Numerous military platforms (land, sea and air) feature serial data bus technology based on the US MIL-STD-1553B data bus standard for integration of their digital systems. Many of these platforms have 15-20 years of operational life remaining, but the installed 1553B data buses (data networks) having only a 1Mbit/sec transfer rate are unable to meet many of the future data networking requirements. Research into new, higher performance data networks has concentrated on modern alternatives with throughput increases of two to three orders of magnitude (100Mbit/sec to 1Gbit/sec). These are generally based on modern commercial-off-the-shelf (COTS) standards, good examples being Ethernet and Fibre Channel. Some are already being employed in military platforms having been ruggedised for the harsh physical and electro-magnetic environment. However these COTS systems while being a natural choice for new platforms may not be cost effective for upgrading older platforms. This paper plots the history of MIL-STD-1553, possibly the most successful military platform standard of all time, and discusses some of the options for increasing its performance and economically extending its life into the future.

Aerodynamic performance degradation has been considered a critical hazard to aircraft due to flying in heavy rain. In this work, a two-way momentum coupled Eulerian-Lagrangian approach is developed to study the aerodynamic performance of a two-dimensional (2D) transport-type NACA 64-210 cruise and landing configuration aerofoil as well as a three-dimensional (3D) NACA 64-210 cruise configuration rectangular wing in heavy rain environment. Raindrop impacts, splashback and formed water film are modeled. The steady-state incompressible air flow field and the raindrop trajectory are calculated alternately by incorporating an interphase momentum coupling term through a curvilinear body-fitted grid surrounding the aerofoil/wing. Our simulation results agree well with the experimental results and show significant aerodynamic penalties for all the test cases in heavy rain. Rain-induced premature boundary-layer transition and separation are observed and details of the raindrop splashback effect on the boundary layer are examined. A 1° rain-induced decrease in stall angle-of-attack is predicted. An uneven water film upon the wing surface is observed and its effect on the wing surface roughness is also examined.

This paper presents a novel method for quantifying the effect of ambient, environmental and operating conditions on the progression of degradation in aircraft gas turbines based on the measured engine and environmental parameters. The proposed equivalent operating time (EOT) model considers the degradation modes of fouling, erosion, and blade-tip wear due to creep strain, and expresses the actual degradation rate over the engine clock time relative to a pre-defined reference condition. In this work, the effects of changing environmental and engine operating conditions on the EOT for the core engine booster compressor and the high-pressure turbine were assessed by performance simulation with an engine model. The application to a single and multiple flight scenarios showed that, compared to the actual engine clock time, the EOT provides a clear description of component degradation, prediction of remaining useful life, and sufficient margin for maintenance action to be planned and performed before functional failure.

Since the 1970s, large passenger aircraft design has evolved by modest, but commercially significant, incremental change following the underwing pod concept pioneered by the Boeing 707. In the 21st century, the market is likely to require greater passenger capacity to deal with air-side congestion and higher performance and operational efficiency to justify investment and conserve finite fossil fuel resources. Public opinion will require far greater emphasis on control of noise and engine exhaust pollution. There is likely to be much increased emphasis on pollution in the upper-atmosphere and its environmental impact. In such a situation, it is questionable whether the evolutionary design route can produce the necessary advances and this must stimulate the search for radical design alternatives.

A revolutionary approach, involving the delta planform combined with wing laminar flow control and its impact on overall design, is given a preliminary study in this paper. This has been carried out by assuming that major improvements in drag can be obtained by extensive laminarisation. Using non-dimensional methods, the resulting broad interactive impacts on airframe and engine design and performance are derived. The effect of assuming varying quantities of low-energy air sucked from the foot of the boundary layer is studied and suction-system performance examined.

In addition to the large potential improvements in range, a strong relationship between lower drag, lower cruising altitude and lower cruise engine size is identified. In the study, several factors emerge which may combine to drive large aircraft design towards low aspect ratio and the integrated delta wing planform. This paper is intended as a stimulus and a basis for further study and research.

Gas path diagnostics is one of the most effective condition monitoring techniques in supporting condition-based maintenance of gas turbines and improving availability and reducing maintenance costs of the engines. The techniques can be applied to the health monitoring of different gas path components and also gas path measurement sensors. One of the most important measurement sensors is that for the engine control, also called the power setting sensor, which is used by the engine control system to control the operation of gas turbine engines. In most of the published research so far, it is rarely mentioned that faults in such sensors have been tackled in either engine control or condition monitoring. The reality is that if such a sensor degrades and has a noticeable bias, it will result in a shift in engine operating condition and misleading diagnostic results.

In this paper, the phenomenon of a power-setting sensor fault has been discussed and a gas path diagnostic method based on a Genetic Algorithm (GA) has been proposed for the detection of power-setting sensor fault with and without the existence of engine component degradation and other gas path sensor faults. The developed method has been applied to the diagnostic analysis of a model aero turbofan engine in several case studies. The results show that the GA-based diagnostic method is able to detect and quantify the power-setting sensor fault effectively with the existence of single engine component degradation and single gas path sensor fault. An exceptional situation is that the power-setting sensor fault may not be distinguished from a component fault if both faults have the same fault signature. In addition, the measurement noise has small impact on prediction accuracy. As the GA-based method is computationally slow, it is only recommended for off-line applications. The introduced GA-based diagnostic method is generic so it can be applied to different gas turbine engines.

The theory of the self-excited oscillation of the hinged blades of a helicopter rotor has been treated by Coleman and Feingold and their paper, although written several years ago, remains a standard reference on the subject. Since then, the scope of theoretical studies and calculations has been broadened to include hingeless rotors, more degrees of freedom and blade aerodynamic forces. All this has led inevitably to a great complexity in the equations of motion, with the result that, although the dynamicist can perform adequate calculations, he loses the insight and physical appreciation that he needs, for example, to identify the important parameters in the system. In order to maintain these qualities a different mathematical model is required, one that retains only the most basic parameters and yet still exhibits the fundamental ground resonance instability.

In recent years many methods have been proposed for the measurement of the frequency response of linear systems but the method most often used in the study of structures is the steady state vector response method, the natural frequencies and damping ratios being measured from the resulting vector diagrams. This method is tedious and time consuming and often cannot be readily applied outside the laboratory. Some effort has been directed in the past to employ quasi-steady state methods such as the slow frequency sweep technique but errors are introduced because the response at resonance is less than the steady state maximum and the frequency at which the maximum occurs is shifted in the direction in which the excitation frequency is changing. Additional errors are introduced in the analysis of the system response because of the necessary averaging time of the function analyser used to derive the vector diagrams. Some refinements in technique have been proposed by Reed who developed the “λ Law” frequency sweep in which the percentage change in frequency per cycle is constant. This offers some reduction in test time when compared with logarithmic and linear frequency sweeps but a knowledge of the minimum damping likely to be encountered at any resonant frequency within the sweep range is essential so that the method may be employed advantageously.

Some experiments on the effect of boat-tailing on the pressure distribution round blunt-based aerofoils are described. The experiments were carried out at low speeds at a Reynolds number of 1.5 X 105. The wake was investigated with attention being paid to the vortex shedding, and to the distance downstream of the base where vortices form.

It is shown that the theory due to Nash predicts the effect of boat-tailing on base pressure quite well and that a correlation of drag coefficient, Strouhal number and base pressure proposed by Bearman applies to the models tested here.

The problem of predicting the turbulent supersonic base pressure in axisymmetric flow is treated by an extension of a method of solution to the two-dimensional problem given in Ref. 1. The solution consists principally in tracing the boundary-layer development from upstream of the base to downstream of the recompression region for a given base pressure. A unique solution is obtained by specifying the shape of the rehabilitated boundary-layer velocity profile.

A comparison with experiment in the case of the step-down cylinder problem (the sting-support problem) yields some very favourable results. It is pointed out that, while it has not been found possible to obtain a solution to the problem of a vanishingly small sting, the base pressure does not vary appreciably while the sting is decreased from about 0-3 of the base diameter down to zero. It would appear that the present analysis is capable of giving accurate results down to sting/diameter ratios of the order of 0·3.

This paper presents a strategy for developing the first orbital spaceplane soon and at low cost and risk. The paper then shows how this vehicle will introduce an aviation approach to orbital space transportation to replace the present missile paradigm, leading to far lower costs and improved safety.

To illustrate the potential benefits, the paper presents preliminary sizing and cost estimates of a simple lunar base. Even including the cost of developing the spaceplanes and other vehicles required, the total cost is about ten times less than that of present plans that use large new expendable launch vehicles. Timescales need not be greatly affected.

As part of a programme of research to improve mission effectiveness by studying pilot workload and task performance in mission-oriented flight tasks, a methodology has been developed in which wavelet analysis is used to extract information from records of vehicle response and of pilot control activity. By decomposing the records into discrete wavelets, components of vehicle agility and pilot workload are derived in the form of wavelet-based ‘quickness’ parameters for vehicle agility and so-called ‘attack’ parameters for pilot workload. It is shown how individual wavelet components in the records of pilot control activity, referred to as ‘worklets', can be associated with the sub-tasks of ‘guidance’ and ‘stabilisation'. It is demonstrated how these concepts can be applied to quantify changes in pilot control activity associated with increasing task difficulty or changes in aircraft handling qualities. Two examples are presented, one from a flight trial in which the task difficulty was increased by changes in a prescribed ground track and the other from a simulation trial in which an increased time delay was introduced into the response of the flight control system.

Windtunnel tests on a two-dimensional model of a three element high lift system in a takeoff mode indicate that airjet vortex generators can produce a significant increase in lift at a given angle of incidence and a substantial increase in CLmax. An associated reduction in profile drag is inferred from measurements of momentum defect on the upper surface.

The improvements are much greater than those typically achieved with vane vortex generators and cannot simply be associated with the suppression of boundary layer separation. Instead, the airjets and the vortices which they generate promote enhanced mixing and momentum transfer across the complex shear layers above the main wing, from the external flow right through to the surface. Detailed surveys reveal, for example, that the slat wake is dispersed — or absorbed into a region of greater shear — and that the growth of the main wing boundary layer is significantly reduced.

It is suggested that neither the relatively low Reynolds numbers . and Mach numbers of the tests, nor the effects of windtunnel interference, detract from the general relevance of the novel principle involved. Attention is drawn, however, to the need for further work to elucidate and quantify the flow mechanisms and to establish the balance between the benefits from the improved aerodynamics and the performance costs of installing and supplying the jets.

The flow around a normal flat plate close to a large plane surface has been investigated for a range of boundary-layer thickness on the plane surface of 0.72 to 2.53 times the plate height, h. As the gap between the plate and the plane surface is reduced below about 0.55h vortex shedding from the plate is inhibited by the presence of the plane surface. Boundary layer thickness has little effect on the gap at which this occurs, but does affect the strength and frequency of vortex shedding at greater gaps. As the gap between the plate lower edge and the plane surface is decreased towards the critical value of 0.55h, both vortex strength and Strouhal number are more significantly affected when the approaching boundary layer is thin. It is not clear whether this is a result of shear or increased turbulent intensity.

The General Elastic Analysis of the bending of a bi-metallic strip subjected to uniform heating has been made by Timoshenko. In this note we extend the theory to that of a tri-metal strip. General formulae for curvature and bending stresses have been developed. These formulae reduce to simple expressions if a particular choice of material constants and thicknesses is made. These latter formulae indicate that the curvature can be independent of the coefficient of linear expansion of the middle material and that it is possible to have the stresses at the inter-surfaces equal if a proper choice of the ratios of the coefficients of linear expansion of the three materials is made.

Workers in the field of continuous combustion have long been aware of the importance of the part played by the aerodynamic factors in the satisfactory operation of combustion systems. Although only part of a complex process, the air flow in a combustion system can, to some extent, be studied as an isolated individual factor, without the complications of fuel injection and heat release. This may appear to involve a rather long step from reality, but observations of the basic “cold” flow conditions in such a system can provide a valuable basis for the prediction of the effects of variations in the equipment, particularly in systems conforming to a particular type.

An advanced geared turbofan with year 2035 technology level assumptions was established and used for the hybridisation study in this paper. By boosting the low-speed shaft of the turbofan with electrical power through the accessory gearbox, a parallel hybrid concept was set up. Focusing on the off-design performance of the hybridised gas turbine, electrical power input to the shaft, defined as positive hybridisation in this context, generally moves the compressor operation towards surge. On the other hand, the negative hybridisation, which is to reverse the power flow direction can improve the part-load operations of the turbofan and minimise the use of compressor handling bleeds. For the pre-defined mission given in the paper, negative hybridisation of descent, approach and landing, and taxi operations with 580 kW, 240 kW and 650 kW, respectively was found sufficient to keep a minimum compressor surge margin requirement without handling bleed.

Looking at the hybridisation of key operating points, boosting the cruise operation of the baseline geared turbofan is, however, detrimental to the engine efficiency as it is pushing the cruise operation further away from the energy optimal design point. Without major modifications to the engine design, the benefit of the hybridisation appears primarily at the thermomechanical design point, the hot-day take-off. With the constraint of the turbine blade metal temperature in mind, a 500kW positive hybridisation at hot-day take-off gave cruise specific fuel consumption (SFC) reduction up to 0.5%, mainly because of reduced cooling flow requirement. Through the introduction of typical electrical power system performance characteristics and engine performance exchange rates, a first principles assessment is illustrated. By applying the strategies discussed in the paper, a 3% reduction in block fuel burn can be expected, if a higher power density electrical power system can be achieved.

This paper examines the influence of exit separation, exit availability and seating configuration on aircraft evacuation efficiency and evacuation time. The purpose of this analysis is to explore how these parameters influence the 60-foot exit separation requirement found in aircraft certification rules. The analysis makes use of the airEXODUS evacuation model and is based on a typical wide-body aircraft cabin section involving two pairs of Type-A exits located at either end of the section with a maximum permissible loading of 220 passengers located between the exits. The analysis reveals that there is a complex relationship between exit separation and evacuation efficiency. A main finding of this work is that for the cabin section examined, with a maximum passenger load of 220 and under certification conditions, exit separations up to 170ft will result in approximately constant total evacuation times and average personal evacuation times. This practical exit separation threshold is decreased to 114ft if another combination of exits is selected. While other factors must also be considered when determining maximum allowable exit separations, these results suggest it is not possible to mandate a maximum exit separation without taking into consideration exit type, exit availability and aircraft configuration.

In this paper, a theoretical study of the turn manoeuvre of an agricultural aircraft is presented. The manoeuvre with changeable altitude is analyzed, together with the, effect of the load factors on the turn manoeuvre characteristics during the field-treating flights. The mathematical model used describes the procedure for the correct climb and descent turn manoeuvre. For a typical agricultural aircraft, the numerical results and limitations of the climb, horizontal and descending turn manoeuvre are given. The problem of turning flight with changeable altitude is described by the system of differential equations which describe the influence of the normal and tangential load factors on velocity, the path angle in the vertical plane and the rate of turn, as a function of the bank angle during turning flight. The system of differential equations of motion was solved on a personal computer with the Runge-Kutta-Merson numerical method. Some analytical and numerical results of this calculation are presented in this paper.

The noise radiated from the intakes of turbojet engines during landing is already causing a serious nuisance. This paper, which is largely based on the deliberations of an ad hoc working panel of academic experts formed by Rolls-Royce in 1961, describes the characteristics of the noise and discusses possible mechanisms of generation. The noise appears to be dipole in origin, and to come mainly from the first two compressor stages. It consists of a few discrete tone components, generated mainly by the interactions which occur when one row of aerofoils passes through the wakes of a preceding row; and broad band noise resulting from the turbulence in the wakes and the inherent instability of the boundary layer on the aerofoils.

Some discrete tones may be eliminated over part of the running range by suitable choice of blade numbers and disposition. The broad band noise is unlikely to be reduced without major design modifications involving either a reduction of the air speed relative to the front stage blades, or removal of the boundary layer at their trailing edges.

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.