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.

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.

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 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 manoeuvring 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.

This paper provides an overview of techniques developed for the application of support vector regression in the domain of simulation and system identification of helicopter dynamics. A generic high fidelity FLIGHTLAB helicopter model is used to train and validate a number of pitch response SVR models. These models are then trained using flight data from a Sikorsky Seahawk helicopter. The SVR simulation results show significant promise in the ability to represent aspects of a helicopter’s dynamics at a high fidelity. To achieve this, it is important to provide the SVR kernel with knowledge of past inputs that encompass the delay characteristics of the helicopter dynamic system. In this case, the use of nonlinear auto regressive eXogenous input network architecture achieves this goal. Good performance was achieved using input data that encompassed between 300 to 500ms worth of historic response.

The use of active trailing edge flaps on rotors may lead to performance benefits as well as noise and vibration reduction. In this work, computational fluid dynamics, using the HMB2 solver, is used to assess the effect of the trailing edge flaps on the whole flight domain of a modern main rotor. Starting from a baseline blade design, multiple techniques are demonstrated. The flap is first assessed using 2D pitching aerofoil simulations, followed by dMdt simulations, that account for the simultaneous variations of pitch and Mach around the azimuth. It was shown that enhanced lift was obtained while inspection of the moment coefficient showed negative damping for the flap for a limited set of conditions. Due to the 2D formulation, dMdt computations are fast to perform and can be used to inform codes predicting the rotor performance. The flap was then assessed in hover, and only allowed for limited improvement in blade performance at high thrust. In forward flight, the flap was actuated at a frequency of 1 per revolution, and was found to have a strong effect on the loads on the retreating side. The effect on the moments was even stronger. The flight envelope of the blade was explored, and clean and flapped cases were compared. The most noticeable changes occur at high and medium thrust. The CFD method was found to be efficient and robust, without any substantial penalties in CPU time, due to the flap modelling, over the tested conditions.

The design and building of accurately scaled aeroelastic wind-tunnel models is difficult, time consuming and very costly. With the increasing usefulness of computational methods for predicting aeroelastic phenomena, few complex models have been built in recent years. New fighter aircraft projects are also becoming more and more scarce, and transport aircraft have essentially the same configuration since half a decade. This also significantly reduces the need for aeroelastic wind-tunnel models. However, there still is an interest in the results from aeroelastic testing. In some cases new and radical configurations may warrant wind-tunnel testing and in other cases complex phenomena arising in fight testing may need carefully performed experiments to resolve problems. However, there is definitely a trend towards building models and performing testing in the support of the development of computational methods.

The developments in computer technology do not only improve the computational methods for aeroelasticity. Modern Computer Aided Design and Manufacturing techniques can significantly improve the quality and efficiency of the design and build process for aeroelastic models. There have also been some recent improvements in measurement techniques which have proven very useful for testing of aeroelastic wind-tunnel models. The paper will present some new design and build techniques developed for the manufacturing of a large scale wind-tunnel model of a canard delta wing fighter aircraft configuration. In the build process fiber-reinforced composites will be used, hence, challenges and possible solutions concerning the ability to produce a model with well defined material properties and fiber angles will be discussed. Further challenges arise when both measurement equipment and adjustable control surfaces should be attached inside the model using techniques that are possible to describe with computational methods. In addition, equipment, such as pressure taps, and control surface mechanics need to fit and function in a flexible structure. As a result, the above requirements will lead to necessary compromises in the design, hence, the paper will present the choices taken during the build process and for which reasons. The use of an optical positioning measurement system will also be discussed for both the validation of model properties and non-contact measurement of model deformations during wind-tunnel testing.