U.S. patent number 3,706,203 [Application Number 05/085,629] was granted by the patent office on 1972-12-19 for wall structure for a gas turbine engine.
This patent grant is currently assigned to United Aircraft Corporation. Invention is credited to Perry Goldberg, Irwin Segalman.
United States Patent |
3,706,203 |
Goldberg , et al. |
December 19, 1972 |
WALL STRUCTURE FOR A GAS TURBINE ENGINE
Abstract
A combustion wall or liner construction which utilizes a unique
geometry, primarily a plurality of flow channels with predetermined
length-to-diameter ratio which are positioned between the walls of
the liner, to maintain the liner wall at an acceptable operating
temperature. Cooling air flows through the flow channels and
control of the frictional pressure losses of this stream provide
the control of the wall temperature.
Inventors: |
Goldberg; Perry (West Hartford,
CT), Segalman; Irwin (Bloomfield, CT) |
Assignee: |
United Aircraft Corporation
(East Hartford, CT)
|
Family
ID: |
22192885 |
Appl.
No.: |
05/085,629 |
Filed: |
October 30, 1970 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
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812793 |
Apr 2, 1969 |
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Current U.S.
Class: |
60/757;
431/353 |
Current CPC
Class: |
F23R
3/08 (20130101); Y02T 50/675 (20130101); Y02T
50/60 (20130101) |
Current International
Class: |
F23R
3/04 (20060101); F23R 3/08 (20060101); F02c
007/18 () |
Field of
Search: |
;60/39.65,39.66
;431/353 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Hart; Douglas
Parent Case Text
This is a continuation-in-part of Ser. No. 812,793, filed Apr. 2,
1969, now abandoned.
Claims
We claim:
1. A liner for a combustion chamber for use in a gas turbine engine
in which combustion occurs in a gas stream moving axially through
the combustion chamber, said liner including a plurality of liner
segments arranged circumferentially, each segment having spaced
inner and outer wall elements and means interconnecting the
elements and defining closely spaced longitudinally extending flow
passages therebetween, the outer wall of one segment being
connected to and overlying the inner wall of the next adjacent
downstream segment to cause cooling air to flow into the upstream
ends of the passages from the chamber space outside the liner and
to enter the space inside the liner at the downstream ends of said
passages, successive liner segments increasing in length in a
downstream direction as the operating temperature within the liner
decreases.
2. An annular combustion chamber for a gas turbine engine in which
combustion occurs in a stream of gas moving axially through said
chamber, said chamber having inner and outer walls forming an
annular space therebetween and a liner spaced from one of the said
walls to define between said one wall and the liner a passage for
cooling air and on the other side of the liner a combustion space,
said liner comprising a plurality of segments extending in a
circumferential direction within the chamber and located in
side-by-side relation axially of the chamber and secured together,
each segment having spaced inner and outer wall elements and spacer
means between said elements serving to hold said elements in spaced
relation and forming closely spaced parallel passages between said
wall elements extending in an axial direction, the inner wall
element of one liner segment overlying and secured to the outer
wall element of the adjacent upstream element and said successive
segments increasing in length toward the downstream end of the
combustion chamber for increasing the lengths of the flow passages
near the areas of decreasing temperature within the combustion
space, each of said passages being relatively small in effective
diameter on the order of between 0.02 and 0.06 inches.
3. An annular combustion chamber as in claim 2 in which the
length-to-diameter ratio is between about 20 and 85.
4. A combustion chamber as in claim 2 in which the liner is in the
form of a burner can with the segments extending peripherally
around the can and forming the wall thereof.
Description
BACKGROUND OF THE INVENTION
The present invention relates to a heat resistant wall
construction, and more particularly to a wall construction which
has particular utility in the hot temperature environment of a gas
turbine engine.
In most gas turbine applications the combustion chambers, either
main burner or afterburner, utilize a cooled combustion chamber
liner or wall construction to limit the temperature of the main
load carrying structures of the engine such as the inner and outer
walls of the burner section of the engine. In the main burner
application this liner also serves to control the fuel-air
distribution in the primary combustion zone and to permit
controlled mixing of unburned air with the combustion products to
achieve desired turbine inlet temperature profiles. The cooling for
the liner or wall construction, which actually defines a combustion
zone within which the combustion or burning occurs, is intended to
limit the temperature of the liner wall itself and thereby the
temperature of the surrounding main load carrying structural
members.
Conventional combustion chamber liners have a number of drawbacks.
For example, these constructions require large quantities of
cooling air hence reducing the air available for combustion and/or
dilution in the main burners. This limits the combustor performance
in terms of temperature rise capability and quality of exit
temperature profile. It can also result in longer combustion
chambers or increased combustor pressure loss to achieve the
desired turbine inlet temperature profiles with the reduced
dilution air. These obviously result in weight and performance
penalties for the gas turbine engine.
SUMMARY OF THE INVENTION
The present invention provides a burner construction which avoids
the principal problems encountered by the prior art in a
construction in which the cooling air requirements are reduced by
approximately 50 percent. Additionally, the construction provides a
liner or a wall structure which is substantially free from severe
thermal gradients, therefore, improving the life characteristics of
the burner.
The description and discussion heretofore has been directed and
limited to a construction which has particular utility in the hot
environs of a gas turbine engine; however, it should be noted that
the present invention has utility in many other environments, e.g.,
anywhere a double wall structure confining a fluid between the two
walls is utilized.
It is a primary objective to provide a wall construction which has
particular utility in the combustion section of a gas turbine
engine, the construction being such as to maintain the wall and
surrounding structural members at an acceptable operating
temperature without subjecting these members to adverse thermal
gradients.
The burner or combustion chamber includes a liner or wall
construction which comprises a pair of radially spaced walls. The
radial space is actually a flow passageway which extends over the
entire length of a liner wall segment. The flow passageway in turn
comprises a plurality of relatively small diameter flow channels,
the flow channels being formed by either openings or through the
use of a corrugated strip member. The walls separating the
individual cooling passages increase the surface area on the
coolant side and thereby enhance the ability of this wall
construction to maintain itself at an acceptable wall temperature
with a given amount of cooling. In the preferred burner embodiment,
a plurality of double wall segments are joined to one another to
form sets of flow channels extending over the entire length of each
of the wall segments, with the length of each segment controlled
with respect to the temperature in the adjacent portion of the
combustion area. Each of the flow channels has a relatively large
length-to-diameter ratio.
The inner wall of the liner actually confines or is the boundary
member for the hot combustion gases, and it is this wall in the
liner construction which has to be maintained at an acceptance
operating temperature. To do this a cooling stream, with a
temperature relatively cooler than the combustion gases, is
introduced into the flow channels. As a result of the unique
geometry of the flow channels, i.e., the large length-to-diameter
ratio, the frictional pressure losses of the cooling stream in each
of the flow channels can be established and controlled, and thus
the operating temperature of the inner wall can be maintained at an
acceptable level. Additionally this is achieved with an amount of
cooling flow which is approximately 50 percent less than other
burner construction. Another feature of the present invention is
that since the flow channels extend over the entire length, and
since the control of the cooling flow can be accurately controlled,
the temperature of the inner wall is maintained at a uniform level.
Therefore the adverse thermal gradient and stress loads are
substantially eliminated.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a partial cross-sectional view of the burner apparatus,
the burner apparatus including a liner wall means utilizing the
construction of the present invention.
FIG. 2 is a sectional view taken substantially along line 2--2.
FIG. 3 is a sectional view similar to FIG. 2 showing a
modification.
FIG. 4 is a plot showing diagrammatically the relation of burner
temperature, liner temperature, and length of wall segments.
DESCRIPTION OF THE PREFERRED EMBODIMENT
The invention is shown in a burner can or liner for a combustion
chamber which is located between the compressor and turbine of a
gas turbine power plant and in which fuel is burned in the high
pressure gas discharged from the compressor to provide a hot gas
under pressure for expansion through the turbine. A power plant to
which this type of combustion chamber is applicable is disclosed
for example in the Savin U.S. Pat. No. 2,747,367.
As best shown in FIG. 1 the combustion chamber is of a can-annular
type only one can being shown. It is understood that any type of
combustion chamber may be employed whether it be a can-annular type
or an annular type. The can-annular combustion chamber has an inner
cylindrical wall 16 and an outer wall 17 both attached at their
upstream ends to a diffuser 18. Within the diffuser is mounted a
fuel nozzle 19 for each burner can 20 or liner. The liner or can
includes a dome-shaped head 22 at the inlet end with an opening
therein to receive the fuel nozzle. Combustion occurs within this
burner can or liner and the products thereof discharge from the
open downstream end 24 into the turbine.
High pressure air from the compressor is discharged into and
through the diffuser 18 and flows into the combustion chamber. A
portion of this air enters the liner or can 20 through a swirler 26
around the fuel nozzle and also through a plurality of combustion
holes 28 which are positioned within the liner wall. This liner is
made up of a plurality of rings or segments 20a, 20b, 20c, 20d, and
20e which are attached to one another and through the walls of
which cooling air flows for the purpose of maintaining the wall
temperature within the desired limits. It is essential that the
liner and the surrounding structural members of the engine be
maintained within acceptable operating temperature ranges and to
this extent the liner 20 has a cooled wall of novel
construction.
Referring to FIG. 2 which shows a sectional view through one
embodiment of the liner wall, this wall is double thickness and has
an inner wall element 40 and radially spaced therefrom, an outer
wall element 42. The inner wall element 40 is directly exposed to
the combustion gases within the liner or can and because of the
temperature of these gases it becomes necessary to cool the wall to
a suitable operating temperature to prevent damage to this wall
during operation of the engine. Cooling flow passages 44 are
provided between the wall 40 and 42 by circumferentially spaced,
longitudinally extending ribs 46 integral with and projecting
outwardly from the inner wall and into contact with the outer wall,
the spaces between the ribs defining the flow passages.
The burner can segments 20a, 20 b, 20c, 20d, and 20 e are each of
the double thickness described and with the downstream end 52 of
each segment attached to and positioned within the upstream end 54
of the next adjacent liner segment. The arrangement of these
segments is such that the cooling air enters the upstream end of
the passages 44 from outside of the liner, flows through the
passages and is discharged into the space inside of the liner in a
direction substantially parallel to the liner and directly within
the liner. This arrangement is shown in FIG. 1 in which each of the
liner sections tapers slightly from the upstream end to the
down-stream end as shown and the purpose of this is so that the
successive liner sections will telescope one with respect with
another. As shown, the inner wall element 40 of the first liner
section extends over the edge of the dome 22 for the burner can and
the downstream end of the outer wall element 42 of this liner
section fits within the upstream end of the inner wall element 40
of the next liner section. Also as shown in FIG. 1, the inner wall
element 40 extends forwardly somewhat beyond the passages within
the liner so that this wall is exposed to permit welding or other
attachment to the underlying end of the outer wall of the adjacent
liner segment. Thus, it will be apparent that adjacent liner
sections may be suitably attached one to another by welding the
overlapping and contacting inner wall of one segment with the outer
wall of the adjacent segment.
As shown in FIG. 1, the liner segments vary in length depending
upon the temperature of the combustion gases within the burner can
in order that the wall temperature of the liner segments may remain
substantially constant and at no time reach a point above that at
which the burner can may operate successfully without damage to the
material of the liner wall. Thus, the first liner segment is made
only 1 inch in length. The second segment is one and three eighths
of an inch, the third segment is 11/2 and the successive segments
are longer since at this time the gas temperature within the burner
can is diminishing as shown and less cooling is neccessary to
accomplish the desired result of maintaining a workable temperature
for the liner wall.
In the modified wall structure of FIG. 3, the wall of the burner
can consists of an inner wall element 60, an outer parallel wall
element 62 radially spaced therefrom, and a corrugated sheet 64
positioned therebetween and bonded to both to hold them in spaced
relation and to define a plurality of longitudinally extending
circumferentially spaced passages 66. The effective diameter of
these passages for the particular burner can shown, which has a
diameter of approximately 6.5 inches was 0.04 inches. These double
thickness wall elements are made up in the same manner as shown in
FIG. 1 with the outer wall extending beyond the passages at the
downstream end and with the inner wall extending beyond the
passages at the upstream end to provide overlapping flanges by
which successive segments may be secured together. With either
construction, FIG. 2 or FIG. 3, the additional coolant side surface
area as a consequence of the dividers between individual coolant
channels, significantly enhances the thermal effectiveness of the
wall construction and thereby reduces the cooling air
requirement.
As described above, each segment may have a small taper from end to
end so that when the several segments are secured together, the
resulting can will have a substantially constant diameter from end
to end.
Referring now to FIG. 4, this plot represents a burner can that was
built and tested and shows the temperature variation within the can
and the temperatures at which the successive sections of the liner
operated. As shown, by selecting the proper length of burner can
segments to produce the desired amount of cooling, it was possible
to keep the wall temperature from going above 1,600.degree.F, where
the temperature within the burner can was extremely high. The
particular wall construction of the burner can used was the
arrangement of FIG. 3 above described, and was of such dimension
that the effective diameter of each of the cooling passages was
0.04 inch, so that the length-to-diameter ratio for the first
section was 25. It is thus apparent that by selecting the relative
lengths of the successive liner segments, the greatest amount of
cooling accomplished was in those portions of the liner that were
subjected to the greatest heat from the combustion gases within the
liner and therefore the liner wall may be kept at appropriate
temperatures regardless of the temperature within the liner.
Furthermore, with the successive burner can segments lengthening as
the temperature within the can decreases the amount of cooling air
needed is materially reduced, thereby minimizing the amount of
cooling air required. By making the passages very small, on the
order, as above indicated of from 0.02 inch to 0.06 inch in
effective diameter, the coolant side heat transfer is significantly
enhanced. The smaller the area of the channel, the greater the
surface area of the channel that is exposed to the air passing
through the channel, and thereby, the greater cooling effect. The
ability to cool the structure with a minimum of cooling air is
enhanced by significantly increasing the heat transfer coefficients
and a significant increase in coolant side surface area. The liner
or burner can segments extend circumferentially but obviously the
flow passages are longitudinally or axially of the combustion
chamber, and the length of each segment determines the length of
the flow passages in that segment.
As shown in FIG. 4, the liner segments increase in length toward
the downstream end of the liner or burner can and thus the length
of segments increases as the operating temperature within the can
decreases. The essential feature is to provide only enough cooling
of the wall to keep the temperature of the wall from exceeding the
safe operating temperature, normally established for the particular
alloy used as 1,600.degree.F in the particular arrangement shown.
This has been accomplished in the construction shown as evidenced
by the chart of FIG. 4.
* * * * *