Flame tolerant secondary fuel nozzle

A combustor for a gas turbine engine includes a plurality of primary nozzles configured to diffuse or premix fuel into an air flow through the combustor; and a secondary nozzle configured to premix fuel with the air flow. Each premixing nozzle includes a center body, at least one vane, a burner tube provided around the center body, at least two cooling passages, a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube. The cooling passages prevent the walls of the center body, the vane(s), and the burner tube from overheating during flame holding events.

FIELD OF THE INVENTION

The present invention relates to a flame tolerant secondary fuel nozzle in a premixer that includes cooling.

BACKGROUND OF THE INVENTION

Secondary nozzles in a combustor of a gas turbine may be permanently damaged when a flame is held in the premixing section of the nozzle. The use of high reactivity fuels makes this possibility more likely and confines operability of the gas combustor in a limited fuel space.

Use of high reactivity fuels increases flame holding risk that causes hardware damage and makes it more difficult to operate these fuels under premix operation. This has been previously addressed by so-called partially premixed design concepts that compromise mixing versus flame holding risk and increases NOx emissions.

Referring toFIG. 1, an exemplary gas turbine12includes a compressor14, a dual stage, dual mode combustor16and a turbine18represented by a single blade. Although not specifically shown, the turbine18is drivingly connected to the compressor14along a common axis. The compressor14pressurizes inlet air which is then turned in direction or reverse flowed to the combustor16where it is used to cool the combustor and also used to provide air to the combustion process. The gas turbine12includes a plurality of the combustors16(one shown) which are located about the periphery of the gas turbine12. A transition duct20connects the outlet end of its particular combustor16with the inlet end of the turbine18to deliver the hot products of the combustion process to the turbine18.

Referring toFIGS. 1 and 2, each combustor comprises a primary or upstream combustion chamber24and a second or downstream combustion chamber26separated by a venturi throat region28. The combustor is surrounded by a combustor flow sleeve30which channels compressor discharge air flow to the combustor. The combustor is further surrounded by an outer casing31which is bolted to the turbine casing32.

Primary nozzles36provide fuel delivery to the upstream combustion chamber24and are arranged in an annular array around a central secondary diffusion nozzle38. Each combustor may include six primary nozzles and one secondary nozzle, although it should be appreciated that other arrangements may be provided. Fuel is delivered to the nozzles through plumbing42. Ignition in the primary combustor is caused by spark plug48and in adjacent combustors by crossfire tubes50.

Referring toFIG. 2, a primary diffusion nozzle36includes a fuel delivery nozzle54and an annular swirler56. The nozzle54delivers only fuel which is then subsequently mixed with swirler air for combustion. The centrally located secondary nozzle38contains a major fuel/air premixing passage and a pilot diffusion nozzle.

During base-load operation, the dual stage, dual mode combustor is designed to operate in a premix mode such that all of the primary nozzles36are simply mixing fuel and air to be ignited by the secondary premixed flame supported by the secondary nozzle38. This premixing of the primary nozzle fuel and ignition by the secondary pilot diffusion nozzle leads to a lower NOx output in the combustor.

Referring still toFIG. 2, a diffusion piloted premix nozzle100includes a diffusion pilot having a fuel delivery pipe. The diffusion pilot further includes an air delivery pipe coaxial with and surrounding the fuel delivery axial pipe portion. The air input into the air delivery pipe is compressor discharge air which is reverse flowed around the combustor16into the volume76defined by the flow sleeve30and the combustion chamber liner78. The diffusion pilot includes at its discharge end a first or diffusion pilot swirler for the purpose of directing air delivery pipe discharge air to the diffusion pilot flame.

A premix chamber84is defined by a sleeve-like truncated cone which surrounds the diffusion pilot and includes a discharge end (as shown by the flow arrows) terminating adjacent the diffusion pilot discharge end. Compressor discharge air is flowed into the premix chamber84from volume76in a manner similar to the manner in which air is supplied to the air delivery pipe. The plurality of radial fuel distribution tubes extend through the air delivery pipe and into the premix chamber84such that the injected fuel and air are mixed and delivered to a second or premix chamber swirler annulus between the diffusion pilot and the premix chamber truncated cone. Further details of the combustor and gas turbine engine shown inFIGS. 1 and 2are disclosed in, for example, U.S. Pat. No. 5,193,346

BRIEF DESCRIPTION OF THE INVENTION

According to one embodiment of the invention, a combustor for a gas turbine engine comprises a plurality of primary nozzles configured to diffuse fuel into an air flow through the combustor; and a secondary nozzle configured to premix fuel with the air flow, the secondary nozzle comprising a fuel passage, a center body provided around the fuel passage, a burner tube provided around the center body and defining an annular air-fuel mixing passage between the center body and the burner tube, at least one vane in the annular air-fuel mixing passage configured to swirl the air flow, and at least two cooling passages comprising a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube.

According to another embodiment of the invention, a method of operating a combustor of a gas turbine engine is provided. The combustor comprises a plurality of primary nozzles provided in a primary combustion chamber and configured to diffuse fuel of a fuel supply to the combustor into an air flow through the combustor; and a secondary nozzle provided in a secondary combustion chamber and configured to premix fuel of the fuel supply with the air flow, the secondary nozzle comprising a fuel passage, a center body provided around the fuel passage, a burner tube provided around the center body and defining an annular air-fuel mixing passage between the center body and the burner tube, at least one vane in the annular air-fuel mixing passage configured to swirl the air flow, and at least two cooling passages comprising a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube. The method comprises providing an air flow to the combustor; and providing a fuel supply to at least one of the plurality of primary nozzles and the secondary nozzle; diffusing any fuel supplied to the primary nozzles into the air flow; premixing any fuel supplied to the secondary nozzle with the air flow; cooling the center body and the at least one vane with a portion of the fuel in the fuel cooling passage; and cooling the burner tube with a portion of the air flow between the burner tube and an outer peripheral wall.

DETAILED DESCRIPTION OF THE INVENTION

Referring toFIG. 3, a combustor2according to an embodiment includes a combustor head end4having an array of primary nozzles6and a secondary nozzle102. A combustion chamber liner10comprises a venturi46provided between a primary combustion chamber40and a secondary combustion chamber44. The combustion chamber liner10is provided in a combustor flow sleeve8. A transition duct22is connected to the combustion chamber liner10to direct the combustion gases to the turbine. Dilution holes34may be provided in the transition duct22for late lean injection.

Referring toFIG. 4, the combustor head end4comprises the array of primary nozzles6and the secondary nozzle102. As shown inFIG. 4, the primary nozzles6are provided in a circular array around the secondary nozzle102. It should be appreciated, however, that other arrays of the primary nozzles6may be provided.

The combustion chamber liner10comprises a plurality of combustion chamber liner holes52through which compressed air flows to form an air flow54for the primary combustion chamber40. It should also be appreciated that compressed air flows on the outside of the combustion chamber liner10to provide a cooling effect to the primary combustion chamber40.

The secondary nozzle102comprises a plurality of swirl vanes108that are configured to pre-mix fuel and air as will be described in more detail below. The secondary nozzle102extends into the primary combustion chamber40, but not so far as the venturi46.

Referring toFIG. 5, the combustor head end4comprises an end cover60having an end cover surface62to which the primary nozzles6are connected by sealing joints64. The secondary nozzle102comprises a fuel passage66that is supported by the end cover60. The secondary nozzle102further comprises an air flow inlet68for the introduction of air into the secondary nozzle102.

A nozzle center body106surrounds the end portion of the fuel passage66. The nozzle center body106comprises an end wall114. In the fuel passage66, the fuel flows downstream until it contacts the end wall114. The fuel flow then enters a reverse flow passage116and flows upstream as explained further below. As used herein, the term downstream refers to a direction of flow of the combustion gases through the combustor toward the turbine and the term upstream may represent a direction away from or opposite to the direction of flow of the combustion gases through the combustor.

The nozzle center body106may comprise annular ribs118to enhance heat transfer and cool the outer surface of the center body106. It should also be appreciated that the fuel passage66may comprise ribs, for example on the outer circumferential surface. The fuel passage66may comprise a plurality of holes110that bypass fuel directly to the swirling vanes108to control cooling and the pressure drop in the secondary nozzle102.

The fuel flows upstream in the reverse flow passage116into a cooling chamber70. The fuel then flows around a divider74into an outlet chamber72. The divider74may, for example, be a piece of metal that restricts the direction of flow of the fuel into the outlet chamber72, thus causing the fuel to internally cool all surfaces of the vanes108. The cooling chamber70and the outlet chamber72may be described as a non-linear coolant flow passage, e.g., a zigzag coolant flow passage, a U-shaped coolant flow passage, a serpentine coolant flow passage, or a winding coolant flow passage. A portion of the fuel may also flow directly from the cooling chamber70to the outlet chamber72through a by-pass hole88formed in the divider74.

The by-pass hole88may allow, for example, approximately 1-50%, 5-40%, or 10-20%, of the total fuel flow flowing from the cooling chamber70into the outlet chamber72to flow directly between the chambers70,72. Utilization of the by-pass hole88may allow for adjustments to any fuel system pressure drops that may occur, adjustments for conductive heat transfer coefficients, or adjustments to fuel distribution to fuel injection ports86. The by-pass hole88may improve the distribution of fuel into and through the fuel injection ports86to provide more uniform distribution. The by-pass hole88may also reduce the pressure drop from the cooling chamber70to the outlet chamber72, thereby helping to force the fuel through the fuel injection ports86. Additionally, the use of the by-pass hole88may allow for tailored flow through the fuel injection ports86to change the amount of swirl that the fuel flow contains prior to injection into a fuel-air mixing passage112via the injection ports86.

The fuel is ejected from the outlet chamber72through the fuel injection ports86formed in the swirl vanes108. The fuel is injected from the fuel injection ports86into the fuel-air mixing passage112for mixing with the air flow from the air flow inlet68of the secondary nozzle102. The swirl vanes108swirl the air flow from the air flow inlet68to improve the fuel-air mixing in the passage112.

Referring still toFIG. 5, the secondary nozzle102includes a burner tube122that surrounds the nozzle center body106. The fuel-air mixing passage112is provided between the nozzle center body106and the burner tube122. An outer peripheral wall104is provided around the burner tube122and defines a passage96for air flow. The burner tube122includes a plurality of rows of air cooling holes120to provide for cooling by allowing the coolant to form a film on the burner tube, protecting it from hot combustion gases. Coolant is also directed axially upstream within an annular cavity formed between the burner tube122and the outer peripheral wall104, in order that coolant may exit the cooling holes120upstream of the leading half of vanes108. The holes120may be angled in the range of 0° to 45° degree with reference to a downstream wall surface. The hole size, the number of holes in a circular row, and/or the distance between the hole rows may be arranged to achieve the desired wall temperature during flame holding events.

Operation of the combustor will now be described with reference toFIGS. 6-9. As shown inFIG. 6, during primary operation, which may be from ignition up to, for example, 20% of the load of the gas turbine engine, all of the fuel supplied to the combustor is primary fuel80, i.e. 100% of the fuel is supplied to the array of primary nozzles6. Combustion occurs in the primary combustion chamber40through diffusion of the primary fuel80from the primary fuel nozzles6into the air flow54through the combustor4.

As shown inFIG. 7, a lean-lean operation of the combustor occurs when the gas turbine engine is operated at, for example, 20-50% of the load of the gas turbine engine. Primary fuel80is provided to the array of primary nozzles6and secondary fuel82is provided to the secondary nozzle102. For example, about 70% of the fuel supplied to the combustor is primary fuel80and about 30% of the fuel is secondary fuel82. Combustion occurs in the primary combustion chamber40and the secondary combustion chamber44.

As used herein, the term primary fuel refers to fuel supplied to the primary nozzles6and the term secondary fuel refers to fuel supplied to the secondary nozzle102.

In a second-stage burning, shown inFIG. 8, which is a transition from the operation ofFIG. 7to a pre-mixed operation described in more detail below with reference toFIG. 9, all of the fuel supplied to the combustor is secondary fuel82, i.e. 100% of the fuel is supplied to the secondary nozzle102. In the second-stage burning, combustion occurs through pre-mixing of the secondary fuel82and the air flow from the inlet68of the secondary nozzle102. The pre-mixing occurs in the pre-mixing passage112of the secondary nozzle102.

As shown inFIG. 9, the combustor may be operated in a pre-mixed operation at which the gas turbine engine is operated at, for example, 50-100% of the load of the gas turbine engine. In the pre-mixed operation ofFIG. 9, the primary fuel80to the primary nozzles6is increased from the amount provided in the lean-lean operation ofFIG. 7and the secondary fuel82to the secondary nozzle102is decreased from the amount from provided in the lean-lean operation shown inFIG. 7. For example, in the pre-mixed operation ofFIG. 9, about 80-83% of the fuel supplied to the combustor may be primary fuel80and about 20-17% of the fuel supplied to the combustor may be secondary fuel82.

As shown inFIG. 9, during the pre-mixed operation, combustion occurs in the secondary combustion chamber44and damage to the secondary nozzle102is prevented due to the cooling measures. Referring toFIG. 4, flashback may occur in the event that the flame speed58is greater than the velocity of the air flow54in the primary combustion chambers40. Control of the air-fuel mixture in the secondary nozzle102, i.e. control of the secondary fuel82, provides control of the flame speed and prevents the flame from crossing the venturi46into the primary combustion chamber40.

Referring toFIGS. 10 and 11, secondary nozzle124comprises an inlet flow conditioner (IFC)126, an air swirler assembly132with natural gas fuel injection, and a diffusion gas tip146. A shroud extension134extends from the air swirler assembly132.

Air enters the secondary nozzle124from a high pressure plenum90, which surrounds the entire secondary nozzle124except the discharge end, which enters the combustor reaction zone94. Most of the air for combustion enters the premixer via the IFC126. The IFC126includes a perforated cylindrical outer wall128at the outside diameter, and a perforated end cap130at the upstream end. Premixer air enters the IFC126via the perforations in the end cap130and the cylindrical outer wall128.

The function of the IFC126is to prepare the air flow velocity distribution for entry into the premixer. The principle of the IFC126is based on the concept of backpressuring the premix air before it enters the premixer. This allows for better angular distribution of premix air flow. The perforated wall and endcap128,130perform the function of backpressuring the system and evenly distributing the flow circumferentially around the IFC annulus. Depending on the desired flow distribution within the premixer, appropriate hole patterns for the perforated wall and endcap128,130are selected.

Referring toFIG. 11, the air swirler assembly of the secondary nozzle124comprises a plurality of swirling vanes140and a plurality of spokes, or pegs,142provided between the swirling vanes140. Each spoke142comprises a plurality of fuel injection holes144for injecting fuel into the air swirled by the vanes140. Natural gas inlet ports136allow natural gas to be introduced into fuel passages138that are in communication with the spokes142. A nozzle extension148is provided between the air swirler assembly and the diffusion gas tip146. A bellows150may be provided to compensate for differences in thermal expansions.

Although the various embodiments described above include diffusion nozzles as the primary nozzles, it should be appreciated that the primary nozzles may be premixed nozzles, for example having the same or similar configuration as the secondary nozzles.

The flame tolerant nozzle enhances the fuel flexibility of the combustion system. The flame tolerant nozzle as the secondary nozzle in the combustor makes the combustor capable of burning full syngas as well as natural gas. The flame tolerant nozzle may be used as a secondary nozzle in the combustor and thus make the combustor capable of burning full syngas or high hydrogen, as well as natural gas. The flame tolerant nozzle, combined with a primary dual fuel nozzle, will make the combustor capable of burning both natural gas and full syngas fuels. It expands the combustor's fuel flexibility envelope to cover a wide range of Wobbe number and reactivity, and can be applied to oil and gas industrial programs.

The cooling features of the flame tolerant nozzle, including for example, the fuel cooled center body, the tip of the center body, the swirling vanes of the pre-mixer, and the air cooled burner tube, enable the nozzle to withstand prolonged flame holding events. During such a flame holding event, the cooling features protect the nozzle from any hardware damage and allows time for detection and correction measures that blow the flame out of the pre-mixer and reestablish pre-mixed flame under normal mode operation.