This work presents a comparative study of design and development, in addition, of analyses of variable span morphing of the tapered wing (VSMTW) for the unmanned aerial vehicle (UAV). The proposed concept consists in the sliding of the inner section into the fixed part along the wing with varying the angle of the inner section inside the fixed part (parallel with the leading edge and the moving-wing axis is coincident to the fixed-wing axis) within two configurations. The wing design is based on a NACA 4412 aerofoil with the root chord of 0.675m and the tip chord of 0.367m for the fixed segment and 0.320m for the moving segment. Morphing wing analysis occurs at three selected locations that have been specified for extending and modifying span length by (25%, 50%, and 75%) of its original length to fulfill various flight mission requirements. The main objective of this paper is to compare the aerodynamic characteristics for several span lengths and sweep angles and to find their most efficient combinations. The wing is optimised for different velocities during all phases of flight (min speed, loiter, cruise, and max speed) which are 17, 34, 51, and 68m/s, respectively. The analyses are performed by computing forces (drag and lift) and moments at various altitudes, such as at the sea level, at 5,000 and 10,000ft. Two-dimensional aerodynamic analyses are carried out using XFLR5 code, and the ANSYS Fluent solver is used for investigating the flow field on the three-dimensional wing structure. It has been observed that a variable span morphing of tapered wing technology with a variable sweep angle can deliver up to 32.93% improved aerodynamic efficiency. This concept design can also be used for the aircraft roll motion technique instead of conventional control devices. Furthermore, the range flight mission increases up to 46.89% when the wing is placed at its full length compared to an original position. Finally, it has been concluded from this study that the wing design is more sensitive to the changing angle of the inner section and more efficient in terms of aerodynamic characteristics.

This paper dwells upon investigating the effect of aspect ratio (AR) variation on the aerodynamic performance of unconventional control surfaces called grid fins by virtue of a series of subsonic experiments on a simplified grid fin variant called the cascade fin. Wind tunnel tests were performed for different AR (variable span) grid fins. The same had been investigated for different gap-to-chord ratio (g/c) variants. Results demonstrated a tangible increase in the aerodynamic efficiency as well as stall angle reduction for higher AR. Moreover, higher AR leads to increased pitching moment, which emphasizes elevated hinge moment requirements. The study ensued the presence of higher deviation between the low AR fins, that is $AR<2$ compared to the pertinent deviations between the high AR fins, that is $AR\geq2$ . The effect associated with these variations was termed as span effect in this paper. It was established that, the deviations arising due to this phenomena were lesser for higher g/c and higher AR. The analysis of AR variation for different g/c presented a limiting value of AR reduction for stall performance enhancement. Thus, optimised selection of the g/c and AR values can lead to enhanced aerodynamic efficiency alongside an improved stalling characteristic.

This work presents an alternative methodology for monitoring flight performance during airline operations using the available inboard instrumentation system. This method tries to reduce the disadvantages of the traditional specific range monitoring technique where instrumentation noise and cruise stabilisation conditions affect the quality of the performance monitoring results. The proposed method consists of using an unscented Kalman filter for aircraft performance identification using Newton’s flight dynamic equations in the body X, Y and Z axis. The use of the filtering technique reduces the effect of instrumentation and process noise, enhancing the reliability of the performance results. Besides the better quality of the monitoring process, using the proposed technique, additional results that are not possible to predict with the specific range method are identified during the filtering process. An example of these possible filtered results that show the advantages of this proposed methodology are the aircraft fuel flow offsets, as predicted in the specific range method, but also other important aircraft performance parameters as the aircraft lift and drag coefficients (CL and CD), sideslip angle (β) and wind speeds, giving the operator a deeper understanding of its aircraft operational status and the possibility to link the operational monitoring results to aircraft maintenance scheduling. This work brings a cruise stabilisation example where the selected performance monitoring parameters such as fuel flow factors, lift and drag bias, winds and sideslip angle are identified using only the inboard instrumentation such as the GPS/inertial sensors, a calibrated anemometric system and the angle-of-attack vanes relating each flight condition to a specific aircraft performance monitoring result. The results show that the proposed method captures the performance parameters by the use of the Kalman filter without the need of a strict stabilisation phase as it is recommended in the traditional specific range method, giving operators better flexibility when analysing and monitoring fleet performance.

The background and rationale for the research reported in this paper is detailed in the companion paper subtitled Part A: Modelling the on-deck helicopter Reserve of Stability (ROS).

This paper (Part B) derives the fundamental principles underpinning the safe operational envelope and the definition of Motion Severity and Wind Severity Indices (MSI/WSI), and then presents the probabilistic modelling methodology developed for calculating safe limits for these parameters (MSI/WSI limits curves). The introduction of the Relative Wind Direction (RWD) as an additional limiting parameter is also discussed, together with an explanation of how RWD limits have been derived and how they are linked with the MSI/WSI limits calculation.

The MSI/WSI limits modelling methodology presented in this paper builds on the analytical helicopter stability modelling discussed in Part A. It uses a probabilistic approach to manage the variability in helicopter operations across the UK and to deal with the uncertainty of predicting helideck motion and wind conditions for the period the helicopter is landed on the helideck based on measurements taken prior to landing. This has been developed with the assistance of industry experts and has been subjected to consultation with stakeholders in the UK with the aim of identifying the right balance between enhancing safety and preserving operability.

A new Helideck Monitoring System (HMS) standard has been developed which incorporates the MSI/WSI/RWD functionality. The standard was published by the Helideck Certification Agency in April 2018 with a compliance date of 31 March 2021. Operations to moving decks not equipped with HMS meeting the new standard will be restricted to stable helideck conditions.

This article presents a potential analytic redundancy approach to detect faults in the air data sensor of an aircraft. In modern aircraft, fault detection of air data sensors is performed using a complex voting mechanism, which requires the availability of redundant air data sensor in all situations. However, to continuously monitor operation and performance of these sensors, the analytic redundancy-based air data estimation and fault detection is highly preferred than estimation with air data probe measurements. The proposed algorithm uses the kinematics of aircraft to estimate air data and detect air data sensor fault. In this paper, a simple mathematical model is developed, which does not consider the forces and moments acting on aircraft and uses measurements only from the Inertial Measurement Unit (IMU) and Navigation System Data (NSD). In order to implement this approach, the Iterated Optimal Extended Kalman Filter (IOEKF) is developed to estimate air data, which provides an accurate and stable estimation. With the estimated states, the physical air data sensor measurements are compared and the residual is calculated to track each sensor performance and to detect the occurrence of a fault. The key advantage of this approach is that it does not require complex dynamic equations and is free from system uncertainties. The proposed algorithm is simulated in MATLAB software using flight simulator flight data and validated using the real-time flight data of Cessna Citation II transport aircraft.

Extensive research in recent years has focused on improving the current loss-of-control prevention systems and developing new strategies for safe path planning of the impaired aircraft. Success in developing such systems requires a comprehensive perception of the influence of damage on the aircraft’s dynamic behaviour and performance, and the effect of various failure degrees on the flight envelope confinement and the remaining safe maneuvers. This paper comprehensively describes the effects of lateral control surface failure on the NASA Generic Transport Model (GTM) flight envelope, defined by a set of attainable steady-state maneuvers herein referred to as trim points. The study utilises a large database of high-fidelity maneuvering flight envelopes computed for the unimpaired case and wide ranges of the aileron and rudder failure cases at different flight conditions. Flight envelope boundary is rigorously investigated, and the key parameters confining the trim points at different boundary sections are identified. Trend analyses of the impaired flight envelopes and the corresponding limiting factors demonstrate the effect of various failure degrees on the remaining feasible trim points. Results can be employed in emergency path planning with potential uses in the development of aircraft resilient control and upset recovery systems.

In recent years, the aviation industry has taken a leading role in the integration of composite structures to develop lighter and more fuel efficient aircraft. Among the leading concepts to achieve this goal is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The focus of most PRSEUS studies has been on developing an hybrid wing body structure, with only a few discussing the application of PRSEUS to a tube-wing fuselage structure. Additionally, the majority of investigations for PRSEUS have focused on experimental validation of anticipated benefits rather than developing a methodology to capture the behavior of stitched structure analytically. This paper presents an overview of a numerical methodology capable of accurately describing PRSEUS’ construction and how it may be implemented in a barrel fuselage platform resorting to high-fidelity mesoscale modeling techniques. The methodology benefits from fresh user defined strategies developed in a commercially available finite element analysis environment. It further proposes a new approach for improving the ability to predict deformation in stitched composites, allowing for a better understanding of the intricate behavior and subtleties of stitched aerospace structures.

This paper deals with modelling the performance of an air transport network operated by existing subsonic and the prospective supersonic commercial aircraft. Analytical models of indicators of the infrastructural, technical/technological, operational, economic, environmental, and social performance of the network relevant for the main actors/stakeholders involved are developed. The models are applied to the given long-haul air route network exclusively operated by subsonic and supersonic aircraft according to the specified “what-if” scenarios.

The results from application of the models indicate that supersonic flights powered by LH2 (Liquid Hydrogen) could be more feasible than their subsonic counterparts powered by Jet A fuel, in terms of about three times higher technical productivity, 46% smaller size of the required fleet given the frequency of a single flight per day, 20% lower sum of the aircraft/airline operational, air passenger time, and considered external costs, up to two times higher overall social-economic feasibility, and 94% greater savings in contribution to global warming and climate change. These flights could be less feasible in terms of about 70-85% higher aircraft/airline operational costs, 70% and 19% higher fuel consumption and emissions of Green House Gases, respectively, and 6-13% higher noise compared to the specified acceptable levels.

The objective of this study is to identify, evaluate, and provide recommendations towards the realisation of near-term hypersonic flight hardware through the consideration of carrier vehicle constraints. The current rush of available funds for hypersonic research cannot cause a program to ignore growth potential for future missions. The prior NB-52 carrier vehicles, famous for the X-15 and X-43A missions, are retired. Next generation hypersonic demonstrator requirements will necessitate a substitution of carrier vehicle capability. Flight vehicle configuration, technology requirements, and recommendations are arrived at by constructing and evaluating a hypersonic technology demonstrator design matrix. This multi-disciplinary parametric sizing investigation of hypersonic vehicle demonstrators focuses on the evaluation of the combined carrier platform, booster, and hypersonic cruiser solution space topography. Promising baseline configurations are evaluated against operational requirements by trading fuel type, endurance cruise time, and payload weight. The multi-disciplinary study results are constrained with carrier payload mass and geometry limitations. The multi-disciplinary results provide physical insights into near-term hypersonic demonstrator payload and cruise time requirements that will stretch the capability of existing carrier aircraft. Any growth in hypersonic research aircraft size or capability will require new carrier vehicle investments.

In this work, consideration is given to a novel concept for aerofoil lift enhancement and delaying flow separation. Here, lift enhancement is attained by preventing the growth of the boundary layer through the elimination of the zero-slip condition between the wing surface and the air stream. The concept would simulate all the effects of a moving wall, leading to the appearance of a slip velocity at the gas–fluid interface, including the injection of momentum into the air boundary layer, but with one exception: here there is no moving wall but instead a ferrofluid thin film pumped parallel and attached to the wall by a magnetic field. Utilising a simplified physical model for the velocity profile of the ferrofluid film and based on ferrohydrodynamic stability considerations, an analytical expression for the interfacial velocity is derived. Finally, from the available experimental data on moving walls, the expected lift and angle-of-attack enhancement are found as well as the weight penalty per unit surface area of the wing is estimated. Additional research and development is required to explore the possibilities of using ferrofluid thin films.

Large-eddy simulations (LES) have been employed to investigate the far-field four-vortex wake vortex evolution over 10min behind an aircraft formation. In formation flight scenarios, the wake vortex behaviour was found to be much more complex, chaotic and also diverse than in the classical single aircraft case, depending very sensitively on the formation geometry, i.e. the lateral and vertical offset of the two involved aircraft. Even though the case-by-case variability of the wake vortex behaviour across the various formation flight scenarios is large, the final plume dimensions after vortex dissolution are in general substantially different from those of single aircraft scenarios. The plumes are around 170 to 250m deep and 400m to 680m broad, whereas a single A350/B777 aircraft would produce a 480m deep and 330m broad plume. Formation flight plumes are thus not as deep, yet they are broader, as the vortices do not only propagate vertically but also in span-wise direction. Two different LES models have been employed independently and show consistent results suggesting the robustness of the findings. Notably, $CO_{2}$ emissions are only one contribution to the aviation climate impact among several others like contrails and emission of water vapour and nitrogen oxides, which would be all affected by the implementation of formation flight. Thus, we also highlight the differences in ice microphysical and geometrical properties of young formation flight contrails relative to the classical single aircraft case.

Flight-deck Interval Management (FIM) is a modern airborne self-spacing technology that improves arrival route throughput and runway utilisation and increases hourly arrival capacity by up to four aircraft per hour and per runway, compared to conventional air traffic controller guided arrivals. The National Aeronautics and Space Administration (NASA) has been the leader in FIM research and formulated a logic that was put to an actual flight test in 2017. Despite the overall success of the project, operational deficiencies concerning the number of speed commands, which led to several recommendations for future research before operational implementation, were discovered. In this study, a new logic that implements a two-stage rule-based selection algorithm was developed to overcome those deficiencies. The proposed logic was compared to NASA’s logic on an arrival in Tokyo International Airport with multiple induced error patterns. The results indicate that the new logic significantly decreases the number of speed commands with only minor aggravations in spacing performance. The results that highlight the strengths and weaknesses of both concepts are discussed, and an outlook on and ideas for future research on FIM and the proposed logic are presented.

In the context of future human spaceflight exploration missions, Rendezvous and Docking (RVD) activities are critical for the assembly and maintenance of cislunar structures. The scope of this research is to investigate the specifics of orbits of interest for RVD in the cislunar realm and to propose novel strategies to safely perform these kinds of operations. This paper focuses on far rendezvous approaches and passively safe drift trajectories in the Ephemeris model. The goal is to exhibit phasing orbit requirements to ensure a safe far approach. Ephemeris representations of Near Rectilinear Halo Orbits (NRHOs) were derived using multiple-shooting and adaptive receding-horizon targeting algorithms. Simulations showed significant drift and overlapping properties for phasing and target orbits of interest, motivating the search for safe natural drift trajectories and using impact prediction strategies.

This paper presents a review of the various aeronautical air conditioning systems that are currently available and discusses possible system configurations in the context of the aeronautical environmental control systems. Descriptions of the standard vapor compression cycle and air cycles are provided. The latter includes, simple-cycle, bootstrap-cycle, simple-bootstrap cycle (3-wheel) and condensing cycle (4-wheel). Water separation and air recirculation systems are also explored. A comparison between vapor compression cycles and air cycles is provided, as well as a comparison between different air cycles. Air cycle units are far less efficient than vapor compression cycle units, but they are lighter and more reliable for an equivalent cooling capacity. Details regarding the aircraft conceptual design phase along with general criteria for the selection of an air conditioning system are provided. Additionally, industry trends and technological advances are examined. Conclusions are compiled to guide the systems engineer in the search for the most appropriate design for a particular application.

Turbo-electric propulsive fuselage aircraft featuring Boundary-Layer Ingestion (BLI) are considered promising candidates to achieve the emissions reduction targets set for aviation. This paper presents an analytical method capable of estimating the BLI benefit at aircraft level, enabling a quick exploration of the propulsive fuselage design space. The design space exploration showed that the assumptions regarding the underwing turbofans and BLI fan mass estimation can have an important impact on the final fuel burn estimation. The same applies to the total efficiency assumed for the electric transmission, the range of the aircraft mission, and the propulsive efficiency of the engines used as benchmark. The regional jet and short- to medium-range aircraft classes seem to be the most promising as the ingested drag and power saving are among the largest, with long-range aircraft being just behind. The future introduction of advanced technologies, which target the reduction of vortex and wave dissipation at aircraft level, could increase the potential benefit of propulsive fuselage BLI. On the other hand, the potential benefit would be decreased if more efficient and lighter ultra high bypass ratio engines were used as benchmark.

The gas temperature of the supersonic heat airflow simulated test system is mainly determined by the fuel and air flow rates which enter the system combustor. In order to realise a high-quality control of gas temperature, in addition to maintaining the optimum ratio of fuel and air flow rates, the dynamic characteristics of them in the combustion process are also required to be synchronised. Aiming at the coordinated control problem of fuel and air flow rates, the mathematical models of fuel and air supply subsystems are established, and the characteristics of the systems are analysed. According to the characteristics of the systems and the requirements of coordinated control, a fuzzy-PI cross-coupling coordinated control strategy based on neural sliding mode predictive control is proposed. On this basis, the proposed control algorithm is simulated and experimentally studied. The results show that the proposed control algorithm has good control performance. It cannot only realise the accurate control of fuel flow rate and air flow rate, but also realise the coordinated control of the two.

A novel second-order sliding-mode-based impact angle and autopilot lag guidance law for engaging manoeuvering targets with unknown acceleration is presented in this study. A backstepping technique is applied to the design of the sliding surface. The proposed guidance law is based on a new sliding surface. It exhibits the advantage of ensuring that the sliding surface and its derivative will converge to zero in finite time while guaranteeing that the sliding surface will not cross zero until the ultimate time. The method effectively eliminates the undesired chattering of the sliding surface. To compensate for the uncertainty caused by target manoeuvering, a new observer is developed to estimate target manoeuvering. The convergence of the system is proven through a Lyapunov function and finite time convergence theory. Lastly, mathematical simulations results show that the proposed guidance law can achieve precise interception with a wide range of impact angles, thereby verifying the effectiveness of the guidance law.

Based on the missiles with adjustable thrust, a distributed synergetic guidance law for multiple missiles with angle-of-attack constraint is designed to achieve a cooperative attack by multiple missiles on a moving target. Divide the guidance law into two parts: the line of sight (LOS) direction and the normal direction of LOS. The guidance law is designed in LOS direction based on the multi-agent system cooperative control theory and super-twisting algorithm, which can control missiles’ time-to-go converging in finite time. In the normal direction of LOS, the other guidance law is designed to control missiles hitting the target with impact angle constraint based on zeroing LOS rate thought and finite-time sliding mode control theory. The two non-homogeneous disturbance observers are designed to estimate the target manoeuvring information in the two parts of the guidance law. The simulation results show that the guidance law designed by this paper can complete the cooperative attack task while suppressing the chattering phenomenon effectively.

Air-breathing propulsion has the potential to decrease the cost per kilogram for access-to-space, while increasing the flexibility of available low earth orbits. However, to meet the performance requirements, fuel-air mixing inside of scramjet engines and thermal management still need to be improved.

An option to address these issues is to use intrinsically generated vortices from scramjet inlets to enhance fuel-air mixing further downstream, leading to shorter, less internal drag generating, and thus more efficient engines. Previous works have studied this vortex-injection interaction numerically, but validation was impractical due to lack of published experimental data. This paper extends upon these previous works by providing experimental data for a canonical geometry, obtained in the T4 Stalker Tube at Mach 8 flight conditions, and assesses the accuracy of numerical methodologies such as RANS CFD to predict the vortex-injection interaction.

Focus is placed on understanding the ability of the numerical methodology to replicate the most important aspects of the vortex-injection interaction. Results show overall good agreement between the numerical and experimental results, as all major features are captured. However, limitations are encountered, especially due to a localised region of over predicted heat flux.

This work investigates the propeller’s influence on the stability of High Altitude Long Endurance aircraft, incorporating all resultant loads at the propeller hub, propeller slipstream, and gyroscopic loads. Such effects are usually neglected in the aeroelastic simulation of HALE aircraft. For that goal, a previously developed framework, which couples a geometrically nonlinear structural solver with an Unsteady Vortex Lattice method (uVLM) for lifting surfaces and a Viscous Vortex Particle (VVP) method for propeller slipstream, was employed to generate time-data series. Also, a method, based on a combination of Proper Orthogonal Decomposition and system identification, to extract dynamic information (frequencies, damping, and modes) of the aircraft from a time-series signal is proposed and successfully tested for a purely structural case, for which reference data is available. The method is then applied to investigate the stability of aeroelastic cases. The results demonstrate that the presence of propellers can influence the aeroelastic stability of a Very Flexible Aircraft.