U.S. patent application number 13/974523 was filed with the patent office on 2014-02-27 for method for mixing a dilution air in a sequential combustion system of a gas turbine.
This patent application is currently assigned to ALSTOM Technology Ltd. The applicant listed for this patent is ALSTOM Technology Ltd. Invention is credited to Mirko Ruben Bothien.
Application Number | 20140053569 13/974523 |
Document ID | / |
Family ID | 46967967 |
Filed Date | 2014-02-27 |
United States Patent
Application |
20140053569 |
Kind Code |
A1 |
Bothien; Mirko Ruben |
February 27, 2014 |
METHOD FOR MIXING A DILUTION AIR IN A SEQUENTIAL COMBUSTION SYSTEM
OF A GAS TURBINE
Abstract
The invention concerns a method for mixing a dilution air with a
hot main flow in sequential combustion system of a gas turbine,
wherein the gas turbine essentially comprises at least one
compressor, a first combustor which is connected downstream to the
compressor, and the hot gases of the first combustor are admitted
to at least one intermediate turbine or directly or indirectly to
at least one second combustor. The hot gases of the second
combustor are admitted to a further turbine or directly or
indirectly to an energy recovery, wherein at least one combustor
runs under a caloric combustion path having a can-architecture. At
least one dilution air injection is introduced into the first
combustor, and wherein the direction of the dilution air injection
is directed against or in the direction of the original swirl flow
inside of the first combustor.
Inventors: |
Bothien; Mirko Ruben;
(Zurich, CH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ALSTOM Technology Ltd |
Baden |
|
CH |
|
|
Assignee: |
ALSTOM Technology Ltd
Baden
CH
|
Family ID: |
46967967 |
Appl. No.: |
13/974523 |
Filed: |
August 23, 2013 |
Current U.S.
Class: |
60/776 ;
60/39.23 |
Current CPC
Class: |
F23R 3/12 20130101; F23R
3/28 20130101; F23R 3/06 20130101; Y02E 20/14 20130101; F23R 3/26
20130101; F23R 2900/03341 20130101; F23R 2900/00013 20130101 |
Class at
Publication: |
60/776 ;
60/39.23 |
International
Class: |
F23R 3/28 20060101
F23R003/28 |
Foreign Application Data
Date |
Code |
Application Number |
Aug 24, 2012 |
EP |
12181748.0 |
Claims
1. A method for mixing a dilution air with a hot main flow in a
sequential combustion system of a gas turbine, wherein the gas
turbine essentially comprises at least one compressor, a first
combustor which is connected downstream to the compressor, and the
hot gases of the first combustor are admitted to at least one
intermediate turbine or directly or indirectly to at least one
second combustor, wherein the hot gases of the second combustor are
admitted to a further turbine or directly or indirectly to an
energy recovery, wherein at least one combustor runs under a
caloric combustion path having a can architecture, and wherein at
least one dilution air injection is introduced into the first
combustor, and wherein the resulting swirl flow through the
dilution air injection is directed against or in the direction of
the original swirl flow inside of the first combustor.
2. The method as claimed in claim 1, wherein the first and second
combustor run under a caloric combustion path having a
can-architecture.
3. The method as claimed in claim 1, wherein the first combustor
runs under a caloric combustion path having an annular
architecture, and the second combustor runs under a caloric
combustion path having a can-architecture.
4. The method as claimed in claim 1, wherein the first combustor
runs under a caloric combustion path having a can-architecture, and
the second combustor runs under a caloric combustion path having an
annular architecture.
5. The method as claimed in claim 1, wherein that at least one
combustor runs under a caloric combustion path having an annular
architecture.
6. A dilution air injector for implementing a method for mixing a
dilution air with a hot main flow in a sequential combustion system
of a gas turbine, wherein the gas turbine essentially comprises at
least one compressor, a first combustor which is connected
downstream to the compressor, and the hot gases of the first
combustor are admitted to at least one intermediate turbine or
directly or indirectly to at least one second combustor, wherein
the hot gases of the second combustor are admitted to a further
turbine or directly or indirectly to an energy recovery, wherein at
least one combustor runs under a caloric combustion path having a
can architecture, and wherein the first combustor comprising
tangential air inlet slots forming a swirl flow directed against or
in direction of the original main swirl flow inside of the first
combustor.
7. The dilution air injector as claimed in claim 6, wherein the
first combustor runs under a caloric combustion path having an
annular architecture, and the second combustor runs under a caloric
combustion path having a can-architecture.
8. The dilution air injector as claimed in claim 6, wherein the
first combustor runs under a caloric combustion path having a
can-architecture, and the second combustor runs under a caloric
combustion path having an annular architecture.
9. The dilution air injector as claimed in claim 6, wherein that at
least one combustor runs under a caloric combustion path having an
annular architecture.
10. The dilution air injector as claimed claim 6 wherein the first
combustor comprising at least one injector, wherein the direction
and/or intensity of the injected air along the first combustion
chamber are subject to regulation.
11. The dilution air injector as claimed in claim 6 wherein that
the injector comprising means for regulating the intensity of the
selected dilution air injection or for an additional supporting
dilution air.
12. A combustor as claimed in claim 6 wherein at least one
combustor comprising a burner consisting of hollow part-cone bodies
making up a complete body, having tangential air inlet slots and
feed channels for gaseous and liquid fuels, wherein in that the
centre axes of the hollow part-cone bodies have a cone angle
increasing in the direction of flow and run in the longitudinal
direction at a mutual offset, wherein a fuel nozzle, which fuel
injection is located in the middle of the connecting line of the
mutually offset centre axes of the part-cone bodies, is placed at
the burner head in the conical interior formed by the part-cone
bodies.
13. A combustor as claimed in one of claims 6 to 11, characterized
in that at least one combustor comprising a burner for a combustion
air flow and means for injection of fuel, substantially consisting
of a swirl generator, which substantially consisting of hollow
part-cone bodies making up a complete body, having tangential air
inlet slots and feed channels for gaseous and liquid fuels, wherein
in that the centre axes of the hollow part-cone bodies have a cone
angle increasing in the direction of flow and run in the
longitudinal direction at a mutual offset, wherein a fuel nozzle,
which fuel injection is located in the middle of the connecting
line of the mutually offset centre axes of the part-cone bodies, is
placed at the burner head in the conical interior formed by the
part-cone bodies, and as well of a mixing path provided downstream
of said swirl generator, wherein said mixing path comprises
transaction ducts extending within a first part of the path in the
flow direction for transfer of a flow formed in said swirl
generator into the cross-section of flow of said mixing path, that
joins downstream of said transition ducts.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application claims priority to European Application
12181748.0 filed Aug. 24, 2012, the contents of which are hereby
incorporated in its entirety.
TECHNICAL FIELD
[0002] The invention refers to a method for mixing dilution air in
a sequential combustion system of a gas turbine. The invention
refers additionally to a dilution air injector for implementing the
aforementioned method. Furthermore, the invention is related to
mixing of dilution air with a hot main flow in a gas turbine or in
a "CPSC" (Constant Pressure Sequential Combustion) for a can as
well as annular combustor design in a reliable and uniform way.
BACKGROUND
[0003] Beforehand, some general considerations which allow a better
understanding of the invention:
[0004] CO emissions of gas turbine engines need reductions for the
sake of saving the environment. Such emissions are known to appear,
when there is not sufficient time in the combustion chamber to
ensure the CO to CO.sub.2 oxidation, and/or this oxidation is
locally quenched due to contact with cold regions in the combustor.
Since firing temperatures are smaller under part load conditions
CO, and the CO to CO.sub.2 oxidation gets slower, thus CO emissions
usually tend to increase under these conditions.
[0005] A reduction of CO emissions in turn might be invested in
lowering the gas turbine load at the parking point of a gas
turbine. This reduces the environmental impact due to reduced
CO.sub.2 emissions and overall cost of electricity due to less fuel
consumption during engine parking. Finally the CO emission
reduction might be invested in a reduction of first costs due to
savings on a CO catalyst. In this case a CO catalyst might be
avoided (or at least reduced). At the same time losses, which
appear due to a catalyst will be removed (or at least reduced), and
thereby the overall efficiency of the power plant increased.
[0006] According to the US 2012/0017601 A1 the basic of this state
of art is a method for operating the gas turbine, which keeps the
air ratio .lamda. of the operating burner of the second combustor
below a maximum air ratio .lamda..sub.max during part load
operation. This method is characterized essentially by three new
elements and also by supplementing measures which can be
implemented individually or in combination.
[0007] The maximum air ratio .lamda..sub.max in this case depends
upon the CO emission limits which are to be observed, upon the
design of the burner and of the combustor, and also upon the
operating conditions, that is to say especially the burner inlet
temperature.
[0008] The first element is a change in the principle of operation
of the row of variable compressor inlet guide vanes, which allows
the second combustor to be put into operation only at higher part
load. Starting from no-load operation, the row of variable
compressor inlet guide vanes is already opened while only the first
combustor is in operation. This allows loading up to a higher
relative load before the second combustor has to be put in
operation. If the row of variable compressor inlet guide vanes is
opened and the hot gas temperature or turbine inlet temperature of
the high-pressure turbine has reached a limit, the second combustor
is supplied with fuel.
[0009] In addition, the row of variable compressor inlet guide
vanes is quickly closed. Closing of the row of variable compressor
inlet guide vanes at constant turbine inlet temperature TIT of the
high-pressure turbine, without countermeasures, would lead to a
significant reduction of the relative power.
[0010] In order to avoid this power reduction, the fuel mass flow,
which is introduced into the second combustor, can be increased.
The minimum load at which the second combustor is put into
operation and the minimum fuel flow into the second combustor are
therefore significantly increased.
[0011] As a result, the minimum hot gas temperature of the second
combustor is also increased, which reduces the air ratio .lamda.
and therefore reduces the CO emissions.
[0012] The second element for reducing the air ratio .lamda. is a
change in the principle of operation by increasing the turbine
exhaust temperature of the high-pressure turbine TAT1 and/or the
turbine exhaust temperature of the low-pressure turbine TAT2 during
part load operation. This increase allows opening of the row of
variable compressor inlet guide vanes to be shifted to a higher
load point.
[0013] Conventionally, the maximum turbine exhaust temperature of
the second turbine is determined for the full load case and the gas
turbine and possibly the downstream waste heat boiler are designed
in accordance with this temperature. This leads to the maximum hot
gas temperature of the second turbine not being limited by the TIT2
(turbine inlet temperature of the second turbine) during part load
operation with the row of variable compressor inlet guide vanes
closed, but by the TAT2 (turbine exhaust temperature of the second
turbine). Since at part load with at least one row of variable
compressor inlet guide vanes closed the mass flow and therefore the
pressure ratio across the turbine is reduced, the ratio of turbine
inlet temperature to turbine exhaust temperature is also
reduced.
[0014] Correspondingly, with constant TAT2 the TIT2 is also reduced
and in most cases lies considerably below the full load value. A
proposed slight increase of the TAT2 beyond the full load limit,
typically within the order of magnitude of 10.degree. C. to
30.degree. C., admittedly leads to an increase of the TIT2, but
this remains below the full load value and can practically be
achieved without service life losses, or without significant
service life losses. Adaptations in the design or in the choice of
material do not become necessary or can be limited typically to the
exhaust gas side. For increasing the TIT2, the hot gas temperature
is increased, which is realized by an increase of the fuel mass
flow and a reduction of the air ratio .lamda., which is associated
therewith. The CO emissions are correspondingly reduced.
[0015] A further possibility for reducing the air ratio .lamda. of
the burner in operation is the deactivating of individual burners
and redistribution of the fuel at constant TIT2.
[0016] In order to keep the TIT2 constant on average, the burner in
operation has to be operated hotter in proportion to the number of
deactivated burners. For this, the fuel feed is increased and
therefore the local air ratio .lamda. is reduced.
[0017] For an operation which is optimized for CO emissions, in a
gas turbine with split line, a burner (for example for the second
combustor) which is adjacent to the split line is typically
deactivated first of all. In this case, the plane in which a casing
is typically split into upper and lower halves is referred to as
the split line. The respective casing halves are connected in the
split line by a flange, for example.
[0018] Its adjacent burners are subsequently then deactivated or a
burner, which is adjacent to the parting plane on the opposite side
of the combustor is deactivated and in alternating sequence the
adjacent burners, which alternate on the two sides of the
combustor, starting from the parting plane, are deactivated.
[0019] A burner which is adjacent to the split line is preferably
deactivated first of all since the split line of a gas turbine is
typically not absolutely leak proof and in most cases a leakage
flow leads to a slight cooling and dilution (see below mentioned
considerations) of the flammable gases and therefore to locally
increased CO emissions. As a result of deactivating the burners
which are adjacent to the parting plane, these local CO emissions
are avoided.
[0020] The combustion instabilities, which are to be avoided by
means of staging, typically no longer occur at low load, or are
negligibly small, or at part load combustion occur. In one
exemplary embodiment, it is proposed, therefore, to carry out the
restricting not by means of a fixed restrictor but by means of at
least one control valve. This at least one control valve is opened
at low load so that all the activated burners can be operated
virtually homogenously with a low air ratio .lamda.. At high load,
the at least one control valve is throttled in order to realize the
staging.
[0021] Referring to the aforementioned aspects for an optimized
operation for CO emissions and in connection with the currently
proceeding, cooling air from the reheat combustor and any remaining
air from the premix combustor, or fresh air from plenum can be
supplied as dilution air to the main hot gas flow.
[0022] Existing solutions to solve this problems consists in an
injection of secondary medium without swirl. Additionally burners
generating swirl in opposite directions to minimize swirl of main
flow.
[0023] Accordingly, the technical problem consists in a rapid and
good mixing of hot gas products with fresh dilution air to obtain
uniform inlet temperatures and flow field upstream of a reheat
burner. Control of swirl of main flow is, additionally,
mandatory.
SUMMARY
[0024] The present invention is based on the object of proposing a
method to improve mixing of dilution air and hot combustion
products of first stage combustor by injecting dilution air with a
swirl.
[0025] In addition, control of the existing main swirl is possible
by either injecting in direction of the main swirl flow, to amplify
existing swirl flow, or against it, to suppress integral or at
various stages or levels the existing swirl flow.
[0026] Additionally, the present invention is based as example on
the concept of constant pressure sequential combustion system. In
this concept, hot combustion products from premix combustor are
cooled down by a dilution air introduction and subsequently enter a
reheat combustor.
[0027] Dilution air is responsible for mixing of premix and reheat
cooling air with hot combustion products from the premix combustor.
Primary requirements from such a dilution air are uniform
temperature distribution at the inlet to reheat burner, as well as
low pressure drop for performance reasons.
[0028] Accordingly, in the above identified gas turbine combustors,
the main flow usually exhibits a swirling flow pattern. In an
annular combustion chamber, this can be due that all burners
generate a swirling flow in the same direction. In can combustors,
usually more than one burner nozzle is used to inject the fuel and
air into the combustion chamber. This also can result in a main
swirl of the mean flow.
[0029] If downstream of this swirling flow air or fuel, and also
dilution air, is injected, the challenge consists to obtain a good
mixture with the hot gases as fast as possible. This is crucial to
achieve uniform temperature and flow profile at the inlet of the
reheat burner.
[0030] The present invention is in this sense related to mixing of
dilution air with a hot main flow in a constant pressure sequential
combustion system for a can as well as annular combustor designs in
a reliable and uniform way.
[0031] In details, the invention describes below a procedure for
mixing a dilution air with hot combustion products inside of first
combustor, additionally, by injecting dilution air with a swirl,
furthermore, control of the existing main swirl flow by either
injecting in direction of the main swirl flow and finally amplify
or suppress at various stages the existing main swirl flow.
[0032] Generic sketches of such gas turbines are shown in FIGS. 1
to 3, especially in FIG. 1.
[0033] Therein a compressor is followed by a combustor section,
which can consist of a number of cans. Within these cans a first
combustor is followed by a second combustor. Between or
intermediate these two combustors dilution air might be injected in
order to control the inlet temperature of the second combustor and
therefore the self-ignition time of the fuel injected therein.
Finally the hot combustion gases are fed into a turbine.
[0034] A can-architecture is also given, when an annular first
and/or second combustion chamber having or comprising to each
burner in flow direction an independent can or a separating flow
combustion area which is wall-isolated from each other of the
adjacent combustion areas or burners.
[0035] The basic idea of current invention is based on two
basically concepts:
1. The gas turbine is equipped with two combustors in series with
an injection of dilution air against direction of main swirl flow.
2. The gas turbine is equipped with two combustors in series with
an injection of dilution air in direction of main swirl flow.
[0036] By injecting the dilution air with a defined swirl the
following objectives can be achieved:
1. Enhance mixing between dilution air and hot gases from first
burner. 2. Suppress integral or in part swirl of main flow by
injecting the dilution air against the main swirl flow direction.
3. Amplify swirl of main flow by injecting the dilution air in
direction of the main swirl flow direction.
[0037] Advantages associated with the present invention are as
follows: [0038] Better and faster mixing of dilution air and hot
gas from first combustor. [0039] Control of swirl of main flow in
connection with inlet profile of the reheat combustor. [0040]
Possible application for all gas turbine concepts including feature
swirling main flows and injection of air downstream of the first
burner.
[0041] To ensure this final purpose it is also necessary that the
geometries and/or flow coefficients of the various components of
the gas turbine are measured and components with high flow rates
and components with low flow rates are combined inside the
combustor cans or annular combustion chamber.
[0042] The gas turbine comprises essentially at least one
compressor a first combustor which is connected downstream to the
compressor. The hot gases of the first combustor are admitted at
least to an intermediate turbine or directly or indirectly to a
second combustor. The hot gases of the second combustor are
admitted to a further turbine or directly or indirectly to an
energy recovery, for example to a steam generator.
[0043] Further advantages associated with the present invention are
as follows: [0044] Reduced total combustor pressure drop, thus
increased thermodynamic efficiency. [0045] Simple design of the
injection of dilution air. [0046] Uniform temperature distribution
at reheat burner inlet, thus a homogenous combustion process can
act on the pulsations in the combustor and can act on an
over-proportional increase of CO production of the reheat burner.
[0047] Reliable operation without local backflow or
overheating.
[0048] Based on these findings the concept can be expected to work
for an engine, which runs under sequential combustion (with or
without a high pressure turbine) in a can-architecture, but not
only.
[0049] Referring to a sequential combustion the combination of
combustors can be disposed as follows: [0050] At least one
combustor is configured as a can-architecture, with at least one
operating turbine. [0051] Both, the first and second combustors are
configured as sequential can-can architecture, with at least one
operating turbine. [0052] The first combustor is configured as an
annular combustion chamber and the second combustor is built-on as
a can configuration, with at least one operating turbine. [0053]
The first combustor is configured as a can-architecture and the
second combustor is configured as an annular combustion chamber,
with at least one operating turbine. [0054] Both, the first and
second combustor are configured as annular combustion chambers,
with at least one operating turbine. [0055] Both, the first and
second combustor are configured as annular combustion chambers,
with an intermediate operating turbine.
[0056] Accordingly, in terms of injection of dilution air with a
swirl flow for a can-architecture the interaction between
individual cans is minimal or inexistent. Therefore for a can
variant the described concept will be even more effective than for
annular engine architecture.
[0057] In addition to the method, a gas turbine for implementing
the method is a subject of the invention. Depending upon the
concept of the injection of dilution air, the design of the gas
turbine has to be adapted and/or the fuel distribution system
and/or the cooling air system have to be adapted in order to ensure
the feasibility depending on the used dilution air for reducing the
locally combustor pressure drop. All the components of a gas
turbine lie within the range of permissible tolerances. These
tolerances lead to slightly different geometries and
characteristics for each component and for the used concept of
injection of dilution air.
[0058] This, especially, also leads to different pressure losses
and flow rates during operation. The tolerances are selected so
that they have practically no influence upon the operating behavior
during normal operation, especially at high part load and full
load. For this, the geometries and/or flow coefficients of the
various injection of dilution air are measured with existing flow
rates in connection with the operating dilution air swirls.
[0059] Additional advantages associated with this invention are as
follows:
[0060] CO emissions are reduced especially at lower part-load
conditions. Therefore, the gas turbine can be parked at lower
values during such a period. [0061] Thereby the power plant
operator can save fuel and therefore reduce the overall cost of
electricity. [0062] Environmental benefit due to reduced CO
emissions, lower parking point (thus less fuel consumption and
CO.sub.2 production) or a combination of both advantages. [0063]
Possibility of eliminating an expensive CO catalyst. Therefore
first costs are reduced.
[0064] When using a setup including dilution air swirl between
subsequent operating combustors further advantages arise: [0065]
Further CO reduction, with all advantages described above, due to
increased volume for CO oxidation with origin in the first
combustor. [0066] Reduction of circumferential temperature
gradients between the different can combustors. Therefore the
turbine inlet profile is improved and lifetime of turbine parts is
improved.
BRIEF DESCRIPTION OF THE DRAWINGS
[0067] The invention is shown schematically in FIGS. 1 to 3 (1, 2,
2a, 2b, 3, 3a, 3b) based on exemplary embodiments.
[0068] In the drawings:
[0069] FIG. 1 shows a gas turbine equipped with two combustors in
series forming a sequential combustion;
[0070] FIG. 1a shows a section of a can combustor with respect to
FIG. 1;
[0071] FIG. 1b shows a section of an annular combustor with respect
to FIG. 1;
[0072] FIG. 2 shows a gas turbine equipped with two combustors in
series and dilution air injection;
[0073] FIG. 2a shows a section of a can combustor with respect to
FIG. 2;
[0074] FIG. 2b shows a section of an annular combustor with respect
to FIG. 2;
[0075] FIG. 3 shows a gas turbine equipped with two combustors in
series and dilution air injection;
[0076] FIG. 3a shows a section of a can combustor with respect to
FIG. 3;
[0077] FIG. 3b shows a section of an annular combustor with respect
to FIG. 3.
DETAILED DESCRIPTION
[0078] FIGS. 1, 2 and 3 show a part of the gas-turbine group 100,
200, 300, namely the part which includes the sequential combustion,
referring to a "CPSC" system (Constant Pressure Sequential
Combustion).
[0079] Compressed air flows out of a compressor system (not shown)
into a premixing burner 101, which can be operated with a fuel. The
initial generation of hot gases takes place in a first combustion
chamber 102 designed as a can combustor (see FIGS. 1a, 2a, 3a) or
as an annular combustion chamber (see FIGS. 1b, 2b, 3b). The
following generation of hot gases then take place in a second
combustion chamber 104 designed as a can combustor (see FIGS. 1a,
2a, 3a) or as an annular combustion chamber (see FIGS. 1b, 2b, 3b).
Typically, the gas turbine system includes a generator (not shown),
which at the cold end of the gas turbine, that is to say at the
compressor, is coupled to a shaft of the gas turbine.
[0080] Accordingly, FIGS. 1, 2 and 3 show gas turbine systems with
sequential combustion for implementing the method according to the
invention. The gas turbine system comprises a compressor (not
shown), a first combustor 102, a second combustor 104 with a reheat
burner and downstream of the second combustor a turbine (106, not
shown).
[0081] FIG. 1 shows a first combustion chamber 102 having a premix
burner 101 as disclosed for example by EP 0321 809 A1 or EP 0 704
657 A1. This publication forms an integral part of this
description. The hot gases 103 generated in the first combustion
chamber 102, designated as a can combustor or as an annular
combustion chamber, stream to a second combustion chamber 104. The
second combustion chamber 104 essentially has the form of a can
(see FIG. 1a) or an annular duct (see FIG. 1b) through which flow
occurs and in which preferably a gaseous fuel (not shown) is
injected. A self-ignition of the injected fuel takes place starting
at a temperature of the exhaust gases coming from the first
combustion chamber 102 of at least 850 DEG C.
[0082] The second combustion chamber 104 has as burner 105, as
discloses for example by EP 0 620 362 A1, a number of fuel lances
roughly at the end of the premixing zone, which fuel lances are
distributed over the periphery and assume the function of injecting
the fuel. The entire configuration of the gas-turbine group,
excluding the generator, is mounted on a single common rotor
shaft.
[0083] The can architecture comprises a plurality of cans arranged
in an annular array about the circumference of the turbine shaft
(see FIG. 1a), which enables an individual combustion operation of
each can and which will be no harmful interactions among individual
cans during the combustion process.
[0084] If premix burners 101 for the can's combustion or annular
concept are provided, these should preferably be formed by the
combustion process and objects according to the documents EP 0 321
809 A1 and/or EP 0 704 657 A2, wherein these documents forming
integral parts of the present description.
[0085] In particular, said premix burners 101 can be operated with
liquid and/or gaseous fuels of all kinds. Thus, it is readily
possible to provide different fuels within the individual cans.
This means also that a premix burner can also be operated
simultaneously with different fuels.
[0086] The second or subsequent can combustor or annular combustor
is preferably carried out by EP 0 620 362 A1 or DE 103 12 971 A1,
wherein these documents forming integral parts of the present
description.
[0087] Additionally, the following mentioned documents forming also
integral parts of the present description: [0088] EP 0 321 809 A1
and B1 relating to a burner consisting of hollow part-cone bodies
making up a complete body, having tangential air inlet slots and
feed channels for gaseous and liquid fuels, wherein in that the
centre axes of the hollow part-cone bodies have a cone angle
increasing in the direction of flow and run in the longitudinal
direction at a mutual offset. A fuel nozzle, which fuel injection
is located in the middle of the connecting line of the mutually
offset centre axes of the part-cone bodies, is placed at the burner
head in the conical interior formed by the part-cone bodies. [0089]
EP 0 704 657 A2 and B1, relating to a burner arrangement for a heat
generator, substantially consisting of a swirl generator,
substantially according to EP 0 321 809 A1 and B, for a combustion
air flow and means for injection of fuel, as well of a mixing path
provided downstream of said swirl generator, wherein said mixing
path comprises transaction ducts extending within a first part of
the path in the flow direction for transfer of a flow formed in
said swirl generator into the cross-section of flow of said mixing
path, that joins downstream of said transition ducts.
[0090] Furthermore, it is proposed fuel injector for use within a
gas turbine reheat combustor, utilising auto-ignition of fuel, in
order to improve the fuel air mixing for a given residence time.
The specific embodiments of this injector are envisaged: [0091] The
oscillating gaseous fuel is injected normal to the flow of oxidant
in sense of a cross-flow configuration. [0092] The oscillating
gaseous fuel is injected parallel to the flow of oxidant in sense
of an in-line configuration. [0093] The oscillating gaseous fuel is
injected at an oblique angle, between 0.degree. and 90.degree. to
the flow of oxidant. [0094] EP 0 646 705 A1 and B1, relating to a
method of establishing part load operation in a gas turbine group
with a sequential combustion. [0095] EP 0 646 704 A1 and B1,
relating to a method for controlling a gas turbine plant equipped
with two combustor chambers. [0096] EP 0 718 470 A2 and B1,
relating to method of operating a gas turbine group equipped with
two combustor chambers, when providing a partial-load
operation.
[0097] Other relevant published documents, which include one or
more improvements of the above identified documents forming also
integral parts of the present description.
[0098] Referring to a sequential combustion the combination of
combustors can be disposed as follows:
[0099] At least one combustor is configured as a can-architecture,
with at least one operating turbine. [0100] Both, the first and
second combustors are configured as sequential can-can
architecture, with at least one operating turbine. [0101] The first
combustor is configured as an annular combustion chamber and the
second combustor is built-on as a can configuration, with at least
one operating turbine. [0102] The first combustor is configured as
a can-architecture and the second combustor is configured as an
annular combustion chamber, with at least one operating turbine.
[0103] Both, the first and second combustor are configured as
annular combustion chambers, with at least one operating turbine.
[0104] Both, the first and second combustor are configured as
annular combustion chambers, with an intermediate operating
turbine.
[0105] In both cases, relating to can combustor 120 or annular
combustion chamber 130, the azimuthal main flow 121, 131 is unitary
in each system.
[0106] FIG. 2 shows a gas turbine, according to FIG. 1, having a
first combustion chamber equipped with at least one dilution air
injection 201 at appropriate place, downstream of the first burner
system 101 and upstream of the second burner system 105, and having
a second combustion chamber 104 downstream of the second burner
system 105. More dilution air injections at different places along
the first combustion chamber 102 are possible. Furthermore, the
direction and the intensity of the single injected air along the
first combustion chamber 102 can be regulated.
[0107] FIG. 2a shows a can combustor 220 having tangential air
inlet slots 222 forming a swirl flow 223 directed against the
predominant direction of the original main swirl flow 221 from the
operation of the first burner 101. The result of this impact
consists in the fact that the existing swirl flow intensity from
the first burner can be reduced or completely suppressed, depending
on the intensity of the selected dilution air injection 222. FIG.
2a shows a reduced resulting main swirl flow 224.
[0108] FIG. 2b shows an annular combustion chamber 230 having
tangential air inlet slots 232 forming a swirl flow 233 directed
against the predominant direction of the original swirl flow 231
from the operation of the first burner 101. The result of this
impact consists in the fact that the existing swirl flow intensity
from the first burner can be reduced or completely suppressed,
depending on the intensity of the selected dilution air injection
232. FIG. 2b shows a reduced resulting main swirl flow 234.
[0109] FIG. 3 shows a gas turbine, according to FIG. 2, having a
first combustion chamber equipped with a dilution air injection 301
at appropriate place, downstream of the first burner system 101 and
upstream of the second burner system 105, and having a second
combustion chamber 104 downstream of the second burner system 105.
More dilution air injections at different places along the first
combustion chamber 102 are possible. Furthermore, the direction and
the intensity of the single injected air along the first combustion
chamber 102 can be regulated.
[0110] FIG. 3a shows a can combustor 320 having tangential air
inlet slots 322 forming a swirl flow 323 in direction of the
original main swirl flow 321 from the operation of the first burner
101. The result of this feeding consists in the fact that the
existing swirl flow intensity from the first burner can be
amplified, depending on the intensity of the selected dilution air
injection 322. FIG. 2a shows an amplified resulting main swirl flow
324.
[0111] FIG. 3b shows an annular combustion chamber 330 having
tangential air inlet slots 332 forming a swirl flow 333 directed
against the predominant direction of the original swirl flow 331
from the operation of the first burner 101. The result of this
impact consists in the fact that the existing swirl flow intensity
from the first burner can be amplified, depending on the intensity
of the selected dilution air injection 332. FIG. 3b shows an
amplified resulting main swirl flow 334.
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