Dual fuel ultra-low NOX combustor

Claim 2

A gas turbine, comprising: a) a compressor for compressing air; b) a combustor for producing a hot gas by burning a fuel in said compressed air, said combustor having: (i) a combustion zone, (ii) a centrally disposed first fuel nozzle in flow communication with said combustion zone and having a first discharge port for discharging a liquid fuel and a second discharge port for discharging a gaseous fuel, (iii) first and second concentrically arranged annular passages surrounding said first fuel nozzle and in flow communication with said combustion zone, (iv) means for introducing a liquid fuel into said first passage so as to circumferentially distribute said liquid fuel around said first passage, (v) first means for introducing a gaseous fuel into said second passage so as to circumferentially distribute said gaseous fuel around said second passage; and c) a turbine for expanding said hot gas produced by said combustor.

Claim 3

The gas turbine according to claim 1, further comprising second means for introducing a gaseous fuel into said first passage so as to circumferentially distribute said gaseous fuel around said first passage.

Claim 4

The gas turbine according to claim 2, further comprising a plurality of second fuel nozzles circumferentially distributed around said first passage, and wherein: a) said second gas fuel introducing means comprises a first gas fuel discharge port formed in each of said second fuel nozzles; and b) said liquid fuel introducing means comprises a second liquid fuel discharge port formed in each of said nozzles.

Claim 5

The gas turbine according to claim 3, further comprising a plurality of first swirl vanes distributed circumferentially around each of said second fuel nozzles.

Claim 6

The gas turbine according to claim 4, wherein: a) said first passage has an inlet in flow communication with said compressor, whereby said inlet receives a first portion of said compressed air from said compressor; and b) said first swirl vanes are disposed downstream of said first passage inlet and between said first gas fuel discharge ports and said combustion zone, whereby said first swirl vanes create pre-mixing of said first portion of said compressed air and a gaseous fuel from said first gas fuel ports prior to said compressed air and said gas fuel entering said combustion zone.

Claim 7

The gas turbine according to claim 4, further comprising a plurality of second swirl vanes distributed circumferentially around said second passage, and wherein: a) said first gas fuel introducing means comprises a plurality of second gas fuel discharge ports circumferentially arrayed within said second passage; b) said second passage has an inlet in flow communication with said compressor, whereby said inlet receives a second portion of said compressed air from said compressor; and c) said second swirl vanes are disposed between said second gas fuel discharge ports and said combustion zone, whereby said second swirl vanes create pre-mixing of said second portion of said compressed air and a gaseous fuel from said second gaseous fuel discharge ports prior to said second portion of said compressed air and said gas fuel entering said combustion zone.

Claim 8

The gas turbine according to claim 6, further comprising: a) a third annular passage concentrically arranged with respect to said first and second passages, said third passage surrounding said first fuel nozzle and in flow communication with said combustion zone; b) third means for introducing a gaseous fuel into said third passage so as to circumferentially distribute said gaseous fuel around said third passage; c) a plurality of third swirl vanes distributed circumferentially around said third passage, and wherein: (i) said third gas fuel introducing means comprises a plurality of third gas fuel discharge ports circumferentially arrayed within said third passage; (ii) said third passage has an inlet in flow communication with said compressor, whereby said inlet receives a third portion of said compressed air from said compressor; and (iii) said third swirl vanes are disposed between said third gas fuel discharge ports and said combustion zone, whereby said third swirl vanes create pre-mixing of said third portion of said compressed air and a gaseous fuel from said third gaseous fuel discharge ports prior to said third portion of said compressed air and said gas fuel entering said combustion zone.

Claim 9

The gas turbine according to claim 1, wherein said liquid fuel introducing means comprises a plurality of fuel nozzles circumferentially arrayed within said first annular passage, each of said fuel nozzles having a liquid fuel discharge port formed therein.

Claim 10

The gas turbine according to claim 1, wherein said first gas fuel introducing means comprises a plurality of gas fuel discharge ports circumferentially arrayed within said second passage.

Claim 11

The gas turbine according to claim 9, further comprising a plurality of swirl vanes distributed circumferentially around said second passage, and wherein: a) said second passage has an inlet in flow communication with said compressor, whereby said inlet receives said compressed air; and b) said swirl vanes are disposed between said gas fuel discharge ports and said combustion zone, whereby said swirl vanes create pre-mixing of said compressed air and a gas fuel from said gaseous fuel discharge ports prior to said compressed air and said gas fuel entering said combustion zone.

Claim 12

The gas turbine according to claim 9, wherein said first gas fuel introducing means further comprises a toroidal gas fuel manifold extending around said second passage, said gas fuel discharge ports being distributed around said toroidal manifold.

Claim 13

The gas turbine according to claim 1, wherein said second passage surrounds said first passage.

Claim 14

The gas turbine according to claim 1, wherein said first passage surrounds said second passage.

Claim 15

The gas turbine according to claim 1, further comprising a third annular passage concentrically arranged with said first and second passages, said third passage surrounding said first fuel nozzle and in flow communication with said combustion zone.

Claim 16

The gas turbine according to claim 14, further comprising means for introducing a gaseous fuel into said third passage so as to circumferentially distribute said gaseous fuel around said third passage.

Claim 17

The gas turbine according to claim 15, wherein said third passage surrounds said first and second passages.

Claim 18

The gas turbine according to claim 15, wherein said first passage is disposed between said second and third passages.

Claim 19

In a gas turbine having a compressor for producing compressed air, a combustor comprising: a) a shell forming a combustion zone in which a fuel is burned in compressed air; b) a fuel/air pre-mixing zone enclosed by a first liner, said first liner enclosing second and third liners, a first annular passage formed between said first and second liners and a second annular passage formed between said second and third liners, each of said annular passages having an inlet in flow communication with said compressor and an outlet in flow communication with said shell, whereby a portion of said compressed air from said compressor flows through each of said annular passages, said pre-mixing zone having a flow area defined by an inner diameter of said first liner at said first passage outlet, said shell having a flow area adjacent said passage outlets defined by an inner diameter of said shell, said shell inner diameter being at least about 40% greater that said first liner inner diameter, whereby said compressed air flowing through said passages undergoes an expansion upon exiting said passages; c) means for introducing a fuel into each of said passages; and d) means, disposed within said passages, for mixing said fuel introduced into each of said annular passages with said compressed air flowing through said passages.

Claim 20

In a gas turbine having a compressor for producing compressed air, a combustor comprising: a) a shell forming a combustion zone in which a fuel is burned in compressed air; b) a fuel/air pre-mixing zone enclosed by a first liner, said first liner enclosing second and third liners, a first annular passage formed between said first and second liners and a second annular passage formed between said second and third liners, each of said annular passages having an inlet in flow communication with said compressor and an outlet in flow communication with said shell, whereby a portion of said compressed air from said compressor flows through each of said annular passages, said pre-mixing zone having a flow area defined by an inner diameter of said first liner at said first passage outlet, said shell having a flow area adjacent said passage outlets defined by an inner diameter of said shell, said shell inner diameter being at least about 40% greater that said first liner inner diameter, whereby said compressed air flowing through said passages undergoes an expansion upon exiting said passages; c) means for introducing a fuel into each of said passages; and d) means, disposed within said passages, for mixing said fuel introduced into each of said annular passages with said compressed air flowing through said passages.

Claim 21

The combustor according to claim 18, wherein said means for mixing said fuel comprises a plurality of swirl vanes disposed in each of said passages.

Claim 22

The combustor according to claim 18, further comprising means for introducing a fuel centrally within said shell.