— A great deal of publicity has been given in recent years to the problems caused by microbiological attack on aviation fuels, although there had been considerable experience of the subject before the first serious problems arose in aviation. Intensive studies of the nature of the attack have shown that although limited use of biocides has been successful, the best control is by the elimination of stagnant water from the fuel supply system and the aircraft fuel system.

The 837th Lecture to be given before the Royal Aeronautical Society, was delivered at the Institution of Civil Engineers, Great George Street, London, S.W.I, on 6th December 1951. Mr. G. R. Edwards, M.B.E., F.R.Ae.S., Vice-President of the Society, presided and introduced the Lecturer, Dr. J. Seddon, A.F.R.Ae.S., of the Royal Aircraft Establishment.

A recently developed transitional model for boundary-layer flows has been examined on a flat plate and the well-known S809 wind turbine blade. Proposed numerical model tries to simulate streamwise fluctuations, induced by freestream turbulence, in pre-transitional boundary-layer flows by introducing an additional transport equation for laminar kinetic energy term. This new approach can be used for modeling of transitional flows which are exposed to both the freestream turbulence intensity and streamwise pressure gradient, which are known as the most dominant factors in occurrence of transition. Computational method of this model is based on the solution of the Reynolds averaged Navier-Stokes (RANS) equations and the eddy-viscosity concept. The model includes three transport equations of laminar kinetic energy, turbulent kinetic energy and dissipation rate frequency. The present model is capable of predicting either natural or bypass transitional mechanisms, which may occur in attached boundary-layer flows. In addition, the model can simulate transition in the separated free shear layers and the subsequent turbulent re-attachment to form a laminar separation bubble. Flat plate was exposed to different freestream turbulence intensities and streamwise pressure gradients. Wind turbine blade was examined under two different Reynolds numbers, with one of them suitable for the occurrence of laminar separation bubbles on its surfaces. To evaluate the performance of this new model in resolving transitional boundary-layer flows, final results have been compared to those obtained through application of conventional turbulence models. Comparison of final results for the flat plate and the S809 aerofoil with available experimental data show very close agreements.

Tonight we commemorate the achievements of Juan de la Cierva, the great pioneer of rotary aviation. His development of the autogyro stimulated the development of the early rotor theories and paved the way for the modern helicopter. In turning the autogyro into a practical aircraft, Cierva proved to be an innovator prepared to take an unconventional approach. He was at the same time pragmatic in that he really wanted to build a helicopter but realised that the required technology was not yet mature, and settled on the autogyro as representing what was an achievable concept. His concept was thus unconventional, dramatic and achievable — an admirable combination. However, in his first attempts to fly an autogyro he rapidly came up against the problem that faces all rotory wing designers — which is that the advancing blade has a much greater lifting potential than the retreating blade, due to the difference in dynamic head. This led him to introduce the flapping hinge that allowed the rotor to achieve roll balance as well as providing the backward tilt of the rotor that is necessary for autorotation. Having made this step forward, he realised that the introduction of blade camber would allow the retreating blade to achieve higher angles of incidence before stalling, but in so doing he built in a potentially dangerous control characteristic.

The friction force in many physical systems has been found to depend on a power of the velocity, the magnitude of which varies according to the damping mechanism. This note employs an approximate theory applicable to small damping to examine the free vibration of a single degree of freedom system in which the damping force is proportional to an unspecified power of the velocity.

Correlation with experimental data on the internal damping of solid materials shows that this phenomenon also tends to obey a power law although discrepancies exist between the expected and observed behaviours. These arise from the interpretation of the damping coefficient and from the frequency dependence of the dissipated energy.

The scramjet is a rather a new technology and there are many issues related to their operation, especially when it comes to the combustion processes. Combustion in high-speed flows causes various problems such as flame instability and poor fuel–air mixing efficiency. One of the methods used to overcome these problems is to recess a cavity in the combustor wall where a secondary flow is generated. In this study, a computational fluid dynamics (CFD) code is developed to analyse the reacting flow passing through the cavity-based scramjet combustor. The developed code is based on three-dimensional coupled Navier–Stokes and finite rate chemistry equations. An ethylene-air reduced chemical reaction model is used as a fuel–air combination. The Spalart–Allmaras model is utilised for turbulence closure. The non-dimensional form of the flow and chemical reaction equations are discretised using a finite volume method. The Jacobian-Free Newton–Krylov (JFNK) method is used to solve the coupled system of non-linear equations. The JFNK is a matrix-free solution method which improves the computational cost of Newton’s method. The parameters that affect the performance of the JFNK method are studied in the analysis of a scramjet combustor. The influence of the forcing term on the convergence of the JFNK method is studied in the analysis of scramjet combustor. Different upwind flux vector splitting methods are utilised. Various flux limiter techniques are employed for the calculations of higher order flux vectors. The effects of flux vector splitting and flux limiter methods on the convergence and accuracy of the JFNK method are evaluated. Moreover, the variations of the mixing efficiency with fuel injection angles are studied.

Comprehensive understanding of propulsive jet aerodynamics and aeroacoustics is key to engine design for reduced jet noise and infra-red signature in civil and military aerospace, respectively. Illustrated examples are provided of other aerodynamic/aeroacoustic problems involving jet development, including chevron nozzles, increased jet/wing/flap interference (as fan diameter increases), high acoustic environment (and potentially damaging screech) of supersonic jets on carrier decks and the strongly Three-Dimensional (3D) unsteady flow during the in-ground effect operation of Short Take-Off and Vertical Landing (STOVL) aircraft. To date, laboratory/rig test measurements have primarily been used to identify design solutions; increased use of Computational Fluid Dynamics (CFD) would help achieve cost/time reductions, but Reynolds-Average Navier–Stokes (RANS) CFD with statistical turbulence modelling has proven inadequate for such flows. The scenarios described are far removed from flows used to calibrate model constants, and predictive accuracy demands detailed insight into unsteady flow. Large-Eddy Simulation (LES), whilst computationally more demanding, offers a potential solution. Research undertaken to assess LES capability to address the challenges described is reviewed here. This demonstrates that tremendous progress has been made, indicating that LES can provide sufficiently accurate predictions, representing high value for engineering design. A series of validation studies of increasing realism to practical engineering systems is presented to underpin this conclusion. Finally, areas for further work are suggested to support the combined application of RANS and LES that is probably the optimum way forward.

This paper presents a performance analysis on a novel engine concept, currently under development, in order to achieve hybrid air-breathing rocket technology. A component-level approach has been developed to simulate the performance of the engine at Mach 5, and the thermodynamic interaction of the different working fluids has been analysed. The bypass ramjet duct has also been included in the model. This facilitates the improved evaluation of performance parameters. The impact of ram drag induced by the intake of the engine has also been demonstrated. The whole model is introduced into a multi-platform application for aeroengine simulation to make it accessible to the interested reader. Results show that the bypass duct modelling increases the overall efficiency by approximately 7%. The model calculates the specific impulse at approximately 1800 seconds, which is 4 times higher than any chemical rocket.

We review the development of the contra-rotating propellers from the origins to the present. Initially, these systems were proposed to increase speed, then to increase propulsive efficiency, and thus reduce fuel burn. Ultimately, they hit another environmental limit: too much noise. Acoustics has been in fact the main focus of the development in the past 30 years. Pioneering work done across countries demonstrated several unique features of this propulsor. Various embodiments are available, namely contra-rotating, counter-rotating, co-axial, tandem, open rotor and prop-fans, collectively named contra-rotating propellers. This review only considers concepts that have been applied to real aircraft, prototypes that are known to have been flight tested (about 70 vehicles), or representative laboratory models. Five classifications are proposed: pioneers (before 1940), golden years (1940–1950), Western airplanes (1950s onwards), Soviet-Russian airplanes (1950s onwards) and modern developments (1980s onwards). Selected experimental aircraft and laboratory concepts are mentioned, where these appear to advance the state-of-the-art. Power plants evolved from internal combustion engines to the modern gas turbine engines requiring new solutions. Engine layouts and propulsion configurations are analysed where appropriate. It is concluded that propulsive efficiency can only be achieved at a cost of multiple engineering problems, some of which remain unsolved.

A generalised conical field theory is developed and is applied to delta wings in a non-uniform stream. It is shown that a non-uniform stream may be characterised by the downwash at all points in space. The lift of a delta wing is found when the downwash in the wing plane is given as a power series in the co-ordinates in the wing plane. The basis of the conical field theory is described in some detail but the results only of the calculation of the lift distribution for various down washes are given. The solutions of certain integral equations, required in the calculations, are given in the Appendix.

Wind-tunnel experiments were conducted to determine the interference characteristics of an inclined turbulent jet exhausting from a flat plate into a turbulent subsonic crosswind. The jet was not yawed with respect to the undisturbed free stream. The suction force (lift loss), the jet penetration into the crosswind, the jet deflection and the rate of total pressure decay along the jet centre line were all reduced by an increase in the inclination of the jet. The centre of pressure moved downstream. These results were attributed to a decrease in the entrainment rate of the jet as the inclination increased in a downstream direction.

Tests have been carried out on a series of convergent-divergent nozzles and the results compared with those from a test on a convergent nozzle. At the design pressure ratio, reductions in the noise were found, particularly at 90° to the jet direction. For lower forward pressures the noise is also less, but above the design pressure the noise rises rapidly to the same level as that with a convergent nozzle. In the experiments cold jets were used, but the results are probably applicable in certain respects to hot jets.

In this paper the Janzen-Rayleigh method is used to calculate the velocity potential for the steady subsonic flow of a compressible, inviscid fluid past a prolate spheroid. The fluid velocity at a point on the body is calculated. The analytic form obtained for this velocity differs, from that giving the velocity which an incompressible fluid would possess at the same point on the body, by a correction factor. The factor is an infinite series of first derivatives of Legendre functions of the first kind and odd order. The first three coefficients in this series are computed for bodies of certain axis ratios, and graphs of the values of these coefficients against axis ratio are plotted. The behaviour of the nth coefficient for large values of n is given. Results for slender ellipsoids, considering these as a limiting case of the family of ellipsoids just referred to, are obtained and are found to agree with the usual slender-body theory. Using these an attempt is made to continue the graphs of the first three coefficients in the correction factor series for the whole range of axis ratios of the ellipsoids in the system, namely zero to unity. The results obtained for the bluff-nosed ellipsoids may be used to estimate the effects of compressibility on the pressure distribution over the front of a general bluff-nosed body in steady flow.

The CHAIRMAN, in opening the meeting, said that Mr Zimmerman, a member of the scientific staff of the National Advisory Committee for Aeronautics, at Langley Field, Virginia, had long been concerned with the development of S T O L aircraft, having originated and developed the Chance-Vought circular-wing project during the war at Stratford, Connecticut In the rotary wing field, he was perhaps best known as the originator of the Flying Platform, having designed and built at Nicholls, Connecticut, shortly after the war, his so-called “Aerial Motor-cycle,” which had two side-by-side lifting propellers of disc loading about 20, interconnected by a tiny platform which he strapped to his feet like skis By 1947, he was able to demonstrate to his satisfaction that this device required no controls other than a throttle for the two engines, and that he could hover safely by applying the same reactions to his feet as he used for balancing himself when standing on the ground

A method of calculating turbulent boundary layers on infinite yawed wings is given, making use of a method of calculating turbulent boundary layers due to Spence and of an analogy between three-dimensional and axi-symmetric boundary layers. It is also shown that the displacement thickness is equal to that computed using chordwise components and that the streamwise momentum thickness is approximately equal to the chordwise momentum thickness. Shock-free flow and small boundary layer cross-flow are assumed.

Design success mostly depends on understanding and realising product requirements, which are nested with different levels of stakeholders and specific market demands. Emerging markets are volatile and non-linear for understanding user-specific needs, particularly related to aircraft design. We attempt herein to elicit specific requirements through a proposed requirements elicitation process based on functional, physical and behavioural viewpoints. An evaluation of those critical requirements is also performed through Quality Function Deployment (QFD) to reveal the fitness of the customer requirements (raw requirements) in relation to the technical measures. After selecting all projected potential users, a Systems Modelling Language (SysML) use-case diagram is also summarised to illustrate the user’s relation to the system functions. Some raw requirements are verified through constraint analysis and cost examination as an exemplary approach to present the transformation from raw to final requirements. The data are collected and analysed to construct the aircraft’s key specifications. Most importantly, how to identify good requirements that define the needs, attainability, clarity and verifiability is demonstrated.

In addition, it is demonstrated through this study that there is a need of a specific type of general aviation aircraft that may fulfil the unique demands of emerging economies by satisfying their socio-economic conditions.

The ‘double integral’ method, employed by Squire and Trouncer to calculate the flow of a round turbulent jet in a moving stream, is generalised. Representation of the velocity profile in simple form and use of an arbitrary upper limit of integration in the second application of the momentum integral equation permits the recovery of the well-known similarity solutions for plane and axi-symmetric laminar jets issuing into a quiescent medium. Additionally, an approximation for the decay of centre-line velocity is obtained for the non-similar situation when there is an ambient co-flowing stream. This agrees well with Wygnanski’s perturbation solution for the plane laminar jet. The results of Squire and Trouncer are re-examined in the light of this generalised approach. The work has application to the operation of fluidic sensors.

Similar solutions of an equation governing the flow in the plane of symmetry of a corner boundary layer with favourable pressure gradient are extended to a “critical” value of the pressure gradient for which no solution could be found previously. It is shown that the failure to achieve this result in the past was due to the “singular” nature of this solution rather than to its non-existence as one was tempted to suspect. The existence and uniqueness of this crucial solution are demonstrated and the solution itself is obtained to a high degree of numerical accuracy. Criticism of the theory on which this corner boundary-layer equation is based is discussed in an Appendix.

The combustion instability characteristics in a model dump combustor with an exhaust nozzle were experimentally investigated. The first objective was to understand the effects of operating conditions and geometric conditions on combustion instability. The second objective was to examine more generalised parameters that affect the onset of combustion instability. Three different premixed gases consisting of air and hydrocarbon fuels (C2H4, C2H6, C3H8) were burnt in the dump combustor at various inlet velocity, equivalence ratio and combustion chamber length. Dynamic pressure transducer and photomultiplier tube with a bandpass filter were used to measure pressure fluctuation and CH* chemiluminescence data. Peak frequencies and their maximum power spectral densities of pressure fluctuations at same equivalence ratios showed different trends for each fuel. However, the dynamic combustion characteristics of pressure fluctuations displayed consistent results under similar characteristics chemistry times regardless of the used hydrocarbon fuels. The results showed that characteristic chemistry time and characteristic convection time influenced combustion instabilities. It was found that the convective-acoustic combustion instability could be prevented by increasing the characteristic chemistry time and characteristic convection time.

Quasi-one-dimensional (quasi-1D) tools developed for capturing flow and acoustic dynamics in non-segmented solid rocket motors are evaluated using multi-dimensional computational fluid dynamic simulations and used to characterise damping of modal perturbations. For motors with high length-to-diameter ratios (of the order of 10), remarkably accurate estimates of frequencies and damping rates of lower modes can be obtained using the the quasi-1D approximation. Various grain configurations are considered to study the effect of internal geometry on damping rates. Analysis shows that lower cross-sectional area at the nozzle entry plane is found to increase damping rates of all the modes. The flow-turning loss for a mode increases if the more mass addition due to combustion is added at pressure nodes. For the fundamental mode, this loss is, therefore, maximum if burning area is maximum at the centre. The insights from this study in addition to recommendations made by Blomshield(1) based on combustion considerations would be very helpful in realizing rocket motors free from combustion instability.