Two numerical methods are presented for the computation of steady and unsteady Euler flows. These are applied to steady and unsteady flows about the NACA 0012 aerofoil, using structured grids generated by the transfinite interpolation technique. An explicit central-difference scheme is produced based on the cell-vertex method of Ni modified by Hall. The method is second-order accurate in time and space, and with flow quantities stored at boundaries the boundary conditions are simple to apply. This is a definite advantage over the cell-centred approach of Jameson, where extrapolation of the flow quantities is required at the boundaries, making unsteady boundary conditions difficult to apply. An explicit upwind-biased scheme is also produced, based on the flux-vector splitting of van Leer. The method adopts a three stage Runge-Kutta time-stepping scheme and a high-order spatial discretisation which is formally third-order accurate for one-dimensional calculations. The upwind scheme is shown to be slightly more accurate than the central-difference scheme for steady aerofoil flows, but it is not clear which is the more accurate for unsteady aerofoil flows. However, the central-difference scheme requires less than half the CPU time of the upwind-difference scheme, and hence is attractive, especially when considering three-dimensional flows. The transfinite interpolation technique is ideal for generating moving structured grids due to its simplicity, and grid speeds are available algebraically by the same interpolation as grid points. The method is also ideal for use in a multi-block approach.

The flow separation on a swept-back wing with 44 degrees leading edge sweep at 18 degrees incidence has been investigated by means of detailed pressure distribution measurements over the leading edge, boundary layer flow determination with liquid film technique, and yawmeter traverses. A wool-tuft grid was also used, and a spin detector was developed to search for regions of vorticity. These tests established that the flow separates on the leading edge; over the inboard sections it re-attaches behind a “ short” separation bubble, while outboard it only re-attaches near the trailing edge, thus forming a “ long ” separation bubble, or else fails to attach. The separated flow in what has been commonly called the tip stall does in fact take the form of a “ ram's horn “ vortex with the origin, or “ tip,” located at the junction of the two bubbles on the leading edge. The vortex lies outwards across the wing surface at approximately 20 to 25 degrees to the line-of-flight before curving aft to be shed into the wake, where it extends almost from mid semi-span to the wing tip. This vortex induces considerable changes in flow direction, both on and over the wing, and also in the wake. Thus in the wake a maximum downwash of 23 degrees is induced aft of the mid semi-span, and there is an upwash of 17 degrees at the outer edge of the vortex, almost aft of the tip. A good correlation between yawmeter results and the boundary layer flow direction indications from the liquid film technique was obtained.

In this article a magnetic suspension and balance system suitable for wind tunnel application is discussed. General considerations are presented that illustrate the nature of the problems to be solved as well as one solution of these problems.Finally some initial wind tunnel data are presented.

Over the last five years or so both in the USA (especially by the active group at Berkeley under R. W. Clough(7) ) and the UK, there have been a number of attempts to incorporate the plane triangular element under bending and membrane action into a general matrix displacement analysis of shells. All investigators stressed the positive aspects of the triangular element, which arise from its adaptability to any geometrical form of the shell, but noted at the same time some serious shortcomings due to a partial satisfaction of the kinematic compatibility conditions at the boundary. Thus we have within the Matrix Displacement Method the rather unusual example of a type of element whose deficiency with respect to kinematic compatibility does not automatically ensure convergence to the exact result as the number of elements is increased.

The paper presents the results of the aerodynamic optimisation of an Unmanned Aerial System's wing using a morphing approach. The shape deformation of the wing is achieved by placing actuator lines at several positions along its span. For each flight condition, the optimal displacements are found by using a combination of the new Artificial Bee Colony algorithm and a classical gradient-based search routine. The wing aerodynamic characteristics are calculated with an efficient nonlinear lifting line method coupled with a two-dimensional viscous flow solver. The optimisations are performed at angles of attack below the maximum lift angle, with the aim of improving the Hydra Technologies UAS-S4 wing lift-to-drag ratio. Several configurations of the morphing wing are proposed, each with a different number of actuation lines, and the improvements obtained by these configurations are analysed and compared.

An unusual philosophical approach is proposed here to decarbonise larger civil aircraft that fly long ranges and consume a large fraction of civil aviation fuel. These inject an important amount of carbon emissions into the atmosphere, and holistic decarbonising solutions must consider this sector. A philosophical–analytical investigation is reported here on the feasibility of an airliner family to fly over long ranges and assist in the elimination of carbon dioxide emissions from civil aviation.

Backed by state-of-the-art correlations and engine performance integration analytical tools, a family of large airliners is proposed based on the development and integration of the body of a very large two-deck four-engine airliner with the engines, wings and flight control surfaces of a very long-range twin widebody jet. The proposal is for a derivative design and not a retrofit. This derivative design may enable a swifter entry to service.

The main contribution of this study is a philosophical one: a carefully evaluated aircraft family that appears to have very good potential for first-generation hydrogen-fuelled airliners using gas turbine engines for propulsion. This family offers three variants: a 380-passenger aircraft with a range of 3,300nm, a 330-passenger aircraft with a range of 4,800nm and a 230-passenger aircraft with a range of 5,500nm. The latter range is crucially important because it permits travel from anywhere in the globe to anywhere else with only one stop. The jet engine of choice is a 450kN high-bypass turbofan.

A brief narration on significant accomplishments in computational fluid dynamics (CFD) for basic research and aerospace application is attempted to highlight the outstanding achievements by scientists and engineers of this discipline. To traverse such a vast domain, numerous and excellent contributions to CFD will be unintentionally overlooked by the author’s limited exposure. Nevertheless it is an ardent hope that the present abridged literature review will aid to reaffirm excellence in research and to identify knowledge shortfalls both in aerodynamics and its modeling and simulation capability. The future modeling and simulation technology needs, as well as potential and fertile research areas, are humbly put forth for consideration.

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).

Experiments were carried out on jets issuing from circular nozzles with grooved exits and the results compared with those of the plain nozzle. The plain nozzle had an exit diameter of 10mm. Because of the introduction of semi-circular grooves at the exit, the effective or equivalent diameter of the grooved nozzles was 10·44mm. The groove lengths were varied as 3, 5 and 8mm. The nozzles were operated at fully expanded sonic and underexpanded exit conditions. The corresponding fully expanded Mach numbers were 1·0 and 1·41. The shock cell structure of the underexpanded jets from grooved nozzles appeared to be weaker than that of the plain nozzle, as indicated by lesser amplitudes of the cyclic variation of the Pitot pressure. The iso-Mach contours indicate that the jet spread along the grooved plane is significantly higher than that along the ungrooved plane. Off-centre peaks were observed in the mean pressure profile of underexpanded jets from grooved nozzles. They were probably due to the streamwise vortices shed from the grooves.

The development of a parametric-based indirect aircraft structural usage monitoring system using artificial neural networks is described. Flight parametric data, captured during Operational Loads Measurement have been used to predict strains or stresses at key structural locations for several military aircraft types, using mapping relationships determined by artificial neural networks. A framework for the development of a neural network-based structural usage monitor is discussed and the basic architecture of the multilayer perceptron artificial neural network is described. Additionally, results from case studies are presented. It is concluded that this technology could provide the basis for accurate, cost-effective structural usage monitoring systems across the range of military aircraft types and roles.

The natural frequencies and mode shapes of vibration of cantilever aerofoil cross-section blades of pre-twist angle in the range 0 to 90 degrees are obtained. The beams are 152·4 mm long and the width / thickness ratio is such that they may be regarded as slender. Theoretical frequency ratios and mode shapes of vibration, neglecting shear and rotary inertia effects, are obtained for two sets of beams, one with clockwise pre-twist relative to the root cross-section and the other with anti-clockwise pre-twist. The effect of variation in the value of the centre-of-flexure coordinates upon the natural frequency ratios and mode shapes of vibration is investigated. The theoretical results are compared to corresponding experimental results.

A method has been given by Klitchieff for analysis of grillages of intersecting beams under lateral loading. Klitchieff's solution was in the form of infinite Fourier series and the expressions for bending moments converged approximately as Σ n–2, making numerical computation difficult. This paper shows that, for grillages in which one set of beams are evenly spaced, Klitchieff’s solution can be adapted to give finite series for deflections arid bending moments at the intersections of the beams. The number of terms in the series is not greater than the number of beams in the equally spaced set (referred to in this paper as set “ A ”). A numerical example is given and an accurate solution is obtained with appreciably less computation than other known methods would entail.

The theory of modal control is concerned with the design of control loops for systems with governing equations which can be adequately approximated by linear vector-matrix equations of the form

where x is the system state vector, y is the control vector, A is the matrix of the uncontrolled system, and B is the control matrix.

The current understanding of the aerodynamics of birds in formation flights is mostly based on field observations. The interpretation of these observations is usually made using simplified aerodynamic models. Here, we investigate the aerodynamic aspects of formation flights. We use a potential flow solver based on the unsteady vortex lattice method (UVLM) to simulate the flow over flapping wings flying in grouping arrangements and in proximity of each other. UVLM has the capability to capture unsteady effects associated with the wake. We demonstrate the importance of properly capturing these effects to assess aerodynamic performance of flapping wings in formation flight. Simulations show that flying in line formation at adequate spacing enables significant increase in the lift and thrust and reduces power consumption. This is mainly due to the interaction between the trailing birds and the previously-shed wake vorticity from the leading bird. Moreover, enlarging the group of birds flying in formation further improves the aerodynamic performance for each bird in the flock. Therefore, birds get significant benefit of such organised patterns to minimise power consumption while traveling over long distances without stop and feeding. This justifies formation flight as being beneficial for bird evolution without regard to potential social benefits, such as, visual and communication factors for group protection and predator evasion.

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.

This paper presents infra-red signature data for a small-scale, low pressure ratio turbojet engine typical of that used in unmanned air vehicle applications. The aim of the study was to test a number of different convergent nozzle designs concentrating on those with trailing edge modifications. The engine used in the tests has a single stage centrifugal compressor and radial inflow turbine and is designed to produce approximately 150N of thrust at 103,500rpm using liquid propane fuel. The test rig consisted of a calibrated thrust stand whilst the engine was controlled through an electronic engine control unit and laptop PC. The jet plume was visualised using an infra-red spectroradiometer which yielded qualitative data across the infra-red spectrum. Simultaneous measurements were also made of the engine thrust. A Pitot probe was used to take pressure readings across different sections of the exhaust flow. Analysis of the infrared signature of the engine exhaust plume and any thrust penalty yielded a performance comparison for each of the nozzles tested. Correlation of engine thrust with engine rpm showed that, within the accuracy of the measurements, there was no significant thrust penalty associated with the notched nozzles. Infra-red imagery of the plain and 60° notched nozzles indicated that the latter reduced the length of the hottest part of the exhaust plume by approximately 33%. The spectroradiometer data shows a significant reduction in spectral radiance for the CO2 wavelength of approximately 4·3µm when the notched nozzles are used. The 60° notched nozzle appeared to perform best in reducing the spectral radiance at this wavelength. Centreline total pressure measurements in the exhaust plume correlated well with the infra-red imagery in that a potential core length reduction of up to 30% could be achieved using the 60° notched nozzle. Total pressure contours recorded 20mm (0·43D) downstream of the nozzle exit plane suggest that the notched nozzles are promoting increased mixing through radial spreading of the jet possibly associated with increased streamwise vorticity (although the latter could not be confirmed). There were also signs that the jet plumes being investigated were swirling.

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.

The aerodynamics of insect-like flapping are dominated by the production of a large, stable, and lift-enhancing leading-edge vortex (LEV) above the wing. In this paper the phenomenology behind the LEV is explored, the reasons for its stability are investigated, and the effects on the LEV of changing Reynolds number or angle-of-attack are studied. A predominantly-computational method has been used, validated against both existing and new experimental data. It is concluded that the LEV is stable over the entire range of Reynolds numbers investigated here and that changes in angle-of-attack do not affect the LEV’s stability. The primary motivation of the current work is to ascertain whether insect-like flapping can be successfully ‘scaled up’ to produce a flapping-wing micro air vehicle (FMAV) and the results presented here suggest that this should be the case.

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.

This is a discussion of the various types of flow which may occur on thin swept wings when the air stream is not attached to the wing surface everywhere. In particular, the behaviour of vortex sheets which arise from such mixed flows is examined.