Page 323 - Aeronautical Engineer Data Book

Appendix 5:
Shock wave data

Table A5.1 Normal shock wave data
Pressure, Mach number and temperature changes through
shock waves ( = 7/5).
Notation:
M 1 = Mach number of flow upstream of shock wave
M 2 = Mach number of flow behind the shock wave
= Prandtl–Meyer angle, (deg), for expanding flow at M 1
µ = Mach angle, (deg), (sin(–1)(1/M 1 ))
P 2 /P 1 = Static pressure ratio across normal shock wave
d 2 /d 1 = Density ratio across normal shock wave
T 2 /T 1 = Temperature ratio across normal shock wave
P o2 /P o1 = Stagnation pressure ratio across normal shock
wave

M 1 M 2 P 2 /P 1 d 2 /d 1 T 2 /T 1 P o2 /P o1
1.00 0.000 90.000 1.0000 1.000 1.0000 1.0000 1.0000
1.01 0.045 81.931 0.9901 1.023 1.0167 1.0066 1.0000
1.02 0.126 78.635 0.9805 1.047 1.0334 1.0132 1.0000
1.03 0.229 76.138 0.9712 1.071 1.0502 1.0198 1.0000
1.04 0.351 74.058 0.9620 1.095 1.0671 1.0263 0.9999
1.05 0.487 72.247 0.9531 1.120 1.0840 1.0328 0.9999
1.06 0.637 70.630 0.9444 1.144 1.1009 1.0393 0.9998
1.07 0.797 69.160 0.9360 1.169 1.1179 1.0458 0.9996
1.08 0.968 67.808 0.9277 1.194 1.1349 1.0522 0.9994
1.09 1.148 66.553 0.9196 1.219 1.1520 1.0586 0.9992
1.10 1.336 65.380 0.9118 1.245 1.1691 1.0649 0.9989
1.11 1.532 64.277 0.9041 1.271 1.1862 1.0713 0.9986
1.12 1.735 63.234 0.8966 1.297 1.2034 1.0776 0.9982
1.13 1.944 62.246 0.8892 1.323 1.2206 1.0840 0.9978
1.14 2.160 61.306 0.8820 1.350 1.2378 1.0903 0.9973
1.15 2.381 60.408 0.8750 1.376 1.2550 1.0966 0.9967
1.16 2.607 59.550 0.8682 1.403 1.2723 1.1029 0.9961
1.17 2.839 58.727 0.8615 1.430 1.2896 1.1092 0.9953
1.18 3.074 57.936 0.8549 1.458 1.3069 1.1154 0.9946
1.19 3.314 57.176 0.8485 1.485 1.3243 1.1217 0.9937
1.20 3.558 56.443 0.8422 1.513 1.3416 1.1280 0.9928
1.21 3.806 55.735 0.8360 1.541 1.3590 1.1343 0.9918
1.22 4.057 55.052 0.8300 1.570 1.3764 1.1405 0.9907
1.23 4.312 54.391 0.8241 1.598 1.3938 1.1468 0.9896
1.24 4.569 53.751 0.8183 1.627 1.4112 1.1531 0.9884
1.25 4.830 53.130 0.8126 1.656 1.4286 1.1594 0.9871