Exact solutions for the steady circulatory flow about a circular cylinder with suction applied at its surface have been obtained for the following cases: —

• (a) the cylinder at rest and a circulation at infinity,

• (b) the cylinder rotating and zero circulation at infinity,

• (c) the cylinder rotating in either direction with a circulation at infinity,

provided that, in general, the suction velocity is greater than a certain limiting value.

There is a particular case of (c) when the circulation at the cylinder equals that at infinity in magnitude and direction, for which the circulation is constant at all radii, for all values of R—presumably for blowing as well as suction.

The addition of a uniform stream parallel to the cylinder axis does not affect the circulatory motion, and the solution for the combined flow, which is exact, has some practical interest in that it demonstrates that the boundary layer on a long, circular, rotating boss can be kept thin and that the circulatory motion outside the boundary layer can be prevented by the suitable application of suction before starting the rotation.

The flow (a) is important in connection with the withdrawal of the Thwaites flap. The existence of steady state solutions, coupled with the arguments of Section 5.2, make it appear highly probable that a wing employing this device will retain its lift indefinitely, when the flap is withdrawn.

For high rates of suction flow, the flows exhibit the characteristics of boundary layer flows, in that the vorticity is confined to a narrow annulus enveloping the cylinder and the velocity distribution near the wall tends to the well-known asymptotic distribution of Griffith and Meredith.

It would be interesting to investigate these flows experimentally and to examine the stability of the laminar flow under suction conditions.

The analysis throughout is simple and expressions are obtained for the velocity distribution, the stresses, the vorticity and the torque, and so on.

The paper identifies speed, agility, human interface, generation of sensitivity information, task decomposition, and data transmission (including storage) as important attributes for a computer environment to have in order to support engineering design effectively. It is argued that when examined in terms of these attributes the presently available environment can be shown to be inadequate. A radical improvement is needed, and it may be achieved by combining new methods that have recently emerged from multidisciplinary design optimisation (MDO) with massively parallel processing computer technology. The caveat is that, for successful use of that technology in engineering computing, new paradigms for computing will have to be developed - specifically, innovative algorithms that are intrinsically parallel so that their performance scales up linearly with the number of processors. It may be speculated that the idea of simulating a complex behaviour by interaction of a large number of very simple models may be an inspiration for the above algorithms; the cellular automata are an example. Because of the long lead time needed to develop and mature new paradigms, development should begin now, even though the widespread availability of massively parallel processing is still a few years away.

The method of parameter differentiation is applied to the solution of the non-linear, two-point, ordinary differential equations resulting from an analysis of the finite bending of a thin-walled tube. Starting from a given set of solutions of the differential equations for a particular value of the curvature parameter α, solutions for a range of values of α can be obtained by this method non-iteratively. Very close agreement was obtained with solutions using iterative numerical methods.

This report gives extended data for the flow past roughness trips to provoke boundary-layer transition. This first Part reports work on wire trips giving data for the forces on the wire, the positions of recovery and effective origin of the turbulent layer, the length of the separation bubble and the beginning and end of transition. Parts 2 and 3 describe similar work on spherical roughness trips, used singly and in rows, looking in more detail at the spanwise variations in the downstream boundary-layer.

A study has been made of the effect of inlet conditions on the performance of conical diffusers with 4:1 area ratio and 5 and 10 degrees total angle of expansion. The conditions at entry were varied by using different approach lengths of diffuser inlet diameter, and by means of projecting annular screens of woven wire cloth. With this new technique it was possible to vary the velocity distribution substantially within moderate settling lengths, and to produce velocity profiles with inflections. The suitability of the annular screen method of boundary layer generation for diffuser investigations was confirmed from the comparison with the approach length results.

Energy and pressure coefficients, as well as diffuser energy efficiency and the conversion efficiency, were found to depend on the diffuser angle β and the momentum thickness ratio at inlet θ/D0. The latter emerged as one of the chief parameters controlling diffuser performance. Variation in the inlet shape parameter H of the order of 20 per cent did not significantly affect the pressure recovery or the losses in the diffuser. For moderately thick boundary layers, θ/D0 X β<0·1, the diffuser angle could be eliminated as a parameter by plotting the pressure coefficient against θ/D0 X β. The Reynolds number based on diffuser inlet diameter was, for essentially incompressible flow, 2·5 X 105 in all tests.

The “reference enthalpy” and local flat plate concepts are used to derive some simple expressions for the turbulent heat transfer rate distribution over an arbitrary body at supersonic and hypersonic speeds. The connection between pressure and heat transfer rate is established and tested against a number of experimental data.

This paper is concerned with the development of a failure initiation and progressive failure analysis (PFA) method for advanced composite structures. The present PFA model is capable of predicting interactive out-of-plane and in-plane failure modes observed in fiber reinforced composite laminates including interlaminar behavior and matrix microdamage at the mesoscale. A probability analysis tool is coupled with the PFA to account for uncertainty in modelling parameters caused by material variability and manufacturing inconsistencies. The progressive damage response of a laminated composite panel with an initial delamination is studied and used to demonstrate the PFA modelling framework that is presented here.

In this work, we qualify and quantify the advantages of using SODAR (sonic detection and ranging) from current scenarios of Aeronautical Meteorology, with the goal of establishing a conceptual model for runway change procedures at Guarulhos International Airport (São Paulo, Brazil). The methods consist of analysing data from the Department of Airspace Control (DECEA) and Brazilian Airport Infrastructure (Infraero) about the reports of runway changes in Guarulhos in addition to SODAR data from September 2011 and December 2013. It is noted that in 234 analysed cases of runway change, there were significant periods of weak-intensity wind on the surface as well as the anticipated modification wind direction at altitude detected by sonic detection and ranging (SODAR), indicating future changes in levels closer to the surface. By examining the intersection of both scenarios, it is possible to observe that there is enough time for the air traffic controller to anticipate the needed runway change while minimising the impact on the aircraft flow, and this period has an average duration of 1 hour and 24 minutes. This confirms that the preliminary analysis of the information provided by SODAR can help predict alterations in wind direction, requiring redirection and bringing advantages in economic and security terms.

The paper is primarily concerned with the modelling of aircraft manufacturing cost. The aim is to establish an integrated life cycle balanced design process through a systems engineering approach to interdisciplinary analysis and control. The cost modelling is achieved using the genetic causal approach that enforces product family categorisation and the subsequent generation of causal relationships between deterministic cost components and their design source. This utilises causal parametric cost drivers and the definition of the physical architecture from the Work Breakdown Structure (WBS) to identify product families. The paper presents applications to the overall aircraft design with a particular focus on the fuselage as a subsystem of the aircraft, including fuselage panels and localised detail, as well as engine nacelles. The higher level application to aircraft requirements and functional analysis is investigated and verified relative to life cycle design issues for the relationship between acquisition cost and Direct Operational Cost (DOC), for a range of both metal and composite subsystems. Maintenance is considered in some detail as an important contributor to DOC and life cycle cost. The lower level application to aircraft physical architecture is investigated and verified for the WBS of an engine nacelle, including a sequential build stage investigation of the materials, fabrication and assembly costs. The studies are then extended by investigating the acquisition cost of aircraft fuselages, including the recurring unit cost and the non-recurring design cost of the airframe sub-system. The systems costing methodology is facilitated by the genetic causal cost modeling technique as the latter is highly generic, interdisciplinary, flexible, multilevel and recursive in nature, and can be applied at the various analysis levels required of systems engineering. Therefore, the main contribution of paper is a methodology for applying systems engineering costing, supported by the genetic causal cost modeling approach, whether at a requirements, functional or physical level.

Mean velocity and first order turbulence measurements were obtained from a three-dimensional upwash fountain flow generated by the impingement of two compressible axisymmetric turbulent jets onto a normal plane. The jet impingement area and fountain formation regions were examined with data obtained through the use of particle image velocimetry. Seven configurations with different nozzle pressure ratios were considered to ascertain the influence of jet compressibility on the fountain development. Results indicate that the mixing of the fountain is dependent on the nozzle pressure ratio, leading to an increase in the fountain spreading rate with increase in nozzle pressure ratio.

The analysis of the aerodynamic environment of the re-entry vehicle attaches great importance to the design of the novel drag reduction strategies, and the combinational spike and jet concept has shown promising application for the drag reduction in supersonic flows. In this paper, the drag force reduction mechanism induced by the combinational spike and lateral jet concept with the freestream Mach number being 5.9332 has been investigated numerically by means of the two-dimensional axisymmetric Navier-Stokes equations coupled with the shear stress transport (SST) k-ω turbulence model, and the effects of the lateral jet location and its number on the drag reduction of the blunt body have been evaluated. The obtained results show that the drag force of the blunt body can be reduced more profoundly when employing the dual lateral jets, and its maximum percentage is 38.81%, with the locations of the first and second lateral jets arranged suitably. The interaction between the leading shock wave and the first lateral jet has a great impact on the drag force reduction. The drag force reduction is more evident when the interaction is stronger. Due to the inclusion of the lateral jet, the pressure intensity at the reattachment point of the blunt body decreases sharply, as well as the temperature near the walls of the spike and the blunt body, and this implies that the multi-lateral jet is beneficial for the drag reduction.

Detailed numerical simulations have been carried out on a spiked blunt body with multiple hemispherical disks using a commercial CFD code in order to investigate their effectiveness in reducing the aerodynamic drag and heating. The base configuration is a hemispherical cylinder whose diameter is 40 mm with an overall length of 70 mm. The lengths of the aerospikes investigated are 1, 1.5, 2 and 2.5 times the base diameter of the cylinder and the radii of the aerodisks are varied between 0.05, 0.1, 0.15 and 0.2 times the diameter of the cylinder. Besides these, the position of the aerodisks is varied with the rearmost aerodisk placed at 25%, 50% and 75% along the length of the aerospike and the intermediate aerodisk for three-disk cases, positioned at 25%, 50% and 75% of the distance between the front and the rearmost disk. All the investigations have carried out at a freestream Mach number of 6.2 and Reynolds number of 2.64 × 107/m. It has been observed that the multidisk spikes are advantageous for the purpose of reduction of both aerodynamic drag and heating at hypersonic speed. The two aerodisk spiked configurations show better results in terms of aerodynamic heating and drag in comparison to the single-disk aerospikes while the three-disk spikes yield only a marginal reduction in aerodynamic drag over the two-disk configurations. For reduction of heat fluxes and heat transfer rates though, the three-disk configurations are extremely advantageous and give much larger reductions are compared to the two-disk configurations.

In previous papers the basis of a matrix approach to the initial buckling or vibration of prismatic structures consisting of a series of thin flat rectangular plates rigidly connected together along their longitudinal edges has been developed. The present paper provides numerical results for the buckling, under uniform longitudinal compression, of a series of panels with unflanged or flanged integral stiffeners, or with Z-section stiffeners; the results are compared with those obtained by other methods and with published results, including an appropriate Royal Aeronautical Society Structures Data Sheet. It was found that considerable inaccuracies can arise from the usual methods of analysis and the authors believe that these justify the use of computer programmes, such as the ones used to obtain the results of this paper, at least at the final design stage.

Effects of vibrational excitation and dissociation of air on inviscid high speed flow past a circular cone, at zero incidence, with an attached shock wave, are studied on the assumption of thermal equilibrium. A numerical solution of the problem is outlined and an approximate analytic solution for the flow between the surface of the cone and the shock wave is developed. Two numerical examples are given as an illustration and compared with the corresponding solutions assuming constant air properties.

This paper describes a study to characterise the impact threat to aircraft from runway debris. This work investigated the likelihood of an event exceeding a critical impact energy; it did not address the probability of an event occurring. Runway debris was collected from six airfields and parameters such as size fraction, mass, density, bluntness and material type were characterised. The distribution of the stone mass against frequency was accurately modelled and based on this, an expression was developed for the probability of an impact exceeding a critical level during a single take-off or landing. By considering the probability of a 50J impact from runway debris as a once in a lifetime event, the implication is that the actual threat is significantly more than the currently perceived threat. This study provides the foundations for more comprehensive studies and will give the designers improved understanding of the threat to aircraft from runway debris.

The discussion generated by the Roskill Commission and by airport planning inquiries in different parts of the United Kingdom has increasingly emphasised that airports are merely a sub-system within a larger system. They cannot be planned and developed except with reference to the other elements that make up that system. There are essentially four critical elements:

1. (i) The passengers or goods requiring transport by air.

2. (ii) The physical, social and economic environment around the airport.

3. (iii) The airport itself.

4. (iv) The aircraft and the airlines which operate them.

The total system is concerned with enabling passengers and freight to travel by air in order to reach a different environment.

The eddy dissipation model (EDM) is analysed with respect to the ability to address the turbulence–combustion interaction process inside hydrogen-fuelled scramjet engines designed to operate at high Mach numbers (≈7–12). The aim is to identify the most appropriate strategy for the use of the model and the calibration of the modelling constants for future design purposes. To this end, three hydrogen-fuelled experimental scramjet configurations with different fuel injection approaches are studied numerically. The first case consists of parallel fuel injection and it is shown that relying on estimates of ignition delay from a 1D kinetics program can greatly improve the effectiveness of the EDM. This was achieved through a proposed zonal approach. The second case considers fuel injection behind a strut. Here the EDM predicts two reacting layers along the domain which is in agreement with experimental temperature profiles close to the point of injection but not the case any more at the downstream end of the test section. The first two scramjet test cases demonstrated that the kinetic limit, which can be applied to the EDM, does not improve the predictions in comparison to experimental data. The last case considered a transverse injection of hydrogen and the EDM approach provided overall good agreement with experimental pressure traces except in the vicinity of the injection location. The EDM appears to be a suitable tool for scramjet combustor analysis incorporating different fuel injection mechanisms with hydrogen. More specifically, the considered test cases demonstrate that the model provides reasonable predictions of pressure, velocity, temperature and composition.

Even before 1939 the problems of high-speed flow were receiving a great deal of attention in the German aeronautical world. Their leading aerodynamicists, Prandtl, Busemann and Schlichting among others, had contributed largely to the theory of compressible fluid flow, and basic experimental research in the high-speed field had begun both at Aachen and at Göttingen. The activity in Germany in this particular field before the war was, in fact, much greater than in Britain, where Mr. Lock and his co-workers at the National Physical Laboratory carried practically the entire responsibility for such work.

The great increase in aircraft speeds during the war, and in particular the arrival of jet propulsion, completely changed our attitude to the subject in Britain. What had been an absorbing fundamental study became, quite suddenly, an important operational problem, looming large in every new fighter design.

The benefit to be obtained by using an interference fit between the pin and plate in a pin-jointed connection has already been established. An examination of the published results shows that some non-linearity occurs in the mechanism of load transference from the pin to the plate since, except at very high initial interference, doubling the load on the joint more than doubles the maximum shear stress in the plate. An examination of the stress-load relationship shows a distinct discontinuity, the load at which this discontinuity occurs being dependent upon both the initial interference and the coefficient of friction between the pin and the plate. It is shown that the results hitherto published correspond to a coefficient of friction between the pin and the plate of 0.3 and results for lower and higher coefficients are given.

Aspects of numerical simulation of aircraft program flight along a predetermined trajectory are presented. An emphasis is laid on the problem of describing the required trajectory by means of a succession of defined points. A spline function method is used for interpolation and smoothing routines. Some numerical illustrations of aircraft trajectory flight simulation (only planar motion) are demonstrated and discussed.