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DROOPING OF LEADING EDGE

  During the time that the tests were being made on the notch and leading edge extension for pitch-up we were following very closely the work being done on the F.102 with regard to a reduction in induced drag by drooping the wing leading edge, and also the work that was going on at Avro Manchester on the Vulcan. They were drooping the leading edge to increase the buffet boundary by preventing leading edge breakaway at high angles of attack.
  This also influenced us in choosing the 10 per cent increase in leading edge outboard, to cure pitch-up, since we realised that, if after investigation we found that it would be advantageous to droop the leading edge, the extension would increase the amount of effective droop.
  Droop was installed on the wind tunnel model, 8deg inboard and approximately 4deg outboard. This increased the buffet boundary considerably. For instance, at M=0.925, which is the normal subsonic cruise Mach number, the CL at which we estimate the onset of buffet is increased from 0.26 with the extension alone, to 0.41 with the extension plus droop. The buffet, or flow separation, was indicated by pressure plots on the ailerons in the Cornell Laboratory tunnel tests. The supersonic drag did not appear to be increased appreciably.
  We were also cognisant at this time of the work on Vortex Generators which Avro Manchester were doing for alleviating the shock-induced rear separation, but from the evidence we had from N.A.C.A., and also from the Manchester reports, it was felt that for a t/c ratio of under 5 per cent this would not be a problem, and Vortex Generators would not be required for this
reason on the CF-105.

ANHEDRAL

  Another peculiarity of the CF-105 wing is the 4deg anhedral. This was established entirely to reduce the length of the undercarriage, and has no appreciable aerodynamic effect or significance.

HIGH WING

  A high wing arrangement was adopted because of the greater flexibility with this layout. For example, it allowed a relatively simple engine installation and any changes in engines and armament can be made without affecting the basic wing structure, which is not always the case with an integrated wing-fuselage structure.
  It also allowed us to carry the wing structure straight through without a break at the fuselage and simplified the wing to fin attachment, since there was no necessity to carry the fin structure down through the engines. The fin is 4 per cent t/c.

AREA RULE

  A great deal of theoretical work was done on the application of Area Rule to the CF-105 and during the early design stages certain changes were incorporated in the aircraft to take advantage of the results of our area rule work.
  Eleven plastic models were made at 1/30th scale and cuts were taken on these to represent various Mach numbers. The cuts were then checked on a planimeter, the results fed into a digital computer, and plots were made around the aircraft at 0deg., 45deg., 90deg., 135deg. and 180deg.. Most of the results were obtained around a Mach number of 1.5 and, as a result of this extensive investigation, we sharpened the radar nose, thinned down the intake lips, reduced the cross-section area of the fuselage below the canopy, and added an extension fairing at the rear, to smooth out the bumps in the area rule curve (Fig. 6).

ENGINE AND INTAKES

  The CF-105 is undoubtedly the most " re-engined" of any aircraft at this stage of development since, one by one the engines slated for the project fell by the wayside. However, I will not attempt to go into the history of the " thousand and one " installations but will deal mainly with the final (?) installation on the production Mark 2. The first five aircraft are fitted with Pratt and Whitney J.75 engines, and the sixth aircraft is the first Mark 2 with Orenda Iroquois engines.
  The Iroquois power unit is an axial flow gas turbine of twin spool configuration. The compressor is designed for a high air mass flow, and a pressure ratio of 8 to 1 at sea level static. Compressor delivery air bleed is used for driving the air turbine centrifugal fuel pumps, and is also available for aircraft services.
  The engine incorporates an afterburner which is built as an integral part of the basic engine. The afterburner operation is fully automatic, the engine having a modulated final nozzle to produce the desired thrust to temperature relationship at the selected power lever setting.
  Figure 7 shows the engine cooling system at speeds greaterthanM=0-5.
  The intake gills immediately adjacent to the compressor inlet open up at M= 0. 5 and allow air to by-pass around the engine for cooling purposes, and to alleviate spillage at high Mach number. By this means, it is possible to achieve near optimum performance with this fixed geometry intake, in the subsonic. transonic, and supersonic speed ranges. At very high Mach numbers, if the air which could not be swallowed by the engines were allowed to spill from the intake lips, there would be a high drag penalty, bad pressure recovery characteristics within the intake itself, and possible de-stabilising effects from the components of spillage.
  The technique of by-passing air over the engine between the engine and compartment sidewalls not only takes care of the spillage and cools the engine but, by acting as a beat exchanger, collects heat from the afterburner casing and passes it into the ejector exit annulus, providing a small percentage of additional thrust.

For fire protection, the critical compartments containing the fuel system, and so on, are enclosed by titanium shrouds and stainless steel insulated blankets.
  The gills are automatic, and when the aircraft has reached a forward speed high enough to create a static pressure higher than ambient within FIGURE 7. Engine cooling. Flight case M=0.50 and upwards

the intake duct, the by-pass gills open due to ram intake pressure, and allow the air to by-pass over the engine.
  With the twin-engine configuration on the Arrow, there has been no requirement for either bifurcated intakes or nozzles, and the flow is relatively clean.

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