1. The problem statement, all variables and given/known data Air flows through a symmetric converging-diverging nozzle where the cross-sectional area of the nozzle (in meters^2) varies accoding to the relationship: A(x) = 1.0 - 0.8X +.80X^(2) ; where X is in meters and the nozzle is 1 meter long. P2/P1 = .90 (there is subsonic inflow) A normal shock also occurs near the exit. A) Determine the inflow and outflow mach numbers B) The stagnation pressure ratio across the normal shock C)The X location of the shock (in meter) from the inflow to the normal shock. 2. Relevant equations I have been using the isentropic and normal shock tables to solve this problem which involve ratios of pressures, areas, and temperature. 3. The attempt at a solution I assume Mach number at the throat will be 1 Mth (throat) = 1 L =1 Area of the throat: X=0.5 A(x) = 1.0-0.8(.5) + 0.8(.5^(2)) = 0.8 Throat area: 0.8 = A* Area of the inlet = 1 = A1 A1/A* = 1.25 From this I find the Mach number at the inlet equal to .5533 M1 = .5533 I honestly need help with only finding the Mach numbers at the inlet and exit. I'm a little confused due to the equation for finding the area given. I could really really use some help. Thanks!