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Copyright of Chinese Journal of Explosives & Propellants is the property of Chinese Journal of Explosives & Propellants Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Aero Weaponry is the property of Aero Weaponry Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Journal of Ordnance Equipment Engineering is the property of Chongqing University of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Space rocket engine nozzle usually has large area ratio, and the improvement in its performance is one of the goals of engine design. A design method for divergent section considering the changes in properties of the real fuel gas in nozzle is proposed based on the Rao’s method. For an apogee rocket engine, the Rao’s method is used to design an original nozzle contour, and then the function of specific heat of the real fuel gas is calculated and a new nozzle contour of divergent section of nozzle is designed.The vacuum specific impulses of the new nozzle contour and the original one are compared through numerical simulations. Simulated results show that the calculation method of fuel gas specific heat can be used for the nozzle flow simulation; the specific heat of fuel gas has great effect on the nozzle contour design; for the nozzles with the same length, the nozzle contour designed by considering the change in specific heat capacity has smaller exit area ratio, but has higher vacuum specific impulse; and for the space engine which has not too large thrust, the increase in vacuum specific impulse is less then 1s, and the gain of performance by the boundary layer displacement is very small. [ABSTRACT FROM AUTHOR]

Ultrasonic measurement of burning rate is an advanced technique. For application of ultrasonic measurement in solid rocket motor test,some typical materials used in solid rocket motor（ SRM） are measured by ultrasonic technique,and their ultrasonic signal signatures were obtained. The real-time thickness variation of solid propellant at ordinary pressure is measured by an ultrasonic transducer mounted on the outside of the motor shell. A data processing method is proposed to manipulate the echo wave for real-time measurement of burning rate. The proposed method was successfully used in a hot fire test in which the propellant burns at ordinary pressure. The variation in thickness of propellant and its burning rate at ordinary pressure was obtained. The influence of temperature distribution near burning surface on burning rate test is analyzed. The results show that for motor with metal shell,a window must be made on the metal shell to let enough ultrasonic wave energy penetrate the interface between shell and isolator,and for composite shell motor the probe can be mounted outside the shell directly. The influence of temperature variation near burning face caused by combustion on measurement is negligible. The proposeddata processing method can be used effectively to obtain the variation of grain thickness. [ABSTRACT FROM AUTHOR]

It is inevitable to evaluate the aerodynamic heat and the ablation of thermal protection material for designing a hypersonic aircraft. A coupling method is presented for the calculation of aerodynamic heat, material ablation and transient temperature distribution. The method and program are validated by experimental data. The results indicate that the method has higher precision and efficiency. For a given trajectory, the aerodynamic heat, ablation performance of thermal protection layer and temperature filed are predicted by the coupling method. The proposed method can be used to calculate the ablation and temperature distribution in given operation conditions in design phase which is referable for the design of multi-layer thermal protection material for hypersonic aircrafts. [ABSTRACT FROM AUTHOR]