The design of liquid propellant rocket engine (LPRE) is a very complicated process; this is due to two main concerns: First, the requirements to satisfy the issues of performance, stability and compatibility. Second, the complicated, interacting processes inside thrust chamber. In this paper, an attempt to illustrate the importance of different parameters affecting performance, stability and compatibility is performed, followed by extensive study of processes inside thrust chamber. The result of processes study is developing the concept of “rate limiting process” which means that the process that can be considered the most important hence the design can be done mainly by considering it alone. This is done by developing a 1D vaporization-controlled model with its application to two case studies to illustrate model validation and application. It was found that the 1D model is valid as long as the vaporization process is the slowest process in this case the error in computing chamber cylindrical length is ∼15%. However, if the mixing process is slow, or the reaction process in gas phase is slow as in the second case study of RFNA/Tonka250 case, the error grow and may reaches 50%
In many applications, a constant thrust of the solid propellant rocket motor (SPRM) throughout the operating time is desired. Such thrust-time scenario is known as neutral burning. Typically, tubular grains with internal burning surface yield a progressive thrust-time history and, therefore, if a neutral thrust-time history is required, the progression in the burning surface should be compensated by including the side faces as well. However, neutrality of tubular grain that burns at both of the internal and end surfaces is subject to a limited range of slenderness ratio. Tapering the port at the end faces of the tubular grain would expand the range of the slenderness ratio that ensures neutrality. In the present paper, the issue of neutrality of taper-ended tubular grain is explored. For the studied case, it was found that with normal-ended tube grains neutrality can be realized for slenderness ratio 1.8 and web-to-radius ratio 0.4. However, keeping the same web-to-radius ratio and tapering the ends at angle 25 would slightly worsen the volumetric filling coefficient ( 15% less), but burning neutrality could be attained for slenderness ratios as high as 2.54 ( 40% higher). With this practice, it is possible to maintain neutrality where much longer grains may be adopted for certain applications.
The combustion phenomena inside the thrust chamber of the liquid propellant rocket engine are very complicated because of different paths for elementary processes. In this paper, the characteristic length (L*) approach for the combustion chamber design will be discussed compared to the effective length (Leff) approach. First, both methods are introduced then applied for real LPRE. The effective length methodology is introduced starting from the basic model until developing the empirical equations that may be used in the design process. The classical procedure of L* was found to over-estimate the required cylindrical length in addition to the inherent shortcoming of not giving insight where to move to enhance the design. The effective length procedure was found to be accurate within ± 10%.
The problem of accuracy of fire and its impact on weapon effectiveness presents a major concern of munitions designer. Many factors can be considered to improve the missile dispersion. Some are related to the missile itself, others are related to the launcher or the mutual effects between the missile and the launcher. The environmental launching conditions have a prime impact on the firing accuracy. The present work introduces simulation and experimental studies dealing with individual parameters and their influential weights concerning their impact on dispersion of a finned missile. The effects of the initial firing parameters, namely initial mass, initial launching angle, initial velocity and location of center of gravity are investigated. For each case the parameter was varied within certain limits to simulate the production inaccuracy, or approximations during design and the initial errors that can be impeded in the launcher and aiming devices. The trajectory was calculated using PRODAS software and compared with the real flight data. Comparisons of predicted and experimental results have proven satisfactory matching.
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