bluuebunny/crossref_metadata_2025 huggingface dataset:
An optimal-fuzzy two-phase CLOS guidance law design using ant colony optimisation
Authors: Nobahari H. , Pourtakdoust S. H.
Published: February 2016

The well-known ant colony optimisation (ACO) meta-heuristic is applied to optimise the parameters of a new fuzzy command to line-of-sight (CLOS) guidance law. The new guidance scheme includes two phases, a midcourse and a terminal phase. In the first phase, a lead strategy is utilised which reduces the acceleration demands. A proportional derivative (PD) fuzzy sliding mode controller is used as the main tracking controller of the first phase. Moreover, a supervisory controller is coupled with the main tracking controller to guarantee the missile flight within the beam. In the terminal phase, a pure CLOS guidance law without lead angle is utilised. For this phase, a new hybrid fuzzy proportional-integral-derivative (PID) fuzzy sliding mode controller is proposed as a high precision tracking controller. The parameters of the proposed controllers for the first and the second phases are optimised using ACO. In this regard, the recently developed continuous ant colony system (CACS) algorithm is extended to multi-objective optimisation problems and utilised to optimise the parameters of the pre-constructed fuzzy controllers. The performance of the resulting guidance law is evaluated at different engagement scenarios and compared with the well-known feedback linearisation method. The comparison is also made in the presence of measurement noise.
Survivability of helicopter with individual blade primary control failure
Authors: Ganguli R. , Jehnert B. , Wolfram J. , Voersmann P.
Published: February 2016

The effect of actuator damage on a helicopter rotor with an IBC based primary control system is studied. Such a system eliminates the swashplate and can be accomplished by trailing-edge flaps, active twist or full authority IBC, especially with smart material actuators. Damage to the collective, longitudinal and lateral cyclic are simulated for one blade, both individually and in combinations ranging from partial damage to complete failure. Numerical results are obtained using a dissimilar blade aeroelastic analysis based on finite elements in space and time for hover and forward speed conditions. It is found that the helicopter can be trimmed for all cases with all three controls having failed on the blade with actuator damage thereby showing that the IBC actuated rotor can survive an actuator failure and can be reconfigured by the pilot using the controls on the other blades. However, in case the collective fails and the longitudinal cyclic is present, there are problems in achieving trim at high damage levels at high forward speeds. Physical explanations of this phenomenon are given. The response (especially flap) for the damaged rotor blades can become high and 1/rev and 2/rev are transmitted by the reconfigured rotor to the hub. Results show that IBC based primary controls provide redundancy which can improve the survivability of a helicopter in case of actuator failure in one blade.
Experimental and computational study of unsteady hypersonic cavity flows
Authors: Creighton S. , Hillier R.
Published: February 2016

This paper presents a combined experimental and computational study of annular cavities on a semi-angle cone in a Mach 8·9 flow. A range of cavity length-to-depth ratios has been considered, and a parameter has been determined that distinguishes between ‘weak oscillations’ and ‘strong oscillations’ of the cavity flow. Essentially the work identifies the transition from the case where the flow can be regarded as ‘pure cavity flow’ to that where the flow behaviour is tending towards that of a ‘spiked blunt body’. The CFD simulations also suggest that, for a certain range of cavity scale, the limiting cavity flow state depends upon the flow initialisation process; it may be weak or strongly oscillating.
Flow control with active dimples
Authors: Dearing S. , Lambert S. , Morrison J.
Published: February 2016

The long-term goal is to design and manufacture optimal ‘on-demand’ vortex generators, ‘dimples’ that can produce vortices of prescribed strength and duration for the real-time control of aerodynamic flows that are either undergoing transition or are fully turbulent, attached or separating. Electro-active polymers (EAP) are ideal for a dimple control surface, offering high strain rate, fast response, and high electromechanical efficiency. EAP can also be used as the basis of a resistanc – or capacitance – change pressure sensor, development of which has just begun. In terms of manufacture, inkjet printing of EAP also offers a paradigm shift such that a monolithic control surface is a very real possibility. Important features for integration into a control system are robustness and a predictable, repeatable motion. With these objectives in mind, the suitability of EAP-based actuators is assessed both mechanically and aerodynamically. The ultimate goal is to integrate these devices, along with shear-stress and pressure sensors and distributed control, also under development, into a flexible ‘smart skin’ which could be incorporated into an airframe structure. The response of a laminar boundary layer to forcing is investiagted using mechanical dimples.
A study of secondary instabilities in postbuckling composite aerostructures
Authors: Falzon B. G. , Cerini M.
Published: February 2016

A number of experimental studies have shown that postbuckling stiffened composite panels, loaded in uniaxial compression, may undergo secondary instabilities, characterised by an abrupt change in the buckled mode-shape of the skin between the supporting stiffeners. In this study high-speed digital speckle photogrammetry is used to gain further insight into an I-stiffened panel’s response during this transient phase. This energy-dissipating phenomenon will be shown to be able to cause catastrophic structural failure in vulnerable structures. It is therefore imperative that an accurate and reliable methodology is available to predict this phenomenon. The shortcomings of current non-linear implicit solution schemes, found in most commercially-available finite element codes, are discussed. A robust and efficient strategy, which utilises an automated quasi-static/pseudo-transient hybrid scheme, is presented in this paper and validated using a number of experimental tests. This approach is shown to be able to predict mode-jumping with good accuracy.
A turbulence model study of separated 3D
jet/afterbody flow
Authors: Hasan R. G. M. , McGuirk J. J. , Apsley D. D. , Leschziner M. A.
Published: February 2016

Three-dimensional RANS calculations and comparisons
with experimental data are presented for subsonic
and transonic flow past a non-axisymmetric
(rectangular) nozzle/afterbody typical of those
found in fast-jet aircraft. The full details of the
geometry have been modelled, and the flow domain
includes the internal nozzle flow and the jet
exhaust plume. The calculations relate to two
free-stream Mach numbers of 0-6 and 0-94 and have
been performed during the course of a collaborative
research programme involving a number of UK
universities and industrial organisations. The close
interaction between partners contributed greatly to
the elimination of computational inconsistencies and
to rational decisions on common grids and boundary
conditions, based on a range of preliminary
computations. The turbulence models used in the
study include linear and non-linear eddy-viscosity
models. For the lower Mach number case, the flow
remains attached and all of the turbulence models
yield satisfactory pressure predictions. However,
for the higher Mach number, the flow over the
afterbody is massively separated, and the effect of
turbulence model performance is pronounced. It is
observed that non-linear eddy-viscosity modelling
provides improved shock capturing and demonstrates
significant turbulence anisotropy. Among the linear
eddy-viscosity models, the SST model predicts the
best surface pressure distributions. The standard
k -ε model gives reasonable
results, but returns a shock location which is too
far downstream and displays a delayed recovery. The
flow field inside the jet nozzle is not influenced
by turbulence modelling, highlighting the
essentially inviscid nature of the flow in this
region. However, the resolution of internal shock
cells for identical grids is found to be dependent
on the solution algorithm -specifically, whether it
solves for pressure or density as a main dependent
variable. Density-based time-marching schemes are
found to return a better resolution of shock
reflection. The paper also highlights the urgent
need for more detailed experimental data in this
type of flow.
The effect of flap-end additions on aircraft trailing vortices
Authors: Graham W. R. , Berteny T. , David L.
Published: February 2016

The influence of flap-end geometry modifications on the near-field wake behind a generic high-lift wing/flap configuration is investigated experimentally with a single-tube yawmeter. It is found that flap-end additions tend to split the concentrated area of streamwise vorticity shed by the flap into two less intense regions, associated with the tips of the flap and the addition. For shorter additions, these regions tend to roll up together, resulting in a slightly more diffuse flap vortex of almost unchanged circulation. For larger additions, the outer region is captured by the tip vortex, which correspondingly gains in circulation at the expense of the flap vortex. In all cases, the vorticity centroid of the rolled-up wake is shifted outboard, and the overall circulation slightly increased. However, the associated lift increase implies a decrease in overall circulation at a given lift coefficient.
Consistent and up-to-date aviation safety targets
Authors: Brooker P.
Published: February 2016

There is a need for consistency between aviation safety targets – target levels of safety (TLS). Consistent ‘risk philosophy’ ensures that resources can be allocated in areas where they will be most beneficial in reducing the number of potential accidents and fatalities. Many existing aviation targets were devised decades ago, have not been brought up to date to take account of the considerable improvement in the sector’s safety performance, and are not targeted at some specific future year. It appears feasible to construct consistent and up-to-date sets of TLSs covering the full range of aviation needs, in particular for air traffic management. However, this requires significant, more demanding, changes to some TLS values. One caveat to note is that the safety of ‘loosely coupled’ aviation systems may be better described and managed by the Health and Safety Executive version of risk assessment rather than by a TLS approach.
In-flight Ice formation simulation on finite wings and air intakes
Authors: Özgen S. , Canıbek M.
Published: February 2016

In the present article, in-flight ice formation on finite wings and air intakes of low-speed aircraft are numerically studied. The approach to the problem involves calculation of the velocity field using a three-dimensional panel method. Using the calculated velocity field, the droplet trajectories and droplet impact locations are computed yielding the droplet collection efficiency distribution. In the next step, convective heat transfer coefficient distributions around the geometries are calculated using a two-dimensional Integral Boundary-Layer Method, which takes surface roughness due to ice accretion into account. A thermodynamic analysis employing the Extended Messinger Method yields the ice growth rates. Integration of these rates over time yields the ice shapes, hence the modified geometry. Predicted ice shapes are compared with experimental shapes reported in the literature and good agreement is observed. Ice shapes around vastly varying geometries including complex shapes are successfully computed. As such, the developed tool may be used for academical purposes or for airworthiness certification efforts.
Influence analysis of measurement errors in satellite attitude determination based on extended Kalman filter
Authors: Jiao Y. , Wang J. , Pan X. , Zhou H.
Published: February 2016

The satellite attitude determination approach based on the Extended Kalman Filter (EKF) has been widely used in many real applications. However, the accuracy of this method largely depends on the fitness of measurement model. We aim to analyse the influence of measurement errors to the accuracy of EKF based attitude determination approach in this paper. The measurement errors, which are divided into structural error and nonstructural error by their influences, are analysed in principle. In the setting of the combination of star sensors and gyros, according to the property of innovation, we employ the technique of correlation test to analyse the influences of different kinds of measurement errors. Experimental results demonstrate the effectiveness of our previous analysis.
Aerodynamics of an aerofoil in transonic ground effect: numerical study at full-scale Reynolds numbers
Authors: Doig G. , Barber T. J. , Neely A. J. , Myre D. D.
Published: February 2016

The potential positive effects of ground proximity on the aerodynamic performance of a wing or aerofoil have long been established, but at transonic speeds the formation of shock waves between the body and the ground plane would have significant consequences. A numerical study of the aerodynamics of an RAE2822 aerofoil section in ground effect flight was conducted at freestream Mach numbers from 0·5 to 0·9, at a range of ground clearances and angles of incidence. It was found that in general the aerofoil’s lifting capability was still improved with decreasing ground clearance up until the point at which a lower surface shock wave formed (most commonly at the lowest clearances). The critical Mach number for the section was reached considerably earlier in ground effect than it would be in freestream, and the buffet boundary was therefore also reached at an earlier stage. The flowfields observed were relatively sensitive to changes in any given variable, and the lower surface shock had a destabilising effect on the pitching characteristics of the wing, indicating that sudden changes in both altitude and attitude would be experienced during sustained transonic flight close to the ground plane. Since ground proximity hastens the lower surface shock formation, no gain in aerodynamic efficiency can be gained by flying in ground effect once that shock is present.
Kinematic and experimental aerodynamic characterisation of the RotaFlap – a novel flapping wing mechanism
Authors: Ania A. , Poirel D. , Potvin M.-J.
Published: February 2016

A unique patented mechanism, termed ‘RotaFlap’, which can move its wings in a figure-eight shape kinematically similar to insects or hummingbirds, has been investigated through the design, construction, integration and testing of various prototypes. In this paper, the most recent prototype is presented whereby the RotaFlap kinematics is characterised to understand some of its most pertinent parameters. A host of variations have been identified and in this study a subset of these have been tested. A preliminary characterisation of the force production, especially the vertical lift coefficient, has been completed. It is concluded that this mechanism produces vertical lift coefficient values similar to insect and hummingbird flight for similar Reynolds numbers.
Thermomechanical behaviour of a damaged thermal protection system: experimental correlation and influence of hypersonic flow
Authors: Ng W. H. , Friedmann P. P. , Waas A. M. , McNamara J. J.
Published: February 2016

This paper describes a combined experimental and numerical study on damaged and undamaged space shuttle tile thermal protection system (TPS). The principal objective of the study is to determine its thermomechanical behaviour and assess the structural integrity of the TPS. The TPS tile specimens are subjected to a temperature profile corresponding to the thermal loads of the Access to Space reference vehicle. Experiments are conducted in a vacuum chamber that allows re-entry static pressure to be simulated. Temperatures on the top and bottom surfaces of the specimen, and the strains in the underlying structure are recorded. The experimental results are used to guide the development of a refined finite element model, which is subsequently used to simulate the interactions between the high speed external flow past the cavity that represents damage. Using this model, the relative effects of damage on the thermal protection capability and the induced thermal stresses are determined by comparing the response of the damaged configurations with the undamaged configuration. Damage increases the thermal loads and significantly reduces the radiation heat loss from the surface of the tile, resulting in elevated temperatures. Results indicate that damage can raise the maximum temperature in the tile to values that exceed its melting point.
Heat transfer characteristic modelling and the effect of operating conditions on re-cooled cycle for a scramjet
Authors: Bao W. , Qin J. , Zhou W. X.
Published: February 2016

A re-cooled cycle has been proposed for a regeneratively cooled scramjet to reduce the hydrogen fuel flow for cooling. Upon the completion of the first cooling, fuel can be used for secondary cooling by transferring the enthalpy from fuel to work. Fuel heat sink (cooling capacity) is thus repeatedly used and fuel heat sink is indirectly increased. Instead of carrying excess fuel for cooling or seeking for any new coolant, the cooling fuel flow is reduced, and fuel onboard is adequate to satisfy the cooling requirement for the whole hypersonic vehicle. A performance model considering flow and heat transfer is build. A model sensitivity study of inlet temperature and pressure reveals that, for given exterior heating condition and cooling panel size, fuel heat sink can be obviously increased at moderate inlet temperature and pressure. Simultaneously the low-temperature heat transfer deterioration and Mach number constrains can also be avoided.
On the generation of a helicopter aerodynamic database
Authors: Raffel M. , De Gregorio F. , de Groot K. , Schneider O. , Sheng W. , Gibertini G. , Seraudie A.
Published: February 2016

The GOAHEAD (Generation of an Advanced Helicopter Experimental Aerodynamic Database for CFD code validation) consortium was created in the frame of an EU-project in order to create an experimental database for the validation of 3D-CFD and comprehensive aeromechanics methods for the prediction of unsteady viscous flows. This included the rotor dynamics for complete helicopter configurations, i.e. main rotor – fuselage – tail rotor configurations with emphasis on viscous phenomena like flow separation and transition from laminar to turbulent flow. The wind tunnel experiments have been performed during two weeks in the DNW-LLF on a Mach-scaled model of a modern transport helicopter consisting of the main rotor, the fuselage, control surfaces and the tail rotor. For the sake of controlled boundary conditions for later CFD validation, a closed test section has been used. The measurement comprised global forces of the main rotor and the fuselage, steady and unsteady pressures, transition positions, stream lines, position of flow separation, velocity profiles at the test section inlet, velocity fields in the model wake, vortex trajectories and elastic deformations of the main and tail rotor blades.
Subsonic aerofoil design 2010
Authors: Somers D. M.
Published: February 2016

The state of the art as practiced by a handful of aerofoil designers is discussed. The methods used, both theoretical and experimental, are described. Aerofoil/application design integration and expansion of the design envelope to lower and higher Reynolds numbers are illustrated by examples, including the slotted, natural-laminar-flow aerofoil.
Study of jets from rectangular nozzles with square grooves
Authors: Kumar P. Arun , Verma S. B. , Elangovan S.
Published: February 2016

An experimental study has been carried out to understand jet flow development from plain and grooved rectangular nozzles of aspect ratio 2:1 using two-component hotwire anemometry. Grooves of square configuration (side 4mm) and length 5mm were introduced in the (i) minor-axis, (ii) major-axis and, (iii) in both minor- and major-axes directions. The equivalent diameter of the plain rectangular nozzle is 37·5mm. Studies were carried out for a nominal jet exit velocity of 20ms−1 and for Reynolds number based on equivalent diameter of 54,000. The introduction of grooves in either plane does not show any influence on the potential-core length but results in faster jet-decay thereafter. It is observed that the grooves when introduced in minor-axis direction inhibit the jet growth in that plane while promoting the jet growth along major-axis plane and hence, delays the phenomena of axis-switching. However when introduced in major-axis direction, the grooves promote jet growth along major-axis plane while inhibiting jet-growth in minor-axis plane. Cross-sectional contours of mean-velocity suggest that the grooves modify the process of overall jet development significantly in the plane in which they are introduced relative to the plain jet and hence, significantly affect the axis-switching location in each case.
Dynamics of the wing-tip vortex in the near field of a NACA 0012 aerofoil
Authors: del Pino C. , López-Alonso J.M. , Parras L. , Fernandez-Feria R.
Published: February 2016

The dynamics of the wing tip vortex in the near-field of a NACA 0012 aerofoil has been analysed by means of flow visualisations in a water tunnel. Different axial distances near the wing up to four chords, Reynolds numbers up to 42,000 and three angles-of-attack are studied to characterise the behaviour of the vortex meandering. The spatio-temporal vortex centre positions show distorted elliptical shapes in a (x,y)-plane. The Reynolds number has no significant influence on the axial evolution of the meandering amplitude. In addition, the flow visualisations obtained with a low speed camera are analysed by the singular value or proper orthogonal decomposition. Thus, the most energetic displacement modes are obtained. The frequency associated to these modes is computed by FFT. In all the cases studied, our results show that the most unstable mode corresponds to the azimuthal wavenumber |n| = 1 in the so-called Kelvin helical modes and the frequency is lower or close to 1Hz.
New helicopter model identification method based on flight test data
Authors: De Jesus Mota S. , Botez R. M.
Published: February 2016

A helicopter model has been identified and validated for flight conditions defined by altitudes, speeds, loadings and centre of gravity positions. To identify the helicopter models, 2-3-1-1 multistep control inputs were performed by the pilot to excite all helicopter modes. Then, each estimated signal has to remain in tolerance margins defined by the Federal Aviation Administration. Three methods were used to observe the system outputs from its states: a fuzzy logic method, a linear method optimised with a neural network algorithm and a classical method. Because of random effects when gathering data, classical method did not give good enough results. The fuzzy logic method was not robust enough so that output plots showed peaks that could be felt by the pilot. Then, because the model could be implemented in a simulator for the pilot training, the pilot feedback is very useful in order to compare the reality with the results of the mathematical model. When the outputs are obtained from the measured state variables, the linear method gave the best results.
Wind-tunnel tests of a tilt-rotor aircraft
Authors: Gibertini G. , Auteri F. , Campanardi G. , Macchi C. , Zanotti A. , Stabellini A.
Published: February 2016

A wide aerodynamic test campaign has been carried out on the tiltrotor aircraft ERICA at the Large Wind Tunnel of Politecnico di Milano by means of a modular 1:8 scale model in order to produce a dataset necessary to better understand the aerodynamic behaviour of the aircraft and to state its definitive design.

The target of the tests was the measurement of the aerodynamic forces and moments in several different configurations and different attitudes. The test program included some conditions at very high incidence and sideslip angles that typically belong to the helicopter-mode flight envelope and measurements of forces on the tail and on the tilting wings.

A large amount of data has been collected that will be very useful to refine the aircraft design. In general the aircraft aerodynamics do not present any critical problems, but further optimisation is still possible. From the viewpoint of drag in the cruise configuration, the sponsons of the landing gear seem to be worth some further design refinement since they are responsible for a 20% drag increase with respect to the pure fuselage configuration. On the contrary, the wing fairing has proved to work well when the aircraft longitudinal axis is aligned with the wind, providing just a slight drag increase.

Two other interesting aspects are the quite nonlinear behaviour of the side force for the intermediate sideslip angles as well as the noticeable hysteresis in the moment coefficient at very high incidence angles.
Modification impact on aerodynamic performance of hypersonic waverider
Authors: Xiao-Qing C. , Zhong-Xi H. , Jian-Xia L. , Xian-Zhong G.
Published: February 2016

Waverider serves as a good candidate for hypersonic vehicles. The typical waverider has sharp leading edge and no control face, which is inappropriate for practical use. This paper puts forward a method modifying the waverider, and the modification impact on the performance of waverider at hypersonic flow conditions is studied. The modification is based on blunted waverider, includes cutting the tip and introducing two control wings. The modification’s effect on aerodynamic performance is obtained and analysed through Computational Fluid Dynamics (CFD) techniques. When blunted with 2cm radius, the waverider retains its good aerodynamic performance and the heat flux at the stagnation point can be managed. Three factors of the introduced wing are argued, the fixed angle, aspect ratio and wing area. Results show that influence on the aerodynamic coefficient is slight and the vehicle retains its high lift-to-drag ratio. The main influences of the modification are the control ability and trim efficiency, which is the motivation of this work and can be adapted when designing a practical waverider.
Emission comparison of turbofan and open rotor engines under special consideration of aircraft and mission design aspects
Authors: Seitz A. , Schmitt D. , Donnerhack S.
Published: February 2016

An integrated parametric model involving the design of propulsion system, airframe and flight mission is presented. Based hereon, the carbon dioxide (CO2) emission characteristics of advanced direct-drive turbofan and open rotor powered aircraft are analysed against pertinent aircraft and propulsion system design parameters. In addition, initial concept-specific trend statements on nitrogen oxides (NOx) as well as propulsor noise emission characteristics are derived. The obtained results contribute to a better understanding of more appropriate aircraft design attributes for advanced system architectures.
Applying heat pipes to a novel concept aero engine: Part 1 – Design of a heat-pipe heat exchanger for an intercooled aero engine
Authors: Camilleri R. , Ogaji S. , Pilidis P.
Published: February 2016

Civil aviation has instilled new perceptions of a smaller world, creating new opportunities for trade, exchange of cultures and travelling for leisure. However, it also brought with it an unforeseen impact on the environment. Aviation currently contributes to about 3·5% of the global warming attributed from human activities. With the forecasted rate of growth, this is expected to rise to about 15% over the next 50 years. Although it is projected that the annual improvements in aircraft fuel efficiency are of the order of 1-2%, it is suggested that the current gas turbine design is fully exploited and further improvements are difficult to achieve. A new generation of aero engine core concepts that can operate at higher thermal efficiencies and lower emissions is required. One possibility of achieving higher core efficiencies is through the use of an inter-cooled (IC) core at high overall pressure ratios (OPR). The concept engine, expected to enter into service around 2020, will make use of a conventional heat exchanger (HEX) for the intercooler. This paper seeks to introduce a heat pipe heat exchanger (HPHEX) as an alternative design of the intercooler. The proposed HPHEX design takes advantage of the convenience of the geometry of miniature heat pipes to provide a reduction in pressure losses and weight when compared to conventional HEX. The HPHEX will be made of a number of stages, each stage being made of a large number of miniature heat pipes in radial configuration, that will extend from the inter-compressor duct to the bypass split, thus eliminating any ducting to and from the intercooler. This design offers up to 32% reduction in hot pressure losses, 34% reduction in cold pressure losses and over 41% reduction in weight.
Pilot task demand load during RNAV approaches with a Cessna Citation
Authors: Heiligers M. M. , van Bennekom K. T. P. , van Tuinen T. J. , van Holten Th. , Mulder M.
Published: February 2016

This research aims to develop a method which predicts the task demand load as experienced by pilots while flying an area navigation (RNAV) approach. First, this will yield insight in which aspects of an approach actually influence pilot task demand load. And second, during the design of approaches this method can be used to rapidly evaluate a potential approach and to ‘optimise’ an approach with respect to pilot task demand load. During previous research, focusing on approaches flown with a B747, a list of factors that influence pilot task demand load has been obtained, as well as a method to keep pilot task demand load at an acceptable level. The method consists of seven guidelines to be adhered to during approach design. This paper shows that the list of factors and the method do not only apply to a B747 aircraft but are generally applicable to other aircraft as well. This is underpinned by results from both flight simulator tests and real flight tests with TU Delft’s Cessna Citation laboratory aircraft. Additionally, it is shown that there are no discrepancies between the list of factors influencing pilot task demand load resulting from the flight simulator tests and the list of factors resulting from the real flight tests.
Investigation of the three-dimensional flow over a 40° swept wing
Authors: Zhang S. , Jaworski A. J. , Turner J. T. , Wood N. J.
Published: February 2016

Three-component laser Doppler anemometry (LDA) has been used to measure the complex flow distributions over the suction surface of a symmetrical 40°swept wing model at an angle of incidence of 9° and for the relatively low Reynolds number (based on the root chord) of 2·1 × 105. Emphasis was placed on the separation and reattachment of the boundary-layer flow, and the formation of vortical structures.

The experimental programme now described utilised a large wind tunnel and several advanced measurement techniques to produce an unusually detailed collection of results. Thus, data are presented here for the behaviour of the three-dimensional boundary layer developing on the suction surface at positions from 30% to 90% semi-span. These results show the spatial variations of the time-averaged mean and fluctuating velocity components in three orthogonal directions, including the distributions of the normal and shear stress levels. Further analysis has enabled the time-averaged vortical structures to be identified and compared with the results of surface flow visualisation.

Flow over a swept wing poses many challenges for the computational modeller. The underlying aim, here, therefore, was to produce a data archive for the validation of computer predictions. As will be shown, some of the experimental results have already been used in this way and the data base is available for use by other workers.
Aerodynamic shape optimisation of hovering rotors using compressible CFD
Authors: Allen C. B. , Rendall T. C. S.
Published: February 2016

Generic wrap-around shape optimisation technology is presented, and is applied to a helicopter rotor in hover, using compressible CFD as the aerodynamic model. An efficient domain element shape parameterisation method is used as the surface control and deformation method, and is linked to a radial basis function global interpolation, to provide direct transfer of domain element movements into deformations of the design surface and the CFD volume mesh. This method is independent of mesh type and size, and optimisation independence from the flow solver is achieved by obtaining sensitivity information for an advanced parallel gradient-based algorithm by finite-difference. The method is applied here to a two-bladed hovering rotor, with transonic tip speed, comparing the effects of global twist parameters and more local planform parameters. Significant torque reductions are achieved in both cases, and it is shown that using only three global twist parameters are extremely effective. Using 63 local and global parameters, large geometric changes are demonstrated, with both induced power and wave drag reduced.
Reactive route selection from pre-calculated trajectories – application to micro-UAV path planning
Authors: Hall J. , Anderson D.
Published: February 2016

Operating micro-UAVs autonomously in complex urban areas requires that the guidance algorithms on-board are robust to changes in the operating environment. Limited endurance capability demands an optimal guidance algorithm, which will change as the environment does. All optimal path-planning routines are computationally intensive, with processor load a function of the environmental complexity. This paper presents a new algorithm, the reactive route selection algorithm, for storing a bank of optimal trajectories computed off-line and blending between these optimal trajectories as the operating environment changes. An example is presented using a mixed-integer linear program to generate the optimal trajectories.
Computational study of flowfield characteristics in cavities with stores
Authors: Khanal B. , Knowles K. , Saddington A. J.
Published: February 2016

In this paper, the results of computational studies on the unsteady flow features in three-dimensional empty cavities and cavities with a representative store are presented. Flow simulations with a turbulence model based on a hybrid method, which behaves as a standard Reynolds-averaged Navier-Stokes (RANS) model within the attached boundary layer and as a Large-Eddy Simulation LES sub-grid scale model in the rest of the flow (commonly known as Detached-Eddy Simulation (DES)) are used in this study. The time-mean flow study showed the presence of three-dimensional effects inside the cavities. The mean flowfield visualisation also clearly showed the presence of a pair of ‘tornado-like’ vortices in the upstream half of the cavity which merge to a single, large recirculation further downstream. Visualisation for the cavity-with-store case revealed that the mean flowfield was effectively divided into two halves with significant reduction of the spanwise flow across the cavity width. In the unsteady flow study, near-field acoustic spectra were computed for the empty cavity and cavity-with-store cases. Study of unsteady pressure spectra for the cavity-with-store case found the presence of many peaks and the corresponding mode frequencies were found to agree well with the Rossiter modes. The blockage effect of store and strut on the spanwise flow is thought to have reduced the interaction, and subsequent non-linear coupling, between the Rossiter modes. This may be the reason for the co-existence of multiple modes without the coupling among them.
Parametric studies of aerofoil-vortex interaction; a numerical approach using LES
Authors: Ilie M. , Nitzsche F. , Matida E.
Published: February 2016

The influence of angle-of-attack and vortex-aerofoil vertical miss-distance on the aerofoil-vortex mechanism of interaction is investigated numerically using the large-eddy simulation approach. The study concerns the aerofoil-vortex interaction (AVI) phenomena encountered in the rotorcraft aerodynamics. The studies are performed for a Reynolds number of 1·3 × 106 based on the chord (c = 0·2m) of the aerofoil and free-stream velocity U = 100ms−1. The studies are conducted for three different angles-of-attack (α = 0°, and α = 5°, α = 10°) and vertical miss-distances (h = 0·0m, h = −0·01, h = −0·02m) respectively. The present study shows that the magnitude of lift coefficient and flow separation, at the instant of interaction, decay with the increase of angle-of-attack and vertical miss-distance.
A neuro-fuzzy approach to the weight estimation of aircraft structural components
Authors: Hannon C. , Toropov V. V. , Querin O. M.
Published: February 2016

This paper explores the issues related to the application of fuzzy logic techniques to aid the process of weight estimation for aircraft structures at preliminary design stages. The focus lays on the design of a neuro-fuzzy system for the weight analysis, through the use of the Neuro-Fuzzy Function Approximator (NEFPROX) algorithm. The paper introduces a three-level process designed around Mamdani fuzzy systems derived through NEFPROX and analyses its application to the sizing and weight estimation of spoiler attachment ribs. Problems such as structure parameterisation, variable selection and model optimisation are explored as part of the process validation phase on the selected structural case study. The model performance is evaluated with respect to modelling accuracy, generalisation capabilities and the interdependencies between the variables the model is able to derive from the analysis of the given structural examples. The results are then compared to the performance obtained from the application of Takagi-Sugeno-Kang (TSK) fuzzy models derived using Adaptive Network-based Fuzzy Inference System (ANFIS) on the same sample of spoiler attachment ribs. Results highlight the benefits of adopting NEFPROX for the derivation of fuzzy systems to be applied in structural weight estimation problems, from the point of view of both accuracy of the approximation provided and the quality and interpretability of the final rulebase derived by the system.
Damage arrest design approach using stitched composites
Authors: Velicki A. , Thrash P.
Published: February 2016

Although the deployment of carbon fibre structural components has enjoyed wide success on smaller aircraft, their acceptance on larger transport airframes is less sanguine – especially in the case of primary structure applications where the increasing out of plane loads found on the larger airframes have exposed the weak interlaminar properties of the layered material system. This has led to an overreliance on mechanical attachments to suppress these through thickness failures, which ultimately degrades structural performance and increases manufacturing costs. Until these resin dominated failure modes can be dealt with more effectively, without adversely affecting the in plane properties of the laminate, the true weight-savings potential afforded by carbon fibre material systems will be difficult to attain. This paper describes how researchers at NASA-LaRC and The Boeing Company are working to develop a next generation stitched composite design solution that addresses the fundamental challenges in achieving improved structural performance and reduced fabrication costs for large carbon fibre airframe structures.
Robust geometric sizing of a small flying wing planform based on evolutionary algorithms
Authors: Rodríguez-Cortés H. , Arias-Montaño A.
Published: February 2016

In this paper a geometric sizing method for a small electric powered flying wing is proposed. The geometric sizing method aims to reduce the effects of variations in the power plant characteristics on endurance. This results in a single-objective design optimisation problem where the sensitivity to power plant characteristics of the endurance equation is minimised, constrained to Reynolds number, wing load, wing taper ratio, aircraft size and wing sweep angle. As a result, geometric characteristics of the flying wing such as span, tip chord and root chord are obtained. Flying wing aerodynamic characteristics are obtained by means of an inviscid fluid flow analysis program of the type low-order panel methods, known as CMARC. The optimisation problem involves a non convex function so that it is necessary to rely on heuristic programming methods. In particular an Evolutionary Algorithm based on differential evolution is considered.
An optimal control approach for alleviation
of tiltrotor gust response
Authors: Muro D. , Colella M. Molica , Serafini J. , Gennaretti M.
Published: February 2016

The alleviation of gusts effects on a tiltrotor in
aeroplane and helicopter operation modes obtained by
an optimal control methodology based on the
actuation of elevators, wing flaperons and
swashplate is examined. An optimal observer for
state estimate is included in the compensator
synthesis, with the Kalman-Bucy filter applied in
the presence of stochastic noise. Tiltrotor dynamics
is simulated through an aeroelastic model that
couples rigid-body motion with wing and proprotor
structural dynamics. An extensive numerical
investigation examines effectiveness and robustness
of the applied control procedure, taking into
account the action of both deterministic and
stochastic vertical gusts. In addition, a passive
pilot model is included in the aeroelastic loop and
the corresponding effects on uncontrolled and
controlled gust response are analysed.
Performance of Glare panels subjected to
intense pressure pulse loading
Authors: Soutis C. , Mohamed G. , Hodzic A.
Published: February 2016

A robust and efficient computational model has been
developed which is capable of modelling the dynamic
non-linear behaviour of Glare panels subjected to
blast loadings. High strain rate material
characterisation and modelling of interfacial
debonding between adjacent sublaminates have been
taken into consideration. Numerical model validation
have been performed considering case studies of
Glare panels subjected to a blast-type pressure
pulse for which experimental data on the back face-
displacement and post-damage observations were
available. Excellent agreement of mid-point
deflections and evidence of severe yield line
deformation were shown and discussed against the
performed blast tests. A further parametric study
identified Glare as a potential blast attenuating
structure, exhibiting superior blast potential
against monolithic aluminium plates. The results
were normalised and showed that for a given impulse,
Glare exhibited a smaller normalised displacement,
outperforming monolithic Aluminium 2024-T3 plates.
It was concluded that further work needed to be
carried out to take into account the influence of
geometry (cylindrical structures),
pre-pressurisation effects and boundary
conditions
Parametric damage tolerance design of metallic aeronautical stiffened panels
Authors: Molinari G. , Meneghin I. , Melega M. , Troiani E.
Published: February 2016

On the basis of well-known literature, an analytical tool named LEAF (linear elastic analysis of fracture) was developed by the authors to predict the damage tolerance (DT) proprieties of aeronautical stiffened panels. The tool is based on the linear elastic fracture mechanics and the displacement compatibility method. By means of LEAF, an extensive parametric analysis of stiffened panels, representative of typical aeronautical constructions, was performed to provide meaningful design guidelines. The effects of riveted, integral and adhesively bonded stringers on the fatigue crack propagation performances of stiffened panels were investigated, as well as the crack retarder contribution using metallic straps (named doublers) bonded in the middle of the stringers bays. The effect of both perfectly bonded and partially debonded doublers was investigated as well. Adhesively bonded stiffeners showed the best DT properties in comparison with riveted and integral ones. A great reduction of the skin crack growth propagation rate can be achieved with the adoption of additional doublers bonded between the stringers.
Implementation of vision-based automatic guidance system on a fixed-wing unmanned aerial vehicle
Authors: Lee C-S , Hsiao F-B
Published: February 2016

This paper presents the design and implementation of a vision-based automatic guidance system on a fixed-wing unmanned aerial vehicle (UAV). The system utilises a low-cost ordinary video camera and simple but efficient image processing techniques widely used in computer-vision technology. The paper focuses on the identification and extraction of geographical tracks such as rivers, coastlines, and roads from real-time aerial images. The image processing algorithm primarily uses colour properties to isolate the geographical track of interest from its background. Hough transform is eventually used to curve-fit the profile of the track which yields a reference line on the image plane. A guidance algorithm is then derived based on this information. In order to test the vision-based automatic guidance system in the laboratory without actually flying the UAV, a hardware-in-the-loop simulation system is developed. Description regarding the system and significant simulation result are presented in the paper. Finally, an actual test flight where the UAV successfully follows a stretch of a river under automatic vision-based guidance is also presented and discussed.
Improved extrapolation of steady turbulent aerodynamics using a non-linear POD-based reduced order model
Authors: Zimmermann R. , Görtz S.
Published: February 2016

A reduced-order modelling (ROM) approach for predicting steady, turbulent aerodynamic flows based on computational fluid dynamics (CFD) and proper orthogonal decomposition (POD) is presented. Model-order reduction is achieved by parameter space sampling, solution space representation via POD and restriction of a CFD solver to the POD subspace. Solving the governing equations of fluid dynamics is replaced by solving a non-linear least-squares optimisation problem. The method will be referred to as LSQ-ROM method. Two approaches of extracting POD basis information from CFD snapshot data are discussed: POD of the full state vector (global POD) and POD of each of the partial states separately (variable-by-variable POD). The method at hand is demonstrated for a 2D aerofoil (NACA 64A010) as well as for a complete industrial aircraft configuration (NASA Common Research Model) in the transonic flow regime by computing ROMs of the compressible Reynolds-averaged Navier-Stokes equations, pursuing both the global and the variable-by-variable POD approach. The LSQ-ROM approach is tried for extrapolatory flow conditions. Results are juxtaposed with those obtained by POD-based extrapolation using Kriging and the radial basis functions spline method. As a reference, the full-order CFD solutions are considered. For the industrial aircraft configuration, the cost of computing the reduced-order solution is shown to be two orders of magnitude lower than that of computing the reference CFD solution.