1. The problem statement, all variables and given/known data The small scale drawing given in the Annexure shows the general arrangement and dimensional data on an early version of the English Electric “Canberra” B.1 bomber, powered by two Rolls Royce “Avon” jet engines. It was and is an extremely successful aircraft, designed by W E W Petter. It first flew shortly after the end of WW2 and is still in service, well over forty years later. The drawing and the Aircraft Data will give you most of the information you need, but you may have to do some measuring and make some assumptions. If you do this, give your reasoning. You are tasked with estimating the zero lift and lift-dependent drags of the aircraft and to present these in the form of a drag polar:- CD = CDo + k CL 2 Where, k = 1/ ΠAR with CL from O to 1.0, using the given Reference Wing Area for the coefficients. You must also estimate the Mach Number at which there is a steep drag rise. In calculating the zero-lift drag (DO) you should use a Reynolds Number based on flight at 0.7 Mach No. at 11000 m ISA. Full details of the calculation must be given and it is most important that you state clearly what assumptions you have made and give the references to any data you may have used. Your report must be presented as a formal document, neatly typed and properly edited - containing a Summary Sheet, Introduction, Sections, Figures etc., all with page numbers and with a list of references quoted ( and referred to in the text as necessary). You must give full details of your assumptions and calculations, including a. a breakdown showing the DO / q for each item and the total DO / q (Give the corresponding value of CDo), b. State the makeup of the drag due to lift factor k in the expression k CL2 , c. A plot of the drag polar, d. State the value of max lift/drag ratio, and the CL, at which it occurs, e. State what, in your opinion, is the Mach Number at which the drag coefficient begins to rise steeply. Aircraft Data - English Electric “Canberra” B. 1 JET bomber 1. Wing Span, wingtip to wingtip 64 ft. Inner wing, between fuselage & nacelles, Aerofoil section NACA 64012 Wing chord 18 ft. 6 in Outer wing, outboard of nacelles, Aerofoil section NACA 64012 at nacelle side Wing chord 18 ft. 6 in. NACA 64009 at wing tip Wing chord 6 ft. 10 in. Reference Wing Area, used for the coefficients, 960 sq. ft. Wing Aspect Ratio, [Spain]2 ÷ [Wing Ref Area] = 4.267 2. Fuselage Length, nose to tail 65 ft. Max. Diameter (all sections circular) 6 ft. 6 in. Canopy length 6 ft. Canopy frontal area 3.63 sq. ft. Length of forebody 16 ft. 2 in Length of parallel body 16 ft. Length of after body 32 ft. 10 in. Wetted Area, fuselage and canopy 985 sq. ft. 3. Tail Plane and Elevators Span, measured along dihedral plane 28 ft. Chord at centerline (theoretical) 9 ft. 9 in Chord at tail/body intersection 9 ft. 4 in. Chord at theoretical tip 3 ft. 6 in. Dihedral angle 10 degrees Aerofoil section NACA 64009 4. Fin and Rudder Height of fin above fuselage 7 ft. 4 on. Root chord at fuselage 13 ft. 2 in. Tip chord 5 ft. 6 in. Aerofoil section NACA 64010 5. Engine Nacelles Length 23 ft. 2 in Entry diameter 27 in. Intake “highlight” diameter 30. 75 in. Maximum diameter 48. 75 in Position of max. diam. 10 ft. from the intake nose Exhaust pipe diameter 24 in. Wetted area 194 sq. ft. per Nacelle.