The problem of finding the temperature and velocity distribution inside the boundary layer of a compressible flow over a flat plate is a well-known problem which has been discussed by many distinguished investigators in the past five decades.

Excellent reviews of the work done on this classical hydrodynamic problem are available, for example, in the monographs of Schlichting, Loitsyanskii and Stewartson.

Bragg, in ref. 1, gives a theoretical treatment for the variation of the discharge coefficient in compressible flow of any shaped nozzle with applied pressure ratio given one set of values. This shows that the discharge coefficient of any nozzle with a value less than unity at a pressure ratio approaching unity, rises continuously as the pressure ratio increases and reaches the maximum at a pressure ratio of infinity (zero back pressure). This is based on the assumption that the flow conditions at the Vena contracta are uniform and at unity Mach number once the applied pressure ratio is equal to/ or above the critical (sonic) value.

The following extremely important table, compiled by Major Linton Hope, was inadvertently omitted from the reprint of his lecture in the August issue of the Journal. It should also be noted that the last set of Flying Boat lines, on page 470, is the design which Major Linton Hope made in 1916, after the trials of the 41ft. design, Fig. 2, and to which he refers in his reply to Mr. Baker. It should be noted that the angles of the step tangents to the foreplaning bottom have been increased and the back step reduced in size and moved aft, both of which features appear in C.E.I. designed some considerable time afterwards.

The effect of backlash and trailing-edge strips on the flutter speed of a two-dimensional model of a tailplane with tab has been investigated in a series of wind tunnel tests. When backlash is present, a change in preload, which is achieved by varying the incidence of the tab, can dramatically alter the flutter speed. The addition of trailing-edge strips to the tab lowered the flutter speed.

The Aerodrome, as the point of contact between the aircraft and the world of reality, has to a large extent been taken for granted. The designer regards it as the base for prototype tests and the airline operator as his aerial railway station. It is my intention to show that aerodrome development is at a critical stage of evolution on this side of the Atlantic, that many of the related problems do concern the aircraft designer and operator, and that if these problems are not given consideration, difficulties will be encountered in civil operations.

To quote an example; the optimum aircraft type to serve a system of air routes can be worked out taking into consideration the range required, payload and speed, but if the aircraft ultimately ordered and produced cannot safely operate from the aerodromes to be served because of the runway length and approaches, then the operations will either have to be reorganised or continued without due regard for safety.

A Fire Prevention Sub-Committee of the Aeronautical Research Council published its findings in 1922. At that time the major risk was from crash fires in which the fuel was ignited by the engine either directly or indirectly through the medium of oil, wood, fabric or dry grass. Preventive measures recommended included crash-proof tanks, with a suggestion that these should be carried on the wings; fire-proof bulkheads; fire-proofing of inflammable materials; cooling of exhaust pipes; and fire-proofing of magnetos. Fires in flight were usually in the carburettor and a successful carbon tetrachloride extinguisher was designed at the Royal Aircraft Establishment and tested in flight by the late Squadron Leader Norman. However, the weight of the system, about 35 lb. per engine, was considered excessive and the action recommended was to turn off the fuel and dive the aeroplane.

This paper describes the construction of a simple electronic simulator for the solution of flutter problems in two degrees of freedom. It was intended as a prototype for a much larger machine to solve problems in six degrees of freedom.

Details of the construction and circuits are given, together with some typical solutions obtained on the machine. As a result of the successful operation of this prototype a larger machine in six degrees of freedom has now been built.

Many authors have established that a finite wing in inviscid flow has a theoretical maximum lift coefficient. One formula based on a theoretical model of a rolled-up vortex system aft of an elliptically loaded wing is

where AR is the aspect ratio. Ref. 6 criticises the validity of eqn. (1) but at least eqn. (1) gives the order of magnitude of an inviscid CLmax.

In the technical note in the August JOURNAL by Morgan and Saunders the values given in the last column of Table I on p. 572 are actually the square roots of the open area ratio.

The 969th Lecture to be given before the Society and the 22nd Main Lecture to be held at a Branch Centre, “ Free Flight Techniques for High Speed Aerodynamic Research ” by J. A. Hamilton, M.B.E., B.Sc, A.M.I.Mech.E., A.F.R.Ae.S., and P. A. Hufton, M.Sc, was given under the auspices of the Boscombe Down Branch on 1st December 1955. Air Commodore R. A. Ramsay Rae, O.B.E., President of the Branch welcomed the guests and said how delighted he was to see the President of the Society, Mr. N. E. Rowe and members of the Council and Main Society at the Aeroplane and Armament Experimental Establishment.

The well-known Routh-Hurwitz stability criteria apply to polynomial characteristic equations with real coefficients. Lesser known criteria were devised by Hurwitz for the case of polynomials with complex coefficients. However, when polynomials with real coefficients and even powers only are directly subjected to the first-mentioned criteria, all the test determinants obtained vanish identically and thus fail to indicate the state of stability. In the present paper, the stability test functions for such characteristic polynomials are derived.

Vortex shedding is an important factor in the design of structural elements subject to flow induced vibrations. Several reviews have appeared on this topic. However, in model testing, the vortex shedding frequency of a bluff body may change when the model mounting is eccentric (Fig. 1). This has not been studied in detail. Available information in this regard is restricted to the data of Shaw whose tests were conducted in a towing tank.

This technical note presents experimental data on the vortex shedding characteristics of equiangular prisms (Fig. 2) mounted eccentrically in a wind tunnel. Explanations for the possible flow behaviour are also presented.

Many complex machines and mechanical systems have an ideal and abstract form which has been evolved from fundamental theory without reference to those engineering principles that are so essential in practice. Thus the heat engine has its Carnot cycle and the propeller its actuator disc. With these, and numerous other examples, the object is to provide a standard of reference, to give an ideal which can be approached but never reached and, most important, to indicate the fundamental limitations imposed by theory and prevent profitless investigations.

Experimental results are presented of the measurement of the ratio of the wall shear stresses at the inner and outer surfaces of concentric annuli. Five radius ratios were investigated with Reynolds numbers in the range 2000-89 000 with air. The Reynolds number is defined as where ū is the bulk velocity.

It is concluded that the ratio of the shear stresses is very different from the corresponding laminar flow value and is a function of both radius ratio and Reynolds number.

The design of structures for operation at increasingly higher temperatures and under exposure to severe heating requires more detailed and more exact analysis to determine the most severe conditions under which a given structure is operationally effective. To meet this need, the following analysis has been devised, with the determination of stresses and deflections in plates due to non-uniform heating as its purpose. The analysis results also in the determination of the maximum loads which may be imposed on the supporting members by the plate due to its thermal deformations.

In using the method of stage stacking to compute the off-design performance of multi-stage axial compressors, it has been observed that the limitation on performance at speeds above the design speed has been set by the stall and the choke points of the rear stages(1). Thus if the rear stages can absorb a wide range of mass flows between stalled conditions and choked conditions, a better performance could be obtained.

Compressor stages using low stagger blades will absorb a large range of mass flow between stalled and choked condition; but because of the high axial velocity involved in their use, they tend to be unsuitable for low pressure stages because of the high Mach number obtained. In the higher pressure stages the increased gas temperature will lower the Mach number for the same velocity and give more efficient operation.