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An improved constant volume cycle (CVC) model is developed to analyze the nozzle effects on the thrust and specific impulse of pulse detonation rocket engine (PDRE). Theoretically, this model shows that the thrust coefficient/specific impulse of PDRE is a function of the nozzle contraction/expansion ratio and the operating frequency. The relationship between the nozzle contraction ratio and the operation frequency is obtained by introducing the duty ratio, by which the key problem in the theoretical design can be solved. Therefore, the performance of PDRE can be accessed to guide the preliminary shape design of nozzle conveniently and quickly. The higher the operating frequency of PDRE is, the smaller the nozzle contraction ratio should be. Besides, the lower the ambient pressure is, the larger the expansion ratio of the nozzle should be. When the ambient pressure is 1.013 × 10 Pa, the optimal expansion ratio will be less than 2.26. When the ambient pressure is reduced to vacuum, the extremum of the optimal thrust coefficient is 2.236 9, and the extremum of the specific impulse is 321.01 s. The results of the improved model are verified by numerical simulation. [ABSTRACT FROM AUTHOR]

Rocket motors that are designed to operate at high altitudes need a nozzle with a large expansion ratio to maximize the thrust at much lower atmospheric pressure than that of the sea level pressure. Accurate performance of these nozzles cannot be obtained when static tested on ground. Computational fluid dynamics (CFD) analyses have been performed to characterise the supersonic exhaust diffuser (SED). The results obtained from the CFD analyses have been found to be in good agreement with experimental and numerical values reported in the published literature. Started and un-started regions of the SED have been identified with the CFD results. [ABSTRACT FROM AUTHOR]

The pintle controlled solid rocket motor shows distinct non-steady characteristic when nozzle aera is changed by pintle motion. In order to understand the non-steady characteristic, simulation of inflow field during the thrust controlling process with pintle motion was implemented by using the unique dynamic mesh layering model in Fluent software. Velocity of pintle move, nozzle geometry and propellant parameter were investigated as the main influence factors on the non-steady characteristic. It is found that the velocity and pressure index affect the non-steady characteristic greatly, the effect of nozzle convergence half angle on the non-steady characteristic takes the second and expansion half angle and pintle radius have almost no affect to the non-steady characteristic. [ABSTRACT FROM AUTHOR]

Abstract: Simulation of the temperature distribution on the rocket nozzle wall in high temperature environment is most important for assessing safety and reliability of the nozzle. For that purpose, the coupling simulation of heat transfer and transient temperature of the rocket nozzle wall is carried out in this paper. Equations of radiative heat flux on surfaces in an enclosure with inhomogeneous, participating media is developed to compute the radiative heat flux. The convective heat transfer considering ablation of the C/C material has been studied. And radiative transfer between the plume and nozzle wall has been analyzed with the net radiation computing method in one enclosure. Also the differential equation for two-dimensional transient conduction in the cylindrical coordinate has been established, discretized and solved. The convection and radiation heat fluxes, also the transient temperature along the wall length and thickness of the composite nozzle wall after startup of the engine are investigated. Two benchmark problems have been taken, one is the nozzle of an experimental rocket engine fabricated with stainless substrate and a ZrO2 coating, the other is constructed with multi-layer composite materials. The computing results show that: plume heating on the throat section is most serious and wall temperature in that place is much higher than the average temperature, and heat insulation of the throat is very important for rocket nozzle. The typical character of transient temperature is steady increase with time after startup of the rocket engine, both on the internal and outside walls of the rocket nozzle while the former increases rapidly. On the same section, the temperature decreases along the thickness from internal to outside of the wall, the difference is over 1000 K due to the heat insulating layer in the nozzle material. The coupled simulation method and related results in the present study can be best used to predict the safety performance of the rocket nozzle in startup stage. Results of coupling simulation method are compared with available results in references which show reasonable agreement. [Copyright &y& Elsevier]

Abstract: The modelling of ablation of carbon/carbon (C/C) composites utilized as rocket engine hot parts is addressed under the angle of the competition between bulk transport of reactants and heterogeneous mass transfer, associated to reactivity contrasts between constituent phases. A numerical solver based on a simple model and built on a VOF technique allows direct simulation at two scales. Its application to actual complex materials is performed; the results are consistent with experimental data and help understanding the origin of the material behaviour, either in terms of acquired surface morphology or in terms of effective recession rate. [Copyright &y& Elsevier]

Abstract: The Department of Mechanical Engineering at the California Polytechnic State University, San Luis Obispo, has developed an innovative program of experimental research and development on hybrid rocket motors (where the fuel and the oxidizer are in different phases prior to combustion). One project currently underway involves the development of aerospike nozzles for such motors. These nozzles, however, are even more susceptible to throat ablation than regular converging–diverging nozzles, due the nature of their flow expansion mechanism. This paper presents the result of a recent development project focused on reducing throat ablation in hybrid rocket motor nozzles. Although the method is specifically targeted at increasing the life and operating range of aerospike nozzles, this paper describes its proof-of-concept implementation on conventional nozzles. The method is based on a regenerative cooling mechanism that differs in practice from that used in liquid propellant motors. A series of experimental tests demonstrate that this new method is not only effective at reducing damage in the most ablative region of the nozzle, but that the nozzle can survive multiple test runs. [Copyright &y& Elsevier]

Abstract: This paper presents the study of a concept of Ariane-5 evolution by means of replacement of two solid-propellant boosters EAP with two liquid-propellant reusable fly-back boosters (RFBBs) called “Bargouzine”. The main design feature of the reference RFBB is LOX/LH2 propellant, the canard aerodynamic configuration with delta wings and rocket engines derived from Vulcain-2 identical to that of the central core except for the nozzle length. After separation RFBBs return back by use of air breathing engines mounted in the aft part and then landing on a runway. The aim of the study is a more detailed investigation of critical technology issues concerning reliability, re-usability and maintenance requirements. The study was performed in three main phases: system trade-off, technical consolidation, and programmatic synthesis. The system trade-off includes comparative analysis of two systems with three and four engines on each RFBB and determination of the necessary thrust level taking into account thrust reservation for emergency situations. Besides, this phase contains trade-off on booster aerodynamic configurations and abort scenario analysis. The second phase includes studying of controllability during the ascent phase and separation, thermo-mechanical design, development of ground interfaces and attachment means, and turbojets engine analysis taking into account reusability. [Copyright &y& Elsevier]

Abstract: One of the ways to increase the performance of liquid rocket engines (LRE) is to use nozzle extensions (NE) made of carbon–carbon composites (CCC). The following works were performed during development of the NE: a fabrication method was verified, physicomechanical properties of the material were defined, effects of high temperature on a CCC oxidizing resistance were studied, a method to protect CCC against oxidation was developed. Prototype NE were manufactured and tested in a LRE test chamber. The test firings demonstrated good performance characteristics of CCC. [Copyright &y& Elsevier]

X-ray radiography techniques are widely used for the non-destructive evaluation of composite nozzle liners of rocket motors. In conventional radiography it is difficult to distinguish the images corresponding to excess resin and resin lean regions in the composite, since the variation in the optical density is very small. Also, porous layers and minor delaminations will cast similar images in the radiographic projection. In this work, theoretical modelling is carried out for various defect conditions to study their effects on the X-ray images. The results are useful for characterisation of the defects in the nozzle liners by radiographic nondestructive evaluation. The observations provide very good input for the interpretation of radiographs in silicaphenolic and carbon-phenolic rocket nozzles. The details of the software developed and its application to solve specific problems are discussed. The changes in optical density with respect to the width of the defect, location of the defect and circumferential extension are investigated. [ABSTRACT FROM AUTHOR]