Efforts to develop the next generation of aircraft with ever-increasing levels of performance – higher, farther, faster, cheaper – face great technical challenges. One of these technical challenges is to reduce structural weight of the aircraft. Another is to look to aircraft configurations that have been unrealizable to date. Both of these paths can lead to a rigid flex coupling phenomenon that can result in anything from poor flying qualities to the loss of an aircraft due to flutter. This has led to a need to develop an integrated flight and aeroelastic control capability where structural dynamics are included in the synthesis of flight control laws. Studies have indicated that the application of an integrated flight and aeroelastic control approach to a SensorCraft high-altitude long-endurance vehicle would provide substantial performance improvement(1,2). Better flying qualities and an expanded flight envelope through multi-flutter mode control are two areas of improvement afforded by integrated flight and aeroelastic control. By itself, multi-flutter mode control transforms the flutter barrier from a point of catastrophic structural failure to a benign region of flight. This paper discusses the history and issues associated with the development of such an integrated flight and aeroelastic control system for the X-56A aircraft.

The aeronautical industry has set for itself important environmental goals, creating the need of improved tools for measuring the polluting emissions generated by fuel burn. This paper describes a new method to estimate the fuel burn and the pollution generated by a landing approach and also for a missed approach procedure. Fuel emission estimations can be used to compare the costs of different routes. The method developed in this paper uses the Emissions Guide Book developed by the European Environment Agency as the needed database for the computations. This method gives estimations of the fuel burn and emissions, including the amounts of CO, NOx, HC, EICO, EINOx and EIHC for a given flight as its output. The flight computation is divided into two sections: one section for the aircraft's travelled distance and a second section for the time an aircraft flies under certain flight modes. Since this method computes the missed approach fuel and emissions contribution, it computes the burn for a given descent approach for a successful landing, as well as for the same descent approach with a missed approach procedure followed by a successful landing. These two landings are verified in a complete flight to study the missed approach contribution for a conventional mid-range flight. The results show that a descent with the missed approach procedure requires 5.7 times more fuel than a normal, successful descent. This extra fuel burn increases the pollution released to the atmosphere and would impact the airlines’ profit margin due to the added fuel costs and longer flight times, as well as any future economic measures imposed on the increased air and noise pollution.

This paper describes the design optimisation study used to aerodynamically optimise the fairings that cover the rear wheels of the Land Speed Record vehicle, BLOODHOUND SuperSonic Car (SSC). Initially, using a Design of Experiments approach, a series of Computational Fluid Dynamics simulations were performed on a set of parametric geometries, with the goal of identifying a fairing geometry that was aerodynamically optimised for the target speed of 1,000 mph. Several aerodynamic properties were considered when deciding what design objectives the fairings would be optimised to achieve; chief amongst these was the minimisation of aerodynamic drag. A parallel, finite-volume Navier–Stokes solver was used on unstructured meshes in order to simulate the complex aerodynamic behaviour of the flow around the vehicle’s rear wheel structure, which involved a rotating wheel, and shockwaves generated close to a supersonic rolling ground plane. It was found that the simple response surface fitting approach did not sufficiently capture the complexities of the optimisation objective function across the high-dimensional design space. As a result, a Nelder–Mead optimisation approach was implemented, coupled with Radial Basis Function design space interpolation to find the final optimised fairing design. This paper presents the results of the optimisation study as well as indicating the likely impact this optimisation will have on the ultimate top speed of this unique vehicle.

The cycloidal propeller for a Micro-Aerial Vehicle (MAV)-scale cyclogyro in hover was studied using a 2D Reynolds-averaged Navier-Stokes equations solver. The effects of the blade dynamic stall, parallel Blade Vortex Interaction (BVI), inflow variation and flow curvature were discussed, based on the results of numerical simulation. The results from the 2D Computational Fluid Dynamics simulation indicated that the blade of the cycloidal rotor is actually performing a pitching oscillation, if observed in a moving reference frame. The dynamic stall vortices shed from the upstream blade cause intense parallel BVI on the downstream blade. The interaction will induce upwash and downwash on the downstream blade. This changes the effective reduced frequency and actually delays the stall of the blade, which is beneficial to the thrust generation. There is also strong downwash in the rotor cage and it changes the inflow velocity experienced by the blade. The downwash and flow curvature can either be beneficial or harmful to the thrust generation. The combined effects of dynamic stall, parallel BVI, inflow variation and flow curvature cause large aerodynamic force peaks and ensure the cycloidal rotors work at very low rotation speeds with high thrust. This guarantees that the cycloidal rotors possess at least the same level of hover efficiency as screw propellers.

The flow and noise created by sawtooth trailing-edge serrations has been studied experimentally at a low Reynolds number. Experiments have been performed on a flat-plate model with an elliptical leading edge and an asymmetrically bevelled trailing edge at Reynolds numbers of Rec = 1 × 105–1.3 × 105, based on chord. Wide serrations with a wavelength (λs) to amplitude (2h) ratio of λs/h = 0.6 were found to reduce the overall sound pressure level by up to 11dB. In contrast, narrower serrations with λs/h = 0.2 produce tonal noise and increase the overall noise level by up to 4dB. Intense vortices across the span of the trailing edge with narrow serrations are shown to be the source of tonal noise. Wide serrations reduce turbulent velocity fluctuations at low frequencies which explains the lower radiated noise. The narrow serrations that produce low Reynolds number tonal noise were shown previously to be effective at higher Reynolds numbers (Rec > 2 × 105), demonstrating that care is needed to fully understand the flow field over serrations for all intended operating conditions.

Composite materials have been increasingly used in aircraft structures. However, these composite structures are susceptible to damage from external low-velocity impacts. In this paper, an impact identification algorithm is proposed to estimate the impact location and force time history simultaneously. A localisation method based on basis vectors is proposed, and the impact force time history is reconstructed by simplified transfer functions. The basis vector stands for the relationship between the impact location and the sensor signals, and the transfer function shows the relationship of the sensor signal and the force time history. An experiment is conducted on a flat glass fibre-epoxy matrix composite plate to verify the developed algorithm using only four sensors. The soft impactor and hard impactor are two typical impactors for impact events; therefore, the impact experiment is performed by the rubber and the steel impactors, respectively. The experimental results indicate that the proposed algorithm is feasible for the identification of impact events on plate-like composite structures.

Aircraft damage modelling was conducted on a Boeing 747 to examine the effects of asymmetric horizontal stabiliser loss on the flight dynamics of a commercial Fly-by-Wire (FBW) aircraft. Robustness of the control system is investigated by analysing how characteristic eigenvalues move as a result of damage and comparison to the non-FBW aircraft is made. Furthermore, the extent of stabiliser loss that the system can successfully handle without loss of stability and acceptable performance is identified. The presented analysis of the results gives insightful knowledge to aid in the design of an improved FBW system with increased damage tolerance. A handling qualities evaluation is presented to provide an understanding of how the pilot perceives the damaged aircraft. The results of the study show that a generic FBW system improves robustness such that the aircraft is stable with 50% horizontal stabiliser loss. With 50% damage, the aircraft is controllable but unsafe to fly and may be unable to effectively complete its mission task.