On unswept wings, or wings with small amounts of sweep, the favourable pressure gradient round the leading edge, where the flow is rapidly accelerated away from the stagnation line, ensures a certain amount of laminar flow, provided the wing surface is sufficiently smooth. On highly swept wings, however, it has been found that turbulent flow can exist on the attachment line itself and there are therefore no naturally occurring regions of laminar flow. This trouble arises from the turbulence at the root of the wing, which sweeps along the attachment line. If the Reynolds number of this turbulent attachment line boundary layer is greater than some critical value, the whole attachment line boundary layer remains turbulent and the complete wing is contaminated. But if the Reynolds number is below the critical value, the turbulence decays along the leading edge and the boundary layer on the attachment line reverts back to the laminar state. This situation arises when the leading edge radius is small and the wing is only slightly swept. The attachment line boundary layer Reynolds number, Rθ, is given by the following equation:

On behalf of my Company, which manufactures the RAE Head-up Display Equipment referred to in K. J. Holden's paper in the December 1964 issue of the Journal (pp 833-6), I would like to correct some misunderstandings which seem to exist and to amplify certain statements in this otherwise interesting and informative contribution to the techniques of instrument displays.

Dr. Holden states: “The essence of the (RAE) display is an attempt to reproduce the appearance of the runway as it would appear to the pilot.” This statement is incorrect as the flight director symbol to which he refers, is made up of a “target spot” (a’) and associated director lines (a) in Fig. 1. This symbol is a pure integrated flight director and is not and never has been intended to represent the runway.

—Measurements of pressure distribution and skin friction have been made on an 11% thick aerofoil in the presence of a variable base bleed flow. These tests show that the effects of bleed decay rapidly upstream of the base and are measurable only over the rear 40% of the chord. The integrated profile drag values also confirm that wake traverse measurements are subject to error in a strong vortex street.

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.