ABSTRACT
The present study is concerned with the analytical solution of frozen flows of
combustion gases through the supersonic region of an axisymmetric nozzle. The
product gases of C12H26–O2 , H2–O2 and H2–F2 reaction systems are considered. A
predictor-corrector method is employed for approximating the differential terms of the
governing equations; while the method of characteristics is used to solve the resulting
equations for variable specific heat ratio. The study includes a comparison of the results
of the three reaction systems. The results involve the flow field parameters; namely ,
the temperature T, the pressure P, the velocity V and Mach number M. The results of
the propulsion Parameters namely; thrust force F, the specific impulse Isp, specific
impulse based on fuel combustion Isp,f and thrust efficiency ηf are also presented. All
results are presented for a combustion chamber dimensionless pressure , Pc=30 , an
equivalence ratio, Φ=1, a dimensionless throat radius of curvature of the nozzle, ρt=2 ,
and a straight nozzle wall angle, ψ=20°, as a controlling factors. Results show that the
point of tangency at nozzle wall is a source of creating oblique shock waves. The
C12H26–O2 , H2–O2 reaction systems produce approximately equal thrusts which are
higher than that produced by H2–F2 reaction system. Besides, the specific impulse
attained by H2–F2 system is of highest magnitude; and there are negligible influences of
the three systems on the thrust efficiency.