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    • 1. 发明申请
    • Compact aircraft combustor
    • 紧凑型飞机燃烧器
    • US20110079016A1
    • 2011-04-07
    • US12924671
    • 2010-09-30
    • Shahrokh EtemadBenjamin D. BairdSubir RoychoudhuryWilliam C. Pfefferle
    • Shahrokh EtemadBenjamin D. BairdSubir RoychoudhuryWilliam C. Pfefferle
    • F02C6/18F23R3/04F02C6/04F02C7/264
    • F23R3/40F23C6/045F23C13/06F23C2900/03002F23C2900/9901F23R3/286F23R3/32Y02T50/678
    • The present invention provides a combustor for an aerospace gas turbine engine comprising two stages wherein each stage defines an inlet and an exit. The second stage inlet is in fluid communication with the first stage exit such that a first flowpath is defined and it passes substantially through the second stage. A plurality of flow channel tubes is positioned within the second stage and each flow channel tube passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath that also passes substantially through the second stage. The first flowpath exit and the second flowpath exit are positioned adjacent and proximate to one another to provide for the generation of microflames or microflame jets exiting the second stage from between and around the flow channel tube exits. The first stage of the combustor provides a gasifier and a reformer. The present invention also may comprise an igniter for further combustion of the reacted products or an external heat source for start-up. The second stage also may comprise a microflame combustor.
    • 本发明提供一种用于航空航天燃气涡轮发动机的燃烧器,其包括两个阶段,其中每个阶段限定入口和出口。 第二级入口与第一级出口流体连通,使得限定第一流动路径,并且其基本上通过第二级。 多个流动通道管定位在第二阶段内,并且每个流动通道管密封地穿过位于第二阶段入口上游的集管板,从而限定也基本上通过第二阶段的第二流动路径。 第一流路出口和第二流路出口相邻并且靠近彼此定位,以提供从流动通道管出口之间和周围离开第二阶段的微火焰或微火焰射流的产生。 燃烧器的第一阶段提供气化器和重整器。 本发明还可以包括用于进一步燃烧反应产物的点火器或用于启动的外部热源。 第二阶段还可以包括微火焰燃烧器。
    • 6. 发明授权
    • Two stage combustor with reformer
    • 二级燃烧器与重整器
    • US08739550B2
    • 2014-06-03
    • US12924671
    • 2010-09-30
    • Shahrokh EtemadBenjamin D. BairdSubir RoychoudhuryWilliam C. Pfefferle
    • Shahrokh EtemadBenjamin D. BairdSubir RoychoudhuryWilliam C. Pfefferle
    • F02C7/22
    • F23R3/40F23C6/045F23C13/06F23C2900/03002F23C2900/9901F23R3/286F23R3/32Y02T50/678
    • The present invention provides a combustor for an aerospace gas turbine engine comprising two stages wherein each stage defines an inlet and an exit. The second stage inlet is in fluid communication with the first stage exit such that a first flowpath is defined and it passes substantially through the second stage. A plurality of flow channel tubes is positioned within the second stage and each flow channel tube passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath that also passes substantially through the second stage. The first flowpath exit and the second flowpath exit are positioned adjacent and proximate to one another to provide for the generation of microflames or microflame jets exiting the second stage from between and around the flow channel tube exits. The first stage of the combustor provides a gasifier and a reformer. The present invention also may comprise an igniter for further combustion of the reacted products or an external heat source for start-up. The second stage also may comprise a microflame combustor.
    • 本发明提供一种用于航空航天燃气涡轮发动机的燃烧器,其包括两个阶段,其中每个阶段限定入口和出口。 第二级入口与第一级出口流体连通,使得限定第一流动路径,并且其基本上通过第二级。 多个流动通道管定位在第二阶段内,并且每个流动通道管密封地穿过位于第二阶段入口上游的集管板,从而限定也基本上通过第二阶段的第二流动路径。 第一流路出口和第二流路出口相邻并且靠近彼此定位,以提供从流动通道管出口之间和周围离开第二阶段的微火焰或微火焰射流的产生。 燃烧器的第一阶段提供气化器和重整器。 本发明还可以包括用于进一步燃烧反应产物的点火器或用于启动的外部热源。 第二阶段还可以包括微火焰燃烧器。