The future air traffic management (ATM) concept envisaged by the Single European Sky ATM Research – SESAR – and the USA equivalent NextGen, mark a paradigm shift from the current reactive approach of ATM towards holistic strategic collaborative decision making. The core of the future ATM concept relies on common situational awareness over potentially large time-horizons, based upon the user operational intent. This is beyond human capabilities and requires the support of automation tools to predict aircraft state throughout the operation and provide support to optimal decision making long before any potential conflict may arise. This is achieved with trajectory predictors and conflict detectors and resolvers respectively. Numerous tools have been developed, typically geared towards addressing specific airborne applications. However, a comprehensive literature search suggests that none of the tools was designed to predict trajectories throughout the entire operation of an aircraft, i.e. gate-to-gate. Yet, such functionality is relevant in the holistic optimisation of aircraft operations. To address this gap, this paper builds on an existing en route trajectory prediction (TP) model and develops novel techniques to predict aircraft trajectories for the transitions between the ground- and enroute-phases of operation and for the ground-phase, thereby enabling gate-to-gate (or enroute -to-enroute) TP. The model is developed on the basis of Newtonian physics and operational procedures. Real recorded data obtained from a flight data record (FDR) were used to estimate some of the input parameters required by the model. The remaining parameters were taken from the BADA 3.7 model. Performance results using these flight data demonstrate that the proposed TP model has the potential to accurately predict gate-to-gate trajectories and to support future ATM applications such as gate-to-gate synchronisation.

In the present work natural laminar flow (NLF) and hybrid laminar flow (HLF) wing designs are presented which were obtained by combining new methodologies with experience and knowledge obtained with traditional laminar wing design methods. The NLF wing design is performed for wing-body configurations with backward swept wing (BSW) and forward swept wing (FSW). Initial aerofoil sections were obtained by using a new sectional conical wing method which allows the design of transonic wing sections, taking into account the effects of sweep and taper for the computational cost of a 2D method. Except for flow regions with strong 3D influence, wings constructed with these aerofoils showed an acceptably large region with laminar boundary layer and small shocks at design and specified off-design conditions. For regions close to the body and the tip a 3D inverse design method was further required. For the BSW case, due to cross flow a premature transition occurred. Therefore, a HLF panel was required to obtain a larger laminar region. A suction distribution was obtained using the suction distribution module of the automated target pressure generator (ATPG). This generator optimises the pressure distributions in terms of minimising drag while keeping certain boundary conditions constant, e.g. lift and momentum. Using the ATPG, the laminar extent of the BSW NLF design could be further improved for the inboard wing. With the new methodologies design work was reduced. They lead to a design with reserves that allow for acceptable off-design performance qualities by keeping the wing laminar over a wide range of flight conditions.

This paper presents findings from research conducted at the University of Liverpool aimed at optimising the motion cues available from a short-stroke hexapod motion platform. Piloted simulation experiments were conducted to gather both objective and subjective measures of pilot performance and motion fidelity for a typical helicopter low-speed yaw axis capture task. These measures included subjective ratings taken for the first time using a new motion fidelity rating scale. During the experiment four baseline configurations were presented to the pilot; – No motion, Rotational-only motion, Translational-only motion and Rotational + Translational motion, and within these baseline configurations various values of motion filter gains were examined. The key findings from the experiment show that the most acceptable motion fidelity was obtained when congruent cues were presented in both the rotational and translational axes. A possible explanation for this result is postulated to lie within the relative motion perception thresholds of the human semicircular canals. These results contrast with previous research which suggested that yaw axis motion cues were of less importance than translational motion cues.

The premise is that in the future civil Air-to-Air Refuelling (AAR) will become an economic necessity if popular mass air travel is to continue. What is attempted is to provide a contemporary view of how such future operations could be safely undertaken. The intention is to predict the certification basis for demonstrating safe AAR operation of Cruiser-Feeder concepts. The necessary systems and aircraft functions are treated very much as they are today when civil certifying a large aeroplane type. The compliance demonstration required for environmental conditions, flight envelope, systems providing the necessary functionality, structural integrity, weight and balance are discussed.

Applicable existing civil certification requirements are identified and where necessary expanded in scope to accommodate AAR operation. Where contemporary material does not supply appropriate guidance then corresponding safety criteria are proposed to address the deficiency. Lessons learnt from military AAR include the drive for interoperability. This has resulted in extensive efforts to standardise equipment and systems, which are equally applicable to civil AAR. Extremely useful advisory material exists, ranging from flight testing techniques to related safety.

The importance of ensuring the consistency of failure condition categorisation at system and aircraft level is highlighted. The treatment of failures when two aircraft are in close proximity is something not considered by civil functional hazard analysis. The concept of AAR as an additional flight phase is introduced and affected system safety analyses identified. Examples of failure conditions that are not catastrophic at system level, but potentially could be at aircraft level during AAR are provided. Rendezvous scenarios are described to illustrate their influence on the certification basis. Combining such considerations with the factors that influence aircraft design leads to ramifications for handling qualities, performance and fuel system design. A viable and certifiable AAR configuration is consequently proposed. Consideration is given to treating operational certification in a progressive manner similar to existing LROPS (Long Range Operations).

Peridynamics is a recently proposed continuum theory based on a non local approach and formulated with integral equations. The theory is suitable for dealing with crack propagation in solid materials. The original peridynamic formulation regarded dynamic problems and was adapted to the static case mainly using a relaxation method by introducing a substantial amount of numerical damping in the time integration. In the present work the implementation of the theory within an implicit code for static crack propagation phenomena based on the Newton-Raphson method is presented and applied to several examples of static crack propagation equilibrium problems. Results obtained with the newly developed procedure are presented for various structural configurations, with different boundary and load conditions and quantitatively compared to published data.

This paper contributes to the prediction of aircraft noise. The framework of our study is based on a comprehensive analysis, which includes sub-models for airframe, propulsion and propeller noise. The sources models are complemented by various accounts of interference phenomena, including fuselage scattering, propeller shielding, wing and flap effects. The propagation of the acoustic pressures from the noise sources is done through a number of additional sub-models, which include spherical spreading, Doppler effects, ground reflection, shear winds and temperature gradients through the atmosphere. We present some fly-over noise data. The data include measurements by 13 microphones at community locations, meteorological data, and flight recorder data packs, which provide the complete state of the aircraft along the noise trajectory. Examples of verification are shown for the turboprop airplane Dash8-Q400, both on approach and departure. A sensitivity analysis is carried out to investigate the importance of the uncertainty effects, which are split between internal (aircraft and propulsion), external (atmospheric environment), and aircraft position.

The analysis of risks per phases of flight is fundamental for safe helicopter operations, of which night-time offshore oil- and gas-related missions form an important part. The safe execution of such missions also depends on pilots’ recent flying practice and a stable visual approach segment prior to landing. However, the poor quality of the safety data currently available prevents accurate analysis of risk on a per-phase-of-flight basis, establishment of a meaningful flying recency requirement and identification of any preferable visual approach design. To redress these problems, this paper develops a bespoke taxonomy of phases of offshore helicopter flights and uses it as the basis for a questionnaire survey on the phase-specific risk levels experienced by pilots in the night-time, perceived optimal flying recency requirement and preferred visual approach design. With the responses obtained from pilots located in seven countries, extensive statistical hypothesis testing shows that the phases involving visual scan techniques at high speed regimes are problematic, especially the visual segment of instrument approaches. Moreover, the between-night-flights time gaps required for assured flying recency were found considerably shorter than currently standardised across the industry. Finally, no preferred visual approach technique was identified. A number of important implications have been highlighted and should form the basis for future safety interventions.

The paper focuses on the modelling, simulation and control of an electrical miniature actuator integrated in the actuation mechanism of a new morphing wing application. The morphed wing is a portion of an existing regional aircraft wing, its interior consisting of spars, stringers, and ribs, and having a structural rigidity similar to the rigidity of a real aircraft. The upper surface of the wing is a flexible skin, made of composite materials, and optimised in order to fulfill the morphing wing project requirements. In addition, a controllable rigid aileron is attached on the wing. The established architecture of the actuation mechanism uses four similar miniature actuators fixed inside the wing and actuating directly the flexible upper surface of the wing. The actuator was designed in-house, as there is no actuator on the market that could fit directly inside our morphing wing model. It consists of a brushless direct current (BLDC) motor with a gearbox and a screw for pushing and pulling the flexible upper surface of the wing. The electrical motor and the screw are coupled through a gearing system. Before proceeding with the modelling, the actuator is tested experimentally (stand alone configuration) to ensure that the entire range of the requirements (rated or nominal torque, nominal current, nominal speed, static force, size) would be fulfilled. In order to validate the theoretical, simulation and standalone configuration experimental studies, a bench testing and a wind-tunnel testing of four similar actuators integrated on the real morphing wing model are performed.

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)-0417MOD 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.

Water entry tests are conducted with a scale model capsule to explore the sensitivity of splashdown accelerations and motions to impact conditions, to support estimation of splashdown loads and behaviours of the proposed Titan Mare Explorer (TiME) mission. A 3D printed capsule with off-the-shelf USB accelerometer loggers enabled tests to be performed on a tight schedule. Contact impulse and peak loads are presented for a range of impact speeds and geometries, and predicted loads at full-scale on Titan are derived. The observed variation of peak load with impact speed is broadly consistent with the theoretically-expected square law but is surprisingly also consistent with a linear function (in common with some results from the literature). Test execution procedures and the performance of the data acquisition system are reviewed.

In this review paper, several take-off techniques of different species of animal flyers and gliders, both extinct and extant, are analysed. The methods they use vary according to animal group and size. Smaller animals, such as insects, rely on the use of transient aerodynamic techniques or the use of stored elastic energy. Medium-size flyers such as birds, bats, and other mammal gliders initiate flight by a jump which involves leg and wing movement coordination. The largest animals to fly, the extinct pterosaurs, are believed to have used a combination of aerodynamic and mechanic techniques in order to become airborne. The information presented here can be used as a resource for novel biomimetic unmanned aircraft design.

A combined experimental and computational study on a rectangular lobed mixer is performed. A series of simulations based on a steady Reynolds-averaged Navier-Stokes Simulation (RANS) are conducted to analyse the mixing mechanisms of large-scale streamwise structure shed by the trailing edge of lobed mixer, with emphasis being placed on the effect of turbulence modeling and inflow conditions. The simulations are validated in respect of velocity and scalar distribution against the data obtained through Particle Image Velocimetry (PIV) and Nanoparticle-based Planar Laser Scattering (NPLS) technique. The computational results predicted by the SST k –ω turbulence model show better agreement with the experimental data. But the small-scale turbulence structures are not captured accurately by these turbulence models. The convoluted shear layer shed from trailing edge is stretched and rotated by the large-scale streamwise vortices, forming an unstable ‘pinching-off’ structure, which increases the interfacial area. And at the interface of two streams, a large number of small-scale turbulence structures are formed, which contribute a lot to the mixing enhancement along with the increased interfacial area. The streamwise vorticity decays more rapidly with the decrease of velocity ratio and total pressure ratio of two streams. The scalar thickness which reflects the mixing rate of two streams increases with the decreasing velocity ratio and total pressure ratio.

This technical article discusses design and integration associated with distributed propulsion as a means of providing motive power with significantly reduced emissions and external noise for future aircraft concepts. The technical work reflects activities performed within a European Commission funded Framework 7 project entitled Distributed Propulsion and Ultra-high By-Pass Rotor Study at Aircraft Level, or, DisPURSAL. In this instance, the approach of distributed propulsion includes a Distributed Multiple-Fans Concept driven by a limited number of engine cores as well as one unique solution that integrates the fuselage with a single propulsor (dubbed Propulsive-Fuselage Concept) – both targeting entry-in-service year 2035+. Compared to a state-of-the-art, year 2000 reference aircraft, designs with tighter coupling between airframe aerodynamics and motive power system performance for medium-to-long-range operations indicated potentially a 40-45% reduction in CO2-emissions. An evolutionary, year 2035, conventional morphology gas-turbine aircraft was predicted to be –33% in CO2-emissions.

Smart materials that are directly embedded in the rotor blade structure are an attractive concept for active blade control. A promising approach is the use of anisotropic piezoelectric strain actuators embedded in the rotor blade skin. Especially in Europe and the US this concept has been intensively investigated over the past years. A major drawback of all configurations studied so far is the high operation voltage of up to 2,000V of state of the art piezoelectric actuators. Within the Green Rotorcraft Project of the European Joint Technology Initiative Clean Sky, a new approach with a low voltage piezoelectric actuation system is investigated to demonstrate the feasibility of this technology.

A first major step in this direction was completed by conducting a centrifugal test with a model rotor blade. The objective of the centrifugal test was to demonstrate the performance of the actuation system and the structural concept under centrifugal loads by showing that the expected twist deformation can be achieved at the nominal rotation speed and different actuation frequencies.

It was demonstrated that the new actuation system is capable of operating under representative centrifugal loads. In comparison to state-of-the-art actuators (operation voltage 500V to +1500V) the new actuation system (operation voltage -20V to 120V) exhibits higher active twist performance per active area.

Box-like slung loads exhibit periodic yaw response instabilities, while carried externally by a helicopter. When coupled with the slung load longitudinal and lateral pendulum motions, these instabilities result in significant pendulum oscillations of the load. High amplitude oscillations lead in many cases to the limiting of a load’s flight envelope. Using wind tunnel and flight tests, rear mounted fins were previously demonstrated as efficient means for stabilisation of a problematic load. However, the lack of a proper analytical model of the stabilised load’s aerodynamic characteristics, led to a trial and error development process, without an appropriate physical understanding of the stabilisation problem. The present paper describes a method for the aerodynamic modeling of fins stabilised slung loads based on a limited number of simple static wind-tunnel tests. The resulting database is incorporated in a dynamical slung load simulation that shows good agreement with dynamic wind-tunnel tests. The applicability of the proposed method is demonstrated, by the calculation of stabilised loads aerodynamic databases for interim fin inclination angles not covered by tests.

In this review paper, different landing strategies of diverse species of animal flyers and gliders, both extinct and extant, are analysed. These methods vary depending on the animal group and the sensory system used by the animal to detect its landing site. In almost all species the use of delayed stall during the landing manoeuvre was observed. Sometimes wing flapping was used to aid in deceleration. With respect to guidance and navigation, most insect, bird and mammal gliders use their vision to guide them to landing via optical flow or motion parallax. Bats, which are nocturnal creatures, rely on their auditory system as they use echolocation to find their nesting site. Some butterfly and moth species guide themselves to landing using their olfactory sense as they follow pheromone trails. The information presented here can be used as a source of information for novel bio-inspired unmanned aircraft design.

The hovering performance and the lifting capability of tiltrotor aircraft are strongly affected by the aerodynamic interaction between wing and rotors. The tiltwing concept represents a promising technology to increase the hover performance by reducing the wing-rotor interference. The present work describes an experimental activity carried out on a ¼ scaled tiltwing aircraft half-model to achieve a detailed insight about the main issues characterising the aerodynamic interaction between wing and rotor in hover. The results of the experimental campaign, including both force measurements and Particle Image Velocimetry surveys, enabled to evaluate both the aircraft performance for different configurations of the tilting wing and to achieve a detailed insight about the flow physics of the rotor wake in the interaction with the wing. The test activity provided a comprehensive experimental database that was obtained over a not confidential aircraft configuration.

We present an autonomous visual landmark recognition and pose estimation algorithm designed for use in navigation of spacecraft around small asteroids. Landmarks are selected as generic points on the asteroid surface that produce strong Harris corners in an image under a wide range in viewing and illumination conditions; no particular type of morphological feature is required. The set of landmarks is triangulated to obtain a tightly fitting mesh representing an optimal low resolution model of the natural asteroid shape, which is used onboard to determine the visibility of each landmark and enables the algorithm to work with highly concave bodies. The shape model is also used to estimate the centre of brightness of the asteroid and eliminate large translation errors prior to the main landmark recognition stage. The algorithm works by refining an initial estimate of the spacecraft position and orientation. Tests with real and synthetic images show good performance under realistic noise conditions. Using simulated images, the median landmark recognition error is 2m, and the error on the spacecraft position in the asteroid body frame is reduced from 45m to 21m at a range of 2km from the surface. With real images the translation error at 8km to the surface increases from 107m to 119m, due mainly to the larger range and lack of sensitivity to translations along the camera boresight. The median number of landmarks detected in the simulated and real images is 59 and 44 respectively. This algorithm was partly developed and tested during industrial studies for the European Space Agency’s Marco Polo-R asteroid sample return mission.

This paper presents findings from research conducted at the University of Liverpool aimed at optimising the motion cues available from a short-stroke hexapod motion platform. Piloted simulations were conducted for a typical helicopter low-speed sidestep manoeuvre. To correctly simulate the sidestep manoeuvre the motion platform must translate laterally at the same time as it rolls. If the motion in these two axes is not properly harmonised then the pilot can experience significant false motion cues. This is a particular concern for short-stroke hexapod platforms, where displacement limits can severely constrain the available lateral travel particularly during motion in multiple axes (e.g. roll and sway). During the experiment the motion filter gains in the roll and sway axes and the roll-axis motion filter break-frequency were varied. Objective and subjective measures of pilot performance and motion fidelity were gathered for each motion filter configuration, the latter using a new motion fidelity rating scale. The key findings show that acceptable motion cues could only be achieved by careful harmonisation of the motion filter gains in the roll and sway axes. A high gain in the roll axis coupled with a low gain in the sway axis resulted in motion which was abrupt and uncomfortable. On the other hand, too large a gain in the sway axis resulted in extreme lateral displacements of the motion platform leading to undesirable side-effects. The phase distortion between the visual and platform motion cues, introduced by the roll-axis motion filter, also had a significant impact on the pilot’s perception of motion fidelity. These results are presented in the form of proposed motion fidelity criteria for short-stroke hexapod platforms and compared with results from previous research conducted on a range of large motion systems.