首页> 外文OA文献 >Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber
【2h】

Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber

机译:火箭燃烧室高长宽比冷却通道设计的比较

摘要

An analytical investigation on the effect of high aspect ratio (height/width) cooling channels, considering different coolant channel designs, on hot-gas-side wall temperature and coolant pressure drop for a liquid hydrogen cooled rocket combustion chamber, was performed. Coolant channel design elements considered were: length of combustion chamber in which high aspect ratio cooling was applied, number of coolant channels, and coolant channel shape. Seven coolant channel designs were investigated using a coupling of the Rocket Thermal Evaluation code and the Two-Dimensional Kinetics code. Initially, each coolant channel design was developed, without consideration for fabrication, to reduce the hot-gas-side wall temperature from a given conventional cooling channel baseline. These designs produced hot-gas-side wall temperature reductions up to 22 percent, with coolant pressure drop increases as low as 7.5 percent from the baseline. Fabrication constraints for milled channels were applied to the seven designs. These produced hot-gas-side wall temperature reductions of up to 20 percent, with coolant pressure drop increases as low as 2 percent. Using high aspect ratio cooling channels for the entire length of the combustion chamber had no additional benefit on hot-gas-side wall temperature over using high aspect ratio cooling channels only in the throat region, but increased coolant pressure drop 33 percent. Independent of coolant channel shape, high aspect ratio cooling was able to reduce the hot-gas-side wall temperature by at least 8 percent, with as low as a 2 percent increase in coolant pressure drop. The design with the highest overall benefit to hot-gas-side wall temperature and minimal coolant pressure drop cooling can now be done in relatively short periods of time with multiple iterations.
机译:考虑到不同的冷却液通道设计,对高纵横比(高度/宽度)冷却通道对液氢冷却火箭燃烧室的热气侧壁温度和冷却液压降的影响进行了分析研究。所考虑的冷却剂通道设计元素为:应用高长宽比冷却的燃烧室长度,冷却剂通道数量以及冷却剂通道形状。使用火箭热评估代码和二维动力学代码的耦合,研究了七个冷却剂通道设计。最初,在不考虑制造的情况下开发了每种冷却剂通道设计,以从给定的常规冷却通道基线降低热气侧壁温度。这些设计使热气侧壁温度降低了22%,而冷却剂的压降却比基线降低了7.5%。铣削通道的制造约束条件已应用于这七个设计。这些使热气侧壁温度降低了20%,而冷却剂的压降降低了2%。在燃烧室的整个长度上使用高长宽比的冷却通道,与仅在喉咙区域使用高长宽比的冷却通道相比,对热气侧壁温度没有额外的好处,但是冷却剂压降增加了33%。与冷却剂通道形状无关,高长宽比的冷却能够将热气侧壁的温度降低至少8%,而冷却剂压降的降低幅度仅为2%。现在,可以在相对较短的时间内通过多次迭代完成对热气侧壁温度具有最高总体收益的设计,并且可以将冷却剂压降最小化。

著录项

  • 作者

    Wadel Mary F.;

  • 作者单位
  • 年度 1997
  • 总页数
  • 原文格式 PDF
  • 正文语种
  • 中图分类

相似文献

  • 外文文献
  • 中文文献
  • 专利

客服邮箱:kefu@zhangqiaokeyan.com

京公网安备:11010802029741号 ICP备案号:京ICP备15016152号-6 六维联合信息科技 (北京) 有限公司©版权所有
  • 客服微信

  • 服务号