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In the early era of aviation, Frederick Lanchester was both an inventor and a theoretician driven by the need for a theory of flight that would reduce the guesswork in designing new aircraft. His book Aerodynamics in 1907 laid down the early foundations of such a theory. The theory with contributions from others, notably Ludwig Prandtl, was refined to become the basis for the sleek designs of WWII aircraft brought about with little guesswork. New technology changed aircraft design radically with the increased speed of jet propulsion reaching into the transonic range with nonlinear aerodynamics. In the late 1940s and early 1950s substantial guesswork returned to aircraft design. The legacy of Lanchester et al., however, lived on with the development of computational fluid dynamics (CFD) that could guide designers through nonlinear transonic effects. This article presents a historical sketch of how CFD developed, illustrated with examples explaining some of the difficulties overcome in the design of the first-generation swept-wing transonic fighters. The historical study is forensic CFD in search for the likely explanation of the designer's choice for the wing shape that went into production a long time ago. The capability of current CFD applied to the aerodynamics of aircraft with slender wings is surveyed. The cases discussed involve flow patterns with coherent vortices over hybrid wings and wings of moderate sweep. Vortex-flow aerodynamics pertains to understanding the interaction of concentrated vortices with aircraft components. Modern Reynolds-Averaged Navier-Stokes (RANS) technology is useful to predict attached flow. But vortex interaction with other vortices and breakdown lead to unsteady, largely separated flow which has been found out of scope for RANS. Direct simulation of the Navier-Stokes equations is out of computational reach in the foreseeable future, and the need for better physical modeling is evident. Both cruise performance and stalling characteristics are influenced by strong interactions. Two important aspects of wing-flow physics are discussed: separation from a smooth surface that creates a vortex, and vortex bursting, the abrupt breakdown of a vortex with a subsequent loss of lift. Vortex aerodynamics of not-so-slender wings encounter particularly challenging problems, and it is shown how the design of early-generation operational aircraft surmounted these difficulties. Through use of forensic CFD, the article concludes with two case studies of aerodynamic design: how the Saab J29A wing maintains control authority near stall, and how the Saab J32 mitigates pitch-up instability at high incidence. [ABSTRACT FROM AUTHOR]

The locus of aerodynamic centres of a finite wing is the collection of all spanwise section aerodynamic centres, and depends on aspect ratio, wing sweep and planform shape. This locus is of great importance in the positioning of vortex elements in lifting-line theory. Traditionally, these vortex elements are placed along the quarter-chord of a wing, leading to inaccurate predictions of aerodynamic coefficients for swept wings due to the discontinuity in the line of vorticity at the wing root. An analytical solution was presented by Küchemann in 1956 to determine the locus of aerodynamic centres as a function of sweep. While experimental studies have been performed to visualise this locus, no large amount of data is available to fully evaluate the accuracy of Küchemann's analytical solution. In the present study, a numerical approach is taken using a high-order panel method for inviscid, incompressible flow to calculate the locus of aerodynamic centres for elliptic wings over a wide range of sweep angles, aspect ratios and profile thicknesses. An inviscid panel method is chosen over full CFD solutions because of their ability to isolate the inviscid phenomena. Küchemann's prediction is compared to this numerical data. The root mean square error is calculated for each wing in a broad design space to determine the accuracy of Küchemann's theory. It is shown to be remarkably accurate over the range of cases studied, with the root mean square error staying below 4% for all wings with aft sweep and aspect ratios higher than $R_A=5$. The actual difference between Küchemann's prediction and numerical data is lower than that for the majority of the span for many of the wing designs considered, with the RMS error being skewed by the results at the tip. Results demonstrate that Küchemann's analytical equations can be used as an accurate approximation for the locus of aerodynamic centres and could be used in modern numerical lifting-line algorithms. [ABSTRACT FROM AUTHOR]

A spacetime formulation is presented to solve unsteady aerodynamic problems involving large deformation or topological change such as store separation, slat and flap deployment or spoiler deflection. This technique avoids complex CFD meshing methods, such as Chimera, by the use of a finite-volume approach both in space and time, and permits a locally varying real timestep. The use of a central-difference scheme in the time direction can yield non-physical transient solutions as a consequence of information travelling backwards in time. Therefore, an upwind formulation is provided and validated against one-dimensional and two-dimensional test cases. A hybrid formulation (central in space, upwind in time) is also given and unsteady cases are computed for a spoiler and spoiler/flap deployment, with all three formulations compared, demonstrating that the use of an upwind time stencil yields more representative physical solutions and improves the rate of convergence. [ABSTRACT FROM AUTHOR]

This paper describes a new efficient method for the construction of an approximately balanced aerodynamic Reduced Order Model (ROM) via the frequency domain using Computational Fluid Dynamics data. Time domain ROM construction requires CFD data, which is obtained from the DLR TAU RANS or Euler Linearised Frequency Domain (LFD) solver. The ROMs produced with this approach, using a small number of frequency simulations, are shown to exhibit a strong ability to reconstruct the system response for inviscid flow about the NLR7301 aerofoil and the FFAST wing; and viscous flow about the NASA Common Research Model. The latter demonstrates that the reduced order model approach can reconstruct the full order frequency response of a viscous aircraft configuration with excellent accuracy using a strip wise approach. The time domain models are built using the frequency domain, but also give promising results when applied to reconstruct non-periodic motions. Results are compared to time domain simulations, showing good agreement even with small ROM sizes, but with a substantial reduction in calculation time. The main advantage of the current model order reduction approach is that the method does not require the formation and storage of large matrices, such as in POD approaches. [ABSTRACT FROM AUTHOR]

Compression systems of modern, civil aircraft engines consist of three components: Fan, low-pressure compressor (LPC) and high-pressure compressor (HPC). The efficiency of each component has improved over the last decades by means of rising computational power which made high level aerodynamic optimisations possible. Each component has been addressed individually and separated from the effects of upstream and downstream components. But as much time and effort has been spend to improve performance of rotating components, the stationary inter-compressor duct (ICD) has only received minor attention. With the rotating compression components being highly optimised and sophisticated their performance potential is limited. That is why more aggressive, respectively shorter, ICDs get more and more into the focus of research and engine manufacturers. The length reduction offers high weight saving and thus fuel saving potential as a shorter ICD means a reduction in aircraft engine length. This paper aims at evaluating the impact of more aggressive duct geometries on LPC and HPC performance. A multi objective 3D computational fluid dynamics (CFD) aerodynamic optimisation is performed on a preliminary design of a novel two spool compressor rig incorporating four different operating line and two near-stall (NST) conditions which ensure operability throughout the whole compressor operating range. While the ICD is free to change in length, shape and cross-section area, the blades of LPC and HPC are adjusted for changing duct aerodynamics via profile re-staggering to keep number of free parameters low. With this parametrisation length, reductions for the ICD of up to 40% are feasible while keeping the reduction in isentropic efficiency at aerodynamic design point for the compressor below 1%pt. Three geometries of the Pareto front are analysed in detail focusing on ICD secondary flow behaviour and changes of aerodynamics in LPC and HPC. In order to asses changes in stall margin, speedlines for the three geometries are analysed. [ABSTRACT FROM AUTHOR]

This article focuses on the aerodynamic design of a morphing aerofoil at cruise conditions using computational fluid dynamics (CFD). The morphing aerofoil has been analysed at a Mach number of 0.8 and Reynolds number of $3 \times 10^{6}$ , which represents the transonic cruise speed of a commercial aircraft. In this research, the NACA0012 aerofoil has been identified as the baseline aerofoil where the analysis has been performed under steady conditions at a range of angles of attack between $0^{^{\kern1pt\circ}}$ and $3.86^{^{\kern1pt\circ}}$. The performance of the baseline case has been compared to the morphing aerofoil for different morphing deflections ( $w_{te}/c = [0.005 - 0.1]$) and start of the morphing locations ( $x_{s}/c = [0.65 - 0.80]$). Further, the location of the shock wave on the upper surface has also been investigated due to concerns about the structural integrity of the morphing part of the aerofoil. Based upon this investigation, a most favourable morphed geometry has been presented that offers both, a significant increase in the lift-to-drag ratio against its un-morphed counterpart and has a shock location upstream of the start of the morphing part. [ABSTRACT FROM AUTHOR]

This study is focused on the development of longitudinal aerodynamic models for steady flight conditions. While several commercial solvers are available for this type of work, we seek to evaluate the accuracy of an open source software. This study aims to verify and demonstrate the accuracy of the OpenFoam solver when it is used on basic computers (32–64GB of RAM and eight cores). A new methodology was developed to show how an aerodynamic model of an aircraft could be designed using OpenFoam software. The mesh and the simulations were designed only using OpenFoam utilities, such as blockMesh , snappyHexMesh , simpleFoam and rhoSimpleFoam. For the methodology illustration, the process was applied to the Bombardier CRJ700 aircraft and simulations were performed for its flight envelope, up to M0.79. Forces and moments obtained with the OpenFoam model were compared with an accurate flight data source (level D flight simulator). Excellent results in data agreement were obtained with a maximum absolute error of 0.0026 for the drag coefficient, thus validating a high-fidelity aerodynamic model for the Bombardier CRJ-700 aircraft. [ABSTRACT FROM AUTHOR]

Comprehensive understanding of propulsive jet aerodynamics and aeroacoustics is key to engine design for reduced jet noise and infra-red signature in civil and military aerospace, respectively. Illustrated examples are provided of other aerodynamic/aeroacoustic problems involving jet development, including chevron nozzles, increased jet/wing/flap interference (as fan diameter increases), high acoustic environment (and potentially damaging screech) of supersonic jets on carrier decks and the strongly Three-Dimensional (3D) unsteady flow during the in-ground effect operation of Short Take-Off and Vertical Landing (STOVL) aircraft. To date, laboratory/rig test measurements have primarily been used to identify design solutions; increased use of Computational Fluid Dynamics (CFD) would help achieve cost/time reductions, but Reynolds-Average Navier–Stokes (RANS) CFD with statistical turbulence modelling has proven inadequate for such flows. The scenarios described are far removed from flows used to calibrate model constants, and predictive accuracy demands detailed insight into unsteady flow. Large-Eddy Simulation (LES), whilst computationally more demanding, offers a potential solution. Research undertaken to assess LES capability to address the challenges described is reviewed here. This demonstrates that tremendous progress has been made, indicating that LES can provide sufficiently accurate predictions, representing high value for engineering design. A series of validation studies of increasing realism to practical engineering systems is presented to underpin this conclusion. Finally, areas for further work are suggested to support the combined application of RANS and LES that is probably the optimum way forward. [ABSTRACT FROM AUTHOR]

The addition of wedge-like fairings onto the side of missiles and space launch vehicles, to shield devices such as cameras and reaction jet nozzles, creates additional drag, particularly when in supersonic and hypersonic freestream flow. An experimental and computational study was performed to obtain aerodynamic data on simple representative configurations to test the accuracy of simple theories for the drag increment due to these types of fairings. A semi-empirical method to estimate drag on wedge-shaped projections is presented, which may be used by missile designers to provide predictions of the drag increment due to wedge-like fairings. The method is shown to be valid where the wedge width is much smaller than body diameter, and across the Mach number range 4–8.2 but is likely to be valid for higher Mach numbers. Drag coefficient is found to increase with increasing wedge angle and reducing wedge slenderness, although increasing slenderness tends to increase skin friction drag. [ABSTRACT FROM AUTHOR]

Active flow control for aerofoils has been proven to be an effective way to improve the aerodynamic performance of aircraft. A conceptual hybrid design with surfaces embedded with Shape-Memory Alloy (SMA) and trailing Macro Fibre Composites (MFC) is proposed to implement active flow control for aerofoils. A Computational Fluid Dynamics (CFD) model has been built to explore the feasibility and potential performance of the proposed conceptual hybrid design. Accordingly, numerical analysis is carried out to investigate the unsteady flow characteristics by dynamic morphing rather than using classical static simulations and complicated coupling. The results show that camber growth by SMA action could cause an evident rise of Cl and Cd in the take-off/landing phases when the Angle-of-Attack (AoA) is less than 10°. The transient tail vibration behaviour in the cruise period when using MFC actuators is studied over wide ranges of frequency, AoA and vibration amplitude. The buffet frequency is locked in by the vibration frequency, and a decrease of 1.66–2.32% in Cd can be achieved by using a proper vibration frequency and amplitude. [ABSTRACT FROM AUTHOR]

Lack of flexibility limits the performance enhancement of man-made flapping wing Micro Air Vehicles (MAVs). Active chordwise deformation (bending) is introduced into the flapping wing model at low Reynolds number of Re = 200 in the present study. The lattice Boltzmann method with immersed boundary is adopted in the numerical simulation. The effects of the bending amplitude, bending frequency and phase lag between bending and flapping on the propulsive performance are analysed. The numerical results show that all the chordwise deformation parameters including the bending amplitude, bending frequency and phase lag have a great influence on the flow field, Leading-Edge Vortex (LEV), Trailing-Edge Vortex (TEV) and previous Leading-Edge Vortex (pLEV) of the deformable flapping wing, which leads to the variation of the propulsive performance. With decreasing bending amplitude and increasing bending frequency, both the thrust and energy dissipation coefficients increase. The highest thrust coefficient and highest energy dissipation coefficient occur at a phase lag of 180°. On the other hand, strong dependence of the propulsive efficiency on the vortex tangle is found. The highest propulsive efficiency is obtained for the present model at a dimensionless bending amplitude of 0.2, bending frequency of 0.7Hz, and phase lag of 0°. [ABSTRACT FROM AUTHOR]

Shock control bumps can be used to control and weaken the shock waves that form on engine intakes at high angles of attack. In this paper, it is demonstrated how shock control bumps applied to an engine intake can reduce or eliminate shock-induced separation at high incidence, and also increase the incidence at which critical separation occurs. Three-dimensional Reynolds-average Navier–Stokes (RANS) simulations are used to model the flow through a large civil aircraft engine intake at high incidence. The variation in shock strength and separation with incidence is first studied, along with the flow distribution around the nacelle. An optimisation process is then employed to design shock control bumps that reduce shock strength and separation at a fixed high incidence condition. The bump geometry is allowed to vary in shape, size, streamwise position and circumferential direction around the nacelle. This is shown to be key to the success of the shock control geometry. A further step is then taken, using the optimisation methodology to design bumps that can increase the incidence at which critical separation occurs. It is shown that, by using this approach, the operating range of the engine intake can be increased by at least three degrees. [ABSTRACT FROM AUTHOR]

In this work, consideration is given to a novel concept for aerofoil lift enhancement and delaying flow separation. Here, lift enhancement is attained by preventing the growth of the boundary layer through the elimination of the zero-slip condition between the wing surface and the air stream. The concept would simulate all the effects of a moving wall, leading to the appearance of a slip velocity at the gas–fluid interface, including the injection of momentum into the air boundary layer, but with one exception: here there is no moving wall but instead a ferrofluid thin film pumped parallel and attached to the wall by a magnetic field. Utilising a simplified physical model for the velocity profile of the ferrofluid film and based on ferrohydrodynamic stability considerations, an analytical expression for the interfacial velocity is derived. Finally, from the available experimental data on moving walls, the expected lift and angle-of-attack enhancement are found as well as the weight penalty per unit surface area of the wing is estimated. Additional research and development is required to explore the possibilities of using ferrofluid thin films. [ABSTRACT FROM AUTHOR]

Aircraft handling qualities may be influenced by wing-tip flow separations and horizontal tail (HT) reduced efficiency caused by loss of local dynamic pressure or local tailplane flow separations in high angle-of-attack manoeuvres. From the flight tester's perspective, provided that the test aircraft presents sufficient longitudinal control authority to overcome an uncommanded nose-up motion, this characteristic should not be a safety factor. Monitoring and measuring the local airflow in the aircraft's HT provides information for safe flight-test envelope expansion and data for early aerodynamic knowledge and model validation. This work presents the development, installation and pre-flight calibration using computational fluid dynamics (CFD), flight-test calibration, results and benefits of differential pressure based local angle-of-attack and total pressure measurements through 20 static pressure ports and a Kiel pitot. These sensors were installed in a single-aisle, four-abreast, full fly-by-wire medium-range jet airliner with twin turbofan engines and conventional HT (low vertical position). [ABSTRACT FROM AUTHOR]

Previous studies demonstrated that laminar separation bubbles (LSBs) in the absence of external disturbances or forcing are intrinsically unstable with respect to a three-dimensional instability of centrifugal nature. This instability produces topological modifications of the recirculation region with the introduction of streamwise vorticity in an otherwise purely two-dimensional time-averaged flows. Concurrently, the existence of spanwise inhomogeneities in LSBs have been reported in experiments in which the amplification of convective instability waves dominates the physics. The co-existence of the two instability mechanisms is investigated herein by means of three-dimensional parabolised stability equations. The spanwise waviness of the LSB on account of the primary instability is found to modify the amplification of incoming disturbance waves in the linear regime, resulting in a remarkable enhancement of the amplitude growth and a three-dimensional arrangement of the disturbance waves in the aft portion of the bubble. Present findings suggest that the oblique transition scenario should be expected in LSBs dominated by the convective instability, unless high-amplitude disturbances are imposed. [ABSTRACT FROM AUTHOR]

RANS models remain an attractive turbulence simulation method which could provide some open jet aerofoil interaction analysis at a fraction of the cost of a high-fidelity LES approach. The present work explores the potential and limitations of RANS in this context by simulating an open jet aerofoil noise experiment using the aerospace oriented Menter SST RANS model. This model's tendency to transition at a critical Reynolds number lower than the experimental value was found to impact the boundary layer development. However, the introduction of a low-Re correction improved the prediction of surface pressure and skin friction, enabling the suction surface separation bubble to be captured. The free shear layer's virtual origin characteristics exhibited sensitivity to the interaction with the aerofoil, which can be developed into a metric of the interaction. The main challenge for RANS was accounting for the rise in background disturbance level in the working section, which is caused by the high-turbulence intensity in the free shear layers. [ABSTRACT FROM AUTHOR]

An important prerequisite for the design, assessment and certification of aircraft and their associated control systems is a quantitative specification of the environment in which the aircraft is intended to operate, for example, atmospheric gust. Gust loads on aircraft may induce detrimental influences such as increased aerodynamic and structural loads, structural deformation and decreased flight dynamic performance. This paper presents a systematic and comprehensive overview of important concepts and applications of gust loads on aircraft. This overview includes a brief research background, concepts, research techniques, influences and load alleviation measures of gust. Finally, we summarise some potential improvements in the future work. It is also recommended to learn from previous experiences to avoid aviation accidents due to flight through atmospheric gusts and turbulence. [ABSTRACT FROM AUTHOR]

High-speed aircraft often develop separation-induced leading-edge vortices and vortex flow aerodynamics. In this paper, the discovery of separation-induced vortex flows and the development of methods to predict these flows for wing aerodynamics are reviewed. Much of the content for this article was presented at the 2017 Lanchester Lecture and the content was selected with a view towards Lanchester's approach to research and development. [ABSTRACT FROM AUTHOR]

To accurately predict the probabilities of impact damage to aircraft from runway debris, it is important to understand and quantify the aerodynamic forces that contribute to runway debris lofting. These lift and drag forces were therefore measured in experiments with various bodies spun over a range of angular velocities and Reynolds numbers. For a smooth sphere, the Magnus effect was observed for ratios of spin speed to flow speed between 0.3 and 0.4, but a negative Magnus force was observed at high Reynolds numbers as a transitional boundary layer region was approached. Similar relationships between lift and spin rate were found for both cube- and cylinder-shaped test objects, particularly with a ratio of spin speed to flow speed above 0.3, which suggested comparable separation patterns between rapidly spinning cubes and cylinders. A tumbling smooth ellipsoid had aerodynamic characteristics similar to that of a smooth sphere at a high spin rate. Surface roughness in the form of attached sandpaper increased the average lift on the cylinder by 24%, and approximately doubled the lift acting on the ellipsoid in both rolling and tumbling configurations. [ABSTRACT FROM AUTHOR]

Scaled model test is an effective means to verify the design of a stiffened cylindrical shell. However, there is a problem of similarity distortion by use of the traditional dimensional analysis to design scaled models. In this present study, an equivalent similar method is proposed to solve the problem. The method is applied to an axial stiffened cylindrical shell, and the equivalent criteria and scaling laws satisfying the equivalent similarity of global bending mode are derived and verified by numerical examples. The results indicate that the similarity distortion caused by practical conditions for the stiffened cylindrical shell can be solved and the parameters of scaled model can be designed more freely with the proposed equivalent similar method. [ABSTRACT FROM AUTHOR]

Airframe–propulsion integration concepts that use boundary-layer ingestion (BLI) have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA's STARC-ABL aircraft configuration, which offers the potential for fuel burn reduction by using a turboelectric propulsion system with an aft-mounted electrically driven BLI propulsor. So far, attempts to quantify this potential fuel burn reduction have not considered the full coupling between the aerodynamic and propulsive performance. To address the need for a more careful quantification of the aeropropulsive benefit of the STARC-ABL concept, we run a series of design optimisations based on a fully coupled aeropropulsive model. A 1D thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation to model the aft propulsor at a cruise condition and the effects variation in propulsor design on overall performance. A series of design optimisation studies are performed to minimise the required cruise power, assuming different relative sizes of the BLI propulsor. The design variables consist of the fan pressure ratio, static pressure at the fan face, and 311 variables that control the shape of both the nacelle and the fuselage. The power required by the BLI propulsor is compared with a podded configuration. The results show that the BLI configuration offers 6–9% reduction in required power at cruise, depending on assumptions made about the efficiency of power transmission system between the under-wing engines and the aft propulsor. Additionally, the results indicate that the power transmission efficiency directly affects the relative size of the under-wing engines and the aft propulsor. This design optimisation, based on computational fluid dynamics, is shown to be essential to evaluate current BLI concepts and provides a powerful tool for the design of future concepts. [ABSTRACT FROM AUTHOR]

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine's net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system. [ABSTRACT FROM AUTHOR]

A new wing-tip concept with morphing upper surface and interchangeable conventional and morphing ailerons was designed, manufactured, bench and wind-tunnel tested. The development of this wing-tip model was performed in the frame of an international CRIAQ project, and the purpose was to demonstrate the wing upper surface and aileron morphing capabilities in improving the wing-tip aerodynamic performances. During numerical optimisation with ‘in-house’ genetic algorithm software, and during wind-tunnel experimental tests, it was demonstrated that the air-flow laminarity over the wing skin was promoted, and the laminar flow was extended with up to 9% of the chord. Drag coefficient reduction of up to 9% was obtained when the morphing aileron was introduced. [ABSTRACT FROM AUTHOR]

Wind-tunnel tests were carried out to evaluate the performance of the Computational Fluid Dynamics (CFD)-based air-intake duct shape optimisation of the European platform tiltrotor ERICA. A 1/2.5 scale model including the nacelle, the external portion of the wing and two interchangeable internal ducts reproducing the baseline and optimised shape were manufactered to be tested in the large wind tunnel of Politecnico di Milano. Moreover, tests were carried out with the model equipped with rotating blade stubs. The comprehensive experimental campaign included tests reproducing different forward flight conditions of the aircraft including cruise and conversion phases. The evaluation of the internal duct performance was carried out by measuring total pressure losses and flow distortion by directional probes at the Aerodynamic Interface Plane (AIP). Additional pressure measurements were carried out on the internal surface of the duct to compare the pressure distributions along the air-intake. The experimental results confirmed that the optimised duct offers significantly improved performance with respect to the baseline configuration not only in cruise, representing the flight condition considered for the CFD optimisation, but also for the conversion condition. In particular, a remarkable reduction of the total pressure drop at the AIP was found with the optimised duct with the only exception for the stubs-on configuration in cruise. Indeed, the present investigation highlighted that the design of the blade stubs, particularly their length, represents a very critical aspect for air-intake performance tests due to significant disturbances that could be induced by the stubs’ wake on the internal duct flow. [ABSTRACT FROM AUTHOR]

As simulation has become an integral part of the overall life-cycle support of aircraft, the need for effective Virtual Engineering (VE) tools to support these activities has increased. FLIGHTLAB is a state-of-the-art, aircraft modelling and simulation software tool, that has been designed to address this need and is widely used in rotorcraft design, analysis, test and evaluation, and full-flight simulation applications. This VE tool supports the development and analysis of both fixed and rotary wing aircraft with an extensive library of modelling components which have been successfully used and validated in numerous, real-world applications. These components provide comprehensive modelling of aerodynamic, structural, control and propulsion disciplines. Analyses include performance, dynamic response, stability and control, airloads, and structural loads. Graphical User Interfaces and an interactive scripting language provide user-friendly operation. This paper describes the capabilities and validation activities that have been undertaken to support the development of the commercial VE toolset FLIGHTLAB over the last 20 years and discusses future rotorcraft challenges that could be addressed by enhancements to current generation VE tools. [ABSTRACT FROM PUBLISHER]

Particle Image Velocimetry (PIV) has been used to study the complex flowfield created by simulated battle damage to a two-dimensional wing. Computational Fluid Dynamics (CFD) predictions have also been used for validation of internal cavity flow. Two damage cases were selected for the study; both cases were simulated using a single hole with diameters equal to 20% and 40% of the chord, located at the wing half-chord. Wind-tunnel tests were conducted at a Reynolds number of 500,000 over a range of incidences from 0 to 10° with twocomponent PIV measurements made on three chordwise and three spanwise planes. The PIV data were analysed and compared to CFD data of the same damage cases. The PIV data have shown lower velocity ratios and lower vorticity in the jet compared to past Jet in Cross-Flow experiments and CFD was used to describe the flow features inside the cavity of the wing. It was seen that the wing cavity has large effects on the external flow features, particularly for the 20% damage case. Finally, the flow field data have been related to force balance data. At higher incidence angles, the larger force coefficient increments in both lift and drag can be attributed to the larger wakes and higher jet strengths. [ABSTRACT FROM AUTHOR]

This work is motivated by the need for low-order aerodynamic models to predict accurately the effect on profile drag of controlling attachment line transition. Head's entrainment method(1), a rapid integral boundary layer technique used for design studies on swept wings, suffers from the governing swept-tapered turbulent integral boundary layer equations being ill-posed in the vicinity of the attachment line. This singularity has been treated using crude extrapolations of the attachment-line similarity solution for over half a century, but this approach is unlikely to deliver accurate predictions of the effect of changes in the attachment line flow on profile drag. An experimental study has been carried out to explore the nature of the turbulent flow in the vicinity of a highly swept swept attachment line and has revealed a quite complex, non-monotonic development of the momentum thickness in this region. It has also revealed lower levels of twist in the boundary layer velocity profiles than anticipated from the highly curved character of the inviscid flow streamlines. These observations have prompted an alternative approach to the modelling of the flow in this region which not only successfully eliminates the lack of robustness in the swept-tapered equations but which also matches the experimental results to within ±5%. [ABSTRACT FROM PUBLISHER]

Cranfield University's National Flying Laboratory Centre (NFLC) has developed a Bulldog light aircraft into a flight test facility. The facility is being used to research advanced in-flight instrumentation including fibre optic pressure and strain sensors. During the development of the test bed, Computational Fluid Dynamics (CFD) has been used to assist the flight test design process, including the sensor requirements. This paper describes the development of the Bulldog flight test facility, including an overview of the design and certification process, the in-flight data taken using the installed fibre optic sensor systems and lessons learned from the development programme, including potential further applications of the sensors. [ABSTRACT FROM PUBLISHER]

This paper provides a set of closed form solutions for the lift and drag of wings flying in ground effect both with and without end plates. The developed theories are based on observations of several independent sources of controlled model tests over ground planes and over water and on previous theories prepared by researchers from the original work by Prandtl and Wieselsberger to the present day. The theories developed cover wings of varying aspect ratio, thickness to chord ratios and angle of attack. The results for wings with end plates include the effect of ground (or surface) clearance height, end plate depth and air gap depths beneath the wings or end plates. Good agreement is found between the developed theory and test. [ABSTRACT FROM PUBLISHER]

This design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral and the radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending and torsion) and the high-g manoeuvre pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1.0% in hover, 10.0% at μ = 0.30 and 17.0% at μ = 0.40, relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0.30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre. Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement. [ABSTRACT FROM PUBLISHER]

This paper presents the design and optimisation of an aerofoil with active continuous trailing-edge flap (CTEF) investigated as a potential rotorcraft active control device. Several structural cross-section models are developed: high-fidelity NASA STRucture ANalysis (NASTRAN) and University of Michigan/Variational Asymptotic Beam Section Code (UM/VABS) models and a reduced-order analysis model. The validation of the reduced-order model is established by comparing its predictions of CTEF deformations with those of NASTRAN and UM/VABS analyses, which both show good agreement. The 2D aerodynamic characteristics of the CTEF aerofoil are evaluated using XFOIL and Computational Fluid Dynamics (CFD) analyses: FUN3D and Overset Transonic Unsteady Rotor Navier-Stokes (OVERTURNS). XFOIL, coupled with the reduced-order structure model, is adopted for optimisation study. The accuracy of XFOIL in predicting the aerodynamic pressure of the CTEF aerofoil is verified using CFD simulations, which shows sufficient fidelity. The predicted variations of aerodynamic coefficients with a CTEF angle are compared among the aerodynamic analyses. The optimisation process is developed and applied to two bimorph bender configurations: a Macro-Fibre Composite (MFC) solid bender and an MFC stack bender. The solid bender is used to confirm the functioning of the optimisation procedure and to use its optimal layout as a reference to the stack design, the primary design object. A linear tapered shape is found to be the optimum for a MFC solid bender, which generates an average of 63% more CTEF angles than those of an optimal rectangular bender. An optimised MFC stack bender is shown to resemble the shape of the solid bender. A four-ply bimorph is considered the best choice among the stack layouts because of its large output of CTEF angles and relatively less plies required. The CTEF angle produced by the four-ply optimal layout ranges from 7.6° to 5.3° with speeds from 0 to 200m/s at an angle of attack (AoA) of 6°. The reduction in the CTEF angle with AoA is less steep than that with speed, ranging from 6.5° to 5.8° with AoA from 0 to 8° at speed of 166m/s. An average of 14% increase in CTEF angles is achieved through optimisation for the four-ply bimorph. [ABSTRACT FROM PUBLISHER]

Analysis of the performance of a 1/4.71 model-scale and full-scale Sikorsky S-76 main rotor in hover is presented using the multi-block computational fluid dynamics (CFD) solver of Glasgow University. For the model-scale blade, three different tip shapes were compared for a range of collective pitch and tip Mach numbers. It was found that the anhedral tip provided the highest Figure of Merit. Rigid and elastic full-scale S-76 rotor blades were investigated using a loosely coupled CFD/Computational Structural Dynamics (CSD) method. Results showed that aeroelastic effects were more significant for high thrust cases. Finally, an acoustic study was performed in the tip-path-plane of both rotors, showing good agreement in the thickness and loading noise with the theory. For the anhedral tip of the model-scale blade, a reduction of 5% of the noise level was predicted. The overall good agreement with the theory and experimental data demonstrated the capability of the present CFD method to predict rotor flows accurately. [ABSTRACT FROM PUBLISHER]

Helicopter downwash impact on ship airwake is addressed from a three-pronged approach: (1) Analysis of one-point statistics of autospectrum and two-point statistics of cross-spectrum and coherence from a Computational Fluid Dynamics database of flow velocities effected by helicopter downwash and shipboard airwake; (2) Development of a mathematical framework for extracting interpretive turbulence models in closed form from these autospectral statistics; and (3) Simulation through white-noise-driven filters for the extracted models. The framework begins with an earlier-exercised perturbation-type series expansion of the autocorrelation for all three velocity components, where the first term of the series has a form of the von Karman longitudinal or lateral correlation function. After transformation into equivalent series of autospectrum, the coefficients in the series are evaluated by satisfying theoretical constraints and fitting a curve on a set of selected autospectral data points generated from the database. The framework represents a sensible combination of series expansion, exploitation of a database, and theoretical constraints to provide a foothold on airwake-downwash phenomenon for engineering analysis. It ensures that the extracted model and the autospectral data points have the same time scale, mean square value, and asymptotic decay according to the Kolmogorov –5/3 Law. The framework's strengths and weaknesses, and its major advancement over the earlier series-expansion schemes are also addressed. Finally it is shown that downwash increases airwake energy (mean square value) by one order of magnitude, and almost all of this airwake-downwash energy is concentrated within the bandwidth (0.16 < f(Hz) < 1.6) that affects flight mechanics. [ABSTRACT FROM PUBLISHER]

The fast determination of aerodynamic parameters such as pressure distributions, lift, drag and moment coefficients from the known airflow conditions (angles of attack, Mach and Reynolds numbers) in real time is still not easily achievable by numerical analysis methods in aerodynamics and aeroelasticity. A flight parameters control system is proposed to solve this problem. This control system is based on new optimisation methodologies using Neural Networks (NNs) and Extended Great Deluge (EGD) algorithms. Validation of these new methodologies is realised by experimental tests using a wing model installed in a wind tunnel and three different transducer systems (a FlowKinetics transducer, an AEROLAB PTA transducer and multitube manometer tubes) to determine the pressure distribution. For lift, drag and moment coefficients, the results of our approach are compared to the XFoil aerodynamics software and the experimental results for different angles of attack and Mach numbers. The main purpose of this new proposed control system is to improve, in this paper, wing aerodynamic performance, and in future to apply it to improve aircraft aerodynamic performance. [ABSTRACT FROM PUBLISHER]

An experimental validation of an optimised wing geometry in the Price-Païdoussis subsonic wind tunnel is presented. Two wing models were manufactured using optimised glass fibre composite and tested at three speeds and various angle-of-attack. These wing models were constructed based on the original aerofoil shape of the ATR 42 aircraft and an optimised version of the same aerofoil for a flight condition of Mach number equal to 0.1 and angle-of-attack of 0°. The aerofoil's optimisation was realised using an ‘in-house’ genetic algorithm coupled with a cubic spline reconstruction routine, and was analysed using XFoil aerodynamic solver. The optimisation was concentrated on improving the laminar flow on the upper surface of the wing, between 10% and 70% of the chord. XFoil-predicted pressure distributions were compared with experimental data obtained in the wind tunnel. The transition position was estimated from the experimental pressure data using a second derivative methodology and was compared with the transition predicted by XFoil code. The results have shown the agreement between numerical and experimental data. The wind-tunnel tests have shown that the improvement of the laminar flow of the optimised wing is higher than the value predicted numerically. [ABSTRACT FROM PUBLISHER]

A novel image-based method is presented in this paper to estimate the poses of commercial aircrafts in a runway end safety area. Based on the fact that similar poses of an aircraft will have similar geometry structures, this method first extracts features to describe the structure of an aircraft's fuselage and aerofoil by RANdom Sample Consensus algorithm (RANSAC), and then uses the central moments to obtain the aircrafts' pose information. Based on the proposed pose information, a two-step feature matching strategy is further designed to identity an aircraft's particular pose. In order to validate the accuracy of the pose estimation and the effectiveness of the proposed algorithm, we construct a pose database of two common aircrafts in Asia. The experiments show that the designed low-dimension features can accurately capture the aircraft's pose information and the proposed algorithm can achieve satisfied matching accuracy. [ABSTRACT FROM AUTHOR]

This application paper describes a novel, cluster-based, semi-automatic, stream surface placement strategy for structured and unstructured computational fluid dynamics (CFD) data, tailored towards a specific application: The BLOODHOUND jet and rocket propelled land speed record vehicle. An existing automatic stream surface placement algorithm*81, is extensively modified to cater for large unstructured CFD simulation data. The existing algorithm uses hierarchical clustering of velocity and distance vectors to find potential stream surface seeding locations. This work replaces the hierarchical clustering algorithm, designed to work with small regular grids, with a K-means clustering approach suitable for large unstructured grids. Modifications are made to the seeding curve construction algorithm, improving the smoothness and distribution of the discretised curve in complex cases. A new distance function is described which allows the user to target particular characteristics of simulation data. The proposed algorithm reduces the required memory footprint and computational requirement compared to previous work(8). The performance and effectiveness of the proposed algorithm is demonstrated, and CFD domain expert evaluation is provided describing the value of this approach. [ABSTRACT FROM AUTHOR]

Within an immunity-based architecture for aircraft fault detection, identification and evaluation, a structured, non-self approach has been designed and implemented to classify and quantify the type and severity of different aircraft actuators, sensors, structural components and engine failures. The methodology relies on a hierarchical multi-self strategy with heuristic selection of sub-selves and formulation of a mapping logic algorithm, in which specific detectors of specific selves are mapped against failures based on their capability to selectively capture the dynamic fingerprint of abnormal conditions in all their aspects. Immune negative and positive selection mechanisms have been used within the process. Data from a motion-based six-degrees-of-freedom flight simulator were used to evaluate the performance in terms of percentage identification rates for a set of 2D non-self projections under several upset conditions. [ABSTRACT FROM AUTHOR]

The drag coefficient and the laminar-to-turbulent transition for the aerofoil component of a wing model are optimised using an adaptive upper surface with two actuation points. The effects of the new shaped aerofoils on the global drag coefficient of the wing model are also studied. The aerofoil was optimised with an ‘in-house’ genetic algorithm program coupled with a cubic spline aerofoil shape reconstruction and XFoil 6.96 open-source aerodynamic solver. The wing model analysis was performed with the open-source solver XFLR5 and the 3D Panel Method was used for the aerodynamic calculation. The results of the aerofoil optimisation indicate improvements of both the drag coefficient and transition delay of 2% to 4%. These improvements in the aerofoil characteristics affect the global drag of the wing model, reducing it by up to 2%. The analyses were conducted for a single Reynolds number and speed over a range of angles of attack. The same cases will also be used in the experimental testing of the manufactured morphing wing model. [ABSTRACT FROM AUTHOR]

The article includes study on the influence of a LEX (leading-edge extension) and wing tip plates on aerodynamic characteristics of the Rocket Plane in tailless configuration for manned suborbital flights. The research was conducted only for the low speed regime in the wind tunnel at the Warsaw University of Technology. Tests were carried out for a few configurations of the Rocket Plane. Moreover, two shapes of LEX were investigated. Analysis of results was focused on the aerodynamic synergism effect only. The influence of LEX and the wing tip plates of the Rocket Plane on the aerodynamic coefficients was identified. [ABSTRACT FROM PUBLISHER]

This article examines the increasingly crucial role played by Computational Fluid Dynamics (CFD) in the analysis, design, certification, and support of aerospace products. The status of CFD is described, and we identify opportunities for CFD to have a more substantial impact. The challenges facing CFD are also discussed, primarily in terms of numerical solution, computing power, and physical modelling. We believe the community must find a balance between enthusiasm and rigor. Besides becoming faster and more affordable by exploiting higher computing power, CFD needs to become more reliable, more reproducible across users, and better understood and integrated with other disciplines and engineering processes. Uncertainty quantification is universally considered as a major goal, but will be slow to take hold. The prospects are good for steady problems with Reynolds-Averaged Navier-Stokes (RANS) turbulence modelling to be solved accurately and without user intervention within a decade – even for very complex geometries, provided technologies, such as solution adaptation are matured for large three-dimensional problems. On the other hand, current projections for supercomputers show a future rate of growth only half of the rate enjoyed from the 1990s to 2013; true exaflop performance is not close. This will delay pure Large-Eddy Simulation (LES) for aerospace applications with their high Reynolds numbers, but hybrid RANS-LES approaches have great potential. Our expectations for a breakthrough in turbulence, whether within traditional modelling or LES, are low and as a result off-design flow physics including separation will continue to pose a substantial challenge, as will laminar-turbulent transition. We also advocate for much improved user interfaces, providing instant access to rich numerical and physical information as well as warnings over solution quality, and thus naturally training the user. [ABSTRACT FROM PUBLISHER]

The general picture of research in active flow control for aircraft applications has been continuously changing over the last 20 years. Researchers can now obtain design sensitivities by using numerical flow simulations, and new optical experimental methods can be used that measure flow field data non-intrusively in planes and volumes. These methodological advances enabled significant knowledge increase. The present paper reviews recent progress in active flow control by steady blowing. It appears that two strategies of blowing deserve particular attention. The first uses tangential blowing of thin wall jets to overcome the adverse pressure gradients from locally very large flow turning rates. This approach exploits the potentials of the Coanda effect. The second strategy employs oblique blowing of air jets designed to generate longitudinal vortices in the boundary layer. The longitudinal vortices provide convective redistribution of momentum in the boundary layer, and they also enhance turbulent momentum transport. The sensitivities of these two approaches as observed in fundamental flow investigations and in applications to high-lift aerofoils are described and suited efficiency parameters of blowing are analysed. [ABSTRACT FROM PUBLISHER]

This paper is intended as a general introduction to the requirements for future passenger aircraft design. The needs of the 21st century are addressed to meet the important requirements of the customer airlines as well as those of the general public. In particular, the impact on two traditional major requirements are reviewed, the Design Mission and the operating costs. The effect of aircraft on the environment and the increases in the cost of fuel will have a substantial effect on the way future aircraft are optimised. These demands are summarised before moving on to the basic equations affecting how the aircraft design must respond. Very similar targets driving research work have been set in both Europe and the United States, and some of the new technologies that we can expect to be incorporated are outlined. Finally, a glimpse is given of the possible future aircraft configurations we may see in the skies in response to the new demands. [ABSTRACT FROM PUBLISHER]

Computational aerodynamics, which complement more expensive empirical approaches, are critical for developing aerospace vehicles. During the past three decades, computational aerodynamics capability has improved remarkably, following advances in computer hardware and algorithm development. However, most of the fundamental computational capability realised in recent applications is derived from earlier advances, where specific gaps in solution procedures have been addressed only incrementally. The present article presents our view of the state of the art in computational aerodynamics and assessment of the issues that drive future aerodynamics and aerospace vehicle development. Requisite capabilities for perceived future needs are discussed, and associated grand challenge problems are presented. [ABSTRACT FROM PUBLISHER]

This paper investigates a method for autonomous obstacle avoidance for fixed-wing unmanned aerial vehicles (UAVs), utilising potential fluid flow theory. The obstacle avoidance algorithm needs only compute the instantaneous local potential velocity vector, which is passed to the flight control laws as a direction command. The approach is reactive, and can readily accommodate real-time changes in obstacle information. UAV manoeuvring constraints on turning or pull-up radii, are accounted for by approximating obstacles by bounding rectangles, with wedges added to their front and back to shape the resulting fluid pathlines. It is shown that the resulting potential flow velocity field is completely determined by the obstacle field geometry, allowing one to determine a non-dimensional relationship between obstacle added wedge-length and the corresponding minimum pathline radius of curvature, which can then be readily scaled in on-board implementation. The efficacy of the proposed approach has been demonstrated numerically with an Aerosonde UAV model. [ABSTRACT FROM AUTHOR]

This paper describes a possibility-based multidisciplinary optimisation for electric-powered unmanned aerial vehicles (UAVs) design. An in-house integrated UAV (iUAV) analysis program that uses an electric-powered motor was developed and validated by a Predator A configuration for aerodynamics, weight, and performance parameters. An electric-powered propulsion system was proposed to replace a piston engine and fuel with an electric motor, power controllers, and battery from an eco-system point of view. Moreover, an in-house Possibility-Based Design Optimisation (iPBDO) solver was researched and developed to effectively handle uncertainty variables and parameters and to further shift constraints into a feasible design space. A sensitivity analysis was performed to reduce the dimensions of design variables and the computational load during the iPBDO process. Maximising the electric-powered UAV endurance while solving the iPBDO yields more conservative, but more reliable, optimal UAV configuration results than the traditional deterministic optimisation approach. A high fidelity analysis was used to demonstrate the effectiveness of the process by verifying the accuracy of the optimal electric-powered UAV configuration at two possibility index values and a baseline. [ABSTRACT FROM AUTHOR]

A new method for system stability analysis, the weight functions method, is applied to estimate the longitudinal and lateral stability of a Hawker 800XP aircraft. This paper assesses the application of the weight functions method to a real aircraft and a method validation with an eigenvalues stability analysis of the linear small-perturbation equations. The method consists of finding the weight functions that are equal to the number of differential equations required for system modelling. The aircraft's stability is determined from the sign of the total weight function - the sign should be negative for a stable model. Aerodynamic coefficients and stability derivatives of the mid-size twin-engine corporate aircraft Hawker 800XP are obtained using the in-house FDerivatives code, recently developed at our laboratory of applied research in active controls, avionics and aeroservoelasticity LARCASE. The results are validated with the flight test data supplied by CAE Inc. for all considered flight cases. This aircraft model was chosen because it was part of a research project for FDerivatives code and continued with weight function method for stability analysis in order to develop a design tool, based only on the aircraft geometrical parameters for subsonic regime. The following flight cases are considered: Mach numbers = 0-4 and 0-5, altitudes = 3,000m, 5,000m, 8,000m and 10,000m, and angles-of-attack α = - 5 ° to 20°. [ABSTRACT FROM AUTHOR]

A technique is developed that can be used to predict the effects of battle damage on the aerodynamic performance of an aircraft's wing. The technique is based on results obtained from wind tunnel tests on a NASA LS(1)-0417M0D aerofoil with simulated gunfire damage. The wind tunnel model incorporated an internal cavity to represent typical aircraft construction and this was located between 24% and 75% of chord. The damage was simulated by circular holes with diameters between 20% and 40% of chord. To represent different attack directions, the inclination of the hole axis relative to the aerofoil chord was varied between ±60° pitch and 45° of roll. The aerofoil spanned the wind tunnel to create approximate two-dimensional conditions and balance measurements were carried out at a Reynolds number of 500,000 for incidences, increased in 2° increments, from -4° to 16°. Surface flow visualisation and pressure measurements were also carried out. For a given hole size, the increments in lift, drag and pitching moment coefficients produced trends when plotted against the difference between the upper and lower surface pressure coefficients on the undamaged aerofoil taken at the location of the damage. These trends are used as the basis of the predictive technique. The technique is used to predict the effects of a previously untested damage case, and these are compared with wind tunnel tests carried out on a half model finite aspect ratio wing. For all coefficients the trends in the predicted data are similar to experiment, although there are some discrepancies in absolute values. For the drag coefficient these discrepancies are partly accounted for by limitations in the technique, whilst discrepancies in the lift and pitching moment coefficients are attributed to limitations in the aerofoil test arrangements. [ABSTRACT FROM AUTHOR]

The flight of barnacle geese at airspeeds representing high-speed migrating flight is investigated using experiments and simulations. The experimental part of the work involved the filming of three barnacle geese (Branta Leucopsis) flying at different airspeeds in a wind tunnel. The video footage was analysed in order to extract the wing kinematics. Additional information, such as wing geometry and camber was obtained from a 3D scan of a dried wing. An unsteady vortex lattice method was used to simulate the aerodynamics of the measured flapping motion. The simulations were used in order to successfully reproduce the measured body motion and thus obtain estimates of the aerodynamic forces acting on the wings. It was found that the mean of the wing pitch angle variation with time has the most significant effect on lift while the difference in the durations of the upstroke and downstroke has the major effect on thrust. The power consumed by the aerodynamic forces was also estimated; it was found that increases in aerodynamic power correspond very closely to climbing motion and vice versa. Root-mean-square values of the power range from 100W to 240W. Finally, it was observed that tandem flying can be very expensive for the trailing bird. [ABSTRACT FROM AUTHOR]

Counter-rotating vortices form from the opposite edges of lifting surfaces, and gradually move laterally and dissipate as they travel downstream (as seen in a wing-fixed reference frame). Under ground effect conditions, the vortex from a lifting wing - such as that used in an aircraft application - moves laterally outboard from the wingtip as it progresses downstream; for a downforce wing in ground effect - such as that used in an automotive application - the vortex moves laterally inboard. An interesting case is the situation where the inboard moving vortices become in close proximity to each other. The objective of the present study was to investigate counter-rotating vortices produced from a low aspect ratio downforce wing operating in ground effect. The pair of vortices move towards each other and mutually induce an upwards directed motion which in turn reduces the inboard movement driven by the ground effect. Experimental data gained from three-dimensional Laser Doppler Anemometry in a moving ground wind-tunnel was used to validate a Large Eddy Simulation computational result. [ABSTRACT FROM AUTHOR]