U.S. patent application number 09/759194 was filed with the patent office on 2002-07-04 for combustion cap with integral air diffuser and related method.
Invention is credited to Bourgeois, Richard Scott, Chandler, Christopher Nelson, Dean, Anthony John.
Application Number | 20020083711 09/759194 |
Document ID | / |
Family ID | 25054739 |
Filed Date | 2002-07-04 |
United States Patent
Application |
20020083711 |
Kind Code |
A1 |
Dean, Anthony John ; et
al. |
July 4, 2002 |
Combustion cap with integral air diffuser and related method
Abstract
A combustion cap assembly for closing a forward end of a
combustion chamber includes a radially inner substantially
cylindrical component; a radially outer substantially conical
component, extending substantially along an entire length dimension
of the radially inner component; and an annular airflow passage
therebetween. The invention also provides a method for reducing
pressure loss across a combustion liner cap assembly located in a
gas turbine combustor, the cap assembly supporting a plurality of
premix tubes adapted to receive portions of a like number of
nozzles, and wherein air flows in an annular passage radially
outwardly of the combustor where it reverses direction to flow
through the premix tubes, the method including adding a diffuser to
the forward end of the cap assembly, the diffuser configured to
increase the cross sectional area of the annular flow passage along
an axial length of the cap assembly to thereby cause a reduction in
velocity of the air in the annular flow passage and thereby reduce
pressure loss as the air reverses direction at a forward end of the
combustor.
Inventors: |
Dean, Anthony John; (Scotia,
NY) ; Chandler, Christopher Nelson; (Delmar, NY)
; Bourgeois, Richard Scott; (Albany, NY) |
Correspondence
Address: |
Michael J. Keenan
Nixon & Vanderhye P.C.
1100 North Glebe Road, 8th Floor
Arlington
VA
22201-4714
US
|
Family ID: |
25054739 |
Appl. No.: |
09/759194 |
Filed: |
December 28, 2000 |
Current U.S.
Class: |
60/737 ;
60/752 |
Current CPC
Class: |
F23R 3/286 20130101;
F23R 3/283 20130101; F23R 3/04 20130101 |
Class at
Publication: |
60/737 ;
60/752 |
International
Class: |
F23R 003/30 |
Claims
What is claimed is:
1. A combustion cap assembly for closing a forward end of a
combustion chamber comprising: a radially inner substantially
cylindrical component; a radially outer substantially conical
component, extending substantially along an entire length dimension
of said radially inner component; and an annular airflow passage
therebetween.
2. The combustion cap assembly of claim 1 wherein said radially
inner component and said radially outer component are separated by
a plurality of circumferentially spaced webs.
3. The combustion cap assembly of claim 1 wherein said annular
airflow passage increases in cross sectional area in a flow
direction.
4. The combustion cap assembly of claim 1 wherein said radially
outer component engages a portion of a forward turbine case
component.
5. The combustion cap assembly of claim 1 and further comprising a
plate supporting a plurality of premix burner tubes radially inward
of said radially inner component.
6. The combustion cap assembly of claim 1 including a radial flange
located between forward and aft case components.
7. A combustion cap assembly for closing a forward end of a
combustion chamber comprising: a radially inner substantially
cylindrical component; a radially outer substantially conical
component, extending substantially along an entire length dimension
of said radially inner component: and an annular airflow passage
therebetween; wherein said annular airflow passage increases in
cross sectional area in a flow direction; and further comprising a
plate supporting a plurality of premix burner tubes radially inward
of said radially inner cylindrical component.
8. The combustion cap assembly of claim 7 wherein said radially
inner component and said radially outer component are separated by
a plurality of circumferentially spaced webs.
9. The combustion cap assembly of claim 8 wherein said radially
outer component engages a portion of a forward turbine case
component.
10. The combustion cap assembly of claim 7 including a radial
flange located between forward and aft case components.
11. A method of reducing pressure loss across a combustion liner
cap assembly located in a gas turbine combustor, the cap assembly
supporting a plurality of premix tubes adapted to receive portions
of a like number of nozzles, and wherein air flows in an annular
passage radially outwardly of the combustor where it reverses
direction to flow through the premix tubes, the method comprising
adding a diffuser to the forward end of the cap assembly, said
diffuser configured to increase the cross sectional area of said
annular flow passage along an axial length of the cap assembly to
thereby cause a reduction in velocity of the air in said annular
flow passage and thereby reduce pressure loss as the air reverses
direction at a forward end of the combustor.
Description
[0001] This invention relates generally to gas turbine machines and
specifically, to a combustion cap assembly for a multi-nozzle,
can-annular combustor.
BACKGROUND OF THE INVENTION
[0002] Gas turbines generally include a compressor, one or more
combustors, a fuel injection system and a turbine. Typically, the
compressor pressurizes inlet air which is then turned in direction
or reverse flowed to the combustors where it is used to cool the
combustor and also to provide air to the combustion process. In a
multi-combustor turbine, the combustors are located about the
periphery of the gas turbine, and a transition duct connects the
outlet end of each combustor with the inlet end of the turbine to
deliver the hot products of the combustion process to the
turbine.
[0003] Generally, in Dry Low NOx combustion systems utilized by the
assignee, each combustor includes multiple fuel nozzles, each
nozzle having a surrounding dedicated premix section or tube so
that, in a premix mode, fuel is premixed with air prior to burning
in the single combustion chamber. In this way, the multiple
dedicated premixing sections or tubes allow thorough premixing of
fuel and air prior to burning, which ultimately results in low NOx
levels. See, for example, commonly owned U.S. Pat. No.
5,274,991.
[0004] More specifically, each combustor includes a generally
cylindrical casing having a longitudinal axis, the casing having
fore and aft sections secured to each other, and the casing as a
whole secured to the turbine casing. Each combustor also includes
an internal flow sleeve, and a combustion liner substantially
concentrically arranged within the flow sleeve. Both the flow
sleeve and combustion liner extend between the transition duct at
their downstream ends, and a combustion liner cap assembly (located
within an upstream portion of the combustor) at their upstream
ends. The flow sleeve is attached directly to the combustor casing,
while the liner supports the liner cap assembly which, in turn, is
fixed to the combustor casing. The outer wall of the transition
duct and at least a portion of the flow sleeve are provided with
air supply holes over a substantial portion of their respective
surfaces, thereby permitting compressor air to enter the radial
space between the combustion liner and the flow sleeve, and to be
reverse flowed to the upstream portion of the combustor.
[0005] A plurality (five in the exemplary embodiment) of
diffusion/premix fuel nozzles are arranged in a circular array
about the longitudinal axis of the combustor casing. These nozzles
are mounted in a combustor end cover assembly which closes off the
rearward end of the combustor. Inside the combustor, the fuel
nozzles extend into and through the combustion liner cap assembly
and, specifically, into corresponding ones of the premix tubes that
are secured in the liner cap assembly. The discharge end of each
nozzle terminates within a corresponding premix tube, in relatively
close proximity to the downstream end of the premix tube which
opens to the burning zone in the combustion liner.
[0006] Spacers between the cap's inner body and its outer mounting
flange create an annular passage for premixer air from the
compressor. The premixer air travels through this annular passage,
then again reverses direction within the combustor's forward case
before mixing with gaseous fuel in the inner body of the liner cap
assembly, and proceeding to the reaction zone. This air flow
reversal (commonly referred to as the "cap turn") results in a
pressure loss, which can be as high as 7% of the total combustor
pressure drop. The cap turn pressure loss is a result of two
effects: (1) expansion of premixer air into the forward case area
after passing the cap, and (2) reversal of flow direction within
the forward case to travel through the cap burner tubes.
[0007] Pressure loss in the combustor is a critical contributor to
overall gas turbine performance. Any air the combustion system uses
for cooling or loses to leakage is counted against the budgeted
overall combustion system pressure drop.
[0008] Previous combustor designs have implemented tapered flanges
on the cap assembly to allow some degree of flow expansion prior to
the cap turn. However, the amount of flow expansion was relatively
small, as the diffuser section was only as long as the cap mounting
flange.
BRIEF SUMMARY OF THE INVENTION
[0009] In this invention, the forward portion of the combustion
liner cap assembly is designed as an axial diffuser to reduce the
pressure drop caused by the cap turn as the premixer air passes
between inner and outer bodies or components of the liner cap
assembly and turns toward the fuel nozzles and the combustor
chamber.
[0010] More specifically, this invention provides a combustion
liner cap assembly with a conical outer body that serves to
increase the cross-sectional area of the annular passage between a
cylindrical inner body and the conical outer body in the direction
of airflow, causing a reduction in the velocity of the premixer air
as it passes through the cap assembly. These cap assembly
modifications, in turn, require an enlarged forward case to
accommodate the cap diffuser.
[0011] As mentioned above, the cap turn pressure loss is due to
expansion of premixer air and the reversal of flow at the forward
case. Since the magnitude of the pressure losses is proportional to
the square of the air velocity, the reduction of air velocity
caused by the axial diffuser results in a lower cap turn pressure
loss. In addition, the diffuser improves flow uniformity into the
premixers because the flow begins turning from the forward end of
the diffuser inner cylinder rather than at the inlets to the
premixer tubes. Another expected benefit of this concept is
improved flame holding.
[0012] Accordingly, this invention relates to a combustion cap
assembly for closing a forward end of a combustion chamber
comprising a radially inner substantially cylindrical component: a
radially outer substantially conical component, extending
substantially along an entire length dimension of the radially
inner component; and an annular airflow passage therebetween.
[0013] The invention also relates to a combustion cap assembly for
closing a forward end of a combustion chamber comprising a radially
inner substantially cylindrical component; a radially outer
substantially conical component, extending substantially along an
entire length dimension of the radially inner component; and an
annular airflow passage therebetween; wherein the annular airflow
passage increases in cross sectional area in a flow direction; and
further comprising a plate supporting a plurality of premix burner
tubes radially inward of said radially inner cylindrical
component.
[0014] The invention also relates to a method of reducing pressure
loss across a combustion liner cap assembly located on a gas
turbine combustor, the cap assembly supporting a plurality of
premix tubes adapted to enclose portions of a like number of
nozzles, and wherein air flows in an annular passage radially
outwardly of the combustor where it reverses direction to flow
through the premix tubes, the method comprising adding a diffuser
to the forward end of the cap assembly, the diffuser configured to
increase the cross sectional area of the annular flow passage along
an axial length of the cap assembly to thereby cause a reduction in
velocity of the air in the annular flow passage and thereby reduce
pressure loss as the air reverses direction at the forward end of
the combustor.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] FIG. 1 is a partial cross section of a known gas turbine
combustor;
[0016] FIG. 2 is a perspective view of a combustion liner cap
assembly in accordance with the invention;
[0017] FIG. 3 is a cross section of the combustion liner cap
assembly shown in FIG. 2; and
[0018] FIG. 4 is an upstream end view of the liner cap assembly
shown in FIG. 2.
DETAILED DESCRIPTION OF THE INVENTION
[0019] With reference to FIG. 1, a conventional gas turbine 10
includes a compressor 12 (partially shown), a plurality of
combustors 14 (one shown), and a turbine represented here by a
single blade 16. Although not specifically shown, the turbine is
drivingly connected to the compressor 12 along a common axis. The
compressor 12 pressurizes inlet air which is then reverse flowed to
the combustor 14 where it is used to cool the combustor and to
provide air to the combustion process.
[0020] As noted above, the gas turbine includes a plurality of
combustors 14 located about the periphery of the gas turbine. A
double-walled transition duct 18 connects the outlet end of each
combustor with the inlet end of the turbine to deliver the hot
products of combustion to the turbine. Ignition is achieved in the
combustors by means of a spark plug 20 in conjunction with
crossfire tubes (represented by aperture 22) that transfer the
flame to adjacent combustion in conventional fashion.
[0021] Each combustor 14 includes a substantially cylindrical
combustion casing 24 which is secured at an open aft end to the
turbine casing 26 by means of bolts 28. The forward end of the
combustion casing is closed by an end cover assembly 30 which may
include conventional supply tubes, manifolds and associated valves,
etc. for feeding gas, liquid fuel and air (and water if desired) to
the combustor. The end cover assembly 30 receives a plurality (for
example, five) fuel nozzle assemblies 32 (only one shown for
purposes of convenience and clarity) arranged in a circular array
about a longitudinal axis of the combustor.
[0022] Within the combustor casing 24, there is mounted, in
substantially concentric relation thereto, a substantially
cylindrical flow sleeve 34 that connects at its aft end to the
outer wall 36 of the double walled transition duct 18. The flow
sleeve 34 is connected at its forward end to the combustor casing
24 at a butt joint where fore and aft sections of the combustor
casing are joined.
[0023] Within the flow sleeve 34, there is a concentrically
arranged combustion liner 38 that is connected at its aft end with
the inner wall 40 of the transition duct 18. The forward end of the
combustion liner is supported by a combustion liner cap assembly 42
secured to the combustor casing. It will be appreciated that the
outer wall 36 of the transition duct 18, as well as a portion of
flow sleeve 34 are formed with an array of apertures 44 over their
respective peripheral surfaces to permit air to reverse flow from
the compressor 12 through the apertures 44 into the annular
(radial) space between the flow sleeve 34 and the liner 36 toward
the upstream or forward end of the combustor (as indicated by the
flow arrows shown in FIG. 1).
[0024] At the forward end of the cap assembly 42, the air reverses
direction again, flowing through swirlers 46 surrounding each
nozzle, and into pre-mix tubes 48 that are also supported by the
liner cap assembly, as explained in greater detail in the '991
patent. Note that the nozzles extend into the pre-mix tubes.
[0025] With reference now to FIGS. 2-4, a new combustion liner cap
assembly 50 in accordance with this invention is illustrated. More
specifically, the invention relates to the incorporation of an
extended axial diffuser 52 in the combustion liner cap assembly 50.
The cap assembly 50 includes a radial flange 54 by which the cap
assembly is secured between forward and aft case components 56, 58,
utilizing bolts and locating pins in conventional fashion. The cap
assembly 50 includes a plurality of pre-mix burner tubes 60 as in
the prior construction, with an effusion plate 62 at the aft end
thereof. The premix burner tubes are themselves mounted in a
circular plate 64.
[0026] The axial diffuser 52 of the cap assembly 50 is comprised of
three distinct elements: (1) a conical outer body or component 66
that forms the outer radial surface of the extended cap axial
diffuser, (2) a cylindrical inner body or component 68 that forms
the inner radial surface of the extended cap axial diffuser; and
(3) an enlarged cylindrical portion 70 of the forward case 56 to
accommodate and house the cap diffuser. In this regard, note the
forward end of the outer body 66 engages the cylindrical portion 70
of the forward case 56.
[0027] Air flows in the annular flow passage 72 between the flow
sleeve 74 and combustion liner 76, and that radial space is
maintained between the diffuser inner body 68 and outer body 66 by
means of spacers or webs 78 (best seen in FIG. 4). Because the
cross-sectional area of the annular passage 72 between the inner
body 68 and outer body 66 increases in the direction of airflow in
the expanding region 80 of the flow passage, the velocity of the
air decreases as it passes the cap assembly and turns at the
forward end of the combustor into the inner body 68, and
subsequently into the premix burner tubes 60.
[0028] The novel features of this design are the deliberate
incorporation of an extended axial diffuser section 52 into the cap
assembly 50. The full axial diffuser allows considerable reduction
in air velocities that reduce pressure losses. In addition, the
diffuser inner cylinder is key to the diffuser concept and
contributes improved flameholding margin by making flow into each
premixer more uniform.
[0029] While the invention has been described in connection with
what is presently considered to be the most practical and preferred
embodiment, it is to be understood that the invention is not to be
limited to the disclosed embodiment, but on the contrary, is
intended to cover various modifications and equivalent arrangements
included within the spirit and scope of the appended claims.
* * * * *