Flashback suppression system for a gas turbine combustor
View Patent ↗A gas turbine combustor having improved flashback margin is disclosed. Multiple embodiments of the present invention are disclosed including combustors having a single premix chamber as well as multiple premix chambers. Flashback margin is increased for the combustor by incorporating a means for introducing an inert gas, such as nitrogen, into the premix chamber(s) at the region proximate the boundary layer to purge the boundary layer of combustible mixture, thereby ensuring that no combustion reaction occurs in the boundary layer.
1. A gas turbine combustor having increased flashback margin, said combustor comprising:
a plurality of premix chambers spaced apart a distance both axially and radially, each premix chamber having a first generally annular wall;
a combustion chamber having a second generally annular wall, said combustion chamber in fluid communication with each of said premix chambers;
at least one fuel nozzle for injecting a fuel into each of said premix chambers to form a mixture with compressed air in each of said premix chambers;
a boundary layer located in each of said premix chambers and said combustion chamber generally along said first and second generally annular walls, said boundary layer containing said fuel and compressed air at a lower velocity than said mixture; and,
a means for injecting an inert gas into said boundary layer for each of said plurality of premix chambers, said means for injecting located axially downstream with respect to a fluid flow direction from said at least one fuel nozzle, thereby ensuring that no combustion reaction occurs in said boundary layer.
2. The gas turbine combustor of claim 1 wherein said inert gas is nitrogen.
3. The gas turbine combustor of claim 1 wherein said fuel is selected from the group comprising natural gas, liquid fuel, gas from coal gasification, synthetic fuels, fuels having high hydrogen levels, and fuels having high hydrocarbon levels.
4. A gas turbine combustor comprising:
a plurality of premix chambers spaced a distance axially and radially apart, having a first generally annular wall;
a combustion chamber having a second generally annular wall, said combustion chamber in fluid communication with said one or more premix chambers;
a plurality of fuel nozzles capable of injecting a fuel into said plurality of premix chambers to form a mixture with compressed air in said plurality of premix chambers;
a boundary layer region located in at least said one or more premix chambers generally along said first generally annular wall, said boundary layer containing said fuel and compressed air at a lower velocity than said mixture; and,
a plurality of injection sources for injecting an inert gas into said boundary layer for each of said plurality of premix chambers, each of said plurality of injection means located axially downstream with respect to a fluid flow direction from said at least one fuel nozzle, so as to prevent a combustion reaction in said boundary layer.
5. The gas turbine combustor of claim 4 , wherein said fuel is injected into said one or more premix chambers before said inert gas is injected into said one or more premix chambers.
6. The gas turbine combustor of claim 5 , wherein said fuel is selected from the group comprising natural gas, liquid fuel, gas from coal gasification, synthetic fuels, fuels having high hydrogen levels, and fuels having high hydrocarbon levels.