Fatigue in aircraft structures is a problem for which quantitative and generally accepted solutions are not available as yet. Even in recent years significant fatigue failures in aircraft structures have occurred. Some of these failures can be attributed to either poor design or underestimating of the fatigue environment. Unfortunately such statements are of little help since there are a confusingly large number of aspects involved. For a designer considering fatigue problems, experience, qualitative understanding and a good intuition are therefore essential.

This paper describes a comparative evaluation of two data smoothing algorithms for use in a two step estimation-beforemodelling procedure for aircraft parameter identification. A simple fixed lag smoother is compared with the usual, and more complex, modified-Bryson-Frazier smoother in the first, state estimation, step of the aircraft parameter identification procedure. The comparison is illustrated by application to the analysis of the Dutch Roll motion of the Embraer EMB-312 Tucano. Both algorithms were found to give results of comparable accuracy although the fixed lag smoother is computationally more efficient. It was concluded that the fixed lag smoother algorithm is an acceptable alternative to the modified-Bryson-Frazier algorithm in aircraft parameter identification applications.

This paper investigates a novel method for the control of aircraft. The concept consists of articulated split wing-tips, independently actuated and mounted on a baseline flying wing. The general philosophy behind the concept was that adequate control of a flying wing about its three axes could be obtained through local modifications of the dihedral angle at the wing-tips, thus providing an alternative to conventional control effectors such as elevons and drag rudders. Preliminary computations with a vortex lattice model and subsequent wind tunnel tests and Navier-Stokes computations demonstrate the viability of the concept for co-ordinated turns, with individual and/or combined wing-tip deflections producing multi-axis, coupled control moments. The multi-axis nature of the generated moments tends to over-actuate the flight control system, leading to some redundancy, which could be exploited to optimise secondary objective functions such as drag or bending moment.

The flow past families of two-dimensional cylinders and bodies of revolution with elliptic noses at zero incidence has been examined at free stream Mach numbers between 1·42 and 1·82. The observations include schlieren photographs and measurements of the pressure distributions at the surface.

The measured pressure distributions and positions of the detached bow waves are found to be in fair agreement with values calculated by a number of methods. The drag coefficients of the slender two-dimensional elliptic noses are considerably higher than those calculated for wedges of the same axis ratio with attached shock waves, but the drags become almost equal when the axis ratio approaches the value for subsonic flow behind the bow wave of the wedge. For most axis ratios the drag coefficients of the elliptic-nosed bodies of revolution are lower than those for cones of the same axis ratio.

The noise levels of a jet issuing from a long pipe are compared with those of a jet having a square velocity profile at the exit. A subsonic noise reduction of between 2 and 5 decibels for various conditions is found for the case of the flow emerging with an approximately “turbulent pipe-flow” velocity profile for the same maximum jet velocity, but this is at the expense of a loss in thrust of a quarter. On comparison with a jet of smaller diameter which has an equal thrust for the same maximum jet velocity, it is found that the changes in noise level are rather smaller. For jets of equal diameters, the effects on the subsonic aerodynamic noise generated of a reduction of velocity gradient near the boundary are more than offset by the increased velocities necessary near the centre of the jet to obtain equal thrust. It is concluded that if the effect of differences in initial turbulence can be neglected the use of an auxiliary flow forming a comparatively thin sheath of slower moving fluid at the exit is not likely to result in large decreases in the subsonic noise level, and that a general reduction in jet velocity is more effective.

Above the critical pressure larger reductions of up to 10 decibels are found. These are consistent with a delay of the onset of the self-maintained shock-produced noise.

A method is presented for the calculation of the streamwise component of vorticity, for flows in rotating passages. The method may be regarded as an extension of the methods applied in recent years to the calculation of secondary flows in stationary passages.

Attention is concentrated on the quantity I=p+½ρV2−½ρU2, where V is the fluid velocity relative to the rotor and U is the rotor tangential velocity. I remains constant along a streamline in the rotor, and is found to enter into the equation for the generation of vorticity in much the same way as the total head enters for flow in stationary passages.

An approximate calculation of the streamwise vorticity generated in a simple axial-flow rotor is made, and qualitative consideration is given to the flow in a centrifugal impeller. Just as for the calculation of secondary flows in stationary passages, an approximate shape of the streamline must be assumed before the secondary flows can be calculated.

Flow transition in the developing region of a plane wall jet is studied experimentally by means of hot-wire measurements. The Reynolds number, based on the jet exit velocity and nozzle exit height, varies from 3 x 102 to 3 x 104 in the investigations. The results indicate that the most important factor causing wall jet flow to change from the initial exit state to the final turbulent state is the rapid increase of turbulent intensity from the formation of vortical structures and their interactions with the wall. Data also shows that the transition process of the wall jet depends on the Reynolds number in the operating range. When the Reynolds number is greater than 2000, the mean velocity distribution of the wall jet changes directly from a top-hat profile at the nozzle exit to a turbulent profile in the turbulent developed region. When the Reynolds number is smaller than 2000, Glauert's laminar velocity profile is found between the nozzle exit and the turbulent region. It is also found that the transition will be delayed and the transition region prolonged with the decrease of the Reynolds number.

This paper describes the development of an approach to handling qualities investigation that can be applied at an early stage in t he design of the vehicle. It makes use of inverse simulation techniques, together with a pilot model to provide an integrated description of the man-machine control system. In order to incorporate pilot effects into data generated by inverse simulation, the output from an inverse simulation run is applied as input to a closed-loop system model that includes the vehicle dynamics and a simple parametric model of the pilot. Parameters of the pilot model are determined by optimisation and the pilot effect is added to the system output. Validation of the approach is achieved through a case study involving a predefined mission task involving a lateral manoeuvre. Equalisation characteristics estimated for each pilot are compared with those found by inverse simulation for the same manoeuvre. This approach may be applied using a simple real-time simulation on a desk-top computer and could be of value in identifying any potential deficiencies in a helicopter flight control system at an early stage in its development

The field data characterising aircraft accidental in-service damage was collected, sorted and processed. By means of probabilistic analysis, the wing damageability statistical parameters were determined. The scenarios of wing accidental impacts were described and the qualitative distribution of impact intensity over the wing surfaces was obtained. By means of original analytical method, the metal dent depth data were converted into impact energy data and energy probabilistic distributions were established. It was shown that the functional relationships generated on domestic data are generally consistent with similar foreign results obtained on other types of aircraft with serious differences in operating conditions. Along with realistic impact damage scenarios, the high energy impact events were considered. It was noted that in some cases severe damage events should not be addressed as extremely improbable and should be included into design and certification process.

Each year greenhouse gas emissions remain high the climate mitigation and adaptation challenges grow. The economic downturn was already in train in 2008, yet concentrations of CO2 increased unabated. Without concerted effort across all sectors there will be little chance of avoiding ‘dangerous climate change’ and the aviation sector has a clear role to play. Current and forthcoming technologies, operational practices and behavioural change offer widespread opportunities for other sectors to mitigate their CO2 emissions in absolute terms, but as they do so, aviation will become an increasingly important player. By comparing a range of global cross-sector emission scenarios with existing aviation projections, this paper illustrates the importance of understanding the future context with regard to other sectors when assessing the aviation industry’s potential impact. Given growth projections for aviation and the relatively slow pace of technological change, aviation’s proportion of 2050 global CO2 emissions is low only in those global cross-sector emission scenarios where there is a high probability of ‘dangerous climate change’. For a ‘reasonable’ (>50%) chance of avoiding ‘dangerous climate change’, the most technologically radical scenarios for aviation make up 15% of global CO2 in 2050 and conventional scenarios exceed the carbon budget entirely. Only by recognising that aviation’s currently projected emissions are incompatible with avoiding ‘dangerous climate change’ can the industry fully grasp the challenge of accelerating innovation and managing demand to deliver a more sustainable route to 2050 and beyond.

This paper investigates the effects of the design variables of an aerial deployment mechanism on the robustness of the aerial deployment through a multibody dynamics simulation. The aircraft is modelled as three joined rigid bodies: a right wing, a left wing and a centre body. A spring-loaded hinge is adopted as an actuator for deployment. The design variables are the hinge torque and the deployment timing. The robustness is evaluated using a sigma level method. The margins for the safe deployment conditions are set for the evaluation functions. The dispersive input variables are the initial drop velocity, the surrounding gust velocity, the initial pitch angle and the initial height. The design point with a deployment torque scale value F of 0.7 and a right-wing deployment delay time TSR of 1.0 s can safely deploy in the low-torque deployment condition. This design point is able to accomplish both a safe deployment and a lightweight deployment mechanism.

Multicriteria trajectory optimisation is expected to increase aviation safety, efficiency and environmental compatibility, although neither the theoretical calculation of such optimised trajectories nor their implementation into today’s already safe and efficient air traffic flow management reaches a satisfying level of fidelity. The calibration of the underlying objective functions leading to the virtually best available solution is complicated and hard to identify, since the participating stakeholders are very competitive. Furthermore, operational uncertainties hamper the robust identification of an optimised trajectory. These uncertainties may arise from severe weather conditions or operational changes in the airport management. In this study, the impact of multicriteria optimised free route trajectories on the air traffic flow management is analysed and compared with a validated reference scenario which consists of real flown trajectories during a peak hour of Europe’s complete air traffic in the upper airspace. Therefore, the TOolchain for Multicriteria Aircraft Trajectory Optimisation (TOMATO) is used for both the multicriteria optimisation of txrajectories and the calculation of the reference scenario. First, this paper gives evidence for the validity of the simulation environment TOMATO, by comparison of the integrated reference results with those of the commercial fast-time air traffic optimiser (AirTOp). Second, TOMATO is used for the multicriteria trajectory optimisation, the assessment of the trajectories and the calculation of their integrated impact on the air traffic flow management, which in turn is compared with the reference scenario. Thereby, significant differences between the reference scenario and the optimised scenario can be identified, especially considering the taskload due to frequent altitude changes and rescinded constraints given by waypoints in the reference scenario. The latter and the strong impact of wind direction and wind speed cause wide differences in the patterns of the lateral trajectories in the airspace with significant influence on the airspace capacity and controller’s taskload. With this study, the possibility of a successful 4D free route implementation into Europe’s upper airspace is proven even over central Europe during peak hours, when capacity constraints are already reaching their limits.

In the present work natural laminar flow (NLF) and hybrid laminar flow (HLF) wing designs are presented which were obtained by combining new methodologies with experience and knowledge obtained with traditional laminar wing design methods. The NLF wing design is performed for wing-body configurations with backward swept wing (BSW) and forward swept wing (FSW). Initial aerofoil sections were obtained by using a new sectional conical wing method which allows the design of transonic wing sections, taking into account the effects of sweep and taper for the computational cost of a 2D method. Except for flow regions with strong 3D influence, wings constructed with these aerofoils showed an acceptably large region with laminar boundary layer and small shocks at design and specified off-design conditions. For regions close to the body and the tip a 3D inverse design method was further required. For the BSW case, due to cross flow a premature transition occurred. Therefore, a HLF panel was required to obtain a larger laminar region. A suction distribution was obtained using the suction distribution module of the automated target pressure generator (ATPG). This generator optimises the pressure distributions in terms of minimising drag while keeping certain boundary conditions constant, e.g. lift and momentum. Using the ATPG, the laminar extent of the BSW NLF design could be further improved for the inboard wing. With the new methodologies design work was reduced. They lead to a design with reserves that allow for acceptable off-design performance qualities by keeping the wing laminar over a wide range of flight conditions.

• (a)Purpose of Investigation. To determine the least size of stringers necessary to prevent overall buckling of a flat stiffened panel before buckling of the plates between stringers.

• (b)Range of Investigation. The conditions for buckling of a longitudinally stiffened flat panel are established on the assumption that the rotational stiffness of the stringers is negligible. The results are applied to determine the limiting characteristics of the stringers to ensure that these members remain straight up to the stress at which the plates buckle between stringers. The analysis has been carried through in detail for panels with one, two, or three stringers and is capable of extension to four, five or more stringers. This extension appears unnecessary, because the three stringer case is moderately close to the limiting case of a wide panel with an indefinitely large number of stringers.

• (c)Conclusions. The stiffening effect of the stringers, having each the area of section A, and modulus of section AsK2, when attached to a sheet of thickness t and length a, at spacing b depends upon the three ratios (k/t), (a/b) and (As/bt) In general, the size of stringer necessary to ensure that the stringers remain straight until the plates buckle between stringers may be represented by the conditions

  .

  For a panel stiffened by one stringer only suitable values are λ = 4.5 and μ = 12.5, independent of the value of (a/b), while v=0.366 (a/b)2or 23 if 8>(a/b)>6.

  For two or more stringers the same general conditions hold, but it is impossible to assign values to λ and μ which would be fully satisfactory for all values of the ratio a/b. In these cases, therefore, it is more satisfactory to select the values of λ and μ appropriate to each value of the ratio a/b. These values may be obtained from Figs. 2 and 3 where curves of bt/As against (k/t)2for varying values of (a/b) are shown for two and three stringers respectively. For an indefinitely large number of stringers the corresponding condition is

  

  and the second condition (k/t)2 0.366 (a/b)2 is virtually included in the first.

• (d)Further Developments. In the present analysis the resistance of the stringers. to rotation is neglected. Inclusion of this characteristic in the basic analysis is straightforward, but practical interpretation of the results becomes much more difficult. Moreover, the effect of this stiffness in modifying the condition for buckling between stringers is likely to affect the conclusions more than the direct effect of their torsional stiffness in modifying the condition for buckling of the stiffened panel. This consideration is particularly important in cases of light stringers and heavy sheet, when the effective rotational stiffness of the stringers at buckling of the plates between stringers may be greatly reduced as a result of the end loading and may even be negative. Representation of the effect of both longitudinal stiffeners (stringers) and lateral stiffeners (ribs), including their stiffnesses to rotation is also straightforward; but interpretation of the results in forms suitable for practical application would be extremely difficult.

Complicated oscillatory systems may be broken down into component “ sub-systems ” for the purpose of vibration analysis. These will generally submit more readily to analytical treatment. After an introduction to the concept of receptance, the principles underlying this form of analysis are reviewed.

The dynamical properties of simple systems (in the form of their receptances) may be tabulated. By this means the properties of a complicated system may be found by first analysing it into convenient sub-systems and then extracting the properties of the latter from a suitable table. A catalogue of this sort is given for the particular case of conservative torsional systems with finite freedom.

The properties of the composite system which may be readily found in this way are (i) its receptances and (ii) its frequency equation. Tables are given of expressions for these in terms of the receptances of the component sub-systems. All of the tables may easily be extended. The tabulated receptances may also be used for determining relative displacements during free vibration in any principal mode.

A method of presenting information on the vibration characteristics of machinery, which is effectively due to Carter, is illustrated by means of an example. More general adoption by manufacturers of this method (which requires no more computational effort than must normally be made) would lead to enormous savings of labour in calculating natural frequencies of composite systems.

With the return of British civil aviation to normal operating conditions after the war it was found that several problems which had previously given little cause for concern had become acute. More especially, the increase in the volume of traffic, particularly under instrument flying conditions, together with certain indirect consequences of the greater sizes and weights of aircraft, had brought about greatly increased congestion around major airports, and there was, therefore, an urgent requirement for an effective system of close air traffic control to ensure the safety and rapid movement of the traffic.

The Ministry of Civil Aviation immediately gave much attention to this problem of air traffic control, but at once found itself in serious difficulties on account of lack of sufficiently precise knowledge of the traffic and the traffic pattern. The then Controller of Technical and. Operational Services, Sir Conrad Collier, realised that the obtaining of the necessary data concerning air traffic control and other matters of technical interest was essentially a matter of operational research, and in 1947 a small operational research section was established which began work in the autumn of that year.

In industry, there is an ever-increasing requirement not only to design high performance new products but also to deliver them at lower cost and in shorter time. To meet these demanding engineering challenges, it is not sufficient to treat the different disciplines involved in a product design in isolation; rather they must be considered together as an integrated system that reflects the dependencies and interactions of the different disciplines. The design process must be automated to meet the stringent design time-lines. In spite of promising forays for over a decade, automatic design optimisation (ADO) and multidisciplinary optimisation (MDO) has not been widely adapted by the Turbomachinery design practitioners. This presentation will explore some of the technical and nontechnical barriers such as cultural and organisational issues that must be addressed if ADO/MDO is to be used routinely in industry. Some recent, successful application of automatic optimisation is also reported herein.

The rolling up of the vortex sheet shed from a finite aspect ratio wing into discrete vortices has been extensively investigated in the past. Most of these investigations have, however, considered wings having an elliptic loading.

The early design charts of Silvertstein et al gave downwash angles for a variety of flapped and unflapped aerofoils by considering the trailing vortex sheet to be composed of discrete line vortices but ignored the rolling up process that takes place.

An experimental study was undertaken to measure the instantaneous lift and propulsive forces generated by a rigid, sinusoidally flapping wing at Reynolds numbers ranging from 72 000 to 120 000. The wing was flapped about a fulcrum at the root at frequencies of 2.55 Hz and 3.30 Hz. Instrumentation and a data reduction system were developed to provide quick and reliable assessment of the effects of changes in flapping amplitude, flapping frequency, flapping plane angle, wing incidence, and windspeed. In this way it was attempted to model, however crudely, the forward flapping flight of birds.

The instantaneous values of lift coefficient and thrust coefficient were plotted against position in the cycle while the average CL and CT were plotted against advance ratio (ratio of the windspeed to the maximum tip speed velocity). The experimental results were compared to the theories of Walker(e) and Hoist ' Kuchemann(4) both modified to correspond more closely to the experimental conditions.

Some of the uses of ground based flight simulation as a research tool to aid the design of new aircraft and their equipment are described. The function of the simulator is to provide a method for investigating humon flying tasks in the laboratory, so that the relationship between the pilot's capabilities and the equipment's characteristics can be systematically studied. The paper is presented in two parts describing recent work on two research simulators at the RAE.

Part I deals with the use of simulation for studying aircraft stability and control characteristics. The equipment used at RAE for this work is described, with particular emphasis on methods of presenting to the pilot a simulated view of the outside world, and for reproducing some of the motion cues which he experiences in flight. Experimental evidence of the importance of these simulation cues when making aircraft control assessments is also presented. Several examples of simulation studies into the control of conventional and VTOL aircraft are given, to illustrate the type of research problems in this field which may be tackled and the techniques involved in solving them.