U.S. patent application number 09/848701 was filed with the patent office on 2002-11-07 for partial premix dual circuit fuel injector.
This patent application is currently assigned to Honeywell International, Inc., Law Dept. AB2. Invention is credited to Zelina, Joseph.
Application Number | 20020162333 09/848701 |
Document ID | / |
Family ID | 25304039 |
Filed Date | 2002-11-07 |
United States Patent
Application |
20020162333 |
Kind Code |
A1 |
Zelina, Joseph |
November 7, 2002 |
Partial premix dual circuit fuel injector
Abstract
A low emission fuel injection system and combustion chamber for
use in gas turbine engines comprises one fuel injection body having
a dual circuit to supply both pilot and main fuel systems. Both
pilot fuel circuit and a main fuel circuit inject fuel at
essentially the same axial and radial location. The recessed pilot
fuel injection site is along the combustor centerline into a
swirling air passage produced by axial air swirlers. The main fuel
is injected radially through a plurality of injection sites, at a
compound angle, into the inner diameter of a swirling air passage
produced by radial air swirlers. The fuel/air residence time prior
to entering the combustion chamber is relatively short, minimizing
the likelihood of auto ignition. During pilot circuit only
operation, the flame is stabilized by a swirler produced
recirculation zone, producing high temperatures to completely burn
the fuel producing low CO and UHC emissions. During intermediate
and high engine power conditions, both the main fuel and pilot
circuits discharge fuel into a swirler produced, high air flow,
recirculation zone producing a fuel lean, low temperature flame to
reduce NOx emissions.
Inventors: |
Zelina, Joseph; (Kettering,
OH) |
Correspondence
Address: |
Honeywell International, Inc.
Law Dept. AB2
P.O. Box 2245
Morristown
NJ
07962-9806
US
|
Assignee: |
Honeywell International, Inc., Law
Dept. AB2
Morristown
NJ
|
Family ID: |
25304039 |
Appl. No.: |
09/848701 |
Filed: |
May 2, 2001 |
Current U.S.
Class: |
60/776 ;
60/737 |
Current CPC
Class: |
F23R 3/343 20130101;
F23R 3/286 20130101; Y02T 50/60 20130101; F23R 3/14 20130101; Y02T
50/675 20130101 |
Class at
Publication: |
60/776 ;
60/737 |
International
Class: |
F23R 003/30 |
Goverment Interests
[0001] The invention described herein was made in the performance
of work under NASA Contract No. NAS3-27752 awarded by NASA Glenn
Research Center, and is subject to the provisions of Section 305 of
the National Aeronautics and Space Act of 1958 (42 U.S.C. 2457).
The Government has certain rights in this invention.
Claims
We claim:
1. A gas turbine engine fuel injection and combustor system,
comprising: an outer casing extending from an upstream end to a
downstream end, an internal space of the downstream end defining a
combustion chamber; an annular dome connected to an internal,
upstream end of the outer casing; a radial flow air swirler mounted
to an internal surface of the annular dome and to an external
surface of a fuel injector body, said swirler providing swirled air
to the combustion chamber; a fuel injector body mounted to an
internal surface of the radial flow air swirler, said fuel injector
body comprising a plurality of axially oriented air swirlers, an
axially located pilot circuit fuel nozzle, and a plurality of
radially oriented main circuit fuel nozzles; whereby during low
power gas turbine engine operation, pilot circuit fuel is mixed
with axial swirler produced air flow, and is thereby caused to
rotate in a vortex around a longitudinal axis of the combustion
chamber and upon ignition produces a stable high temperature
emission free flame; whereby during intermediate and high power gas
turbine engine operation, pilot circuit and main circuit fuel is
mixed with radial and axial swirler produced air flow, and is
thereby caused to rotate in a double vortex around the longitudinal
axis of said combustion chamber, and upon ignition produces a
stable, fuel lean, low temperature, low pollutant emission
flame.
2. The gas turbine engine fuel injection and combustor system of
claim 1, whereby a single fuel injector body incorporates both the
pilot circuit and the main circuit fuel injection systems.
3. The gas turbine engine fuel injection and combustor system of
claim 1, whereby the fuel injector body for the pilot and main
circuit fuel injection systems is cylindrical.
4. The gas turbine engine fuel injection and combustor system of
claim 1, whereby pilot circuit fuel and main circuit fuel are
injected in the combustion chamber at essentially the same axial
and radial locations.
5. The gas turbine engine fuel injection and combustor system of
claim 1, whereby a pilot circuit fuel nozzle encircled by a
plurality of axial air swirlers discharges fuel into a combustor
creating a rotating vortex of vaporized fuel and air.
6. The gas turbine engine fuel injection and combustor system of
claim 1, whereby main circuit fuel nozzles discharge fuel at a
compound angle into a radial swirler passage creating a rotating
vortex of vaporized fuel and air within the combustion chamber.
7. The gas turbine engine fuel injection and combustor system of
claim 1, whereby the combined geometry of the pilot circuit and
main circuit fuel injection systems provides short fuel-air
residence times prior to entering the combustion chamber.
8. The gas turbine engine fuel injection and combustor system of
claim 1, whereby at least four radially oriented main circuit fuel
nozzles are employed.
9. The gas turbine engine fuel injection and combustor system of
claim 1, whereby said fuel injection system is usable for both can
type and annular type combustors.
10. The gas turbine engine fuel injection and combustor system of
claim 1, whereby aerospace type fuels can be utilized can be
utilized as the combustion medium.
11. The gas turbine engine fuel injection and combustor system of
claim 1, whereby a variety of non-aerospace type fuels can be
utilized as the combustion medium.
12. A method of producing a low emission gas at low and high power
in a gas turbine engine combustor, comprising the steps of:
injecting pilot circuit fuel spray through a axially disposed
nozzle into a combustor chamber; swirling a first portion of air
axially and flowing said swirled air into said combustor chamber to
form a swirling mixture of pilot circuit fuel and air for low,
intermediate, and high power operations; swirling a second portion
of air radially and flowing said swirled air into an air swirler
passageway; injecting main circuit fuel spray through radially
disposed nozzles into said air swirler passageway and thereafter
into said combustion chamber to form a swirling mixture of main
circuit fuel and air for intermediate and high power operations;
expanding and igniting said swirling mixture of pilot circuit fuel
and air to form said low emission gas for low and high power
operations; and expanding and igniting said swirling mixture of
main circuit fuel and air to form said low emission gas for
intermediate and high power operations.
13. The method of claim 12, wherein the step of expanding pilot
circuit fuel and air forms a fuel rich, high flame temperature,
zone in the combustion chamber.
14. The method of claim 12, wherein the step of expanding main
circuit fuel and air forms a fuel lean, low flame temperature, zone
in the combustion chamber.
Description
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to fuel injectors used in
gas turbine engine combustors and, in particular, to a fuel
injector design primarily for aerospace applications, which
produces a stable flame at low power, generating low CO and UHC
pollutants, and also provides enhanced fuel-air mixing to reduce
NOx emissions at high-power.
[0003] Atmospheric pollution concerns and effects worldwide have
led to enactment of increasingly stricter emission controls and
standards requiring most industries to significantly reduce the
emissions of pollutants. The strict emission controls have required
implementation of numerous design changes in gas turbines used for
industrial, power generation and propulsion applications. Thus
typical gas turbine engines are now required to operate efficiently
over a wide range of conditions while at the same time producing
minimal quantities of noxious emissions. The common precursors to
gas turbine engines cause atmospheric pollution include Carbon
Monoxide (CO), and Unburned Hydrocarbons (UHC) at low engine power
conditions, and Nitrous Oxide (NOx) at intermediate and high engine
power conditions.
[0004] Reductions in gas turbine emissions of NOx have been
obtained primarily by the reduction of flame temperatures in the
combustor. Some of the techniques employed include lean burn
pre-mix combustors in which the fuel to air ratio is reduced as far
as possible in the higher operating range; staged combustors,
whereby fuel is admitted to the combustion chamber at different
axial locations thereby staging the combustion process; lean-direct
injection involving methods of injecting fuel and air into a
combustion chamber where the mixture is fuel-lean, or below the
stoichiometric fuel/air ratio; and other related techniques known
to those skilled in the art as "rich burn" and "quick quench".
[0005] Gas turbine propulsion engines employ annular and can-type
combustors to burn fuel. The fuel is metered and sprayed into the
combustor through a single or a plurality of fuel nozzles along
with combustion air having a designated amount of swirl. In the
typical gas turbine engine, flame stability, variable cycle
operation, and emission control dominate combustor design
requirements. The characteristics of a given fuel injector under
light-up and low speed conditions are different to those under full
power conditions. Consequently a fuel injector is often a
compromise between two designs to enable it to operate under both
of these conditions.
[0006] Typically, fuel is supplied through one supply duct under
starting or low power conditions and through another or through
both fuel supply ducts under high power conditions. During light up
and low speed conditions, only the pilot fuel injectors are used
whereas both the pilot and the main fuel injectors are used under
higher speed conditions. The pilot combustion stage is usually long
in comparison with the main combustion stage. Consequently, the
residence time of the fuel in the pilot stage is comparatively
long, limiting emissions of hydrocarbons and carbon monoxide.
Conversely, the residence time of the fuel in the main stage is
comparatively short, limiting emissions of the oxides of
nitrogen.
[0007] Many techniques have been utilized in the prior art to
enable efficient gas turbine engine operation over a wide range of
conditions while at the same time producing minimal quantities of
noxious emissions. Of particular interest in this regard are the
following references and examples:
[0008] U.S. Pat. No. 4,701,124 discloses a can-type combustor
intended for industrial applications and designed to use oil or
natural gas as the fuel. The fuel is injected at a compound angle
in the main fuel chamber at the air swirler vane exit plane.
[0009] U.S. Pat. No. 5,062,792 also targets industrial burners and
uses oil or natural gas as the fuel. In this design, the oil is
used to fuel the pilot circuit, while the natural gas is used to
fuel the main circuit.
[0010] U.S. Pat. No. 5,069,029 involves a staged combustor, wherein
fuel is admitted to the combustion chamber at different axial
locations, thereby staging the combustion process. Axial-type air
swirlers are utilized for both pilot and main fuel circuits.
[0011] U.S. Pat. No. 5,816,050 discloses a can-type combustor with
pilot fuel system that injects the fuel axially into the swirling
air stream, which is kept separate from the main air stream by a
containment wall. Separate fuel injectors are utilized for the
pilot and main systems, the latter injecting the fuel near the
outer diameter of the air swirler.
[0012] U.S. Pat. No. 5,862,668 teaches a radial-staged combustion
system, whereby the pilot system is operated at low power (inward
and radial location) and the main and pilot are operated at high
power conditions. The combustor is considered a premix type system
due to the longer residence time of fuel and air in the premixing
tubes.
[0013] U.S. Pat. No. 6,151,899 involves a can-type combustion
chamber whereby main fuel is injected axially into the swirling air
passage, and pilot fuel is injected in two discreet axial and
radial locations.
[0014] U.S. Pat. No. 6,158,223 also teaches a can-type combustion
chamber providing for a pilot fuel injection point in the same
plane as the main fuel/air injection point. The design utilizes a
plurality of main air swirlers for each pilot air swirler.
[0015] U.S. patent application Ser. No. 09/492,678, filed Jan. 27,
2000, for Combustor with Fuel Preparation Chambers, by the same
inventor as this application, and also assigned to the assignee of
this application, discloses an annular combustor having fuel
preparation chambers mounted in the dome of the combustor with the
objective of uniformly mixing fuel and air so as to reduce NOx
formed by the ignition of the fuel/air mixture.
[0016] With the exception of the last reference, none of the
above-cited prior art is specifically intended for lightweight,
aerospace type applications, and some suffer from one or more of
the following disadvantages:
[0017] a) aerospace type fuels cannot be utilized.
[0018] b) fuel-air residence times prior to entering the combustion
chamber are relatively long, thus increasing the danger of
autoignition or flashback situations, which are a safety concern
for aerospace applications.
[0019] c) complexity in fabrication and operation.
[0020] d) inability to retrofit the system on existing engines.
[0021] e) the fuel injection system is usable for can type or
annular combustors, but not for both.
[0022] For the foregoing reasons, as can be seen, there is a need
for a simple, retrofittable, gas turbine engine fuel injector
system primarily for aerospace applications, which produces a clean
stable flame at low power, provides enhanced fuel-air mixing to
reduce noxious emissions at high-power, and can be employed on both
can-type and annular type combustion systems.
SUMMARY OF THE INVENTION
[0023] In one aspect of the present invention, a fuel injector
system is provided to produces a stable flame at low power
operations, generating low CO and UHC pollutants.
[0024] In another aspect of the present invention, a fuel injector
system is provided to produce enhanced fuel-air mixing to reduce
NOx emissions at intermediate and high power operations.
[0025] In a further aspect of the present invention, a gas turbine
engine fuel injection system, comprised of a single fuel injection
body incorporating both pilot circuit and main circuit fuel
injectors, is provided that produces minimal quantities of CO, UHC
and NOx emissions.
[0026] In yet another aspect of the present invention, a simple,
low cost gas turbine engine fuel injection system is provided that
is capable of producing minimal quantities of noxious emissions,
and which can be retrofitted into existing products, including
engines for aerospace applications.
[0027] In a still further aspect of the present invention, a gas
turbine engine fuel injection system is provided that is capable of
producing minimal quantities of noxious emissions, and which can be
utilized on annular as well as can-type combustors.
[0028] The present invention is comprised of one fuel injection
body to supply both pilot and main fuel systems. The fuel injection
body is further comprised of a pilot fuel circuit and a main fuel
circuit, both of which inject fuel at essentially the same axial
and radial location. The recessed pilot fuel injection site is
along the combustor centerline into a swirling air passage produced
by axial air swirlers. The main fuel is injected radially through a
plurality of injection sites, at a compound angle, into the inner
diameter of a swirling air passage produced by radial air swirlers.
One main air swirler is utilized for each pilot air swirler. The
fuel/air residence time prior to entering the combustion chamber is
relatively short, minimizing the likelihood of auto ignition.
[0029] A method for producing a low noxious emission gas in a
combustor is also disclosed. The method includes injecting pilot
circuit fuel spray into a swirling passage along the combustor
centerline for low power operations. For intermediate and high
power operations, main circuit fuel spray, together with pilot
circuit fuel spray, is discharged into a radial swirler air
passage, and thereafter fluidly communicates with a single
combustion chamber. The swirling mixture of pilot circuit and main
fuel and air enters the combustion chamber, is expanded, and
ignited to form low emission gas.
[0030] These and other objects, features and advantages of the
present invention, are specifically set forth in, or will become
apparent from, the following detailed description of a preferred
embodiment of the invention when read in conjunction with the
accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0031] FIG. 1 is a longitudinal cross section of a gas turbine
engine combustor assembly having the inventive partial premix dual
circuit fuel injector;
[0032] FIG. 2 is a perspective view of the inventive partial premix
dual circuit fuel injector;
[0033] FIG. 3 is a cross section of the combustor assembly showing
air and fuel flow patterns during low power operation; and
[0034] FIG. 4 is a cross section of the combustor assembly showing
air and fuel flow patterns during intermediate and high power
operation.
DETAILED DESCRIPTION OF THE INVENTION
[0035] The following detailed description is of the best currently
contemplated modes of carrying out the invention. The description
is not to be taken in a limiting sense, but is made merely for the
purpose of illustrating the general principles of the invention,
since the scope of the invention is best defined by the appended
claims.
[0036] Referring to FIG. 1, there is shown a longitudinal cross
section of a gas turbine engine combustor assembly 10, having an
upstream end 11 and a downstream end 12. The combustor fuel mixing
zone 13 is located along the downstream end 12. The combustor
assembly 10 is cylidrically shaped and bounded along its
longitudinal outer periphery by a combustor outer casing housing 14
which is attached along its internal, upstream end 11 to an annular
sheet metal dome 16. The sheet metal dome 16 serves as a connecting
member as well as a shield, protecting upstream end 11 components
from exposure to excessive temperatures generated at the combustor.
Thermal protection is provided by a heat shield 17 and cooling
skirt 18 as well as a plurality of circumferentially disposed dome
cooling nozzles 19 which dispense cooling air flow 19A from the gas
turbine engine compressor stage.
[0037] Still referring to FIG. 1, there is shown an annular swirler
housing 26 rigidly connecting the sheet metal dome 16 to a fuel
injector body 15 which is located along the longitudinal centerline
10A of the combustor assembly 10. The radial air swirler housing 26
accommodates a plurality of radial swirlers 27 peripherally located
along the upstream end 11, each said radial swirler 27 provided to
bring air flow, designated by arrows 28, supplied under pressure
from a compressor of the gas turbine engine (not shown). The radial
swirler 27 is adapted to bring air flowing inwardly radially,
designated by arrows 29, therethrough to rotate through and around
the radial swirler passage 30 and into the combustion chamber's
fuel mixing zone 13.
[0038] The fuel injector body 15 incorporates both the pilot
circuit fuel injection and the main circuit fuel injection systems.
Pilot circuit burner fuel 20 enters the pilot circuit fuel nozzle
21 from the upstream end 11 along the combustor centerline 10A and
into the pilot circuit fuel nozzle 21 where it is atomized and
thereafter dispersed, through the convergent-divergent end 15A,
into the fuel mixing zone 13 in the direction essentially as shown
by the pilot circuit fuel flow arrows 25. A plurality of axial
pilot circuit air swirlers 22, located radially outboard of the
pilot circuit fuel nozzle 21, swirls air flow 23 provided from a
compressor of the gas turbine engine (not shown), and routes it, as
shown by arrows 24, to the convergent-divergent end 15A, where it
will mix with the pilot circuit fuel flow 25, to be thereafter
ignited.
[0039] Main circuit burner fuel 31 enters the main circuit fuel
nozzle 32 from the upstream end 11, is routed through a passageway
33 bored approximately parallel to the combustor assembly
centerline 10A, and rotated approximately 90 degrees to exit the
main circuit fuel nozzle 32 at an outwardly radial direction. The
main circuit fuel flow represented by arrows 34 is atomized as it
exits the main circuit fuel nozzle 32, is thereafter mixed with the
radial swirler air flow 29 in the radial swirler passage 30, and is
then routed to the fuel mixing zone 13 to be thereafter ignited. In
the preferred embodiment, main fuel is introduced through at least
four main circuit fuel nozzles 32 symmetrically spaced around the
outer periphery of the fuel injector body 15.
[0040] FIG. 2 is a perspective view showing various central
elements of the partial premix dual circuit fuel injector assembly.
For clarity purposes, the combustor outer casing 14 housing is not
shown. The radial swirler exit plane 30A and radial swirler
attachment (to combustor dome) flange 30B establish the outer
boundary of the fuel injector assembly. Multiple main fuel circuit
radial swirlers 27 are located along the entire outer periphery
adjacent to the base of the radial swirler passage 30. The fuel
injector body 15 is centrally located and is comprised of the
recessed pilot circuit fuel nozzle 21 which is itself peripherally
surrounded by a plurality of axial pilot circuit air swirlers 22.
Two main circuit fuel nozzles 32 are shown located along the
external periphery of the fuel injector body 15. In the preferred
embodiment, at least four symmetrically spaced main circuit fuel
nozzles 32 are used. Alternate embodiments may use fewer or
additional main circuit fuel nozzles 32.
[0041] Referring now to FIG. 3, there is shown a cross section of
the combustor assembly showing air and fuel flow patterns during
low power operation. Pilot circuit fuel entering the fuel mixing
zone 13 is sprayed as represented by arrows 25. The swirler air
flow 24 produced by the axial pilot circuit air swirlers 22 mixes
with the fuel and is caused to recirculate in the general direction
shown by arrows 35 forming a vortex recirculation zone 36. When the
fuel air mixture is ignited, the fuel rich flame stabilized by the
recirculation zone 36, produces high flame temperatures, helping to
improve combustor efficiency by completely burning the fuel and
eliminating CO and UHC emissions. Flame stabilization is also
somewhat aided by the double vortex effect of radial swirler 27
recirculation air, represented by arrows 37, which forms a radial
swirler air vortex recirculation zone 38. Main circuit fuel does
not flow during low power operation.
[0042] Referring now to FIG. 4, there is shown a cross section of
the combustor assembly showing air and fuel flow patterns during
intermediate and high power operation. Fuel entering the fuel
mixing zone 13 is sprayed as represented by arrows 25 for the pilot
circuit and by arrows 34 for the main circuit. The swirler air flow
24 produced by the axial pilot circuit air swirlers 22 mixes with
the pilot circuit fuel and is caused to recirculate in the general
direction shown by arrows 35 forming a vortex recirculation zone
36. Radial swirler 27 air flow, represented by arrows 29, mixes
with the main circuit fuel flow 34 and the mixture is caused to
recirculate in the general direction shown by arrows 39 forming a
separate vortex recirculation zone 40. When the pilot circuit and
main circuit fuel air mixtures are ignited, the flame stabilized by
recirculation zones 36 and 40 produces relatively low flame
temperatures since the mixture is fuel lean, with the result that
NOx emissions are materially reduced.
[0043] The previously described versions of the present invention
have many advantages, including:
[0044] a. the capability of utilizing not only aerospace type fuels
but other types of fuels as well;
[0045] b. the ability to retrofit the inventive fuel injection
system on existing gas turbine engines;
[0046] c. the simplicity and corresponding low cost of the
apparatus whereby a single fuel injection body incorporating both
pilot circuit and main circuit fuel injectors is provided;
[0047] d. a design which provides for fuel-air residence times
prior to entering the combustion chamber to be relatively short,
thus reducing the danger of autoignition or flashback situations,
which are a safety concern for aerospace applications; and
[0048] e. the capability of utilizing the inventive fuel injection
system for both can type or annular combustors.
[0049] Although the present invention has been described in
considerable detail with reference to certain preferred versions
thereof, other versions are possible. Therefore, the spirit and
scope of the appended claims should not be limited to the
description of the preferred versions contained therein.
* * * * *