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.

Two-Dimensional Stress concentration factors may be obtained more quickly and simply than the corresponding three-dimensional factors, either by experiment or mathematical analysis. It would be convenient to obtain information, for varying geometry in the two-dimensional case of a particular type of stress raiser, e.g. a shoulder, groove or hole, and use this either to predict the three-dimensional stress concentration factors or to extend the range of existing three-dimensional results. Clearly a comparison is only possible if the three-dimensional stress raiser embodies a plane of symmetry (which gives the geometry of the similar two-dimensional stress raiser), and if the loading conditions can be reproduced in both the two- and three-dimensional cases. The latter requirement restricts the correlation to the stress concentration factors obtained in tension and in bending. The three-dimensional torsional loading system has no plane of symmetry which can be simulated in two dimensions.

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.

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.

A rigid, 55° sweep, half delta wing has been oscillated in pitch at subsonic speeds, and the unsteady pressures on both the upper and lower surfaces recorded for pre-stalled conditions. The oscillations were of low amplitude and at frequencies equivalent to a typical wing first bending mode for full scale applications.

When compared to a quasi-steady approximation, the unsteady pressures on the upper surface of the wing lag the steady case along the line of the primary attachment. The lag represents a constant convective time from the leading-edge with increasing frequency of oscillation. A further localised area of lagging flow is observed beneath the vortex burst point, the location of which is a function of mean angle of attack.

The magnitude of the unsteady pressure variations was seen to increase linearly with the amplitude of the pitching oscillation while the phase lag was seen to increase linearly with frequency parameter.

Tables of the characteristic functions of certain types of uniform beams were published by Young and Felgar in 1949. The functions concerned represent (to suitable scales) the principal modes of free flexural vibration, although their uses are far more extensive than is implied by this. Some of the uses to which they can be put are mentioned in Ref. 1 and others are discussed by Felgarand Bishop;they cover both statical and dynamical problems and relate to beams, plates and shells and include problems of elastic instability. In fact these functions can be used in series representations in the same way as circular functions are used in Fourier series.

One of the main uses to which these characteristic functions may be put does not appear to have been described in print, although the relevant formulae are quoted in Ref. 2. The functions can be used in vibration analysis to shorten calculations, sometimes leading to very considerable savings of labour.

Measurements have been made of the surface pressure distributions in the region of the junction between an untapered wing, of NACA 0015 section swept back at 20°, and a flat plate on which a turbulent boundary layer had developed, for several values of wing incidence. It is shown that the section lift coefficient of the wing diminishes, while the pressure drag coefficient increases, as the junction is approached. Oil-flow visualisations on the plate surface show the passage of the horseshoe-like vortex which forms when the retarded boundary layer flow separates as it approaches the leading-edge of the junction. Kinks in the isobars on the plate correlate with the trailing “legs” of this vortex. The surface flow visualisations also show that the turbulence in the junction region spreads onto the wing from the leading edge at an angle of about 10°.

A rotatable X-wire anemometer was used to make measurements of the mean velocity field and of five components of the Reynolds stress tensor, in the wake of the junction, with the wing at incidences of 0° and 9°. Log-law (Clauser) plots were used with the profiles Ū(Y) of longitudinal velocity to estimate the skin friction coefficient on the plate, though adjustments of the zero for Y were necessary to obtain a sensible fit. These corrections were often larger than can be readily explained, but the skin-friction values are consistent with the corresponding, measured velocity correlation, . The Reynolds stresses in the wake region clearly show that the horseshoe-like vortex legs persist beyond the trailing-edge of the wing, the turbulence intensity being larger on the suction side for the wing at incidence.

Average values of the skin-friction coefficient on the plate in the junction region are little different from those away from the junction. No corresponding information is available for the skin-friction on the wing, but if this is likewise unaffected by the junction, then the total drag of the junction region will be greater than the sum of those of the isolated parts, simply because of the increase in pressure drag.

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.

As part of the ongoing development of small unmanned air systems by the University of Southampton, an all laser sintered aircraft has been test flown from the Royal Navy's ice patrol ship HMS Protector to assist with navigating through the Antarctic. These flights were carried out with pre-planned autopilot control with oversight from Andrew Lock, acting as the pilot embarked on HMS Protector. This is the first time the Royal Navy has used unmanned aerial vehicles in this part of the world. In this paper, we set out the trial reports and lessons learnt from this series of test flights.

Recent experiments have demonstrated that humans can fly, with great exertion of the leg muscles using fixed wings and a propeller. However, understanding of aerodynamics and physiology show that using existing technology it is not possible for humans to fly on Earth using their arm muscles to flap wings like birds. Human flapping flight might, however, be possible in a partial gravity environment such as a chamber in Earth orbit or on the surface of the Moon. This paper considers the feasibility of this and provides preliminary estimates of the main parameters involved.

Throughout the years that rotorcraft have been on the British Register of Civil Aircraft, the single reciprocating engined, single main rotor helicopter has been the predominant representative of this particular family of aircraft. These helicopters were certificated to the British Civil Airworthiness Requirements (BCAR), Section G, “Rotorcraft,” issued in January 1954, which was written in terms of such helicopters.

During the past few years there have been significant changes in rotorcraft design—the turbine engine has been introduced and the multi-engined type has become a reality.

Basic changes in the design trends of any type of aircraft must inevitably involve a re-appraisal of the appropriate requirements by the responsible airworthiness authority. The BCAR are subjected to a constant process of refinement and revision and, as a result of the development of new types, the Air Registration Board has been considering the airworthiness implications and preparing amendments to most parts of Section G, but I wish to discuss only Flight Requirements which are in the main contained in one Sub-Section of Section G—Sub-Section G2. From the recent revisions I have extracted only the significant changes, although there are still problem areas I would like to put on record.

The interaction of a helicopter tail rotor blade with the tip vortex system from the main rotor is a significant source of noise and, in some flight states, can produce marked reductions in control effectiveness. This paper describes a series of wind-tunnel tests to simulate tail rotor blade vortex interaction with a view to providing data for the development and validation of numerical simulations of the phenomenon. In the experiments, which were carried out in the Argyll wind-tunnel of Glasgow University, a single-bladed rotor located in the tunnel’s contraction was used to generate the tip vortex which travelled downstream into the working section where it interacted with a model tail rotor. The tail rotor was instrumented with miniature pressure transducers that measured the aerodynamic response during the interaction. The results suggest that the rotor blade vortex interaction is similar in form to that measured at much higher spatial resolution on a fixed, non-rotating blade. The combination of the two datasets, therefore, provides a valuable resource for the development and validation of predictive schemes.

Non-axisymmetric endwalls in turbine stages have shown to be a robust method to improve the performance of turbines in both power generation and aero-derivative applications. Non-axisymmetric endwalls target the control of secondary flows and are designed using detailed computational fluid dynamics coupled with a variety of optimisation algorithms and utilising a number of objective functions according to the engine company or researcher's preference. These numerical predictions are often backed up by detailed measurements in linear and annular cascades and later proven in full-scale engine tests. Relatively little literature is available describing their performance in rotating test rigs or at conditions other than design, apart from that of the authors. This study comprehensively revisits the low-speed, model turbines used in the earlier study, replacing all of the 5-hole probe data with more accurate results and additional hot-film measurements. These results together with computational fluid dynamics solutions are used to show the success of the method across a large incidence range and to compare to earlier cascade results for a similar endwall and blade profile to establish the usefulness of cascade testing in this application. In addition, a comparison to two other off-design studies is made. Results indicate that the endwalls successfully improve the rotor total isentropic efficiency at all test conditions and that the improvement increases with increased turning in the blade row, from 0.5% to 1.8% across the incidence range. The results also compare well to the estimation of isentropic efficiency improvement that can be drawn from the cascade testing which stands at 1.55%.

The computer can be used in many ways to assist in the project and design stages of aircraft development. As would be expected it was used first to assist in the more unwieldy calculations of the design stage. Massive calculation sequences which previously involved much tedious labour, or were effectively beyond the capacity of a design office to perform, suddenly became a few hours work for one or two technicians. Project analysis, being concerned with an earlier stage of development where calculations are simpler and more approximate, tended to remain largely unaffected. However, it became clear that the very simplicity of the preliminary project calculations would, at least in principle, allow the entire project analysis phase to be computerised, so allowing many parametric variations to be done in a short time with minimum effort. The natural follow-on to this was the optimising project analysis program of which examples now exist. With such programs one is able to predict an optimum configuration to meet a given requirement, given only the basic data to define the scope of the calculation.

For the past 150 years man's ingenuity has been exercised in producing machines to transport himself faster and faster until today he is talking of aircraft which can travel faster than the rotating speed of the earth at the equator, and such that any point on the earth's surface is within ten hours flight time from any other point. But such is the contrariness of the situation that these speeds involve the use of extensive runways away from the crowded city centres and, what with the crowded roads and trains, the journey from the airport to the city centre can be long, tedious and uncomfortable.

The helicopter enthusiast suggests that this problem can be overcome by using a machine that can take off vertically from the city centre and carry the passenger rapidly in comfort to the airport. He also suggests that for journeys up to 200 miles or more, the helicopter is the quickest form of transport from city centre to city centre.

It was a great revelation to me to read Lieut.-Comdr. Graham's study of the “Safety Devices on Wings of Birds,” reprinted in the Journal of the Aeronautical Society of January, 1932. I had always believed that the gaps formed by the tip feathers on a bird's wing were not strictly comparable to slots, because from a superficial observation of birds’ wings, it appeared that the slots were running in the wrong direction, namely, from the top to the lower surface of the wing. Lieut.-Comdr. Graham's outstanding contribution is his discovery of the reason for the emargination of the web of the slot forming wing tip feathers. There can be no doubt that this arrangement is aerodynamically equivalent to a multi-slotted wing.

The first Sir Henry Royce Memorial Lecture described the work of a brilliant engineer, who was directly responsible for the design of a long series of great engines. As a contrast, I have been asked to tell you about the work, not of an individual, but of a team. A team, moreover, that has had no responsibility for the final product, but whose task has been that of attempting to clear the ground and make the engine designers’ work less difficult.

The year 1957 has a certain significance in gas turbine history, in that it is the twentieth anniversary of several important early events. It was in March 1937 that the Engine Sub-Committee of the Aeronautical Research Committee discussed the internal combustion turbine and as a result recommended “that the Air Ministry should take up the question of the development of the internal combustion turbine as a matter of urgency and make all possible arrangements for its production at the earliest possible moment …”

Pressure Distribution and liquid-film tests on a 10 per cent, thick aerofoil revealed the presence of separation “bubbles” close to the leading edge. These bubbles are formed beneath the boundary layer which separates near the leading edge and re-attaches farther aft; their existence is usually indicated by localised constant-pressure regions in the pressure distributions. It is also believed that if a thin film of liquid (such as a suspension of lamp-black in paraffin) is spread on the surface, the scrubbing action of the air rotating in the bubble will tend to draw liquid forward into the bubble, and hence the location and extent of the bubble may be indicated approximately by the accumulation of the fluid.

Many boundary layer traverses of bubbles on N.A.C.A. aerofoils have been made, but it was felt that similar measurements of the bubbles on this particular aerofoil would provide useful data, since the separation characteristics of this section appeared to differ from those in the N.A.C.A. tests.

This paper intends to develop a target trajectory estimation algorithm with application to the ballistic target estimation in the terminal phase. The proposed design is based on the application of a second-order extended state observer (ESO) technique using target information acquired from the seeker to estimate the trajectory of manoeuverable ballistic targets. Satisfactory results have been received while applying the design in estimation of either two-dimenional or three-dimentional target evasive acceleration via computer simulation.