4105 衝撃風洞を用いたラバルノズルスロート部における伝熱特性に関する研究(G19 宇宙工学に関する基盤研究,G19 宇宙工学)

概要

論文の詳細を見る
The adequate cooling of nozzles of rocket engines requires design information for predicting a local heat flux. The heat flux has highest value at a nozzle throat area. The purpose of this study is to reveal experimental heat transfer data in the nozzle throat by using a shock tunnel. The shock tunnel was built with a shock tube and instrumented nozzle at the end of the shock tube. In the case of shock tunnel, duration time of supersonic flow is so short that high response instrumentation or methods are necessary for measuring the transient heat flux and temperature. We applied platinum thin film heat flux gauge. Open literatures suggest several experimental equations that may be applicable for predicting the heat transfer coefficients in the nozzle. Bartz's equation is one of them and is a well-known equation. This paper will present comparison between experimental measurement data by using the developed gauges and estimation by Bartz's equation.
社団法人日本機械学会の論文
2007-09-07