Stable separation is a crucial condition that must be met in order for combined aircraft to successfully engage in cooperative flight. In order to achieve the desired fast and controlled separation, this paper proposes a novel design for a torque-driven compliant separation mechanism. By taking into account the compliance characteristics of a sinusoidal acceleration function curve, a mechanical model for the separation mechanism is developed. By utilising the Coulomb friction law, an accurate determination of the aerodynamic load distribution under various conditions is achieved. Subsequently, the relationship between the unlocking moment and the aerodynamic load is derived based on these findings. Through the utilisation of the finite element method, a model of the separation mechanism is generated. To ensure the safety and reliability of the compliant separation mechanism, the mechanical properties of the structural materials are thoroughly analysed under the maximum aerodynamic load. Subsequently, the separation mechanism structure is constructed and subjected to testing in order to showcase the compliance characteristics. In addition, this paper conducts a simulation to analyse the impact of flight speed and angle-of-attack on the separation process. By doing so, the optimal conditions for separation are determined. The methods and findings presented in this study have the potential to contribute valuable insights to the design of combined aircraft.

Topology optimization has been identified as a powerful tool to improve aircraft structures for many years. Yet, innovative layouts have not been successfully implemented in commercial aircraft for several reasons. One reason identified by our research group is the lack of design constraints during topology optimization, such as buckling stability, which yields complex solutions that are not easily manufacturable. Second, the complexity of the resulting layouts makes integration with other systems highly challenging. With respect to these challenges, we propose a new heuristic layout optimization process: complexity-driven layout exploration for aircraft structures (CD-LEAS). The new process addresses the challenges of complexity and nonlinear constraints, such as buckling, in aircraft structure layout optimization. The novelty of CD-LEAS comes from the integration of a relative complexity metric as a driver to navigate the design space efficiently. Two case studies of commonly used stiffened panels are carried out to showcase the performance of the process. The results show that using complexity to navigate an explicit design space allows our process to quickly output a family of simple, light, stiff and buckling-resistant layouts.

The German research Cluster of Excellence SE2A (Sustainable and Energy Efficient Aviation) is investigating different technologies to be implemented in the following decades, to achieve more efficient air transportation. This paper studies the Hybrid Laminar Flow Control (HLFC) using boundary layer suction for drag reduction, combined with other technologies for load and structural weight reduction and a novel full-electric propulsion system. A multidisciplinary design optimisation framework is presented, enabling physics-based analysis and optimisation of a fully electric aircraft wing equipped with HLFC technologies and load alleviation, and new structures and materials. The main focus is on simulation and optimisation of the boundary layer suction and its influence on wing design and optimisation. A quasi three-dimensional aerodynamic analysis is used for drag estimation of the wing. The tool executes the aerofoil analysis using XFOILSUC, which provides accurate drag estimation through boundary layer suction. The optimisation is based on a genetic algorithm for maximum take-off weight (MTOW) minimisation. The optimisation results show that the active flow control applied on the optimised geometry results in more than 45% reduction in aircraft drag coefficient, compared to the same geometry without HLFC technology. The power absorbed for the HLFC suction system implies a battery mass variation lower than 2%, considering the designed range as top-level requirement (TLR).

Design success mostly depends on understanding and realising product requirements, which are nested with different levels of stakeholders and specific market demands. Emerging markets are volatile and non-linear for understanding user-specific needs, particularly related to aircraft design. We attempt herein to elicit specific requirements through a proposed requirements elicitation process based on functional, physical and behavioural viewpoints. An evaluation of those critical requirements is also performed through Quality Function Deployment (QFD) to reveal the fitness of the customer requirements (raw requirements) in relation to the technical measures. After selecting all projected potential users, a Systems Modelling Language (SysML) use-case diagram is also summarised to illustrate the user’s relation to the system functions. Some raw requirements are verified through constraint analysis and cost examination as an exemplary approach to present the transformation from raw to final requirements. The data are collected and analysed to construct the aircraft’s key specifications. Most importantly, how to identify good requirements that define the needs, attainability, clarity and verifiability is demonstrated.

In addition, it is demonstrated through this study that there is a need of a specific type of general aviation aircraft that may fulfil the unique demands of emerging economies by satisfying their socio-economic conditions.

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.

The range of small, electrically powered UAVs is still limited by the mass specific energy of batteries. This paper investigates the idea that, in cases where multiple aircraft must transit to the same location, savings in mass or an extension of achievable range are possible when they join wingtip-to-wingtip. The viability of joined flight is investigated by quantifying the relative magnitude of savings resulting from increased aerodynamic efficiency and that of penalties due to the increased structural and component weights. Through a parametric analysis the level of savings achievable is found to be greatly dependent on the proportion of the flight spent in a joined configuration and aircraft design parameters such as wing loading, aspect ratio and the added weight of the joining mechanism. A custom, multidisciplinary UAV sizing algorithm is presented and utilised to design several sample aircraft, featuring two different joining mechanism architectures. The results verify the findings of the parametric study and indicate that mass savings are possible only for moderate to low aspect ratios, with semi-permanent magnetic joining mechanism performing better than rigid structural ones, even when the joined fight segment accounts for only 30% of the total airborne time.

One of the challenges of modern engineering design is the amount of data that designers must keep track while performing system analysis and synthesis. This task is particularly important in the design process of complex systems such as novel aerospace systems where Modeling and Simulation play an essential role. The Agile philosophy stems from the field of Software Engineering and describes an approach to development in which requirements and solutions gradually develop through collaboration between self-organising cross-functional teams and end users. Agile Model-Based System Engineering (AMBSE) is the application of the Agile philosophy to Model-Based System Engineering. In this paper, AMBSE is accomplished through the application of the Object-Oriented System Engineering Method (OOSEM). OOSEM employs a top-down scenario-driven process that adopts System Modeling Language (SysML) and leverages the object-oriented paradigm to support the analysis, specification, design, and verification of systems. AMBSE assisted by mathematical modelling and safety assessment techniques is applied to the first design iterations of the main aircraft systems, allowing a comprehensive design exploration. The flight control system was chosen to illustrate the procedure in detail, emphasising the synthesis of a six-degrees-of-freedom model augmented by dynamic inversion control for a hypothetical supersonic transport aircraft satisfying class II MIL-F-8785C handling qualities. It is concluded that AMBSE presents promising properties to support future aircraft development within the current regulatory framework for aircraft design, while enabling a smooth transition from conceptual to preliminary design.

This investigation surveyed the potential and established outcomes for future 19-passenger fixed-wing commuter transport aircraft concepts employing battery-based Voltaic-Joule/Brayton motive power systems with no additional electrical energy drawn from generators mechanically coupled to thermal engines. The morphological approach was that of a tri-prop (two on-wing podded turbo-props and one aft-fuselage mounted electric motor configured as a pusher-on-pylon installation). A Battery System-level Gravimetric Specific Energy (referred to as “battery energy density”) of at least 500 Wh/kg yielded 39%, 25% and 10% block fuel reductions for 150-nm (Design Service Goal), 430-nm (85th percentile) and 700-nm (maximum range) stage lengths, respectively. All quoted comparisons are against a suitably projected turbo-prop only year-2030 aircraft. In contrast to the reference Beech 1900D, block fuel reductions of up to 44-49% were predicted, which could facilitate a significantly lower deficit in relation to the Advisory Council for Aviation Research and Innovation in Europe (ACARE) Strategic Research and Innovation Agenda (STRIA) 55% target for year 2030. This investigation also indicated that, in the future, suitably flexible hybrid-electric architectures could be fashioned allowing possibility for the aircraft to complete any required city-pair operations (within the legitimate payload-range working capacity) irrespective of exchangeable batteries being available at a given station. Finally, it was also established, assuming such a tri-prop morphology, Normal conducting machines delivering maximum shaft power output of 1.1 MW would be required.

Aviation brings great commercial and social benefits and, as the global economy develops, the demand for air transport is expected to grow. However, aviation also contributes to climate change and there is increasing international pressure to limit mean global temperature rise. Therefore, the future success of aviation is likely to depend upon the industry's ability to hold its environmental impact within politically acceptable limits. This paper describes aviation's interaction with climate and sets out, in broad terms, the challenge facing the industry. The prospects for mitigating the adverse effects by advanced technology working through aircraft design and manufacture are assessed and some potential restrictions imposed by airport infrastructure are identified. Some consideration is also given to the practicalities imposed by airline economics and the likely impact of the recent ICAO regulations, plus the proposed global-based market measures scheme for aviation carbon offset. It is argued that the environmental problem is not just about carbon dioxide emissions and that aircraft technology improvement alone is unlikely to provide a complete solution. The observation that, in current operations, the total fuel used is almost twice the minimum required suggests a need to recognise a broader ownership of the problem and broader responsibility for the solution. However, improvements in the overall system efficiency will probably need to be driven by additional regulation and the imposition of other financial measures.

The overall conclusion is that the actions currently being taken and currently being proposed are probably not sufficient in themselves to meet the challenge of climate change. However, it appears that there is more that can be done and, provided that action is taken soon, there are some grounds for optimism that aviation will still be able to meet the needs of society in the 21st century.

A recent consideration in aircraft design is the use of folding wing-tips with the aim of enabling higher aspect ratio aircraft with less induced drag while also meeting airport gate limitations. This study investigates the effect of exploiting folding wing-tips in flight as a device to reduce both static and dynamic loads. A representative civil jet aircraft aeroelastic model was used to explore the effect of introducing a wing-tip device, connected to the wings with an elastic hinge, on the load behaviour. For the dynamic cases, vertical discrete gusts and continuous turbulence were considered. The effects of hinge orientation, stiffness, damping and wing-tip weight on the static and dynamic response were investigated. It was found that significant reductions in both the static and dynamic loads were possible. For the case considered, a 25% increase in span using folding wing-tips resulted in almost no increase in loads.

This paper provides a set of closed form solutions for the lift and drag of wings flying in ground effect both with and without end plates. The developed theories are based on observations of several independent sources of controlled model tests over ground planes and over water and on previous theories prepared by researchers from the original work by Prandtl and Wieselsberger to the present day. The theories developed cover wings of varying aspect ratio, thickness to chord ratios and angle of attack. The results for wings with end plates include the effect of ground (or surface) clearance height, end plate depth and air gap depths beneath the wings or end plates. Good agreement is found between the developed theory and test.

Commercial jets usually have relatively low-aspect-ratio wings, in spite of the associated benefits of increasing the wing aspect-ratio, such as higher lift-to-drag ratios and ranges. This is partially explained by the fact that the wing becomes more flexible by increasing the aspect-ratio that results in higher deflections which can cause aeroelastic instability problems such as flutter. An aeroelastic computational framework capable of evaluating the effects of geometric non-linearities on the aeroelastic performance of high-aspect-ratio wings has been developed and validated using numerical and experimental data. In this work, the aeroelastic performance of a base wing model with 20 m span and 1 m chord is analysed and the effect of changing the wing chord or the taper-ratio is determined. The non-linear static aeroelastic equilibrium solutions are compared in terms of drag polar, root bending moment and natural frequencies, and the change in the flutter speed boundary is assessed as a function of aspect-ratio using a time-marching approach.

This design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral and the radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending and torsion) and the high-g manoeuvre pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1.0% in hover, 10.0% at μ = 0.30 and 17.0% at μ = 0.40, relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0.30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre. Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement.

This paper presents a multi-disciplinary analysis methodology for a box-wing aircraft configuration optimised for a given mission scenario. This conceptual design methodology and associated toolchain combines well-established vortex lattice analysis and a newly developed structural analysis tool called WingMASS, allowing the design space to be explored from a combined aerodynamics and structural design perspective. For a given mission scenario, the method optimises a box-wing configuration and compares it with an equivalent conventional configuration. This study shows that, for a given mission, a box-wing configuration can lead to a fuel burn reduction of up to 5% by optimising aspect ratio, horizontal and vertical wing separation.

This paper presents a study of the W3-Sokol main rotor equipped with Gurney flaps. The effect of the active Gurney is tested at low and high forward flight speeds to draw conclusions about the potential enhancement of the rotorcraft performance for the whole flight envelope. The effect of the flap on the trimming and handling of a full helicopter is also investigated. Fluid and structure dynamics were coupled in all cases, and the rotor was trimmed at different thrust coefficients. The Gurney proved to be efficient at medium to high advance ratios, where the power requirements of the rotor were decreased by up to 3.3%. However, the 1/rev actuation of the flap might be an issue for the trimming and handling of the helicopter. The current study builds on the idea that any active mechanism operating on a rotor could alter the dynamics and the handling of the helicopter. A closed loop actuation of the Gurney flap was put forward based on a pressure divergence criterion, and it led to further enhancement of the aerodynamic performance. Next, a generic light utility helicopter was built using 2D aerodynamics of the main aerofoil section of the W3 Sokol blade along with a robust controller, and the response of the rotorcraft to control inputs was tested. This analysis proved that the 1/Rev actuation of the Gurney did not alter the handling qualities of the helicopter, and as a result, it can be implemented as a flow control mechanism for aerodynamic enhancement and retreating blade stall alleviation.

The cycloidal propeller for a Micro-Aerial Vehicle (MAV)-scale cyclogyro in hover was studied using a 2D Reynolds-averaged Navier-Stokes equations solver. The effects of the blade dynamic stall, parallel Blade Vortex Interaction (BVI), inflow variation and flow curvature were discussed, based on the results of numerical simulation. The results from the 2D Computational Fluid Dynamics simulation indicated that the blade of the cycloidal rotor is actually performing a pitching oscillation, if observed in a moving reference frame. The dynamic stall vortices shed from the upstream blade cause intense parallel BVI on the downstream blade. The interaction will induce upwash and downwash on the downstream blade. This changes the effective reduced frequency and actually delays the stall of the blade, which is beneficial to the thrust generation. There is also strong downwash in the rotor cage and it changes the inflow velocity experienced by the blade. The downwash and flow curvature can either be beneficial or harmful to the thrust generation. The combined effects of dynamic stall, parallel BVI, inflow variation and flow curvature cause large aerodynamic force peaks and ensure the cycloidal rotors work at very low rotation speeds with high thrust. This guarantees that the cycloidal rotors possess at least the same level of hover efficiency as screw propellers.

The International Civil Aviation Organization (ICAO) Circular Cir 337 is the first step towards ICAO establishing an Aeroplane CO2 Emissions Standard to form part of Annex 16, Volume III to the Chicago Convention. It describes itself as ‘a work in progress’. This paper reviews Cir 337 against the background of flight physics, the published literature on aircraft fuel burn and CO2 emissions and the current practices of the aircraft and engine manufacturers and the airline operators. We have taken, as our starting point, the aim of ICAO to reduce the fuel used per revenue tonne-kilometre performed and argue that the Breguet range equation, which captures all the relevant flight physics, should be the basis of the metric system underpinning the standard. Our overall conclusion is that Cir 337 provides an excellent basis for the initial regulation of aviation's CO2 emissions and, further in the future, for developing measures to increase the fuel efficiency of the operational side of civil aviation. Our main criticism of the circular in its current form is that it does not address the ICAO goal of reducing fuel used per revenue tonne-kilometre performed and makes no reference to payload. This defect could be eliminated simply by omission of the exponent 0.24 of the Reference Geometric Factor (RGF) in the formula for the metric given in Chapter 2 (paragraph 2.2) of the circular. Retaining the RGF to the power unity in the metric and multiplying it by an appropriate value of the effective floor loading would convert it to what the 37th Assembly of ICAO called for – a statement of fuel used per revenue tonne-kilometre performed. Finally, correlating the amended metric against design range, as determined from the measured specific air range and the key certificated masses, provides a sound scientific basis for an initial regulation to cap passenger aircraft emissions.

The aviation industry is dominated by the domain of heavier-than-air, fixed-wing, subsonic flight, and central to any design in this domain is the wing itself. One of the earliest debates in aviation still centres around the usefulness of the wing volume. On the one hand it is held that the wing, as an inevitable necessity, should provide the volume also for the payload. On the other, it is argued that more efficient wings do not even have sufficient volume for the entire wing structure. This work proposes precise definitions of the Wing Density and the Inflation Factor, two parameters that can quantitatively reflect the economic and technological trends in aviation. The wing volume of a hypothetical Ideal Wing is derived from the Operational Parameters of any given Flight Objective and compared to the volume requirement of that flight objective. We conclude that the dominant aircraft configuration of the future is likely to remain within the same family of the current dominant configuration, in conflict with some older predictions.

The article includes study on the influence of a LEX (leading-edge extension) and wing tip plates on aerodynamic characteristics of the Rocket Plane in tailless configuration for manned suborbital flights. The research was conducted only for the low speed regime in the wind tunnel at the Warsaw University of Technology. Tests were carried out for a few configurations of the Rocket Plane. Moreover, two shapes of LEX were investigated. Analysis of results was focused on the aerodynamic synergism effect only. The influence of LEX and the wing tip plates of the Rocket Plane on the aerodynamic coefficients was identified.