A system of external forces on the structure of a rigid airship can be analysed into two or more of the component systems due to

1. W. The Weights of the members of the structure and the weights of the masses carried by it;

2. L. The Lift and pressure of the gas;

3. A. The Aerodynamic forces consequent on the relative motion of the airship and the air;

4. T. The Thrusts of the airscrews;

5. R. The Reactions due to either the mooring arrangements, the hauling-in guys, the supports used in the shed, or handling on the ground.

As these component systems, though restricted by their interdependence, are individually variable, a very large number of complete systems is represented by the possible combinations and variations that can occur. It becomes necessary, therefore, for the purposes of design, to attempt to limit the number of combinations to be considered, and to restrict the variations of the components, but in such a way that the structure designed to carry the resulting systems with adequate factors of safety, shall be able to carry any other possible system with an appropriate factor of safety.

This paper describes the development of a finite difference method that solves the boundary-layer equations for three-dimensional compressible turbulent flows. The most prominent achievements are the employment of a Newton technique for the simultaneous solution of all governing equations, an option to choose an algebraic or a k-ε eddy-viscosity turbulence model, and the flexible use of curvilinear coordinates. The method is validated by comparisons with a number of experimental and theoretical data sets of three-dimensional, compressible and incompressible, steady and unsteady boundary layers. In parallel, the performance of a three-dimensional compressible industrial integral boundary-layer technique is evaluated by comparisons with experimental test cases and with the results of the field method.

Surface shape optimisation was one of several topics researched under the multi-disciplinary optimisation (MDO) project. Several project partners participated in this activity which set out to develop, validate and demonstrate a number of technical issues related to automated shape design. Each of the participants brought together in-house and/or commercial analysis tools, a variety of different modelling and optimisation techniques, and common database and model generation methods. The objectives of this work were to validate the methodologies established in the preliminary MDO project phase and to demonstrate the optimisation process when applied to a simplified civil aircraft wing design problem. Several issues were central to the investigation, including alternative means of parametric descriptions for the wing surface shape, alternative optimisation strategies for wing MDO, and the impact of different levels of physical and geometrical modelling on MDO performance and resulting product optima. The outcomes of this work demonstrate the need for advanced analysis tools to be incorporated into aerospace vehicle design problems at a relatively early stage and highlight the dangers of over-simplification of the vehicle model.

An experimental investigation was undertaken to determine the effectiveness of distributed surface porosity for the alleviation of pitch-up on a delta wing. Tests were undertaken using a 65° sweep delta wing with distributed porosity evaluated at various locations on the wing. Force balance, on and off surface flow visualisation and flow field surveys using a multi-hole probe were undertaken. The data shows that distributed porosity applied along the wing leading edge at the apex is effective in eliminating pitch-up whilst incurring a minimal performance cost. Trailing edge porosity generally degraded performance.

Surface pressure measurements were obtained during a three-dimensional vortex interaction with a NACA 0015 aerofoil. The upper and lower surfaces of the blade experienced different aerodynamic loads which appear to be controlled by the impact of the vortex axial core flow on the blade surface. On the upper surface of the blade, where the vortex core flow was away from the aerofoil, the interaction was characterised by the generation of a suction peak. On the lower surface, where the axial component was towards the blade, a pressure pulse developed and seemed to be influenced by the vortex approach angle. These features resulted in rapid changes in normal force and quarter chord pitching moment during the interaction. This impulsive loading of the blade may provide some explanation for sound generation and control degradation problems associated with the tail rotor of helicopters.

An experimental investigation of cone cavity flow at M = 8·2 with injection of air and helium has been performed in the Cranfield University gun tunnel. The cone used had a 10° semi-angle and incorporated an annular cavity. The boundary layer ahead of the cavity was laminar. The study concentrated on the effects of increasing the mass injected into the cavity.

With no injection the external flow bridged the cavity. A shock wave and a pressure peak occurred at the rear face corner. By injecting gas, this pressure peak was decreased in intensity and a shock wave was formed at the front face corner. As the mass injected was increased, the shocks initially formed at the front and rear face corners were shifted upstream from their initial positions and boundary layer separation occurred ahead of the cavity. For the highest mass flow rate injected, the pressure peak at the rear face corner was very small. The cavity flow was never more than mildly unsteady.

It is well known that passive vortex generators can be very effective in controlling separation by ‘re-energising’ the low momentum fluids at the boundary layers. They have been used extensively in many practical aerodynamic applications; both in external and internal flows. Typical examples include aerofoil stall alleviation and engine face distortion control in the jet aircraft intake during high angles of incidence. The general flow feature behind a vortex generator is that a pair of contra-rotating streamwise vortices would be formed which will significantly strengthen the flow at the boundary layers. However, the rationale for successful vortex generator designs is often poorly understood. In many cases, vortex generator designs have even been shown to be arbitrary. Anderson et al and Reichert and Wendt used rectangular fin and tapered fin vortex generators respectively, to eliminate the internal flow separation of S-shaped intake ducts. Both geometries were found to be equally effective. Weng and Guo successfully applied aerofoil shape type of vortex generators to suppress the swirl on the engine face of an S-shaped intake duct at high angles of incidence.

A simple technique is proposed that allows the definition of the geometry and characteristics of an air-jet vortex generator to be defined, using the system requirements in practical design. Typically, the aircraft designer is concerned with the mass flow-rate and air pressure requirements of any pneumatic system for inclusion to an airframe. These parameters are not congruent with those for air-jet vortex generator aerodynamic design, and therefore, some tool is required to bridge the gap.

Such a tool is proposed, based on empirical methods for the prediction of air-jet vortex generator behaviour. The technique allows the comparison of the vortex strength, and the system inputs (the jet mass flow-rate and the air-jet plenum pressure) for the air-jet, for a range of jet nozzle diameters and jet velocity ratios. Through this comparison, the optimum air-jet design can be reached for a given system input.

A comparison has been made between the composite beam designs produced by minimum mass optimisation using two different sets of constraints. The first approach constrained the design to have a given separation between fundamental bending and fundamental torsional natural frequencies; the second constrained the design to have a given flutter and divergence speed. The beams are modelled as a series of elements, stepped in thickness at discrete nodes, with the Dynamic Stiffness Method being used for calculation of their natural frequencies. The aeroelastic constraints are obtained from the Fortran program CALFUN. The results show that for similar flutter and divergence speeds, the optima produced using aeroelastic constraints have a slightly lower mass (up to 4% lower) and a less ‘hard’ flutter onset. However, the time taken to produce these optima is significantly longer (in excess of 2 orders of magnitude). A preliminary study discusses the merits of a combined optimisation method where frequency constrained optimisation is used to provide a near-optimum starting point for flutter and divergence constrained optimisation. In addition, a wind-tunnel model of one of the optima has been manufactured and subject to both modal analysis and wind-tunnel tests to validate the flutter speed calculations. This shows that when using strip theory, CALFUN predicts a conservative value of flutter speed for this design. Further investigation has shown CALFUN's lifting surface theory to be more accurate for low aspect ratio models.

The ability of a CFD technique to evaluate the performance of a heat shield separation system has been demonstrated. The evolution of the differential pressure across the heat shield surface has been estimated. The results of the analysis have been compared with the data from the separation test of the heat shield of an Indian launch vehicle. The pressure on the heat shield surface was estimated by solving the Navier Stokes equations using the Eulerian Lagrangian formulation. The results, which show good agreement with the observed results of a heat shield separation test on the ground, were found to be adequate for design purposes.

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.

The performance of German airships during the war, and the double crossing of the Atlantic by the airship R34 in 1919, sufficiently impressed the Air Ministry that it embarked on a British airship programme which produced R38 and R80. The Ministry was undoubtedly deterred by the appalling R38 disaster in 1921 but, presumably as a result of Professor Bairstow’s demonstration of the inexcusable inadequacy of the R38 design, they did not abandon their plans. After all R80, to the design of B.N. Wallis, the great engineer who became famous as Sir Barnes Wallis after the 1939-45 war, had been a success.

So in 1924 the Air Ministry decided to order two airships, each of five million cubic feet capacity. One, to be called R100, was ordered from the Airship Guarantee Company, a Vickers subsidiary headed by Sir Denniston Burney. It was to be built at Howden. The other, to be called R101, was to be built by the Air Ministry itself, at the Royal Airship Works, Cardington. At the head of the Burney design team was B.N. Wallis. At the head of the Air Ministry design team was V.C. Richmond.

The prediction and characterisation of the limit cycle oscillation (LCO) behaviour of non-linear aeroelastic systems has become of great interest recently. However, much of this work has concentrated on determining the existence of LCOs. This paper concentrates on LCO stability. By considering the energy present in different limit cycles, and also using the harmonic balance method, it is shown how the stability of limit cycles can be determined. The analysis is then extended to show that limit cycles can be controlled, or even suppressed, by the use of suitable excitation signals. A basic control scheme is developed to achieve this, and is demonstrated on a simple simulated non-linear aeroelastic system.

Fatigue usage monitoring systems (Fums) offer considerable potential for life extension of aircraft parts. In this work the life extension benefits of Fums is assessed by adopting a probabilistic approach. The roles of damage law type and of service usage variability is explored. It is shown by analysis that in the absence of cycle to cycle load interaction effects, load sequence has no effect on eventual life in either linear or non linear damage laws, provided that the function describing the rate of damage growth has separable variables of stress and damage. This condition includes fracture mechanics crack growth laws. Monte Carlo simulations have been conducted of fatigue life distributions in helicopter rotor components. Variability in manoeuvre damage, when summed over a large number of manoeuvres, has little effect on scatter in overall lives. A fixed manoeuvre usage spectrum will result in very small scatter in lives, whereas keeping the usage constant for each helicopter and allowing it to vary between helicopters, produces a significantly increased variability. The influence of load factors on life is also assessed. The extent of possible maintenance credits is discussed together with the use of Bayesian updating to make optimum use of both prior design data and current loads or damage information provided by Fums.

This paper describes and presents results from static wind tunnel tests conducted on a 60° delta wing at a root chord Reynolds number of 2·7 x 106. In these tests, the wing was instrumented with 192 miniature pressure transducers which, in conjunction with a powerful multi-channel data-logging system, allowed the distribution of time-varying surface pressures to be measured at high temporal resolution. Analysis indicates that the distribution of root mean square pressure on the leeward surface of the wing can provide considerable insight into the behaviour of both the primary and secondary vortex structures. In addition, it has been established that the frequency content of pressure signals measured in the vicinity of these vortex structures is sensitive to the vortex state. It is suggested that these data features can be directly attributed to previously observed behavioral characteristics of the vortex breakdown process.

This paper discusses how the inevitable limitations of computing power available to designers has restricted adoption of optimisation as an essential design tool. It is argued that this situation will continue until optimisation algorithms are developed which utilise the range of available analysis methods in a manner more like human designers. The concept of multi-level algorithms is introduced and a case made for their adoption as the way forward. The issues to be addressed in the development of multi-level algorithms are highlighted.

The paper goes on to discuss a system developed at Southampton University to act as a test bed for multi-level algorithms deployed on a realistic design task. The Southampton University multi-level wing design environment integrates drag estimation algorithms ranging from an empirical code to an Euler CFD code, covering a 150,000 fold difference in computational cost. A simple multi-level optimisation of a civil transport aircraft wing is presented.

This paper presents a short account of the flight control systems used in commercial transport, military combat and general aviation aircraft. The effects of aircraft safety, reliability and weather delays on satisfactory aircraft operations are shown to be significant reasons for the extensive use of flight control systems. The principles of flight control, the sensors and actuators required and the various modes which can be selected are treated, together with a short account of the primary flying controls and the use of manual reversion in emergency situations. The paper concludes with a consideration of the fly-by-wire (FBW) and fly-fby-light concepts, and covers relaxed static stability, carefree manoeuvring and the use of canards before discussing some FBW flight control systems which are used in passenger aircraft.

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.