Establishing aerothermal criteria for swept leading-edge hypersonic vehicle design is the predominant purpose of this work. This study is focused on two different configurations of vehicles based on the swept-back angle ($\varLambda$) viz. minimum drag ($\varLambda$Drag-min), and minimum heat transfer to vehicle ($\varLambda$HT-min). Maximum wall temperatures obtained from the simulation performed in ANSYS 2020 with the k-epsilon turbulence model are 1,013 and 970K for $\varLambda$Drag-min and $\varLambda$HT-min, respectively. These temperatures are used to obtain the corresponding thicknesses of thermal protection systems to maintain inner wall temperature at 323K. Further study is divided into two cases depending on the direction of thickness of thermal protection system with respect to vehicle body. For constant payload capacity, the direction of thickness is outside; whereas for constant overall volume case, direction of thickness is inside. For constant payload volume case, the percentage weight reduction of thermal protection system is 4.8%. For constant overall volume case, the percentage payload capacity increases with design at $\varLambda$HT-min by 4.04% in addition to thermal protection system weight reduction. The lift-induced drag on vehicles with design at $\varLambda$HT-min is significantly reduced for both cases, by 47.68% (for constant payload volume) and 45.27% (for constant overall volume).

In this paper, the results of an experimental investigation for a Y-shaped engine inlet are presented. The experiment is performed at subsonic flow conditions. The main focus is given to time-dependent total pressures measured at the aerodynamic interface plane. Distinctive frequencies carrying high energy contents of the fluctuating total pressures are given and the relation between time-dependent and time-average performance parameters is presented. The cross-correlation coefficients of the high frequency probe readings distributed through the aerodynamic interface plane are also investigated.

In this paper, we present a detailed experimental investigation mainly on the vortical flow fields and the associated vortex breakdown phenomena over a non-slender flying wing (sweep angle, ${\rm{\Lambda }}$ = 53°). In the process, the aerodynamic coefficients were also determined using a six-component force balance. Surface oil flow visualisation, surface pressure measurements and particle image velocimetry (PIV) measurements, in various crossflow planes and in a longitudinal plane passing through the leading-edge vortex core, were carried out at various Reynolds numbers to understand the flow field over the non-slender flying wing. Aerodynamic characteristics of the flying wing show local peaks and valleys in the pitching moment coefficient. The surface flow visualisation reveals that the nonlinearity of the pitching moment curve is due to the complex nature of vortical flow structures. The flow visualisation also demonstrates the presence of a wave-like surface pattern, and its size is found to reduce with increasing Reynolds numbers. The present PIV measurements confirm that this wave-like surface pattern is associated with vortex breakdown phenomena. These measurements also reveal that the vortex breakdown has not reached the apex of the wing, even at post-stall angle-of-attack. For pre-stall ($\alpha $ = 20°) flow regimes, it is observed that the location of the vortex breakdown moves downstream as the Reynolds number increases, but this influence is minimised at near-stall ($\alpha $ = 25°) and post-stall ($\alpha $ = 30°) flow regimes. Reconstructed velocity field using the first 10 dominant proper orthogonal decomposition (POD) modes reveals that the nature of the vortex breakdown over the flying wing is a spiral-type vortex breakdown.

A synthetic flow angle sensor, able to estimate angle-of-attack and angle-of-sideslip, can exploit different methods to solve a set of equations modelling data fusion from other onboard systems. In operative scenarios, measurements used for data fusion are characterised by several uncertainties that would significantly affect the synthetic sensor performance. The off-line use of neural networks is not a novelty to model deterministic synthetic flow angle sensors and to mitigate issues arising from real flight applications. A common practice is to train the neural network with corrupted data that are representative of uncertainties of the current application. However, this approach requires accurate tuning on the target aircraft and extensive flight test campaigns, therefore, making the neural network tightly dependent on the specific aircraft. In order to overcome latter issues, this work proposes the use of neural networks to solve a model-free scheme, derived from classical flight mechanics, that is independent from the target aircraft, flight regime and avionics. It is crucial to make use of a training dataset that is not related to any specific aircraft or avionics to preserve the generality of the scheme. Under these circumstances, global and local neural networks are herein compared with an iterative method to assess the neural capabilities to generalise the proposed model-free solver. The final objective of the present work, in fact, is to select the neural technique that can enable a flow angle synthetic sensor to be used on board any flying body at any flight regime without any further training sessions.

In this paper, an adaptive neural output-constrained control algorithm is proposed for a class of non-affine kinetic kill vehicle (KKV) systems. The key point is that the non-affine control law can be designed and the output of the KKV system conform to the output limit with the aid of the proposed method. Due to the aerodynamic moments, the actual control torque is non-affine, which can be addressed by introducing an integral process to the design of the controller. Besides, in order to improve the control precision, a nonlinear mapping is put forward so that the output constraint can be transformed to the constraint of the introduced dynamic signal that can be simply achieved. From the simulation results it can be concluded that the states of the KKV system can track the desired trajectories in spite of different working conditions and the control precision is higher compared with other control methods.

In this paper, a new approach to modeling and controlling the problems associated with a morphing unmanned aerial vehicle (UAV) is proposed. Within the scope of the study, a dataset was created by obtaining a wide range of aerodynamic parameters for the UAV with Ansys Fluent under variable conditions using the computational fluid dynamics approach. For this, a large dataset was created that considered 5 different angles of attack, 14 different swept angles, and 5 different velocities. While creating the dataset, the analyses were verified by considering studies that have been experimentally validated in the literature. Then, an artificial intelligence-based model was created using the dataset obtained. Metaheuristic algorithms such as the artificial bee colony algorithm, ant colony algorithm and genetic algorithms are used to increase the modeling success of the adaptive neuro-fuzzy inference system (ANFIS) approach. A novel modeling approach is proposed that constitutes a new decision support system for real-time flight. According to the results obtained, all the ANFIS models based on metaheuristic algorithms were more successful than the traditional approach, the multilinear regression model. The swept angle that meets the minimum lift needed by the UAV for different flight conditions was estimated with the help of the designed decision support system. Thus, the drag force is minimised while obtaining the required lift force. The performance of the UAV was compared with the nonmorphing configuration, and the results are presented in tables and graphs.

The present paper describes the results of an experimental wind tunnel test campaign aimed at investigating the aerodynamic performance and flow physics related to a wing section equipped with two propellers mounted on a boom. The configuration investigated is meant to be representative of a full-scale eVTOL aircraft in cruise flight condition. The use of full-scale components of an eVTOL aircraft made this setup a quite advanced experiment in the recent literature. Pressure measurements and an infrared thermography technique were used during the test campaign, respectively, to evaluate localised effects induced by the propeller blowing on the wing and to provide a quantitative evaluation of the amount of laminar flow on the wing surface with and without the influence of the propeller at different thrust conditions.

It is envisaged that future civil aero-engines will operate with ultra-high bypass ratios to reduce the specific fuel consumption. To achieve the expected benefits from the new engine cycles, these new powerplants may mount compact nacelles. For these new configurations the aerodynamic coupling between the powerplant and the airframe may increase. For this reason, it is required to quantify and further understand the effects of aircraft integration for compact aero-engine nacelles. This study provides an insight of the changes in flow aerodynamics as well as quantification of the most relevant performance metrics of the powerplant, airframe and the combined aircraft system across a range of different installation positions. Relative to a conventional architecture, there is an aerodynamic benefit in net vehicle force of about 1.2% for a compact powerplant when installed in forward positions. This is the same improvement that was identified when the aero-engine nacelles were in isolation. However, for close-coupled installation positions, the aerodynamic benefit in net vehicle force erodes to 0.44% due to the larger effects of aircraft integration on compact nacelles.

This study investigates the combustion characteristics (lean blowoff (LBO), flashback, and combustion instability) of sustainable aviation fuels (SAFs) using a two-staged combustor arranged in tandem. The combustor consists of a first (main) burner with an axial swirler and a second burner (afterburner) with a V-gutter bluff-body. The main burner utilises a constant flow rate of CH4/Air as reactants, while the afterburner operates on the combustion product from the first burner, additional dilution air, and one of the six liquid test fuels (Jet-A, A-2, C-1, C-5, C-9, and JP-8/HRJ) as reactants. When test fuels with higher derived cetane number (DCN) are used in the afterburner, it is observed that LBO occurs at lower afterburner global equivalence ratios (${\phi _{af}}$). This is due to the smaller chemical time scale of those test fuels, which enables the flame to be sustained under higher flow speeds. This is consistent with the results of previous studies. Additionally, this study shows that flashback also occurs at lower equivalence ratio for test fuels with higher DCN. This is a new finding and can be explained in a similar way to LBO. As the operating conditions approach the LBO or exceed flashback limits, the reactivity of shear layer also decreases, creating favourable conditions for them to occur. Moreover, for test fuels with low DCN (C-1 and C-5), blowoff occurs instead of flashback as ${\phi _{af}}$ increases. This can be attributed to the stronger pressure perturbation resulting from combustion instability in these flames compared to others. It demonstrates a stronger correlation between the pressure perturbation and the heat release rate fluctuation for C-1 and C-5 than that of other test fuels at the same equivalence ratio. In conclusion, it was found that test fuels with higher DCN have advantages in terms of LBO and stability, however, they are more prone to flashback.

Traditional active flutter suppression controllers are designed based on model. However, as the aircraft becomes more and more powerful, the modeling of aeroelastic system becomes difficult and the model-free requirement of controller design becomes more and more urgent. The complexity of industrial processes has brought about massive operational data generated online. Aviation industry development has entered the era of big data. Breaking through the traditional theoretical framework, mining the correlation, evolution and dynamic characteristics of the system from the data is the inevitable choice to meet this demand. In this paper, a data-driven model-free controller is designed, which relies on ridge regression of the input and output variation at each operating point of the closed-loop controlled system to recursively derive the iterative format of the control signals and ensure the numerical stability of the signals. The controller can only use the real-time measurement of the system’s online input and output data for continuous correction, to achieve the purpose of flutter suppression. Then flutter suppression of a three-degree-of-freedom binary wing with a control surface is studied, and the superiority of model-free controller is demonstrated by comparing it with the optimal controller.

The aerodynamic performance of a wing model with a row of distributed engines are investigated at the vertical take-off condition. The engines are installed near the trailing edge of the wing. During vertical take-off, the jets exit from the engines and impinge perpendicularly to the ground, providing a thrust for the aircraft. Due to the ground effects, complex vortex structures are induced by the jets. The vortices are categorised into the spanwise vortices and the chordwise vortices. The underwing vortices can lead to low-pressure regions on the lower surface of the wing, resulting in an undesirable downward force. The underwing vortex structures are affected by the ratio of the engine distance to the engine diameter ($S/D$). At a small $S/D$ = 1.10, the flow field is dominated by the spanwise vortices; at a large $S/D$ = 2.78, the flow field is dominated by the chordwise vortices. The range and strength of the spanwise vortices are affected by the vortices interaction. Competition mechanism exists between the range and strength effects, which results in the non-linear variation of the wing lift coefficient with engine spacing. The details of the flow physics underneath the wing and its mechanism on the lift of the wing during take-off are investigated.

This paper aims to present a vertical take-off and landing unmanned aerial vehicle (VTOL UAV) configuration and numerically simulate its flight transition from hover to cruise and from cruise to hover. It can tilt the canard and wing along with two attached propellers. Additionally, two fixed front propellers are pointing upwards. Multi-body equations of motion are derived for this concept of aircraft, which are used to compute the flight transition trajectory from hover to cruise configuration. Furthermore, a transition control algorithm based on gain scheduling is described, which stabilises the aircraft while it accelerates from hover to cruise, gradually tilting the wing along with its propellers, sequentially switching between equilibrium states, as the stability cost functions thresholds are reached. The transition control algorithm of the conceptual aircraft model is numerically simulated.

Large truss structures have many potential applications in space, such as antennas, telescopes and space solar power plants. In this scenario, a natural concern is the susceptibility of these lightweight structures to be damaged during their operational life, due to impacts, transient thermal states and fatigue phenomena. The inclusion of active elements, equipped with sensor/actuator systems capable of modulating their shape and strength, makes it possible to transform the truss into a smart structure capable of remedying the damage, once it is detected. In this paper, a procedure is described that is capable of restoring at least the basic functionality of a composite truss for space applications, starting with the observation that damage has occurred, regardless of its specific location. The system eigenstructure is used as a benchmark for damage detection, as well as a target characteristic for the subsequent restoration activity. The observer/Kalman filter identification algorithm (OKID), in cascade with the eigensystem realization algorithm (ERA), is adopted to reconstruct, from sensor recordings, the dynamic response of the truss in terms of system state-space representation and eigen-characteristics. Finally, a static output feedback control is developed to recover the low-frequency dynamic behaviour of the truss. The entire procedure is tested using finite element analysis. All activities are coordinated in an innovative procedure that, within a unique Python language code, automatically generates finite element (FE) models, launches finite element analysis (FEA), extracts output data, implements OKID-ERA, processes the control law and applies it to the final FE simulation.

Lateral jets are used to control the missiles and re-entry vehicles at high altitudes. The objective of the research paper is to investigate the effect of lateral jet interaction with the external flow on a blunted nose cone re-entry vehicle configuration and its flight speed. Structured mesh is used for the simulations, and the computational analysis is carried out by Ansys Fluent solver. The simulation results are validated, and the same methodology used for the parametric analysis. Simulations have been carried out at an external flow Mach number 6, 8.1, 12 and 16 at five degree angle-of-attack for jet-off and jet-on conditions. At 8.1 Mach number, the normal force coefficient is decreased by 45.6% due to jet interaction. The lateral jet interaction effectively reduces the nose down pitching moment. At 8.1 Mach number, the pitching moment coefficient was reduced by 48% with the jet-on condition compared to the jet-off condition.

The present work aims to extend the capabilities of DUST, a mid-fidelity aerodynamic solver developed at Politecnico di Milano, for the aerodynamic simulation of rotorcraft applications. With this aim, a numerical element was implemented in the solver obtained by a coupling between the potential unsteady vortex lattice method and viscous aerodynamic data of aerofoil sections available from two-dimensional high-fidelity computational fluid dynamics (CFD) simulations or experimental wind-tunnel tests. The paper describes the mathematical formulation of the method as well as a validation of the implementation performed by comparison with both high-fidelity CFD simulation results and experimental data obtained over aerodynamics and aeroelastic fixed-wing benchmarks. Then, the method was used for the evaluation of the aerodynamic performance of two rotorcraft test cases, i.e. the full-scale proprotor of the XV-15 tiltrotor operating in different flight conditions and two propellers in tandem with overlapping disks. Simulation results comparison with high-fidelity CFD and data from wind tunnel tests highlighted the potentialities and advantages of the implemented approach to be used for the design and investigation of rotorcraft configurations characterised by consistent viscosity effects.

Blowing is often used to alleviate the intense heating rates on blunt noses of hypersonic vehicles. This flow efflux at the leading edge transforms the flow field in the blunt-nose regions with implications on the dynamic stability of the vehicles. As a demonstrative exercise, the flow fields past blunt-nosed and truncated-nosed conical bodies under blowing and no-blowing conditions were perturbed to obtain the unsteady effects using the shock expansion method to recover the unsteady pressure coefficient. Static and pitching moment derivatives were then duly obtained by integrating the differential of the unsteady pressure coefficient with respect to the pitch angle (α) or the pitch rate ($\dot \theta $) together with the moment arm with reference to the centre of gravity. The results obtained for blunt-nose and truncated conical bodies show a noticeable drop in dynamic stability. Even when the flow is transformed from a tangential blowing at the shoulder of the blunt-nosed vehicle shows some degradation in dynamic stability.

Algeria’s micro-satellite, Alsat-1b, was successfully launched into a 680 km low Earth orbit onboard a PSLV-C35 rocket from Sriharikota, South India, on September 26, 2016. The spacecraft was conceived, built and launched as part of an 18-month technology transfer programme between Algeria’s Algerian Space Agency (ASAL) and the United Kingdom’s Surrey Satellite Technology Limited (SSTL). This document details the Power Conditioning and Distribution Module’s (PCM-PDM) design and performance in orbit, critical component of a satellite electrical power system, responsible for converting, regulating and distributing power to various subsystems and payloads. The PCM-PDM developed and produced by SSTL was subjected to rigorous testing simulating harsh space conditions to assess its performance. The results of this comprehensive analysis indicate that the module can effectively withstand extreme environmental factors and function optimally in challenging settings. The analysis focused on the PCM-PDM’s ability to provide reliable and efficient power conditioning and distribution to the satellite, including its load management capabilities, overcurrent protection, protection against undervoltage and critical mode operations. The results of the performance analysis showed that the PCM-PDM met the required specifications and demonstrated reliable and efficient operation in different modes of the satellite’s mission. The study highlights the importance of careful design and rigorous testing of the PCM-PDM to ensure the reliable and efficient operation of the satellite and its payloads.

This paper studied the back-stepping adaptive sliding mode control (SMC) attitude problem of quaternion aircraft model based on radial basis function (RBF) network approximation. Firstly, a sliding mode controller is designed based on the back-stepping method (BSM) for the nonlinear aircraft model. Secondly, a RBF network algorithm is designed to compensate for the unknown and uncertain parts of the aircraft system. RBF network has simple network structure and good generalisation ability, avoids lengthy and unnecessary calculations, realises adaptive approximation of unknown parts in the aircraft model, and through the adjustment of adaptive weights, the convergence and stability of the entire closed-loop system (CLS) are guaranteed. Finally, the anti-interference performance of the controller is verified by simulation of the actuator fault model. Our proposed method has all-right control performance indicated by the simulation results.

On-demand urban air transportation gains popularity in recent years with the introduction of the electric VTOL (eVTOL) aircraft concept. There is an emerging interest in short/medium range eVTOL air taxi considering the critical advantages of electric propulsion (i.e. low noise and carbon emission). Using several electric propulsion systems (distributed electric propulsion (DEP)) has further advantages such as improved redundancy. However, flight controller design becomes more challenging due to highly over-actuated and coupled dynamics. This study defines and resolves flight control problems of a novel DEP eVTOL air taxi. The aircraft has a fixed-wing surface to have aerodynamically efficient cruise flight, and uses only tilting electric propulsion units to achieve full envelope flight control via pure thrust vector control. The aircraft does not have conventional control surfaces such as aileron, rudder or elevator. Using pure thrust vector control has some design benefits, but the control problem becomes more challenging due to the over-actuated and highly coupled dynamics (especially in transition flight). A preliminary flight dynamics model is obtained considering the dominant effects at hover and high-speed forward flight. Hover and forward flight models are blended to simulate the transition dynamics. Two central challenges regarding the flight control are significant nonlinearities in aircraft dynamics during the transition and proper allocation of the thrust vector control specifically in limited control authority (actuator saturation). The former challenge is resolved via designing a sensor-based incremental nonlinear dynamic inversion (INDI) controller to have a single/unified controller covering the wide flight envelope. For the latter one, an optimisation-based control allocation (CA) approach is integrated into the INDI controller. CA requires special attention due to the pure thrust vector control’s highly coupled dynamics. The controller shows satisfactory performance and disturbance rejection characteristics. Moreover, the CA plays a vital role in guaranteeing stable flight in case of severe actuator saturation.

The aerodynamic response of a NACA0012 wing section was investigated at a Reynolds number of 100,000 in an open return wind tunnel in the presence of a second wing in tandem. The angle-of-attack of the front wing ranged from −5° to 90° while the rear wing remained at zero incidence. The presence of the downstream wing significantly altered the post-stall behaviour of the upstream wing in the form of a secondary stall characterised by a sudden drop in lift and drag for a specific combination of angle-of-attack and the spacing between the wings. The secondary stall was found to be insensitive to the Reynolds number and the aspect ratio of the downstream wing and did not affect the lift-to-drag ratio. Flow visualisation in the water tunnel indicated that the downstream wing effectively suppressed vortex shedding and lift fluctuations of the upstream wing.