@techreport{oai:jaxa.repo.nii.ac.jp:00044726,
 author = {鈴木, 昭夫 and 八柳, 信之 and 五味, 広美 and 坂本, 博 and SUZUKI, Akio and YATSUYANAGI, Nobuyuki and GOMI, Hiromi and SAKAMOTO, Hiroshi},
 month = {Sep},
 note = {An experimental investigation of combustion performance was conducted in a LOX/GH2 rocket combustor with coaxial type injectors at a thrust level of 300 kgf with a nominal chamber pressure of 20 atm. The oxidant-fuel ratios ranged from approximately 2 to 10. The combustion chambers were of the segmented, heat-sink type. The chamber geometry varied in lengths from 11 to 26 cm and contraction rations of 2.03, 3.41 and 4.46, with a constant nozzle throat area. The injector had twelve coaxial elements with no recess and three replaceable faceplates having different hydrogen port areas. The attendant hydrogen-oxygen injection velocity ratio ranged from approximately 10 to 50. Observed C* efficiencies were correlated as a function of an equivalent chamber length, Le, in which the injection velocity ratio, VR, and the chamber length, L, are expressed as  Le=L+αVR  where α was considered nearly constant throughout the experiment. The correlation takes the form of ηc*=1-A exp(-BLe) where the constants A and B depend on the chamber contraction ratio. The variations of all experimental data from the relation were within approximately ± 1.8 percent., 資料番号: NALTR0473000, レポート番号: NAL TR-473},
 title = {液体酸素・ガス水素ロケットの燃焼性能(Ⅰ)},
 year = {1976}
}