Wind Tunnel Studies of Leading Edge Separation Phenomena on a Quarter Scale Model of the Outer Panel of the Handley Page "Victor" Wing, with and without Nose Droop

, Leading edge flows, present in the incidence range embracing the onset of general nose separation, have been studied experimentally. The effect of nose droop of these flows was established during the development of this device as a possible alternative to leading-edge flaps on the aircraft. Discrete air jets and vane type vortex generators were investigated in regard to both boundary-layer and hysteresis control. The influence of trailing-edge flaps on leading edge separation phenomena was also studied. The concepts outlined in regard to the nature of nose stalling problems and the quantitative data presented should have application in respect to general wing design. Drooped nose wing with lower surface air jets Drooped nose wing with lower and upper surface air jets Drooped nose wing with trailing edge flaps Flow studies of hysteresis Effect of inboard leading-edge spoiler Effect of pitot probe 5.3 Normal force coefficients 5.4 Pressure distributions 5.5 Minimum pressure coefficients Discussion 6.1 Flow features 6.1.1 Separated vortex flows on swept wings in general 6.1.2 Flow at extreme tip 6.1.3 Inboard origin of separated flow vortex 6.1.4 Hysteresis 6.1.5 Leading edge droop 6.1.6 Trailing edge flaps 6.1.7 Spanwise control of separation 6.1.8 On flow separation Application of test data to assessment of Victor wing modification 6.2 6.3 Design criteria 6.4 Usefulness of air jets Acknowledgment Notation References Tables 1 to 8 1. Introduction The project had its origin in a request from Handley Page Ltd. for an assessment of the probable effectiveness of discrete air jets as a method of boundary-layer control on the 'Victor' bomber aircraft. This was associated with a desire to replace the leading edge flaps with a less complex device while. retaining an equal degree of effectiveness as regards maximum lift and aircraft stability. The assessment (Ref. 1) was encouraging in its findings and as a consequence low-speed wind tunnel tests in the 13 ft. x 9 ft. wind tunnel at N.P.L. were planned. When designing the model consideration was given to previous experience gained by the author in Australia. It was found that satisfactory results relating to nose stalling phenomena could be obtained only on models possessing moderately large chords; for the present tests the minimum chord was fixed at 3 ft. This limitation prohibits the testing of a complete half-span wing and hence the model is restricted to the outer panel of the wing; the panel has a leading edge sweep of 35 …