In this paper, we investigate the attitude manoeuver planning and tracking control of the flexible satellite equipped with a coilable mast. Due to its flexible beamlike structure, the coilable mast experiences bending and torsional modal vibrations in multi-direction. The complex nonlinear coupling and other external disturbances significantly impact the achievement of high-precision attitude control. To overcome these challenges, a robust attitude tracking controller is proposed for easy implementation by the Attitude Determination and Control System (ADCS). The controller consists of a disturbance compensator, feedforward controller and output feedback controller. The compensator, based on a Nonlinear Disturbance Observer (NDO), effectively compensates for the cluster disturbances caused by vibrations, environmental factors and parameter perturbations. The feedforward controller tracks the desired path in the nominal satellite model. Furthermore, the output feedback controller enables large-angle manoeuver control of the satellite and evaluates the suppression effect of the controlled output on the observation error of cluster disturbances used the ${L_2}$-gain. Simulation results demonstrate that the proposed controller successfully achieves high-precision attitude tracking control during large-angle manoeuvering.

With the wide application of quadrotor unmanned aerial vehicles (UAVs), the requirements for their safety and reliability are becoming increasingly stringent. In this paper, based on the feedback of airframe performance health perception information and the predictive function control strategy, the autonomous maintenance of a quadrotor UAV with multi-actuator degradation is realised. Autonomous maintenance architecture is constructed by the predictive maintenance (PdM) idea and the Laguerre function model predictive pontrol (LF-MPC) strategy. Using the two-stage Kalman filter (TSKF) method, based on the established UAV degradation model, the aircraft state and actuator degradation state are predicted simultaneously. For the predictive perception of system health, on the one hand, the system health degree (HD) based on Mahalanobis distance is defined by the degree of airframe state deviation from the expected state, and then the failure threshold of the UAV is obtained. On the other hand, according to the degradation state of each actuator, a comprehensive degradation variable fused with different weight coefficients of multiple actuators degradation is used to obtain the probability density function (PDF) of remaining useful life (RUL) prediction. For the autonomous maintenance of system health, the LF-MPC weight matrixes are adjusted adaptively in real-time based on the HD evaluation, to achieve a compromise balance between UAV performance and control effect, and greatly extend the working time of UAV. Simulation results verified the effectiveness of the proposed method.

The current paper is focused on the conceptual design of a thermal management system with a liquid working medium for a commuter hybrid-electric aircraft, featuring a series propulsion configuration. Regarding the system’s architecture, parametric analyses are conducted, by altering the number of heat exchangers. To clarify, a centralised and a decentralised thermal management system architecture are examined. Furthermore, a computational model calculates the temperatures during the system’s operation and the required coolant mass flows to sufficiently cool all the compartments. Subsequently, the required heat exchanger surface is determined and the weight of each compartment that comprises the thermal management system can be calculated. It is worth noting, that the compartments’ cold plate weight is integrated. The results indicate that the decentralised configuration results in lower temperature fields for all components compared to the centralised configuration. However, the latter weighs 32.2% lower at 158.22kg while the decentralised configuration weighs 233.48kg.

This research paper presents an application of the integrated process and product design (IPPD) approach for selecting the best joint configuration for dissimilar material joining in the early product design phase. The proposed methodology integrates the multi-criteria decision making (MCDM) approach with quality function deployment (QFD) to identify the key criteria for joint selection, including load-carrying capacity, size, cost per joint, ease of manufacturing, time consumption and deformation. Three types of joints (rivet, weld and adhesive) and two hybrid joints (adhesive-weld and adhesive-rivet) are considered for three dissimilar material configurations: carbon fiber-reinforced plastic (CFRP) aluminum, CFRP steel, and aluminum-steel. QFD is utilised to transform job requirements into design criteria, and in the second phase, the Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) is employed to choose the optimal joint configuration based on the weighted criteria acquired in the previous phase. The selected joint configuration is then validated through experimental study. The findings demonstrate that the proposed IPPD approach with QFD-TOPSIS techniques is highly effective for selecting mechanical joints for dissimilar material joining in the early design phase. The study concludes that the adhesive-rivet hybrid joint is the optimal solution among all alternatives. The proposed methodology can ultimately lead to improved product reliability and performance, as well as reduced development time and cost.

A rapid nonlinear aeroelastic framework for the analysis of the highly flexible wing with distributed propellers is presented, validated and applied to investigate the propeller effects on the wing dynamic response and aeroelastic stability. In the framework, nonlinear beam elements based on the co-rotational method are applied for the large-deformation wing structure, and an efficient cylinder coordinate generation method is proposed for attached propellers at different position. By taking advantage of the relatively slow dynamics of the high-aspect-ratio wing, propeller wake is modeled as a quasi-steady skewed vortex cylinder with no updating process to reduce the high computational cost. Axial and tangential induced velocities are derived and included in the unsteady vortex lattice method. For the numerical cases explored, results indicate that large deformation causes thrust to produce wing negative torsion which limits the displacement oscillation, and slipstream mainly increases the response values. In addition, an improvement of flutter boundary is found with the increase of propeller thrust while slipstream brings a destabilising effect as a result of the increment of dynamic pressure and local lift. The great potential of distributed propellers in gust alleviation and flutter suppression of such aircraft is pointed out and the method as well as conclusions in this paper can provide further guidance.

Carefully designing blade geometric parameters is necessary as they determine the aerodynamic performance of a rotor. However, manufacturing inaccuracies cause the blade geometric parameters to deviate randomly from the ideal design. Therefore, it is essential to quantify uncertainty and analyse the sensitivity of the blade geometric deviations on the compressor performance. This work considers a subsonic compressor rotor stage and examines samples with different geometry features using three-dimensional Reynolds-averaged Navier-Stokes simulations. A method to combine Halton sequence and non-intrusive polynomial chaos is adopted to perform the uncertainty quantitative (UQ) analysis. The Sobol’ index and Spearman correlation coefficient help analyse the sensitivity and correlation between the compressor performance and blade geometric deviations, respectively. The results show that the fluctuation amplitude of the compressor performance decreases for lower mass flow rates, and the sensitivity of the compressor performance to the blade geometrical parameters varies with the working conditions. The effects of various blade geometric deviations on the compressor performance are independent and linearly superimposed, and the combined effects of different geometric deviations on the compressor performance are small.

The deceleration effect of the deployable aerodynamic decelerator is not as good as a parachute in the subsonic region. This paper proposes a novel concept of using a parachute-like configuration (PLC) to enhance the deceleration performance of the mechanically deployable aerodynamic decelerator (MDAD) through structural transformation. The MDAD turned into the PLC from the sphere cone configuration (SCC) at Ma 0.8. The aerodynamic characteristics of the two configurations are analysed deeply. Compared to the SCC, the results show that the drag coefficient increases effectively, and the maximum increases is about 10% in the PLC. The airflow is altered by the MDAD configuration, which can affect the surface pressure and temperature. During the transformation process, the axial and normal force coefficients tend to stabilise. However, the static stability of the PLC deteriorates sharply compared to the SCC when the angle-of-attack exceeds 45°.

Aerodynamic design of a high-efficiency two-stage axial turbine is carried out using a hybrid method through implantation of a two-step design procedure. In the first step, the well-known streamline curvature (SLC) and free vortex (FV) methods are properly combined to establish three-dimensional geometries of the blades at each row and to obtain the flow field properties. The second step is provided to obtain the highest aerodynamic efficiency by optimum clocking of the second stator blades relative to the first ones through executing steady and unsteady computational fluid dynamics (CFD) of three-dimensional viscous flow. Slight discrepancies were observed between gas dynamics results of the SLC and those of CFD. Total pressure and temperature at the turbine outlet, obtained from SLC method, differed from those obtained by 3D-CFD technique by 13.06% and 1.88% respectively. Aerodynamic efficiency of the turbine is obtained about 91.83%, based on 3D-CFD. Time-averaged results showed that under the optimum clocking of the second row stator blades, inlet total pressure and output power of the second rotor increase by 0.23%, and 0.93%, respectively, in comparison to the worst clocking case. These augmentations resulted in increased total to total efficiency of the second stage by 0.444%. Additionally, the total output power of the two stages increased by 0.71% through the optimum clocking. Modeling the unsteady wake flow trajectory within the blades passages confirmed that all of these beneficial effects happen if the upstream wake impinges on the leading edge region of the second stator blades.

This paper is focused on the stability of real-time hybrid aeroelastic simulation systems for flexible wings. In a hybrid aeroelastic simulation, a coupled aeroelastic system is ‘broken down’ into an aerodynamic simulation subsystem and a structural vibration testing subsystem. The coupling between structural dynamics and aerodynamics is achieved by real-time communication between the two subsystems. Real-time hybrid aeroelastic simulations can address the limitations associated with conventional aeroelastic testing performed within a wind tunnel or with pure computational aeroelastic simulation. However, as the coupling between structural dynamics and aerodynamics is completed through the real-time actuation and sensor measurement, their delays may inherently impact the performance of hybrid simulation system and subsequently alter the measured aeroelastic stability characteristics of the flexible wings. This study aims to quantify the impact of actuation and sensor measurement delays on the measured aeroelastic stability, e.g. the flutter boundary, of flexible wings during real-time hybrid simulations, especially when different aerodynamic models are implemented.

Microvibrations originating from onboard disturbance sources can lead to a range of issues, including a decrease in satellite pointing accuracy, image distortion and blurring. Therefore, reaction wheels emerge as the primary sources of disturbance noise. This paper employs an experimental approach based on the real dynamics of rotating reaction wheel assembly, closely simulating on-orbit configurations to measure noise responses transferred to the satellite structure. An assessment of noise response behaviour, incorporating a comprehensive understanding of the factors influencing the levels, was conducted on a proto-flight satellite for three reaction wheels. Initially, reaction wheel assemblies underwent individual iterative balancing to reduce mass deviations. Subsequently, amplitude-time responses at different rotational speeds of reaction wheel assemblies (RWA) disturbance noise were measured. The experimental results demonstrate that each individually balanced reaction wheel generates independent perturbation noise level due to manufacturing imperfections. Hence, the necessity of wheels testing for accurate prediction and mitigation of disturbance levels is crucial, especially for payloads sensitive to microvibrations. Furthermore, increasing wheel speeds proportionally amplify disturbance noise levels. Therefore, implementing an optimised mission attitude control profile with lower rotation speeds of reaction wheels effectively reduces microvibration levels which minimises risks to payload performance and reduce power consumption.

In order to investigate the three-dimensional effects on the flow characteristics of the thin water film for the three-dimensional wings, the numerical simulation of the droplet impingement and film flow on the MS-0317 wing is implemented based on the open-source package OpenFOAM. The simulation focuses on the effects of the angle-of-attack and the angle of sweepback. The movement and impingement of the droplets are calculated using the Lagrangian method, and the film flow is simulated using the thin film assumption and the finite area method. The simulation of the water film flow of the three-dimensional MS-0317 wing shows that there is a spanwise flow of the water film due to the three-dimensional effects. This suggests that more research should be conducted on the warm glaze ice with surface water film of three-dimensional ice accretion on three-dimensional geometries.

Effusion cooling is one of the significant cooling technologies in combustor liners in terms of cooling efficiency and weight reduction. However, effusion cooling technology is difficult to manufacture. In fact this technology requires laser-drilling of thousands of tiny holes with shallow angles on a sheet metal with a thickness generally varying between 0.5 to 1.5 mm. In addition, the use of thermal barrier coating is common in gas turbine engines and is one more challenge for the drilling process. In order to obtain more efficient gas turbine engines, the inlet temperature keeps increasing in the last decades, which induces the combustion chamber to operate in a hotter environment. Therefore, efficient cooling technology is needed, even if it is hard to manufacture. For laser drilling, several parameters have to be explored to obtain acceptable holes. This study includes the microstructure investigations of the holes produced with different laser parameters and the optimal laser parameters determined according to the microstructure of six different effusion cooling hole configurations. The results show that laser process differences affect the metal substrate microstructure and thermal barrier coating structure. Drilling method, peak power, number of pulses, gas type and pressure value have a significant effect on the hole geometry and its microstructure.

Changes in flight stability characteristics at the advanced stage of aircraft design are complex and require thorough investigations. This paper examines the impact of wing strake modification on high-performance aircraft using computational fluid dynamics (CFD). The dynamic behaviour is calculated using the forced oscillation technique, while the effect of geometric variation on longitudinal stability characteristics is extensively studied. Steady-state experimental data is utilised to validate the computational setup. Static aerodynamic coefficients, dynamic stability derivatives and the positions of aerodynamic and pressure centres are employed to quantify the changes. Furthermore, the alterations in stability characteristics are correlated with flow physics. The results indicate a reduction in longitudinal static and dynamic stability at various flight conditions due to the proposed modification. This deliberate reduction was necessary to accommodate the installation of a fly-by-wire system. The discussed design changes have been effectively implemented on an in-service aircraft.

Sensed data from high-value engineering systems is being increasingly exploited to optimise their operation and maintenance. In aerospace, returning all measured data to a central repository is prohibitively expensive, often causing useful, high-value data to be discarded. The ability to detect, prioritise and return useful data on asset and in real-time is vital to move toward more sustainable maintenance activities.

We present a data-driven solution for on-line detection and prioritisation of anomalous data that is centrally processed and used to update individualised digital twins (DT) distributed onto remote machines. The DT is embodied as a convolutional neural network (CNN) optimised for real-time execution on a resource constrained gas turbine monitoring computer. The CNN generates a state prediction with uncertainty, which is used as a metric to select informative data for transfer to a remote fleet monitoring system. The received data is screened for faults before updating the weights on the CNN, which are synchronised between real and virtual asset.

Results show the successful detection of a known in-flight engine fault and the collection of data related to high novelty pre-cursor events that were previously unrecognised. We demonstrate that data related to novel operation are also identified for transfer to the fleet monitoring system, allowing model improvement by retraining. In addition to these industrial dataset results, reproducible examples are provided for a public domain NASA dataset.

The data prioritisation solution is capable of running in real-time on production-standard low-power embedded hardware and is deployed on the Rolls-Royce Pearl 15 engines.

This paper proposes a composite non-singular fast terminal sliding mode attitude control scheme based on a reduced-order extended state observer for the stratospheric airship’s attitude system affected by multiple disturbances. First, the feedback linearisation method is applied to address the nonlinearity of the attitude motion model and achieve decoupling of the model in three channels. Second, the overall disturbances, encompassing airship parameter perturbations and external disturbances, are treated as an aggregate. A reduced-order extended state observer is designed for each channel to formulate a composite non-singular fast terminal sliding mode surface. In the control design phase, the hyperbolic sine function is adopted as replacement for the sign function to ensure the continuity of the control signal. The estimated disturbances are incorporated in the control law design to directly offset the effects of multiple disturbances on the attitude motion of the airship. Third, based on Lyapunov theory, it has been proven that the control law can drive the attitude tracking error to converge to zero within a finite time. Simulation results demonstrate that the proposed control scheme exhibits favorable disturbance rejection capability, as well as higher tracking accuracy and faster response speed.

Aircraft ground taxiing contributes significantly to carbon emissions and engine wear. The electric towing tractor (ETT) addresses these issues by towing the aircraft to the runway end, thereby minimising ground taxiing. As the complexity of ETT towing operations increases, both the towing distance and time increase significantly, and the original method for estimating the number of ETTs is no longer applicable. Due to the substantial acquisition cost of ETT and the need to reduce waste while ensuring operational efficiency, this paper introduces for the first time an ETT quantity estimation model that combines simulation and vehicle scheduling models. The simulation model simulates the impact of ETT on apron operations, taxiing on taxiways and takeoffs and landings on runways. Key timing points for ETT usage by each aircraft are identified through simulation, forming the basis for determining the minimum number of vehicles required for airport operations using a hard-time window vehicle scheduling model. To ensure the validity of the model, simulation model verification is conducted. Furthermore, the study explores the influence of vehicle speed and airport scale on the required number of ETTs. The results demonstrate the effective representation of real-airport operations by the simulation model. ETT speed, airport runway and taxiway configurations, takeoff and landing frequencies and imbalances during peak periods all impact the required quantity of ETTs. A comprehensive approach considering these factors is necessary to determine the optimal number of ETTs.

This contribution focuses on the abatement with hydrogen of CO2 and non-CO2 emissions. It is agenda-setting in two respects. Firstly, it challenges the globally accepted hydrocarbon sustainable aviation fuel (SAF) pathway to sustainability and recommends that our industry accelerates along the hydrogen pathway to ‘green’ aviation. Secondly, it reports a philosophical and analytical investigation of appropriate accuracy on abatement strategies for nitrogen oxides and contrails of large hydrogen airliners. For the second contribution, a comparison is made of nitrogen oxide emissions and contrail avoidance options of two hydrogen airliners and a conventional airliner of similar passenger capacity. The hydrogen aircraft are representative of the first and second innovation waves where the main difference is the weight of the hydrogen tanks. Flights of 1000, 2000, 4000 and 8000 nautical miles are explored. Cranfield’s state of the art simulators for propulsion system integration and gas turbine performance (Orion and Turbomatch) were used for this. There are two primary contributions to knowledge. The first is a new set of questions to be asked of SAF and hydrogen decarbonising features. The second is the quantification of the benefits from hydrogen on non-CO2 emissions. For the second generation of long-range hydrogen-fuelled aircraft having gas turbine propulsion, lighter tanks (needing less thrust and lower gas temperatures) are anticipated to reduce NOx emissions by over 20%; in the case of contrails, the preliminary findings indicate that regardless of the fuel, contrails could largely be avoided with fuel-burn penalties of a few per cent. Mitigating action is only needed for a small fraction of flights. For conventional aircraft this penalty results in more CO2, while for hydrogen aircraft the additional emission is water vapour. The conclusion is that our research community should continue to consider hydrogen as the key ‘greening’ option for aviation, notwithstanding the very significant costs of transition.

Aiming at alleviating the adverse influence of coupling unmodeled dynamics, actuator faults and external disturbances in the attitude tracking control system of tilt tri-rotor unmanned aerial vehicle (UAVs), a neural network (NN)-based robust adaptive super-twisting sliding mode fault-tolerant control scheme is designed in this paper. Firstly, in order to suppress the unmodeled dynamics coupled with the system states, a dynamic auxiliary signal, exponentially input-to-state practically stability and some special mathematical tools are used. Secondly, benefiting from adaptive control and super-twisting sliding mode control (STSMC), the influence of the unexpected chattering phenomenon of sliding mode control (SMC) and the unknown system parameters can be handled well. Moreover, NNs are employed to estimate and compensate some unknown nonlinear terms decomposed from the system model. Based on a decomposed quadratic Lyapunov function, both the bounded convergence of all signals of the closed-loop system and the stability of the system are proved. Numerical simulations are conducted to demonstrate the effectiveness of the proposed control method for the tilt tri-rotor UAVs.

High-temperature non-equilibrium effects are prominent in scramjet nozzle flows at high Mach numbers. Hence, the thermochemical non-equilibrium gas model incorporating the vibrational relaxation process of molecules in the hydrocarbon-air reaction is developed to numerically simulate the flow of a hydrocarbon fuel scramjet nozzle at Mach 10. Besides, the results computed by the models of the thermally perfect gas, chemically non-equilibrium gas, and thermally non-equilibrium chemically frozen gas are applied for comparative studies. Results indicate that chemical non-equilibrium effects are more significant for the flow-field structure and parameters compared to thermal non-equilibrium effects. Meanwhile, vibrational relaxation and chemical reactions interact in the flow-field. The heat released from the chemical reactions in the flow-field of the thermochemical non-equilibrium gas model makes the thermal non-equilibrium effects weaker compared to the thermally non-equilibrium chemically frozen gas model; the chemical reactions in the thermochemical non-equilibrium gas model are more intense than in the chemically non-equilibrium gas model. Due to the slow relaxation of vibrational energy, the thermal non-equilibrium models predicted nozzle thrust lower than the thermal equilibrium models by approximately 1.11% to 1.33%; when considering the chemical reactions, the chemical non-equilibrium models predicted nozzle thrust higher than the chemical frozen models by approximately 7.30% to 7.54%. Hence, the structural design and performance study of the high Mach numbers scramjet nozzle must consider thermochemical non-equilibrium effects.

The attitude-tracking problem of hypersonic morphing vehicles (HMVs) is investigated in this research. After introducing variable-span wings, the optimal aerodynamic shape is available throughout the entire flight mission. However, the morphing wings cause significant changes in aerodynamic coefficients and mass distribution, challenging the attitude control. Therefore, a complete design procedure for the flight control system is proposed to address the issue. Firstly, the original model and the control-oriented model of HMVs are built. Secondly, in order to eliminate the influence caused by the multisource uncertainties, an adaptive fixed-time disturbance observer combined with fuzzy control theory is established. Thirdly, the fixed-time control method is developed to stabilise hypersonic morphing vehicles based on a multivariable sliding mode manifold. The control input can be obtained directly. Finally, the effectiveness of the proposed method is proved with the help of the Lyapunov theory and simulation results.