Peter Thomas

Renewed interest in supersonic air travel has prompted researchers to reconsider the design and operation of supersonic transport aircraft. Previously, such aircraft were restricted to overwater routes due to the disturbances caused by their sonic booms. Now, however, low-boom designs and overland flight at marginally supersonic Mach numbers are seen as potential enablers for widespread supersonic air travel. As a result, the trajectories that next-generation supersonic transports may fly are likely to be less constrained than for previous types, and in the last decade there has been a noticeable increase in research focusing on trajectory planning for such aircraft. This paper reviews the different methods that have been used to generate and optimise the flight paths of past and future supersonic transports. The challenges associated with optimising trajectories for aircraft that do not yet exist are discussed, and suggestions for future research activity are presented. Climate-optimal trajectory planning and development of detailed, non-proprietary supersonic aircraft performance models are identified as two key areas for future work.

In the realm of supersonic design, obtaining data for numerous supersonic configurations amidst intricate flow conditions proves time-consuming due to the excessive costs associated with high-fidelity computational demands. Running iterative simulations over an extended period is often impractical or entails substantial expenses. This inherent challenge necessitates the adoption of low-order potential solvers with reasonable accuracy to generate datasets. In support of this objective, This study addresses the high computational costs of obtaining data for supersonic configurations by developing a low-order solver that combines the Taylor-Maccoll hypervelocity method (TMHM) with the supersonic vortex lattice method. This approach aims to provide accurate drag predictions in supersonic flows while minimizing computational demands. By integrating TMHM to calculate wave drag and skin friction drag and enhancing the vortex lattice method to handle shockwave impacts through panel matching, the solver achieves improved accuracy in lift and drag computations. Validation against experimental data shows a 20% reduction in drag prediction error compared to traditional vortex lattice methods, with a 2.01% error for low-shock angles. The method achieves accuracy rates between 90% and 95% across various configurations, including a 90% accuracy for delta wings, 85% for positive dihedral wings, and 95% for large sweptback angle designs, as confirmed by comparisons with high-fidelity CFD data.

This paper presents a range control approach for implementing hard geofencing for unmanned air vehicles (UAVs), and especially remotely piloted versions (RPVs), via a proposed anticipatory range calculator. The approach employs turning circle intersection tests that anticipate the fence perimeter on approach. This ensures the vehicle turns before penetrating the geofence and remains inside the allowable operational airspace by accounting for the vehicles’ turning dynamics. Allowance is made for general geozone shapes and locations, including those located at the problematic poles and meridians where nonlinear angle mapping is dealt with, concave geozones, narrow corners with acute internal angles, and transient turn dynamics. The algorithm is shown to prevent any excursions using a high-fidelity simulation of a small remotely piloted vehicle. The algorithm relies on a single tuning parameter which can be determined from the closed-loop rise time in the aircraft’s roll command tracking.

In recent years, the air-drop launch technology of near-space UAVs has attracted much attention. Between downfall from the carrier and the flight control system’s initiation, the UAV presents free-fall movement. This free-fall process is very important for the control effect of the flight control system and is also crucial for the safety of the UAV and the carrier. Focus is required on two important dynamic parameters of the UAV: the moment of inertia and the center of mass position. In this paper, we used a quasi-steady model proposed by predecessors to address the flat-plate falling problem with modifications to describe the freely falling motion of the wing. Computational fluid dynamics (CFD) were used to simulate the free-fall movement of the wing with various parameters, and the wing release behavior was analyzed to check the quasi-steady model. Research shows that the movement characteristics of the falling wing are mostly reflected in the longitudinal plane, and the developed quasi-steady analytical model can more accurately describe the dynamic behavior of free-fall to some extent. By using CFD methods, we further investigated the aerodynamic performance of the free-fall wing. The results show that the wing mainly presents tumbling and fluttering motion. Changing the moment of inertia around the tumbling axis changes the tumbling frequency and the time point as the wing enters tumbling. In contrast, changing the position of the center of mass significantly changes the form of falling and makes the free-fall motion more complex. Therefore, it is necessary to carefully configure the center of mass in the UAV design process.

High-altitude pseudo-satellites are an expanding focal area of the aerospace industry which require new technologies and manufacturing processes to reduce weight and increase efficiency with the aim of increasing endurance. One such process has resulted in the occurrence of small deformities along the leading edge of a lightweight unmanned aerial vehicle structure with the application of its skin, which may have a detrimental impact on its performance and efficiency. This paper focuses on the effects of these manufacturing deformities on the aerodynamic performance of the vehicle’s aerofoil when operating in low Reynolds number flow with the intention of identifying any detrimental flow variation. This analysis is achieved by comparing the lift curve, drag polar and pressure coefficient of both the deformed and undeformed cases of two aerofoils: a SG6042 and a GOE 523. This is accompanied with an examination of the local flow conditions scrutinising the near-wall y⁺ and turbulent kinetic energy calculations. The investigation finds that in two-dimensional flow, the deformities replicate the effects of transition trips in the shrinking or elimination of laminar separation bubbles. At Reynolds numbers below 250,000, the deformities reduce the net drag while leaving the lift largely unaffected. However, as a result, there is a slight shift in the minimum power condition in the order of 8% which would produce some performance loss for power efficiency and endurance.

Eye gaze controlled interfaces allow us to directly manipulate a graphical user interface just by looking at it. It has already been explored for assistive technology. Eye gaze controlled interface has great potential use in military aviation. In particular, operating different displays in situations where pilots' hands are occupied with flying the aircraft. This paper reports two studies on analyzing accuracy of eye gaze controlled interface inside aircraft undertaking representative flying missions. We reported that pilots can undertake representative pointing and selection tasks at less than 2 sees on average and the accuracy of COTS eye gaze tracking glass is less than 5° of visual angle up to +3G, although it is less accurate at ‒1G and +5G. We used these results in developing a multimodal head and eye gaze movement based HMDS and in a user study involving standard ISO 9241 pointing task. The average pointing and selection times was significantly lower in the proposed system compared to traditional TDS.

This article investigates the modeling of macro fiber composite-actuated laminate plates with distributed actuator patches. The investigation details an analytical and finite element modeling, with experimental validation of the bending strain and deflection of an epoxy E-glass fiber composite laminate. An analytical approach is also developed to estimate the plate deflection from the experimental strain measurements. The analytical method uses direct integration of single dimensional plate bending moments obtained by strain-induced shear moments from the macro fiber composite actuators. Finite element analysis software was used with the composite laminate modeled in ANSYS ACP. The results from both analytical and numerical models show good agreement with the experimental results, with strain values agreeing within 20 ppm and the maximum difference in deflection not exceeding 0.1 mm between models. Finally, an application of the analytical model for developing morphing aerofoil designs is demonstrated.