U.S. patent number RE34,962 [Application Number 07/890,916] was granted by the patent office on 1995-06-13 for annular combustor with tangential cooling air injection.
This patent grant is currently assigned to Sundstrand Corporation. Invention is credited to John P. Archibald, Colin Rodgers, Jack R. Shekleton, Robert W. Smith.
United States Patent |
RE34,962 |
Shekleton , et al. |
June 13, 1995 |
Annular combustor with tangential cooling air injection
Abstract
The combustion dynamics and efficiency of gas turbine having an
annular combustor 26 provided with fuel injection nozzles 50 that
inject fuel generally tangentially is improved by providing the
walls 32, 34, 39 of the combustion 26 with cooling air film
injectors 70, 86; 72, 88; 74, 90 at substantially equally angularly
spaced locations about each such wall and which are oriented to
generally tangentially inject a film-like air stream on the
associated wall 32, 34, 39.
Inventors: |
Shekleton; Jack R. (San Diego,
CA), Rodgers; Colin (San Diego, CA), Archibald; John
P. (LaJolla, CA), Smith; Robert W. (Lakeside, CA) |
Assignee: |
Sundstrand Corporation
(Rockford, IL)
|
Family
ID: |
22481605 |
Appl.
No.: |
07/890,916 |
Filed: |
May 29, 1992 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
Reissue of: |
138342 |
Dec 28, 1987 |
04928479 |
May 29, 1990 |
|
|
Current U.S.
Class: |
60/804; 60/746;
60/755; 60/756; 60/760 |
Current CPC
Class: |
F23R
3/04 (20130101); F23R 3/28 (20130101); F05B
2220/50 (20130101); F05B 2250/322 (20130101) |
Current International
Class: |
F23R
3/28 (20060101); F23R 3/04 (20060101); F02C
003/08 () |
Field of
Search: |
;60/743,746,748,755,756,757,758,759,760,394.36,737,738,740 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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762596 |
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Nov 1956 |
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GB |
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1060095 |
|
Feb 1967 |
|
GB |
|
Primary Examiner: Thorpe; Timothy S.
Attorney, Agent or Firm: Wood, Phillips, VanSanten, Hoffman
& Ertel
Claims
We claim:
1. A gas turbine comprising:
a rotor including compressor blades and turbine blades;
an inlet adjacent one side of said compressor blades;
a diffuser adjacent the other side of said compressor blades;
a nozzle adjacent said turbine blades for directing hot gasses at
said turbine blades to cause rotation of said rotor;
an annular combustor having radially inner and outer walls
connected by a generally radially extending wall about said rotor
and having an outlet connected to said nozzle and a primary
combustion annulus defined by said walls remote from said outlet, a
plurality of fuel injectors to said primary combustion annulus and
being substantially equally .[.angular.]. .Iadd.angularly
.Iaddend.spaced therearound and configured to inject fuel into said
primary combustion annulus in a nominally tangential direction;
and
means .[.associated with.]. .Iadd.on .Iaddend.each of said walls
for injecting a film-like stream of cooling air into said primary
combustion annulus in a generally tangential direction.
2. The gas turbine of claim 1 wherein said .Iadd.cooling air
.Iaddend.injection means include cooling air openings in fluid
communication with said diffuser to receive compressed gas
therefrom.
3. The gas turbine of claim 2 wherein a compressed gas housing
surrounds said combustor in spaced relation thereto and is in fluid
communication with said diffuser, said cooling air openings
extending to the interface of said housing .[.an.]. .Iadd.and said
.Iaddend.combustor to receive compressed gas therefrom.
4. The gas turbine of claim 1 wherein said fuel injectors comprise
fuel nozzles having ends within said primary combustion annulus,
and atomizing nozzles for said combustion supporting gas
surrounding said ends.
5. A gas turbine comprising
a rotor including compressor blades and turbine blades;
an inlet adjacent one side of said compressor blades:
a diffuser adjacent the other side of said compressor blades:
a nozzle adjacent said turbine blades for directing hot gasses at
said turbine blades to cause rotation of said rotor, and
an annular combustor about said rotor having an outlet to said
nozzle, an inner wall and an outer wall spaced therefrom and a
connecting radial wall defining an annular combustion space, a
plurality of cooling air film injectors at substantially equally
angularly spaced locations about each said wall and oriented to
inject a film-like air stream on the associated wall generally
tangentially to said annular combustion space.
6. The gas turbine of claim 5 wherein each said cooling air film
injector comprises a row of openings in the associated wall and a
cooling strip having an edge overlying and spaced from said
row.
7. The gas turbine of claim 6 wherein said cooling strips each have
a cross section in the shape of a flatted "S".
8. The gas turbine of claim 6 wherein the rows and strips
associated with said inner and outer walls extend generally
axially.
9. The gas turbine of claim 6 wherein the rows and strips
associated with said radial wall extend generally radially.
.Iadd.10. A gas turbine comprising:
a rotary compressor;
a rotary turbine wheel mounted for rotation about an axis and
coupled to said compressor to drive the same;
a nozzle adjacent said turbine wheel for directing hot gases
thereat to rotate the same;
an annular combustor about said turbine wheel having radially inner
and outer walls, a radially extending wall and an outlet connected
to said nozzle and opposite said radially extending wall with a
primary combustion annulus defined by said walls remote from said
outlet: and
a plurality of circumferentially spaced fuel injectors adjacent
said radially outer wall and having ends disposed in said primary
combustion annulus so as to inject fuel thereinto in a direction
nominally tangential to said radially inner wall at radii passing
through said axis. .Iaddend. .Iadd.11. The gas turbine of claim 10
wherein said combustor further includes a plurality of
circumferentially spaced air injector tubes mounted on said
radially outer wall and oriented to direct air into said primary
combustion annulus in a direction nominally tangential to said
radially inner wall. .Iaddend. .Iadd.12. The gas turbine of claim
11 wherein said fuel injectors are located within at least some of
said air injection tubes. .Iaddend. .Iadd.13. A gas turbine
comprising:
a rotary compressor;
a rotary turbine wheel mounted for rotation about an axis and
coupled to said compressor to drive the same;
a nozzle adjacent said turbine wheel for directing hot gases
thereat to rotate the same;
an annular combustor about said turbine wheel having radially inner
and outer walls, a radially extending wall and an outlet connected
to said nozzle and opposite said radially extending wall with a
primary combustion annulus defined by said walls remote from said
outlet;
a plurality of circumferentially spaced fuel injectors for
injecting fuel into said primary combustion annulus in a nominally
tangential direction; and
a plurality of circumferentially spaced air injector tubes mounted
on said radially outer wall and oriented to direct air into said
primary combustion annulus in a direction nominally tangential to
said radially inner wall at radii passing through said axis.
.Iadd.14. The gas turbine of claim 13 wherein said fuel injectors
are located within at least some of said air injection tubes.
.Iadd.15. A gas turbine comprising:
a rotary compressor;
a rotary turbine wheel mounted for rotation about an axis and
coupled to said compressor to drive the same;
a nozzle adjacent said turbine wheel for directing hot gases
thereat to rotate the same about said axis;
an annular combustor about said turbine wheel having radially inner
and outer walls, a radially extending wall and an outlet connected
to said nozzle and opposite said radially extending wall with a
primary combustion annulus defined by said walls remote from said
outlet;
a plurality of circumferentially spaced fuel injectors adjacent
said radially outer wall and having ends disposed in said primary
combustion annulus so as to inject fuel thereinto in a direction
nominally tangential to said radially inner wall at radii passing
through said axis;
a plurality of circumferentially spaced air injector tubes mounted
on said radially outer wall and oriented to direct air into said
primary combustion annulus in a direction nominally tangential to
said radially inner wall at radii passing through said axis;
and
at least some of said air injection tubes being in surrounding
relation to one of said fuel injector so that air flowing through
said air injection tubes assists in atomizing fuel injected by said
injectors. .Iaddend.
Description
.Iadd.This application is a a reissue of Ser. No. 07/138,342 filed
Dec. 28, 1987, now U.S. Pat. No. 4,928,479. .Iaddend.
FIELD OF THE INVENTION
This invention relates to gas turbines, and more particularly, to
an improved combustor for use in gas turbines.
BACKGROUND OF THE INVENTION
It has long been known that achieving uniform circumferential
turbine inlet temperature distribution in gas turbines is highly
desirable. Uniform distribution minimizes hot spots and cold spots
to maximize efficiency of operation as well as prolongs the life of
those parts of the turbine exposed to hot gasses.
To achieve uniform turbine inlet temperature distribution in gas
turbines having annular combustors, one has had to provide a large
number of fuel injectors to assure that the fuel is uniformly
distributed in the combustion air. Fuel injectors are quite
expensive with the consequence that the use of a large number of
them is not economically satisfactory. Moreover, as the number of
fuel injectors increases in a system, with unchanged fuel
consumption, the flow area for fuel in each injector becomes
smaller. As the fuel flow passages become progressively smaller,
the injectors are more prone to clogging due to very small
contaminants in the fuel.
This in turn creates the very problem sought to be done away with
through the use of a number of fuel injectors. In particular, a
fouled fuel injector will result in a non uniform turbine inlet
temperature in an annular combustor with the result that hot and
cold spots occur.
To avoid this difficulty, the prior art has suggested that by and
large axial injection using a plurality of injectors be modified to
the extent that such injectors inject the fuel into the annular
combustion chamber with some sort of tangential component. The
resulting swirl of fuel and combustion supporting gas provides a
much more uniform mix of fuel with the air to provide a more
uniform burn and thus achieve more circumferential uniformity in
the turbine inlet temperature. However, this solution deals only
with minimizing the presence of hot and/or cold spots when one or
more injectors plug and does not deal with the desirability of
eliminating a number of fuel injectors to reduce cost and/or
avoiding the use of injectors having very small fuel flow passages
which are prone to clogging.
The present invention is directed to overcoming one or more of the
above problem.
SUMMARY OF THE INVENTION
It is the principal object of the invention to provide a new and
improved annular combustor for a gas turbine. More specifically, it
is an object of the invention to provide such a combustor wherein
the number of fuel injectors may be minimized and yet uniform
circumferential turbine inlet temperature distribution retained
along with a minimization the possibility of the fuel injectors
plugging.
An exemplary embodiment of the invention achieves the foregoing
objects in a gas turbine including a rotor having compressor blades
ana turbine blades. An inlet is located adjacent one side of the
compressor blades and a diffuser is located adjacent the other side
of the compressor blades. A nozzle is disposed adjacent the turbine
blades for directing hot gasses at the turbine blades to cause
rotation of the rotor and an annular combustor having spaced
radially inner and outer, axially extending wails connected by a
radially extending wall is disposed about the rotor and has an
outlet connected to the nozzle and a primary combustion annulus
remote from the outlet. A plurality of fuel injectors to the
primary combustion annulus are provided and are substantially
equally angular spaced about the same. They are configured to
inject fuel into the primary combustion annulus in a nominally
tangential direction. Cooling air for one or more of the walls of
the annular combustor is introduced tangentially in a film-like
fashion along the interior side or sides of one or more of the
combustor walls. The use of a tangentially flowing film of cooling
air serves to reduce the tendency of injected fuel from moving in
the axial direction allowing complete evaporation within the
primary combustion annulus to increase operational efficiency. In
addition, annular momentum of the air stream from the compressor is
conserved to reduce the overall pressure loss and again increase in
operational efficiency.
Injection of air for film cooling is accomplished through the use
of cooling air openings in one or more of the walls of the annular
combustor.
Where the air film injection is accomplished through the radially
inner and/or radially outer walls of the combustor, it is
preferably accomplished through the provision of a plurality of
axially extending rows of openings while cooling air film injection
through the radially extending wall of the combustor is
accomplished through the use of radially extending rows of
openings.
In either case, elongated cooling strips having a shape somewhat
akin to that of a flattened "S" are utilized. The cooling strips
have one edge secured to the corresponding wall of the annular
combustor and the opposite edge spaced therefrom. The opposite
edges overlie corresponding ones of the rows of cooling air
openings and in the case of the radially inner and outer walls are
axially directed and in the case of the radially extending wall are
generally radially directed. The opposite edges are downstream in
the direction of swirl within the annular combustor from the edges
that are attached to the respective walls. As a consequence, air
enter the combustor through the cooling air opening is directed by
the cooling strip in the tangential direction and in close
proximity to the associated wall to thereby generate the cooling
air firm.
According to a preferred embodiment, the cooling air openings are
in fluid communication with the diffuser to receive compressed air
therefrom.
In a highly preferred embodiment, the fuel injectors comprise fuel
nozzles having ends within the primary combustion annulus and air
atomizing nozzles for the combustion supporting air surround each
of the ends of the fuel injector fuel nozzles.
The invention contemplates the use of a compressed air housing
surrounding the combustor in spaced relation thereto and in fluid
communication with the diffuser. The cooling air openings open to
the interface of the housing and combustor to receive compressed
air therefrom.
Other objects and advantages will become apparent from the
following specification taken in connection with the accompanying
drawings.
DESCRIPTION OF THE DRAWINGS
FIG. 1 is a somewhat schematic, fragmentary, sectional view of a
turbine made according to the invention;
FIG. 2 is a fragmentary sectional view taken approximately along
the line 2--2 in FIG. 1; and
FIG. 3 is a fragmentary enlarged sectional view of a cooling strip
that may be used in the invention.
DESCRIPTION OF THE PREFERERD EMBODIMENT
An exemplary embodiment of a gas turbine made according to the
invention is illustrated in the drawings in the form of a radial
flow gas turbine. However, the invention is not so limited, having
applicability to any form of turbine or other fuel cornbusting
device requiring an annular combustor.
The turbine includes a rotary shaft 10 journalled by bearings not
shown. Adjacent one end of the shaft 10 is an inlet area 12. The
shaft 10 mounts a rotor, generally designated 14 which may be of
conventional construction. Accordingly, the same includes a
plurality of compressor blades 16 adjacent the inlet 12. A
compressor blade shroud 18 is provided in adjacency thereto and
just radially outwardly of the radially outer extremities of the
compressor blades 18 is a conventional diffuser 20.
Oppositely of the compressor blades 16, the rotor 14 has a
plurality of turbine blades 22. Just radially outwardly of the
turbine blades 22 is an annular nozzle 24 which is adapted to
receive hot gasses of combustion from a combustor, generally
designated 26. The compressor system including the blades 16,
shroud 18 and diffuser 20 delivers compressed air to the combustor
26, and via dilution air passages 27 and 28, to the nozzle 24 along
with the gasses of combustion. That is to say, hot gasses of
combustion from the combustor 26 are directed via the nozzle 24
against the blades 22 to cause rotation of the rotor 14 and thus
the shaft 10. The latter may be, of course, coupled to some sort of
apparatus requiring the performance of useful work.
A turbine blade shroud 29 is interfitted with the combustor 26 to
close off the flow path from the nozzle 24 and confine the
expanding gas to the area of the turbine blades 22.
The combustor 26 has a generally cylindrical inner wall 32 and a
generally cylindrical outer wall 34. The two are concentric and
merge to a necked down area 36 which serves as an outlet from the
interior annulus 38 of the combustor to the nozzle 24. A third wall
39, generally radially extending and concentric with the walls 32
and 34, interconnects the same to further define the annulus
38.
Oppositely of the outlet 36, and adjacent the wall 39, the interior
annulus 38 of the combustor 26 includes a primary combustion zone
40. By primary combustion zone, it is meant that this is the area
in which the burning of fuel primarily occurs. Other combustion
may, in some instances, occur downstream from the primary
combustion area 40 in the direction of the outlet 36. As mentioned
earlier, provision is made for the injection of dilution air
through the passageways 27 and 28 into the combustor 26 downstream
of the primary combustion zone 40 to cool the gasses of combustion
to a temperature suitable for application to the turbine blades 22
via the nozzle 24. It should be noted that the passageways 27 and
28 are configured so that the vast majority of dilution air flow
into the combustor 26 occurs through the passageways 28. This, of
course, requires the vast majority of dilution air to pass about
the generally radially outer wall 34, the third wall 39 and the
radially inner wall 32 which in turn provides excellent convective
cooling of these combustor walls and avoids the formation of hot
spots on any of the walls 32, 34 and 39.
In any event, it will be seen that the primary combustion zone 40
is an annulus or annular space defined by the generally radially
inner wall 32, the generally radially outer wall 34 and the wall
39.
A further wall 44 is generally concentric to the walls 32 and 34
and is located radially outwardly of the latter. The wall 44
extends to the outlet of the diffuser 20 and thus serves to contain
and direct compressed air from the compressor system to the
combustor 26.
As best seen in FIG. 2, the combustor 26 is provided with a
plurality of fuel injection nozzles 50. The fuel injection nozzles
50 have ends 52 disposed within the primary combustion zone 40 and
which are configured to be nominally tangential to the inner wall
32. The fuel injection nozzles 50 generally but not necessary
utilize the pressure drop of fuel across swirl generating orifices
(not shown) to accomplish fuel atomization. Tubes 54 surround the
nozzles 50. High velocity air from the compressor flows through the
tubes 54 to enhance fuel atomization. Thus the tubes 54 serve as
air injection tubes. When swirl generating orifices are not used as
in the embodiment illustrated, high velocity air flowing through
the tubes 34 is the means by which fuel exiting the nozzles 50 is
atomized.
The fuel injecting nozzles 50 are equally angularly spaced about
the primary combustion annulus 40 and optionally disposed between
each pair of adjacent nozzles 50 there may be a combustion
supporting air jet 56. When used, the jets 56 are located in the
wall 34 and establish fluid communication between the air delivery
annulus defined by the walls 34 and 44 and the primary combustion
annulus 441. These jets 56 may be somewhat colloquially termed
"bender" jets as will appear. They are also oriented so that the
combustion supporting air entering through them enters the primary
combustion annulus 40 in a direction nominally tangential to the
inner wall 32.
Preferably the injectors 50 and jets 56 are coplanar or in
relatively closely spaced planes remote from the outlet area 36.
Such plane or planes are transverse to the axis of the shaft
10.
When the intended use of the engine requires the delivery of large
quantities of bleed air, the wall 44 is provided with a series of
outlet openings 58 which in turn are surrounded by a bleed air
scroll 60 secured to the outer surface of the wall 44. Thus, bleed
air to be used for conventional purposes may be made available at
an outlet (not shown) from the scroll 60.
To prevent the formation of undesirable hot spots on the walls 32,
34 and 39 for any of a variety of reasons, the invention
contemplates the provision of means for flowing a cooling air film
over the walls 32, 34 and 39 on the surfaces thereof facing the
annulus 38. Further, the invention provides means whereby the
cooling air film is injected into the annulus 38 in a generally
tangential, as opposed to axial, direction.
Preferably, the injection is provided along each of the walls 32,
34 and 39 but in some instances, such injection may occur on less
than all of such walls as desired.
In the case of the radially inner wall 32, the same is provided
with a series of apertures 70. Preferably, the apertures 70 are
arranged in a series of equally angularly spaced, generally axially
extending rows. Thus, the three apertures 70 shown in FIG. 2
constitute one aperture in each of three such rows while the
apertures 70 illustrated in FIG. 1 constitute the apertures in a
single such row.
A similar senes of equally angularly spaced, axially extending rows
of apertures 72 is likewise provided in the wall 34.
Similarly, in the case of the wall 39, there are a series of
generally radially extending rows of apertures 74. As can be
readily appreciated, the apertures 70, 72 and 74 establish fluid
communication between the annulus defined by the wall 44 and the
wall 34, a radially extending annulus defined by the wall 39 and a
wall 80 connected to the wall 44, and the connecting annulus
defined by the wall 32 and a connecting wall 82.
Thus tangential and film-like streams of cooling air enter the
annulus 38 through the openings 70, 72 and and cooling strips 86,
88, and 90 are applied respectively to the walls 32, 34 and 39.
As a consequence of this construction, the air flowing in the
annuli about the combustor 26 will remove heat therefrom by
external convective cooling of the walls 32, 34 and 39. Similarly
the cooling air film on the sides of the walls 32, 34 and 39
fronting the annulus 38 resulting from film-like air flow into the
annulus 38 through the apertures 70, 72 and 74 minimizes the input
of heat from the flame within the combustor 26 to the walls 32, 34
and 39.
Thus, in the preferred embodiment, the entirety of the internal
surface of all of the walls, 32, 34 and 39 is completely covered
with a film of air. The ability to completely cool the internal
walls of a combustor is difficult to accomplish, particularly as
combustor size decreases. However, utilizing the novel technique of
tangential injection of air as herein disclosed readily
accomplishes the establishment of a complete wall covering film to
provide improved wall cooling. The film further serves to minimize
carbon build-up and the elimination of hot spots on the combustor
walls.
These advantages are enhanced by reason of the jets of air which
result from air flow through the apertures 70, 72 and 74. Such jets
of air impact upon the cooling strips to cool them. The cooling
strips 86, 88 and 90 are further cooled by the aforementioned film
of air flowing over them. The cooling strips also act as a local
barrier to convective and radiative heating of the walls 32, 34 and
39 by the flame burning within combustor 26.
The cooling strips 86, 88 and 90 are generally similar one to the
other and accordingly, it is believed that a complete understanding
of the operation of the same can be achieved simply from
understanding the operation of one. Thus, only the cooling strip 86
will be described.
With reference to FIG. 3, the cooling strip 86 is seen to be in the
shape of a generally flattened "S" having an upstream edge 92
bonded to the wall 32 just upstream of a corresponding row of the
openings 70 by any suitable means as brazing or, for example, a
weld 94. Because of the S shape of the cooling step 86, this
results in the opposite or downstream edge 96 being elevated above
the opening 70 with an exit opening 98 being present. The exit
opening 98 is elongated in the axial direction along with the edge
96 and also opens generally tangentially to the wall 32.
Consequently, air entering the annulus 38 through the openings in
the direction of arrows 100 (FIGS. 2 and 3) will flow in a
film-like fashion in a generally tangential direction along the
wall 32 on its interior surface to cool the same. The air flow
indicated by arrows 102 in FIG. 2 illustrate the corresponding
tangential, film-like flow of cooling air on the interior of the
wall 34 while additional arrows 104 in FIG. 2 illustrated a
similar, tangential film-like air flow of air entering the openings
74 in the wall 39.
Operation is generally as follows. Fuel emanating from each of the
nozzles 50 will enter along a line such as shown at "F". This line
will of course be straight and it will be expected that the fuel
will diverge from it somewhat. Assuming bender jets 56 are used, as
the fuel approaches the adjacent bender jet 56 in the clockwise
direction, the incoming air from the diffuser 20 and compressor
blades 16 will tend to deflect or bend the fuel stream to a
location more centrally of the primary combustion annulus 40 as
indicated by the curved line "S". There will, of course, be a
substantial generation of turbulence at this time and such
turbulence will promote uniformity of burn within the primary
combustion annulus 40 and this in turn will result in a uniform
circumferential turbine inlet temperature distribution at the
nozzle 24 and at radially outer ends of the turbine blades 22. Such
uniform turbine inlet temperature distribution is achieved in a
combustor made according to the invention utilizing many fewer fuel
injecting nozzles 50 than would be required according to prior art
teachings. As a result of the invention, and even without the use
of the bender jets 56, through the use of tangential fuel injection
and cooling film introduction, a combustor made according to the
invention will require about half the number of fuel injector
nozzles 50 as would a conventional combustor of equal volume. In
particular, the two will have approximately the same so-called
"pattern factor".
If the bender jets 56 are added without adding nozzles 50, an
improvement in pattern factor will be obtained over the
conventional combustor.
In any event, resulting elimination of a number of fuel injector
nozzles 50 provides a substantial cost savings. Moreover, in
engines having an increased combustor volume, a further substantial
reduction in the number of fuel injectors by as much as 80% of
those required according to conventional practice may be
obtained.
It will also be observed that where the number of fuel injections
nozzles 50 is halved using the principals of the invention, the
fuel flow passages of the remaining fuel injection nozzles,
assuming they are cylindrical, can be increased in diameter
slightly over 40%. This increase in diameter reduces the
possibility of plugging of the fuel injection nozzles 50 to provide
a more trouble free apparatus. This characteristic of the invention
assumes extreme importance in small engines which utilize small
combustors and thus have relatively small fuel flows, particularly
at low engine speeds or while starting at high altitudes.
In addition, the injection of cooling air in a film-like manner
achieved by means of the openings 70, 72 and 74 and associated
cooling strips 86, 88 and 90 further minimizes the possibility of a
hot spot on a wall coming into existence and thereby prolongs the
life of the apparatus. Significantly, the tangential injection of
the cooling air film in the same direction as the swirl within the
annulus 38 does not provide an axial impetus to fuel droplets
entering the primary combustion zone 40 from the nozzles 50. As a
consequence, there is ample time for such fuel to fully and
completely vaporize within the primary combustion zone 40 and
thereby achieve highly efficient combustion. For example, in one
combustor made according to the invention tested at 10% of rated
engine speed with a combustor pressure drop of only 0.8 inches of
water, a short efficient flame was obtained using No. 2 diesel
fuel. In contrast, a conventional annular combustor using
conventional swirl air blast inject on would typically be unable to
sustain combustion under similar circumstances. Thus, an engine
employing the invention is more easily started, a feature that may
be particularly critical when high altitude operation is used as,
for example, when the engine is used as part of an auxiliary power
unit or an emergency power unit. Because a high degree of
tangential motion or swirl is found desirable in a turbine made
accordingly to the invention, desire vanes such as those somewhat
schematically illustrated at 106 in FIG. 1 may be relatively
minimal thereby reducing the complexity of the invention. The swirl
that is thus permitted conserves the angular velocity of the
compressed air as it leaves the diffuser 20 so that the pressure
drop is minimized, thereby enhancing operational efficiency.
Furthermore, since the turbine nozzle 24 is desire, need to impart
swirl to the hot gases directed against the turbine blades 22, the
fact that the gases are already swirling as a result of tangential
air and fuel injection minimizes the directional change applied to
such gases by the nozzle 24 to provide a further increase in
efficiency.
At the same time, the use of minimal deswirl vanes 106 allows the
initial swirl that is typically imparted to the compressed air by
the compressor 16 and diffuser 20 to be retained outside the
combustor 26 allowing bleed air, which is commonly obtained from a
circumferential vent enclosed by a scroll, to be obtained with a
high degree of efficiency.
According to the invention, the combustor is sized by an equation
of the form: ##EQU1## Where K is a constant;
W.sub.a is the combustor air flow in pounds per second;
T.sub.3 is the turbine inlet temperature in degrees Rankine;
T.sub.2 is the combustor inlet temperature in degree Rankine;
.DELTA.P/P is the combustor pressure drop.times.100;
P is the combustor air inlet pressure in psia;
.DELTA.P is the combustor pressure drop in psia;
D is the mean combustor height in inches;
H is the mean combustor width in inches;
N is the number of fuel injectors; and
R is the engine pressure ratio.
The present invention provides a trade-off between combustor volume
and the number of injectors. It is a trade-off that cannot be
achieved in conventional combustors. Specifically, in a
conventional combustor, the number of injectors is determined
generally by the expression N=.pi.D/H.
If the number of injectors as defined by the preceding equation is
reduced, there is a senous increase in turbine inlet hot spots. In
one combustor made according to the invention, only four injectors
were required whereas normal practice would require about thirteen
such injectors. Further, in the combustor made according to the
invention, a pattern factor of 0.095 was obtained. The pattern
factor is a measure of the uniformity of temperature throughout the
combustion area and is defined by the formula ##EQU2## where
T.sub.h is a temperature of the hottest spot in degrees
Rankine.
In any event, the pattern factor of 0.095 obtained in a combustor
made according to the invention is twice as good as the pattern
factor that would be obtained in normal practice with thirteen
injectors.
Further, when one of the fuel injectors in the four injector
structure made according to the invention was plugged up to
simulate a typical field failure. the pattern factor increased only
to 0.11, a negligible increase. Converely, extensive experience in
turbine engines has indicated that if one injector plugs up in a
conventional combustor, the resulting hot spot will seriously
damage or even destroy the turbine engine.
Similarly, when a combustor employing two diametrically opposite
injectors with two intermediate bender jets was employed, a pattern
factor of 0.2 was obtained. This pattern factor its comparable to
that which would be obtained in a conventional combustor utilizing
13 injectors. The improvements in pattern factors along with the
ability to tolerate plugging as well as the elimination of a large
number of injectors clearly the demonstrates the superiority of the
invention.
In addition, in a combustor made according to the invention, a test
was run with fuel flowing only out of one injector of the four
provided. The injector from which fuel was flowing was the
lowermost one and the test was to simulate start-up of the engine
at very high altitudes when, due to so-called "manifold head"
effects, at low fuel flow rates. Substantially all fuel flows into
the combustor through the lowermost injector. The resulting time
visually observed spread about the entire combustor and the pattern
factor was a tolerable 0.33. Conversely, in a conventional
combustor wherein fuel is flowed only through one injector, a very
localized of lame with inefficient burning is observed and starting
at altitudes is poor.
Thus, in addition to the previously stated advantages, the
invention is ideally suited for use in turbine engines,
particularly small turbine engines, that may be operated at high
altitudes and require starting at such altitudes as well.
* * * * *