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ASE 162M - Lab 3

# ASE 162M - Lab 3 - D rew Rosecrans Lab 3 Abstract Pressure...

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Drew Rosecrans Lab 3 Abstract Pressure measurements and schlieren images were taken on a diamond airfoil with a 10° half angle, in flows between Mach 2 and 3, with angles of attack between 0 and 8. The differences from the pressure measurements were greater than expected and can be attributed to the carryover of the error between the two top or bottom surfaces. Also there is certainly experimental error. The schlieren images showed important characteristics of the flow at different Mach numbers and angles of attack. Introduction The diamond airfoil is a very simple airfoil and thus it can be the basis of future studies to compare. It is important to understand how this simple airfoil behaves in supersonic flow at different Mach numbers and different angles of attack. It might be counter intuitive on what is going on in the flow, so it is useful to take images and measurements of the flow to make sure it is known what is happening in it. Also the comparing the measured values to the calculated values will give some insight on how accurate our assumptions can be in the real world. Theory To find the pressure drop over the bow shock on the diamond airfoil, the oblique shock theory is used. This theory depends on the turning angle, theta, that the flow must turn and the Mach number, M, of the freestream. Here is a diagram explaining the oblique shock. The diamond airfoil used here is has a half angle of 10°. Theta is related to the angle of attack of the airfoil. 1

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Drew Rosecrans Lab 3 With theta and the freestream Mach number known, beta, the oblique shock angel, can be found using the theta-beta-M chart. Once beta is known, geometry is used to find the portion of the Mach number that is normal to the oblique shock = M1n M sinβ Now with M1n, the normal shock table can be used to find the Mach number after the shock, M2n, and the static pressure ratio over the shock, P2/P1. Using the static pressure before the shock the pressure after the shock, P2, can be found, which is the pressure at the surface of the airfoil. Now M2 can be found using this equation.
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