Leading Edge Sweep Research Articles

Triglobal resolvent-analysis-based control of separated flows around low-aspect-ratio wings

We perform direct numerical simulations of actively controlled laminar separated wakes around low-aspect-ratio wings with two primary goals: (i) reducing the size of the separation bubble and (ii) attenuating the wing tip vortex. Instead of preventing separation, we modify the three-dimensional (3-D) dynamics to exploit wake vortices for aerodynamic enhancements. A direct wake modification is considered using optimal harmonic forcing modes from triglobal resolvent analysis. For this study, we consider wings at angles of attack of $14^\circ$ and $22^\circ$ , taper ratios $0.27$ and $1$ , and leading edge sweep angles of $0^\circ$ and $30^\circ$ , at a mean-chord-based Reynolds number of $600$ . The wakes behind these wings exhibit 3-D reversed-flow bubble and large-scale vortical structures. For tapered swept wings, the diversity of wake vortices increases substantially, posing a challenge for flow control. To achieve the first control objective for an untapered unswept wing, root-based actuation at the shedding frequency is introduced to reduce the reversed-flow bubble size by taking advantage of the wake vortices to significantly enhance the aerodynamic performance of the wing. For both untapered and tapered swept wings, root-based actuation modifies the stalled flow, reduces the reversed-flow region and enhances aerodynamic performance by increasing the root contribution to lift. For the goal of controlling the tip vortex, we demonstrate the effectiveness of actuation with high-frequency perturbations near the tip. This study shows how insights from resolvent analysis for unsteady actuation can enable global modification of 3-D separated wakes and achieve improved aerodynamics of wings.

Surrogate-based shape optimization and sensitivity analysis on the aerodynamic performance of HCW configuration

With an additional wing upon the fuselage, the novel high-pressure capturing wing (HCW) configuration exhibits remarkable aerodynamic characteristics at hypersonic speeds under beneficial aerodynamic interference. This bi-wing structure can also enhance the lift at subsonic speeds, positioning HCW configuration as an excellent aerodynamic layout under wide-speed range conditions. In this paper, a single-wing principle HCW configuration is carried out to analyze the influence of the variations in the geometric parameters of the HCW on the aerodynamic performance. Drawing on extant research, four key geometric parameters of the HCW are chosen as design variables, and the hypersonic as well as supersonic conditions are selected for surrogate-based shape optimization. Utilizing polynomial response surface method and method of moving asymptotes, the single-objective optimization studies are carried out with the objectives of maximum lift-to-drag ratio at Ma = 6 and minimum drag coefficient at Ma = 3. Subsequently, the sensitivity analysis is performed for each design parameter. The above methods exhibit notable precision and favorable results. The results show that except for the leading-edge sweep angle, the other three optimization results exhibit divergent trends in their variations. The setting angle has the most significant influence on the aerodynamic forces, owing to the influences of its variation on the shock wave intensity and reflection angle. Based on the stronger intensity of shock wave, this sensitivity indices are higher at Ma = 6. The other three parameters (half-span, leading-edge sweep angle, and trailing-edge sweep angle) modify the aerodynamic forces by adjusting the area of high-pressure region on HCW. Due to the different flow field structures, optimum parameters exhibit diverse effects on lift coefficient and lift-to-drag ratio.

Laminar post-stall wakes of tapered swept wings

While tapered swept wings are widely used, the influence of taper on their post-stall wake characteristics remains largely unexplored. To address this issue, we conduct an extensive study using direct numerical simulations to characterize the wing taper and sweep effects on laminar separated wakes. We analyse flows behind NACA 0015 cross-sectional profile wings at post-stall angles of attack $\alpha =14^\circ$ – $22^\circ$ with taper ratios $\lambda =0.27$ – $1$ , leading-edge sweep angles $0^\circ$ – $50^\circ$ and semi aspect ratios $sAR =1$ and $2$ at a mean-chord-based Reynolds number of $600$ . Tapered wings have smaller tip chord length, which generates a weaker tip vortex, and attenuates inboard downwash. This results in the development of unsteadiness over a large portion of the wingspan at high angles of attack. For tapered wings with backward-swept leading edges, unsteadiness emerges near the wing tip. On the other hand, wings with forward-swept trailing edges are shown to concentrate wake-shedding structures near the wing root. For highly swept untapered wings, the wake is steady, while unsteady shedding vortices appear near the tip for tapered wings with high leading-edge sweep angles. For such wings, larger wake oscillations emerge near the root as the taper ratio decreases. While the combination of taper and sweep increases flow unsteadiness, we find that tapered swept wings have more enhanced aerodynamic performance than untapered and unswept wings, exhibiting higher time-averaged lift and lift-to-drag ratio. The current findings shed light on the fundamental aspects of flow separation over tapered wings in the absence of turbulent flow effects.

Numerical Investigations on Low-Speed Aerodynamic Characteristics of Generic Unmanned Combat Aerial Vehicle Configurations

The leading-edge sweep angle significantly influences the flight performance of swept wings. The vortex flow structures affect the flow behavior of the moderate to highly swept wings, which form at low angles of attack because of leading-edge flow separation. This topic of investigation gained momentum due to its application to unmanned combat aerial vehicle (UCAV) configurations that are significantly affected by flow separation at high angles of attack. The present work investigates the flowfield and low-speed aerodynamic characteristics of the generic UCAV models with a constant and nonconstant leading-edge sweep for a wide range of angles of attack. The open-source computational fluid dynamics code OpenFOAM 8.0 is employed for the numerical simulations. The present numerical simulation results agree well with the experimental data in the literature. It is demonstrated that the lift and drag characteristics of the UCAV are sensitive to the leading-edge sweep angle, and a nonconstant leading-edge sweep angle variant exhibits delayed stall and enhanced aerodynamic performance.

Flow structure and performance analysis of Darrieus vertical axis turbines with swept blades: A critical case study on V-shaped blades

Blade geometric modification is one favorite way of enhancing the performance of vertical axis turbines. In this regard, V-shaped blades are a novel design featuring some leading-edge sweep angle. Currently, the relevant literature only focuses on finding the optimum parameters of the V-shaped blades and lacks a deep understanding of how they alter flow structure and whether they offer any performance improvement compared to other conventional types. To this aim, using three-dimensional numerical simulations, the flow structure and performance of two different V-shaped blade turbines, twisted and untwisted, are compared with straight and helical-blade turbines at tip speed ratios (TSRs) ranging from 0.69 to 1.5. Turbine torque average and fluctuation amplitude are chosen as the performance metrics. It is shown that the sweeping configuration of the V-shaped and helical types causes an inevitable spanwise flow, resulting in a performance loss, which makes neither type of V-shaped blade turbine advantageous at low TSRs.

Vortex Interaction Characteristics of Multiswept Wings at Subsonic Speeds

Despite the sufficient knowledge on the physics of vortices formed over slender delta wings, there is a lack of understanding about how corotating and counter-rotating vortices near the wing leading edge interact with each other. The vortex interaction will depend on the wing planform, leading-edge sweeps, leading-edge shapes, angle of attack, angle of sideslip, and the Reynolds number. In this study, experiments were performed on a range of planform configurations to determine the influence of multisweep wing configurations on the vortex interactions and breakdown thereof. Force/moment measurements were conducted to understand globally the effect of the planform configuration, whereas particle tracking surface oil flow and stereoscopic particle image velocimetry were employed to detect interactions locally in the flowfield. Results were validated against corresponding numerical simulations and past experiments. Comprehensive results indicated a strong correlation between the chord ratios of the initial sweep on the suppression of vortex breakdown. Specifically, two cases with varying chord ratios were identified that exhibited unique vortex breakdown progression (gradual and rapid). Flowfield measurements elucidated key flow features responsible for the differing breakdown progression.

Experimental investigation of synthetic jet control of wing rock for a flying wing aircraft

Flying wing aircraft easily experience wing rock due to the lack of lateral-directional stability, which causes serious challenges to flight control and safety. Thus, it is necessary to reduce the wing rock amplitude or reduce the mean roll angle by additional control. For a flying wing model with a 65° leading-edge sweep, we propose a strategy using an array of synthetic jet actuators to control the wing rock. The control effect and mechanism are studied by attitude measurement and particle image velocimetry measurement in a wind tunnel; the results confirm that the synthetic jet can effectively change the trim position of the wing rock. The control effect is affected by the angle of attack, Reynolds number, actuation position, actuation voltage, and frequency. In general, downstream actuators perform better at low angles of attack, while upstream actuators perform better at high angles of attack; the actuators positioned at the downward rolling side have a better effect than those positioned at the upward side. Furthermore, continuously variable control of the trim position can be achieved by changing the actuation voltage or modulation frequency, which provides a base for attitude manipulation by using active flow control instead of a mechanical control surface. Quantitative analysis of the flow field indicates that the leading-edge vortex on the upward side provides a rolling moment, while the recirculation zone on the downward side also contributes to the wing rock. This is a dynamic process, causing the flying wing to balance at a nonzero mean roll angle. The synthetic jet positioned at the downward rolling side can transport high-momentum fluids to the near-wall region, thereby suppressing flow separation and reducing the size of the recirculation zone. This enhances the lift on the control side and thus reduces the mean roll angle of the wing rock.

Conceptual Design of a Nonconstant Swept Flying Wing Unmanned Combat Aerial Vehicle

The design of unmanned combat aerial vehicles (UCAVs) is primarily governed by the low-observability requirement for military applications rather than aerodynamic performance. The conceptual design and optimization of UCAV models via size and shape variables for different missions in different flow regimes form a research area for military vehicle design. Flying wing UCAVs experience flow separation during takeoff and landing, and furthermore exhibit stability issues. The aerodynamic performance of these UCAVs can be significantly improved by redesigning their leading-edge sweep angle and wing planform. In the present work, the initial weight determination, aerodynamic sizing, and planform with and without inlet lip integration, and the conceptual design of a nonconstant leading-edge flying wing UCAV configuration are performed. Next, the obtained conceptual design is downscaled 1:20 to be used as a wind-tunnel model and optimized for low-speed conditions using a kriging-based surrogate model with a vortex-lattice method to maximize the lift-to-drag ratio. Later, the optimized design is validated using an open-source computational fluid dynamics code, OpenFOAM 8.0, to verify the accuracy of the surrogate model and to investigate the aerodynamic characteristics. The optimized UCAV design exhibited improved aerodynamic characteristics in terms of the lift-to-drag ratio. Furthermore, the aerodynamic performance and flowfield of the optimized UCAV model with and without inlet lip integration have been evaluated at low and high speeds.

Swept-Wing Active Flow Control with a Streamwise Row of Vortex Generating Jets

A novel swept-wing active flow control (AFC) technique was developed and tested on a NACA wing with 30° leading-edge sweep. A streamwise row of 20 vortex generating jets (VGJs), pitched 30° from the wing surface and skewed 60° toward the wing root, was applied at three different spanwise locations. Performance was assessed using load-cell measurements and surface oil flow visualizations at a Reynolds number of 100,000. More inboard placement of the AFC produced higher values of , but further outboard placement resulted in a higher stall angle. With a momentum coefficient of only 2%, lift was enhanced up to 15% and stall was delayed up to 12°, depending on the AFC location. The number of jets used in the row was reduced, resulting in nearly the same performance benefits with half of the momentum coefficient required for the full row of jets. The performance enhancements of this active flow control technique are comparable to a passive boundary-layer fence without the 25% increase in drag caused by the fence. The mass flow cost of the streamwise row of VGJs is 75% lower than the mass flow required for a previously studied swept-wing active flow control technique involving a streamwise slot.

Flow features of swept shock/turbulent boundary-layer interaction due to a gap beneath a sharp fin

A numerical study of the shock/turbulent boundary-layer interaction (STBLI) generated by a sharp fin with a 60deg leading-edge sweep mounted above a flat plate at Mach 2.5 was performed. Incipient separation due to this highly swept fin was determined by varying the fin angle to the incoming freestream. The effect of the fin/plate gap on STBLI incipient separation was established. Sweeping the fin increases the deflection required for incipient separation. Sweeping the fin reduces the interaction strength. A gap causes flow leakage under the fin, further reducing the interaction strength on the windward side. The incipient angle of attack is increased when the gap is large enough. In this study, all gaps investigated produced incipient separation. For the leeward side of the fin, a vortex at the fin leading edge was generated due to the pressure difference across the fin leeward and windward sides. In the cases with gaps, an additional vortex is generated at the fin leading edge and a vortex at the plate surface is also produced. It was found that the effect of the gap is to change the location and shape of these additional vortices.

Hydrofoil geometry and leading-edge modifications: Influence of section profile, aspect ratio, and sweep

Modifications mimicking the leading-edge protuberances on the pectoral fin of humpback whales have been widely adopted to the designs of foils and control surfaces to delay stall and provide post-stall superior lifting characteristics. However, the efficacy of these modifications in flow control and lift augmentation depends on the wing geometry. The present work investigates the lift, drag, and flow characteristics of finite-span hydrofoils with different section profiles-both symmetric and cambered (NACA 0012, NACA 0018, NACA 634-021, NACA 2412, and NACA 4415) having twin-protuberances over their leading-edge at a Reynolds number of 2 × 105. The influence of aspect ratio (1, 2, and 3) and leading-edge sweep angle (0 deg, 15 deg, and 26.1 deg) on the hydrodynamic performance of modified foils is also investigated. The characteristic feature of twin-protuberance foil designs is the restriction of flow separation between the chordwise vortices shed from the two protuberances in certain post-stall regimes. The use of leading-edge modifications is observed to be more conducive for lift enhancement in thin foil sections, especially NACA 0012, having a smaller stall angle. Considering the other parameters, leading-edge protuberances are observed to be advantageous for hydrofoils with aspect ratios 2 and 3 at post-stall angles of attack, and for 0 deg and 15 deg sweep angles for post-stall lift enhancement. The influence of Reynolds number on the performance modifications is also investigated.

Aerodynamic Characteristics and Topology of Interfering Vortex Systems at Hybrid-Delta Wings

High-agility aircraft often perform flight maneuvers at high angles of attack and angles of sideslip. For these configurations, a comprehension of the vortex development, interaction, and breakdown at sideslip conditions is crucial. In this study, two generic, low-aspect-ratio, hybrid-delta-wing configurations are investigated. A triple-delta-wing and a double-delta-wing configuration are discussed. Force and moment measurements, particle image velocimetry, and oil-flow visualization tests are conducted to resolve the vortex-dominated flowfield extensively. The flow separates at the sharp leading edges, and vortices are formed. For an angle of sideslip, the effective leading-edge sweep on the windward side is reduced and on the leeward side increased. This leads to a more stable vortex system on the leeward side and a destabilization on the windward side. Therefore, significant lateral instabilities occur at high angles of attack.

The flow mechanism of the shock structure and secondary flow control in transonic centrifugal compressors with leading edge sweep

Limitations still exist in the research on the effects of leading edge sweep on transonic centrifugal compressors. The lack of understanding of this mechanism also leads to confusion regarding the magnitudes of the performance changes that occur due to the use of leading edge sweep. According to a detailed numerical simulation verified by experiments on two different compressors, the effect and flow mechanism of leading edge sweep on transonic centrifugal compressors were investigated in detail to alleviate the confusion as clearly as possible. The conclusions show that leading edge sweep has different impacts on shock waves at different positions; a forward sweep leading edge can weaken the strength of the shock wave in front of the leading edge, while a backward sweep leading edge can weaken the strength of the passage shock wave that is swallowed in the blade passage. The improved compressor performance attributed to leading edge sweep is due to the weakening of the shock waves and the interaction between the shock waves and the secondary flow of the inducer tip. Furthermore, the intensity of the inducer tip secondary flow influences the potential for improved performance induced by using sweep leading edges in transonic centrifugal compressors. This study provides a valuable reference for applying leading edge sweep to existing transonic centrifugal compressors or future advanced centrifugal compressor designs.

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.

Effect of the Leading-Edge Sweep on Wrap-Around Fins

In this paper, the investigation was performed to determine the effect of leading-edge angle and leading-edge sweep on the aerodynamic coefficients of the projectile geometry and analyze the effect of roll moment on the fins at various operating conditions. A group of four models were considered, standard TTCP with the blunt leading edge and with 45° leading-edge along with modified TTCP with the blunt leading edge along with 30° sweep and with 45° leading-edge along with 30° sweep. The flow field solutions were obtained and considered as a function of roll moment coefficients, which are then compared to other numerical models and experimental results. The standard wrap-around TTCP models were subjected to varying velocities ranging from Mach 1.5 to 2.5. WAFs are considered a choice for spinning tube-launched projectiles because of their high packing factor and instant deployment. The computational grids have been built to accurately reflect the fin edge shape of the experimental model. The significant parameter of the WAF is the rolling moment, that was computed using CFD and compared with available experimental test data.

Investigation on a Flow Coupling Rudder for Directional Control of a Low-Aspect Tailless Configuration with Diamond-Shaped Wing

To solve the issues of directional control and aerodynamic moment coupling for a low-aspect tailless configuration with a diamond-shaped wing, we herein propose the concept of a flow coupling rudder (FCR). The FCR was composed of two basic control surfaces: a spoiler slot deflector (SSD) on the upper surface of the wing and the corresponding elevon. With the constraints of the fixed area and spanwise position of the SSD, the effects of the leading-edge sweep angle, chord position of the SSD, and collocating deflection angles of the SSD and elevon on the control characteristics of the configuration were analyzed using a numerical method. Based on the analysis, the selection principle of the key parameters for designing the FCR for the configuration was proposed. This proves that the leading edge of the selected SSD should be swept back instead of being parallel to the trailing edge of the wing to design an FCR with favorable aerodynamic performance for such a configuration. By accurately adjusting the parameters of the SSD and elevon, the FCR on the one-side wing could supply an effective yawing moment while simultaneously weakening or even eliminating the coupling pitching and rolling moments of traditional drag-type control surfaces. In this study, we provide a method for the directional control surface design of a low-aspect tailless configuration with a diamond-shaped wing.

Experiment on longitudinal aerodynamic characteristics of flying wing model with plasma flow control

Horizontal tail is eliminated from the flying wing layout for improving the low observable and aerodynamic efficiency, resulting in degrading longitudinal maneuverability and fight stability. The low speed wind tunnel test study of improving the longitudinal aerodynamic characteristics of large aspect ratio flying wing model is carried out by using plasma flow control technology. The flying wing model has a leading-edge sweep angle of 34.5° and an aspect ratio of 5.79. The reasons for deteriorating the static maneuverability and stability of the flying wing model and the mechanism of plasma control of the flow field and longitudinal aerodynamic characteristics are studied by particle image velocimetry (PIV) flow visualization and static force measurement test. The control law of plasma control of the flight maneuverability and stability of the flying wing model is studied through flight test. The fact that the flow separation of the outer wing of the flying wing model occurs earlier than the inner wing and the wing is swept back can result in the forward movement of the aerodynamic center and the deterioration of the longitudinal static stability. The shock disturbance induced by plasma can suppress the flow separation of the suction surface, thereby extending the linear section of the lift curve of the model, preventing the aerodynamic center from moving forward, and improving the longitudinal static stability. When the wind speed is 50 m/s, the plasma control improves the horizontal rudder efficiency at a high angle of attack of the flying wing model, increases the maximum lift coefficient of the model by about 0.1, and postpones the stall angle of attack by more than 4° at different rudder angles. The plasma control allows the flying model to follow the command movement better while flying, increases the flying pitch limit angle from 11.5° to 15.1°, reduces the amplitude of longitudinal disturbance motion by 2°, and reduces the oscillation attenuation time from 15 to 8 s, thereby improving the longitudinal flight maneuverability and stability of the flying wing model. It can be seen that plasma flow control technology has great potential applications in improving the flight quality of flying wing layout.

Low-Noise Blade Design Optimization for a Transonic Fan Using Adjoint-Based Approach

The target of reducing the environmental footprint of the aviation industry has continually driven the need to design more efficient and quieter aircraft engines. In this paper, the aeroacoustic adjoint formulization for low-shock tone noise fan blade design is first proposed and combined with the aerodynamic adjoint analysis to develop an adjoint-based, multi-objective design optimization process for transonic fan blade design. High-fidelity numerical simulations are employed in the optimization loop to predict aeroacoustic and aerodynamic objectives and gradients. Aeroacoustic and aerodynamic design optimizations of an industrial transonic research fan are conducted using the proposed adjoint-based approach, demonstrating that the noise performance and the efficiency of the fan can be improved simultaneously. Results indicate that the significant reduction in the sound power level of the fan shock-associated tone noise is achieved by a compound leading-edge sweep pattern generated by the optimizations. Flowfields and acoustic results of the optimized blades are then analyzed to understand the noise reduction mechanism.

Оптимизация формы крыла и выбор рациональных параметров траектории многоразового космического аппарата туристического класса

The article considers solving the interrelated problems of wing shape optimization and synthesis of the re-entry trajectory control law for the reusable spacecraft of tourist class. To ensure a high aerodynamic quality of the wing, increase its bearing properties, as well as improve the maneuverability and controllability of the spacecraft as a whole, the wing shape has been optimized for sub- and supersonic flight modes. The problem of minimizing the wing area is solved for subsonic flight speed while ensuring the level of lift sufficient for landing, with the introduction of restrictions on the minimum wing sweep angle. For supersonic flight speed, maximization of the aerodynamic quality of the wing is used as an objective function. The length and taper of the wing, leading-edge sweep angle, the size of the root and tip chords, and the position of the wing relative to the fuselage were chosen as variables. For the wing shape selected on the basis of the parametric analysis, the calculation of the dependences of the spacecraft aerodynamic coefficients on the Mach number, used for selecting a rational program for the descent control in the atmosphere, was carried out. The choice of a rational control program is made with restrictions on the level of overloads, kinetic pressure and maximum heat flux.

Fin sweep angle does not determine flapping propulsive performance.

The importance of the leading-edge sweep angle of propulsive surfaces used by unsteady swimming and flying animals has been an issue of debate for many years, spurring studies in biology, engineering, and robotics with mixed conclusions. In this work, we provide results from three-dimensional simulations on single-planform finite foils undergoing tail-like (pitch-heave) and flipper-like (twist-roll) kinematics for a range of sweep angles covering a substantial portion of animals while carefully controlling all other parameters. Our primary finding is the negligible 0.043 maximum correlation between the sweep angle and the propulsive force and power for both tail-like and flipper-like motions. This indicates that fish tails and mammal flukes with similar range and size can have a large range of potential sweep angles without significant negative propulsive impact. Although there is a slight benefit to avoiding large sweep angles, this is easily compensated by adjusting the fin’s motion parameters such as flapping frequency, amplitude and maximum angle of attack to gain higher thrust and efficiency.