KR20140082659A - Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines - Google Patents

Abstract

A combustion apparatus for use in a gas turbine engine for rotating a power generating shaft or generating propulsion includes a can-annular type combustion chamber having a fuel and air inlet passageway and a system of nozzles so that the optimum combustion of the premixed fuel and air Environment. The fuel-air inlet is located in various longitudinal positions and distributed in the circumferential direction, directing the flow tangentially or substantially tangentially to the can liner. Combustors provide effective mixing of fuel and air, reduce the need for costly contaminant suppression devices, improve ignition and flame stability, reduce steering problems, and improve vibration reduction.

Description

CAN CAN ANNULAR COMBUSTOR WITH PREMIXED TANGENTIAL FUEL-AIR NOZZLES FOR USE ON GAS TURBINE ENGINES WITH PRE-MIXED TANKLED FUEL-AIR NOZZLE USED IN A GAS TURBINE ENGINE

The present invention relates to an apparatus in a gas turbine engine that serves to receive a fuel and air mixture and to cause combustion of the fuel and air mixture. Such devices include, but are not limited to, flow transitions, fuel-air nozzles, combustion chamber liners and casings used in military and commercial aircraft, power generation, and other applications related to gas turbines.

Gas turbine engines include machines that extract work from flue gases flowing at very high temperatures, pressures and speeds. The extracted work can be used to provide the required thrust to the aircraft or to drive the generator to generate power. A typical gas turbine engine consists of a multi-stage compressor in which the atmosphere is compressed to high pressure. The compressed air is mixed at a specific fuel / air ratio in a combustion chamber where the temperature is raised. The high temperature and high pressure combustion gas is expanded through the turbine to extract the work to provide the required thrust according to the application or to drive the generator. The turbine includes at least one stage in which each stage consists of a row of blades and a row of vanes. The blades are circumferentially distributed on the rotating hub, and the height of each blade covers the hot gas flow path. The non-rotating blades of each stage are arranged in the circumferential direction, which also extends across the hot gas flow path. The present invention includes a combustion chamber of a gas turbine engine, and a component for injecting fuel and air into the apparatus.

The combustion chamber portion of a gas turbine engine can be divided into two types of combustion chambers that form several different types: can / tubular, annular, and can-annular combustion chambers . ≪ / RTI > In the component, a compressed fuel-air mixture passes through the fuel-air swirler and a combustion reaction of the mixture occurs, resulting in a hot gas flow that lowers the density and accelerates the downstream. The can-type combustion chamber typically has a respective circumferentially spaced can accommodating the flames of each nozzle separately. The flow from each can is directed through the duct and bound in an annular transition piece before the flow enters the first stage vane. In the annular combustion chamber, the fuel-air nozzles are typically circumferentially distributed and inject the mixture into a single annular chamber in which combustion occurs. The flow does not require a transition to join the flow and simply drains the downstream end of the annulus into the first stage turbine. An important difference between the last type can-annular type combustion chamber is that the can-annular type combustion chamber has individual cans surrounded by an annular casing that receives the air supplied into each can. Each change has advantages and disadvantages depending on the application.

In a combustion chamber for a gas turbine, it is common for a fuel-air nozzle to inject a swirl into the mixture for a number of reasons. One reason is to improve mixing and consequently the combustion, and the other reason is to add a vortex to stabilize the flame to prevent the flame from turning off and to reduce the leaner fuel-air mixture So that it can be used. The fuel air nozzles may have different configurations, such as one to several annular inlets with vortex vanes, respectively.

As with other gas turbine components, it is necessary to implement a cooling method to prevent melting of the combustion chamber material. A typical method of cooling the combustion chamber is effusion cooling, which is carried out by surrounding the combustion liner with an additional offsetted liner, through which the compressor discharge air passes, and a dilution hole and a cooling passage Lt; RTI ID = 0.0 > flow path. ≪ / RTI > This technique not only forms a thin boundary film of cooling air between the liner and the combustion gas, but also removes heat from the components to prevent heat transfer to the liner. The dilution ball serves two purposes depending on the axial position on the liner: First, the dilution ball, which is closer to the fuel-air nozzle, not only provides unburned air for combustion, but also improves combustion Secondly, the dilution holes, which are located closer to the turbine, can be designed to cool the hot gas flow and adjust the combustion chamber outlet temperature profile.

It can be seen that several methods and techniques can be incorporated into the design of the combustion chamber for a gas turbine engine to improve combustion and reduce emissions. Gas turbines tend to discharge less pollutants than other generation methods, but there is still room for improvement. In situations where government regulations on exhaust gases in some countries are becoming more stringent, it may be necessary to improve these techniques to meet these requirements.

In connection with the present invention, new and improved combustion chamber design forms are provided that can operate in a conventional manner with minimal pollutant emissions resulting from the combustion of fuel and air mixtures. The present invention relates to a typical can-annular combustion chamber having a dilution hole and / or a premixed fuel-air nozzle for injecting compressor discharge air and compressed fuel into combustion chambers at various positions in the longitudinal and circumferential directions, . An inventive feature of the invention is that the fuel and air nozzles are arranged to create an environment in which mixing of the combustion reactants and product is improved. Arranging the premix fuel and air nozzles to allow more fuel to be upstream from the other set of nozzles can improve the mixing of the combustion reactants and reduce the specific oxygen concentration in the combustion zone . Additionally, by injecting compressor discharge air downstream of the combustion zone, CO generated during the combustion process can be burned / consumed before entering the first stage turbine. Indeed, the combustion chamber improves the gas turbine emission level, and as a result not only minimizes the environmental impact of the exhaust gas control device, but also reduces the need for an exhaust gas control device. In addition to these improvements, the ignition process of the combustion chamber is greatly improved by directing the initial flame front toward the adjacent combustion device nozzles in each can by tangentially directing the fuel nozzles and fuel-air nozzles .

The present invention relates to a can-annular combustion chamber having a premixed tangential fuel-air nozzle for use in a gas turbine engine, and in accordance with the present invention, can improve combustion and reduce exhaust gas and greatly improve the combustion process of the combustion chamber .

Referring to the drawings:
BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 is a two-dimensional sketch showing a can-annular array with nozzles attached to an outer can liner that injects fuel and air into a common plane.
2 is a two-dimensional sketch showing the concept of a tangential nozzle in contact with a can in a can-annular combustion chamber.
Figure 3 is an isometric view as a side view of an upstream portion of an exemplary configuration of the invention.
4A is a cross-sectional view of the present invention taken along an isometric view.
4B is an enlarged view of the shape of Fig. 4A.
5 is a cross-sectional view taken along line AA as shown in Fig.
Fig. 6 is a cross-sectional view taken along line BB in Fig.

1 shows an example of a general arrangement of a can-annular type combustion chamber with a can 1 spaced circumferentially on a common radius, all of the cans in the can-annular type combustion chamber are arranged in a cylindrical outer line 2, Shaped inner line (3). This figure also shows the tangential nozzle arrangement of the can. Figure 2 shows the can in more detail. The can liner 4 forms a can space, and the fuel / air nozzle 5 injects fuel or air. The nozzle forms an angle 8 between the nozzle center line 6 and a line in contact with the can liner 4 intersecting the nozzle center line 6. This angle defines the circumferential direction of the nozzle.

2 shows the general operation of the can in an exemplary can-annular combustion chamber configuration, in which case the premixed fuel-air mixture 9 is injected into the can 1 at a constant angle 8. The flame 10 is formed throughout the can and moves across the can in a path 11 along the can liner. These tangential nozzles cause a flow from each nozzle that interacts with the adjacent nozzles downstream. This important feature reduces the need for control of the combustion device nozzles and improves ignition, which is possible by allowing flames from the nozzles to ignite the fuel at the downstream, adjacent nozzles.

3 shows the beginning or upstream portion of an exemplary can with the downstream portion omitted. The above invention may have a plurality of nozzle rows spaced along the length of the can. Each row of nozzles may have at least one nozzle and may be offset from the adjacent nozzle array by a circumferential angle. The can also have several rows of holes 14 or passages circumferentially spaced apart, which are intended to cool the air entering the can at any location.

4A and 4B show the top surface 13 of the can, which may have holes 14 similar to a dilution hole that allows compressor discharge air to enter the can. Figures 5 and 6 show how the nozzles in each set of rows can be offset by a constant circumferential angle. The other row of nozzles allows the injection of the fuel-air mixture near the front wall to produce a fuel-air batch effect and a target mix that can create an optimal combustion environment that reduces the emissions of NOx and CO from the combustion chamber, The fuel-air mixture herein may combine with the mixture injected downstream of the fuel nozzle 5 to have a greater fuel / air ratio than the second set of nozzles.

The present invention has been described above with reference to preferred embodiments. However, it will be appreciated by those of ordinary skill in the art that changes and modifications may be made to the embodiments described without departing from the spirit and scope of the invention. Various changes and modifications to the embodiments herein selected for clarity will be apparent to those of ordinary skill in the art. Such modifications and variations are intended to be included within the scope of the present invention without departing from the spirit of the present invention.

The present invention has been fully described in such clear and concise language, so that those skilled in the art will be able to understand and to practice the invention. The claims of the present invention are as follows.

Claims (19)

A can-annular combustion chamber for a gas turbine for use in surface-based power generation, land or marine-based transportation means or aircraft engine applications,
A plurality of can liners spaced apart in the circumferential direction,
The plurality of can liners are cylindrically shaped,
Each can has a plurality of circumferentially spaced fuel-air nozzles directed in a tangential direction,
The plurality of fuel-air nozzles share a common plane perpendicular to the direction of the can center line,
Characterized in that all liners are made of a high temperature alloy or a ceramic material. The method according to claim 1,
The nozzles may be spaced circumferentially in a common plane perpendicular to the longitudinal direction, near the front wall of the can, and may inject a fuel-air mixture having a greater fuel / air ratio than a downstream set of nozzles that may be present, Characterized in that it has a radial direction and / or a longitudinal direction predominantly in the circumferential direction. The method according to claim 1,
Wherein the nozzles are spaced circumferentially in at least one common plane perpendicular to the longitudinal direction and downstream from the nozzle referred to in claim 2 and having a fuel / air ratio of less than the fuel / air ratio at the nozzle described in claim 2 - air mixture and can have a radial and / or longitudinal direction predominantly in the circumferential direction. The method according to claim 1,
Characterized in that the nozzle can have an angle of constant value or a varying angle from plane to plane in the range of 0 to 90 degrees, as indicated by reference numeral 8 in Fig. The method according to claim 1,
Wherein the nozzles in the other plane can have the same value of fuel / air ratio or a varying value of fuel / air ratio. The method according to claim 1,
Wherein the fuel air nozzles in the coplanar can have the same fuel / air ratio or a varying fuel / air ratio. The method according to claim 1,
Characterized in that the nozzles are capable of greatly improving the ignition process of the tangentially directed combustion chamber and reducing the need to control a plurality of combustion devices since the adjacent nozzle flames are directed to adjacent nozzles in the plane, combustion chamber. 6. The method of claim 5,
Wherein the improved ignition process produces an essentially stable combustion device wherein the essentially stable combustion device reduces flame induced vibration and noise resulting from flame instability in partial load level operation and full load level operation, Annular combustion chamber. The method according to claim 1,
Wherein the tangential fuel-air nozzle arrangement improves mixing of reactants for effective combustion at very low load levels. The method according to claim 1,
Low-reactivity fuels, such as low BTU gases, are readily available and combustible in the combustion chamber due to increased flame stability. The method according to claim 1,
Wherein the vortex (an important result of the tangential fuel-air nozzle) is produced around the can-can line which promotes a stable flame at the combustion device outlet. The method according to claim 1,
The residence time required to burn the fuel-air mixture is reduced, resulting in a reduction in the combustion space, resulting in an increase in engine size (important for all applications), and consequently, weight to thrust ratio (important in aircraft gas turbine applications ) Of the cannula. The method according to claim 1,
Wherein a more uniform temperature distribution is achieved at the outlet of the combustion chamber to allow the combustion chamber to operate at a higher combustion (ignition) temperature without reducing the life of the combustion chamber and turbine components. . The method according to claim 1,
A can-annular combustion chamber characterized in that the ability to operate at a higher combustion temperature as described in claim 13 results in an increase in engine efficiency and power output, resulting in a reduction in the emission level of carbon dioxide. The method according to claim 1,
The can liner front wall may have at least one hole or nozzle that allows the compressor discharge air to pass through the liner at a speed that is less than the size of the nozzles referred to in claims 2 and 3. [ Annular combustion chamber. The method according to claim 1,
Characterized in that the radius and length of the can can vary in the longitudinal direction depending on the shape and size of the gas turbine engine. The method according to claim 1,
Cooling methods that can be used to cool the gas turbine components include impact cooling, exudative cooling, steam cooling, and the like. The method according to claim 1,
Wherein the nozzles sharing a common plane can be offset from a different set of nozzles in another plane by a circumferential angle around the can center line. The method according to claim 1,
Characterized in that the air passages (12, 14) are straight holes or can have bell mouth inlets and are formed using electrical discharge machining (EDM).

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