a) Develop a relation between local static pressure P and freestream static pressure P.. Assume the stagnation pressure remains unchanged (i.e.,isentropic flow). b) Write the local pressure coefficient C, in terms of free stream Mach number M, and the ratio P/P . c) Combining your results from a) and b), write an expression for the local pressure coefficient C, in terms of local and free stream Mach numbers. d) If the peak C, in incompressible flow is -0.43, estimate the critical Mach number. Hint: place all terms on one side of the equation and use a trial and error approach.

Evaluation and Discussion 8. Task A: A1: A2: A3: A4: A5: A4: A5: Why pressure distribution on the upper and lower surface are the same for NACA0015 airfoil at zero AOA? Why there is a difference in pressure distribution between the upper and lower surface for cases you simulated in this Task? Which sides (upper or lower surface) has higher pressure, and why? Do you see the difference in pressure (between upper and lower surface) changes with AOA? Why? Describe the pressure distribution on the upper surface by identifying the stagnation point, suction peak, and adverse pressure region. What is Cp? How do we compute the lift coefficient of the aerofoil from the Cp distribution? (15 marks)

Question 1. Calculate the temperature (in Kelvin) at an altitude of 9665 metres above sea level, assuming ISA atmosphere: O -47.82 O 0.0065 O 350.82 O 12 O 225.18 Not answered Question 2. Calculate the temperature (in Kelvin) at an altitude of 11471 metres above sea level, assuming ISA atmosphere: 0.0065 216.50 O 3 O 362.56 O-59.56 Not answered SESSMEN sity The Questions Question 3. Calculate the pressure (in Pascals) at an altitude of 5192 metres above sea level, assuming ISA atmosphere: O -19 O 52611 O 90423 O 53 101325 [a2-alkarbi] UFMFRU-15-1 DEWIS E-ASSE Not answered The Questions Question 4. Calculate the air density (in kg per cubic metre) at an altitude of 12504 metres above sea level, assuming ISA atmosphere: O 81.64 O 0.29 O 0.36 O 288.00 O-80.99 [a2-alk UFMFRU-15-1 DEWIS E-AS Not answered The Questions Question 5. An aircraft is flying at an altitude of 8364 m above sea level. Its airspeed with respect to the surrounding air is 137 m/s. Assuming ISA conditions, calculate the dynamic pressure (in Pascals). O 1965.1 O 288 -39.37 4716.52 O-928.7 OFMPRU-15-1 DEWIS E-ASSESSMEN Not answered

Question 5: An aircraft climbs at an angle of 15°. The weight of the aircraft is 114 kN and its wing area is 68 m². The aircraft's drag equation is given by CD 0.035 +0.025 C. If the engines of the aircraft produce 55 kN of thrust during the climb what is the fastest equivalent airspeed the aircraft could be climbing at? The answer should have units of m s 1. =

Question 1 2 pts Question 1: In a wind tunnel test of flow over a wing section, air enters the test section at a speed of 54 m s1 and a pressure of 942 kPa. Assuming Bernoulli's principle applies to the flow, what is the pressure at the stagnation point on the wing surface if the air density is 1.16 kg m 3? The answer should have units of kPa.

t Figure 8.4 shows a drag polar plot of a glider, with C₁/CD and C2/C3 superimposed as func- ons of the lift coefficient. For a given weight, W, wing area, S and altitude, h (this determines the density), by selecting a range of values for angle of attack, it is possible to compute: C₁/CD, C/C, V and RD. The corre- sponding performance data are shown in Table 8.1. The glider selected for this example is a relatively poor one by modern standards. Its lowest flight path angles are high compared to those achievable with modern gliders. Trimmed lift-to-drag ratios of 40.0 to 50.0 have already been achieved, yielding descent angles as low as 1.2 degrees! Despite this, the reader is asked to verify that the approximations of Eqns (8.22) - (8.24) are quite good! The performance results of Table 8.1 are plotted in Figure 8.5. The reader is asked to check out the significance of the points labeled A, B and C in Figures 8.4 and 8.5. These very important points are discussed next.

Part 1: Assume the A320neo is currently flying with an airspeed (TAS) of 250 kn at an altitude of 5000 ft and is climbing. The throttle factor is set at F, = 0.85. The mass of the aircraft is 78.5 tonnes. Determine: a) the climb angle and rate of climb. b) the steepest climb angle and maximum rate of climb. c) compare the current airspeed of 250 kn with those required to achieve the steepest climb angle and maximum rate of climb.

As "task1.m" script runs from "P2amain.m" script, it must output the following two plots (no user inputs required): (i) The "stagnation streamline" plot (the shape of the semi-Rankine body: r/R) in the range of: 09⁰, 179° (upper half surface) and = 181°, 191°, 201°,... 351° (lower half surface). 19°, 29°,. (ii) The "pressure coefficient" plot (the surface pressure distribution over the semi-Rankine body: C₂) in the range of: 0=9°, 19°, 29°,... 179° (upper half surface) and = 181°, 191°, 201°,... 351° (lower half surface). (Note) the C, plot in aerodynamics is "negative values in positive y coordinate direction" (rule of thumb). This will require you to modify the y axis data direction ("flip-flop" of the y coordinate) when you perform C, plot in MATLAB.

For a particular airfoil section the pitching moment coefficient about an axis 1/3 chordbehind the leading edge varies with the lift coefficient in the following manner: a) Find the acrodynamic center and the value of Cmac ·b) Find the relation between the position of the center of pressure and the lift coefficient;plot a curve showing the variation of the position of center of pressure with liſtcoefficient.

10. Task C: C1: C2: C3: Based on your simulation result, identify regions where boundary layer is laminar, turbulent and transition. How do you identify transition point and from you simulation result discuss how they are affected by AOA and Re. How to you identify separation point and discuss how they are affected by AOA and Re. (20 marks)