Reverse-flow gas turbine combustion system

A gas turbine combustion system for burning air and fuel into exhaust gases. The gas turbine combustion system may include a combustor, a turbine nozzle integral with the combustor for providing the air to the combustor, and a fuel injector for providing the fuel to the combustor. The turbine nozzle and the fuel injector are positioned within the combustor such that a mixture of the air and the fuel flows in a first direction and the exhaust gases flow in a second direction.

TECHNICAL FIELD

The present invention relates generally to gas turbine combustion engines and relates more particularly to a reverse-flow gas turbine combustion system with an integral first stage turbine nozzle.

BACKGROUND OF THE INVENTION

Generally described, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Known turbine engines have developed into highly complex and sophisticated devices. For example, known turbine combustor systems alone may require more than 14,000 parts.

Another issue with known gas turbine engines is promoting operation at high efficiency without producing undesirable air emissions. The primary air emissions usually produced by gas turbine engines include nitrogen oxides (NOx). Lean premixed combustors also have a susceptibility to dynamics that can cause catastrophic damage or accelerated wear in current dry, low emissions combustion systems.

There is a desire, therefore, for a gas turbine engine with increased reliability, efficiency, and performance. Preferably, the number and size of the parts of the gas turbine engine can be reduced while maintaining or improving on performance and emissions output.

SUMMARY OF THE INVENTION

The present application thus describes a gas turbine combustion system for burning air and fuel into exhaust gases. The gas turbine combustion system may include a combustor, a turbine nozzle integral with the combustor for providing the air to the combustor, and a fuel injector for providing the fuel to the combustor. The turbine nozzle and the fuel injector are positioned within the combustor such that a mixture of the air and the fuel flows in a first direction and the exhaust gases flow in a second direction.

The second direction may be the reverse direction to the first direction. The combustor may include a reverse flow annular combustor. The turbine nozzle may include a stage one nozzle. The combustor may include one or more reaction zones and one or more stagnation points. The turbine nozzle and the fuel injector may be positioned about a tube within the combustor. The combustor may include two reverse turns for the exhaust gases. A number of fuel injectors may be used. An impingement sleeve may be positioned about the combustor.

An airflow channel may be positioned about the combustor. Air may flow through the airflow channel in a reverse direction to the exhaust gases in the second direction. The airflow channel may lead to one or more intake channels positioned within the combustor. The fuel injector may be positioned within one of the intake channels. The airflow channel may lead to one or more turbine plenums. An impingement/film cooling sleeve may be positioned about the airflow channel.

The present application further describes a gas turbine combustion system for burning air and fuel into exhaust gases. The gas turbine combustion system may include a reverse flow combustor, a turbine nozzle integral with the combustor for providing the air to the combustor, a fuel injector for providing the fuel to the combustor, and a cooling airflow channel positioned about the combustor. The cooling airflow channel may include an impingement sleeve positioned about the combustor. Air flows through the airflow channel in a reverse direction to an exhaust gases direction.

The present application further describes a method for burning air and fuel into exhaust gases in a gas turbine combustion system. The method may include injecting an air stream into a combustor in a first direction, injecting a fuel stream into the combustor in the first direction, reacting the air stream and the fuel stream in a reaction zone to create the exhaust gases, and reversing the flow of the exhaust gases in a second direction so as to exit the combustor.

These and other features of the present invention will become apparent to one of ordinary skill in the art upon review of the following detailed description of the preferred embodiments when taken in conjunction with the drawings and appended claims.

DETAILED DESCRIPTION

Referring now to the drawings, in which like numerals refer to like elements throughout the several views,FIGS. 1 and 2show a combustion system100as is described herein. The combustion system100includes a combustor110. As will be described in more detail below, the combustor110may be a reverse flow annular combustor. The combustor110includes an annular combustor liner120. The liner120shapes the flow field to provide the reverse flow.

The combustion system100includes a turbine nozzle130that is integral with the combustor110. In this example, the turbine nozzle130is a stage-one nozzle. Other designs also may be used herein. The turbine nozzle130supplies air for combustion and for cooling. The nozzle130also may be cooled via a flow of the fuel.

The combustion system100further includes a fuel injector140. In this example, the fuel injector140is positioned within the turbine nozzle130and is integral therewith. The fuel injector140may be in communication with a manifold150so as to deliver the fuel thereto.

As is shown inFIG. 2, the fuel and the air are injected into a reaction zone160within the combustor110. The fuel and the air are injected in a direction opposite to the direction that the hot combustion products exit, thus creating a stagnation point165in the reaction zone160. This stagnation point165creates a high shear counter flow between the incoming fresh mixture and the outgoing combustion products. In this example, the fuel and the combustion air are injected in a co-flow manner via the concentric tubes of the turbine nozzle130and the fuel injector140. Concentric elongated ducts and similar structures also may be used. The fuel and the air streams also can be premixed prior to injection.

FIG. 3show a combustion system200that is a further embodiment as is described herein. The combustion system200provides a single tube, vane fed injection system. The combustion system200includes the combustor110, the turbine nozzle130, and the fuel injector140. The fuel injector140is positioned within a tube205. The turbine nozzle130also is in communication with the tube205so as to provide a co-flow as described above. The combustion process produces the reaction zones160, the stagnation points165, and the reverse flows as described above.

The combustion system200also provides for air/fuel cooling. The combustion system200may include a liner210. The liner210is similar to the liner120described above but may be cooled by impingement and/or convection. The liner210may be positioned within an impingement sleeve220so as to create an airflow channel230. Air passes through the impingement sleeve220and into the airflow channel230. The air is directed over the liner210about the reaction zones160for cooling of both the liner210and the turbine nozzle130. Alternatively, air also may be delivered from the gas turbine compressor (not shown) and may be directed into the airflow channel230. Flow turning or swirl vanes may be used in the airflow channel230. After the air passes through the airflow channel230, the air passes through the turbine nozzle130at the hub side and the shroud side so as to cool the vanes and airfoil surfaces of the nozzle130by enhanced convection. The fuel supplied to the turbine nozzle130also may be used to cool parts of the nozzle130prior to being injected into the reaction zones160.

The tube205, also known as a vane, can take any desired shape.FIGS. 9A-9Eshow various vane configurations. Each tube205may be designed for enhanced interaction between the combustion air and fuel.FIGS. 9D and 9Eshow concentric tube type geometries.

FIG. 4shows a combustion system250that is a further embodiment as is described herein. The combustion system250provides a forward mounted fuel injector. The combustion system250includes the combustor110, the turbine nozzle130, the liner210, the impingement sleeve220, and the airflow channel230as described above. In this example, however, a fuel injector255is now positioned at the front of the combustor110. As before, the fuel and the air are injected into the combustor110and form the reaction zones160. In this example, the exhaust gases take two reverse turns before exiting the combustor110. The incoming air again travels in a direction opposite the exhaust flow so as to provide cooling.

FIG. 5shows a combustion system260that is a further embodiment as is described herein. The combustion system260provides for a multi-tube, aft-fed injection system. The combustion system260includes the combustor110, the turbine nozzle130, the liner210, the impingement sleeve220, and the airflow channel230as described above. In this example, the fuel injector140is positioned on one side of the combustor110while the airflow channel230leads to a pair of intake air channels270positioned about the combustor110. The fuel flows and the air flows are directed towards and ignite within the reaction zones160. The dual fuel flows also may be used herein.

FIG. 6shows a combustion system280that is a further embodiment as is described herein. The combustion system280provides for a single tube, aft fed injection system. The combustion system280includes the combustor110, the turbine nozzle130, the liner210, the impingement sleeve220, and the airflow channel230as described above. In this example, a portion of the air entering the impingement sleeve220exits via a turbine plenum290while the remaining air is directed towards the intake air channel270. The fuel injector140is positioned within the intake air channel270so as to provide the single tube injection process.

FIG. 7shows a combustion system300that is a further embodiment as is described herein. The combustion system300provides for a wall fed injection system. The combustion system300includes the combustor100, and the turbine nozzle130as described above. The combustion system300also includes an inner liner310. The inner liner310may not extend all the way about the combustor110, but may terminate before the reaching the front end. The impingement sleeve220may surround the inner liner310. The impingement sleeve220creates the airflow channel230about the inner liner310. One end of the impingement sleeve220may be surrounded by an impingement/film cooling sleeve320. The fuel injector140may be positioned at about the end of the inner liner310. Air may enter from the airflow channel230from the impingement sleeve220or otherwise and may merge with air from the impingement/film cooling sleeve320.

FIG. 8shows a combustion system330that is a further embodiment as is described herein. The combustion system330provides for a staged wall feed injection system. The combustion system330includes the combustor110and the nozzle130as is described above. The combustion system330also includes the inner liner310, the impingement sleeve220, the airflow channel230, and the impingement/film cooling layer320ofFIG. 7. The combustion system330also includes a second inner liner340and one or more secondary fuel supplies350. In this example, the fuel injector140is again positioned at about the end of the inner liner310. One of secondary fuel supplies350is positioned at about the end of the second inner liner340on one side of the combustor110while a further secondary fuel supply350may be positioned on the other side. The fuel and air from the second inner liner340may mix with the exhaust products at the other end of the combustor110.

FIG. 10shows a further embodiment as is described herein, a combustion system400. The combustion system400is a retrofitted can design. The combustion system400may use an existing transition piece410with a reverse flow combustor concept. An outer liner420replaces the existing cap and flow sleeve. Fuel is directed from the manifold150through an inner liner430. A swozzle440is positioned on the other end of the liner420for injecting premixed air/fuel. (The term “swozzle” is a combination of the words “swirler” and “nozzle”.) The combustion system400thus can be incorporated into many existing devices.

The combustion systems described herein provide significant benefits in terms of reduced physical size, complexity, part count, costs, and exhaust emissions relative to current combustion systems. These benefits are obtained in part by the close integration of the turbine nozzle130and the combustor110and by using the combustion airflow and fuel flow as a means of cooling the turbine nozzle130. At the same time, system efficiencies as a whole are improved by using the turbine nozzle130for preheating the incoming fuel and air. The use of the reverse flow and the stagnation point165provides efficient fuel/air mixing, mixing of exhaust gases and incoming air, preheating of fuel and air streams, and short reaction zone residence time.

The combustion systems described herein also provide low NOx emissions without complex fuel/air mixtures and/or fuel staging. Single digit NOx ranges are possible with both gas and liquid fuels, potentially less than five (5) parts per million NOx. The small reaction zones160also promote “flameless” combustion (or widely distributed reaction zones), further lowering NOx emissions. Combustion noise also may be reduced. Damaging combustion noise also may be reduced or eliminated, improving combustion system durability and reliability.

With respect to size, about eighty percent (80%) of the combustion parts may be eliminated. Generally described, current end covers, casings, flow sleeves, inner/outer crossfire tubes, caps, liners, transition pieces, and individual fuel injectors may be eliminated. Fewer seals also may be required. In other words, while a typical combustion system may include more than 14,000 parts, the combustion systems described herein may have about 500 parts. The combustion systems described herein thus providing cost savings and size reduction. Existing turbine equipment also may be retrofitted.

Another benefit found herein is in the reduction of the surface area and the improvement in the surface to volume ratio for the liner210and other elements. This reduces the amount of dedicated liner cooling required and the associated metal temperatures. This further improves the overall durability of the elements described herein.

It should be apparent that the foregoing relates only to the preferred embodiments of the present invention and that numerous changes and modifications may be made herein without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.