This paper describes a new efficient method for the construction of an approximately balanced aerodynamic Reduced Order Model (ROM) via the frequency domain using Computational Fluid Dynamics data. Time domain ROM construction requires CFD data, which is obtained from the DLR TAU RANS or Euler Linearised Frequency Domain (LFD) solver. The ROMs produced with this approach, using a small number of frequency simulations, are shown to exhibit a strong ability to reconstruct the system response for inviscid flow about the NLR7301 aerofoil and the FFAST wing; and viscous flow about the NASA Common Research Model. The latter demonstrates that the reduced order model approach can reconstruct the full order frequency response of a viscous aircraft configuration with excellent accuracy using a strip wise approach. The time domain models are built using the frequency domain, but also give promising results when applied to reconstruct non-periodic motions. Results are compared to time domain simulations, showing good agreement even with small ROM sizes, but with a substantial reduction in calculation time. The main advantage of the current model order reduction approach is that the method does not require the formation and storage of large matrices, such as in POD approaches.

The stabilisation and control mechanisms of an Unmanned Aerial System (UAS) must be properly designed to ensure acceptable flight performance. During their operation, these mechanisms are subjected to unknown and random environmental effects, making it imperative that all available information should be taken into consideration during the mechanisms’ design process (e.g. system dynamics, actuators, flight conditions and certain criteria requirements such as phugoid and short modes for longitudinal dynamics, and roll subsidence, spiral and Dutch-roll modes for lateral dynamics) in order to guarantee flight stability. Therefore, this paper introduces a novel methodology for the stabilisation and control of the UAS-S45 Bálaam, designed and manufactured by Hydra Technologies. This methodology uses composite controllers that combine feedback Linear Quadratic Regulators (LQR) and Proportional Integral Feed-Forward (PI-FF) compensation controller for stabilisation and tracking tasks, respectively. Furthermore, a Generalised Extended State Observer was implemented to provide robustness to the closed loop dynamics by introducing disturbance compensation. Furthermore, an Adaptive Neuro-Fuzzy Inference System (ANFIS) was adopted to perform a gain scheduling by computing the gains of each composite controller for certain unknown trim conditions within a given flight domain. Finally, several numerical assessments were performed to highlight the efficiency of the proposed methodology.

Single pilot passenger aircraft concepts are being developed by several manufacturers. Various technological approaches are being explored: One concept is to use a Harbour Pilot dedicated to providing support for departures and arrivals. The Harbour Pilot has comprehensive knowledge of the terminal area airspace, procedures and operations. If a single pilot aircraft is to be viable, however, the number of supporting personnel needs to be significantly smaller than the number of First Officers normally employed for a two-crew aeroplane, but the number such staff has yet to be determined. This study models operations by a UK low-cost operator at a regional airport to determine the optimum number of Harbour Pilots required to support operations throughout the day. The model uses a simplified timeline analysis with task data incorporated into a dynamic discrete event modelling system allowing for multiple replications using various configurations. Results suggest that for this operation six Harbour Pilots per shift used flexibly to support both departures and arrivals would be required.

Intense acoustic loads from jet noise cause noise pollution and induce failures, such as the malfunctioning of electronic devices and fatigue failure of internal/external structures. Consequently, the prediction of jet noise characteristics is crucial in the development of high-speed vehicles. This study presents acoustic experiments and predictions for an under-expanded, unheated jet using a small-scale prototype. Outdoor measurements are carried out using a vertical ejection setup. Acoustic characteristics are measured using both linear and circular microphone arrays. Additionally, numerical prediction of the same jet noise is performed using a detached eddy simulation and the permeable Ffowcs-Williams and Hawkings acoustic analogy. The vertical experimental setup exhibits the typical acoustic characteristics of a supersonic jet in terms of directivity and broadband shock-associated noise. Moreover, the numerical prediction exhibits satisfactory accuracy for the jet downstream, where the large-scale turbulence structures of the directivity predominate. However, discrepancy increases in the domain of lower directivity. The presented experiment and prediction will be extended to future studies regarding the noise of various deflector duct configurations impinging on supersonic jets.

This study presents a methodology to estimate the battery consumption of an electric powerplant, based on brushless motors, typically used in light unmanned aerial systems. The methodology models brushless motors through an equivalent circuit obtained from their dynamic behaviour. Propellers’ data are taken from an experimental database. Furthermore, a variable speed controller efficiency is considered in the methodology. All the parameters involved in the model are adjusted by minimising the mean quadratic error of measurements taken in both direct and alternating currents. This model allows designers to predict energy consumption, also if any element of the powerplant changes, such as battery or propeller. Thus, it is useful for selecting the best powerplant for an actual RPAS operation. The results obtained to predict the current consumption of several electric powerplants show a coefficient of determination higher than 0.96. Finally, the methodology is validated by means of a case study of an actual RPAS, where the best powerplant is selected in terms of endurance.

This article reports on the development of a fast method for generating the aerodynamic database for subsonic flow over a missile. At present, this is typically achieved using RANS-based CFD, which is expensive for the complex missile geometries and the multiplicity of operating conditions to be evaluated. The presented reduced-order model (ROM) provides a reasonably accurate prediction of the aerodynamic coefficients of the missile (and, in fact, the full flow field around it) within half a minute. In particular, in the interpolative regime, prediction errors for all coefficients are typically less than 1% of their respective maximum values encountered in the database, with extrapolation incurring more error. The empirical approach ‘learns’ from the CFD solutions calculated for a few operating conditions, and then is able to make predictions for any other condition within a feasible set. The learning employs proper orthogonal decomposition (POD) to characterise the most important features of the flow. The prediction is posed as an optimisation problem that aims to find the flow solution as a linear combination of the above POD modes that minimises the residual of the governing equations. Innovations on the prevailing POD-ROM approach include novel implementation of boundary conditions, simplified computation of the aerodynamic coefficients, and a procedure for choosing modelling parameters based on extensive cross-validation. Challenges overcome in application to a problem of industrial relevance are discussed.

Elevator failure may have fatal consequences for fighter aircraft that are unstable due to their high manoeuvrability requirements. Many studies have been conducted in the literature using active and passive fault-tolerant control structures. However, these studies mostly include sophisticated controllers with high computational load that cannot work in real systems. Considering the multi-functionality and broad operational prospects of fighter aircraft, computational load is very important in terms of applicability. In this study, an integrated fault-tolerant control strategy with low computational load is proposed without sacrificing the ability to cope with failures. This control strategy switches between predetermined controllers in the case of failure. One of these controllers is designed to operate in a non-failure condition. This controller is a basic controller that requires very little computational effort. The other controller operates when an asymmetric elevator failure occurs. This controller is a robust fault-tolerant controller that can fly the aircraft safely in case of elevator failure. The switching is decided by a failure detection system. The proposed integrated fault-tolerant control system is verified by non-linear F-16 flight simulations. These simulations show that the proposed method can cope with failures but requires less computational load because it uses a conventional controller in the case of no failure.

For the safety problems caused by the limited landing space of the deck during the arresting process of the carrier-based aircraft, a dynamic model of the carrier-based aircraft’s landing and arresting is built. Based on the batch simulation method, the lateral dynamics safety envelope of the aircraft during the arresting was defined, and the dynamic response of the key points in the envelope during the arresting process was investigated. Subsequently, the influence of engine thrust and aircraft quality on the arresting safety envelope was studied based on reasonable safety evaluation indicators, and the safety status envelope of the deck arresting was given. Then, the particular Hamilton-Jacobi partial differential equation is used to obtain the lateral dynamics safety envelope of the carrier-based aircraft in the process of landing and arresting by backward inversion. Results indicate that engine thrust and landing quality have little effect on the yaw angle in the arresting safety boundary during the arresting. Additionally, with the engine thrust and landing quality increase, the maximum safe off-centre distance gradually decreases, and the safety boundary decreases accordingly. During the phase of landing glide, the engine thrust and quality have little effect on the maximum safe eccentric distance. When the engine thrust is increased by 40%, the maximum safe yaw angle is reduced from 0.3°, and the safety boundary is reduced by 4.2%. When the aircraftquality increases by 40%, the maximum safe yaw angle is reduced by 0.4°, and the safety boundary is reduced by 2.8%. The findings of this paper can provide framework for the research on theaircraft-to-carrier dynamic matching characteristics of the carrier-based system, and is of great significance to the research on improving the safety of the carrier-based aircraft landing arresting.