U.S. patent application number 13/974687 was filed with the patent office on 2017-10-19 for method for operating a lean premix burner of an aircraft gas turbine and device for carrying out the method.
This patent application is currently assigned to Rolls-Royce Deutschland Ltd & Co KG. The applicant listed for this patent is Rolls-Royce Deutschland Ltd & Co KG. Invention is credited to Imon Kalyan BAGCHI, Waldemar LAZIK.
Application Number | 20170299183 13/974687 |
Document ID | / |
Family ID | 49084773 |
Filed Date | 2017-10-19 |
United States Patent
Application |
20170299183 |
Kind Code |
A1 |
BAGCHI; Imon Kalyan ; et
al. |
October 19, 2017 |
Method for operating a lean premix burner of an aircraft gas
turbine and device for carrying out the method
Abstract
The present invention relates to a method for operating a lean
premix burner of an aircraft gas turbine, where fuel and primary
supporting air are supplied by means of a supporting burner (pilot
burner) arranged centrically to the burner axis, where secondary
air surrounding the supporting burner is supplied, and where fuel
and air are supplied by means of a main burner, characterized in
that the primary supporting air is supplied in an amount of 5 vol %
to 10 vol % of the total air quantity, that the secondary
supporting air is supplied in an amount of 5 vol % to 20 vol % and
that 35 vol % to 75 vol % of the total air quantity are supplied
via the main burner in the partial load range and in the full load
range.
Inventors: |
BAGCHI; Imon Kalyan;
(Berlin, DE) ; LAZIK; Waldemar; (Teltow,
DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Deutschland Ltd & Co KG |
Blankenfelde-Mahlow |
|
DE |
|
|
Assignee: |
Rolls-Royce Deutschland Ltd &
Co KG
Blankenfelde-Mahlow
DE
|
Family ID: |
49084773 |
Appl. No.: |
13/974687 |
Filed: |
August 23, 2013 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/60 20130101;
F23C 2900/06043 20130101; F23R 2900/03343 20130101; F23N 3/002
20130101; F23R 3/16 20130101; F23C 2201/20 20130101; F23N 2237/16
20200101; F23R 3/14 20130101; Y02T 50/675 20130101; F23N 2241/20
20200101; F23R 3/26 20130101; F23D 2900/11101 20130101; F23N
2237/10 20200101; F23R 3/286 20130101; F23R 3/343 20130101; Y02T
50/677 20130101 |
International
Class: |
F23R 3/28 20060101
F23R003/28 |
Foreign Application Data
Date |
Code |
Application Number |
Aug 28, 2012 |
DE |
10 2012 017 065.3 |
Claims
1. Method for operating a lean premix burner of an aircraft gas
turbine, where fuel and primary supporting air are supplied by
means of a supporting burner (pilot burner) arranged centrically to
the burner axis, where secondary air surrounding the supporting
burner is supplied, and where fuel and air are supplied by means of
a main burner, characterized in that the primary supporting air is
supplied in an amount of 5 vol % to 10 vol % of the total air
quantity, that the secondary supporting air is supplied in an
amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the
total air quantity are supplied via the main burner in the partial
load range and in the full load range.
2. Method in accordance with claim 1, characterized in that
adjacent to the supporting burner a rich zone is formed, that the
rich zone is enclosed by an intermediate admixing zone, that the
intermediate admixing zone is enclosed and that the intermediate
admixing zone is enclosed by a lean zone.
3. Aircraft gas turbine lean premix burner for carrying out the
method in accordance with claim 1, characterized in that a flame
stabilizer concentrically surrounding the supporting burner is
provided with secondary air recesses.
4. Premix burner in accordance with claim 3, characterized in that
the secondary air supply recesses are provided in the form of
straight or V-shaped slots.
5. Premix burner in accordance with claim 3, characterized in that
the secondary air supply recesses are provided in the form of tubes
(chutes).
Description
[0001] This invention relates to a method for operating a lean
premix burner of an aircraft gas turbine, where fuel and primary
supporting air are supplied by means of a supporting burner (pilot
burner) arranged centrically to the burner axis and where fuel and
air are supplied by means of a main burner.
[0002] It is known from the state of the art to use two fuel
atomizers, i.e. a supporting burner and a main burner, in lean
premix burners. The supporting burner is arranged centrically in
the main burner. The supporting burner is here usually designed as
a pressure swirl atomizer. The lean premix burner includes here two
fuel lines for supplying the supporting burner and the main burner.
In operation, the supporting burner is used for igniting the
aircraft gas-turbine engine and in low load conditions, whereas the
main burner is put into operation at partial load and is used up to
maximum load. The supporting burner is designed here for the
ignition operation and for a stable combustion during the engine
starting phase.
[0003] The state of the art is described in the following in light
of FIG. 2, which shows a burner 32 arranged on a combustion chamber
head 31 and supplying the combustion chamber with fuel and
approximately 10 vol % to 20 vol % of the total air (strictly
speaking this is mass %, but in this case they are identical since
the air has a constant temperature.) As a result, a rich zone 33 is
formed, which is arranged directly downstream of the burner 32. A
further 30 vol % to 40 vol % of air are supplied through mixing air
openings 34 to 37. This results in an air admixture to the rich
flame in the flame zone 38. Downstream of this flame zone 38, a
lean zone 39 is provided. The remaining air of 40 vol % to 50 vol %
is used for cooling and flows through an inner combustion chamber
wall 40 and an outer combustion chamber wall 41, which maintain the
flame. FIG. 2 thus shows a standard burner with a rich zone
supplied with air and followed by a lean zone.
[0004] FIG. 3 shows an embodiment according to the state of the
art, where the burner 32 includes means for mixing air and fuel. A
direct or further flame zone 38 as shown in FIG. 2 can be dispensed
with. The entire burner (supporting burner and main burner) passes
50 vol % to 80 vol % of the total air into the combustion chamber.
The remaining air quantity of 20 vol % to 50 vol % is used for
cooling. The burner includes two fuel circuits and thus permits the
supply of fuel through two concentric fuel atomizers. The
supporting burner 42 with the associated atomizer supplies 5 vol %
to 15 vol % of the total air and creates a small rich zone 33 which
is used for starting the engine and for flame stability. The
concentric main fuel atomizer 43 supplies fuel at medium to maximum
load conditions and 40 vol % to 75 vol % of the total air 44. This
creates a lean zone 39 surrounding the rich zone 33. This lean zone
39 is responsible for low pollutant emissions, in particular of
NOx.
[0005] The main drawback of the solution shown in FIG. 2 is that
high pollutant emissions result, in particular of NOx, and that in
high load conditions soot is emitted, since the combustion
conditions approximate to a stoichiometric or rich combustion
state. It was therefore attempted in the solution described in FIG.
3 to optimize the combustion process by a lean burner concept. This
however has the disadvantage that the supporting burner (pilot
burner) has a reduced flame stability at low output of the aircraft
gas turbine. In medium load conditions, the combustion by the main
burner is too lean to operate effectively, leading to increased
fuel consumption by an aircraft. Additionally, the absence of high
air admixing results in low oxidation of soot, so that considerable
soot quantities are emitted from the aircraft gas turbine at medium
load.
[0006] The object underlying the present invention is to provide a
method--and a device for carrying out the method--for operating a
lean premix burner which avoid the disadvantages of the state of
the art and enable, in particular, a good, stable and low-pollutant
combustion.
[0007] It is a particular object of the present invention to
provide solution to the above problematics by a combination of the
features of the independent Claims. Further advantageous
embodiments of the present invention become apparent from the
sub-claims.
[0008] It is thus provided in accordance with the invention that
between the rich zone and the lean zone an intermediate admixing
zone with a high jet-like admixture of air is formed, into which
zone an additional fuel/air flow is introduced. This results in
optimized combustion in partial load areas too, and has the
advantage that the soot emissions are reduced by improved oxidation
of the soot. Furthermore, there is improved combustion with better
efficiency, since the very lean zones known from the state of the
art are avoided. Due to the intermediate admixing zone, a
combustion zone is created which is closer to stoichiometric
fuel/air ratios. Although this zone is still lean, it avoids the
disadvantages of a too-lean combustion zone.
[0009] In accordance with the invention, there is an enrichment of
the rich zone with a lower air proportion by the supporting burner.
Instead, the additional air is introduced into the intermediate
admixing zone. This leads to good flame stability and good
ignitability of the aircraft gas turbine.
[0010] The present invention is described in the following in light
of the accompanying drawing, showing exemplary embodiments. In the
drawing,
[0011] FIG. 1 shows a schematic representation of a gas-turbine
engine in accordance with the present invention,
[0012] FIG. 2 shows a longitudinal sectional view of a combustion
chamber in accordance with the state of the art,
[0013] FIG. 3 shows a schematic representation of a further variant
of a combustion chamber in accordance with the state of the art by
analogy with FIG. 2,
[0014] FIG. 4 shows a simplified longitudinal sectional view of a
combustion chamber in accordance with a first exemplary embodiment
of the invention by analogy with the representation of FIG. 3,
[0015] FIG. 5 shows a representation of a further exemplary
embodiment by analogy with FIG. 4,
[0016] FIG. 6 shows a sectional view of a further exemplary
embodiment by analogy with FIGS. 4 and 5,
[0017] FIG. 7 shows an enlarged partial representation of the flow
conditions of the inventive solution in accordance with FIG. 6,
[0018] FIGS. 8 to 11 show sectional, front and perspective views of
differing exemplary embodiments of flame stabilizers and secondary
air recesses,
[0019] FIG. 12 shows a graphic representation of the equivalence
ratio as a function of the thrust in accordance with the state of
the art,
[0020] FIG. 13 shows a graphic representation of a lean premix
burner, by analogy with FIG. 12, and
[0021] FIG. 14 shows a representation of the inventive solution by
analogy with FIGS. 12 and 13.
[0022] The gas-turbine engine 10 in accordance with FIG. 1 is a
generally represented example of a turbomachine where the invention
can be used. The engine 10 is of conventional design and includes
in the flow direction, one behind the other, an air inlet 11, a fan
12 rotating inside a casing, an intermediate-pressure compressor
13, a high-pressure compressor 14, a combustion chamber 15, a
high-pressure turbine 16, an intermediate-pressure turbine 17 and a
low-pressure turbine 18 as well as an exhaust nozzle 19, all of
which being arranged about a center engine axis 1.
[0023] The intermediate-pressure compressor 13 and the
high-pressure compressor 14 each include several stages, of which
each has an arrangement extending in the circumferential direction
of fixed and stationary guide vanes 20, generally referred to as
stator vanes and projecting radially inwards from the engine casing
21 in an annular flow duct through the compressors 13, 14. The
compressors furthermore have an arrangement of compressor rotor
blades 22 which project radially outwards from a rotatable drum or
disk 26 linked to hubs 27 of the high-pressure turbine 16 or the
intermediate-pressure turbine 17, respectively.
[0024] The turbine sections 16, 17, 18 have similar stages,
including an arrangement of fixed stator vanes 23 projecting
radially inwards from the casing 21 into the annular flow duct
through the turbines 16, 17, 18, and a subsequent arrangement of
turbine blades 24 projecting outwards from a rotatable hub 27. The
compressor drum or compressor disk 26 and the blades 22 arranged
thereon, as well as the turbine rotor hub 27 and the turbine rotor
blades 24 arranged thereon rotate about the engine axis 1 during
operation.
[0025] FIG. 4 shows in a schematic representation a longitudinal
sectional view of a burner in accordance with the invention. It
includes a supporting burner 42 and a fuel atomizer 43 surrounding
the latter, both forming part of a burner 32 mounted on a
combustion chamber head 31. The combustion chamber includes an
inner combustion chamber wall 40 and an outer combustion chamber
wall 41. Air and fuel are supplied by the supporting burner 42 for
forming a rich zone 33 immediately adjoining the supporting burner
42. In total, an air quantity of approx. 50 vol % to 80 vol % of
the total air is supplied to the combustion chamber by the burner.
The fuel is supplied via two concentric atomizers. Only a small
amount of air (5 vol % to 10 vol % of the total combustion chamber
air) is supplied via the atomizer of the supporting burner 42, thus
ensuring that the rich zone 33 is created.
[0026] The rich zone 33 is delimited and partially enclosed by an
intermediate admixing zone 45. An air quantity of 5 vol % to 20 vol
% of the total combustion air of the combustion chamber is
introduced into the intermediate zone in order to provide a second
zone or secondary supporting zone (intermediate admixing zone) 45
forming a further admixing zone (quenching zone).
[0027] The main fuel atomizer 43 supplies fuel and air. The air
quantity supplied is 35 vol % to 75 vol % of the total combustion
chamber air. The main fuel atomizer 43 is used during medium to
maximum load states of the aircraft gas turbine. By supplying air
and fuel via the main fuel atomizer 43, a lean zone 39 is created
which surrounds the intermediate admixing zone 45 and adjoins the
latter in the axial direction (flow direction).
[0028] FIG. 6 shows a detailed representation of a further
exemplary embodiment of the invention, by analogy with the
representation in FIG. 4. Identical parts are provided with the
same reference numerals, as is the case in the following exemplary
embodiments.
[0029] FIG. 5 shows in detail the supporting burner 42 with a fuel
outlet 47. The supporting burner is concentrically surrounded by an
annular air passage 48, in which a swirler element 49 is arranged.
The escaping air/fuel mixture creates the rich zone 33.
[0030] The intermediate admixing zone 45 is formed by the further
supply of air and fuel. A concentric annulus 50 is provided for
this. The design permits a greater pressure drop, in order to
generate higher air velocities at the place where air is introduced
into the combustion chamber. This results in good mixing with the
rich zone 33. The secondary air supply 51, 52 and 53 can take place
through suitable recesses described in the following in conjunction
with FIGS. 8 to 10.
[0031] The main fuel is supplied through a concentric main air
supply 54 and atomized by the inner air supply 55 and mixed with
the latter. A swirl is imparted by an inner main swirler element
56. The main fuel is also guided through an outer air supply 57 and
atomized and mixed with it, with a swirl being imparted to this air
supply by means of an outer main swirler element 58. The flame
resulting from the main burner surrounds the intermediate admixing
zone 45 and forms a lean zone 39.
[0032] The secondary air can be supplied at different points
(secondary air supply 51, 52 or 53). This supply can take place
singly or in combination.
[0033] FIG. 11 shows an axial longitudinal sectional view plus a
front view of an exemplary embodiment of a burner in accordance
with the invention. It is shown here that the secondary air
recesses 52 can be designed in the form of round holes provided on
a flame stabilizer.
[0034] In the exemplary embodiment in FIG. 9, the secondary air is
supplied by tubes (chutes) provided on the flame stabilizer 59. It
can be supplied in either an axial or a tangential alignment in
order to impart a swirl to the secondary air. Between 4 and 36 of
these outlet tubes (chutes) can be provided, being at angles of
0.degree. and 60.degree. to the burner axis.
[0035] FIG. 11 shows a further design variant in which the
secondary air recesses 52 are designed in the form of slots.
Between 4 and 36 slots can be provided, and can have an angle
between 0.degree. and 60.degree. relative to the burner axis in
order to impart an additional swirl to the air.
[0036] FIG. 8 shows a further exemplary embodiment with V-shaped
slots 52, which can also be provided in a number between 4 and 36.
Here too it is possible to incline the V-shaped slots relative to
the burner axis between 0.degree. and 60.degree. for further
swirling of the air.
[0037] The burner described above can also be designed with an
onflow supporting burner, as is shown in FIG. 6. The supporting
fuel is supplied through a fuel outlet 47. The supporting air is
supplied through an inner air passage 48 with a swirler element 49
and an outer annulus 50 with a swirler element 56.
[0038] FIG. 7 shows that in accordance with the invention a second
supporting flame stabilization zone Y is formed additionally to
zone X and to the rich zone 33. The zone Y leads to an improved
interaction between the supporting burner and the main burner. In
certain operating conditions, the main flame can also be stabilized
in zone Y.
[0039] In accordance with the invention, an additional intermediate
zone is thus created by which combustion in the combustion chamber
can take place in a controlled and optimized way. This leads to the
supporting burner zone being able to operate in a stable manner,
without any fear of the supporting burner being extinguished. The
intermediate zone can be operated even in relatively high load
conditions without soot emissions. Furthermore, the intermediate
admixing zone improves combustion efficiency (total combustion)
during staged operation of the main burner. This leads to a minimum
drop in the efficiency of combustion during a transition from
operation of the supporting burner to combined operation of the
supporting burner and of the main burner.
[0040] In the following, the invention is again explained in
respect of the method in accordance with the invention in light of
FIG. 14, where the illustrations in FIGS. 12 and 13 reflect the
underlying state of the art.
[0041] FIG. 12 shows a diagram in which the thrust is plotted in
percentages as a function of the equivalence ratio between air and
fuel. With an equivalence ratio of 1, there is a stoichiometric
ratio, below 1 to 0 the result is a rich combustion, while above 1
a lean combustion is obtained. These illustrations are also shown
in FIGS. 13 and 14.
[0042] Furthermore, the information for the combustion zones
relates to FIGS. 2 to 5.
[0043] FIG. 12 shows an illustration from the state of the art
which has high emission values. In particular, at high thrust or
high output, respectively, the NOx values are high and there is a
lot of soot. The respective equivalence ratios of the individual
zones achieve here, as is indicated in FIG. 12 for the zones 33, 38
and 39, values having an equivalence ratio of 1 or a rich
equivalence ratio close to the stoichiometric value. By contrast,
the result for the supporting burner is good flame stability.
[0044] To avoid the drawbacks of increased soot formation and high
NOx emissions, solutions were proposed as shown in FIG. 13. While
FIG. 12 in particular relates to the representation in FIG. 2, the
values shown in FIG. 13 are based especially on an embodiment in
accordance with FIG. 3. As shown in FIG. 13, the supporting burner
is operated somewhat more leanly. This leads to a good combustion,
but at the same time generates a lot of soot. At the same time a
reduced stability at low load results from the leaner supporting
burner. As also shown in FIG. 13, the burner is set very lean at
medium thrust, so that in this partial load area or transition area
there is no good combustion in particular in zone 39. A low
efficiency thus applies, and this leads to increased fuel
consumption of an aircraft.
[0045] Furthermore, the absence of the flame zone 38 and the poor
interaction between the mixing air 36 lead to a poor oxidation of
soot, resulting in high soot emissions.
[0046] The drawbacks of the mode of operation shown in FIG. 13 can
be reduced according to the state of the art in that the supporting
burner is enlarged to pass a larger air quantity through the
supporting burner. Soot formation could be reduced by this, but
this has the negative effect that higher NOx emissions result.
Moreover, a leaner operation of the supporting burner leads to a
lower stability. Furthermore, a second supporting burner circuit
with a total of three fuel circuits could be introduced, but this
would increase the complexity of the overall system, involving
additional costs for fuel injection nozzles, fuel systems and
control systems.
[0047] Based on the procedures described above, a completely
different solution was created in accordance with the invention,
and is explained in light of FIG. 14.
[0048] The solution in accordance with the invention was described
above in particular in conjunction with the design solution
according to FIG. 4.
[0049] In accordance with the invention, a secondary supporting
zone or intermediate admixing zone 45 is formed, as explained
above, which is achieved by diverting air/fuel from the rich zone
33. Furthermore, there is a diversion of fuel and air from the
total airflow 44. By doing so, an additional flow 46 is used, as is
shown in FIG. 4.
[0050] The solution in accordance with the invention results in the
following advantages:
[0051] As shown in FIGS. 4 and 5, the addition of the secondary
supporting zone/intermediate admixing zone 45 leads to a reduction
in the soot emissions, caused by an improved oxidation of the soot.
Furthermore, there is an improved combustion efficiency due to the
reduction of very lean zones. The zone of the main burner remains
lean, but zone 45 forms a secondary supporting zone or intermediate
admixing zone which is closer to the stoichiometric fuel/air ratio.
Furthermore, the zone 45 leads to a reduction in soot formation.
The zone 33 (zone of the supporting burner 42) can be operated with
a richer fuel/air mixture than in the solutions known from the
state of the art. This leads to an improved flame stability.
[0052] As explained above, the core of the invention is the
additional introduction of a secondary supporting zone or
intermediate admixing zone 45. This leads to the supporting burner
being capable of operation with a constant combustion zone, thereby
ensuring stable operation and preventing the flame from being
extinguished (flame-out). Both the supporting burner zone and the
secondary supporting zone/intermediate admixing zone 45 can be
operated without problems resulting with regard to soot emissions.
Furthermore, the secondary supporting zone/intermediate admixing
zone 45 improves the efficiency of combustion during a staged
operation of the main burner. This leads to a minimum reduction in
the combustion efficiency during the transition from operation with
the supporting burner to combined operation of the supporting
burner and of the main burner.
LIST OF REFERENCE NUMERALS
[0053] 1 Engine axis [0054] 10 Gas-turbine engine/core engine
[0055] 11 Air inlet [0056] 12 Fan [0057] 13 Intermediate-pressure
compressor (compressor) [0058] 14 High-pressure compressor [0059]
15 Combustion chamber [0060] 16 High-pressure turbine [0061] 17
Intermediate-pressure turbine [0062] 18 Low-pressure turbine [0063]
19 Exhaust nozzle [0064] 20 Guide vanes [0065] 21 Engine casing
[0066] 22 Compressor rotor blades [0067] 23 Stator vanes [0068] 24
Turbine blades [0069] 26 Compressor drum or disk [0070] 27 Turbine
rotor hub [0071] 28 Exhaust cone [0072] 31 Combustion chamber head
[0073] 32 Burner [0074] 33 Rich zone [0075] 34, 35, 36, 37 Mixing
air [0076] 38 Flame zone [0077] 39 Lean zone [0078] 40 Inner
combustion chamber wall [0079] 41 Outer combustion chamber wall
[0080] 42 Supporting burner [0081] 43 Fuel atomizer [0082] 44 Total
air [0083] 45 Secondary supporting zone/intermediate admixing zone
[0084] 46 Additional flow [0085] 47 Fuel outlet [0086] 48 Air
passage [0087] 49 Swirler element [0088] 50 Annulus [0089] 51, 52,
53 Secondary air supply/secondary air recesses [0090] 54 Concentric
main air supply [0091] 55 Inner air supply of main burner [0092] 56
Inner main swirler element [0093] 57 Outer air supply of main
burner [0094] 58 Outer main swirler element [0095] 59 Flame
stabilizer [0096] 60 Supporting air supply
* * * * *