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Aviation History
1952
1952 - 0038.PDF
10 FLIGHT, 4 January 1952 INSPIRATION . . . back, but also on the shape of the entry. Fig. 2 showed the results of calculation of the variation of approach loss with angle of sweep- back for entries of various shapes, both when isolated and also when adjacent to a body of fixed length. The results assumed no boundary-layer lip or by-pass but, where necessary, special forms of by-pass could be devised to reduce the loss caused by sweep- back. Against the growing disadvantage of a swept intake, from the viewpoint of internal efficiency, had to be set possible advantage in improved external flow. Whilst on a high subsonic swept-wing design the intake in almost any other position was a liability, the wing-root intake provided, potentially, a useful form of blending between wing and body. The flow round it had something of the character both of that on the isolated swept wing and of that on the three-dimensional body, so that the intake was well-suited to make the transition from one to the other. Referring to the N.A.C.A. submerged intake, the lecturer said that the efficiency obtained was very much the same as that given by a protruding side-intake in the corresponding position, equipped with a conventional by-pass. A problematical point was the rela tive external drag of two such intakes. Whilst the more conven tional layout had to pay something for its efficiency by way of drag in the by-pass duct, it was equally undeniable that the sub merged intake must carry a drag penalty in the form of energy- loss in the vortex motion on the ramp and in the mixing process behind the entry. Dr. Seddon next turned his attention to lip losses in ground running, saying that, in order to get the best performance at high speeds, intake, lips needed to be fairly sharp. As entry lips got thinner and entry area smaller, it became increasingly difficult to meet requirements, and nowadays special measures might be needed. One solution was to provide an auxiliary inlet which functioned only under static and low-speed conditions. On a num ber of plenum-chamber installations this had been done by the simple and convenient method of fitting one or more spring- loaded doors in the wall of the plenum chamber. Slotted Intakes An alternative device suitable for fully-ducted intakes was the slotted intake. In addition to providing auxiliary entry area, the slot worked in a manner analagous to that of a wing leading-edge slot for lift control, the inside of the duct corresponding to the upper surface of a wing. The slot should eject backwards into the duct, and be of sufficient length/width to have good directional control of the air. The use of a slotted intake would enable a designer to employ a smaller main entry and thinner lips, i.e., to match the design more closely to the conditions of high-speed flight. A problem of some consequence arose when the plane of the entry was not normal to the axis of the duct, as occurred with an intake in the leading edge of a swept wing. Under static conditions, the average direction of flow at the entry was more or less normal to the entry plane, so that effectively the air required to be turned through an angle roughly equal to the angle of obliquity of the entry. In model tests of twin-intake systems—as, for example, a pair of wing-root or body-side intakes leading into a common duct or plenum chamber—it had been observed that, if the total duct- flow was reduced below a critical value, the distribution of flow between the two intakes became unsymmetrical. The asymmetry developed rapidly, often to the state where the flow in one duct was actually reversed in direction. This resulted in a highly unsymmetrical velocity distribution at the compressor, a reduced pressure-recovery, and the possibility in flight of oscillating flow resulting in aircraft vibration. The appropriate flight condition would be in a dive at high speed, or on suddenly throttling back the engine while flying at moderate speeds. It was an example of flow instability which was always an inherent possibility in a ducted system having a rising pressure- characteristic (one in which the internal pressure increased with increase of flow). The state of equilibrium tended to be unstable, because any momentary change in the flow altered the pressure n a direction tending to increase the disturbance. A pitot-type intake gave full ram-pressure at zero flow, and a falling characteris tic over the whole range. With a side intake, the pressure at zero flow was below the pitot value, and as the flow was increased, the pressure first rose to a maximum and then fell. Instability was possible in the range between zero flow and the value for which the intake pressure was a maximum. It was desirable with twin intakes that the critical flow ratio should be as low as possible. The value could be calculated by use of the intake loss formula, and the approximate result was: \Vjcrit 2 \AJ where A2 was the area of each duct at the mixing section, and the other notation was as before. This led to two important con clusions: (i) A large amount of diffusion in the duct increased the critical value. For this reason, plenum-chamber intakes were more susceptible than direct inlets, (ii) Since J=k (i — rjt)SlAi the effective position ratio was the principal determining factor. Thus the use of a good boundary layer by-pass in addition to increasing the intake efficiency, would reduce the possibility of flow instability. Dr. Seddon stated that he had discussed the problem purely in relation to a wind-tunnel-model intake system, with free flow, i.e., without engines. This left aside the question of the possible effect of an engine in stabilising the flow. On the subject of drag the lecturer suggested that the intake- drag problem obviously had much in common with the drag pro blem of any other principal component, and of the aircraft as a whole. The emphasis of modern research had clearly to be centred on the effects of compressibility at high subsonic and at supersonic Mach numbers. Low-speed wind-tunnel tests were useful in defining the basic pressure-field in subsonic flow, but theory was not yet powerful enough to predict from this the actual drag at transonic speeds, and the great need was for experimental evidence at the appropriate Mach numbers. At the same time, it was important to keep a check on basic items of sub-critical drag, such as the drag of a boundary-layer by-pass. The drag associated with the boundary-layer by-pass or other corresponding device merited a careful examination. Unfortu nately, the conventional type of by-pass, consisting of a narrow ducted slot separating the main intake from the neighbouring wall, was notoriously difficult to install and to lead away efficiently. Friction loss added up quickly, so the duct needed to be short. This usually meant that it was not too well shaped, and probably discharged back into the external stream at a fairly large angle. It was not possible to generalize on the total cost in drag, but model tests of a few cases had shown that, unless care was taken, the drag might nullify the improvement in thrust obtained from the by-pass. On the score of drag from compressibility effects, Dr. Seddon said that in practice, on a subsonic design, it was not so much the rate of drag-rise which had to be controlled, as the critical Mach number at which the rise began. A good subsonic intake would have a critical Mach number at least as high as that of the aircraft wing. This was not easily achieved, because the local conditions at the intake lip were often more stringent than those on the wing. The lip had to meet the general requirements of fineness-ratio and shape common to all parts, but there was the further complication that, owing to the pre-entry retardation, the lip was effectively at a considerable incidence to the flow. This exaggerated the suction peak which was the criterion for critical Mach number. In the early days of model testing, the process of obtaining the best lip shape was one of hit and miss. Nowadays, a more rational procedure could be adopted owing to the work by Kuchemann at the R.A.E., and by the N.A.C.A. in America. The position, stated very briefly, was that an intake fairing designed for a given critical Mach number required to have (a) a certain minimum thickness, (b) a certain optimum length and (c) good shape within fairly narrow limits. The minimum thickness was that which was neces sary to carry the thrust force on the outside of the fairing without exceeding the value of suction coefficient corresponding to the stated critical Mach number. The criterion of good shape was the "constant velocity profile" first derived bv Ruden. This was calculable in terms of flow ratio and critical Mach number, and corresponded both to minimum thickness and optimum length. It might be said that, taking critical Mach number as the criterion, present knowledge showed how to obtain the best design for a given set of conditions—entry area, maximum fairing area, and entry velocity ratio—so long as these conditions were reasonably normal. When deciding on his final compromise, the designer could take this aspect into consideration in a qualitative way. Dr. Seddon concluded with the observation that, however, much more experimental evidence was needed before it could be determined quantitatively how the drag built up above the critical Mach number and in the supersonic range. AIRCRAFT RECOGNITION SOCIETY THERE was a record attendance at the Aircraft Recognition Society's "Brains Trust" held in the R.Ae.S. Library on December 19th, when the "brains" were Capt. J. Lawrence Pritchard, Mr. Arthur Clarke, Mr. Lankester Parker. Prof. A. A. Hall and Mr. N. E. Rowe. Later, prizes were presented to winning contestants in the Society's Christmas Quiz; among the awards were some fine photographs of the P. 1067, given by the Hawker Siddeley Group, and copies of Pierre Clostermann's book The Big Show, pre sented by the publishers.
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