework for active/passive hybrid flutter suppression is proposed. The traditional CFD-based ROMs have been successfully applied to active flutter suppression with high accuracy and efficiency. But, when a structure modification is made such as in aeroelastic tailoring and aeroelastic structural optimisation, the structural model should be updated, and the expensive, time-consuming CFD-based ROMs have to be reconstructed; such a process is impractical for passive flutter suppression. To overcome the realistic challenge, an efficient reduced-order framework for active/passive hybrid flutter suppression is proposed by extending an efficient aeroelastic CFD-based POD/ROM which we have developed. The proposed framework is demonstrated and evaluated using an improved AGARD 445.6 wing model. The results show that the proposed framework can accurately predict the aeroelastic response for active/passive hybrid flutter suppression with high efficiency. It provides a powerful tool for active/passive hybrid flutter suppression, and therefore, is ideally suited to design more effective controllers, and may have the potential to reduce the overall cost of aircraft design.

Predicting pressure coefficients of wing surface based on the transfer of spatial dependency

Multi-conditional holographic pressure coefficients over a wing are crucial for wing design, and a wind tunnel test is an indispensable means to obtain this profile. However, it is resource-consuming to obtain wind tunnel data under different conditions and only a limited number of sensors can be placed on the wing model during one test, which results in sparse pressure coefficient data with distribution inconsistency across cross sections and conditions. Thus, how to obtain pressure coefficients of more cross sections or even the whole wing surface with multiple conditions from the distribution-inconsistent sensor data becomes a challenging problem. Therefore, a deep learning framework based on transfer learning is proposed in this paper, in which the spatial dependency captured by a long short-term memory model between the obtained multi-conditional sensor data is transferred to other cross sections with few-condition data on the wing. The results demonstrate that the proposed framework achieves high accuracy on the pressure coefficients prediction of distribution-inconsistent cross sections on wind tunnel test data, and thus improves data utilization and cuts costs by reducing wind tunnel tests under different design conditions. Our work proves the possibility of reconstructing the holographic flow field from sparse sensor data of wind tunnel tests and puts forward recommendations on the placement of sensors for achieving this goal.

Numerical study of drag force on the UAV tilt and fixed wing model during vertical take off landing

A hybrid vertical take-off landing (VTOL) UAV combines the concepts of fixed-wing and rotary-wing UAV aircraft in one platform while performing in both conventional and vertical take-off landings. This aircraft has a drawback of significant drag force generated due to fixed-wing. Therefore, a Tilt-Wing often utilized to overcome this obstacle whereby it could be adjusted to the vertical and horizontal directions. To enhance the understanding of generated drag force on both wing model, this study was performed by examine the drag characteristic of the VTOL UAV. The simulations were carried out in the wind speed range by 1.4 m/s, 4.17 m/s, and 6.94 m/s. Simulation results showed that the drag of the UAV Hybrid Tilt-Wing and Fixed-Wing increased at the speed of 1.4 m/s to 6.94 m/s while the highest drag value was 177.51 N on a Fixed-Wing UAV aircraft and 1.97 N on a Tilt-Wing UAV aircraft. The result concluded that Tilt-Wing UAV has less drag which was more efficient than fixed-wing.

Real-time feedback stall control of an airfoil at large Reynolds numbers using linear genetic programming

Machine learning control is applied in real-time to an airfoil equipped with variable-velocity jets and pressure sensors in a closed-loop wind tunnel. The objective of the control is to improve the lift-to-drag ratio using an array of variable velocity jets located at the leading edge of the wing model. Pressure sensors are located along the chord, at mid-span, and are used in a feedback control strategy where the symbolic control laws are optimized using a linear genetic programming control (LGPC) algorithm. For angles of attack at the onset of stall, we show that the best control law is able to outperform the best open-loop control strategy in the case of a single-input multiple-output control loop. Results are reported for Reynolds numbers at both half a million and one million. In particular, the lift-to-drag ratio is improved by 4% compared to the best open-loop strategy, which corresponds to a relative increase by a factor 3. In the case of leading-edge separation conditions, LGPC builds a control law performing similarly to the best open-loop strategy while minimizing the actuation power. This study suggests that in order to improve the control authority, the LGPC strategy is sufficiently mature. This control methodology highlights that in order to improve the aerodynamic performances, novel fluidic actuators providing access to the orientation of the jets could enable breakthroughs for high-Reynolds-number experimental demonstrators.

Unsteady interaction of crossflow instability with a forward-facing step

Experiments have been conducted on a swept wing model in a low-turbulence wind tunnel at chord Reynolds number of $2.17 \times 10^{6}$ to investigate the unsteady interaction of a forward-facing step (FFS) with incoming stationary crossflow (CF) vortices. The impact of varying the FFS height on the development and growth of primary and secondary CF disturbances and the ensuing laminar–turbulent transition is quantified through detailed hot-wire anemometry and infrared thermography measurements. The presence of the FFS results in either a critical (i.e. moderate transition advancement) or a supercritical behaviour (i.e. transition advancing abruptly to the FFS location). The arrival of the forced stationary CF vortices at the step is accompanied by their amplification. Unsteady analysis for the critical cases indicates temporal velocity fluctuations following closely the development of the baseline configuration (i.e. agreeing with the development of secondary instabilities). Consequently, laminar breakdown originates from the outer side of the upwelling region of the CF vortices. In contrast, for the supercritical FFS, the laminar breakdown unexpectedly originates from the inner side of the upwelling region. Evidence points to an unsteady mechanism possibly supported by locally enhanced spanwise-modulated shears and the recirculation region downstream of the FFS edge. This mechanism appears to govern the abrupt tripping of the flow in supercritical step cases. The findings in this work provide insight into the unsteady FFS–CF vortex interaction, which is pivotal to understanding the influence of an FFS on the laminar–turbulent boundary-layer transition in swept aerodynamic surfaces.

Low speed aerodynamic characteristics of non-slender delta wing at low angles of attack

Low-speed wind tunnel experiments are conducted to study the aerodynamic performance of a half-span delta wing with 45° leading-edge sweep at subsonic flow regime. The experiments are carried out at a Reynolds number of 8.37 × 105, a free-stream Mach number of 0.1 and angles of attack up to 25°, in steps of 5°. The test model was designed with thirty-two pressure taps fixed on its surfaces (sixteen on each side). Multi-tube manometers were connected to these taps using long tubes to enable recording the pressure readings. Surface pressure distributions and aerodynamic characteristics were calculated at different span-wise locations along the non-dimensional chord-wise distance. Results exhibited that most lift on the studied wing is generated in the region close to the leading edge for all the studied incidence angles. Additional lift is created in the region close to the root chord rather than the tip chord, whereas drag forces increases from tip to root. This can be attributed to the formation of trailing edge vortexes due to the flow separation at the wing leading edge that produces more drag, hence suppressing lift. The study showed also that angle of attack increases the drag coefficient from tip to root, especially at high angle of attack, indicating unfavourable behaviour for manoeuvring. Moreover, the angle of attack increased the pitching moment coefficient up to 10° before it drops sharply until it reaches the tip of the wing model.

Numerical and Experimental Modal Analyses of Inclined Supported Bridge Wing Structures

This paper studies ship bridge extensions called "bridge wings", which are one of the most critical areas of importance in ship vibration. These extensions, which have been modeled at the laboratory scale, have been analyzed using both experimental and finite element analysis methods. Bridge wing models have been constructed with varying inclination angles, and the natural frequencies and mode shapes obtained, both experimentally and numerically, have been compared. In this way, the effects of design properties on the results have been emphasized. Besides, the natural frequency discrepancies between the experimental and finite element analyses have been found to be very low. Another important result of this study has been obtained such that a certain support angle has been detected which the mode shape transition occurs in resonance. Natural frequency and mode shape analyses of the plates which have been supported with beams with varying inclination angles and achievement of mode shape transition phenomenon can be considered as the novelty about the contribution of this study to the state-of-the-art.

Flutter Prediction Using Reduced-Order Modeling with Error Estimation

This paper presents a methodology for dynamic aeroelastic analysis of aircraft based on model order reduction with error estimation. A projection-based model order reduction approach is used to create an aerodynamic reduced-order model (ROM), which is coupled to a structural model to create an aeroelastic ROM. The governing aerodynamic equations are the linearized semidiscrete Euler equations. Flutter analysis is conducted by analyzing the eigenvalues of the aeroelastic ROM. A dual-weighted residual-based error estimator is presented which approximates the error in the eigenvalues obtained from the reduced eigenproblem relative to the eigenvalues from the high-dimensional aeroelastic model. The error estimator thus allows for the construction of aeroelastic ROMs with select eigenvalues that satisfy a user-prescribed accuracy. The aerodynamic ROM is constructed using approximate high-dimensional aeroelastic eigenvectors computed using the two-sided Jacobi–Davidson algorithm. Dynamic aeroelastic analyses are presented for the NACA 64A010 Isogai case, the AGARD 445.6 wing model, and a NACA 0012 benchmark case. The error estimator is shown to have good agreement with the exact error. For the test cases presented in this work, the cost of computing the flutter point at a given Mach number is equivalent to the cost of approximately 5 to 12 steady nonlinear flow evaluations of the high-dimensional Euler equations.

Active robust control for wing vibrations attenuation

In this paper a flexible smart wing is presented. The aileron wing with aileron was tested in the INCAS Subsonic Wind Tunnel with the aim of flutter mitigation and vibrations attenuation. The work takes over some of the results of the active anti-flutter control tests performed in the wind tunnel based on the receptance method. More specifically, the mathematical models in the time domain, necessary for the synthesis of control laws, are obtained from experimentally identified transfer functions. The main part of the paper presents the synthesis and analysis of the robustness of the control laws LQG, LQG/LTR, 𝐻𝐻∞ standard and 𝐻𝐻∞ robust from which the optimal solution will be chosen in a later work for the active vibration control tests on this smart wing model.

Vibration control studies of a high-aspect-ratio wing with geometric nonlinearity

Many challenges of vibration problems due to structure nonlinearity are growing while high-aspect-ratio wings are being widely used. Feasible vibration control method is an effective way to decrease such challenges. At the beginning, a new high-aspect-ratio flexible wing model with swept-back angle and control surface was designed and manufactured. Numerical simulation was employed to find basic dynamic characteristics of the wing. Ground experiments could provide vibration test results of this wing. The numerical method of computational flutter velocity of the wing including geometric nonlinearity was applied to study this wing. Then the open-loop flutter dynamic pressure and the closed-loop flutter speed were compared, it is found that the control law designed by linear quadratic Gaussian theory can increase the flutter speed of the present geometric nonlinear wing model. The computational results of flutter vibration show that the geometric nonlinear flutter speed is lower than the linear flutter speed, and the change in torsional mode is the critical reason for the decrease in flutter speed. Both numerical and experimental method proposed in this paper could be applied to predict the approximate flutter effect of such complex elastic high-aspect-ratio wing via structural nonlinearity.

Analysis of Surface Patterns and Electric Field Simulation of Antireflective Green Lacewing Wings.

The structural coloration and decoloration are problems of scientific interest for a long time. Hence, the fundamental investigations on structures and the optical properties of insect wings have been performed. As a part of such studies, we elucidate the optical properties of green lacewing wings via observation and simulation. First, we elucidate the surface pattern of green lacewing wings using a two-dimensional fast Fourier transform. A cross-shaped pattern of a Fourier spectrum is obtained, and the concise wing model with the surface protrusions arranged in a square grid on a base substrate is constructed in reference to the obtained Fourier spectrum. Next, we perform a finite-difference time-domain (FDTD) simulation to elucidate a light path through wings with and without surface protrusions. The FDTD simulation results indicate that the surface protrusions of a wing increase and decrease the intensity of the transmitted and reflected light, respectively, which is an antireflection behavior. This phenomenon was also observed in the case of 45° incident light. The intensity of transmitted light coupled to wings is induced by surface protrusions with a stepwise refractive index between air and a substrate, which induces antireflection. In particular, transmitted light is increased by the surface protrusions of wings in the range of 500-800 nm wavelength. The intensities of transmitted and reflected light are affected by the direction of incident electric field (polarization) in the case of wings with protrusions arranged in the same direction (parallel). Hence, the surface protrusions are arranged in a square grid to reduce the influence of the polarization direction.

Suction and Pulsed Blowing Applied to Local Separation Downstream of Slat Cutout

An experimental study on the application of active flow control (AFC) to a .4 scale model of a swept wing in a landing configuration was conducted. The wing is fitted with an ultrahigh bypass ratio (UHBR) engine nacelle. The highly efficient UHBR engines are characterized by a large diameter that interferes with the flow around the wing, especially with slats deflected, degrading its performance. An innovative AFC device, creating steady suction and pulsed blowing (PB), was installed in the leading-edge region of the wing, above the nacelle, and its performance was experimentally evaluated. The effects of the suction and PB mechanisms were examined individually and simultaneously using relevant normalized parameters to pave the way to a full-scale wind-tunnel test. It was shown that the AFC actuators increased lift by up to 3%, redirected the flow to the desired downstream direction, and reduced the size of the separation zone created due to the implementation of the UHBR nacelle. The next step is validating the small-scale results of this study in a full-scale wind-tunnel tests that would be the last step before the technology is flight test ready.

Research on dynamic aerodynamic characteristics of flying wing model with low aspect ratio

The flying wing with low aspect ratio removes the horizontal and vertical tails and adopts configuration of wing-body fusion, which has better lift and drag characteristics and stealth than the traditional wing-body split configuration. In this paper, with the help of the semi-reflux transient transonic wind tunnel of China Aerodynamics Research and Development Center, a single-degree-of-freedom pitching dynamic force test is carried out on the standard model of flying wing with low aspect ratio, and the mechanism of some cases is analyzed through numerical calculation. The study found that when the fuselage is forced to pitch at f=1Hz: The normal force coefficient curve will form a hysteresis loop in the shape of a spoon; during the pitching process, the angle of attack of stall is delayed from the 40° which is the angle of attack of stall in static case to 45°; the pitching movement will delay the dissipation of the energy of leading edge vortex along the flow direction and increases the vortex suction at the rear of the wing; the flow pattern of the suction surface is similar during the upward elevation process, and spanwise flow is more severe than the static case under the same angle of attack.