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In aerodynamics, the normal shock tables are a series of tabulated data listing the various properties before and after the occurrence of a normal shock wave. With a given upstream Mach number, the post-shock Mach number can be calculated along with the pressure, density, temperature, and stagnation pressure ratios. Such tables are useful since the equations used to calculate the properties after a normal shock are cumbersome.
The tables below have been calculated using a heat capacity ratio, γ {\displaystyle \gamma } , equal to 1.4. The upstream Mach number, M 1 {\displaystyle M_{1}} , begins at 1 and ends at 5. Although the tables could be extended over any range of Mach numbers, stopping at Mach 5 is typical since assuming γ {\displaystyle \gamma } to be 1.4 over the entire Mach number range leads to errors over 10% beyond Mach 5.