Linear elastic buckling strength assessments are presented for a range of thin plate arrays with optimal strength to weight ratio configurations subject to a range of arbitrary in-plane stress states. The outcomes of the assessment are presented as design curves demonstrating relative buckling strength increases with respect to the classical square plate datum of equal mass.

A stiffness matrix method is adopted for the initial buckling strength predictions. The method is based on exact flat plate theory and assumes that the plate is continuous over supports, whereby deformations in one cell of the plate array influence the deformations in adjacent cells. The supporting webs of each cell are approximated in this study by simple rigid supports that enforce nodal lines or lines of zero out-of-plane displacement in the buckled panel, i.e. the bending and torsional stiffness provided by the supporting webs of the real structure are ignored. Selected results are validated by FEM prediction.

A heat transfer measurement technique has been developed, which utilised a laser to heat a spot of the ‘standard’ luminescent paint on an insulated metal wind tunnel model. The convective heat transfer coefficient was determined from the experimental quasi steady-state surface temperature, and solutions obtained from radial and axial conduction in a numerical heat transfer model. The convective heat transfer coefficient variation over both a flat plate and a NACA 0012 aerofoil have been measured in transonic flow. Measurements obtained from the flat plate were seen to agree well with correlation data from the literature. Measurements on the NACA 0012 aerofoil indicated the point of transition from laminar to turbulent boundary layer as well as the location of shock boundary layer interaction. The luminescent paint provided simultaneous measurements of pressure and temperature (see Part 1). The distribution of pressure over the NACA 0012 aerofoil was shown to be in excellent agreement with conventional transducer data, although the luminescent paint data provided greater spatial resolution. The position of the shock determined from the heat transfer measurements was shown to be in excellent agreement with the pressure measurements.

Although The Aeronautical Journal goes back a long way its main purpose is to look forward by publishing the latest research, design and development ideas for a knowledgeable and critical audience. Launched in January 1897, well before the Wright brothers’ record-making flight, the Journal has now reached Volume 107, number 1072. In this special issue to celebrate the 100th anniversary of the Wright brothers’ extraordinary achievement, it was decided to ask an international group of experts to look ahead to the next 100 years (or at least as far as they thought they could see). Hence in November 2001 the then RAeS President, Professor Ian Poll, wrote to a number of eminent people in the world of aerospace, inviting them to contribute. The response was enthusiastic and the fruits of their labours are on the following pages.

Transverse sonic injection, usually in a staged manner, in a confined environment is a necessity in the design of an efficient combustion chamber. The design requires an analysis of the mixing of the injectant with incoming stream as efficient combustion depends upon good mixing. This kind of analysis needs numerical model/simulation to assess the mixing of the two streams. To determine the suitability of an existing software package for such a study, staged transverse sonic injection into a Mach 2 stream in a confined environment is taken as a test case. The experimental conditions of McDaniel et al are reproduced for this simulation. In this experiment, staged transverse injection of sonic jet behind the backward facing step in Mach 2 stream was carried out and profiles of various flow parameters were measured. The numerical simulation solves the 3D Navier-Stokes equations with a k – ε turbulence model using the PARallel Aerodynamic Simulator PARAS3D. Computed results show a good match for injectant penetration profile although the computations predict slightly higher penetration near the orifice location. Detailed comparison of flow parameter profiles between computed and experimental data reveal that in the zone away from the injection orifice, computations predict the flow field reasonable well. However, in the vicinity of the orifice, there are some differences between experimental data and computed results. These differences could be due to non-uniform inlet profile in the experiment and/or inadequacy of the turbulence model considered in the study.

The evaluation of rotorcraft performance and handling qualities in piloted flight using a desk-top simulation is a desirable facility. The development of the SYCOS pilot model has brought this capability closer by simulating piloted flight along prescribed trajectories such as the Mission Task Elements of ADS-33. This development has overcome some of the limitations of the precise control of inverse simulation, by adopting a structure which includes a corrective component that adjusts the control strategy to counteract departures from the desired flight path. The development and implementation of SYCOS for a range of rotorcraft types and applications is shown to be well defined and straightforward. The capability and reliability of SYCOS is demonstrated through a number of recent applications in performance, control strategy and pilot workload studies. Some of the applications benefit from enhancements to the basic structure and these are described in context. The accumulation of experience with the SYCOS pilot indicates that it is a significant advance as a predictor of human control activity during helicopter manoeuvring flight.

This paper deals with the methods of physical modelling of Flight Control Systems (FCS) by means of Dynamically Similar free-flying Models (DSM) for the investigation of stability and controllability of aircraft at subsonic flight speeds. The subsonic flight regime allows us to avoid Mach number similarity considerations. The large scale of the DSM meets autosimilarity of Reynolds numbers, whilst Froude similarity is assured during the development and manufacturing of the DSM. This paper proves the presence of necessary and sufficient conditions of similarity for the FCS of an aircraft, and that of its DSM. The existence of necessary conditions have been proved both mathematically (by means of the π – Theorem from the theory of dimensions), and with equations involving physical quantities. The generalised scale coefficients for transitioning from the FCS’s aircraft gain factors, to those of the FCS of the DSM have been obtained, and it is shown that the coefficients depend only on the linear scale of the DSM.

A dynamic backpropagation training algorithm for an adaptive fuzzy gain scheduling feedback control scheme with the application to missile autopilots is developed. This novel design methodology uses a Takagi-Sugeno fuzzy system to represent the fuzzy relationship between the scheduling variables and controller parameters. Mach number and angle-of-attack are used as measured, time-varying exogenous scheduling variables injected into the control law. By incorporating scheduling parameter variation information, the adaptation law for controller parameters is derived. Results from extensive simulation studies show that the presented approach offers satisfactory controlled system performance.

The results of an investigation into the effect of both the number of orifices and the geometry of these orifices on the fluid structures exiting from a synthetic jet actuator are presented. The experiments have used smoke and laser visualisation and single hotwire measurement techniques to establish the flow field exiting the various orifices and the fluid velocities in the exit plane of the orifice. The results indicate that the use of two small orifices can perhaps produce two coherent vortex rings with a total circulation greater than that found in the ring produced by a large, single orifice of equivalent area. However, they have also shown that there are three different types of interaction between two adjacent orifices, related to the spacing between them and the level of excitation applied to the diaphragm. In addition, these interactions may affect the overall level of circulation in different ways. Finally the results have shown that a rectangular orifice of a given exit plane area is more likely to produce a turbulent ring than the equivalent circular orifice due to a ‘smoothing’ process which endeavours to create an axisymmetric ring from a non-axisymmetric slug of fluid. This work has shown that when using synthetic jet actuators in flow control applications, care should be taken in selecting the spacing and geometry of the orifices used, as certain set-ups may be more suited to certain applications than others.

The present study demonstrates that cyclically varying the damping coefficient of controllable lag and flap dampers can reduce the 4/rev vibratory hub loads of a four-bladed hingeless rotor helicopter in high speed forward flight. Gradient-based optimization is used to determine the optimal multi-cyclic damping variation inputs that minimise a composite vibration index comprising of all six components of vibratory hub loads. Optimal 2/rev and 3/rev variations in the lag damping coefficient virtually eliminate the vibratory hub drag force and yawing moments, and produce small reductions in the vibratory hub side force. The optimal lag damping variations, interestingly, produce increases in the 3/rev and 5/rev components of the blade root drag shear, that cancel the contributions of the blade root radial shear to the vibratory in-plane hub forces. Despite some increases in higher harmonics of blade response, damper loads, and blade and flexbeam root loads, the lower harmonics and the peak-to-peak values show little change, implying that blade and damper fatigue life would not be adversely affected. When optimal 2/rev and 3/rev variations in flap damping coefficient are introduced in conjunction with the optimal lag damping variations, 30% reductions in the hub vertical vibrations are obtained, in addition to the previous reductions in the vibratory in-plane forces and yawing moment. The cyclic flap damping variations reduce the higher harmonics of the blade root vertical shear. Reductions in hub vibration levels are obtained over a range of forward flight speeds.

Passive imaging using mm-waves offers very significant advantages in scientific and military surveillance. However, the relatively long wavelengths mean that for the resolutions that are typically sought, the input aperture of the imager needs to be quite large; typically in excess of one metre. Deployment of conventional dish antennae of these dimensions on small aircraft is highly problematic. The use of snapshot synthetic aperture interferometric radiometry (SAIR) offers an attractive route to integrating a two-dimensional antenna array into the structure of an aircraft so that the transverse effective dimensions of the antenna can be almost as large as the aircraft. We report here a study into the feasibility of deployment of a SAIR on unmanned airborne vehicles and the achievable performance parameters. Critical considerations are the achievement of acceptable sensitivity and angular resolution from a SAIR that does not require excessive complexity. It is shown that traditional approaches based on fully sampling the spatial frequencies in the scene are unable to simultaneously meet all of these criteria, but that a SAIR based on thinned sampling of the spatial frequencies shows promise.

Aeronautics faces new and increased challenges in the 21st century. These include climate change, long-term fuel supply, traffic congestion, and noise pollution. Radical technological steps will be demanded of the industry in order to maintain a sustainable transportation system.

In setting out its ideas for a Sonic Cruiser, Boeing has recently acknowledged that there is a need to look beyond the high-aspect-ratio wing with separate fuselage and podded engine configuration, which it pioneered at the start of the Jet era.

This paper sets out the case for the broad delta including its advantages as a basis for achieving higher transonic cruising speeds, with very long range and large payloads, while providing progressive and significantly improved environmental impact. The work is based on concept designs drawing on flight and weight data from the Vulcan bomber and the Concorde aircraft – both pioneering designs, by no means at the end of their technical potential.

A broad delta is taken to mean an aircraft of aspect ratio in the range 2 to 4. Not only does such a planform have application to many sizes of airliner, cruising at various high subsonic speeds, but also it is inherently well suited to later application of laminar flow. Wing surfaces undisturbed by flaps and slots and unhampered by very thick wing skins, together with favourable engine inlet locations, could be a basis for improved overall aerodynamic efficiency.

Airliners designed for operation in the region of Mach 1 are likely to pose more difficult problems for the aerodynamicist. Designs must cater for both cruise and maximum design speeds, which may encompass the Mach 1 condition.

Very Large Transport Aircraft (VLTA) pose considerable challenges to designers, operators and certification authorities. Questions concerning seating arrangement, nature and design of recreational space, the number, design and location of internal staircases, the number of cabin crew required and the nature of the cabin crew emergency procedures are just some of the issues that need to be addressed. Other more radical concepts such as blended wing body (BWB) design, involving one or two decks with possibly four or more aisles offer even greater challenges. Can the largest exits currently available cope with passenger flow arising from four or five aisles? Do we need to consider new concepts in exit design? Should the main aisles be made wider to accommodate more passengers? In this paper we demonstrate how computer based evacuation models can be used to investigate these issues through examination of staircase evacuation procedures for VLTA and aisle/exit configuration for BWB cabin layouts.

There is increased interest in new classes of mini- and micro-UAVs with sizes ranging from one metre to ten centimetres. Many envisioned applications of such UAVs require them to be able to fly close to the ground in complex environments. The difficulties associated with flying in such environments coupled with the reduced payload capacity of such airframes means that new methods of sensing and control need to be considered. Good models for such methods are found in the world of flying insects. One particular visual cue used by insects is optic flow, which is the apparent visual motion seen by the insect as a result of its motion through the environment. This paper discusses several research efforts aimed at developing new sensing and control algorithms inspired by insect vision and flight behaviors. These efforts are part of DARPA's controlled biological and biomimetic systems (CBBS) programme. In these (and related) efforts, many elegant control stratagems have been discovered which suggest that simple reflexive schemes combined with the measurement of optic flow may be sufficient to provide many aspects of autonomous navigation in complex environments. Furthermore, these efforts are implementing these behaviors in real flying UAV platforms by using novel hardware and software to measure optic flow, and inserting optic flow measurements into a control loop using a combination of ‘best engineering approaches’ with inspiration taken from biology. This has resulted in fixed and rotary-wing mini-UAVs that are able to hold an altitude and perform terrain following.

The use of solar sail propulsion is investigated for both Mercury orbiter (MO) and Mercury sample return missions (MeSR). It will be demonstrated that solar sail propulsion can significantly reduce launch mass and enhance payload mass fractions for MO missions, while MeSR missions are enabled, again with a relatively low launch mass. Previous investigations of MeSR type missions using solar electric propulsion have identified a requirement for an Ariane V launcher to deliver a lander and sample return vehicle. The analysis presented in this paper demonstrates that, in principle, a MeSR mission can be enabled using a single Soyuz-Fregat launch vehicle, leading to significant reductions in launch mass and mission costs. Similarly, it will be demonstrated that the full payload of the ESA Bepi Colombo orbiter mission can be delivered to Mercury using a Soyuz-Fregat launch vehicle, rather than Ariane V, again leading to a reduction in mission costs.

The method of matched asymptotic expansions has been applied to analyse the asymptotic structure of the incompressible turbulent attachment-line boundary layer at large Reynolds numbers. Inner and outer solutions were obtained that give the ‘wall-unit’ and ‘velocity-defect’ represention of both the spanwise and chordwise velocity profiles. The predictions from the theory agree with experimental and DNS data in the range of fully developed turbulence.

Heat transfer levels have been investigated behind a rearward-facing step in a superorbital expansion tube. The heat transfer was measured along a flat plate and behind 2 and 3mm steps with the same length to step height ratio. Results were obtained with air as the test gas at speeds of 6·76kms-1 and 9·60kms-1 corresponding to stagnation enthalpies of 26MJ/kg and 48MJ/kg respectively. A laminar boundary layer was established on the flat plate and measured heat transfer levels were consistent with classical empirical correlations. In the case of flow behind a step, the measurements showed a gradual rise in heat transfer from the rear of the step to a plateau several step heights downstream for both flow conditions. Reattachment distance was estimated to be approximately 1·6 step heights downstream of the 2mm step at the low enthalpy condition through the use of flow visualisation.

This paper describes an optimisation study concerning arrival trajectories that has been conducted using a recently developed tool for the analysis and design of noise abatement procedures around airports. This new tool combines a noise model, a geographic information system, and a dynamic trajectory optimisation algorithm. The optimisation algorithm generates routings and flight-paths that minimise the noise impact in the residential communities surrounding the airport, while satisfying all imposed operational and safety constraints. The study on arrival trajectories presented herein complements an earlier study involving departure trajectories. Although the numerical results shown pertain to a particular example airport, viz Amsterdam Airport Schiphol, the study actually focuses on the development of a generic methodology that could be applied to any given airport. The results clearly demonstrate the effectiveness and flexibility of the developed tool.

The subsonic doublet lattice method (DLM) has been the industry standard method for the calculation of unsteady air loads for the past three decades. A recent comparison between the DLM and the subsonic constant pressure panel code ZONA6 suggested that the DLM lacked robustness. However, the DLM code used in the comparison was flawed and not representative of the DLM as such. Results from a contemporary DLM code are presented for the same test cases that were used in the misleading comparison. These new results show that the DLM can produce valid results for all the test cases considered. Where feasible, a comparison is made between results from the present DLM code and the published ZONA6 results.

The dye-injection flow visualisation technique was used to investigate the effect of the apex flap on the leading-edge vortex breakdown over a cropped 76°/40° double delta wing. The angle-of-attack of the experimental model varied from 20° to 40°, and the length of the apex flap was 25%c and 50%c respectively. By changing the angle of the apex flap, we found that the apex flap is an efficient method to control the leading-edge vortex breakdown, and that there exists an angle of the apex flap at which the value of the leading-edge vortex breakdown delay reaches maximum. Moreover, it is found that, for α < 28°, the small apex flap is a better choice for delaying the vortex breakdown; for α > 28°, the large apex flap is superior to the small one.