For inviscid, compressible flow past a thin airfoil, shown in the figure, free-stream Mach number and pressure are denoted by $\boldsymbol{M}_{\infty}$ and $\boldsymbol{p}_{\infty}$ respectively. Ratio of pressure at point $\mathbf{A}$ and $\boldsymbol{p}_{\infty}$ is 0.8 and specific heat ratio is 1.4 . If the Mach number at point $\mathbf{A}$ is 1.0 and rest of the flow field is subsonic, the value of $\boldsymbol{M}_{\infty}$ is

(A) 2.95

(B) 0.79

(C) 1.18

(D) 0.64

For inviscid, compressible flow past a thin airfoil, shown in the figure, free-stream Mach number and pressure are denoted by $\boldsymbol{M}_{\infty}$ and $\boldsymbol{p}_{\infty}$ respectively. Ratio of pressure at point $\mathbf{A}$ and $\boldsymbol{p}_{\infty}$ is 0.8 and specific heat ratio is 1.4 . If the Mach number at point $\mathbf{A}$ is 1.0 and rest of the flow field is subsonic, the value of $\boldsymbol{M}_{\infty}$ is

(A) 2.95

(B) 0.79

(C) 1.18

(D) 0.64

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