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.

Compression systems of modern, civil aircraft engines consist of three components: Fan, low-pressure compressor (LPC) and high-pressure compressor (HPC). The efficiency of each component has improved over the last decades by means of rising computational power which made high level aerodynamic optimisations possible. Each component has been addressed individually and separated from the effects of upstream and downstream components. But as much time and effort has been spend to improve performance of rotating components, the stationary inter-compressor duct (ICD) has only received minor attention. With the rotating compression components being highly optimised and sophisticated their performance potential is limited. That is why more aggressive, respectively shorter, ICDs get more and more into the focus of research and engine manufacturers. The length reduction offers high weight saving and thus fuel saving potential as a shorter ICD means a reduction in aircraft engine length. This paper aims at evaluating the impact of more aggressive duct geometries on LPC and HPC performance. A multi objective 3D computational fluid dynamics (CFD) aerodynamic optimisation is performed on a preliminary design of a novel two spool compressor rig incorporating four different operating line and two near-stall (NST) conditions which ensure operability throughout the whole compressor operating range. While the ICD is free to change in length, shape and cross-section area, the blades of LPC and HPC are adjusted for changing duct aerodynamics via profile re-staggering to keep number of free parameters low. With this parametrisation length, reductions for the ICD of up to 40% are feasible while keeping the reduction in isentropic efficiency at aerodynamic design point for the compressor below 1%pt. Three geometries of the Pareto front are analysed in detail focusing on ICD secondary flow behaviour and changes of aerodynamics in LPC and HPC. In order to asses changes in stall margin, speedlines for the three geometries are analysed.

A spacetime formulation is presented to solve unsteady aerodynamic problems involving large deformation or topological change such as store separation, slat and flap deployment or spoiler deflection. This technique avoids complex CFD meshing methods, such as Chimera, by the use of a finite-volume approach both in space and time, and permits a locally varying real timestep. The use of a central-difference scheme in the time direction can yield non-physical transient solutions as a consequence of information travelling backwards in time. Therefore, an upwind formulation is provided and validated against one-dimensional and two-dimensional test cases. A hybrid formulation (central in space, upwind in time) is also given and unsteady cases are computed for a spoiler and spoiler/flap deployment, with all three formulations compared, demonstrating that the use of an upwind time stencil yields more representative physical solutions and improves the rate of convergence.

The reduction of computational costs in the context of the Multidisciplinary Design Optimisation of a typical medium-range aircraft was investigated through an assessment of active constraints and the use of multi-fidelity models-based estimation of drag and structural stress. The results show that for this problem, from the set of considered constraints that includes flutter boundary, the active constraint is a 2.5g pull up Maximum Take Off Weight. Results show that the multi-fidelity approach reduced the required high-fidelity aerodynamic number of evaluations, for both drag assessment and stress assessment with sufficient level of accuracy for the former and conservatively for the latter. Further computational cost reduction can be achieved using a surrogate model based Multidisciplinary Design Optimisation. The best configuration attained shows an Aspect Ratio increase of 16%, a reduction of 4.5% in fuel consumption and wing structural weight increase of 2.7% relative to a predefined baseline configuration.

A control system verification framework is presented for unmanned aerial vehicles using theorem proving. The framework’s aim is to set out a procedure for proving that the mathematically designed control system of the aircraft satisfies robustness requirements to ensure safe performance under varying environmental conditions. Extensive mathematical derivations, which have formerly been carried out manually, are checked for their correctness on a computer. To illustrate the procedures, a higher-order logic interactive theorem-prover and an automated theorem-prover are utilised to formally verify a nonlinear attitude control system of a generic multi-rotor UAV over a stability domain within the dynamical state space of the drone. Further benefits of the procedures are that some of the resulting methods can be implemented onboard the aircraft to detect when its controller breaches its flight envelope limits due to severe weather conditions or actuator/sensor malfunction. Such a detection procedure can be used to advise the remote pilot, or an onboard intelligent agent, to decide on some alterations of the planned flight path or to perform emergency landing.

In the era of Unmanned Aerial Systems (UAS), an onboard autopilot occupies a prominent place and is inevitable for many of their modern applications. The efficacy of autopilot heavily relies upon the accuracy of the sensors employed and the capability of the onboard flight controller. In general, aerodynamic behaviour and flight dynamic capabilities of Unmanned Aerial Vehicles (UAVs) govern the selection and the design of flight controllers. Precise modeling of linear aerodynamic characteristics from flight data can be achieved using many of the existing classical parameter estimation techniques such as Output Error Method (OEM), Equation Error Method (EEM), and Filter Error Method (FEM). However, all the classical methods may not be readily applicable for aerodynamic modeling in nonlinear flight envelopes. The current manuscript is an attempt to exploit the capabilities of the Artificial Intelligence (AI) technique, named Particle Swarm Optimisation (PSO), in combination with Least Squares (LS) cost function to perform linear as well as nonlinear aerodynamic parameter estimation. The aforementioned task is accomplished by considering flight data from manoeuvers pertaining to linear angles of attack, moderate and near stall flight envelopes of two different UAVs with cropped delta planform geometry. Parameters estimated using the proposed LS-PSO method are consistent with minimum standard deviation and are on a par with OEM estimates. The proposed LS-PSO method enhances the capabilities of LS-based EEM while estimating stall characteristic parameters, which was not possible with LS alone. The longitudinal and lateral-directional static parameters estimated from the full-scale wind tunnel testing of the two UAVs were also used to corroborate the results obtained from the flight data using the LS-PSO method.

Although the problem of locked-in deep stall is well documented over many years, there currently exists no consistent procedure that can guarantee recovery. Past studies have suggested that it might be possible to rock the aircraft in pitch to destabilise the statically stable deep stall trim point, thereby gaining enough momentum to push the nose down. However, the methods used in these studies are either of preliminary or empirical nature and cannot guarantee recovery. In this paper, we use bifurcation analysis to derive a recovery manoeuvre, specifically by assessing the aircraft’s nonlinear frequency response under an elevator forcing. The ensuing nonlinear Bode plot detects unstable (divergent) solutions near resonance that contribute to a successful deep stall recovery. Moreover, the nonlinear resonant frequency is slightly lower than the result obtained using linear analysis, and time simulation shows that relying on the linear result does not lead to a successful recovery. It is also found that at the high angles of attack associated with deep stall, the frequency separation between the short period and phugoid mode is significantly reduced, leading to only one visible peak in the frequency response. This feature is also reflected in the time-domain step response.

Reinforcement learning has previously been applied to the problem of controlling a perched landing manoeuvre for a custom sweep-wing aircraft. Previous work showed that the use of domain randomisation to train with atmospheric disturbances improved the real-world performance of the controllers, leading to increased reward. This paper builds on the previous project, investigating enhancements and modifications to the learning process to further improve performance, and reduce final state error. These changes include modifying the observation by adding information about the airspeed to the standard aircraft state vector, employing further domain randomisation of the simulator, optimising the underlying RL algorithm and network structure, and changing to a continuous action space. Simulated investigations identified hyperparameter optimisation as achieving the most significant increase in reward performance. Several test cases were explored to identify the best combination of enhancements. Flight testing was performed, comparing a baseline model against some of the best performing test cases from simulation. Generally, test cases that performed better than the baseline in simulation also performed better in the real world. However, flight tests also identified limitations with the current numerical model. For some models, the chosen policy performs well in simulation yet stalls prematurely in reality, a problem known as the reality gap.

The impact of alternative aviation jet fuels and their properties on lean blowout (LBO) limits has recently raised several questions in the jet fuel area. There is a need for a detailed investigation of the impact of fuel properties on the LBO limit involving actual engine hardware. This study investigates the impact of a range of alternative aviation jet fuels with notable differences in physical and chemical properties and derived cetane number (DCN) on the LBO limit and their effects on key performance indicators. LBO performance results for ten different alternative fuels using a Rolls-Royce single-can Tay combustor are presented in this study. The study also assesses impact of different equivalence ratios and flow rates on LBO, with the aim of determining the impact of a certain range of operating conditions. The results are further analysed to determine the influence of fuel chemical and physical properties on the LBO limit. Finally, based on results in the above experiments, individual fuel properties are adjusted for subsequent experimental analysis of blended fuels. With this approach, 25 additional fuel blends are evaluated and presented, with an emphasis on varying the DCN. This study provides effective data and results to facilitate future fuel optimisation and reduce the risk of a negative performance of new fuels in gas turbines.

The objective of the present work is to estimate the performance of a turbojet engine during Fluidic Thrust Vectoring (FTV) employed by injecting the secondary-jet at the throat of a convergent nozzle. The nozzle performance maps and effective nozzle throat area obtained from experiments are coupled with the performance of a conventional engine (without FTV) using an iterative algorithm developed as a part of this work. The performance is estimated for different flow rates of secondary-jet sourced either from a separate compressor or the engine’s compressor. During FTV, the operating point shifted towards the surge line with increased turbine entry temperature. The desired and obtained vector angles and thrust magnitudes are different. At high secondary-jet flow rates, the turbine operation moved out of its performance map. These aspects should be incorporated while integrating the FTV at the system level, thus, asserting the importance of FTV studies coupled with engine performance.

Airports have been frequently affected by different internal and external disruptive events, which generally deteriorated their planned/regular performances. Their resilience is defined as the ability to withstand and maintain a certain level of functionality of performances compared to their reference regular/planned level during the impact of disruptive events and to recover reasonably rapidly afterwards. Robustness is defined as the level of saved functionality of performances compared to their planned/regular level, enabling continuous operations during the impact of disruptive events. The level of deteriorated functionality compared to that of the planned/regular functionality of performances represents their vulnerability. This paper develops a methodology for assessing resilience, robustness and vulnerability of airports affected by a given disruptive event(s). The methodology consists of analytical models of indicators of operational, economic, social and environmental performances of airports and other main actors/stakeholders involved. These indicators are used as figures-of-merit in analytical models for assessing their cumulative and time-dependent resilience, robustness, and vulnerability. The methodology is applied to assess resilience, robustness and vulnerability of two large airports – LHR (London Heathrow, UK) and NYC JFK (John F. Kennedy, US) – affected by a global and lasting external disruptive event – the COVID-19 pandemic disease. Based on the indicators of operational and economic performances, the results indicate very low resilience and robustness and very high vulnerability of both airports and their other main actors/stakeholders involved. Their resilience and robustness based on the indicators of social and environmental performances were not substantively different from the corresponding vulnerability. In absolute terms, LHR airport has been affected stronger than its NYC JFK counterpart. Savings in costs/externalities during the observed period under the given conditions have modestly compensated for total losses of both airports and their main actors/stakeholders involved.

Spinning/Rolling Airframe Missiles (RAM) mostly use an analog (ON-OFF type) control approach that deflects control surfaces (fins) at minimum and maximum positions continuously, and control is achieved by applying phase shift between the minimum and maximum deflections during a complete roll cycle. Therefore, the control signal shape changes continuously during the roll cycle. In this study, a novel single channel digital controller is designed and tested for a high spinning rate (10–20 Hz) RAM. The digital controller adjusts the amplitude of fin deflections instead of applying a phase shift. In this way, control signal shapes are predetermined in design to completely decouple yaw and pitch dynamics. At the beginning of each roll cycle, the algorithm decides on the control signal shape and amplitude to apply throughout the cycle. Delays on the actuator system and sensor measurements which might lead to instability at high spinning rates are handled effectively thanks to the predetermined control signal shapes that are changing with 90-degree intervals. Detailed geometry of a surface-to-air spinning missile is used to obtain aerodynamic coefficients for the entire flight regime (i.e. from launch to terminal phase) via Missile DATCOM. The 6-DOF flight dynamics model and the controller algorithms are built in MATLAB/Simulink environment. The proposed digital controller is tested systematically for various scenarios and the performance is compared with the conventional analog control approach. The digital controller gives better performance compared to the analog approach under the influence of servo delays and sensor noise.

The electrification of the commuter aircraft is instrumental in the development of novel propulsion systems. The scope of this work aims to explore the design space of a parallel hybrid-electric configuration with an entry into service date of 2030 and beyond and determine the impact of other disciplines on conceptual design, such as components positioning, aircraft stability and structural integrity. Three levels of conceptual sizing are applied and linked with a parametric aircraft geometry tool, to generate the aircraft’s geometry and position the components. Subsequently, the structural optimisation of the wing box is performed, providing the centre of gravity of the components placed inside the wing, that minimise the induced stresses. Furthermore, the stability and trim analysis follow, with the former being highly affected by the positioning of components. Results are compared to a similar aircraft with entry into service technology of 2014 and it is indicated that in terms of block fuel reduction the total electrification benefit increases with the increase of degree of hybridisation, if aircraft mass is kept constant. On the other hand, if battery specific energy is kept constant, similar block fuel reduction is possible with lower hybridisation degrees. The potential for improvement in terms of carbon dioxide emissions and block fuel reduction ranges from 15.73% to 21.44% compared to the conventional aircraft, for levels of battery specific energy of 0.92 and 1.14 kWh/kg respectively. Finally, the component positioning evaluation indicates a maximum weight limitation of 240 kg for the addition of an aft boundary layer ingestion fan to a tube and wing aircraft configuration, without compromising the aircraft static stability.

This paper addresses the potential impact of adverse aircraft-pilot couplings on tiltrotor safety, when a pilot or autopilot attempts to constrain flight dynamics with strong control. The work builds on previously published research on the theory and application of constrained flight to fixed- and rotary-wing aircraft. Tiltrotor aircraft feature characteristics from both types of aircraft and how these determine behaviour in a unique manner is investigated using a FLIGHTLAB simulation model of the XV-15 aircraft. Two different scenarios are explored in detail, using linearised models that reflect the flight-physics of stability for small deviations from trim. First, the control of vertical flight path with longitudinal cyclic pitch and elevator, and the consequences for the stability of the aircraft surge mode and short-period pitch-heave mode. The classical surge-mode instability for flight at speeds below minimum power is shown to apply to the tiltrotor in helicopter mode but alleviated in conversion and airplane modes. The impact on the short–period mode is seen to be a trade-off between the stabilising pitch attitude and destabilising incidence (angle-of-attack) contributions to the flight-path angle. The second example concerns strong control of roll attitude in the presence of adverse aileron-yaw. Here, the yaw-sway motion can be driven unstable, a problem encountered on fixed-wing aircraft with weak weathercock stability, but rare in the rotorcraft world. For both examples, the loss of stability is expressed as the change in sign of effective damping or stiffness stability derivatives. The explanatory theory for these non-oscillatory or low-frequency aircraft-pilot couplings is presented, along with interpretations in terms of handling qualities criteria. The paper also addresses the question of how to translate the findings into a form of aeronautical knowledge useful for the pilot training community.

The simulations and assessment of transient performance of gas turbine engines during the conceptual and preliminary design stage may be conducted ignoring heat soakage and tip clearance variations due to lack of detailed geometrical and structural information. As a result, problems with transient performance stability may not be revealed correctly, and corresponding design iterations would be necessary and costly when those problems are revealed at a detailed design stage. To make an engine design more cost and time effective, it has become important to require better transient performance simulations during the conceptual and preliminary design stage considering all key impact factors such as fuel control schedule, rotor dynamics, inter-component volume effect as well as heat soakage and tip clearance variation effects. In this research, a novel transient performance simulation approach with generically simplified heat soakage and tip clearance models for major gas path components of gas turbine engines including compressors, turbines and combustors has been developed to support more realistic transient performance simulations of gas turbine engines at conceptual and preliminary design stages. Such heat soakage and tip clearance models only require thermodynamic design parameters as input, which is normally available during such design stages. The models have been implemented into in-house transient performance simulation software and applied to a model twin-spool turbojet engine to test their effectiveness. Comparisons between transient performance simulated with and without the heat soakage and tip clearance effects demonstrate that the results are promising. Although the introduced heat soakage and tip clearance models may not be as accurate as that using detailed component geometrical information, it is able to include the major heat soakage and tip clearance effects and make the transient performance simulations and analysis more realistic during conceptual and preliminary engine design stage.

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.

A new type of design domain for a sounding rocket is suggested in this study, which is more intuitive, simplified, and conducive to a single-stage sounding rocket development process. Among the various operation parameters, this study identifies several effective variables, which are also among the most practical in a sounding rocket design process. Of the many design variables considered for peak altitude optimisation, arguably the three most effective in shaping the whole system and conferring the most practicality in the design process of a sounding rocket were determined. A simulation-based study was conducted to establish: the effect of the selected parameters on flight performance, and the optimum design conditions of a single-stage sounding rocket in terms of its peak altitude. The simulation result was compared with randomly chosen experimentally tested flight data and validated. With the clear performance curves varying depending on the variables, the combination of the considered design inputs was effective. The new type of design domain and design procedure proposed is expected to constitute a useful reference and deliver practical benefits to the development process for a target-altitude-optimised single-stage sounding rocket.

The high operational cost of aircraft, limited availability of air space, and strict safety regulations make training of fighter pilots increasingly challenging. By integrating Live, Virtual, and Constructive simulation resources, efficiency and effectiveness can be improved. In particular, if constructive simulations, which provide synthetic agents operating synthetic vehicles, were used to a higher degree, complex training scenarios could be realised at low cost, the need for support personnel could be reduced, and training availability could be improved. In this work, inspired by the recent improvements of techniques for artificial intelligence, we take a user perspective and investigate how intelligent, learning agents could help build future training systems. Through a domain analysis, a user study, and practical experiments, we identify important agent capabilities and characteristics, and then discuss design approaches and solution concepts for training systems to utilise learning agents for improved training value.

Research on propeller performance has been reinvigorated by the development of new classes of vehicles, ranging from electrically powered fixed-wing aircraft, to multi-rotor electrical Vertical Take-off/Landing (eVTOL) vehicles and tilt-rotor aircraft. These types of aircraft utilise a range of modern propellers, often with more advanced planforms and features such as anhedral, and operate in flight envelopes that are outwith the traditional bands of performance. The use of advanced materials (mostly composites), high geometrical sweeps and variable angular velocities are the source of unsteady aerodynamics, that is often coupled with the blade’s structural response. Data from experimental investigations is mostly historic, with the majority of studies conducted before 1960, when aviation shifted rapidly towards jet propulsion. These studies lack in flutter boundary assessment. Modern propellers are likely to be pushed toward their flutter boundaries, but the experimental database published to-date is insufficient to provide flutter boundary assessment. This review examines the value of the available experimental research and the status of the state-of-the-art numerical methods, in order to establish the requirements for modern research on propeller stall flutter.