JPH06502240A - Gas turbine combustion chamber and its operation method - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed description of the invention] Gas turbine combustion chamber and its operation method The present invention provides a gas turbine combustion chamber and a method of operating a gas turbine engine combustion chamber. Regarding.

To meet the emission levels required for industrial low-injection gas turbine engines, The volume of the gin combustion chamber is increased. Currently, industrial gas turbine engines are Alternatively, a can-annular combustion chamber is used. Same axial direction to increase the volume of the combustion chamber assembly while incorporating the combustion chamber in the length. For this purpose, it is necessary to use several tubular combustion chambers, the longitudinal axis of which They are arranged in the radial direction. The inlets of the tubular combustion chambers are located at their radially outer ends. Yes, the transition duct directs the hot exhaust gases to the turbine section of the gas turbine engine. Connect the outlet of the tubular combustion chamber with a row of nozzle guides for axial discharge do.

It also minimizes the amount of nitrogen oxides generated to meet emission level requirements. Staged combustion is required to achieve this. Currently, emission level requirements are 25 or less per million nitrogen in the bottle exhaust. Nitrogen oxide release The basic method to reduce the combustion reaction temperature is to lower the combustion reaction temperature, which , all of the fuel and combustion air must be premixed before combustion occurs. nitrogen acid NOx emissions are reduced by a method that uses 22-stage fuel injection.

British Patent No. 1489339 discloses two-stage fuel injection to reduce NOx. Show. In staged combustion, both stages of combustion produce lean combustion and nitrogen oxides (N The aim is to provide the low combustion temperature necessary to minimize the child The term "lean burn" means that the air-to-fuel ratio is low, i.e. below the stoichiometric ratio. Refers to the combustion of fuel in air.

The present invention provides a new gas turbine room and a new gas turbine engine room operating system. The purpose is to provide a method.

Accordingly, the present invention provides a first air intake device, a first fuel injector device and a first air intake device. a first fuel-air mixing region, the first fuel-air mixing region having at least one ring; formed by a shaped wall and an upstream wall connected to the upstream end of the tubular wall, with the upstream being less the first air intake device has at least one opening; a radial flow swirler and at least one second radial flow swirler; each first radial flow swirler having a first fuel air flow through said opening. Each second radial flow swirler is adapted to supply air to the mixing region. is adapted to supply air through the opening to the first fuel-air mixing region. and each first radial flow swirler has a respective second radius with respect to the axis of the combustion chamber. each first radial flow vortex disposed axially downstream of the directional flow vortex; The winders swirl the air in an opposite direction to each second radial flow winder. and the first fuel injector device has a first radial flow vortex. a second radial passageway formed between each vane of the vessel; supplying fuel to at least one passageway formed between each vane of the flow swirler; A combustion chamber for a gas turbine engine is provided.

Preferably, there is at least one pilot fuel injector, with each pilot fuel injector having at least one pilot fuel injector. The first fuel injector has an opening for supplying fuel to the first fuel-air mixing region. It is consistent with each one.

Preferably, the first fuel injector device includes a first radial flow swirler) It is designed to supply fuel to all the passages formed between the blades.

The first fuel injector device is shaped between the vanes of the second radial flow swirler. It is designed to supply fuel to all passages created.

Preferably, the first fuel injector injects fuel into a radially outer region of the passage between the vanes. It is designed to supply

The first fuel injector device has a hollow cylindrical shape extending axially with respect to the combustion chamber. a member, the cylindrical member having an axis along the cylindrical member for injecting fuel into the passageway; It has a plurality of openings spaced apart in the direction.

The openings are adapted to direct fuel radially inward.

The first fuel injector is adapted to inject gaseous fuel or vaporized liquid fuel. ing.

The pilot fuel injector is designed to inject gas or liquid fuel. ing.

The combustion chamber is tubular and has one opening in its upstream wall.

The combustion chamber includes a second air intake device, a second fuel injector device and a second combustion chamber. the second fuel-air mixing region is tubular and the first fuel-air mixing region is tubular; surrounding the air mixing region, the second fuel air mixing region being radially defined by the second tubular wall. a second fuel injector device having a second fuel injector device formed at a directionally outer end; A second fuel-air mixture is adapted to supply fuel to the upstream end of the air-mixing region. The region is in fluid communication with the interior of the combustion chamber downstream of the first fuel-air mixing region at its downstream end. We are in communication.

The second air intake is downstream of the first air intake device.

A second fuel-air mixing region is formed at a radially inner end thereof by a third tubular wall. ing.

The tubular wall includes a first portion forming a first fuel-air mixing region and a portion downstream of the first portion. a second portion having a larger diameter; and a third portion interconnecting the first and second portions. It has a truncated conical part.

The third conical section acts as a plurality of jets downstream toward the centerline of the combustion chamber. delivering a second fuel-air mixture from the second fuel-air mixture to the second fuel-air mixing region; It has a plurality of openings equally spaced in the circumferential direction such that the openings are spaced apart from each other.

The opening is a slot.

The downstream end of the second tubular wall is secured to the third conical portion I of the tubular wall.

Cooling air is directed into a tubular chamber formed between the tubular wall and the third tubular wall.

The second fuel injector fungus has a plurality of circumferentially equally spaced injectors. It has a vector.

The second combustion injector device is configured to inject gaseous fuel or vaporized liquid fuel. It has become.

The downstream end of the first portion of the tubular wall decreases in diameter to the throat.

The combustion chamber includes a third air intake device, a third fuel injector device and a third combustion chamber. a third fuel-air mixing region is tubular and a second combustion region; and a third fuel-air mixing region is bounded by a radially outer end by a fourth tubular wall. A third fuel injector device is formed at the end and includes a third fuel-air mixing region. The third fuel-air mixing region is adapted to supply fuel to the upstream end of the fluid flow with a third combustion zone inside the combustion chamber downstream of the second combustion zone at the downstream end; It's communicating.

The tubular wall is of a larger diameter downstream of the second section and defines a third combustion zone. a fourth section forming a zone; and a fifth truncated cone interconnecting the second and fourth sections. It has a part.

The downstream end of the second portion of the tubular wall decreases in diameter to the throat.

The third air intake is downstream of the second air intake device.

A third fuel-air mixing region is formed at a radially inner end thereof by a fifth tubular wall. ing.

The fifth conical section is configured as a plurality of jets in a downstream direction toward the centerline of the combustion chamber. Circumference adapted to deliver a third fuel-air mixture from a third fuel-air mixing region It has a plurality of openings equally spaced in the direction.

The opening is a slot.

The downstream end of the fourth tubular wall is secured to the fifth conical portion of the tubular wall.

The third fuel injector device includes a plurality of circumferentially equally spaced injectors. It has a vector.

The third combustion injector device is configured to inject gaseous fuel or vaporized liquid fuel. It has become.

The pyro is used from the start of operation of the gas turbine engine until the specified output power level is obtained. The fuel from the set fuel injector is supplied only to the first fuel-air mixing region and the predetermined to flow the first fuel-air mixing region for an output power level greater than the level of at least one passageway formed between the vanes of the first radial flow swirler; and at least one passageway formed between the vanes of the second radial flow swirler. from a first fuel injector device and below the first fuel-air mixing region. a second fuel injector device for directing a second fuel flow inside the combustion chamber; Simultaneously supplies fuel to the air mixing area.

The pyro is used from the start of operation of the gas turbine engine until the specified output power level is obtained. The fuel from the set fuel injector is supplied only to the first fuel-air mixing region and the predetermined to flow the first fuel-air mixing region for an output power level greater than the level of A first radial flow vortex is formed between the vanes of the first radial flow swirler. and at least one passageway formed between the vanes of the second radial flow swirler. from a first fuel injector device and below the first fuel-air mixing region. a second fuel-air mixing region for flowing the flow to a second combustion region inside the combustion chamber; at an output power level greater than the second predetermined level and at a predetermined temperature. a third combustion chamber inside the combustion chamber downstream of the second combustion zone at an atmospheric temperature greater than A third fuel air mixing region is provided with fuel for flow to the sintering region.

The predetermined output power level is 35% to 40% power.

a gas turbine engine combustion chamber having a first combustion zone and interconnecting the first combustion zone; and a second combustion zone disposed downstream thereof, and interconnected to the second combustion zone; a third combustion zone located downstream thereof; and a third combustion zone that supplies the first air to the first combustion zone. a first air intake device for supplying fuel to the first combustion zone; a pilot fuel injector arrangement and a first combustion zone for supplying fuel to the first combustion zone; a fuel injector device and a second fuel injector device for supplying a second air to the second combustion zone. a second air intake device and a second fuel intake for supplying fuel to the second combustion zone; an injector device and a third air intake for supplying third air to the third combustion zone a third fuel injector device for supplying fuel to the third combustion zone; and has.

The second combustion zone is interconnected to the first combustion zone by a throat.

The third combustion zone is interconnected to the second combustion zone by a second throat. There is.

The invention will be described more fully by reference to the accompanying drawings, in which: FIG.

FIG. 1 shows a gas star having a combustion chamber assembly and fuel injector according to the present invention. It is a sectional view of a bin engine.

FIG. 2 is an enlarged longitudinal section through the combustion chamber shown in FIG. 1;

FIG. 3 shows a further enlarged longitudinal view through the upstream side of the combustion chamber assembly shown in FIG. FIG.

FIG. 4 is a cross-sectional view in the direction of arrow GG in FIG.

FIG. 5 is a cross-sectional view in the direction of arrow HH in FIG. 3. FIG.

Figure 6 is a graph of the percentage of base load fuel flow for the combustion chamber shown in Figure 3 versus the percentage of combustion chamber. It is f.

FIG. 7 shows an enlarged longitudinal view through the upstream end of another combustion chamber assembly according to the present invention. FIG.

FIG. 8 is an enlarged longitudinal section through the upstream end of another combustion chamber according to the invention; Ru.

The industrial gas turbine engine 10 shown in FIG. 12, compressor section 14, combustion chamber assembly 16, turbine section 18, power It has a turbine section 20 and an exhaust pipe 22. The turbine section 18 may include one or more to drive the compressor section 14 via more than one shaft (not shown). It has become. Power turbine portion 20 drives a generator 26 via shaft 24. It has become so. However, the power turbine section 20 may be used for other purposes. It is adapted to provide a driving body. All operations of the gas turbine engine 10 are This is obvious and requires no further explanation. The combustion chamber assembly 16 is shown in FIG. This is shown in more detail in FIGS. The plurality of compressor outlet vanes 28 provided at the downstream end of the compressor section 14 in the axial direction; An inner annular wall 30 forming the inner surface of the tubular chamber 34 is fixed at the radially inner end in the downstream direction. It will be done. A diffuser is formed between the annular wall 32 and the upstream portion of the inner annular wall 30.

The downstream end of the inner annular wall 30 is hot from the combustion chamber assembly 16 to the turbine section 18. It is fixed to the radially inner end of a row of nozzle guide vanes 90 that direct the gas.

Combustion chamber assembly 16 includes a plurality of equally circumferentially spaced annular combustion chambers 36. has. The axis of the annular combustion chamber 36 is arranged to extend approximately in the radial direction. Ru. The inlets of the annular combustion chambers 36 are at their radially outer ends, and their outlets are at their radially outer ends. , at their radially inner ends. Each of the annular combustion chambers 36 is located upstream of the annular wall 37. It has an upstream wall 44 fixed at the end. The first upstream end of the annular wall 37 is connected to the first fuel source. The second downstream portion 42 of the annular wall forms a third portion 40 and an air mixing region 64 . is interconnected with the first portion 38 by. The second portion 42 of the annular wall 1 has a larger diameter than portion 38 of 1. The third portion 40 is frustoconical.

A plurality of equally circumferentially spaced transition ducts 46 are provided. . Each of the transition ducts 46 has a circular cross section at its upstream end. Tiger The upstream end of each of the combustion chambers 46 is downstream of the corresponding end of the annular combustion chamber 36. arranged coaxially around the end, each of the transition ducts 46 has a nozzle. Connect the annular portion of the guide vane 90 and seal it.

A plurality of cylindrical casings 48 are provided, each cylindrical casing 48 having an annular combustion They are arranged coaxially around each end of chamber 36. Each cylindrical casing 48 is It is fixed in each hole 52 on the annular engine casing 50. A number of chambers 54 each A tubular combustion chamber 36 and its respective cylindrical casing 48 are formed between them.

The upstream end of each transition duct 46 is connected to the upstream end of the transition duct 46. The engine casing 50 includes a plurality of pairs of brackets. 58. Each bracket 56 is rotated by a pin 60 to each of the pair of brackets. fixed so that it can be rotated. This is British Patent No. 901 filed on September 1, 1990. No. 9089.3 is more fully disclosed.

The upstream wall 44 of each of the tubular combustion chambers 36 connects the first fuel and air mixing region 64 with a static electric current. and an opening 62 that allows the supply of fuel. multiple first radial flows; Swirls are provided, each first radial flow vortex being L of each tubular combustion chamber 36. It is arranged coaxially with the opening 62 in the flow wall 44 . Similarly, multiple second radial directions directional flow swirlers are disposed, each second radial flow swirler forming a respective second radial flow swirler in each tubular combustion chamber. The openings 62J of the upstream wall 44 of 36 are arranged coaxially. first radial A flow swirler is located axially downstream of the second radial flow swirler. Ru.

Each first radial flow volute includes a first side plate 66, a second side plate 68 and a plurality of It has a first wing 70. The first side plate 66 is coaxially connected to the opening 62 in the upstream wall 44. The side plate 66 is secured to the upstream wall 44 with a central opening located therein.

A first wing 70 extends axially between the first and second side plates 66 and 68 and This is fixed. A number of passageways 72 are formed between the first airflow wings 70. ing. Each second radial flow swirler includes a plurality of second vanes 74 and a third side plate. It has 76. The second wing 74 is arranged between the second side plate 68 and the third side plate 76 in the axial direction. It extends in the direction. A second side plate 68 is disposed coaxially with the opening 62 within the upstream wall 44. It has a central opening 62 extending axially downstream. many A passageway 80 is formed between the second wings 74 for airflow. first and second half The first and second wings 70, 74 of the radial flow swirler can be seen from FIGS. They are arranged in such a way that they spiral in opposite directions. First tubular air intake 82 is a body surrounding the radially outer end of the first side plate 66 and the outer end of each cylindrical casing 48; It is formed in the axial direction between the plate 84 and the plate 84 . Multiple pilot fuel injectors 86 are provided, and each pilot fuel injector 86 injects fuel through opening 62. at one opening of the tubular combustion chamber 36 for supplying the first fuel-air mixing region 64. arranged coaxially. A plurality of first main fuel injectors 88 are connected to the tubular combustion chamber 36 . provided for each. Each of the first fuel injectors 88 is connected to the tubular combustion chamber 36. It has a hollow cylindrical member extending axially relative thereto. Each hollow cylindrical member passes through the third side plate 76 and the second side plate 68 in the axial direction, and passes through the first side plate 66. It is placed in a blind hole. The hollow cylinder member is one between the second vanes 74 and passes axially through one passage between the first vanes 70. It's becoming a sea urchin. A hollow cylindrical member is provided in the radially outer region of the passage 872.80. a first radial flow swirler assembly disposed toward and for injecting fuel into the first radial flow swirler assembly; axially spaced openings 90 for directing the fuel into a second radial flow vortex. It has axially spaced openings 92 that inject into the winder assembly. this Openings 90 and 92 are arranged to allow fuel to flow radially.

The second fuel and air mixing region 94 is the first fuel and air mixing region 6 of each combustion chamber 36. It surrounds 4. It also includes each second tubular fuel wall 96 and third tubular wall 9 8. A second tubular wall 96 defines one half of the second fuel and air mixing region. forming the radial outer end. The third tubular wall 98 is connected to the second fuel and air mixing region 9 form the radially inner end of 4. The axial upstream end 100 of each third tubular wall 98 is It is secured to the first side plate 66 of the first radial flow volute of each tubular combustion chamber. No. The second tubular air intake 102 directs air to the second tubular fuel and air mixing region 94. the upstream end of each second tubular wall 96 and the upstream end 1 of each third tubular wall 98 for supplying 00 in the axial direction.

A plurality of second fuel injectors 104 are provided for each of the tubular combustion chambers 36. . Each of the second fuel injectors 104 extends axially with respect to the tubular combustion chamber 36. It has a hollow cylindrical member that extends. Each of the hollow cylindrical members transfers the fuel to the second fuel. and an axial group passing upstream of the third tubular M98 to supply the air mixing region 94. Penetrates in the direction.

Each second and third tubular wall 96,98 is coaxial about the first portion 38 of each tubular wall. be placed in At the downstream end of each second tubular fuel and air mixing region 94, a second second and third tubular walls 96 and 98 are secured to each third frustoconical portion 40; Each frustoconical portion 40 has a plurality of circumferentially spaced openings 106. The opening 106 extends into the tubular combustion chamber 36 in a downstream direction toward the axis of the tubular combustion chamber 36. The second combustion zone 112 of the combustion chamber is adapted to supply fuel and air to the second combustion zone 112 of the combustion chamber. Opening 1 06 is a circle or a slot.

Each first side plate 66 has an upstream portion 38 of the annular wall and a third annular wall for cooling the tubular wall. a plurality of openings 108 for supplying cooling air to the annular space between the wall 98; ing.

The annular wall has spaced perforated inner and outer sheets that provide exhaust cooling of the annular wall. It is formed from a thin layer structure.

In operation, the first air A passes through the first air intake 82 and into the first and second air inlets 82. The radial flow flows through the vortex. This lip 78 directs the first air 1 fuel-air mixing zone, a first combustion zone 64 . 1st and 2nd The airflow from the radial flow is in the opposite direction, which creates opposing vortices B and C. Form. A shear layer is formed between vortices B and C that improves mixing turbulence.

Pilot injector 86 is only used at low power settings below approximately 40% power. be done. They create a fuel-rich mixture locally on the axis of the annular combustion chamber 36. In order for the first air to pass through the second radial flow swirler only at a small angle Inject gas or pre-vaporized liquid fuel. Fuel is moved into vortex B by diffusion. of the first air. Vortex C leaves an area of only air. Therefore, the first combustion It is created on the centerline of the combustion chamber 36 to maintain combustion within the region 64. Low power operation During operation, the first fuel injector 88 is not used and the first air is injected into the first and second vortices. It exits at the downstream end of the passageway formed between each wing 70 and 74 of the winder assembly.

At high power settings greater than approximately 40% power, the pilot injector 86 is not used, and all fuel supplied to the combustion chamber 36 is transferred to the first and second injectors. 88 and 104, respectively.

At high power settings, i.e., 40% or more, the first fuel injector 88 supplies gas or pre-vaporized liquid fuel to the first and second swirlers. Inject into passages 72 and 80 formed between each vane 70 and 74. 2nd at the same time The fuel injector 104 injects the second fuel through the second annular intake 102. and a second fuel-air mixing region for mixing with secondary air entering air mixing region 94. 94 with gas or pre-vaporized liquid fuel.

The first and second radial flow swirler assemblies are arranged so that the fuel-air mixture flows into the combustion chamber. towards the centerline of the annular combustion chamber 36 before rotating so as to flow parallel to the centerline of the annular combustion chamber 36. to flow the fuel-air mixture. The fuel flows into vortex B and vortex C which have opposite spirals. The shear layer between the two vortices improves the mixed flow. Inside the annular combustion chamber 36 is free of swirl and therefore the gas minimizes the volume of the annular combustion chamber 36 and a first combustion zone 6 that minimizes mixing with the cooling air on the inner surface of the upstream part of the combustion chamber 36; 4 to quickly return to the centerline of the annular combustion chamber 36. This is above the combustion chamber minimizes heat transfer to the flow section 38 and allows efficient use of cooling air; Improve combustion efficiency.

The second air E passes through the second air intake 102 to the second air-fuel mixing region 94. flows into. The second air and fuel are passed axially through a second fuel-air mixing region 94. They are mixed as they flow downstream. As a result, the second air-fuel mixing region 94 is injected into the second downstream portion 42 of the shaped combustion chamber 36 through the opening 106 to produce a second combustion occurs in the second combustion zone 112. Injected from the second fuel-air mixing region 94 The fuel-air mixture separates in a downstream direction toward the centerline of the annular combustion chamber 36. It is in the form of a jet. This means that the second fuel-air mixture leaves the first combustion zone 64. of gases, thus ensuring good mixing.

Cooling flowing over the second fuel-air mixing jet F and the inner surface of the downstream portion 42 of the combustion chamber 36 Interaction with air is minimized by alignment of the jet, F, toward the centerline of the combustion chamber. be converted into

The graph of FIG. 6 shows the pilot, first and second injectors 86, 88 and 104. How the fuel flow to the gas turbine engine depends on the power or load setting of the gas turbine engine. It is a graph showing whether the amount changes.

Only the pilot injector 86 is supplied with a power setting of 35% or less.

At power settings of 35% or higher, fuel flows through the first and second injectors 88 and 10. 4 at the same time, and the supply of fuel to the pilot injector 86 ends.

The power or load setting for 35% and 83% of the fuel supplied to each combustion chamber is determined by the first engine. The remaining 17% of the fuel is supplied to the second injector 104. It will be done. As the power, or load, setting increases, the total amount of fuel delivered to each combustor The total amount of fuel that increases in quantity and is supplied to the first injector and the second injector The amount increases. All fuel supplied to the combustion chamber supplied to the first injector 88 Body percentage gradually decreases from 83% at 35% power setting to 100% power setting. At the force setting it decreases to almost 50%. supplied to the second injector 104 Percentage of total fuel delivered to the combustion chamber starts at 17% at 35% power setting It gradually increases to almost 50% at 100% power setting.

The percentage of fuel supplied to the first injector 88 is 40% It gradually decreases from 78% at a constant power setting, and decreases to almost 50% at a 100% power setting. Ru. On the other hand, the percentage of fuel supplied to the second injector 104 is 4 It gradually increases from 22% at 0% power setting, and almost 5 at 100% power setting. Increase to 0%.

Prevents nitrogen from becoming associated with fuel at high temperatures, reducing the formation of nitrogen oxides (NOx) to have a constant maximum temperature of 1800°K (1527°C) to prevent A first fuel-air mixing region 64 is supplied with fuel.

Also - oxidized carbon having a constant maximum temperature of 1800°K (1527°C) It has a minimum temperature of 1500'K (1227°C) to prevent the formation of The fourth fuel-air mixing region 64 is supplied with fuel so that the fourth fuel-air mixing region 64 is supplied with fuel. Preferably, the most The low temperature is at 1550'. Heat released to first fuel and air mixing region 64 heats the second air in the second fuel-air mixing region 94 .

In the combustion chamber 36 shown in FIGS. 2-5, the flame in the first fuel-air mixing region 64 is Temperatures should be kept at approximately 100 ms, i.e. in places, so that nitrogen oxide emissions are kept low. It is necessary to stay within a range of temperatures. However, the movement of 35% and 100% By changing the power setting between forces, the boundary between the required flame temperature and the temperature at which the flame goes out can be changed. The world changes. In some circumstances, the flame is within the first fuel and air mixing region. disappear. of the fuel to provide a suitable boundary between the flame temperature and the flame extinction temperature. The majority is supplied to the first fuel and air mixing region 64 . However, this solution This method is undesirable because it increases the flame temperature and increases nitrogen oxide emissions.

The other combustion chamber assembly 136 shown in FIG. 7 is substantially similar to that shown in FIGS. 2-5. and the same reference numbers are used to designate the same parts. combustion chamber assembly The downstream end of the first portion 38 of the annular wall 37 is directed toward the throat 122. 2-5 in that it has a truncated conical portion 120 of decreasing diameter. different from that. The third truncated conical portion 40 separates the first portion 38 and the second portion 42. Connected to each other, the second portion 42 has a larger diameter than the first portion 38.

the downstream end of the first portion 38 provided by the frustoconical portion 120 and the throat 112; the first to reignite the fuel-air mixture by reducing the diameter at the directs recirculation of hot combustion products to the fuel and air mixing zone or primary combustion zone 64. make it work. This also means that the second air is transferred from the second fuel-air mixing region 94 to the first fuel. Minimize or prevent air from flowing into the mixing region 64, i.e., the primary combustion region. Stop. By reducing the diameter at the downstream end of the first portion 38, the first In combination with the constant temperature of the fuel air mixing region or combustion region 64, the temperature of the flame and A suitable boundary between the flame extinguishing temperature and the flame extinguishing temperature in the first fuel-air mixing region 64. To prevent this, the power setting will change between 35% and 100% power. maintained. Pilot, first and second injectors 86, 88 and 104 The fuel flow is controlled at the power setting of the gas turbine engine in the same manner as shown in Figure 6. Each changes.

The combustion chambers shown in Figures 2 to 5 and 7 must be -30' to +30'' Suitable for operation over the full power range at ambient temperatures in the range of .

The other combustion chamber assembly 236 shown in FIG. 8 is generally similar to that shown in FIG. The same reference numbers are used to designate the same parts. Combustion chamber assembly 2 36 is such that each of the tubular combustion chambers 236 is connected to the second portion 42 by the fifth portion 132. having a fourth portion 130 interconnected with and located downstream thereof; This is different from that shown in FIG. The fourth portion 130 of the tubular wall is larger than the second portion 40. The fifth portion 132 is frustoconical in shape. Second part of tubular wall 37 The downstream end of 42 has a frustoconical portion 134 whose diameter extends toward the throat 136. The third tubular fuel-air mixing region 138 is connected to the second combustion region 1 of each tubular combustion chamber 236. Surround 12. Each third tubular combustion air mixing region 138 has a fourth tubular wall 140 and the fifth tubular wall 142. The fourth tubular wall 140 is a third fuel chamber. A fifth tubular wall 142 defines a radially outer end of the gas mixing region 138, and a fifth tubular wall 142 defines a third fuel air space. A radially inner end of a gas mixing region 138 is defined. Third tubular air intake 144 a fourth and fifth tube for supplying air to the third tubular fuel-air mixing region 138; is formed between the upstream ends of the shaped wall 140.

A plurality of tertiary fuel injectors 146 are provided for each of the tubular combustion chambers 236 and each Fourth and fifth tubular walls 140, 142 are coaxial about the second portion 42 of each tubular wall. It is located in a central location. At the downstream end of each third fuel-air mixing region 138, a fourth and fifth tubular walls 140 and 142 are secured to the fifth frustoconical portion 132, with each circular The frustum portion 132 is provided with a plurality of radially spaced openings 148. This opening extends downstream toward the axis of the tubular combustion chamber 236. The fuel air is directed to the tertiary combustion zone 150 at 236 . This opening 1 48 is circular or slotted.

In operation, the first air A flows through the first air intake 82 and the first and second Radial flow flows through the volute. The lip 78 converts the first air into the first fuel. to the air mixing zone, ie, the first combustion zone 64. first and second radial The airflow from the flow swirler is in opposite directions to improve mixing turbulence.

This pilot injector 86 is used at lower power settings of less than about 40%. Ru. They create a fuel-rich mixture locally on the axis of the tubular combustion chamber 236. a narrow angle to the first airflow passing through the second radial flow swirler to or pre-vaporized liquid fuel. The fuel is mixed with vortex B by diffusion. let Vortex C is in an air-only region. This locally abundant mixture is In this case, the pilot injector 86 was not used and the fuel was supplied to the combustion chamber 236. All fuel is supplied from primary and secondary or primary, secondary and tertiary injectors. Improved recirculation of combustion products to the first combustion zone 64.

Downstream of second portion 40 provided by frustoconical portion 134 and throat 136 By reducing the diameter at the end, the temperature of the flame in the second combustion zone 112 and the flame are extinguished. By changing the power setting, a suitable boundary between the temperature and the Prevent the 12 flames from going out. This increases the recirculation area J. hot combustion products to the second combustion zone 64 to reignite the fuel-air mixture. Improve recirculation.

The combustion chamber 236 is operated at low ambient temperatures in the range of 160°C to 130'C. 40% and 104, respectively. Fueling the first and second combustion zones 64 and 112 for power settings between 100%. supply The third combustion injector 146 operates at low ambient temperatures at any power setting. The third combustion zone 150 is not supplied with fuel at any time. At low atmospheric temperatures, the first The amount of fuel supplied to the injector 88 increases the temperature of the first combustion zone 64 to 18 Increase to maintain at 00°K. This is the optimal fuel for reducing nitrogen oxides. a second combustion zone to continue the combustion in a second combustion zone 112 to ensure ignition; It is important to ensure that the temperature of 112 is maintained sufficiently high.

If the combustion chamber 236 is operated at high atmospheric temperatures in the range of +30°C or above, the first and second fuel injectors 88 and 104 at 40% and predetermined power, respectively. Fueling the first and second combustion zones 64 and 112 for lower power settings between settings. supply High ambient temperature and high power settings between a given power setting and 100% power In the configuration, the first, second and third fuel injectors 88, 104 and 146 are Supplying fuel to the first, second and third combustion zones 64, 112 and 150, respectively. Ru.

As the ambient air temperature decreases from the higher ambient air temperature, the first, second and third fuel Injectors 88, 104 and 146 direct fuel into the first, second and third combustion zones 6. 4. The lowest power setting supplied to the 112 and 150 is recommended for high ambient temperature operation. from a predetermined power setting. At low atmospheric temperatures as mentioned above, the third Fuel injector 146 is not supplied with fuel at any power setting.

At high power and high ambient temperatures, the temperature of the first fuel-air mixing region 64 is approximately 1 800°K and the temperature within the second combustion zone 112 is maintained at approximately 1740°K. and the temperature of the third combustion zone 150 varies between 1550°K and 1800'K. do. When the temperature within the third fuel region 150 drops below 1550°, the third The combustion injector 146 'does not supply fuel to the third combustion zone 150 and of fuel to the second combustion zone 112 supplied by the fuel injector 104 of the The quantity increases until its temperature rises to 1850'K. The device is a two-stage combustion Acts as a vessel.

a second fuel-air mixing region 94 and a second combustion region 112 and a third fuel-air mixing region The combination of the combustion chamber 150 and the third combustion zone 150 allows for various shapes for the combustion chamber 236. 4 over a wide range of pressure ratios and velocity profiles without the need for air intakes Reduced nitrogen oxide emissions achieved at all power settings between 0% and 100% can do.

Industrial gas turbine engines require combustion chambers shown in FIGS. 2-5 and 7. The power supplied to the pilot, first and second injectors is It is equipped with a control device for controlling.

Industrial gas turbine engines have the required power and atmospheric temperature for the combustion chamber shown in Figure 8. Controls the fuel supplied to the pilot, first and second injectors by It is equipped with a control device to

% of base load fuel flow Submission of translation of written amendment (Article 184-8 of the Patent Law) April Δ, 1993

Claims (1)

[Claims] 1. A first air intake device (82), a first fuel injector device (88) and and a first fuel-air mixing region (64), the first fuel-air mixing region (64) at least one tubular wall (37) and an upstream connected to the upstream end of the tubular wall (37) formed by a wall (44), the upstream wall (44) having at least one opening (62 ), the first air intake device having at least one first radial flow a volute (66, 68, 70) and at least one second radial flow vortex a respective first radial flow swirler (66, 6); 8,70) is emptied into the first fuel-air mixing region (64) through said opening (62). each second radial flow swirler (68, 74 , 76) directs air through the opening (62) into the first fuel-air mixing region (64). each first radial flow swirler (66, 68, 70) includes a respective second radial flow swirler (6) with respect to the axis of the combustion chamber (36). 8, 74, 76) and each first radial flow volute A swirler (66, 68, 70) is connected to each second radial flow swirler (68, 70, 7). 6) Gas turbine engine combustion in which air is swirled in the opposite direction to In the combustion chamber (36), a first fuel injector device (88) is arranged in a first radial direction. Direct flow formed between the vanes of each (70) of the swirlers (66, 68, 70) at least one passageway (72) and a second radial flow swirler (66,6 at least one passageway (72) formed between each vane (74) of ), the gas turbine engine is adapted to supply fuel to the Grilling room. 2. at least one pilot fuel injector (86), each pilot fuel injector (86); a fuel injector (86) for supplying fuel through the opening (62); Combustion chamber according to claim 1, in which the combustion chamber is aligned with each one of the openings (62). 3. The first fuel injector device (88) includes a first radial flow swirler (88). Supply fuel to all passages (72) formed between the blades (66, 68, 70) The combustion chamber according to claim 1, wherein the combustion chamber is adapted to. 4. The first fuel injector device (88) is connected to the second radial flow swirler (88). All the passages (80) formed between the blades (74) of the blades (68, 74, 76) 2. A fuel chamber as claimed in claim 1, adapted to supply fuel. 5. The first fuel injector (88) has a passage (72, 8) between the vanes (70, 74). 2. The fuel according to claim 1, adapted to supply fuel to the radially outer region of the Grilling room. 6. The first fuel injector device (88) is arranged axially with respect to the combustion chamber (36). a hollow cylindrical member extending into the fuel passageway (72, 80); multiple apertures spaced axially along the cylindrical member for spraying Combustion chamber according to claim 1, comprising a section (90). 7. 5. The opening (90) is adapted to direct fuel radially inwardly. 6. The combustion chamber according to 6. 8. The first fuel injection device (88) injects gas fuel or vaporized liquid fuel. 2. The combustion chamber according to claim 1, wherein the combustion chamber is adapted to emit radiation. 9. Pyrobutte fuel injector (86) injects gaseous or liquid fuel The combustion chamber according to claim 2, wherein the combustion chamber is configured as follows. 10. The combustion chamber (36) is tubular and has one opening in its upstream wall (44). The combustion chamber according to claim 1, having (62). 11. a second air intake device (102), a second fuel injector device (10 4) and a second fuel-air mixing region (94); 4) is tubular and surrounds the first fuel and air mixing region (64) and includes a second fuel and air mixing region (64); The air mixing region (94) is defined at the radially outer end by a second tubular wall (96). The second fuel injector device (104) has a second fuel-air mixing region. The fuel is supplied to the upstream end of the region (94), and the second fuel-air mixing region The zone (94) is connected at its downstream end to the combustion chamber ( 3. A second combustion zone (112) as claimed in claim 2, in fluid flow communication with a second combustion zone (112) inside the combustion chamber (36). combustion chamber. 12. The tubular wall (37) is a first portion forming a first fuel-air mixing region (64). (38) and downstream of the first portion (38) forming the second combustion zone (112). a second portion (42) having a larger diameter and first and second portions (38, 42); a third truncated conical section (40) interconnecting the combustion chamber according to claim 11; Room. 13. The downstream end (120) of the first portion (38) of the tubular wall (37) is diametrically Combustion chamber according to claim 12, characterized in that it is reduced to a throat (122). 14. The second air intake (102) is below the first air intake device (82). 12. The combustion chamber of claim 11 in the flow. 15. The second fuel-air mixing region (104) is separated by a third tubular wall (98). The combustion chamber according to claim 11, wherein the combustion chamber is formed at a radially inner end of the combustion chamber. 16. The third conical portion (40) is arranged in a downstream direction toward the centerline of the combustion chamber (36). from the second fuel-air mixture as a plurality of jets to the second fuel-air mixing region (1 04) equally spaced circumferentially to deliver a second fuel-air mixture to the Combustion chamber according to claim 12, having a plurality of openings (106). 17. Combustion chamber according to claim 16, wherein the opening (106) is a slot. 18. The downstream end of the second tubular wall (96) is connected to the third conical portion (37) of the tubular wall (37). 13. The combustion chamber according to claim 12, wherein the combustion chamber is fixed to 40). 19. Cooling air was formed between the tubular wall (37) and the third tubular wall (98) Combustion chamber according to claim 15, which is fed into a tubular chamber (110). 20. A plurality of second fuel injector devices (104) are arranged at equal intervals in the circumferential direction. 12. The combustion chamber of claim 11, comprising a plurality of injectors. 21. The second combustion injector device (104) supplies gaseous fuel or vaporized liquid fuel. Combustion chamber according to claim 11, adapted for injection. 22. a third air intake device (144); a third fuel injector device (14); 6) and a third fuel-air mixing region (138); 138) is tubular and surrounds the second combustion zone (112) and provides a third fuel-air mixture. A mating region (138) is defined at the radially outer end by a fourth tubular wall (140). The third fuel injector device (146) is a tertiary fuel-air mixing region. (138) to supply fuel to the upstream end of the third fuel-air mixing region. The zone (138) at its downstream end connects to the combustion chamber (236) downstream of the second combustion zone (112). ) in fluid flow communication with a third combustion zone (150) inside the combustion chamber. 23. The tubular wall (37) is further downstream of the second portion (42) than the second portion (42). a fourth portion (42) of larger diameter and forming a third combustion zone (150); , a fifth truncated cone (132) interconnecting the second and fourth portions (42, 130). ) portion. 24. The downstream end of the second portion (42) of the tubular wall has a throat (136) in diameter. 23. The combustion chamber of claim 22, wherein the combustion chamber is reduced to . 25. The third air intake (144) is connected to the second air intake device (102). 23. The combustion chamber of claim 22, which is downstream. 26. A third fuel-air mixing region (138) is defined by a fifth tubular wall (142). 23. A combustion chamber according to claim 22, formed at a radially inner end thereof. 27. The fifth conical portion (132) is downstream toward the centerline of the combustion chamber (236). the third fuel-air mixture as a plurality of jets in the direction of the third fuel-air mixing region; (104) A plurality of circumferentially equally spaced openings (104) 24. The combustion chamber according to claim 23, comprising: 48). 28. Combustion chamber according to claim 27, wherein the opening (148) is a slot. 29. The downstream end of the fourth tubular wall (140) is connected to a fifth conical portion of the tubular wall (37). Combustion chamber according to claim 23, wherein the combustion chamber is fixed at a minute (132). 30. A plurality of third fuel injector devices (146) are arranged at equal intervals in the circumferential direction. 23. The combustion chamber of claim 22, comprising a plurality of injectors. 31. A third combustion injector device (146) delivers gaseous fuel or vaporized liquid fuel. 23. Combustion chamber according to claim 22, adapted for injection. 32. From the start of operation of the gas turbine engine until the specified output power level is obtained Fuel from the pyrobutte fuel injector (86) is transferred to the first fuel air mixing region (6). 4) and a first step for an output power level greater than a predetermined level. A first radial flow swirler (66) for flow to the fuel-air mixing region (64). , 68, 70) at least one passageway (72) formed between the vanes (70) and between the vanes (74) of the second radial flow swirler (68, 74, 76). A first fuel injector device (88) is provided in at least one passageway (80) formed therein. ) and a combustion chamber downstream of the first fuel-air mixing region (64). a second fuel injector device (104) for flowing inside the second fuel injector device (104); 12. The gas according to claim 11, comprising the step of simultaneously supplying fuel to the mixing region (94). How to operate a turbine engine combustion chamber. 33. 33. The predetermined output power level is between 35% and 40% power. the method of. 34. The proportion of fuel supplied from the first fuel injector device is 40% to 10 From 0% output power level to 75% to 50% of the total fuel supplied to the combustion chamber. 34. The method of claim 33, wherein: 35. From the start of operation of the gas turbine engine until the specified output power level is obtained Fuel from the pilot fuel injector (86) is transferred to the first fuel air mixing region (6). 4) and a first step for an output power level greater than a predetermined level. A first radial flow swirler (66) for flow to the fuel-air mixing region (64). , 68, 70) at least one passageway (72) formed between the vanes (70) and between the vanes (74) of the second radial flow swirler (68, 74, 76). A first fuel injector device (88) is provided in at least one passageway (80) formed therein. ) and a combustion chamber downstream of the first fuel-air mixing region (64). a second fuel-air mixing region (94) for flow to a second combustion region inside the simultaneously supplying an output power level greater than a second predetermined level and a predetermined temperature. The combustion chamber (23) downstream of the second combustion zone (112) at an atmospheric temperature greater than a third fuel-air mixing region (150) for flow to a third combustion region (150) inside the 138) for operating a gas turbine engine combustion chamber having a step of supplying fuel to How to do it. 36. 36. The predetermined output power level is between 35% and 40% power. the method of. 37. a first combustion zone (64); and interconnected to the first combustion zone (64); a second combustion zone (112) located downstream of the second combustion zone (112); a third combustion zone (150) interconnected and located downstream thereof; and a first air a first air intake device (82) for supplying air to the first combustion zone (64); and a pilot fuel injector for supplying fuel to the first combustion zone (64). a first fuel device (86) for supplying fuel to the first combustion zone (64); an injector device (88) for supplying second air to the second combustion zone (112); a second air intake device (102) for directing fuel to a second combustion zone (112); ) and a third fuel injector device (104) for supplying air to the a third air intake device (144) supplying the third combustion zone (150); a third fuel injector arrangement (150) for supplying fuel to the third combustion zone (150); 146). A gas turbine engine combustion chamber having: 38. The second combustion zone (112) is connected to the first combustion zone by the throat (122). 38. Combustion chamber according to claim 37, wherein the combustion chamber is interconnected with a region (64). 39. The third combustion zone (150) is connected to the second combustion zone (150) by the second throat (136). Combustion chamber according to claim 37, interconnected with a combustion zone (64).