A COMPARISON BETWEEN FLOWS OF THREE REACTION SYSTEMS THROUGH SUPERSONIC PROPULSIVE NOZZLES

Document Type : Original Article

Authors

1 Prof., Mechanical Engineering Department, Al-Azhar University, Cairo, Egypt.

2 Associate Prof. , Mechanical Engineering Department, Al-Azhar University, Cairo, Egypt.

Abstract

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.

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