The recurring discussions regarding human flapping flight (if possible, by muscle-power!) have done more harm than good, because they distract attention from other more important subjects.

For example, the heed for the faster and larger flying creatures (such as large birds) and for high-speed aeroplanes to reduce speed when necessary, without corresponding loss of lift. All kinds of “ high lift devices ” have been invented for aeroplanes. All are based on the principle of making possible a steeper angle of attack or a higher wing-camber, thus giving greater lift without causing eddies that would lead to a break-away of the airflow on the upper surface.

A summary of the ways in which aviation impacts the environment is presented and the ratio of the energy liberated during a flight to the revenue work done (ETRW) is identified as a key indicator in the assessment of environmental impact. Using the ‘Breguet’ range equation, a number of theorems relating to ETRW are derived and discussed. This is followed by an approximate analysis to produce estimates for the ETRW of aircraft currently in service. It is found that the global fleet average value for ETRW is much higher than those estimated for existing individual aircraft. An explanation of the difference is presented, with the contributions from airline operations and air traffic management identified and quantified. Consideration is then given to the potential for future reduction in ETRW through advances in materials, alternative fuels, structures, aerodynamics and propulsion technologies and the likely benefits are quantified. The improvement in ETRW that could be achieved if this parameter was minimised in the design process with the current level of technology is also considered. Finally, the likelihood of performance improvements being introduced in the short, medium and long term is briefly discussed.

This paper describes a simulation technique that has been developed to quantify the unsteady forces and moments that are imposed onto a maritime helicopter by a ship’s airwake during a deck landing. An unsteady CFD-generated airwake, created using a CAD model of the ship, is integrated with a flight dynamics model of a helicopter. By holding the helicopter at a fixed position in the airwake it is possible to quantify the unsteady forces and moments imposed on the aircraft. The technique is therefore a software-based airwake dynamometer, and has been called the virtual AirDyn. As well as determining the mean loads, from consideration of the unsteady loads in the closed-loop pilot response frequency range of 0·2-2Hz it is also possible to quantify the magnitude of the unsteady disturbance in each axis. The loads are also indicators of the control activity the pilot would have to exert to maintain aircraft position and attitude. By placing the virtual AirDyn at different positions around the landing deck in different wind conditions, it is able to quantify the effect of the airwake on the mean and unsteady loads. The quantified loads can be explained by examining the CFD-generated flow field, and the geometric features on the ship’s superstructure that gave rise to them can be identified. The virtual AirDyn is therefore a tool that can be used to evaluate and inform ship design for maritime helicopter operations.

Potential function methods have been applied extensively to many terrestrial and space robot control problems. This paper extends these methods to multiple robot problems for on-orbit assembly. By considering a population of free-flying robotic vehicles in orbit, capable of both rotation and translation, the potential function method reduces complex assembly problems to sets of simple translation and rotational commands. This is achieved using a potential function incorporating both collision avoidance between the free-flyers and connection constraints in the assembled structure. The method also allows a subsumptive type control architecture with the flexibility to carry out many different assembly tasks simultaneously.

A forward facing spike attached to a hemispherical body significantly changes its flow field and influences aerodynamic drag and wall heat flux in a high speed flow. The dynamic pressure in the recirculation area is highly reduced and this leads to the decrease in the aerodynamic drag and heat load on the surface. Consequently, the geometry, that is, the length and shape of the spike, has to be simulated in order to obtain a large conical recirculation region in front of the blunt body to get beneficial drag reduction. It is, therefore, a potential candidate for aerodynamic drag reduction for a future high speed vehicle. Axisymmetric compressible laminar Navier-Stokes equations are solved using a finite volume discretisation in conjunction with a multistage Runge-Kutta time stepping scheme. The effect of the spike length and shape, and the spike nose configuration on the reduction of drag is numerically evaluated at Mach 6 at a zero angle-of-attack. The computed density contours agree well with the schlieren images. Additional modification to the tip of the spike to get different types of flow field such as the formation of a shock wave, separation area and reattachment point are examined. The spike geometries include the conical spike, the flat-disk spike and the hemispherical disk spike of different length to diameter ratios attached to the blunt body.

Results obtained by hot-wire measurements in a turbulent jet are presented for the cases where the jet issues into a co-flowing free stream and, for comparison, where the jet issues into quiescent air. The results obtained in the former show some agreement with other workers, but they are more extensive, including measurements of microscales and macroscales. The results also suggest that the presence of a free stream reduces the rate of decay of the longitudinal turbulent component of velocity, which has the effect of increasing the skewness of the eddy structure.

The rapid frequency sweep technique is the result of a considerable amount of research into the development of a transient method for the rapid measurement of structural frequency response, digital data analysis procedures being used to derive the required dynamic characteristics. Theoretical analyses, together with the results of experimental studies and experience gained during structural testing are presented here which illustrate practical considerations and limitations of the method; these range from discussion of exciter attachment, effects of transducers and choice of instrumentation to signal processing methods and related criteria. A simple guide to practical structural testing is given in the final section which outlines procedures to be followed during testing and signal acquisition.

Variable tail rotor speed is investigated as a method for reducing tail rotor power, and improving helicopter performance. A helicopter model able to predict the main rotor and tail rotor powers is presented, and the flight test data of the UH-60A helicopter is used for validation. The predictions of the main and tail rotor powers are generally in good agreement with flight tests, which justifies the use of the present method in analysing main and tail rotors. Reducing the main rotor speed can result in lower main rotor power at certain flight conditions. However, it increases the main rotor torque and the corresponding required tail rotor thrust to trim, which then decreases the yaw control margin of the tail rotor. In hover, the tail rotor may not be able to provide enough thrust to counter the main rotor torque, if it is slowed to follow the main rotor speed. The main rotor speed corresponding to the minimum main rotor power increases, if the change of tail rotor power in hover is considered. As a helicopter translates to cruise, the induced power decreases, and the profile power increases, with the profile power dominating the tail rotor. Reducing the tail rotor speed in cruise reduces the profile power to give a 37% reduction in total tail rotor power and a 1.4% reduction to total helicopter power. In high-speed flight, varying the tail rotor speed is ineffective for power reduction. The power reduction obtained by the variable tail rotor speed is reduced for increased helicopter weight.

Vortex breakdown phenomenon remains a challenging aspect of the flow over delta wings at high angle of attack. The observations of vortex breakdown and related interpretations for vortex tube experiments and delta wings are summarised in several review articles. The theoretical models of the phenomena are described in detail in References 1, 2, 3 and 7. These theoretical models are not complete enough to predict breakdown location over delta wings. The purpose of this note is to review several existing criteria and correlations for vortex breakdown over delta wings using a large database. The earliest criteria for delta wings was suggested by Lambourne and Bryer.

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.

Although the application of new, reduced aircraft separation minima can directly increase runway throughput, the impact thereof on the traffic flow of aircraft arriving at the destination airport has not been discussed yet. This paper proposes a data-driven and queue-based modeling approach and presents an analysis of the impact on the delay time of arriving aircraft in the airspace within a radius of 100 nautical miles around an airport. The parameters of our queuing model were estimated by analysing the data contained in the radar tracks and flight plans for flights that arrived at Tokyo International Airport during the 2 years of 2016 and 2017. The results clarified the best arrival strategy according to the distance from the arrival airport: The combination of airspace capacity control and reduction of the flight time and separation variance is the most powerful solution to mitigate delays experienced by arriving traffic while also allowing an increase in the amount of arrival traffic. The application of new wake vortex categories would enable us to increase the arrival traffic to 120%. In addition, the arrival delay time could be minimised by implementing the proposed arrival traffic strategies together with automation support for air traffic controllers.

An earlier paper described a method of calculating the turbulent boundary layer flow over the rear of an infinite swept wing. It made use of an entrainment equation and momentum integral equations in streamwise and cross-flow directions, together with several auxiliary assumptions. Here the method is adapted to the calculation of the turbulent boundary layer flow along the attachment line of an infinite swept wing. In this case the cross-flow momentum integral equation reduces to the identity 0 = 0 and must be replaced by its differentiated form. Two alternative approaches are also adopted and give very similar results, in good agreement with the limited experimental data available. It is found that results can be expressed as functions of a single parameter C*, which is evidently the criterion of similarity for attachment-line flows.

We discuss the status, trends and long-term ambitions of Computational Fluid Dynamics (CFD) when applied away from the design point or concept, and therefore in the historically weak areas of CFD. This includes both undesirable flight conditions, such as stall, and undesirable products of the flow, such as noise. All pose the great challenge of turbulence, and accuracy is as dependent on the ideas behind the turbulence treatment as it is on computing power. A measured shift from Reynolds-averaged representations to large-eddy simulations will take place. Empiricism, both turbulence and engineering related, will recede only step by step over many years. Yet, CFD will make full use of every increase in computer power of this century. Increasingly, CFD will compete with flight tests, not just with wind tunnels, and will be validated by flight tests. Integration with other disciplines will allow us to predict crucial phenomena such as flutter, sonic fatigue, and pilot-induced oscillations. A gap will remain at any time between the phenomena amenable to a ‘grand challenge’ calculation and those amenable to a fully CFD-based design process, because certification involves thousands of conditions. The highest demand currently is in community and cabin noise for which industrial-accuracy methods are non-existent. On the other hand, gratifying progress has occurred in the area of stall and spin.

The paper focuses on the modelling, simulation and control of an electrical miniature actuator integrated in the actuation mechanism of a new morphing wing application. The morphed wing is a portion of an existing regional aircraft wing, its interior consisting of spars, stringers, and ribs, and having a structural rigidity similar to the rigidity of a real aircraft. The upper surface of the wing is a flexible skin, made of composite materials, and optimised in order to fulfill the morphing wing project requirements. In addition, a controllable rigid aileron is attached on the wing. The established architecture of the actuation mechanism uses four similar miniature actuators fixed inside the wing and actuating directly the flexible upper surface of the wing. The actuator was designed in-house, as there is no actuator on the market that could fit directly inside our morphing wing model. It consists of a brushless direct current (BLDC) motor with a gearbox and a screw for pushing and pulling the flexible upper surface of the wing. The electrical motor and the screw are coupled through a gearing system. Before proceeding with the modelling, the actuator is tested experimentally (stand alone configuration) to ensure that the entire range of the requirements (rated or nominal torque, nominal current, nominal speed, static force, size) would be fulfilled. In order to validate the theoretical, simulation and standalone configuration experimental studies, a bench testing and a wind-tunnel testing of four similar actuators integrated on the real morphing wing model are performed.

The design of structures to resist jet noise fatigue demands a knowledge of a wide range of subjects from pure acoustics at one end to metal physics at the other. At the present time the various aspects of the problem are not sufficiently well known quantitatively for a purely theoretical design study to be made. Nevertheless a knowledge of the behaviour of typical forms of construction in noise environments can be used with a limited amount of theoretical work to indicate the most efficient types of structure. This approach to the problem is adopted in this lecture as it seems to be the most promising one available at the moment. It must be emphasised, however, that although some progress has been made in discovering the behaviour of a structure subjected to noise it is not possible to estimate the life of any component at the drawing board stage. Some prototype strain measurements and proof testing are therefore essential if one is to prove the integrity of the design.

The results of an experimental programme conducted in the University of Southampton's induced flow supersonic wind tunnel are presented. The pitch static force coefficient and centre of pressure data relate to a series of bodies with a blunt cone nose forebody with afterbodies comprising circular, square and triangular cross section. The blunt cone nose forebody has a forward located, body axis, aerodynamic spike system.

The data presented reveals the well known benefits of aerospikes with blunt nosed axisymmetric vehicles. This is then used as a base to explore the coupled flowfield effects on force coefficients due to spike deployment on bodies of non-axisymmetric cross section. The results identify how the body-only benefits of triangular cross sections are not degraded by the spike presence, with the main aero-spike benefit of drag reduction preserved throughout the angle of attack range investigated.

In a recent article, H. Werlé, has described how the free spiral vortices on delta wings with leading edge separation suddenly expand if the incidence is increased beyond a critical value. His description conforms to a great extent with the results, arrived at during an English investigation of the same phenomenon (called the vortex breakdown), but the interpretations of the observations, suggested by the two sources, are different. Against this background it is felt that some further comments and some pertinent high speed observations, may be of interest.

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

From an analysis of the work of Crocco and others, semi-empirical formulae are derived for the skin friction on a flat plate at zero incidence with a laminar boundary layer. These formulae are

for the general case of heat transfer, and

when there is no heat transfer.

The problem of heat transfer and the effect of radiation are discussed in the light of these formula;. The second formula is then utilised in the development of an approximate method for solving the momentum equation of the boundary layer on a cylinder without heat transfer. The method indicates that with increase of Mach number there is a marked forward movement of separation from a flat plate in the presence of a constant adverse velocity gradient.

Major results and findings of the numerical work package of the European high lift programme EUROLIFT are outlined. The main objective of these studies is to validate and test numerical methods for the prediction of high lift flows for transport aircraft configurations. The activities comprise the assessment of current CFD methods for 3D flows, evaluation of means for code improvement, and transition prediction. All aspects are especially devoted to high lift flow problems. A general capability to predict maximum lift on a simplified wing/fuselage high lift configuration is demonstrated by a variety of different numerical approaches. In general, major shortcomings are the reliability and the accurate simulation of large separation areas and the turn-around time to compute 3D lift polars. Advanced turbulence modelling and numerical solver features, such as the preconditioning technique, show a potential to overcome these deficiencies. Promising results with respect to transition prediction were obtained on a swept high lift wing using a database method. The results obtained in the numerical activities represent major ingredients on the way to a consistent numerical approach for the simulation of transport aircraft high lift configurations including all maximum lift determining effects.