Delta Wing Configuration Research Articles

Physical aspects of vortex-shock dynamics in delta wing configurations

Delta wing configurations with double- and triple-leading edges introduced within the North Atlantic Treaty Organization Applied Vehicle Technology -316 task group are examined to investigate the dynamics of vortices and shocks, with potential implications for the preliminary aircraft design. The numerical simulations are conducted for the configurations at Ma∞=0.85 and Re∞=12.53×106 using the Reynolds-averaged Navier–Stokes k−ω shear stress transport (SST) model across a range of incidence angles. The detailed analysis focuses on the case with α=20° using the scale-adaptive simulation based on the k−ω SST model. This study considers shock-vortex interaction and breakdown with buffeting to study the transient flow physics over the wing. Additionally, insights into vorticity strength and destruction are gained through the enstrophy transport equation. The findings reveal that the inboard vortex (IBV) development is impeded by counter-rotating secondary vortices from IBV and the midboard vortex. A key distinction is observed for the first time between the double-delta and triple-delta wings, in that the double-delta wing experiences shock-induced vortex breakdown, with the transient nature of this breakdown leading to an adjustment in the shock position, causing a shock buffet. In contrast, the breakdown in the triple-delta wing is linked to a stationary shock induced by the kink in the planform. This study highlights the crucial role of the orientation of the shock relative to the vortex axis in characterizing the aerodynamic performance of the planforms.

Analysis of Vortex Dominated Flow over Double Delta Wing

The maneuverability of fighter aircraft designed using double delta wing configuration at low subsonic speed in addition to under high angle of attacks are highly appealing and fascinating to comprehend. Flow pattern formed by double delta wing help to facilitate lifting force under high angle of attack and at such low subsonic speed. In comparison to delta wing, vortex formations by strake part of wing play an important role to create a persistent vortex core capable of delaying stalling. In this investigation, force coefficients of 75°/45° double delta wing configuration is extracted both from wing tunnel experiment and numerical calculation between 0° to 30° angles of attack at Reynolds number of 1.4 × 105 based on root chord length. It has also been attempted to investigate the flow pattern formation over wing leeward surface, interaction between vortices formed by strake/wing based on Q-criterion and pressure distribution along propagation path of vortices obtained from CFD (Computational Fluid Dynamics) calculations. Implications of interplay between vortices and pressure distribution in lift generation have also been analyzed. It has been found that around 15° angle of attack strake and wing vortices merge and gets more intensified. With higher angle of attack the lift generation significantly diminishes because of vortex bursting

Experimental and numerical analysis of the aerodynamics and vortex interactions on multi-swept delta wings

The flow field around a generic multi-swept delta wing configuration is investigated under transonic flow conditions, both experimentally and numerically. A special focus is on the analysis of vortex/vortex and vortex/shock interactions at moderate angles of attack. In the present study, the Mach number is varied between textrm{Ma} = {0.50} and textrm{Ma} = {1.41} and the angle of attack is varied between alpha = 8^circ and alpha = 28^circ. Numerical results are validated using experimental surface pressure data from pressure taps, as well as forces and moments based on strain gauge measurements. For selected cases, velocity field data from particle image velocimetry (PIV) measurements are available as well. Over a broad range of angle of attack and Mach number, strong vortex/vortex interactions, including vortex braiding and vortex merging, occur. The location of vortex merging is moving downstream with increasing angle of attack and increasing Mach number. Additionally, at textrm{Ma} = {0.85}, vortex/shock interaction occurs above the wing. For moderate angles of attack, shock-induced vortex breakdown is observed.

Dynamic Structural Scaling Concept for a Delta Wing Wind Tunnel Configuration Using Additive Manufacturing

Considering aeroelastic effects plays a vital role in the aircraft design process. The construction of elastic wind tunnel models is a critical element in the investigation of occurring aeroelastic phenomena. However, the structural scaling between full-scale and reduced-scale configurations is a complex design and manufacturing task and is usually avoided in wind tunnel testing. This work proposes a numerical approach for a dynamic aeroelastic scaling technique, which is applied to a fictive delta wing configuration. This scaling methodology is designed to optimise the structural layout of wind tunnel models with an integrated rib and spar structure to meet the behaviour of a realistic full-scale equivalent. For the modelling approach of the wing structure, a beam and shell structure is utilised. The applied scaling laws for the relevant quantities and the applied procedures are described. Computational fluid dynamics (CFD) calculations are performed by solving the Reynolds-averaged Navier–Stokes (RANS) equations for the assumption of a rigid full-scale and down-scaled wing. These calculations are used to verify the aerodynamic scaling assumptions, which are applied to the scaling procedure of the wind tunnel model. Global aerodynamic coefficients are evaluated for a variety of angles of attack. The local flow phenomena of the full-scale and the scaled model are compared in more detail for a medium and a high angle of attack. The pressure coefficient distribution shows a proper accordance for the full-scale and the scaled model. To verify the results of the structural scaling optimisation, a high-fidelity structural full-scale model is compared with the scaled model using the ELFINI FEM solver. Therefore, all structural components are modelled by 2D elements. The results for the reduced eigenfrequencies and according modes of the full-scale and the scaled model show a high level of similarity. A static deformation of the structural grids is performed by applying the aerodynamic loads from the CFD simulations. The results show that the deviation of the nondimensional deformation between the scaled and the full-scale model is negligible. Consequently, the applied scaling methodology proves to be a valuable tool for the conceptual approach of designing aeroelastically scaled wind tunnel models considering 3D-printed material.

Edge Blending to Enhance the Flow Over Double delta wing configuration during re-entry

During re-entry of space vehicle, the vehicle undergoes flow variation from hypersonic flow to subsonic flow, which causes huge heat transfer over the wing configuration. The double delta wing of configuration 76 degree - 40 degree is considered and effect of blending the wing at edges showed variation and enhancement of flow over the double delta wing configuration. Heat flux calculations were carried out at regions of shock-shock interactions and the effect of blend radius at edges were analyzed using simulations done in ANSYS Fluent. It was observed that aerodynamic efficiency increases with the edge blending which ultimately enhanced the flow field over the double delta wing.

A Novel Design Approach for Low Speed Recovery of High Performance Fighter Aircraft

In this paper, a novel design approach for low-speed recovery of a high-performance fighter aircraft is presented. It is shown that the phugoid mode has an important bearing on the problem of low-speed departure. Based on the analysis of the phugoid mode trajectories, a novel low-speed protection algorithm is presented in this paper. The proposed low speed recovery is achieved in three phases. The first phase consists of detecting the incipient departure followed in the second phase by the application of suitable recovery controls and finally the third phase ends with the transfer of controls to the pilot. The design of the first and the third phase consist of choosing the correct trigger conditions which ensures safe recovery of the aircraft in all conditions. The proposed Automatic low speed recovery is triggered when the aircraft trajectory crosses a fixed boundary in the region spanned by the dynamic pressure and its rate of decrease. It is observed that this boundary is approximately a straight line, implying that it is equivalent to a forward prediction in time to indicate when the aircraft will reach the lowest controllable airspeed. This Automatic Low Speed Recovery with Forward Prediction (ALSR-FP) algorithm is found to be simpler than other existing design methods and effective in preventing low speed departure for a variety of pilot inputs that result in the aircraft losing airspeed leading to stall. In the second phase control inputs are chosen to align the velocity vector to the direction of local gravity. The recovery phase is considered complete after the aircraft reaches the dynamic pressure which is approximately 10 % higher than the minimum dynamic pressure for control. Performance of the ALSR-FP is demonstrated using the high-performance fighter aircraft ADMIRE model which has a delta wing configuration, canards and multiple redundant controls. It is also shown that the proposed algorithm can be easily implemented on board for any other fighter and civil aircraft.

Evaluating Longitudinal Unsteady Aerodynamic Effects in Stall for a T-Tail Transport Model

Although there have been many proposed methods to model unsteady aerodynamic effects in the stall and poststall region, little work has been done to directly assess the impact of unsteady aerodynamic models on stability and control characteristics. In this paper, the state-space method for unsteady aerodynamic modeling is combined with bifurcation analysis to examine the sensitivity of stall and poststall behavior to the choice of aerodynamic modeling method: quasi-steady or unsteady. It is found that quasi-steady modeling can adequately capture the dynamics of the chosen example of a T-tailed transport aircraft with negligible wing–tail coupling. The study is then expanded to investigate a hypothetical situation with highly unsteady aerodynamic characteristics resembling a delta wing configuration—achieved by increasing the time delay constants in the unsteady model. This results in an aircraft with significantly lower flying qualities as indicated by bifurcation analysis. These findings highlight the need to implement unsteady aerodynamic modeling techniques in high-performance aircraft with significant vortex-related unsteady aerodynamics in order to sufficiently capture their stall and poststall dynamics.

The role of California sea lion (Zalophus californianus) hindflippers as aquatic control surfaces for maneuverability.

California sea lions (Zalophus californianus) are a highly maneuverable species of marine mammal. During uninterrupted, rectilinear swimming, sea lions oscillate their foreflippers to propel themselves forward without aid from the collapsed hindflippers, which are passively trailed. During maneuvers such as turning and leaping (porpoising), the hindflippers are spread into a delta-wing configuration. There is little information defining the role of otarrid hindflippers as aquatic control surfaces. To examine Z. californianus hindflippers during maneuvering, trained sea lions were video recorded underwater through viewing windows performing porpoising behaviors and banking turns. Porpoising by a trained sea lion was compared with sea lions executing the maneuver in the wild. Anatomical points of reference (ankle and hindflipper tip) were digitized from videos to analyze various performance metrics and define the use of the hindflippers. During a porpoising bout, the hindflippers were considered to generate lift when surfacing with a mean angle of attack of 14.6±6.3deg. However, while performing banked 180deg turns, the mean angle of attack of the hindflippers was 28.3±7.3deg, and greater by another 8-12deg for the maximum 20% of cases. The delta-wing morphology of the hindflippers may be advantageous at high angles of attack to prevent stalling during high-performance maneuvers. Lift generated by the delta-shaped hindflippers, in concert with their position far from the center of gravity, would make these appendages effective aquatic control surfaces for executing rapid turning maneuvers.

Performance analysis and design of loitering munitions: A comprehensive technical survey of recent developments

Loitering munitions are increasingly used in armed conflicts. An extensive database of loitering munitions is developed based on information available in the public domain. This database includes dimensions, weights, and performance parameters such as flight endurance and communication range. Based upon this dataset, 6 categories of loitering munitions are identified and statistical trends in the form of equations are provided for each category. The statistical trends are supported by aircraft performance theory tailored to loitering munitions applications. Altogether, the combination of the database, statistical trends and aircraft performance theory can be used to analyse the flight performance and design considerations of new loitering munitions of which only limited non-technical information is available in the public domain such as pictures and news articles. Based on the statistical trends and aircraft performance theory it is concluded that for long range applications, the preferred design solution is the conventional configuration. The cruciform configuration is beneficial in case precision flight path control is of prime importance. The tandem wing configuration combines the benefits of a canister launch and relatively high aspect ratio wings suitable for long range flight. Finally, the delta wing design provides a large internal volume and a high terminal attack airspeed. Two example case studies are included to illustrate the flight performance capabilities of two types of loitering munitions used in the current conflict in Yemen (a long range conventional design and a delta wing configuration).

Numerical Simulation of the Streamwise Transport of a Delta Wing Leading-Edge Vortex

Longitudinal vortices are a common flow phenomenon in the flow around aircraft. They emerge wherever sharp edges are encountered and affect the stability of the boundary layers with which they interact. Hence, for an optimal aircraft design, it is necessary to accurately predict not only the formation of these vortices, but also their downstream evolution. This paper presents a numerical simulation approach for the formation and downstream transport of longitudinal vortices. The approach consists in a hybrid Reynolds-averaged Navier–Stokes/large eddy simulation (RANS/LES) method with synthetic turbulence forcing for which different RANS/LES interface locations are investigated. To provide a quantitative analysis, an isolated sharp-edged delta wing configuration is adopted that allows to get rid of unwanted uncertainties. The presented RANS/LES approach is validated against particle image velocimetry data and compared with a wall-modeled LES and to a differential Reynolds stress model simulation. The results show that the RANS simulation is able to accurately predict the formation of the bulk vortical structure. However, its downstream transport can be accurately predicted only by a scale-resolving simulation. The hybrid RANS/LES approach allows to achieve this while saving computational resources by modeling the roll-up of the vortex. Nevertheless, its predictive accuracy highly depends on the appropriate placement of the RANS/LES interface.

The dynamic vortical flow behaviour on a coplanar canard configuration during large-amplitude-pitching

The pitching oscillations (α from 0° to 60°) of a close-coupled coplanar canard configuration were investigated at various reduced frequencies (k=0.0375, 0.075, 0.15, 0.3, 0.6 and 1.2). Both experiments and numerical simulations were carried out to study the differences of dynamic flow characteristics for the configurations with and without a canard (k=0.0375 and 0.6). It was found that lift hysteresis occurred for both a delta wing and canard configuration at various reduced frequencies. At k≥0.6, the orientation of the lift hysteresis loop of the canard became reversed. During the upstroke pitching, a transition phenomenon occurred, in which the canard-wing vortices switched from a counter-rotating vortex system to a co-rotating one. This is due to variation of the effective angle of attack and the canard vortex formed over the lower surface during pitching at k≥ 0.3. As the wing pitched downstroke, the canard vortex and its mirror brought strong downwash flow close to the symmetric plane, which created a high positive pressure region on the upper surface of the wing and eventually led to substantial lift loss.

Correlations of unsteady vortex burst point and dynamic stability over a pitching double-delta wing

Pitching dynamic stability is significant for the aerodynamic performance of an aircraft at a large angle of attack (AoA). Commonly, the dynamic behaviour of vortex breakdown flows over an aircraft with a delta or double-delta wing (DW or DDW) configuration is closely related to the pitching dynamic stability. An 80°/65° DDW subjected to sinusoidal pitching motions around the stall AoA was numerically investigated using the delayed detached eddy simulation method [1]. This study is focused on the correlation between the unsteady behaviour of the vortex burst point (BP) and pitching dynamic stability. The reverse of the pitching dynamic stability is closely related to the inconsistent movement of the BP during the upstroke and downstroke in one pitching period. When the upstream speed of the BP is higher than the downstream speed, the variation in the pitching moment is ahead of the AoA, which leads to dynamic instability. Conversely, when the downstream speed is higher than the upstream speed, the DDW is dynamically stable. This correlation can be interpreted through the energy transfer relationship between the vortex breakdown flows and DDW model.

Influence of leading edge shapes on vortex behaviour of delta wing

This paper analyzes the vortex behaviour of three delta wing configurations namely plane delta wing, saw-tooth delta wing, and sinusoidal tooth delta wing. Experiments such as flow visualization at low speeds, oil flow visualization at different velocities are performed using a low speed wind tunnel. Experiments are conducted for different Reynolds number 0.25, 0.3 and 0.35 million at various angles of attack 10° to 40° in steps of 10°. CFD studies are also carried out by using commercially available software ANSYS FLUENT to have an insight of unsteady nature of vortices. Oil flow visualization study depicts two symmetric separation lines indicating symmetric vortices for all three delta wings at 10° and 20° angles of attack. It was also noticed that at 30° angle of attack plane delta shows asymmetric separation lines indicating asymmetric nature of vortices. Again at 40° angle of attack random pattern of oil traces are evident of flow separation and wake formation on the top surface of the wing. Plane delta has more area of random pattern compared to sinusoidal-tooth and saw-tooth delta wing. Results were authenticated with computational study and it is found that vortices are highly stable at 10° and 20° angles of attack for all planforms. At the same time vortices are found to be flipflopping if angle of attack is increased to 30°, which in turns are governed by unsteady forces due to uneven pressure distribution.

Leading-Edge Vortex Interactions at a Generic Multiple Swept-Wing Aircraft Configuration

Experimental investigations on a generic low-aspect-ratio aircraft configuration with triple and double delta wing planforms were performed at subsonic speed. Flowfield measurements by means of stereoscopic particle image velocimetry and force and moment measurements are used to investigate the system of interacting leading-edge vortices and its effect on the stability characteristics. The analysis comprises the discussion of the overall flowfield and of the vortex system including breakdown locations, vortex core data, and vortex trajectories. Topology considerations in selected crossflow sections highlight the flow patterns of the vortex system. The stability characteristics are discussed by means of the associated aerodynamic moment derivatives. Both the triple and double delta wing configurations feature a flowfield dominated by inboard and midboard vortices. The inboard vortex develops at a nonslender or slender wing section for the triple or double delta wing configurations, respectively, which entails different vortex characteristics and vortex–vortex interaction characteristics. The breakdown behavior associated with the different vortex types influences the stability of the midboard vortex in different ways. Both configurations exhibit pronounced instabilities in the medium to high angle-of-attack regimes, whereas the onset and magnitude of the instability are attenuated for the triple delta wing configuration.

Effect of delta wing on the particle flow in a novel gas supersonic separator

The present work presents numerical simulations of the complex particle motion in a supersonic separator with a delta wing located in the supersonic flow. The effect of the delta wing on the strong swirling flow is analysed using the Discrete Particle Method. The results show that the delta wings re-compress the upstream flow and the gas Mach number decreases correspondingly. However, the Mach number does not vary significantly from the small, medium and large delta wing configurations. The small delta wing generates a swirl near its surface, but has minor influences on the flow above it. On the contrary, the use of the large delta wing produces a strong swirling flow in the whole downstream region. For the large delta wing, the collection efficiency reaches 70% with 2μm particles, indicating a good separation performance of the proposed supersonic separator.

Aerodynamic Flow Characteristics of Utilizing Delta Wing Configurations in Supersonic and Subsonic Flight Regimes

Computational fluid dynamic tests are performed on delta wing models at different heights and speeds in order to achieve lift and drag coefficient values. Primarily, testing was done at supersonic speeds to reveal the advantages of these wing configurations at supersonic flight regimes at a cruise speed and altitude. The low speed characteristics are also examined, important for take-off and landing regimes where the distinctive vortices become prominent. Throughout the two flight conditions tested, a simple delta wing model (with a straight swept wing) is compared to a delta wing model that exhibited a ‘LERX’ (leading edge root extension). Provided literature describes how the performance of delta wings can be improved through this inclusion. Results obtained from the tests show that the model with the LERX has a small, but significant, performance improvement over the simple delta model, in respect to the maximum achievable lift coefficient and maximum stall angle. Lift to drag ratio is not improved however, due to the large vortices creating pressure drag. Generally, the delta wing models produce relatively small amounts of drag, and slightly less lower lift, when at low angles of attack. This is primarily due to the geometry of the models that have thin leading edges and also low thickness to chord ratios.

Numerical study of aerodynamic forces and flow physics of a delta wing in dynamic ground effect

During takeoff and landing processes, the aerodynamic characteristics of a delta wing configuration aircraft are influenced by the dynamic ground effect (DGE). In this paper, the DGE during the landing process of a 65° sweep delta wing (VFE-2) with sharp leading edge is numerically studied at AOA = 23°. The unsteady compressible Reynolds-Averaged Navier–Stokes equations and Spalart–Allmaras turbulence model are discretized using the finite volume method. With the ride height decreasing, the lift, drag and nose-down pitching moment increase nonlinearly, and the increments become larger with the sink velocity increasing. The DGE can be divided into three regions based on the aerodynamics and flow physics. In the large ride height region, the aerodynamic forces in DGE remain unchanged with the ride height decreasing, and they are equal to those in static ground effect (SGE) as long as the angles of attack (AOAs) are equal; the ground effect (GE) can be neglected. In the medium ride height region, the aerodynamic forces in DGE increase slowly with the ride height decreasing, and they are still equal to those in SGE as long as the AOAs are equal; SGE governs the flow. In the small ride height region, the aerodynamic forces in DGE increase significantly with the ride height decreasing, and they are much larger than those in SGE although the AOAs are equal; SGE and compression work effect govern the flow together. Among them, SGE stagnates the airflow under the delta wing, increasing the pressure on the windward surface; simultaneously, it enhances the primary vortex strength, promotes its breakdown and drives it outward along the span. The compression work effect further increases the pressure on the windward surface; however, it has little influence on the airflow over the delta wing.

Summary and Statistical Analysis of the First AIAA Sonic Boom Prediction Workshop

A summary is provided for the First AIAA Sonic Boom Workshop held 11 January 2014 in conjunction with AIAA SciTech 2014. Near-field pressure signatures extracted from computational-fluid-dynamics solutions are gathered from 19 participants (representing three countries) for the two required cases: an axisymmetric body and a simple delta-wing configuration. Structured multiblock, unstructured mixed-element, unstructured tetrahedral, overset, and Cartesian cut-cell methods are used by the participants. Participants provided signatures computed on a series of uniformly refined workshop provided grids and participant-generated and solution-adapted grids. These submissions are propagated to the ground, and noise measures are computed. This allows the grid convergence of a noise measure and a validation metric (difference norm between computed and wind-tunnel-measured near-field signatures) to be studied for the first time. A statistical analysis is also presented for these measures. An optional configuration includes fuselage, wing, tail, flow-through nacelles, and blade sting. More variation in computed noise measures are observed for this full configuration than the required cases. Recommendations are provided for potential improvements to the analysis methods and a possible subsequent workshop.

The effect of edge profile on delta wing flow

This paper presents flow measurements on four delta wing configurations which are differentiated by their leading edge profiles; sharp-edged, small, medium, and large radius. The experiments were performed as a part of the European Vortex Flow Experiment-2 campaign. Tests were conducted at speeds of 20.63 m/s and 41.23 m/s representing Reynolds numbers of 1 × 106 and 2 × 106, respectively. In this paper, oil flow visualization data are presented for the four wings together with particle image velocimetry results for the large radius wing. The study has identified interesting features of the interrelationship between the conventional leading edge primary vortex and the occurrence and development of the inner vortex on the round-edged delta wings. The effects of Reynolds number, angle of attack, and leading-edge radii on both vortex systems are discussed in detail.

DDES study of the aerodynamic forces and flow physics of a delta wing in static ground effect

During takeoff and landing phases, the performance of an aircraft with delta wing configuration is significantly influenced by the ground effect (GE). In this paper, the static ground effect (SGE) of a 65° sweep delta wing (VFE-2) with sharp leading edge at α=20° is investigated by the Delayed Detached Eddy Simulation (DDES) grounded on Spalart–Allmaras (S–A) turbulence model. With the ride height decreasing, the lift, drag and nose-down pitching moment of the delta wing increase nonlinearly, and the majority of aerodynamic force increments in SGE are from the windward side of the delta wing. With the ride height decreasing, under the windward side, the flow is severely blocked resulting in the augmentation of static pressure, and the increasing span wise flow near the leading edge strengthens the shear layer and the leading edge vortex over the leeward side. The strengthened leading edge vortex induces lower pressure over the leeward side, and moves outside along the span wise direction. With the ride height decreasing, the jet-like flow in the primary vortex core is accelerated before the vortex breakdown, and the wake-like flow in the primary vortex core is decelerated after the vortex breakdown.