U.S. patent number 3,763,650 [Application Number 05/166,095] was granted by the patent office on 1973-10-09 for gas turbine temperature profiling structure.
Invention is credited to Serafino M. DeCorso, Charles E. Hussey.
United States Patent |
3,763,650 |
Hussey , et al. |
October 9, 1973 |
GAS TURBINE TEMPERATURE PROFILING STRUCTURE
Abstract
An axial flow gas turbine having a structure including a fuel
nozzle for providing any desired turbine inlet temperature profile
according to the mechanical stress on the rotating turbine blades.
The fuel nozzle may be a combined multiple fuel gas and liquid type
or a single fuel type, an important feature of both being the
provision of fuel injection asymmetrically into the combustion
chamber to establish a desired ignited fuel temperature pattern
therein which continues down to the turbine inlet. This
asymmetrical fuel supply into the combustion chamber is obtained by
providing a number or size of fuel jets according to the
temperature desired; for example, to provide a temperature gradient
decreasing in an inward turbine radial direction through the
combustor to correspond to the desired turbine blade inlet
temperature profile. The angular jet direction also may be varied
for different parts of the fuel injection pattern to obtain a
further spatial control of the fuel injection distribution into the
combustor. These fuel distribution control variables may be used
singly or in any combination, and the same or different ones may be
used for the gas and liquid jets in multiple fuel nozzles. In
addition, a further modification of the intake temperature profile
may be obtained by providing relatively cool jets of gaseous fluid
through spaced orifices in the radially inner sides of the
transition passages which direct the high temperature working gas
to the turbine blades.
Inventors: |
Hussey; Charles E. (Glenolden,
PA), DeCorso; Serafino M. (Media, PA) |
Family
ID: |
22601804 |
Appl.
No.: |
05/166,095 |
Filed: |
July 26, 1971 |
Current U.S.
Class: |
60/39.463;
60/737; 60/742; 239/561 |
Current CPC
Class: |
F23D
11/104 (20130101); F23R 3/28 (20130101); F23D
11/38 (20130101); F23R 2900/00005 (20130101) |
Current International
Class: |
F23R
3/28 (20060101); F02c 003/14 () |
Field of
Search: |
;60/39.74R,39.6G
;239/423,424,560,561 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Croyle; Carlton R.
Assistant Examiner: Garrett; Robert E.
Claims
The invention claimed is:
1. A high temperature gas supply system for an axial flow turbine
having turbine blades arranged to be driven by the high temperature
gases and having a combustor with means for igniting fuel supplied
thereto, including a fuel supply nozzle comprising a body having a
member defining a plurality of fuel outlet spray passages each
having a terminal orifice for injecting fuel into the combustor,
said terminal orifices being so proportioned that the orifice
disposed at the greatest distance from the axis of turbine is
larger than the orifice disposed closest to the axis of the turbine
and including means for defining a fuel injection pattern for the
nozzle of fuel injected into the combustor from said nozzle
orifices to provide a substantially predetermined asymmetrical
spatial fuel injection pattern transversely of the combustor and a
consequent temperature profile of ignited fuel in the
combustor.
2. A fuel nozzle for a turbine combustor adapted to supply hot
gases to blades of a gas turbine, comprising a body having a member
with a spray face and defining a plurality of angularly outwardly
extending fuel outlet spray passages each having a terminal orifice
in said spray face for injecting fuel into the combustor, said
terminal orifices being so proportioned that the orifice disposed
at the greatest distance from the axis of the turbine is larger
than the orifice disposed closest to the axis of the turbine and
said member defining said orifices arranged relative to each other
so as to provide a substantially predetermined asymmetrical fuel
injection pattern and a consequent temperature profile of ignited
fuel transversely of the combustor.
3. A fuel nozzle as defined in claim 2 wherein said member defines
said orifices substantially uniformly circumferentially angularly
spaced and of a larger size for all orifices in substantially the
outer semicircle of the spray face corresponding to substantially
the outer half of turbine blades fed by the hot gases from said
combustor and of a smaller size for orifices in the complementary
inner semicircle of the spray face.
4. A fuel nozzle as defined in claim 2 wherein said member defines
said orifices in a progressively graduated size from the smallest
at the innermost position of the nozzle orifices relative to the
radially inner ends of the turbine blades to the largest orifice at
the outermost position of the nozzle orifices relative to said
inner ends of the turbine blades.
5. A fuel nozzle as defined in claim 2 wherein said member defines
said spray passages at predetermined different outwardly extending
angles relative to the longitudinal center line of the nozzle to
provide a predetermined larger fuel injection into the combustor
toward the outer side thereof relative to the radially inner ends
of the turbine blades and a graduated lesser fuel injection toward
the opposite side of the combustor.
6. A fuel nozzle as defined in claim 2 wherein said member defines
said orifices to provide a substantially larger fuel injection by
orifices in the outer semicircle of the spray face relative to the
radially inner ends of the turbine blades and a predetermined
lesser fuel injection by orifices in the complementary semicircle
of the spray face.
7. A fuel nozzle as defined in claim 2 wherein said member defines
said orifices to provide a predetermined graduated increase in fuel
injection into the combustor from the side thereof nearest to the
longitudinal axis of the turbine toward the opposite side
thereof.
8. A fuel nozzle as defined in claim 4 wherein said member defines
said orifices substantially uniformly circumferentially angularly
spaced at a predetermined angle in the outer semicircle of the
spray face relative to the longitudinal axis of the turbine and
substantially uniformly circumferentially angularly spaced at a
larger angle than said predetermined angle in the complementary
semicircle of the spray face.
9. A fuel nozzle as defined in claim 5 wherein said member defines
said orifices substantially uniformly circumferentially angularly
spaced and of a larger substantially uniform size for all orifices
in substantially the outer semicircle of the spray face
corresponding to substantially the radially outer half of turbine
blades fed by hot gases from said combustor and of a smaller
substantially uniform size for orifices in the complementary inner
semicircle of the nozzle spray face.
10. A fuel nozzle as defined in claim 5 wherein said orifices are
arranged in a progressively graduated size from the smallest at the
innermost position of the nozzle orifices relative to the
longitudinal axis of the turbine to the largest at the outermost
position of the nozzle orifices to said turbine axis.
11. A fuel nozzle as defined in claim 5 wherein said member defines
said orifices substantially uniformly circumferentially angularly
spaced at a predetermined angle in the outer semicircle of the
spray face relative to the longitudinal axis of the turbine and
substantially circumferentially angularly spaced at a larger angle
than said predetermined angle in the complementary inner semicircle
of the spray face.
12. A fuel nozzle as defined in claim 5 wherein said spray passage
angles are a substantially uniform predetermined value relative to
the said nozzle center line in the outer semicircle of the spray
face relative to the longitudinal axis of the turbine and are a
substantially uniform lesser value than said predetermined value in
the complementary inner semicircle of the spray face.
13. A fuel nozzle as defined in claim 5 wherein said spray passage
angles are of progressively graduated values from the smallest
angle at the innermost passage relative to the longitudinal axis of
the turbine to the largest angle at the outermost passage in the
nozzle relative to said turbine axis.
14. A fuel nozzle as defined in claim 2 having a second body
surrounding and radially spaced from at least a part of the
longitudinal sides of said first mentioned member and defining a
cylindrical air flow passage therebetween, means providing for the
supply of pressurized air to said cylindrical passage, and means
for directing pressurized air from said cylindrical passage around
fuel injected into the combustor from said spray passage
orifices.
15. A fuel nozzle as defined in claim 14 wherein said second body
includes means defining a plurality of fuel gas outlet passages
each having an orifice for injecting pressurized fuel gas into the
combustor around fuel injected thereinto from said spray passage
orifices, and means for supplying fuel gas to said fuel gas outlet
passages.
16. A fuel nozzle as defined in claim 15 wherein said fuel gas
orifices defining means defines said orifices relative to each
other to provide a substantially predetermined asymmetrical fuel
gas injection pattern and consequent temperature profile of ignited
fuel in the combustor.
17. A turbine high temperature gas supply system as defined in
claim 1 having a plurality of combustors arranged in an annular
array and having a plurality of transition structures for
transmitting the high temperature gases from the combustors to the
turbine blades, said transition structures defining a plurality of
circumferentially spaced relatively cool pressurized fluid supply
passages for directing relatively cool fluid into the high
temperature gases to moderate the temperature profile thereof.
18. A turbine high temperature gas supply system as defined in
claim 20 having means for directing relatively cool fluid into said
transition structures' cool fluid supply directing passages for
progressively decreasing the temperature of the high temperature
gas in an inward direction radially of the turbine.
Description
BACKGROUND OF THE INVENTION
1. Field of the invention
The present invention relates to improvements in turbine operating
with high temperature fluids, and particularly to an arrangement
for profiling the temperature of the fluid in accordance with the
mechanical stress occurring on rotating blades of the turbine.
2. Prior Art
It is known that higher initial operating temperatures in a
turbine, such as an axial flow gas turbine, will provide higher
thermal efficiency and specific power output. It is also known that
the allowable stress to which the blades can be subjected for a
given blade life decreases with increasing temperatures. Thus the
main limiting factor in raising gas turbine operating temperatures,
and thereby raising turbine efficiency and power output, is the
physical capability of the rotating blades, the blades being highly
stressed during turbine operation.
Rotating turbine blades are usually made individually and attached
to the rim of a turbine wheel so that they extend radially
outwardly. On rotation, the blades are subjected to a tensile
stress in a radial direction due to the centrifugal force thereon
which is a function of the weight of the blade. This stress is
greatest near the root or hub portion of the blade where it is
attached to the turbine wheel, and the stress decreases outwardly
to the tip of the blade, reaching zero at the blade tip.
In most prior gas turbine designs, the temperature of the working
gas flow is generally uniform along the radial height of the
turbine blades. Since the root portion of the blade is stressed the
greatest, and since an essentially uniform gas temperature flow
sets the allowable stress for a particular blade and blade
material, the root portion stress at a specified temperature is
generally taken as the reference point in fixing the temperature at
which the turbine blades are designed to operate for an efficient
use of the energy from the gas flow. This is so, even though the
outer radial portion of the blades can effectively handle higher
temperatures because of the lower stresses thereon.
This prior approach in designing gas turbines is inherently
wasteful of the potential capabilities of the outer, lesser
stressed portions of the rotating blades to withstand higher
temperatures and thereby increase the efficiency and power output
of the turbine. U.S. Pat. No. 3,490,747-- DeCorso and Carlson,
illustrates one type of improved arrangement for improving the
intake gas temperature profile for increasing the efficiency of
this type turbine.
SUMMARY OF THE INVENTION
According to the present invention, a gas turbine inlet temperature
profile may be obtained having any desired shape, the particular
shape corresponding to the mechanical stress occurring in a
particular set of rotating turbine blades. This is attained by
providing a fuel nozzle which supplies the heat producing materials
to the combustor according to a predetermined distribution pattern
which approximates the asymmetrical temperature profile to be
supplied to the turbine blade intake. Various factors are disclosed
for obtaining the desired combustor fuel intake pattern; such as, a
predetermined angular liquid fuel spray orifice spacing
arrangement, a predetermined spray orifice size arrangement, a
predetermined relative spray orifice supply passage angular
direction for forming the spray pattern, combination of these
features, use of a multiple fuel supply including any of the
foregoing liquid fuel spray features combined with any fuel gas
supply utilizing any of the liquid spray features to control the
fuel gas pattern to provide the desired temperature profile, and
use of asymmetrical nozzle combustor sweep air supply for
efficiently burning the fuel to provide the desired profile. The
final turbine blade intake temperature profile may be further
modified to obtain the desired profile by directing a relatively
cool or lower temperature gaseous fluid, such as air, through
suitably spaced orifices in the radially inner sides of transition
passage members which direct the high temperature working gas to
the turbine blades from the combustor. The cooler gas moderates the
working gas temperature further to assure the flow of the cooler
gas past the radially inner root portions of the rotating blades
where the stresses are the highest and the flow of the highest
temperature gas over the radially outer tips of the rotating
blades, where the centrifugal blade stresses are the lowest. This
provides for the most efficient use of the blade structure by
permitting the operation of the rotating blades at a more uniform
stress and at the maximum practical operable stress throughout the
blade rather than at a stress limited solely by the blade root
stress. This latter feature of cool gaseous fluid supply to the
transition passage to the turbine blade intake can utilize cool gas
jets in connection with the illustrated orifices, such as the jet
structures disclosed in U.S. Pat. No. 3,490,747-- DeCorso and
Carlson, where the refinements of such jets may be of additional
use in providing the desired intake temperature profile.
BRIEF DESCRIPTION OF THE FIGURES OF THE DRAWINGS
In the drawings:
FIG. 1 is a longitudinal sectional view of the upper half of a gas
turbine provided with a combustion apparatus incorporating the
present invention;
FIG. 2 is a longitudinal sectional view of part of the gas turbine
combustion chamber shown in FIG. 1, including a fuel nozzle
incorporating an embodiment of the present invention;
FIG. 3 is an end view of the liquid fuel spray forming portion of a
nozzle such as that shown in FIG. 2;
FIG. 4 is an end view similar to FIG. 2, illustrating another
embodiment of the end of the liquid fuel spray forming portion
thereof;
FIG. 5 is an end view similar to FIGS. 3 and 4, illustrating a
further embodiment of the fluid spray forming portion of the
nozzle;
FIG. 6 is a fragmentary axial sectional view of yet another
embodiment of the nozzle, illustrating variations in the fuel and
air nozzle orifice supply passages; and
FIG. 7 is a longitudinal sectional view of a part of an axial flow
gas turbine showing an embodiment of a further temperature
profiling means according to this invention for use with nozzles of
the types shown in FIGS. 2-6.
DETAILED DESCRIPTION OF THE INVENTION
Referring to the drawings, the upstream end portion of a combustion
chamber 10 of a gas turbine is shown in longitudinal section in
FIG. 2, for a turbine of the type shown in section in FIG. 1,
provided with a multiple fuel nozzle 11 incorporating one
embodiment of an improved nozzle according to the present
invention. The illustrated gas turbine is of the axial flow type
having the usual multi-stage axial flow compressor 12 for supplying
air under pressure to the combustion apparatus through a plenum
chamber 13, partially defined by wall structure 13', FIGS. 1 and 7,
of a turbine housing in which a plurality of combustion chambers or
combustors 10 are arranged in conventional angularly spaced
relation around the longitudinal axis of the turbine, such as is
illustrated in U.S. Pat. No. 3,169,367-- Hussey. The hot gases from
the combustion chambers 10 are fed to the turbine through a
plurality of circumferentially spaced transition members 14, FIG.
7, of the type disclosed in U.S. Pat. No. 3,490,747--DeCorso et
al., each combustion chamber having its own transition member. The
transition member has an end portion 15 axially spaced from and
leading to an annular group of circumferentially spaced stationary
blades 16 mounted on a stationary turbine housing 17. Immediately
downstream of the stationary blades 16 are a corresponding number
of rotor blades 18 suitably mounted in an annular array on the
periphery of a rotor wheel 19, alternately longitudinally spaced
between sets of the stationary blades.
The transition members 14 direct the flow of hot gases from the
combustion chambers 10 to the turbine blades 16 and 18. As
previously explained, the rotating blades 18 are subjected to
tensile stress which is greatest at the roots, where they are
attached to the turbine wheel 19, and decreases outwardly to the
tips of the blades. Since the allowable safe operating stress on
the rotating blades 18 for efficient blade life, for blades of a
particular material and design, varies according to the turbine
power output, which varies according to the initial gas temperature
to the blades, the most efficient use of the blades would be
attained by varying the initial input gas temperature profile
inversely to the centrifugal force stress profile thereof.
According to this invention, a close approximation to the desired
hot gas temperature profile can be attained in the combustion
chamber by controlling the amount and direction of the combustion
materials supplied into the combustor. This combustor hot gas
temperature profile may be represented by the curve t-c-d-e,
superimposed on the conbustor shown in FIG. 2, illustrating the
average diametrical temperature across the combustor, wherein the
abscissas, reading from t as zero on the curve, represent the
average temperatures across the combustor on a diameter thereof
extending outwardly from any point t in the combustor along a
longitudinal line joining the inner peripheral points thereof
nearest to the longitudinal axis of the turbine, and the ordinates
t-e represent the points along the diameter as shown. The burning
and burned materials then pass longitudinally from the fuel supply
end of the combustor to the transition member 14 with this
substantially desired hot gas temperature profile. A final
refinement of this profile can be made in passing the gases through
the transition into the intake to the gas turbine blade section to
provide a gas temperature blade intake profile corresponding to the
desired profile O"-a'-b', superimposed on the structure shown in
FIG. 7, as will be explained later.
In the FIG. 2 embodiment of this invention, the illustrated
combustor is of the canister type and comprises a plurality of
cylindrical combustion basket or liner members 20 of graduated
cross-sectional area, disposed in slightly overlapping relation and
forming a primary combustion zone 21. Each basket or liner member
20 has an array of circumferentially spaced apertures or primary
air holes 22 for admitting primary combustion air from the plenum
chamber 13 into the combustion zone 21 to support combustion of
fuel injected thereinto by a fuel supply system comprising the
nozzle 11 mounted on a cover plate 23 of the upstream part 13' of
the main turbine housing 17. The illustrated fuel supply nozzle 11
is of the multiple fuel type and includes a central liquid fuel
spray body member formed by a central axially extending supply
passage body 24 defining a liquid fuel passage 25 therein, to which
liquid fuel is supplied under pressure from a suitable source, not
shown, by a conduit 26. The amount and pressure of the liquid fuel
supplied to the nozzle may be controlled in any suitable manner, as
by a control valve 27 in the conduit 26. Liquid fuel passes through
the supply passage 25 into an annular manifold distribution chamber
28 in a manifold unit 29, fitted in fluid-tight engagement with the
inner end face of the body 24 and held rigidly in position by a
nozzle orifice spray cap 30, tightly secured to the body 24 in any
suitable manner, as by tightly drawn up threaded engagement
therewith. A plurality of circumferentially angularly spaced fuel
distribution passages 31 extend through the manifold unit 29 from
the distribution chamber 28 to the opposite inner end face of the
unit. Each manifold passage 31 preferably has an enlarged mouth 32
communicating with a fuel outlet passage 33 in the spray cap 30 by
way of a passage 34.
In order to atomize the liquid fuel spray into minute droplets to
facilitate ignition thereof, steam or air under pressure from a
suitable source, not shown, is supplied through a conduit 36 into
an atomizing fluid supply passage 37 extending longitudinally
through the nozzle body 24. This passage 37 communicates, at its
inner end, with a distribution passage 38 extending through the
manifold unit 29. The flow and pressure of the atomizing fluid may
be controlled in any suitable manner, as by a valve 39 in the
conduit 36, and may be varied according to the rate of supply of
the liquid fuel to assure the most efficient atomization thereof.
The atomizing fluid is sprayed out of the nozzle through the
terminal spray orifices 35 of the spray passages 34 in which it is
mixed with the liquid fuel from the passages 33.
In order to obtain the desired temperature profile of the hot
gaseous products of combustion of the burned fuel in the combustor,
the fuel nozzle features can be varied so as to provide desired
fuel distribution patterns which will produce this profile. One of
these variables is the fuel supply spray passages and the terminal
orifices in the spray cap 30. As shown in FIGS. 2 and 4, the spray
passages 34 and orifices 35 may be of uniform size and the passages
34 may extend angularly at the same angle with reference to the
center line of the nozzle, FIG. 2, so that these factors will not
vary the spray pattern or the temperature profile in this
embodiment. As shown in FIG. 4, the number and spacing of spray
passages and orifices may be defined by the spray cap 30 to provide
the desired fuel spray pattern. For example, the spray passages and
orifices 35a in the radially outer semicircle or 180.degree. of arc
of the nozzle spray face relative to the longitudinal axis of the
turbine, corresponding substantially to the radially outer half of
the turbine blades 18, are substantially uniformly spaced at a
predetermined angle circumferentially, illustrated at 30.degree.,
but any other suitable angle could be used. The spray passages and
orifices 35b in the inner semicircle or 180.degree. of arc of the
nozzle face complementary to the set of orifices 35a also are
defined by the spray cap 30 substantially uniformly angularly
spaced, but at a larger angle than the spacing of the orifices 35a,
and are shown spaced at 45.degree., although any other suitable
angle could be used. This definition of the spray passages and
orifices provides a predetermined larger fuel injection into the
combustor toward the outer side thereof relative to the
longitudinal axis of the turbine and a lesser fuel injection toward
the opposite inner side thereof. The provision of a suitable supply
of air and ignition of the fuel thus supplied will provide the
desired temperature profile.
Primary combustion air is supplied into the combustion chamber 10
through the primary air holes 22. An additional air supply is
provided in this nozzle through passages 40 in the cylindrical
support 41 of a combustion baffle 42, and passages 43, aligned with
passages 40 and extending through a fuel gas nozzle body 44, from
which it passes axially of the nozzle through a cylindrical chamber
45 between the liquid fuel nozzle body member 24 and the fuel gas
nozzle body 44, and is blown into the combustion chamber
peripherally around the complete circumference of the spray cap 30.
This air further aids in providing an intimate combustible mixture
of the atomized liquid fuel.
In a multiple fuel type nozzle, as illustrated, the fuel gas is
supplied under pressure from a suitable source, not shown, through
a conduit 46 connected to an inlet passage 47 defined by the fuel
gas nozzle body 44. Fuel gas passes from the passage 47 into a
manifold chamber 48, from which it is fed by passages 49 to a
distribution chamber 50 and out through terminal orifices 51
defined by the face portion 52 of the nozzle body 44. The flow and
pressure of the fuel gas can be controlled in any suitable manner,
as by a valve 53 in the supply conduit 46, so that the most
desirable and efficient proportion of fuel gas to liquid fuel can
be attained. All fuel gas orifices 51 may extend as passages
through the face 52 at substantially equal angles relative to the
center line of the nozzle and be of substantially the same size, as
shown in FIG. 2. Also, these orifices may be equally angularly
spaced circumferentially or they may be spaced in the same angular
arrangement as the orifices 35a and 35b, shown in FIG. 4, in order
to accentuate the larger fuel injection into the outer semicircle
or 180.degree. of arc of the combustor relative to the longitudinal
axis of the turbine and the lesser fuel injection into the opposite
side thereof. Improved mixture of the fuel and air is provided by
supplying combustion air admitted by passages 40 and through a
cylindrical chamber 54, extending between the radially spaced walls
of the fuel gas nozzle body 44 and the combustion baffle supporting
wall 41, and blown into the combustion chamber peripheraly around
the complete circumference of the fuel gas nozzle body face 52.
Additional combustion sweep air is introduced under pressure from
the plenum chamber 13 through a plurality of angularly spaced
passages 55 through the combustion baffle 42 and peripherally
around the outwardly flared sides of the combustion baffle and
inner side of the combustion dome 56 of the combustion chamber.
In starting up combustion, the fuel is ignited in any conventional
manner, as by a suitable igniter or spark plug 57, suitably mounted
with its spark gap in the path of the injected fuel and air in the
combustion chamber.
In order further to refine the temperature profile of the hot gases
as they pass from the transition member 14 to the turbine blades,
to jets of relatively cool gas are directed into the discharge end
portion 15 of the transition member. The turbine air compressor 12
discharges relatively cool air under high pressure into the plenum
chamber and partly into an axially extending passage 58, FIG. 7,
between the combustor 10 and transition member 14 and the
compressor drive shaft coupling 59 to the turbine rotor wheel 19.
As shown in FIG. 7, the inner side of the end portion 15 of the
transition member defines a plurality of circumferentially spaced
orifices 60 through which the cool pressurized air is blown
outwardly into the path of the hot gas discharge. These jets of
relatively cool air modify the temperature profile from that of
t-c-d-e, FIG. 2, to that of O"-a'-b', FIG. 7. These cool air jets
may, in some instances, be further modified in accordance with the
teaching of U.S. Pat. No. 3,490,747--DeCorso and Carlson.
In accordance with this invention, the desired combustor
temperature profile t-c-d-e, FIG. 2, can also be obtained by
defining the fuel inlet orifices of the nozzle substantially
uniformly angularly spaced, as shown in FIG.3, and of a larger
substantially uniform size for all orifices 61 in substantially the
outer semicircle or 180.degree. of arc of the nozzle spray face
corresponding to the radially outer half of the turbine blades, and
of a smaller substantially uniform size for orifices 61" in the
complementary inner semicircle or 180.degree. of arc of the nozzle
spray face. These two sets of different size orifices can be used
for the liquid fuel spray orifices combined with two similar sets
of correspondingly arranged different size orifices for the fuel
gas, or the fuel gas orifices may be of size and spacing types
similar to FIGS. 4 or 5, or even uniform in size and spacing. In
some instances, it may be found that the latter can be used
advantageously to provide a large part of the heat represented by
t-c and e-d in the temperature profile, FIG. 2. All of these nozzle
orifice combinations provide for a substantially predetermined
larger fuel injection toward the side of the combustor
corresponding to the radially outer ends of the turbine blades and
lesser fuel injection toward the opposite side, corresponding to
the inner ends or roots of the blades.
Another embodiment of an improved nozzle according to this
invention is illustrated by FIG. 5. In this nozzle, the main body
defines the fuel inlet orifices in a progressively graduated size
from the smallest orifice 62 at the innermost position of the
nozzle orifices relative to the longitudinal axis of the turbine to
the largest orifice 62' at the outermost position of the orifices
relative to the longitudinal axis of the turbine. In this type of
structure, the fuel injection pattern is controlled primarily by
the relative size arrangement of the orifices, and it may be used
for both the liquid fuel nozzle orifices and the fuel gas orifices,
or the latter may be either the FIG. 3 or FIG. 4 types or of
uniform size.
A yet further embodiment of an improved nozzle according to this
invention is illustrated in FIG. 6 in which parts corresponding to
those of FIGS. 1 and 2 BEAR the same reference numbers. The
relative size arrangement bear be any of the other types. In
addition, the fuel injection angle is modified for the different
orifices by the provision of fuel passages leading to these
orifices which extend at predetermined different angles relative to
the center line of the nozzle to provide a predetermined larger
fuel injection into the combustor toward the outer side thereof
relative to the longitudinal axis of the turbine and a lesser fuel
injection toward the opposite inner side thereof. As shown in FIG.
6, the angle .alpha. of the outer spray passage 64 is larger than
the angle .alpha.' of the inner spray passage 64'. These two
passages could respectively correspond to those leading to the
orifices 62' and 62 of FIG. 5, or 61 and 61' of FIG. 3.
Furthermore, the angle .alpha.' may be progressively enlarged from
its innermost orifice to the outermost orifice angle .alpha.. The
same relative change in the angle .beta. for the passage 65 for the
outermost fuel gas orifice to the smaller angle .gamma.' for the
innermost fuel gas orifice passage 65' may be used. As shown, the
variation in the angles for the liquid fuel passages need not be
the same as for the fuel gas passages.
In all of the illustrated nozzles, the air flow through and around
the nozzle may be of the type shown in FIG. 2, and the pressurized
air or steam flow for atomizing the liquid fuel spray from the
nozzle preferably also is of the type explained with reference to
FIG. 2. Also, the refinement of the temperature profile as
explained with reference to FIG. 7 to obtain a profile O"-a'-b',
preferably is used with all types of nozzles disclosed, and may be
of the types disclosed in U.S. Pat. No. 3,490,747--DeCorso and
Carlson.
While particular embodiments of this invention have been
illustrated and described, modifications thereof will occur to
those skilled in the art. It is to be understood, therefore, that
the invention is not to be limited to the exact details
disclosed.
* * * * *