Swept Wing Research Articles

Study on the Influence of Swept Angle on the Aerodynamic Characteristics of the Cross-Section Airfoil of a Variable Swept-Wing Aircraft

The variable forward swept-wing aircraft is a typical variant aircraft. The airfoil is a variable airfoil during the wing swept, and its aerodynamic characteristics have a huge impact on the flight performance of the variable swept-wing aircraft. It is also an aircraft. One of the difficult problems to be solved first in aerodynamic layout design. In order to explore the unsteady aerodynamic characteristics of the variable airfoil, a variable airfoil model with different relative thicknesses is established. Then the S-A turbulence model is used and the CFD simulation is used to calculate the variable airfoil unsteady. The variation of the angle of attack and the Mach number when the aerodynamic parameters change in the forward sweep angle. The results show that the airfoil lift coefficient and moment coefficient are negatively correlated with the relative thickness, while the drag coefficient is positively correlated with the relative thickness. These aerodynamic characteristics can provide a useful reference for the aerodynamic layout design and dynamic analysis of the variant aircraft.

Flow around the wing models with straight and swept leading edge in case of contact with turbulent wake

The paper presents the results of experimental studies of the flow around wing models with a straight and swept leading edge in contact with a turbulent wake. The experiments were carried out in subsonic wind tunnels at low Reynolds numbers. A fluffy thread was used to simulate a turbulent wake in the flow. The flow patterns on the surface of the wings were obtained before and after the impact of the turbulent wake at various angles of attack, wing sweep angles and free-stream velocity values. It was found that the wake has a significant effect on flow separation and on vortex structures arising above the wings. The evolution of the vortex flow patterns in the areas of separation of the boundary layer and with a change in the flow regime is shown.

Model of Distributed Receptivity to Kinetic Fluctuations in High-Speed Boundary Layers

The receptivity of high-speed compressible boundary layers to kinetic fluctuations is considered within the framework of fluctuating hydrodynamics. The formulation is based on the idea that KF-induced dissipative fluxes may lead to the generation of unstable modes in the boundary layer. Fedorov and Tumin (“Receptivity of High-Speed Boundary Layers to Kinetic Fluctuations,” AIAA Journal, Vol. 55, No. 7, 2017, pp. 2335–2348) solved the receptivity problem using an asymptotic matching approach, which used a resonant inner solution in the vicinity of the neutral point. This study adopts a method of multiple scales approach, which requires fewer assumptions about the locus of primary excitation. The approach is modeled after the one taken by Luchini (“Receptivity to Thermal Noise of the Boundary Layer over a Swept Wing,” AIAA Journal, Vol. 55, No. 1, 2017, pp. 121–130) to study low-speed incompressible boundary layers over a swept wing. The new framework is used to study examples of high-speed flat plate boundary layers whose spectra exhibit nuanced behavior near the generation point, such as first-mode instabilities and near-neutral evolution over moderate length scales.

Morphing aircraft control based on switched nonlinear systems and adaptive dynamic programming

This paper investigates the control problem of a morphing aircraft with variable sweep wings based on switched nonlinear systems and adaptive dynamic programming (ADP). The longitudinal altitude motion of the morphing aircraft is first modeled as switched nonlinear systems in lower triangular form. Then, the designed controller is comprised of the basic part and supplementary part. For the basic part, the backstepping technique is applied and a modified dynamic surface is introduced to overcome the ‘explosion of complexity’ problem. Disturbance observers inspired from the idea of extended state observer are designed to obtain the estimations of the internal uncertainties and external disturbances. The common virtual control laws of the backstepping method are developed by the disturbance observers and radial basis function neural networks. On the other hand, for the supplementary part, an ADP approach with the name of action-dependent heuristic dynamic programming is used to further decrease the altitude tracking error, which generates an additional control input by observing the differences between the actual and desired values in the backstepping design. Finally, comparative simulations are conducted to demonstrate the improved control performance of the proposed approach.

The influence of moderate angle-of-attack variation on disturbances evolution and transition to turbulence in supersonic boundary layer on swept wing

The paper is devoted to an experimental and theoretical study of effect of moderate angle-of-attack variation on disturbances evolution and laminar-turbulent transition in a supersonic boundary layer on swept wing at Mach 2. Monotonous growth of the transition Reynolds numbers with angle of attack increasing from −2° to 2.7° is confirmed. For the same conditions, calculations based on linear stability theory are performed. The experimental and computational results show a favourable comparison.

Experimental Aerodynamic Simulation of a Scallop Ice Accretion on a Swept Wing

<div class="section abstract"><div class="htmlview paragraph">Understanding the aerodynamic impact of swept-wing ice accretions is a crucial component of the design of modern aircraft. Computer-simulation tools are commonly used to approximate ice shapes, so the necessary level of detail or fidelity of those simulated ice shapes must be understood relative to high-fidelity representations of the ice. Previous tests were performed in the NASA Icing Research Tunnel to acquire high-fidelity ice shapes. From this database, full-span artificial ice shapes were designed and manufactured for both an 8.9%-scale and 13.3%-scale semispan wing model of the CRM65 which has been established as the full-scale baseline for this swept-wing project. These models were tested in the Walter H. Beech wind tunnel at Wichita State University and at the ONERA F1 facility, respectively. The data collected in the Wichita St. University wind tunnel provided a low-Reynolds number baseline study while the pressurized F1 facility produced data over a wide range of Reynolds and Mach numbers with the highest Reynolds number studied being approximately <i>Re</i> = 11.9×10<sup>6</sup>. Past work focused on only three different fidelity variations for ice shapes based on multiple icing conditions. This work presents a more detailed investigation into several fidelity representations of a single highly three-dimensional scallop ice accretion. Sensitivity to roughness size and application technique on a low-fidelity smooth ice shape is described. The data indicate that the aerodynamic performance is not especially sensitive to the grit variations. An ice accretion code was also used to generate ice shapes for aerodynamic testing and comparisons. These ice shapes have a general appearance like the low-fidelity smooth ice shapes, but in this case, the computer-generated ice shape is significantly smaller. As such, the impact of that ice shape on the aerodynamic performance of the wing is reduced compared to the smooth ice shape based on the icing experiment for those same conditions. Spanwise discontinuities were also introduced to a low-fidelity ice shape in an attempt to quantify the impact of those variation in the high-fidelity ice shape. While the lift data indicate good agreement between the high-fidelity ice shapes and the low-fidelity ice shapes with spanwise discontinuities, a closer investigation of the data suggests potential, significant differences in the flowfield. These results were similar at both facilities over the wide range of test conditions utilized.</div></div>

Additional Comparison of Iced Aerodynamic Measurements on a Swept Wing from Two Wind Tunnels

<div class="section abstract"><div class="htmlview paragraph">Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University’s Walter H. Beech Memorial Wind Tunnel utilizing an 8.9% scale model, and high Reynolds number tests were conducted at ONERA’s F1 wind tunnel utilizing a 13.3% scale model. Several identical geometrically-scaled ice shapes were tested at both facilities, and the results were compared at overlapping Reynolds and Mach numbers. This was to ensure that the results and trends observed at low Reynolds number could be applied and continued to high, near-flight Reynolds number. The data from Wichita State University and ONERA F1 agreed well at matched Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations. This confirmed results from previous tests with other ice shapes that indicated the data from the low Reynolds number tests could be used to understand iced-swept-wing aerodynamics at high Reynolds number. This allows ice aerodynamics testing to be performed at low Reynolds number facilities with much lower operating costs and generate results that are applicable to flight Reynolds number.</div></div>

Aerodynamics, Stability and Performance Characteristics of X-Tandem Aircrafts

This paper details design, analysis and validation of tandem aircraft with one wing swept forward and the other wing swept backwards. The design objective is to investigate the aerodynamic and stability characteristics of this configuration created with the motive of exploiting the manoeuvring and post stall characteristics of a forward swept wing along with the structural robustness and reliable performance of the conventional aft-swept wing. Parameters such as wing sweep, wing position, anhedral and dihedral were varied to develop a range of designs. This gave considerable information regarding the aerodynamic and stability characteristics which enabled a preliminary design of a military combat aircraft exploiting this configuration. All performance characteristics and parameters of the final design compared with current operational military aircraft give a favourable picture regarding the effectiveness of this design.

Aerodynamic Characteristics of Conventional and Innovative High Lift Swept Wings

The new bird inspired wing sweep was introduced and compared with straight and conventional swept wings in gliding flight by an experimental test setup. Due to the similarity with the birds’ wing, all test models have S1223 airfoil. Swept models inspired from the bird consist of two parts: the straight part near the root and the swept part near the tip. Aerodynamic forces on each wing were measured from 0° to 24° angles of attack and Reynolds numbers of 4.3×104, 8.6×104, 1.3×105, and 1.7×105. Wind tunnel test results show that wings with an innovative sweep at α > 0 have more lift for Reynolds numbers between 4.3×104 and 8.6×104. Also, innovative sweep increases the stall angle, and the wing did not stall until α = 24° for Reynolds between 1.3×105 and 1.7×105. An increase in lift and having sufficient aerodynamic performance in low Reynolds numbers for birds’ inspired wing sweep in gliding flight may be the answer to why the wing sweep of birds is not like conventional sweeps.

Coordinate Descent Optimization for Winged-UAV Design

In this paper, a powerful optimization framework is proposed to design highly efficient winged unmanned aerial vehicle (UAV) that is powered by electric motors. In the proposed approach, the design of key UAV parameters including both aerodynamic configurations, (e.g. wing span, sweep angle, chord, taper ratio, cruise speed and angle of attack) and the propulsion systems (e.g. propeller, motor and battery) are cast into an unified optimization problem, where the optimization objective is the design goal (e.g. flight range, endurance). Moreover, practical constraints are naturally incorporated into the design procedures as constraints of the optimization problem. These constraints may arise from the preliminary UAV shape and layout determined by industrial design, weight constraints, etc. The backend of the optimization based UAV design framework are highly accurate aerodynamic models and propulsion system models proposed in this paper and verified by actual experiment data. The optimization framework is inherently non-convex and involves both continuous variables (e.g. the aerodynamic configuration parameters) and discrete variables (e.g. propulsion system combinations). To solve this problem, a novel coordinate descent method is proposed. Trial designs show that the proposed method works rather efficiently, converging in a few iterations. And the returned solution is rather stable with different initial conditions. Finally, the entire approach is applied to design a quadrotor tail-sitter VTOL UAV. The designed UAV is validated by both CFD simulations and intensive real-world flight tests.

Prescribed performance control of morphing aircraft based on switched nonlinear systems and reinforcement learning

The prescribed performance control of a morphing aircraft with variable sweep wings is investigated based on switched nonlinear systems and reinforcement learning. Switched nonlinear systems in lower triangular form are first adopted to describe the longitudinal altitude motion, and an error transformation is applied to handle the prescribed performance bound. Then, the designed controller is divided into the basic part and supplementary part. For the basic part, the backstepping method with involvement of the modified dynamic surface control technique is utilized to avoid the “explosion of complexity” problem. Improved disturbance observers inspired from the idea of extended state observer are then designed to estimate the disturbances and combined with radial basis function neural networks to develop the common virtual control laws. Moreover, by using the error variables defined in the backstepping design, the reinforcement learning–based supplementary part controller is devised with the critic-action neural network structure, which can adjust its parameters online and further decrease the altitude tracking error. It is proved that all signals of the closed-loop system are uniformly ultimately bounded, and the prescribed performance bound for convergence of the altitude tracking error can be satisfied. Finally, comparative simulations demonstrate the effectiveness of the proposed control approach.

CFD Analysis of the Sideslip Angle Effect around a BWB Type Configuration

In this study, we conducted numerical simulations for a nonslender BWB type planform with a rounded leading edge and span of 2.0 m to analyze the effect of the sideslip angle on the planform at a freestream velocity of 60 m/s. The Reynolds number based on the mean chord length was 2.9×106, and we considered the angle of attack ranging from -4° to 16° and sideslip angles up to 20°. We used an unstructured mesh with a prism layer for the boundary layer with 1.11×107 grid points, and the k−ω SST turbulence model. We analyzed force and moment coefficients with respect to variation of angle of attack and sideslip angles. Side force and rolling/yawing moment coefficients had highly nonlinear relationships with the sideslip angle while lift and drag coefficients were not significantly affected. We interpreted the mechanism of these aerodynamic characteristics based on pressure and skin friction contours. Suction pressure near the leading edge had a marked effect on the pitching and rolling moment. We identified five flow types on the blunt leading edge swept wing by skin friction lines and off-body streamlines at a high angle of attack and sideslip angles.

Transonic static aeroelastic and longitudinal aerodynamic characteristics of a low-aspect-ratio swept wing

Based on the CFD/CSD coupling method, the transonic static aeroelastic characteristics of the low-aspect-ratio swept wing was studied by numerical simulation. The Navier-Stokes equation was used as the fluid computational control equations. The flow field mesh deformation was performed using the Transfinite Interpolation method (TFI) method, and the data exchange on the CFD/CSD coupling interface was realized by the Constant Volume Transform (CVT) method. Based on the above method, the transonic static aeroelastic coupling calculation program was developed. The results of ONERA M6 wing static aeroelastic calculated by the coupling program were compared with wind tunnel test and reference results, and the effectiveness of the program was verified. Through numerical simulation, the influence of different angles of attack and Mach number on the static aeroelastic deformation and torsion angle of the wing are studied. Finally, by comparing the surface pressure cloud and pressure coefficient of the rigid wing and the elastic wing, the influence of static aeroelastic on the longitudinal aerodynamic characteristics and stability of the wing is mainly analyzed.

Multipoint Design Optimization of Vortex Generators on Transonic Swept Wings

Multipoint design optimization of the shape of vane-type vortex generators (VGs) and their arrangement on swept wings are conducted using computational fluid dynamics in order to provide designs that are suitable for both cruise and critical conditions. The optimization is divided into two phases: the VG arrangement is optimized using the NASA Common Research Model (CRM) after the VG shape is optimized on a swept infinite wing with a cross-sectional airfoil of the CRM. The shape optimization reveals that VG height and incidence angle control the aerodynamic performance under cruise and critical conditions, and VGs with higher length-to-height ratios show better performance. One of the shape-optimized VGs has been used in this study to optimize the VG arrangement to ensure static stability of the pitching moment with minimally increased drag under cruise conditions. Optimized VG arrangements can prevent shock-induced separation from propagating to wing tips around the Yehudi break of the CRM wing. One optimal solution reduces conventional VG drag penalties by 50%.

Sukhoi SU-47 Berkut and Eurofighter Typhoon Models Flow Visualization and Performance Investigation Using GAMA Water Tunnel

Changes and modifications to the wings of fighter aircraft were carried out, one of which was the forward swept wing which was a moderate wing that continues to develop. There were also types of delta wings that had been applied to many fighter planes. Both types of aircraft wings had certainly different aerodynamic characteristics. This research would study the flow visualization that occurs in the aircraft model body to determine the aerodynamic characteristics of the forward swept wing and delta wing. This study used a water tunnel to observe the aerodynamic flow and forces that occurred in both types of wings. This visualization test used similar aircraft models: SU-47 Berkut and Eurofighter Typhoon. The results provided flow visualization, coefficient of lift (Cl), and coefficient of drag (Cd) which showed that the stall that occurred on the aircraft model similar to the SU-47 Berkut occurred at an angle of attack (AoA) 500 with a Cl max value of 2.66. Meanwhile, the Eurofighter Typhoon stall model occurred at an angle of attack 450 with a Cl max value of 1.48.

Shape Optimization and Sensitivity Analysis of a Morphing-Wing Aircraft

Morphing-wing concepts have received growing interests in the recent years for enhancing the multi-mission performance of unmanned aerial vehicle. However, to obtain a feasible morphing strategy, the optimization design studies are required. In this paper, an aerodynamic optimization study for a morphing-wing aircraft with variable sweep, span, and chord length was conducted to obtain the optimum configurations at subsonic, transonic and supersonic conditions. The optimization objective is to obtain maximum lift-to-drag ratios subject to lift coefficients and static stability constraints at each flight condition. A genetic algorithm in conjunction with surrogate models was employed to search optimum solutions in the entire design space. The aerodynamic forces are calculated by an Euler-based solver and friction drag estimation code. The optimum configuration corresponding to any flight condition can be determined through the optimization. A global sensitivity analysis based on the surrogate model was also carried out, hence the contribution of each design variable to the optimization objective can be analyzed. The results indicated that the optimum wing at subsonic speeds have a lower sweep angle and high aspect ratio, and the wing sweep has a primary contribution to lift-to-drag ratios, and the span is secondary; at transonic conditions, the medium-sweep wing is the optimum, and the contribution of the span is far more than other variables; at supersonic conditions, the optimum configuration becomes a high-sweep cropped delta wing, and the span has a dominant contribution, and the sweep is secondary.

Intermittent Gliding Flight Control Design and Verification of a Morphing Unmanned Aerial Vehicle

The intermittent gliding flight method takes advantage of the updraft is adopted to increase the flight endurance time of a type of multifunctional morphing unmanned aerial vehicle (UAV). This type of UAV is able to be folded before launch, launch from a cylindrical launch tube, and adjust the wing span and sweep angle during flight. It has the advantages of convenient storage and transportation and can use the energy of the updraft. In order to effectively utilize the updraft, the mathematical model of updraft is established, and an updraft judging method is proposed. In order to complete the intermittent gliding flight, the flight progress is divided into two stages and a segmentation control law is proposed. During the intermittent gliding stage switching, the wing sweep angle and wing span length are varied to accommodate requirements of different stages. H ∞ robust control design method for gain scheduling based on linear parameter varying model is used to ensure the stability of the morphing UAV during the wings change. The mathematical simulation verifies the effectiveness of the morphing control algorithm. After that, a semi-physical simulation verification involving a mathematical UAV model with a physical wing changing mechanism, a physical embedded flight control module, the real flight attitude, and the virtual route are implemented. It has been verified that the morphing UAV can remain stable throughout the entire cycle of the intermittent gliding. Moreover, the intermittent gliding flight with updraft can reduce the fuel consumption of the morphing UAV and increase its flight endurance time.

Experimental investigation of the flow structures over a 40° swept wing

ABSTRACTFlow over a moderately swept wing is characterised by complex localised flow vortex topologies such as ‘closed’ separation bubbles or ‘open’ separation structures. A model of a complex cambered, twisted, tapered wing with 40° leading edge sweep, representative of those designed for manoeuvre at high subsonic Mach numbers, was investigated using the oil-film visualisation, stereo particle image velocimetry and force moment measurements. Wind-tunnel tests were conducted at a range of Reynolds number from 2.1×105 to 8.4×105 and at angles of incidence from −1° to 22°. Still images combined with video clips enable flow patterns over the wing model to be interpreted more clearly and accurately. Using successive images extracted from the video of flow visualisation, the movement of the oil pigment has been estimated. The influence of the Reynolds number and incidence angle was discussed through analysing the flow pattern over the wing surface. Additionally, the link between the flow structures present and the wing aerodynamic performance was studied.

Analysis of Morphing Modes of Hypersonic Morphing Aircraft and Multiobjective Trajectory Optimization

Aiming at improving the range of wing-body combination aircraft at hypersonic flow conditions and exploring the application of morphing technology in hypersonic aircraft, morphing aircrafts with different morphing modes have been proposed. The aerodynamic characteristics and wing efficiency of morphing aircraft with different morphing modes have been studied for a further application. In order to verify the significance of morphing technology on the trajectory, the trajectory of the glide phase has been optimized through multiobjective optimization method. Through a 3 degree-of-freedom dynamic model and a heat flux model, the range of glide trajectory and the total heat of the leading edge of the wing are calculated and selected as optimization objectives in the multiobjective optimization problem. The optimization variables include the Mach numbers when the aircraft is morphing (morphing timing) and the angle of attack of different phases. Based on the multiobjective evolutionary algorithm based on decomposition, the multiobjective trajectory optimization problem is solved and a uniform Pareto Front is obtained, and through the analysis of typical solutions, it can be seen that a compromise is made to balance the two objectives. Through the comparison of morphing and non-morphing aircraft, morphing aircraft can fly further with a smaller total heat of the leading edge of the wing. Also, it seems that the variable sweep wing morphing mode has a better overall performance. The result in this paper is a verification of the application prospects of the morphing technology under hypersonic environment, which will provide a reference for further application of morphing technology in hypersonic aircraft.

Experiments on Discrete Roughness Element Technology for Swept-Wing Laminar Flow Control

Micrometer-sized spanwise-periodic discrete roughness elements (DREs) were applied to and tested on a 30 deg swept-wing model in order to study their effects on boundary-layer transition in flight in which stationary crossflow vortices were the dominant instability. Significant improvements have been made to previous flight experiments in order to more reliably determine and control the model angle of attack and unit Reynolds number in order to minimize the uncertainties that DREs have on laminar–turbulent transition on a swept wing. Two interchangeable leading-edge surface-roughness configurations were tested: highly polished and painted. The baseline transition location for the painted leading edge (increased surface roughness) was unexpectedly farther aft than the polished. Infrared thermography, coupled with a postprocessing code, was used to globally extract a quantitative boundary-layer transition location. Linear stability guided the DRE configuration test matrix. Although the results have not reliably confirmed the use of DREs as a viable laminar flow control technique in the flight environment, it has become clear that significant computational studies, specifically direct numerical simulation of these particular DRE configurations on this model geometry and flight conditions, are a necessity in order to better understand the influence that DREs have on laminar–turbulent transition