# The stagnation temperatures at the inlet and exit of a combustion chamber are $600 \mathrm{~K}$ and $1200 \mathrm{~K}$, respectively. If the heating value of the fuel is $44 \mathrm{MJ} / \mathrm{kg}$ and specific heat at constant pressure for air and hot gases are $1.005 \mathrm{~kJ} / \mathrm{kg} . \mathrm{K}$ and $1.147 \mathrm{~kJ} / \mathrm{kg} . \mathrm{K}$ respectively, the fuel-to-air ratio is (A) 0.0018 (B) 0.018 (C) 0.18 (D) 1.18

## Question ID - 156245 :- The stagnation temperatures at the inlet and exit of a combustion chamber are $600 \mathrm{~K}$ and $1200 \mathrm{~K}$, respectively. If the heating value of the fuel is $44 \mathrm{MJ} / \mathrm{kg}$ and specific heat at constant pressure for air and hot gases are $1.005 \mathrm{~kJ} / \mathrm{kg} . \mathrm{K}$ and $1.147 \mathrm{~kJ} / \mathrm{kg} . \mathrm{K}$ respectively, the fuel-to-air ratio is (A) 0.0018 (B) 0.018 (C) 0.18 (D) 1.18

3537

(B) 0.018

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A solid propellant of density $1800 \mathrm{~kg} / \mathrm{m}^{3}$ has a burning rate law $r=6.65 \times 10^{-3} p^{0.45} \mathrm{~mm} / \mathrm{s}$, where $p$ is pressure in Pascals. It is used in a rocket motor with a tubular grain with an initial burning area of $0.314 \mathrm{~m}^{2}$. The characteristic velocity is $1450 \mathrm{~m} / \mathrm{s}$. What should be the nozzle throat diameter to achieve an equilibrium chamber pressure of 50 bar at the end of the ignition transient?
(A) $35 \mathrm{~mm}$
(B) $38 \mathrm{~mm}$
(C) $41 \mathrm{~mm}$
(D) $45 \mathrm{~mm}$ 