Power plants – Combustion products used as motive fluid – Combined with regulation of power output feature
Reexamination Certificate
2001-05-10
2002-05-07
Freay, Charles G. (Department: 3746)
Power plants
Combustion products used as motive fluid
Combined with regulation of power output feature
C060S734000, C060S039780, C060S039810, C137S625170
Reexamination Certificate
active
06381947
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to a method for fuel injection into a staged or steped gas turbine combustion chamber with separate fuel injection nozzles for each stage, whereby at least one stage is able to be switched off for specific operating conditions by interrupting the fuel supply. Furthermore, the invention relates to a fuel injection mechanism for execution of the fuel injection method according to the invention.
BACKGROUND OF THE INVENTION
For the state of the art, reference may be had to WO 95/17632 as an example.
Gas turbine combustion chambers, in particular annular combustion chambers of gas turbines, which operate with staged combustion/staged fuel injection, are increasingly gaining importance for the purpose of reducing the oxides of nitrogen. Typically, a pilot combustion chamber as well as a main combustion chamber is provided which each form constituting a so-called stage or step. Of course, further gradations/stages may also be provided in addition to these two stages. The pilot combustion chamber has as a first stage one or more pilot burners which, in the preferred case of application, comprises an annular combustion chamber and includes fuel injection nozzles in an annular arrangement; likewise, the second stage, namely the main combustion chamber, has several main burners also in the form of several injection nozzles preferably also in an annular arrangement, but optimized for reducing the oxides of nitrogen.
The attached
FIG. 2
shows a basic illustration for such a staged gas turbine combustion chamber. In this case, the combustion chamber outer wall is marked with reference number
20
and the combustion chamber inner wall with reference number
21
. In addition, these two walls
20
,
21
are surrounded by enveloping walls
20
a
,
21
a
which also define on the left side the combustion chamber entrance
22
a
and on the right side the combustion chamber exit
22
b
. Typically, several sets of pilot and main combustion chambers such as are shown in
FIG. 2
are arranged symmetrically about the center line or axis
23
in a gas turbine engine.
A separating wall structure
24
is provided within the left half of each combustion chamber. The so-called pilot combustion chamber
25
a
is situated between this separating wall structure
24
and the center line
23
, while the so-called main combustion chamber
25
b
is below this separating wall structure
24
, that is, radially outwardly of the pilot chamber
25
a
. Assigned to the pilot combustion chamber
25
a
are pilot nozzles
26
a
, while main nozzles
26
b
are provided for the main combustion chamber
25
b
. Fuel and/or an air-fuel mixture is introduced via these nozzles
26
a
,
26
b
into the combustion chambers, while a main air current
27
makes its way via the combustion chamber entrance
22
a
into the individual combustion chambers
25
a
,
25
b
. Furthermore, admixed air
28
can enter via openings in the outer wall
20
, in the inner wall
21
, as well as in the separating wall structure
24
into the individual combustion chambers
25
a
,
25
b
. The air-fuel mixture burned in the pilot burner combustion chamber
25
a
and/or in the main combustion chamber
25
b
as well as in the junction of these two combustion chambers is finally carried off via the combustion chamber exit
22
b.
Only the pilot nozzles
26
a
are operated in lower stress points (low load operations) of the gas turbine, that is to say, the injection nozzles of the main burner
26
b
are not supplied with fuel. In higher load points of the gas turbine, the main burners
26
b
are operated in addition to the pilot burners
26
a
, in such a way that their injection nozzles are then supplied with fuel. Typically, the pilot combustion chamber
25
a
, which is also operated singly for starting the gas turbine and for raising the engine speed up to idle, is operated throughout the entire operating performance range of the gas turbine, particularly in an airplane gas turbine, in order to create an ignition source for the main burners
26
b
which are only switched on when necessary. The purpose of the staged combustion lies in the minimizing of harmful substance emissions, in particular NO
x
. This is achieved in that the respective burner sizes can be better adapted to the given power requirement. Thus, to reduce NO
x
the combustion chamber temperature should be as low as possible, which can be achieved by targeted air supplying (admixed air
28
) into the combustion chamber zone. In this connection, the respective stages, namely the pilot burner
26
a
/the main burners
26
b
are designed for special air-fuel ratios. In the case of low load points of the gas turbine, in which altogether only relatively little fuel is burned, the air-fuel ratio reaching the main burners
26
b
would be too great to be able to support a reasonable combustion at all. For this reason, the main burners
26
b
are switched on only in higher load conditions of the gas turbine.
FIG. 3
shows graphically the strategy according to which the individual burners, namely, the pilot burners
26
a
as well as the main burners
26
b
, are supplied with fuel in this connection. The total fuel flow for the two burners is plotted on the abscissa of this diagram, and the percentage of the pilot burners
26
a
and/or of the main burners
26
b
in this total fuel flow is plotted on the ordinate. The corresponding characteristic curve of the pilot burner
26
a
is marked with the letter A and that of the main burners
26
b
with the letter B. One recognizes that with only a slight total fuel flow at first, that is, in the left section of this diagram, only the pilot burners
26
a
are operated, in such a way that their share of the total fuel flow is 100%. As total fuel flow increases, the main burners
26
b
are then switched on, namely at the switch-on point Z. In so doing, however, there should not be a sudden power increase. Rather, a smooth power increase is desired, in such a way that with a relatively slight supply to the main burners
26
b
, the pilot burners
26
a
are supplied at the same time with a smaller fuel quantity. This switch-on point Z is therefore extremely critical with regard to its setting because there must always be a suitable air-fuel ratio in the pilot burners
26
a
as well as in the main burners
26
b
. In this regard, the same considerations also apply with respect to a reduction in or withdrawal of power of the gas turbine, that is, if the main burners
26
b
after being operated at first are switched off again. To avoid instabilities in the immediate surroundings of this switch-on point Z, a control that contains a hysteresis is proposed for this in WO 95/17632 mentioned above. As thrust increases, the main burners are switched on only at a higher total fuel throughput than when they are switched off as thrust decreases.
But since it is desirable to always have a defined fuel throughput in a defined load point or thrust status of the gas turbine, i.e., regardless of whether it is a matter of a thrust increase or a thrust reduction, the invention addresses the technical problem of providing another solution for the above-described problems in connection with the operation of a second stage with a first stage.
SUMMARY OF THE INVENTION
This technical problem is solved in that at least the stage which is able to be switched off can be operated with pulsed fuel injection. Appropriate fuel injection mechanisms for execution of this fuel injection method according to the invention are described in claims
5
and
6
, while the further subclaims contain advantageous designs and further improvements.
The objectives of the invention are twofold: firstly, to pulse the fuel flow hence the combustion in the main combustion chamber and, secondly, to extend the operation region of the main burner stage further into the lean operating region. Pulsing the fuel flow is desirable since it is well known that pulsed combustion results in lower emissions of oxides of nitrogen.
According to the invention, at least the stage which is
Freay Charles G.
Gartenberg Ehud
Manelli Denison & Selter PLLC
Rolls-Royce Deutschland Ltd & Co KG
Taltavull W. Warren
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