Modern technology places great demands on structural material performance. Until the 1950s, structures were invariably designed according to safe-life principles, for which a crack-free service life was required and assumed. However, the tendency to much longer utilisation of aircraft, and latterly the introduction of high strength materials, have focused attention on damage tolerant design in which allowance must be made for possible cracking and its certain detection before catastrophic failure. Especially for high strength materials is this change in philosophy necessary, because the improved static strengths (and hence potentially increased structural efficiency) rarely result in comparable improvements in fatigue properties, and are usually accompanied by a decline in fracture toughness and in greater susceptibility to environment-induced cracking.

By using the methods of the Calculus of Finite Differences, expressions are obtained for the nodal moments and deflections of a simply-supported grillage, subjected to a loading constant along one set of beams and having a sinusoidal variation along the other set of beams. A simple example verifies the expressions and illustrates their use.

The paper presents a solution to the buckling of infinitely long plates when they are reinforced by transverse stiffeners possessing both torsional and flexural rigidity. The cases of both edges being clamped and simply-supported are dealt with. Numerical results are presented for the ratio of torsional rigidity to flexural rigidity as obtained with a thin-walled circular tube. When the stiffeners are completely rigid, in which case the individual panels are clamped along the transverse edges, the results obtained are in agreement with existing solutions for isolated rectangular plates.

The shear stiffness of a beam in bending is commonly determined by measuring the central deflection δ in a three-point loading test for a range of values of the span L under a constant central load. The value of bending stiffness, EI, and shear stiffness, GA, can then be obtained by plotting δ/L against L2 and drawing the best straight line through the points so obtained. The slope of this line for unit central load is then 1/48EI and its intercept on the δ/L axis is 1/4GA on the simple “Engineer's” theory.

The aerodynamics of a circular jet emerging normally from a wall into a cross flow is critically assessed. Some general trends are obtained when viscous effects are neglected. But viscous effects dominate the real flow behaviour. Attempts to predict these real effects are reviewed and their anomalies highlighted. The importance of understanding and specifying the jet exit conditions is stressed, which is relevant to complete numerical solutions.

A simplified approximate solution for clamped skewed plates under uniform pressure is presented. Variations in maximum edge and centre moments with skew angle and sides ratio are shown plotted. The convergency of the solution is checked and results compared with the limited available data. Results were verified by an experimental investigation.

As a result of the failure of parts of the airframe on several types of aircraft having high intensity acoustic environments, the military and civil certification authorities are requiring assurances that new designs have adequate strength to resist the acoustic loads. In most cases failures have been detected early enough to prevent serious impairment of the aircraft's structural strength. In some cases this has led to unsatisfactory maintenance and inspection requirements and there is also the possibility of more serious failures arising from the extension of undetected damage. The purpose of this paper is to review the test data available and to present a design procedure for structures in the region of high noise levels. As it is impractical to apply the full theoretical treatment in design because of computational complexity, an approximation is presented and the predicted results are compared with the experimental values.

In the spring of 1980, a long range derivative of the L-1011 TriStar entered commercial airline service with an extended wing span and a digital Active Control System (ACS). The extended wing span improves fuel efficiency by 3 to 3½ % due to a higher wing aspect ratio and lower induced drag. The digital ACS provides load alleviation functions by symmetric motions of the outboard ailerons which relieve the loads induced by the increased wing span. This load relief minimises the amount of structural modifications required to the basic L-1011 wing to accommodate the increased wing span.

This contribution is based on the Lanchester Lecture of the Royal Aeronautical Society held in Manchester, UK, in October 2009. The lecture reflected on further possible perspectives for aeronautical research, in particular with respect to aerodynamics. The air transport system of today has matured to very high performance and safety standards. However, due to the ongoing global developments of economy and ecology, new challenges have to be faced. Thus, the relevance of all engineering aeronautical disciplines contributing to the final aircraft configuration is now closely scrutinised, and especially aerodynamics is challenged whether any investment in further research is still justified. In this contribution, it is argued that, when carefully exploited, aerodynamics still offers a substantial technological perspective, well beyond the concepts for the next generation of aircraft as currently being pursued. As an example, the aerodynamics of a forward swept-wing are discussed and it is outlined that, when combining the transonic characteristics of a forward swept-wing with the advantages of natural laminar-flow, considerable improvements in efficiency may be gained. However, this potential may only be leveraged through a strategic, long-term orientation of fundamental and applied research.

The semi-analytical prediction of pollutants emissions from gas turbines in the conceptual design phase is addressed in this paper. The necessity of this work arose from an urgent need for a comprehensive model that can quickly provide data in the conceptual design phase. Based on the available inputs data in the initial phases of the design process, a chemical reactor network (CRN) is defined to model the combustion with a detailed chemistry. In this way, three different chemical mechanisms are studied for Jet-A aviation fuel. Furthermore, the droplet evaporation for liquid fuel and the non-uniformity in fuel-air mixture are modelled. The results of a developed augmented modelling tool are compared with the pollutants data of two annular engine's combustors. The CRN results have good agreement with the actual engine test rig emissions output. In conclusion, the augmented CRN has shown to be efficient in predicting engine emissions with a very short executing time (few seconds) using a small CPU requirement such as a personal computer.

During the decay process the shape of wake vortices changes significantly which has an influence on the encounter characteristics, hence on the encounter hazard. In order to evaluate the influence of vortex deformation on the wake encounter hazard, in-flight simulations with the DLR research aircraft ATTAS were carried out. For a realistic analysis of wake encounters flow fields of matured vortices were generated with large eddy simulations. These flow fields were used for the determination of histories of induced forces and moments acting on the wake encountering aircraft. The force and moment histories were then fed into the equations of motions of the non-linear six degree-of-freedom in-flight simulation of the DLR research aircraft ATTAS. In order to compare different stages of vortex deformation, encounters were simulated in flight with wavy vortices and vortex rings. The most important benefit of the in-flight-simulation is the realistic environment, which enables a realistic assessment of pilots’ encounter acceptance. The analysed scenario was of a wake encounter during final approach. The encounter conditions correspond to separation distances of about 4nm and 5nm behind an aircraft of the ‘heavy’ category. During the encounters the ATTAS was flown under manual control. Altogether 31 encounters were simulated in flight, 9 with wavy vortices and 22 with vortex rings.

Multidisciplinary design and innovative highly automated manufacturing methods are increasingly important to today’s aircraft industry: multidis-ciplinary design because it reduces lead-time and results in a better design, and automated manufacturing methods because they are more capable and reduce manufacturing cost. In this paper a cost estimation model is presented that integrates the manufacturing cost of friction stir welded connections within a multidisciplinary design decision tool. Due to the fact that friction stir welding is a new manufacturing method, the cost estimation model is based on the actual process physics, meaning what the process looks like in terms of processing speeds and characteristics. As an integral part of a multidisciplinary design framework, the developed cost estimation model contributes to a design support tool that assesses not only manufacturing but also structural and aerodynamic issues. It is shown that the cost model developed can be integrated into this more holistic design process support architecture. The predicted costs are accurate to the historical data and allow tradeoff of manufacturing and economic considerations within the context of the multidisciplinary design tool. The tradeoff capability is highlighted through a presented case study that compares the friction stir welding process as an alternative solution to more tradition riveting. Most importantly, this results in a quantitative tradeoff between two processes that shows the manufacturing cycle time of friction stir welding to be reduced by 60% and the recurring assembly cost by 20%.

Motivated by the idea of turbomachinery active subspace performance maps, this paper studies dimension reduction in turbomachinery 3D CFD simulations. First, we show that these subspaces exist across different blades—under the same parametrisation—largely independent of their Mach number or Reynolds number. This is demonstrated via a numerical study on three different blades. Then, in an attempt to reduce the computational cost of identifying a suitable dimension reducing subspace, we examine statistical sufficient dimension reduction methods, including sliced inverse regression, sliced average variance estimation, principal Hessian directions and contour regression. Unsatisfied by these results, we evaluate a new idea based on polynomial variable projection—a non-linear least-squares problem. Our results using polynomial variable projection clearly demonstrate that one can accurately identify dimension reducing subspaces for turbomachinery functionals at a fraction of the cost associated with prior methods. We apply these subspaces to the problem of comparing design configurations across different flight points on a working line of a fan blade. We demonstrate how designs that offer a healthy compromise between performance at cruise and sea-level conditions can be easily found by visually inspecting their subspaces.

Computer based mathematical models describing the aircraft evacuation process have a vital role to play in aviation safety. However, such models have a heavy dependency on real evacuation data in order to (a) identify the key processes and factors associated with evacuation, (b) quantify variables and parameters associated with the identified factors/processes and finally (c) validate the models. The Fire Safety Engineering Group of the University of Greenwich is undertaking a large data extraction exercise from three major data sources in order to address these issues. This paper describes the extraction and application of data from one of these sources — aviation accident reports. To aid in the storage and analysis of the raw data, a computer database known as AASK (aircraft accident statistics and knowledge) is under development. AASK is being developed to store human observational and anecdotal data contained in accident reports and interview transcripts. AASK comprises four component sub-databases. These consist of the ACCIDENT (crash details), FLIGHT ATTENDANT (observations and actions of the flight attendants), FATALS (details concerning passenger fatalities) and PAX (observations and accounts from individual passengers) databases. AASK currently contains information from 25 survivable aviation accidents covering the period 4 April 1977 to 6 August 1995, involving some 2415 passengers, 2210 survivors, 205 fatalities and accounts from 669 people. In addition to aiding the development of aircraft evacuation models, AASK is also being used to challenge some of the myths which proliferate in the aviation safety industry such as, passenger exit selection during evacuation, nature and frequency of seat jumping, speed of passenger response and group dynamics. AASK can also be used to aid in the development of a more comprehensive approach to conducting post accident interviews, and will eventually be used to store the data directly.

In this note it is suggested that the three-tube yaw probe be placed on the boundary and used as a Preston tube to measure the boundary shear stress when both its magnitude and direction are unknown. The calibration results necessary for this method are also presented herein.

A general-purpose low turbulence wind tunnel was constructed using the design method of Bradshaw. Sound absorbent material was used in every four corners to decrease sound intensity produced by a fan. The longitudinal component of turbulence intensity at the centre of the closed working section is less than 0.02% of the mean velocity in the speed range between 18 m/s arid 53 m/s. The mean velocity variations across the working section are within ±0.1% of the mean velocity.

Performance measurements have been done at representative tunnel cross sections to clarify the behaviour of flow in the tunnel. This work differs from previous studies in the sense that emphasis is placed not only on velocity distributions, but also on turbulence intensity distributions at several cross sections of the tunnel.

The critical Reynolds number for a flat plate at zero incidence, Rec = 3.5 x 106, measured in a stream of the very low turbulence intensity of 0.016%, is larger than that reported by Schubauer and Skramstad.

A family of alternative expressions is presented suitable for the representation of the lateral deflection of rectangular plate elements in bending. Such expressions are extensions of a simple polynomial representation assumed in earlier work. The new expressions are such that not all displacement continuity conditions are met completely but, nonetheless, a criterion ensuring convergence of numerical results to true stiffness levels is satisfied. Deflection and natural frequency estimates based on one expression of the proposed family demonstrate rapid convergence and high numerical accuracy.

A meeting of the Royal Aeronautical Society was held in the Library of the Society at 4 Hamilton Place, London, W.l, on Tuesday 25th September 1945, at which M. L. Malavard, of the Laboratoire des Analogies Electriques, Paris, lectured on “The Use of Rheo-electrical Analogies for the solution of Aerodynamical Problems.” In the chair, Mr. E. F. Relf, C.B.E, F.R.S., F.R.Ae.S. (at that time Superintendent of the Aeronautics Division, National Physical Laboratory).

The integration inverse method has been extensively investigated in the past decade; however, none of the researches fully addresses the stability analysis of the method that is crucial to successful implementation. This paper presents a stability test to analyse the global stability of the integration inverse method for linear time-invariant systems. A stable solution may be obtained from careful selection of the discretisation interval using the proposed stability test. A discrete model is derived to approximate the Newton's scheme in the inverse method. With this approximate model, the stability of the inverse method can be examined. The stability test is exact for linear systems and can be extended to the inverse method for non-linear aircraft simulations by considering an equivalent linear model for each point of the flight envelopes. Guidelines for selection of appropriate reference points in the inverse simulation are given.