US3842597A

US3842597A - Gas turbine engine with means for reducing the formation and emission of nitrogen oxides

Info

Publication number: US3842597A
Application number: US00342041A
Authority: US
Inventors: F Ehrich
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1973-03-16
Filing date: 1973-03-16
Publication date: 1974-10-22
Anticipated expiration: 1991-10-22
Also published as: BE812374A; FR2221622B3; IT1007396B; DE2411950A1; FR2221622A1; JPS5025916A

Abstract

Means for reducing the formation and emission of nitrogen oxides in a gas turbine engine provide for bleeding and cooling a portion of the airflow pressurized by the compressor. The cooled compressor bleed airflow is then introduced into the primary combustion zone of the combustor in order to reduce the flame temperature effecting a reduction in the rate of formation of oxides of nitrogen.

Description

Unite Sttes Patent [191 Ehrich I45] Oct. 22, 1974 GAS TURBINE ENGINE WITH MEANS FOR REDUCING THE FORMATION AND EMISSION OF NITROGEN OXIDES [75] Inventor: Frederic Franklin Ehrich,

Murblehead, Mass.

[73] Assignec: General Electric Company, Lynn,

Mass,

[22] Filed: Mar. 16, 1973 [21] Appl. No.: 342,041

[52] US. Cl 60/226 R, 60/266, 60/39.65, 60/3966, 60/3974 R, 60/DIG. ll [51] Int. Cl F02c 7/l8,'F02c 7/22 [58] Field of Search 60/3965, 39.66, 39.67, 60/3923, 39.07, DIG. 10, 226 R, 262, 266, 39.74 R

[56] References Cited UNITED STATES PATENTS Bowdcn (ail/DIG. l0

3,528,250 9/1970 Johnson 60/261 3,648,45 3/1972 Bobo 60/3974 R 3,705,492 l2/l972 Vickcrs 60/DIG. l0

Primary ExaminerC. .I. Husar Assistant Examiner-Robert E. Garrett [57] ABSTRACT Means for reducing the formation and emission of nitrogen oxides in 21 gas turbine engine provide for bleeding and cooling :1 portion of the airflow pressurized by the compressor. The cooled compressor bleed airflow is then introduced into the primary combustion zone of the combustor in order to reduce the flame temperature effecting a reduction in the rate of formation of oxides of nitrogen.

5 Claims, 4 Drawing Figures GAS TURBINE ENGINE WITH MEANS FOR REDUCING THE FORMATION AND EMISSION OF NITROGEN OXIDES BACKGROUND OF THE INVENTION This invention relates generally to a gas turbine engine having means for significantly reducing the formation and emission of nitrogen oxides and, more particularly, to a gas turbine engine having means for cooling a portion of the compressor discharge air and then introducing the precooled air into the primary combustion zone to reduce the flame temperature and effect a corresponding reduction in the formation of nitrogen oxides.

The present day emphasis on the elimination of air pollution has resulted in a great deal of work and effort by aircraft gas turbine engine manufacturers who have succeeded in significantly reducing most forms of polluting emissions. Nitrogen oxides, however, are one form of pollutant emitted from a gas turbine engine which have not been satisfactorily reduced. Although it is not fully understood how oxides of nitrogen are formed, it is believed that such oxides are produced by the direct combination of atmospheric nitrogen and oxygen at the high temperatures occurring in primary combustion zones. The rates with which nitrogen oxides form depend upon the flame temperature and consequently a small reduction in flame temperature will result in a large reduction in the nitrogen oxides.

One proposed solution for reducing the emission of nitrogen oxides involves the introduction of more air to the critical primary combustion zone during peak periods of nitrogen oxide formation. The introduction of more air operates to reduce the fuel air ratio and corresponding flame temperature. The formation of oxides of nitrogen are generally most severe at the high power settings of the engine such as during takeoff. However, if the engine were designed to provide an excess of air during the peak power period at takeoff, then the fuel to air ratio at low power settings or at lightoff would likely be too lean to sustain or initiate combustion. Therefore, it becomes necessary to provide variable geometry apparatus to modulate the flow of air to the combustor in a manner which provides an excess of air during high power settings of the engine and reduces the airflow at low power setting to prevent the combustor flame from blowing out. Variable geometry, however, generally adds weight and complexity to a gas turbine engine and consequently is not an entirely satisfactory solution to the problem of eliminating nitrogen oxide emissions.

Another proposed solution for reducing the emission of nitrogen oxides relates to the effect of vitiating the combustion air with inert products such as recirculated cooled exhaust products or steam, wherein the flame temperature is reduced mainly by dilution and by the increased specific heat of the mixture. One difficulty with recirculating cooled exhaust products or steam, however, is that the combustor isv generally at a higher pressure than the exhaust necessitating the addition of a pump or blower to introduce the exhaust products directly into the combustor. The addition of a pump or blower again adds weight and complexity to the gas turbine engine with an attendant reduction in engine efficiency.

The addition of a pump or blower could be eliminated by reintroducing the exhaust products directly into the compressor inlet; however, this method also incurs certain disadvantages. Foremost is the increased risk that air contaminated by exhaust products will be circulated through the aircraft cabin from malfunctioning airconditioners which utilize compressor bleed air. Also, introducing exhaust products into the compressor inlet would accelerate the overall corrosion within the compressor.

Still another proposed solution for reducing the emission of nitrogen oxides relates to the introduction of a spray of water into the primary combustion zone in order to lower the flame temperature. This, however, necessitates that water storage tanks and pumps be added to the aircraft which could significantly increase the weight of the aircraft.

Therefore, it is a primary object of this invention to provide a simple means for reducing the formation and emission of nitrogen oxides from the combustor of a gas turbine engine without the use of variable geometry, blowers, pumps or water storage tanks.

It is also an object of this invention to provide a simple means for lowering the flame temperature within the primary combustion zone of a gas turbine engine in order to reduce the formation of nitrogen oxides therein.

It is a further object of this invention to provide a simple means for cooling a portion of the compressor discharge air and then introducing the precooled air into the primary combustion zone to quench the thermal reaction and reduce the formation of nitrogen oxides.

SUMMARY OF THE INVENTION Briefly stated, the above and other objects are achieved in the present invention by providing a gas turbine engine which may be of the bypass front fan type having a core compressor, combustor and turbine in serial flow relation, with a simplified means for'reducing the formation of oxides of nitrogen within the combustor. In the gas turbine engine, there is included means for bleeding a portion of the airflow pressurized by the compressor together with means for directing the airflow bled from the compressor into heat exchange relation with the bypass airflow in order to cool the compressor bleed airflow. Means are further included for introducing the cooled compressor bleed airflow into the primary combustion zone of the combustor in order to reduce the flame temperature and rate of formation of oxides of nitrogen.

BRIEF DESCRIPTION OF THE DRAWINGS While the specification concludes with claims distinctly claiming and particularly pointing out the invention described herein, it is believed that the invention will be more readily understood by reference to the discussion below and the accompanying drawings in which:

FIG. 1 is a side view, partly in cross-section, of a gas turbine engine embodying means for reducing the formation of nitrogen oxides within the engine combustor.

FIG. 2 is a partial cross-sectional view of the forward end of a gas turbine engine combustor embodying means for reducing the formation of nitrogen oxides therein.

FIG. 3 is a perspective view, partly cut away, of the nitrogen oxide reducing means of FIG. 2.

FIG. 4 is a side view, partly in cross-section, of a gas turbine engine embodying an alternate embodiment of the means of FIG. 1 for reducing the formation of nitrogen oxides within the engine combustor.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to FIG. 1, there is shown a gas turbine engine of the bypass front fan type comprising a core engine 12 having a compressor 14, a combustor 16, a gas generator turbine 18 for driving the compressor l4, and a power turbine 20 arranged in axially spaced serial flow relationship. The inner turbomachine, or core engine 12, is enclosed within a cylindrical casing 22 which terminates at its downstream end in an exhaust nozzle 24 through which the combustion products of the core engine 12 may be discharged to produce thrust. To provide additional thrust a fan 26 is mounted upstream of the core engine 12 and is driven by the power turbine 20. The fan 26 is comprised of a plurality of fan blades 28 which extend radially outward from a fan wheel 30, which is coupled for rotation with the power turbine 20 through the interconnecting shaft 31. The fan blades 28 extend radially across a bypass duct or passageway 32 defined between an outer cylindrical casing 34 and a bullet nose" 36 located upstream of the fan blades 28. Downstream of the fan blades 28, the passageway 32 is divided into two

passages

38 and 40 by the casing 22. Radially positioned between the casing 34 and casing 22 are a plurality of fan stator vanes 42 which are followed by a plurality of fan outlet guide vanes 44. Thus a portion of the air entering the passageway 32 flows through the fan blades 28 into the passageway 38, through the stator vanes 42, and through the outlet guide vanes 44 and thereafter exits through an outlet opening 46 formed by the casing 34 and the casing 22. Since this air is pressurized in flowing through the fan blades 28, it provides forward thrust to the turbofan engine 10.

The remainder of the air flowing through the passageway 32 and fan blades 28 enters the passageway 40. Located within this passageway 40 are a plurality of inlet guide vanes 48 for the core engine 12 which may be followed by a plurality of rotatable booster blades 50, extending from a disc 52 and coupled for rotation with the fan blades 28 by means of the disc 52 and a shaft 54. Located downstream of the booster blades is a row of stator vanes 56. Air passing through the stator vanes 56'next flows into the core engine 12 through passageway 58.

The gas turbine engine 10 may be either a high bypass ratio machine or a low bypass ratio machine wherein the bypass ratio refers to the ratio of the mass flow of fluid in the bypass passageway 38 to the mass flow in the core engine 12 (or passageway 40).

The compressor 14 discharges pressurized air through a plurality of circumferentially spaced apart outlet guide vanes 60 which extend radially between the walls of a diffuser passageway 62. A portion of the pressurized air exiting from the diffuser passageway 62 is bled through a conduit 64 to a heat exchanger 66. The heat exchanger 66 receives a portion of the cool bypass airflow from passageway 38 through inlet 68 whereupon the bypass airflow is directed through conduit 70 in heat exchange relation with the compressor discharge air and then discharged back into the bypass passageway 38 through outlet 72. Whereas the temperature and pressure of the compressor discharge air has been raised by the compressor 14, the heat exchanger 66 operates to reduce the temperature of that portion of the compressor discharge air bled through conduit 64.

The heat exchanger herein described is of conventional design and may take on many different forms with the only requirement being that the bypass airflow be used to cool a portion of the compressor discharge air. In practice the heat exchanger 66 may be annular with the fluids passed through the heat exchanger in paths as illustrated. If desired, however, alternative flow arrangements can be used for providing more effective heat transfer. For example, each of the streams could be directed through the heat exchanger twice in radial outflow passes. It is thus essential that there be two sets of independent passages disposed in heat exchange relationship within the heat exchanger, the bypass air flowing through one set of passages and the compressor discharge air flowing through the other set of passages.

The cooled compressor discharge air exiting from the heat exchanger 66 is directed to the combustor 16 by way of conduit 74 from whence it is dumped into the primary combustion zone 76 of the combustor 16. Fuel from a source of pressurized fuel (not shown) is introduced into the primary zone 76 of combustor 16 through fuel nozzle 78. Once the fuel is introduced into combustor 16, it may be ignited by igniter 80. The primary combustion zone 76 is generally defined within liner 81 in the area adjacent fuel nozzle 78.

In operation, combustor 16 will generally emit the following combustion products: carbon monoxide, carbon dioxide, water vapor, smoke and particles, unburned hydrocarbons, nitrogen oxides and sulfur oxides. Of these, carbon dioxide and water vapor may be considered normal and unobjectionable. Smoke can be a problem; however, it is generally controlled by design modifications in the primary combustion zone. Particulates which are one form of smoke are generally associated with coal burning plants and are not of immediate concern to gas turbines. Sulfur oxides can be limited by the careful selection of fuels low in total sulfur. This leaves carbon monoxide, unburned hydrocarbons and nitrogen oxides as the emissions of primary concern in the gas turbine engine.

As previously discussed, it is not fully understood how oxides of nitrogen are formed; however, it is believed that such oxides are produced by the direct combination of atmospheric nitrogen and oxygen at the high temperatures occurring in primary combustion zones. The presence of organic nitrogen in the fuel may also aid in the production of nitrogen oxides together with the atmospheric nitrogen. The rates with which nitrogen oxides form depend upon the flame temperature and, consequently, a small reduction in flame temperature will result in a large reduction in the nitrogen oxides.

Previously suggested means for reducing the maximum temperature in the primary zone of a gas turbine combustor have included schemes for introducing more air at the primary combustion zone, recirculating cooled exhaust products or steam back into the primary combustion zone and injecting a water spray back into the primary zone. All of these previously suggested methods, however, involve the addition of either complex or heavy hardware such as variable geometry, blowers, pumps or water storage tanks.

The invention herein described, however, provides for a heat exchanger 66 which precools a portion of the compressor discharge air and directs the precooled compressor discharge air back to the primary combustion zone 76 through conduit 74 in order to reduce the flame temperature thereby inhibiting the formation of oxides of nitrogen. Although it is preferred that a portion of the compressor discharge air be precooled, it would not be outside the scope of invention to bleed and cool a portion of the compressor interstage air and redirect the precooled compressor interstage air back to the primary zone of the combustor.

Referring now to FIGS. 2 and 3 where like numerals refer to previously described elements, there is shown one arrangement by which the precooled compressor discharge air may be introduced into the primary combustion zone 76. An outer shell 82 is provided to enclose the liner 81 and to cooperate therewith to form passageways 84, 86 surrounding the liner 81. As will be understood, the passageways 84, 86 are adapted to deliver a flow of pressurized air from the compressor 14 through suitable apertures or louvers 88. In this manner, the passageways 84, 86 act to both cool the liner 81 and to provide dilution air to the gaseous products of combustion formed within the primary combustion zone 76.

The upstream end of liner 81 is adapted to function as a flow splitter to divide the pressurized air delivered from the compressor 14 between the passages 84, 86 and an upstream end opening 90. Located within the liner 8] is a conventional fuel injecting apparatus shown generally at 92 which may be of the atomizing type as is well known in the art. Pressurized fuel may be supplied to the fuel injection apparatus 92 through a conduit 94 which extends through the outer shell 82 and communicates with a source of pressurized fuel (not shown).

Surrounding the fuel injection apparatus 92, there is a housing 96 which comprises an involute outer wall 97 and generally planar, spaced upstream and downstream

annular end walls

98 and 100, respectively, which are peripherally joined to the outer wall 97. A plurality of circumferentially spaced apart swirl vanes 101 extend between the upstream and downstream end walls in spaced apart relation to the outer wall 97. The housing 96 defines a conventional spin chamber 102 having an annular outlet 104 which surrounds the fuel injection apparatus 92. The outer wall 97 is of spiral shape with progressively decreasing radius from an inlet 106 to a terminal edge or lip 108 which in part defines the opening from inlet 106 to the spin chamber 102. The inlet 106 is in direct flow communication with the conduit 74 for receiving a flow of precooled compressor discharge air directly from the heat exchanger 66. In this manner precooled compressor discharge air may be directed from inlet 106 into the spin chamber 102 in a circular motion of ever decreasing radius so as to generate a vortical flow around the fuel injection apparatus 92.

In operation, a portion of the compressor discharge air is bled through inlet 110 of conduit 64 and directed to the heat exchanger 66 whereupon it is circulated in heat exchange relationship with the fan bypass airflow, and then directed to the spin chamber 102 via conduit 74. The precooled compressor discharge air is dumped into the primary combustion zone 76 in a vortical flow from outlet 104 of the spin chamber 102 whereupon it reduces the flame temperature in the primary combustion zone 76 and inhibits the formation of oxides of nitrogen.

Turning now to FIG. 4 where like numerals refer to previously described elements, there is shown a modified gas turbine engine 10 which is also of the bypass front fan type, and which includes an extended cowling or casing 34' The cowling 34 is spaced apart from the core engine 12 so as to define a bypass passageway 38' which extends substantially the length of the core engine 22 and terminates in an outlet opening 46'. A heat exchanger 66 is disposed within the passageway 38' and receives a portion of the pressurized compressor discharge air exiting from the diffuser passageway 62 through conduit 64. The compressor discharge air is directed into heat exchange relation with the comparatively cooler fan air flow in passageway 38, and then returned to the combustor 16 through conduit 74' in order to reduce the flame temperature and inhibit the formation of oxides of nitrogen as previously discussed.

From the foregoing it will be appreciated that the formation of nitrogen oxide in the combustor of a gas turbine engine are inhibited by simple means without the addition of variable geometry, blowers, pumps or water storage tanks. It will also be appreciated that a portion of the compressor discharge air which is bled Y and cooled by the heat exchanger 66 may be directed to the turbine for cooling purposes. Accordingly, while preferred embodiments of the present invention have been depicted and described, it will be appreciated by those skilled in the art that many modifications, substitutions, and changes may be made thereto without departing from the inventions fundamental scheme.

Having thus described one embodiment of the invention, what is desired to be secured by Letters Patent is as follows:

1. In a gas turbine engine having a compressor, combustor and turbine in serial flow relation, means for reducing the formation of oxides of nitrogen within the combustor comprise:

means within said combustor defining a primary combustion zone, a secondary combustion zone, and a dilution zone;

means for bleeding a portion of the airflow pressurized by the compressor;

means for cooling the airflow bled from the compressor;

means for directing the unbled portion of compressor discharge airflow to the primary combustion, secondary combustion, and dilution zones of the combustor in order to support combustion in the primary combustion zone;

and means for introducing the cooled compressor bleed airflow into the primary combustion zone of the combustion in order to reduce the flame temperature.

2. In a gas turbine engine of the bypass front fan type having a core compressor, combustor and turbine in serial flow relation, means for reducing the formation of oxides of nitrogen within the combustor comprise:

means for bleeding a portion of the airflow pressurized by the compressor; means for directing the airflow bled from the compressor into heat exchange relation with the bypass,

airflow in order to cool the compressor bleed airflow;

means for directing the unbled portion of compressor discharge airflow to the primary combustion, secondary combustion, and dilution zones of the combustor in order to support combustion in the primary combustion zone, and

means for introducing the cooled compressor bleed airflow into the primary combustion zone of the combustor in order to reduce the flame temperature.

3. The gas turbine engine of claim 2 wherein the means for directing the compressor bleed airflow into heat exchange relation with the bypass airflow includes a heat exchanger having at least two passages arranged in heat exchange relationship wherein one of the passages receives and discharges bypass airflow and the other passage receives and discharges compressor bleed airflow.

4. The gas turbine engine of claim 2 wherein the means for reintroducing the cooled compressor bleed airflow back into the primary zone of the combustor includes:

a spin chamber in general surrounding relation to a fuel injection apparatus wherein the spin chamber receives the cooled compressor bleed airflow and directs the airflow in a circular motion of ever decreasing radius so as to generate a vortical flow around the fuel injection apparatus, and,

conduit means communicating from the spin chamber to the heat exchange means for directing the cooled compressor bleed airflow therebetween.

5. The gas turbine engine of claim 4 wherein the spin chamber includes:

an involute outer wall and generally planar spaced upstream and downstream annular end walls, the outside edges of which are peripherally joined to the outer wall and the inside edges of which define an annular outlet which surrounds the fuel injection apparatus; plurality of circumferentially spaced apart swirl vanes extending between the upstream and downstream end walls in spaced apart relation to the outer wall; and an inlet in direct communication with the conduit means for receipt of the cooled compressor bleed airflow.

Claims

1. In a gas turbine engine having a compressor, combustor and turbine in serial flow relation, means for reducing the formation of oxides of nitrogen within the combustor comprise: means within said combustor defining a primary combustion zone, a secondary combustion zone, and a dilution zone; means for bleeding a portion of the airflow pressurized by the compressor; means for cooling the airflow bled from the compressor; means for directing the unbled portion of compressor discharge airflow to the primary combustion, secondary combustion, and dilution zones of the combustor in order to support combustion in the primary combustion zone; and means for introducing the cooled compressor bleed airflow into the primary combustion zone of the combustion in order to reduce the flame temperature.

2. In a gas turbine engine of the bypass front fan type having a core compressor, combustor and turbine in serial flow relation, means for reducing the formation of oxides of nitrogen within the combustor comprise: means within said combustor defining a primary combustion zone, a secondary combustion zone, and a dilution zone; means for bleeding a portion of the airflow pressurized by the compressor; means for directing the airflow bled from the compressor into heat exchange relation with the bypass airflow in order to cool the compressor bleed airflow; means for directing the unbled portion of compressor discharge airflow to the primary combustion, secondary combustion, and dilution zones of the combustor in order to support combustion in the primary combustion zone, and means for introducing the cooled compressor bleed airflow into the primary combustion zone of the combustor in order to reduce the flame temperature.

4. The gas turbine engine of claim 2 wherein the means for reintroducing the cooled compressor bleed airflow back into the primary zone of the combustor includes: a spin chamber in general surrounding relation to a fuel injection apparatus wherein the spin chamber receives the cooled compressor bleed airflow and directs the airflow in a circular motion of ever decreasing radius so as to generate a vortical flow around the fuel injection apparatus, and, conduit means communicating from the spin chamber to the heat exchange means for directing the cooled compressor bleed airflow therebetween.

5. The gas turbine engine of claim 4 wherein the spin chamber includes: an involute outer wall and generally planar spaced upstream and downstream annular end walls, the outside edges of which are peripherally joined to the outer wall and the inside edges of which define an annular outlet which surrounds the fuel injection apparatus; a plurality of circumferentially spaced apart swirl vanes extending between the upstream and downstream end walls in spaced apart relation to the outer wall; and an inlet in direct communication with the conduit means for receipt of the cooled compressor bleed airflow.